CN114519232A - Arrow body motion equation coefficient calculation method and system - Google Patents
Arrow body motion equation coefficient calculation method and systemDownload PDFInfo
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- CN114519232A CN114519232ACN202011310223.XACN202011310223ACN114519232ACN 114519232 ACN114519232 ACN 114519232ACN 202011310223 ACN202011310223 ACN 202011310223ACN 114519232 ACN114519232 ACN 114519232A
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Abstract
Description
Technical Field
The invention relates to the technical field of aerospace engineering, in particular to a method and a system for calculating arrow body motion equation coefficients.
Background
The carrier rocket is a non-uniform continuous elastomer with infinite multiple degrees of freedom, has variable rigidity and variable mass distribution along a longitudinal axis, and generates deformation and bending vibration, but mainly transverse bending vibration of a slender arrow body under the action of external load. The sensitive element of the control system will sense the additional signal caused by bending deformation, and the change of control force and moment is caused by the automatic stabilizing device, and in severe cases, the additional signal will affect the effective work of the control system, even lead to the failure of flight. In order to ensure the safety of the aircraft, the elastic rocket body motion equation coefficient needs to be obtained through calculation in the process of developing the carrier rocket, so that the control stability index in the flight process can be evaluated.
The conventional elastic rocket body motion equation coefficient calculation method needs a great deal of effort of designers to carry out detailed design and extraction processing on various data such as structural data, missile way data, pneumatic data, finite element models and the like of a carrier rocket, and the design method is low in efficiency, high in cost and long in period due to the fact that the data types are large, the middle quantity is large, and comprehensive factors in all aspects are caused.
Disclosure of Invention
Aiming at the defects in the prior art, the invention aims to provide an arrow body motion equation coefficient calculation method and system, which can solve the problems of low efficiency, high cost and long period of a design method caused by the fact that designers need to spend a great deal of energy on detailed processing of various data such as structural data, missile way data, pneumatic data, finite element models and the like of a carrier rocket in the prior art.
In order to achieve the above purposes, the technical scheme adopted by the invention is as follows:
the invention provides a method for calculating arrow body motion equation coefficients, which comprises the following steps:
loading design trajectory initial parameters, design pneumatic initial parameters and a parameterized arrow finite element model under various flight control modes of the arrow; inputting control parameters and structural parameters of an arrow body;
determining all characteristic time points according to the control parameters;
determining ballistic operating parameters of each characteristic time point according to the designed ballistic initial parameters and each characteristic time point; determining pneumatic operation parameters according to the designed pneumatic initial parameters and ballistic operation parameters; determining a real-time arrow body finite element model of each characteristic time point according to the parameterized arrow body finite element model and the trajectory operation parameters; calling finite element software, and simulating and determining kinetic parameters through the finite element software according to the real-time arrow body finite element model of each characteristic time point;
and determining each parameter of the elastic rocket body motion equation according to the ballistic operation parameter, the pneumatic operation parameter, the kinetic parameter and the structural parameter of the carrier rocket.
In some alternative embodiments, the various flight control modes include: one or more of attitude control engine, gas rudder, air rudder and swinging jet pipe;
the initial design trajectory parameters comprise time, rocket body mass, mass center, relative speed, flight Mach number, dynamic pressure, attack angle, trajectory inclination angle, active machine effective thrust, atmospheric density, pitching rudder deflection angle and attitude control engine effective thrust corresponding to different flight control modes;
the design of the pneumatic initial parameters correspondingly comprises the following steps according to different flight control modes: a plurality of combinations of a plurality of items in a sectional missile body normal force gradient, a sectional missile body normal force coefficient to attack angle derivative, a sectional missile body lateral force coefficient to slip angle derivative, an air rudder normal force coefficient to rudder deflection angle derivative, a sectional missile body characteristic area, a gas rudder lift gradient and an air rudder lift gradient;
the control parameters comprise flight stages of the carrier rocket, ignition time and fuel exhaustion time of an engine in the flight process corresponding to the flight stages, and time intervals and vibration mode stages of each characteristic time point.
In some optional embodiments, the determining the trajectory operating parameter of each characteristic time point according to the design trajectory initial parameter and each characteristic time point under various control modes specifically includes:
performing interpolation calculation on the designed ballistic initial parameters of the corresponding flight control modes according to each characteristic time point, and determining the ballistic operating parameters of each characteristic time point, wherein the method comprises the following steps: time, rocket mass, centroid, relative velocity, flight mach number, dynamic pressure, angle of attack, ballistic inclination, active machine effective thrust, atmospheric density, pitch rudder deflection angle, and attitude control engine effective thrust.
In some optional embodiments, the determining the aerodynamic operation parameter according to the designed aerodynamic initial parameter and the designed ballistic operation parameter specifically includes:
determining the normal force gradient of the segmented projectile body and/or the lift gradient of the air rudder at each moment according to the Mach number, the attack angle and the rudder deflection angle in the ballistic operation data of each characteristic time point;
and performing interpolation calculation according to the normal force gradient of the segmented projectile body and/or the lift gradient of the air rudder to the designed pneumatic initial dynamic parameters of the corresponding control mode, and determining the pneumatic operation parameters of each characteristic time point, wherein the parameters comprise the normal force gradient of the segmented projectile body, the derivative of the normal force coefficient of the segmented projectile body to the attack angle, the derivative of the lateral force coefficient of the segmented projectile body to the slip angle, the derivative of the normal force coefficient of the air rudder to the rudder deflection angle, the characteristic area of the segmented projectile body, the lift gradient of the gas rudder and/or the lift gradient of the air rudder.
In some optional embodiments, the determining a real-time arrow finite element model of each feature time point according to the parameterized arrow finite element model and the trajectory operating parameter, and simulating to determine the kinetic parameter specifically includes:
calculating a parameterized arrow finite element model according to the arrow mass and the mass center of each characteristic time point, and determining a real-time arrow finite element model of each characteristic time point;
and determining the dynamic parameters of each characteristic moment point according to the real-time arrow body finite element model and the vibration mode order, wherein the dynamic parameters comprise frequency, vibration mode and section quality data of the arrow body.
In some optional embodiments, the parameters of the elastic arrow motion equation comprise the circular frequency of each characteristic time point, the mode shape and slope at the inertial set, and the influence coefficient of the generalized aerodynamic force of the equivalent rigid arrow body on each order mode vibration in proportion to the rotation angular velocity.
In some optional embodiments, the structural parameters include: the distance from the installation base surface of the inertial unit to the actual sharp point of the whole rocket, the distance from the action point of the engine spray pipe to the actual sharp point of the whole rocket and the deviation of the thrust action line of the engine.
On the other hand, the invention also provides a system of the method for calculating the arrow body motion equation coefficients, which comprises the following steps:
the data input module is used for loading design trajectory initial parameters, design pneumatic initial parameters and a parameterized arrow body finite element model under various flight control modes of the arrow body; inputting control parameters and structural parameters of an arrow body;
the data processing module is used for determining all characteristic time points according to the control parameters; determining ballistic operating parameters of each characteristic time point according to the designed ballistic initial parameters and each characteristic time point; determining pneumatic operation parameters according to the designed pneumatic initial parameters and ballistic operation parameters; determining a real-time arrow body finite element model of each characteristic time point according to the parameterized arrow body finite element model and the trajectory operation parameters; calling finite element software, and simulating and determining dynamic parameters through the finite element software according to the real-time arrow body finite element model of each characteristic time point; and determining each parameter of the elastic rocket body motion equation according to the ballistic operation parameter, the pneumatic operation parameter, the kinetic parameter and the structural parameter of the carrier rocket.
In some optional embodiments, the data input module comprises:
a data loading unit for loading design trajectory initial parameters, design aerodynamic initial parameters and a parameterized rocket body finite element model under various flight control modes of the carrier rocket;
-a data input unit for control parameters and structural parameters of the arrow body;
the data processing module comprises:
-a characteristic moment processing unit for control parameters and structural parameters of the arrow body;
-a ballistic parameter processing unit for determining ballistic operational parameters for each characteristic time point from the design ballistic initial parameters and each characteristic time point;
-a pneumatic parameter processing unit for determining pneumatic operational parameters from design pneumatic initial parameters and ballistic operational parameters;
a model processing unit for determining a real-time arrow finite element model for each characteristic time instant point based on the parameterized arrow finite element model and the ballistic operational parameters,
-a calling unit for calling finite element software for determining kinetic parameters by means of finite element software simulation based on the real-time arrow finite element model of each characteristic time instant;
-a coefficient solving unit for determining the parameters of the elastic rocket body equations of motion from ballistic, aerodynamic, kinetic and structural parameters.
In some optional embodiments, the data input module is further configured to receive a command from the data processing module.
Compared with the prior art, the invention has the advantages that: determining all characteristic time points according to control parameters of an arrow body; determining ballistic operating parameters of each characteristic time point according to the designed ballistic initial parameters and each characteristic time point; determining pneumatic operation parameters according to the designed pneumatic initial parameters and ballistic operation parameters; determining real-time rocket body finite element models of each characteristic time point according to the parameterized rocket body finite element models and the trajectory operation parameters, and simulating and determining kinetic parameters by combining the vibration mode order; and determining each parameter of the elastic rocket body motion equation according to the ballistic operation parameter, the pneumatic operation parameter, the kinetic parameter and the structural parameter of the carrier rocket. The system only needs to load design trajectory initial parameters, design pneumatic initial parameters and a parameterized arrow body finite element model under various flight control modes of the arrow body through a data input module, and input control parameters and structural parameters of the arrow body; the data processing module carries out automation, flow and integration on the calculation process of the elastic rocket body motion equation coefficient by calling finite element software and processing data in the background, so that the elastic rocket body motion equation coefficient of the carrier rocket is simply, efficiently and quickly calculated. The invention can accelerate the development progress, improve the design quality and reduce the development cost.
Drawings
In order to more clearly illustrate the technical solutions in the embodiments of the present application, the drawings needed to be used in the description of the embodiments are briefly introduced below, and it is obvious that the drawings in the following description are only some embodiments of the present application, and it is obvious for those skilled in the art to obtain other drawings based on these drawings without creative efforts.
FIG. 1 is a flowchart of a method for calculating coefficients of an arrow body equation of motion according to an embodiment of the present invention.
Detailed Description
In order to make the objects, technical solutions and advantages of the embodiments of the present application clearer, the technical solutions in the embodiments of the present application will be clearly and completely described below with reference to the drawings in the embodiments of the present application, and it is obvious that the described embodiments are some embodiments of the present application, but not all embodiments. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present application.
Embodiments of the present invention are described in further detail below with reference to the accompanying drawings. As shown in fig. 1:
the invention provides a method for calculating arrow body motion equation coefficients, which comprises the following steps:
s1: loading design trajectory initial parameters, design pneumatic initial parameters and a parameterized arrow finite element model under various flight control modes of the arrow; and inputting control parameters and structural parameters of the arrow body.
In some embodiments, the various flight control modes include: and one or more of attitude control engines, gas rudders, air rudders and swinging spray pipes.
The adopted flight control mode can be one or the combination of a plurality of control modes of an attitude control engine, a gas rudder, an air rudder and a swinging spray pipe.
The initial design trajectory parameters comprise time, rocket body mass, mass center, relative speed, flight Mach number, dynamic pressure, attack angle, trajectory inclination angle, active machine effective thrust, atmospheric density, pitching rudder deflection angle and attitude control engine effective thrust corresponding to different flight control modes.
In the embodiment, when the air rudder control is adopted in the flight control mode, the initial design parameters of the pitching rudder deflection angle are only included in the initial design parameters of the trajectory; when the attitude control engine is controlled in a flight control mode, the initial design parameters of the effective thrust of the attitude control engine are designed in the initial parameters of the trajectory. The design of the pneumatic initial parameters correspondingly comprises the following steps according to different flight control modes: the system comprises a plurality of combinations of a plurality of items in a sectional missile body normal force gradient, a sectional missile body normal force coefficient to attack angle derivative, a sectional missile body lateral force coefficient to slip angle derivative, an air rudder normal force coefficient to rudder deflection angle derivative, a sectional missile body characteristic area, a gas rudder lift gradient and an air rudder lift gradient.
In the embodiment, when the air rudder control is adopted in the flight control mode, the air rudder normal force coefficient, the rudder deflection angle derivative and the air rudder lift gradient are only designed in the aerodynamic initial parameters; when the flight control mode is adopted and the gas vane control is adopted, the lift gradient of the gas vane is only in the designed pneumatic initial parameters.
The control parameters comprise flight stages of the carrier rocket, the flight stages correspond to the ignition time of an engine, the fuel exhaustion time and the time intervals of all characteristic time points in the flight process.
S2: and determining all characteristic time points according to the control parameters.
In this embodiment, according to the flight control requirement of the launch vehicle, the number of flight stages of the launch vehicle to be calculated, and the time intervals of the engine ignition time, the fuel exhaustion time, and the feature time points to be calculated during the flight of the rocket body of the launch vehicle under the flight technology are selected, and the data processing module determines the feature time points to be calculated of the equation of motion coefficients in the background.
S3: determining ballistic operating parameters of each characteristic time point according to the designed ballistic initial parameters and each characteristic time point; determining pneumatic operation parameters according to the designed pneumatic initial parameters and ballistic operation parameters; determining a real-time arrow body finite element model of each characteristic time point according to the parameterized arrow body finite element model and the trajectory operation parameters; and (4) calling finite element software, and simulating and determining the kinetic parameters through the finite element software according to the real-time arrow body finite element model of each characteristic time point.
S31: determining ballistic operating parameters of each characteristic time point, specifically comprising:
performing interpolation calculation on the designed ballistic initial parameters of the corresponding flight control modes according to each characteristic time point, and determining the ballistic operating parameters of each characteristic time point, wherein the method comprises the following steps: time, rocket mass, centroid, relative speed, flight mach number, dynamic pressure, attack angle, trajectory inclination angle, active machine effective thrust, atmospheric density, pitch rudder deflection angle (determined according to a flight control mode, namely adopted air rudder control) and attitude control engine effective thrust (determined according to a flight control mode, namely adopted attitude control engine control).
S32: determining the pneumatic operation parameters specifically comprises the following steps:
firstly, determining the normal force gradient of the segmented projectile and/or the lift gradient of the air rudder at each moment according to the Mach number, the attack angle and the rudder deflection angle in the trajectory operation data of each characteristic moment, wherein the lift gradient of the air rudder is related in a flight control mode adopting the air rudder.
And then, carrying out interpolation calculation according to the design pneumatic initial dynamic parameters of the sectional missile body normal force gradient and/or the air rudder lift force gradient to the corresponding control mode, and determining the pneumatic operation parameters of each characteristic time point, wherein the parameters comprise the sectional missile body normal force gradient, the sectional missile body normal force coefficient to attack angle derivative, the sectional missile body lateral force coefficient to side slip angle derivative, the air rudder normal force coefficient to rudder deflection angle derivative, the sectional missile body characteristic area, the gas rudder lift force gradient (only used under the gas rudder control mode) and/or the air rudder lift force gradient (only used under the air rudder control mode).
S33: simulating and determining kinetic parameters, specifically comprising:
and calculating the parameterized arrow finite element model according to the mass and the mass center of the arrow in the trajectory operation parameters of each characteristic time point, and determining the real-time arrow finite element model of each characteristic time point.
And calling finite element software, and determining the kinetic parameters of each characteristic time point, including frequency, vibration mode and section quality data of the arrow body, according to the real-time arrow body finite element model and the input vibration mode order of each characteristic time point to be calculated as simulation calculation control parameters.
In this embodiment, the parameterized arrow finite element model in step S1 is based on parameterized modeling, and the mass parameters of each component of the engine are parameterized, and the mass parameters of each component of the engine can be adaptively adjusted according to the change of the mass and the center of mass of the engine. Engine mass and centroid calculation: and calculating the mass and mass center parameters of the main engine in each characteristic time point by taking the time in each characteristic time point in the designed ballistic initial parameters as input.
The mass and the mass center of the rocket body in the trajectory operation parameters of each characteristic time point are used as input, the fuel consumption of the engine is mainly considered, the parameterized engine mass and mass center parameters in the finite element model are calculated by the data processing module, the mass parameters of each component of the engine at each characteristic time are obtained by calculation, and the real-time rocket body finite element model at each characteristic time is obtained by the automatic output of the data processing module at the background. The finite element software includes commercial software such as ANSYS, NASTRAN, etc.
S4: and determining each parameter of the elastic rocket body motion equation according to the ballistic operation parameter, the pneumatic operation parameter, the kinetic parameter and the structural parameter of the carrier rocket.
Preferably, the parameters of the elastic arrow motion equation comprise the circular frequency of each characteristic time point, the mode shape and the slope at the inertial set, and the influence coefficient of the generalized aerodynamic force of the equivalent rigid arrow body on the vibration of each mode shape, wherein the generalized aerodynamic force is proportional to the rotation angular velocity.
Preferably, the structural parameters include: the distance from the installation base surface of the inertial unit to the actual sharp point of the whole rocket, the distance from the action point of the engine spray pipe to the actual sharp point of the whole rocket and the deviation of the thrust action line of the engine.
The invention also provides an arrow body motion equation coefficient calculation system, which comprises:
the data input module is used for loading design trajectory initial parameters, design pneumatic initial parameters and a parameterized arrow body finite element model under various flight control modes of the arrow body; inputting control parameters and structural parameters of an arrow body;
the data processing module is used for determining all characteristic time points according to the control parameters; determining ballistic operating parameters of each characteristic time point according to the designed ballistic initial parameters and each characteristic time point; determining pneumatic operation parameters according to the designed pneumatic initial parameters and ballistic operation parameters; determining a real-time arrow body finite element model of each characteristic time point according to the parameterized arrow body finite element model and the trajectory operation parameters; calling finite element software, and simulating and determining kinetic parameters through the finite element software according to the real-time arrow body finite element model of each characteristic time point; and determining various parameters of the elastic rocket body motion equation according to the ballistic operation parameters, the pneumatic operation parameters, the kinetic parameters and the structural parameters of the carrier rocket.
Preferably, the data input module includes: the data loading unit is used for loading design trajectory initial parameters, design pneumatic initial parameters and a parameterized rocket body finite element model under various flight control modes of the carrier rocket; the device also comprises a data input unit, a data output unit and a control unit, wherein the data input unit is used for controlling parameters and structural parameters of the arrow body;
the data processing module comprises: a characteristic time processing unit for controlling parameters and structural parameters of the arrow body; the system also comprises a ballistic parameter processing unit, a ballistic parameter processing unit and a ballistic parameter processing unit, wherein the ballistic parameter processing unit is used for determining ballistic operation parameters of each characteristic time point according to the designed ballistic initial parameters and each characteristic time point; the pneumatic parameter processing unit is used for determining pneumatic operation parameters according to the designed pneumatic initial parameters and the designed ballistic operation parameters; the system also comprises a model processing unit used for determining the real-time arrow body finite element model of each characteristic time point according to the parameterized arrow body finite element model and the trajectory operation parameters, and a calling unit used for calling finite element software and determining the dynamic parameters through the finite element software simulation according to the real-time arrow body finite element model of each characteristic time point; the system also comprises a coefficient solving unit which is used for determining each parameter of the elastic rocket body motion equation according to the ballistic operation parameter, the pneumatic operation parameter, the dynamic parameter and the structural parameter.
The system also comprises a storage module which is used for storing the data of the data input module and the data processing module.
In summary, the method determines all feature time points according to the control parameters of the arrow body; determining ballistic operating parameters of each characteristic time point according to the designed ballistic initial parameters and each characteristic time point; determining pneumatic operation parameters according to the designed pneumatic initial parameters and ballistic operation parameters; determining real-time rocket body finite element models of each characteristic time point according to the parameterized rocket body finite element models and the trajectory operation parameters, and simulating and determining kinetic parameters by combining the vibration mode order; and determining each parameter of the elastic rocket body motion equation according to the ballistic operation parameter, the pneumatic operation parameter, the kinetic parameter and the structural parameter of the carrier rocket.
The system only needs to load design trajectory initial parameters, design pneumatic initial parameters and a parameterized arrow body finite element model under various flight control modes of the arrow body through a data input module, and input control parameters and structural parameters of the arrow body; the data processing module carries out automation, flow and integration on the calculation process of the elastic rocket body motion equation coefficient by calling finite element software and processing data in the background, so that the elastic rocket body motion equation coefficient of the carrier rocket is simply, efficiently and quickly calculated. The invention can accelerate the development progress, improve the design quality and reduce the development cost.
In the description of the present application, it should be noted that the terms "upper", "lower", and the like indicate orientations or positional relationships based on the orientations or positional relationships shown in the drawings, which are only for convenience in describing the present application and simplifying the description, and do not indicate or imply that the referred device or element must have a specific orientation, be constructed in a specific orientation, and operate, and thus, should not be construed as limiting the present application. Unless expressly stated or limited otherwise, the terms "mounted," "connected," and "connected" are intended to be inclusive and mean, for example, that they may be fixedly connected, detachably connected, or integrally connected; can be mechanically or electrically connected; they may be connected directly or indirectly through intervening media, or they may be interconnected between two elements. The specific meaning of the above terms in the present application can be understood by those of ordinary skill in the art as appropriate.
It is noted that, in the present application, relational terms such as "first" and "second", and the like, are used solely to distinguish one entity or action from another entity or action without necessarily requiring or implying any actual such relationship or order between such entities or actions. Also, the terms "comprises," "comprising," or any other variation thereof, are intended to cover a non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements does not include only those elements but may include other elements not expressly listed or inherent to such process, method, article, or apparatus. Without further limitation, an element defined by the phrase "comprising an … …" does not exclude the presence of other identical elements in a process, method, article, or apparatus that comprises the element.
The previous description is only an example of the present application, and is provided to enable any person skilled in the art to understand or implement the present application. Various modifications to these embodiments will be readily apparent to those skilled in the art, and the generic principles defined herein may be applied to other embodiments without departing from the spirit or scope of the application. Thus, the present application is not intended to be limited to the embodiments shown herein but is to be accorded the widest scope consistent with the principles and novel features disclosed herein.
Claims (10)
1. A method for calculating arrow body motion equation coefficients is characterized by comprising the following steps:
loading design trajectory initial parameters, design pneumatic initial parameters and a parameterized arrow finite element model under various flight control modes of the arrow; inputting control parameters and structural parameters of an arrow body;
determining all characteristic time points according to the control parameters;
determining ballistic operating parameters of each characteristic time point according to the designed ballistic initial parameters and each characteristic time point; determining pneumatic operation parameters according to the designed pneumatic initial parameters and ballistic operation parameters; determining a real-time arrow body finite element model of each characteristic time point according to the parameterized arrow body finite element model and the trajectory operation parameters; calling finite element software, and simulating and determining kinetic parameters through the finite element software according to the real-time arrow body finite element model of each characteristic time point;
and determining each parameter of the elastic rocket body motion equation according to the ballistic operation parameter, the pneumatic operation parameter, the kinetic parameter and the structural parameter of the carrier rocket.
2. The method of calculating arrow body equation of motion coefficients of claim 1, wherein:
the various flight control modes include: one or more of attitude control engine, gas rudder, air rudder and swinging jet pipe;
the initial design trajectory parameters comprise time, rocket body mass, mass center, relative speed, flight Mach number, dynamic pressure, attack angle, trajectory inclination angle, active machine effective thrust, atmospheric density, pitching rudder deflection angle and attitude control engine effective thrust corresponding to different flight control modes;
the design of the pneumatic initial parameters correspondingly comprises the following steps according to different flight control modes: the combination of multiple items in the sectional missile normal force gradient, the sectional missile normal force coefficient to attack angle derivative, the sectional missile lateral force coefficient to sideslip angle derivative, the air rudder normal force coefficient to rudder deflection angle derivative, the sectional missile characteristic area, the gas rudder lift gradient and the air rudder lift gradient;
the control parameters comprise flight stages of the carrier rocket, ignition time and fuel exhaustion time of an engine in the flight process corresponding to the flight stages, and time intervals and vibration mode stages of each characteristic time point.
3. The method for calculating rocket body equation of motion coefficients according to claim 2, wherein the determining ballistic operating parameters of each characteristic time point according to the design ballistic initial parameters and each characteristic time point under various control modes specifically comprises:
performing interpolation calculation on the designed ballistic initial parameters of the corresponding flight control modes according to each characteristic time point, and determining the ballistic operating parameters of each characteristic time point, wherein the method comprises the following steps: time, rocket mass, centroid, relative velocity, flight mach number, dynamic pressure, angle of attack, ballistic inclination, active machine effective thrust, atmospheric density, pitch rudder deflection angle, and attitude control engine effective thrust.
4. The method for calculating rocket body equation of motion coefficients according to claim 3, wherein determining pneumatic operating parameters according to design pneumatic initial parameters and ballistic operating parameters specifically comprises:
determining the normal force gradient of the segmented projectile body and/or the lift gradient of the air rudder at each moment according to the Mach number, the attack angle and the rudder deflection angle in the ballistic operation data of each characteristic time point;
and carrying out interpolation calculation according to the normal force gradient of the sectional missile body and/or the lift gradient of the air rudder to the designed pneumatic initial dynamic parameters of the corresponding control mode, and determining the pneumatic operation parameters of each characteristic time point, wherein the parameters comprise the normal force gradient of the sectional missile body, the coefficient of the normal force of the sectional missile body to an attack angle derivative, the coefficient of the lateral force of the sectional missile body to a slip angle derivative, the coefficient of the normal force of the air rudder to a rudder deflection angle derivative, the characteristic area of the sectional missile body, the lift gradient of the gas rudder and/or the lift gradient of the air rudder.
5. The method for calculating arrow motion equation coefficients of claim 3, wherein the determining a real-time arrow finite element model for each characteristic time point based on the parameterized arrow finite element model and ballistic operating parameters, and the simulating determination of the kinetic parameters specifically comprises:
calculating a parameterized arrow finite element model according to the arrow mass and the mass center of each characteristic time point, and determining a real-time arrow finite element model of each characteristic time point;
and determining the dynamic parameters of each characteristic moment point according to the real-time arrow body finite element model and the vibration mode order, wherein the dynamic parameters comprise frequency, vibration mode and section quality data of the arrow body.
6. The method for calculating the coefficients of an arrow motion equation of claim 1, wherein the parameters of the elastic arrow motion equation include the circular frequency of each characteristic time point, the mode shape and slope at the inertial set, and the influence coefficient of the generalized aerodynamic force of the equivalent rigid arrow body on the vibration of each mode shape in proportion to the rotation angular velocity.
7. The method of calculating rocket body equation of motion coefficients as recited in claim 1, wherein said structural parameters comprise: the distance from the installation base plane of the inertial unit to the actual sharp point of the whole rocket, the distance from the action point of the engine spray pipe to the actual sharp point of the whole rocket and the deviation of the thrust action line of the engine are obtained.
8. A system for implementing the method of calculating arrow body equation of motion coefficients of claim 1, comprising:
the data input module is used for loading design trajectory initial parameters, design pneumatic initial parameters and a parameterized arrow body finite element model under various flight control modes of the arrow body; inputting control parameters and structural parameters of an arrow body;
the data processing module is used for determining all characteristic time points according to the control parameters; determining ballistic operating parameters of each characteristic time point according to the designed ballistic initial parameters and each characteristic time point; determining pneumatic operation parameters according to the designed pneumatic initial parameters and ballistic operation parameters; determining a real-time arrow body finite element model of each characteristic time point according to the parameterized arrow body finite element model and the trajectory operation parameters; calling finite element software, and simulating and determining kinetic parameters through the finite element software according to the real-time arrow body finite element model of each characteristic time point; and determining each parameter of the elastic rocket body motion equation according to the ballistic operation parameter, the pneumatic operation parameter, the kinetic parameter and the structural parameter of the carrier rocket.
9. The arrow body equation of motion coefficient calculation system of claim 8, wherein the data input module includes:
a data loading unit for loading design trajectory initial parameters, design aerodynamic initial parameters and a parameterized rocket body finite element model under various flight control modes of the carrier rocket;
-a data input unit for control parameters and structural parameters of the arrow body;
the data processing module comprises:
-a characteristic moment processing unit for control parameters and structural parameters of the arrow body;
-a ballistic parameter processing unit for determining ballistic operational parameters for each characteristic time point from the design ballistic initial parameters and each characteristic time point;
-a pneumatic parameter processing unit for determining pneumatic operational parameters from design pneumatic initial parameters and ballistic operational parameters;
a model processing unit for determining a real-time arrow finite element model for each characteristic time instant point based on the parameterized arrow finite element model and the ballistic operational parameters,
-a calling unit for calling finite element software for determining kinetic parameters by means of finite element software simulation based on the real-time arrow finite element model of each characteristic time instant;
-a coefficient solving unit for determining the parameters of the elastic rocket body equations of motion from ballistic, aerodynamic, kinetic and structural parameters.
10. The rocket body equation of motion coefficient calculation system of claim 8 further comprising a storage module for storing data from the data input module and the data processing module.
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| CN115455790A (en)* | 2022-06-01 | 2022-12-09 | 北京理工大学 | A finite element intelligent post-processing method for launch vehicle dynamics |
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