CN112711815B - Aircraft modeling and model characteristic analysis system - Google Patents

Aircraft modeling and model characteristic analysis system Download PDF

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Publication number
CN112711815B
CN112711815B CN202110318172.3A CN202110318172A CN112711815B CN 112711815 B CN112711815 B CN 112711815B CN 202110318172 A CN202110318172 A CN 202110318172A CN 112711815 B CN112711815 B CN 112711815B
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aircraft
model
aerodynamic
state
module
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CN112711815A (en
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刘振
吴士广
张天乐
蒲志强
丘腾海
易建强
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Institute of Automation of Chinese Academy of Science
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    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F30/00—Computer-aided design [CAD]
    • G06F30/10—Geometric CAD
    • G06F30/15—Vehicle, aircraft or watercraft design
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F30/00—Computer-aided design [CAD]
    • G06F30/20—Design optimisation, verification or simulation
    • G06F30/28—Design optimisation, verification or simulation using fluid dynamics, e.g. using Navier-Stokes equations or computational fluid dynamics [CFD]
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F2119/00—Details relating to the type or aim of the analysis or the optimisation
    • G06F2119/14—Force analysis or force optimisation, e.g. static or dynamic forces
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T90/00—Enabling technologies or technologies with a potential or indirect contribution to GHG emissions mitigation

Abstract

The invention belongs to the field of aircraft design and analysis, particularly relates to an aircraft modeling and model characteristic analysis system, and aims to solve the problem that the existing aircraft modeling and model characteristic analysis system cannot efficiently and accurately perform aircraft modeling and model characteristic analysis. The system comprises: the system management module is configured to carry out initialization setting on the system; the aerodynamic/aerodynamic moment module is configured to calculate an aerodynamic force acting on/about the aircraft center of mass; the six-degree-of-freedom motion equation module is configured to construct a full-dimensional motion state equation set of the aircraft; the reference motion state solving module is configured to solve the full-dimensional motion state of the aircraft under the reference motion condition; the motion equation linearization module is configured to generate an aircraft all-state linearization model in a state space form; the characteristic analysis and display module is configured to perform aircraft model characteristic analysis. The invention realizes the high-efficiency and accurate modeling of the aircraft and the characteristic analysis of the model.

Description

Aircraft modeling and model characteristic analysis system
Technical Field
The invention belongs to the field of aircraft design and analysis, and particularly relates to an aircraft modeling and model characteristic analysis system.
Background
In recent years, with the rapid development of flight technology, various aircrafts are widely applied to military and civil fields. Modeling and model characterization is an extremely important task in the development of aircraft, especially in the initial design phase of the aircraft. In the preliminary design stage of the aircraft, designers need to perform modeling and model characteristic analysis based on a preliminary design scheme, and modify the aircraft structure, the aerodynamic parameters and the like selected by the preliminary design scheme on the basis of the characteristic analysis, and the process is repeated until the design scheme meeting the requirements is obtained. On the other hand, with the complexity and diversification of the task and environment of aircraft operation, the complexity of aircraft design work is also increasing continuously. Therefore, how to perform rapid and accurate aircraft modeling and model characteristic analysis is an urgent need of aircraft engineering designers.
In the aircraft design engineering practice, in order to measure different flight quality parameters, designers often adopt different methods for evaluation, but the evaluation methods usually have certain subjective colors, lack uniform measurement standards, and are difficult to comprehensively coordinate different flight quality parameters. Therefore, in order to conveniently evaluate the aircraft characteristics and realize rapid and efficient aircraft modeling and model characteristic analysis, a comprehensive system integrating aircraft modeling and model characteristic analysis is required to be constructed.
Disclosure of Invention
In order to solve the above problems in the prior art, that is, to solve the problem that the existing aircraft modeling and model characteristic analysis system cannot efficiently and accurately perform aircraft modeling and model characteristic analysis, a first aspect of the present invention provides an aircraft modeling and model characteristic analysis system, including: the system comprises an aerodynamic module, an aerodynamic moment module, a six-degree-of-freedom motion equation module, a reference motion state solving module, a motion equation linearization module, a longitudinal/lateral motion module, a characteristic analysis and display module and a system management module;
the system management module is configured to perform initialization setting on the system; further configured to store aircraft state information and model characteristics; the initialization setting comprises aircraft structure parameter setting, reference motion condition setting and transfer function input/output quantity setting;
the aerodynamic module is configured to calculate aerodynamic forces acting on the center of mass of the aircraft;
the aerodynamic moment module is configured to calculate an aerodynamic moment acting around a center of mass of the aircraft;
the six-degree-of-freedom motion equation module is configured to combine the aerodynamic force and the aerodynamic moment to construct an aircraft full-dimensional motion state equation set; the full-dimensional motion state equation set comprises a moment equation set, an angular displacement equation set, a force equation set, a linear displacement equation set, an actuating mechanism model and an atmospheric model;
the reference motion state solving module is configured to solve the full-dimensional motion state of the aircraft under the reference motion condition;
the motion equation linearization module is configured to linearize the aircraft motion equation at the reference motion state to generate an aircraft all-state linearization model in a state space form;
the longitudinal/lateral motion module is configured to extract a longitudinal/lateral motion equation set based on the aircraft all-state linearized model and generate an aircraft longitudinal/lateral motion state linearized model in a state space form;
the characteristic analysis and display module is configured to perform aircraft model characteristic analysis based on a pneumatic stability derivative, a full-state linearized model zero-pole, an aircraft longitudinal/lateral motion state linearized model zero-pole, and a corresponding transfer function and stability margin between any one of the control input quantity and the flight state output quantity, and output a model analysis result;
the derivative of aerodynamic stability comprises an aerodynamic coefficient of one dimension along the three axes of the aircraft body coordinate system
Figure 576339DEST_PATH_IMAGE001
、
Figure 152814DEST_PATH_IMAGE002
、
Figure 954548DEST_PATH_IMAGE003
And a aerodynamic moment of one dimension in three axial directions around the coordinate system of the aircraft bodyCoefficient of performance
Figure 34500DEST_PATH_IMAGE004
、
Figure 778334DEST_PATH_IMAGE005
、
Figure 310946DEST_PATH_IMAGE006
。
In some preferred embodiments, the aerodynamic module "calculates aerodynamic forces acting on the center of mass of the aircraft" by: based on the real-time state of the aircraft and aerodynamic force data, solving the aerodynamic force of the aircraft at the current moment through an interpolation algorithm; the method comprises the following specific steps:
Figure 650792DEST_PATH_IMAGE007
Figure 850829DEST_PATH_IMAGE008
Figure 768494DEST_PATH_IMAGE009
Figure 53982DEST_PATH_IMAGE010
Figure 197518DEST_PATH_IMAGE011
Figure 252062DEST_PATH_IMAGE012
wherein the content of the first and second substances,
Figure 72119DEST_PATH_IMAGE001
、
Figure 579324DEST_PATH_IMAGE002
、
Figure 651185DEST_PATH_IMAGE003
respectively are aerodynamic coefficients with the dimension of one along the three axial directions of the aircraft body coordinate system,
Figure 435601DEST_PATH_IMAGE013
、
Figure 567505DEST_PATH_IMAGE014
、
Figure 686640DEST_PATH_IMAGE015
representing the aerodynamic force component in the aircraft body coordinate system,
Figure 562192DEST_PATH_IMAGE016
in order to generate a dynamic pressure,
Figure 466694DEST_PATH_IMAGE017
for the purpose of reference area, the area of the reference,
Figure 503920DEST_PATH_IMAGE018
、
Figure 110351DEST_PATH_IMAGE019
and
Figure 524015DEST_PATH_IMAGE020
respectively an aileron, an elevator and a rudder deflection of the aircraft,
Figure 283024DEST_PATH_IMAGE021
and
Figure 225572DEST_PATH_IMAGE022
respectively a flight attack angle and a sideslip angle,
Figure 194665DEST_PATH_IMAGE023
caused by deflection of elevators
Figure 539583DEST_PATH_IMAGE013
The coefficient of variation of the axial aerodynamic force,
Figure 277732DEST_PATH_IMAGE024
generated when pitch rate is not zero
Figure 797706DEST_PATH_IMAGE013
The coefficient of variation of the axial dynamic aerodynamic force,
Figure 988516DEST_PATH_IMAGE025
caused by deflection of ailerons and rudder
Figure 868616DEST_PATH_IMAGE014
The coefficient of variation of the axial aerodynamic force,
Figure 726851DEST_PATH_IMAGE026
generated when the yaw rate is not zero
Figure 417726DEST_PATH_IMAGE014
The coefficient of variation of the axial dynamic aerodynamic force,
Figure 830253DEST_PATH_IMAGE027
generated when the roll rate is not zero
Figure 389410DEST_PATH_IMAGE014
The coefficient of variation of the axial dynamic aerodynamic force,
Figure 226785DEST_PATH_IMAGE028
caused by deflection of elevators
Figure 947616DEST_PATH_IMAGE015
The coefficient of variation of the axial aerodynamic force,
Figure 253964DEST_PATH_IMAGE029
generated when pitch rate is not zero
Figure 351233DEST_PATH_IMAGE015
And (4) axial dynamic aerodynamic force variation coefficient.
In some preferred embodiments, the aerodynamic moment module "calculates the aerodynamic moment about the center of mass of the aircraft" by: solving the aerodynamic moment of the aircraft at the current moment by an interpolation algorithm based on the real-time state of the aircraft and the aerodynamic moment data; the method comprises the following specific steps:
Figure 43114DEST_PATH_IMAGE030
Figure 200426DEST_PATH_IMAGE031
Figure 587545DEST_PATH_IMAGE032
Figure 363871DEST_PATH_IMAGE033
Figure 51205DEST_PATH_IMAGE034
Figure 506981DEST_PATH_IMAGE035
wherein the content of the first and second substances,
Figure 381396DEST_PATH_IMAGE036
、
Figure 695834DEST_PATH_IMAGE037
and
Figure 237674DEST_PATH_IMAGE038
representing the aerodynamic moment components in three axes around the aircraft body coordinate system,
Figure 471209DEST_PATH_IMAGE039
for the purpose of reference to the length of the strip,
Figure 223133DEST_PATH_IMAGE004
、
Figure 528213DEST_PATH_IMAGE005
、
Figure 314772DEST_PATH_IMAGE006
respectively are aerodynamic moment coefficients with the dimension of one around the three axial directions of the coordinate system of the aircraft body,
Figure 719209DEST_PATH_IMAGE040
caused by deflection of ailerons and rudder
Figure 443582DEST_PATH_IMAGE013
The coefficient of variation of the axial aerodynamic moment,
Figure 490036DEST_PATH_IMAGE041
generated when the yaw rate is not zero
Figure 740888DEST_PATH_IMAGE013
The axial dynamic pneumatic moment variation coefficient,
Figure 721088DEST_PATH_IMAGE042
generated when the roll rate is not zero
Figure 57391DEST_PATH_IMAGE013
The axial dynamic pneumatic moment variation coefficient,
Figure 517322DEST_PATH_IMAGE043
caused by deflection of elevators
Figure 888261DEST_PATH_IMAGE014
The coefficient of variation of the axial aerodynamic moment,
Figure 759134DEST_PATH_IMAGE044
generated when pitch rate is not zero
Figure 848312DEST_PATH_IMAGE014
The axial dynamic pneumatic moment variation coefficient,
Figure 846355DEST_PATH_IMAGE045
caused by deflection of ailerons and rudder
Figure 337380DEST_PATH_IMAGE015
The coefficient of variation of the axial aerodynamic moment,
Figure 644733DEST_PATH_IMAGE046
generated when the yaw rate is not zero
Figure 690049DEST_PATH_IMAGE015
The axial dynamic pneumatic moment variation coefficient,
Figure 616417DEST_PATH_IMAGE047
generated when the roll rate is not zero
Figure 837314DEST_PATH_IMAGE015
And (4) axial dynamic aerodynamic moment variation coefficient.
In some preferred embodiments, the actuator model is:
Figure 190935DEST_PATH_IMAGE048
Figure 848181DEST_PATH_IMAGE049
wherein the content of the first and second substances,
Figure 578240DEST_PATH_IMAGE050
in order to input a command for the actuator,
Figure 653643DEST_PATH_IMAGE051
in order to be the actual output of the actuator,
Figure 178166DEST_PATH_IMAGE052
and
Figure 198074DEST_PATH_IMAGE053
respectively the frequency and the damping, respectively,
Figure 859387DEST_PATH_IMAGE054
a magnitude limiting function representing the actuator,
Figure 913931DEST_PATH_IMAGE055
、
Figure 484721DEST_PATH_IMAGE056
the maximum and minimum limiting amplitudes, respectively,
Figure 991925DEST_PATH_IMAGE057
representing the amount of the second derivative actually output by the actuator.
In some preferred embodiments, in the reference motion state solving, "solving the full-dimensional motion state of the aircraft under the reference motion condition", the method includes:
carrying out square weighted sum on the derivative values of the full-dimensional motion state of the aircraft to construct a cost function;
and solving the full-dimensional motion state of the aircraft with the minimum cost function based on a preset reference motion condition.
In some preferred embodiments, the reference motion condition is:
Figure 188420DEST_PATH_IMAGE058
Figure 97470DEST_PATH_IMAGE059
Figure 104741DEST_PATH_IMAGE060
wherein the content of the first and second substances,
Figure 833662DEST_PATH_IMAGE061
and
Figure 709214DEST_PATH_IMAGE062
respectively the set flight speed value and the flight altitude value,
Figure 597405DEST_PATH_IMAGE022
、
Figure 900210DEST_PATH_IMAGE051
representing the sideslip angle, roll angle of the aircraft,
Figure 991794DEST_PATH_IMAGE063
、
Figure 405458DEST_PATH_IMAGE064
、
Figure 413734DEST_PATH_IMAGE065
representing the roll rate, pitch rate and yaw rate of the aircraft in a body coordinate system,
Figure 621862DEST_PATH_IMAGE066
the speed of flight is indicated as a function of,
Figure 590955DEST_PATH_IMAGE067
representing the altitude of the aircraft in the ground coordinate system,
Figure 683676DEST_PATH_IMAGE068
the derivative of the flight speed is represented as,
Figure 156245DEST_PATH_IMAGE069
representing the angle of attack derivative and the sideslip angle derivative of the aircraft,
Figure 928417DEST_PATH_IMAGE070
the roll rate derivative, pitch rate derivative, and yaw rate derivative of the aircraft are represented.
In some preferred embodiments, the cost function model is:
Figure 384806DEST_PATH_IMAGE071
wherein the content of the first and second substances,
Figure 15638DEST_PATH_IMAGE072
in order to be the weight coefficient,
Figure 608294DEST_PATH_IMAGE073
the derivative of the flying height is indicated.
In some preferred embodiments, the full-state linearized model of the aircraft is:
Figure 158224DEST_PATH_IMAGE074
Figure 226543DEST_PATH_IMAGE075
wherein the content of the first and second substances,
Figure 520121DEST_PATH_IMAGE076
is the full state quantity of the aircraft,
Figure 108228DEST_PATH_IMAGE077
the derivative of the full state quantity of the aircraft is represented,
Figure 94639DEST_PATH_IMAGE078
as input quantities, output quantities of the model
Figure 384675DEST_PATH_IMAGE014
In the state quantity
Figure 216364DEST_PATH_IMAGE013
On the basis of the three axial directions of the aircraft
Figure 49191DEST_PATH_IMAGE079
Acceleration, flight mach number and dynamic pressure value of the aircraft,
Figure 81869DEST_PATH_IMAGE080
，
Figure 734567DEST_PATH_IMAGE081
、
Figure 494582DEST_PATH_IMAGE082
、
Figure 181915DEST_PATH_IMAGE083
and
Figure 244549DEST_PATH_IMAGE084
respectively representing the state matrix, control matrix, observation matrix and feedforward matrix of the model.
In some preferred embodiments, the aircraft longitudinal motion state linearization model is:
Figure 259910DEST_PATH_IMAGE085
Figure 698981DEST_PATH_IMAGE086
wherein the content of the first and second substances,
Figure 368385DEST_PATH_IMAGE087
is the longitudinal motion state quantity of the aircraft,
Figure 601920DEST_PATH_IMAGE088
the derivative of the longitudinal motion state quantity of the aircraft and the longitudinal motion output quantity of the model
Figure 104576DEST_PATH_IMAGE089
，
Figure 347339DEST_PATH_IMAGE090
、
Figure 478106DEST_PATH_IMAGE091
、
Figure 272756DEST_PATH_IMAGE092
And
Figure 387342DEST_PATH_IMAGE093
and the state matrix, the control matrix, the observation matrix and the feedforward matrix respectively represent the longitudinal motion state linearization model.
In some preferred embodiments, the aircraft lateral motion state linearized model is:
Figure 309162DEST_PATH_IMAGE094
Figure 560014DEST_PATH_IMAGE095
wherein the content of the first and second substances,
Figure 259986DEST_PATH_IMAGE096
is the lateral motion state quantity of the aircraft,
Figure 861869DEST_PATH_IMAGE097
the derivative of the state quantity of lateral motion of the aircraft, the lateral motion output quantity of the model
Figure 321800DEST_PATH_IMAGE098
，
Figure 692738DEST_PATH_IMAGE099
、
Figure 438978DEST_PATH_IMAGE100
、
Figure 652790DEST_PATH_IMAGE101
And
Figure 775467DEST_PATH_IMAGE102
and the state matrix, the control matrix, the observation matrix and the feedforward matrix respectively represent the lateral motion state linearization model.
The invention has the beneficial effects that:
the invention realizes the high-efficiency and accurate modeling of the aircraft and the characteristic analysis of the model.
The method can quickly form an aerodynamic force/moment model and an aircraft six-degree-of-freedom motion equation according to different aircraft structures, aerodynamics and other design schemes, solve a linearized equation and longitudinal/lateral equation decomposition under the set reference motion condition, output various data and curves of aerodynamic stability derivative curves, a zero pole point diagram, a bode diagram, stability margin and other characteristic model characteristics, comprehensively analyze and evaluate the aircraft model characteristics in a comprehensive form, and improve the accuracy of evaluation and analysis. And moreover, repeated iterative improvement is convenient for designers aiming at key parameters, and the method is particularly suitable for rapid updating iteration of a design scheme in an initial design stage of the aircraft.
Drawings
Other features, objects and advantages of the present application will become more apparent upon reading of the following detailed description of non-limiting embodiments thereof, made with reference to the accompanying drawings.
FIG. 1 is a block diagram of an aircraft modeling and model characterization system according to an embodiment of the present invention;
FIG. 2 is a simplified flow diagram of a method for modeling and analyzing model characteristics of an aircraft in accordance with an embodiment of the present invention;
FIG. 3 is a detailed flow diagram of a method for modeling and analyzing model characteristics of an aircraft in accordance with an embodiment of the invention;
fig. 4 is a schematic structural diagram of a computer system suitable for implementing an electronic device according to an embodiment of the present application.
Detailed Description
In order to make the objects, technical solutions and advantages of the present invention clearer, the technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings, and it is apparent that the described embodiments are some, but not all embodiments of the present invention. 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 invention.
The present application will be described in further detail with reference to the following drawings and examples. It is to be understood that the specific embodiments described herein are merely illustrative of the relevant invention and not restrictive of the invention. It should be noted that, for convenience of description, only the portions related to the related invention are shown in the drawings.
It should be noted that the embodiments and features of the embodiments in the present application may be combined with each other without conflict.
The aircraft modeling and model characteristic analysis system of the present invention, as shown in fig. 1, includes: the system comprises an aerodynamic module, an aerodynamic moment module, a six-degree-of-freedom motion equation module, a reference motion state solving module, a motion equation linearization module, a longitudinal/lateral motion module, a characteristic analysis and display module and a system management module;
the system management module is configured to perform initialization setting on the system; further configured to store aircraft state information and model characteristics; the initialization setting comprises aircraft structure parameter setting, reference motion condition setting and transfer function input/output quantity setting;
the aerodynamic module is configured to calculate aerodynamic forces acting on the center of mass of the aircraft;
the aerodynamic moment module is configured to calculate an aerodynamic moment acting around a center of mass of the aircraft;
the six-degree-of-freedom motion equation module is configured to combine the aerodynamic force and the aerodynamic moment to construct an aircraft full-dimensional motion state equation set; the full-dimensional motion state equation set comprises a moment equation set, an angular displacement equation set, a force equation set, a linear displacement equation set, an actuating mechanism model and an atmospheric model;
the reference motion state solving module is configured to solve the full-dimensional motion state of the aircraft under the reference motion condition;
the motion equation linearization module is configured to linearize the aircraft motion equation at the reference motion state to generate an aircraft all-state linearization model in a state space form;
the longitudinal/lateral motion module is configured to extract a longitudinal/lateral motion equation set based on the aircraft all-state linearized model and generate an aircraft longitudinal/lateral motion state linearized model in a state space form;
the characteristic analysis and display module is configured to perform aircraft model characteristic analysis based on a pneumatic stability derivative, a full-state linearized model zero-pole, an aircraft longitudinal/lateral motion state linearized model zero-pole, and a corresponding transfer function and stability margin between any one of the control input quantity and the flight state output quantity, and output a model analysis result;
the derivative of aerodynamic stability comprises an aerodynamic coefficient of one dimension along the three axes of the aircraft body coordinate system
Figure 876278DEST_PATH_IMAGE001
、
Figure 58998DEST_PATH_IMAGE002
、
Figure 485738DEST_PATH_IMAGE003
And aerodynamic moment coefficient of dimension one in three axial directions around the coordinate system of the aircraft body
Figure 412106DEST_PATH_IMAGE004
、
Figure 633003DEST_PATH_IMAGE005
、
Figure 986624DEST_PATH_IMAGE006
。
In order to more clearly illustrate the aircraft modeling and model characterization system of the present invention, the following description will be made in detail with reference to the accompanying drawings.
The system management module is configured to perform initialization setting on the system; further configured to store aircraft state information and model characteristics; the initialization setting comprises aircraft structure parameter setting, reference motion condition setting and transfer function input/output quantity setting;
in this embodiment, when the system is initialized, the system management module performs initialization setting. The initialization settings comprise aircraft structure parameter settings, reference motion condition settings and transfer function input/output quantity settings. And storing the aircraft state information and the model characteristics after the aircraft model characteristic analysis is completed.
The aerodynamic module is configured to calculate aerodynamic forces acting on the center of mass of the aircraft;
in the embodiment, the aerodynamic module comprises an aerodynamic model, an aerodynamic data packet and an interpolation algorithm;
the aerodynamic force model is used for constructing aerodynamic force components along the three axial directions of an aircraft body coordinate system;
the aerodynamic data packet is used for storing aerodynamic data in a table form;
and the interpolation algorithm is used for solving the aerodynamic force at the current moment according to the real-time state of the aircraft and the aerodynamic force data.
Three axial aerodynamic force components
Figure 519236DEST_PATH_IMAGE013
、
Figure 373929DEST_PATH_IMAGE014
And
Figure 573966DEST_PATH_IMAGE015
the calculation process of (2) is shown in the formulas (1), (3), (4), (5) and (6):
Figure 973854DEST_PATH_IMAGE007
（1）
Figure 993763DEST_PATH_IMAGE008
（2）
Figure 386567DEST_PATH_IMAGE009
（3）
Figure 441111DEST_PATH_IMAGE103
（4）
Figure 136534DEST_PATH_IMAGE104
（5）
Figure 784684DEST_PATH_IMAGE105
（6）
wherein the content of the first and second substances,
Figure 590966DEST_PATH_IMAGE001
、
Figure 890229DEST_PATH_IMAGE002
、
Figure 490975DEST_PATH_IMAGE003
respectively are aerodynamic coefficients with the dimension of one along the three axial directions of the aircraft body coordinate system,
Figure 626421DEST_PATH_IMAGE016
in order to generate a dynamic pressure,
Figure 236394DEST_PATH_IMAGE017
for the purpose of reference area, the area of the reference,
Figure 924252DEST_PATH_IMAGE018
、
Figure 102424DEST_PATH_IMAGE019
and
Figure 584220DEST_PATH_IMAGE020
respectively an aileron, an elevator and a rudder deflection of the aircraft,
Figure 122518DEST_PATH_IMAGE021
and
Figure 740581DEST_PATH_IMAGE022
respectively a flight attack angle and a sideslip angle,
Figure 214288DEST_PATH_IMAGE023
、
Figure 58747DEST_PATH_IMAGE024
、
Figure 10523DEST_PATH_IMAGE025
、
Figure 873305DEST_PATH_IMAGE026
、
Figure 517913DEST_PATH_IMAGE027
、
Figure 849669DEST_PATH_IMAGE028
and
Figure 605135DEST_PATH_IMAGE029
and (4) carrying out interpolation solution according to an interpolation algorithm based on variables in brackets respectively.
Figure 322424DEST_PATH_IMAGE023
Caused by deflection of elevators
Figure 872354DEST_PATH_IMAGE013
Axial pneumaticsThe coefficient of variation of the force is,
Figure 816039DEST_PATH_IMAGE024
generated when pitch rate is not zero
Figure 250563DEST_PATH_IMAGE013
The coefficient of variation of the axial dynamic aerodynamic force,
Figure 963304DEST_PATH_IMAGE025
caused by deflection of ailerons and rudder
Figure 546120DEST_PATH_IMAGE014
The coefficient of variation of the axial aerodynamic force,
Figure 977101DEST_PATH_IMAGE026
generated when the yaw rate is not zero
Figure 949736DEST_PATH_IMAGE014
The coefficient of variation of the axial dynamic aerodynamic force,
Figure 782563DEST_PATH_IMAGE027
generated when the roll rate is not zero
Figure 798929DEST_PATH_IMAGE014
The coefficient of variation of the axial dynamic aerodynamic force,
Figure 451628DEST_PATH_IMAGE028
caused by deflection of elevators
Figure 352587DEST_PATH_IMAGE015
The coefficient of variation of the axial aerodynamic force,
Figure 649708DEST_PATH_IMAGE029
generated when pitch rate is not zero
Figure 977921DEST_PATH_IMAGE015
Axial dynamic aerodynamic force variationA coefficient;
the 'carrying out interpolation solution according to interpolation algorithm' means that aerodynamic force data packets are respectively solved according to the bilinear interpolation principle by utilizing aerodynamic force data packets
Figure 976970DEST_PATH_IMAGE023
、
Figure 681621DEST_PATH_IMAGE024
、
Figure 98827DEST_PATH_IMAGE025
、
Figure 597941DEST_PATH_IMAGE026
、
Figure 959652DEST_PATH_IMAGE027
、
Figure 327049DEST_PATH_IMAGE028
And
Figure 723395DEST_PATH_IMAGE029
the fitting function of (1).
The aerodynamic moment module is configured to calculate an aerodynamic moment acting around a center of mass of the aircraft;
in this embodiment, the aerodynamic moment module includes three parts, namely an aerodynamic moment model, an aerodynamic moment data packet and an interpolation algorithm;
the aerodynamic moment model is used for constructing aerodynamic moment components around the three axial directions of an aircraft body coordinate system;
the aerodynamic torque data packet is used for storing aerodynamic torque data in a table form;
and the interpolation algorithm is used for solving the aerodynamic moment at the current moment according to the real-time state of the aircraft and the aerodynamic moment data.
Three-axial aerodynamic moment component around aircraft body coordinate system
Figure 268777DEST_PATH_IMAGE036
、
Figure 383363DEST_PATH_IMAGE037
And
Figure 26222DEST_PATH_IMAGE038
the calculation process of (2) is shown in the formulas (7), (8), (9), (10), (11) and (12):
Figure 542654DEST_PATH_IMAGE030
（7）
Figure 258937DEST_PATH_IMAGE031
（8）
Figure 860820DEST_PATH_IMAGE032
（9）
Figure 570019DEST_PATH_IMAGE033
（10）
Figure 675378DEST_PATH_IMAGE106
（11）
Figure 687196DEST_PATH_IMAGE035
（12）
wherein the content of the first and second substances,
Figure 386162DEST_PATH_IMAGE036
、
Figure 774418DEST_PATH_IMAGE037
and
Figure 858917DEST_PATH_IMAGE038
representing three axes around the aircraft body coordinate systemThe aerodynamic moment component of (a) is,
Figure 41637DEST_PATH_IMAGE039
for the purpose of reference to the length of the strip,
Figure 493478DEST_PATH_IMAGE004
、
Figure 419846DEST_PATH_IMAGE005
、
Figure 234218DEST_PATH_IMAGE006
respectively are aerodynamic moment coefficients with the dimension of one around the three axial directions of the coordinate system of the aircraft body,
Figure 978052DEST_PATH_IMAGE040
、
Figure 776244DEST_PATH_IMAGE041
、
Figure 116089DEST_PATH_IMAGE042
、
Figure 50547DEST_PATH_IMAGE043
、
Figure 979931DEST_PATH_IMAGE044
、
Figure 265419DEST_PATH_IMAGE045
、
Figure 268010DEST_PATH_IMAGE046
and
Figure 197920DEST_PATH_IMAGE047
and (4) carrying out interpolation solution according to an interpolation algorithm based on variables in brackets respectively.
Figure 158923DEST_PATH_IMAGE040
Caused by deflection of ailerons and rudder
Figure 525182DEST_PATH_IMAGE013
The coefficient of variation of the axial aerodynamic moment,
Figure 597043DEST_PATH_IMAGE041
generated when the yaw rate is not zero
Figure 647039DEST_PATH_IMAGE013
The axial dynamic pneumatic moment variation coefficient,
Figure 513363DEST_PATH_IMAGE042
generated when the roll rate is not zero
Figure 507864DEST_PATH_IMAGE013
The axial dynamic pneumatic moment variation coefficient,
Figure 242471DEST_PATH_IMAGE043
caused by deflection of elevators
Figure 6028DEST_PATH_IMAGE014
The coefficient of variation of the axial aerodynamic moment,
Figure 918620DEST_PATH_IMAGE044
generated when pitch rate is not zero
Figure 665996DEST_PATH_IMAGE014
The axial dynamic pneumatic moment variation coefficient,
Figure 204294DEST_PATH_IMAGE045
caused by deflection of ailerons and rudder
Figure 822357DEST_PATH_IMAGE015
The coefficient of variation of the axial aerodynamic moment,
Figure 296064DEST_PATH_IMAGE046
generated when the yaw rate is not zero
Figure 140523DEST_PATH_IMAGE015
The axial dynamic pneumatic moment variation coefficient,
Figure 92298DEST_PATH_IMAGE047
generated when the roll rate is not zero
Figure 223590DEST_PATH_IMAGE015
Axial dynamic aerodynamic moment variation coefficient;
the 'carrying out interpolation solution according to interpolation algorithm' means that the pneumatic torque data packets are respectively solved according to the bilinear interpolation principle by utilizing the pneumatic torque data packets
Figure 337039DEST_PATH_IMAGE040
、
Figure 668795DEST_PATH_IMAGE041
、
Figure 158682DEST_PATH_IMAGE042
、
Figure 16916DEST_PATH_IMAGE043
、
Figure 957060DEST_PATH_IMAGE044
、
Figure 635166DEST_PATH_IMAGE045
、
Figure 804110DEST_PATH_IMAGE046
And
Figure 516851DEST_PATH_IMAGE047
the fitting function of (1).
The six-degree-of-freedom motion equation module is configured to combine the aerodynamic force and the aerodynamic moment to construct an aircraft full-dimensional motion state equation set; the full-dimensional motion state equation set comprises a moment equation set, an angular displacement equation set, a force equation set, a linear displacement equation set, an actuating mechanism model and an atmospheric model;
in the embodiment, the full-dimensional motion state equation set comprises a moment equation set, an angular displacement equation set, a force equation set, a linear displacement equation set, an actuator model and an atmospheric model.
The moment equation set is constructed as shown in equations (13), (14) and (15):
Figure 627895DEST_PATH_IMAGE107
（13）
Figure 58877DEST_PATH_IMAGE108
（14）
Figure 890566DEST_PATH_IMAGE109
(15）
wherein the content of the first and second substances,
Figure 333180DEST_PATH_IMAGE110
representing the roll rate derivative, the pitch rate derivative and the yaw rate derivative of the aircraft in a body coordinate system,
Figure 490492DEST_PATH_IMAGE063
、
Figure 533403DEST_PATH_IMAGE064
、
Figure 903205DEST_PATH_IMAGE065
representing the roll rate, pitch rate and yaw rate of the aircraft in a body coordinate system,
Figure 465904DEST_PATH_IMAGE111
the inertia moments of the aircraft around the three axial directions of the aircraft body,
Figure 794117DEST_PATH_IMAGE112
for aircraft relative to airframeIn a coordinate system
Figure 668532DEST_PATH_IMAGE113
The product of the inertia of the face,
Figure 500747DEST_PATH_IMAGE114
，
Figure 42587DEST_PATH_IMAGE115
，
Figure 417067DEST_PATH_IMAGE116
,
Figure 778778DEST_PATH_IMAGE117
，
Figure 880595DEST_PATH_IMAGE118
，
Figure 542521DEST_PATH_IMAGE119
，
Figure 946957DEST_PATH_IMAGE120
，
Figure 936910DEST_PATH_IMAGE121
，
Figure 717784DEST_PATH_IMAGE122
，
Figure 93271DEST_PATH_IMAGE123
。
the system of angular displacement equations is constructed as shown in equations (16) (17) (18):
Figure 934188DEST_PATH_IMAGE124
（16）
Figure 411437DEST_PATH_IMAGE125
（17）
Figure 996002DEST_PATH_IMAGE126
（18）
wherein (A), (B), (C), (D), (C), (
Figure 491574DEST_PATH_IMAGE127
) Representing the roll, pitch and yaw derivatives of the aircraft,
Figure 237813DEST_PATH_IMAGE051
、
Figure 61413DEST_PATH_IMAGE128
、
Figure 325035DEST_PATH_IMAGE129
representing the roll, pitch and yaw of the aircraft.
The force equation set is constructed as shown in equations (19) (20) (21):
Figure 284901DEST_PATH_IMAGE130
（19）
Figure 595184DEST_PATH_IMAGE131
（20）
Figure 171659DEST_PATH_IMAGE132
（21）
wherein the content of the first and second substances,
Figure 973393DEST_PATH_IMAGE133
representing the aircraft flight speed derivative, angle of attack derivative and sideslip angle derivative,
Figure 53344DEST_PATH_IMAGE134
and
Figure 406965DEST_PATH_IMAGE135
respectively the track pitch angle and the track roll angle of the aircraft,
Figure 329791DEST_PATH_IMAGE066
the speed of flight is indicated as a function of,
Figure 794270DEST_PATH_IMAGE084
the indication of the flight resistance is that,
Figure 869673DEST_PATH_IMAGE136
representing aircraft engine thrust.
The linear displacement equation set is constructed as in equations (22) (23) (24):
Figure 394196DEST_PATH_IMAGE137
（22）
Figure 538738DEST_PATH_IMAGE138
（23）
Figure 72488DEST_PATH_IMAGE139
（24）
wherein the content of the first and second substances,
Figure 2397DEST_PATH_IMAGE140
representing the longitudinal displacement derivative, the lateral displacement derivative and the altitude derivative of the aircraft in a ground coordinate system,
Figure 697821DEST_PATH_IMAGE141
、
Figure 205026DEST_PATH_IMAGE142
、
Figure 401521DEST_PATH_IMAGE067
representing the longitudinal displacement, lateral displacement and altitude of the aircraft in a ground coordinate system,
Figure 310571DEST_PATH_IMAGE143
is the track azimuth of the aircraft.
The constructed actuator model is shown as formula (25):
Figure 317841DEST_PATH_IMAGE048
（25）
wherein the content of the first and second substances,
Figure 224924DEST_PATH_IMAGE050
in order to input a command for the actuator,
Figure 975843DEST_PATH_IMAGE051
in order to be the actual output of the actuator,
Figure 739399DEST_PATH_IMAGE052
and
Figure 776626DEST_PATH_IMAGE053
respectively the frequency and the damping, respectively,
Figure 648635DEST_PATH_IMAGE054
a magnitude limiting function representing the actuator,
Figure 62299DEST_PATH_IMAGE057
representing the amount of the second derivative actually output by the actuator. The amplitude limiting function of the actuator is shown in equation (26):
Figure 821308DEST_PATH_IMAGE144
（26）
wherein the content of the first and second substances,
Figure 763856DEST_PATH_IMAGE055
、
Figure 857583DEST_PATH_IMAGE056
maximum and minimum limiting amplitudes, respectively.
The constructed atmosphere model is shown as formula (27):
Figure 74938DEST_PATH_IMAGE145
（27）
wherein the content of the first and second substances,
Figure 688453DEST_PATH_IMAGE146
corresponding to a height of
Figure 333061DEST_PATH_IMAGE147
The density of (a) of (b),
Figure 523871DEST_PATH_IMAGE148
in order to normalize the parameters for the height,
Figure 403971DEST_PATH_IMAGE149
is expressed as a height of
Figure 262205DEST_PATH_IMAGE067
The density of (c).
The reference motion state solving module is configured to solve the full-dimensional motion state of the aircraft under the reference motion condition;
in the embodiment, the derivative values of the full-dimensional motion state of the aircraft are subjected to square weighted sum to construct a cost function; and solving the full-dimensional motion state of the aircraft with the minimum cost function based on a preset reference motion condition. Namely, the reference motion state solving module comprises an optimizing algorithm module and a cost function module.
The reference motion condition is shown in equations (28), (29) and (30):
Figure 267595DEST_PATH_IMAGE058
（28）
Figure 712745DEST_PATH_IMAGE059
（29）
Figure 193274DEST_PATH_IMAGE060
（30）
wherein the content of the first and second substances,
Figure 220529DEST_PATH_IMAGE061
and
Figure 331573DEST_PATH_IMAGE062
respectively, the set flight speed value and the flight height value.
The cost function module is shown in equation (31):
Figure 292050DEST_PATH_IMAGE071
(31)
wherein the content of the first and second substances,
Figure 310690DEST_PATH_IMAGE072
are weight coefficients.
The motion equation linearization module is configured to linearize the aircraft motion equation at the reference motion state to generate an aircraft all-state linearization model in a state space form;
in this embodiment, the full-state linearized model of the aircraft is shown in equation (32) (33):
Figure 877938DEST_PATH_IMAGE074
（32）
Figure 645037DEST_PATH_IMAGE075
（33）
wherein the content of the first and second substances,
Figure 297735DEST_PATH_IMAGE076
is the full state quantity of the aircraft,
Figure 326258DEST_PATH_IMAGE077
the derivative of the full state quantity of the aircraft is represented,
Figure 13592DEST_PATH_IMAGE078
as input quantities, output quantities of the model
Figure 76226DEST_PATH_IMAGE014
In the state quantity
Figure 91586DEST_PATH_IMAGE013
On the basis of the three axial directions of the aircraft
Figure 530658DEST_PATH_IMAGE079
Acceleration, flight mach number and dynamic pressure value of the aircraft,
Figure 931552DEST_PATH_IMAGE080
，
Figure 430667DEST_PATH_IMAGE081
、
Figure 933323DEST_PATH_IMAGE082
、
Figure 176086DEST_PATH_IMAGE083
and
Figure 572432DEST_PATH_IMAGE084
respectively representing the state matrix, control matrix, observation matrix and feedforward matrix of the model.
The longitudinal/lateral motion module is configured to extract a longitudinal/lateral motion equation set based on the aircraft all-state linearized model and generate an aircraft longitudinal/lateral motion state linearized model in a state space form;
in the present embodiment, the aircraft longitudinal motion state linearized model is shown in equation (34) (35):
Figure 367082DEST_PATH_IMAGE085
（34）
Figure 216089DEST_PATH_IMAGE086
（35）
wherein the content of the first and second substances,
Figure 872329DEST_PATH_IMAGE087
is the longitudinal motion state quantity of the aircraft,
Figure 388761DEST_PATH_IMAGE088
is the derivative of the longitudinal state of motion quantity of the aircraft,
Figure 354312DEST_PATH_IMAGE150
for the longitudinal motion input, the longitudinal motion output of the model
Figure 690615DEST_PATH_IMAGE089
，
Figure 9601DEST_PATH_IMAGE090
、
Figure 255906DEST_PATH_IMAGE091
、
Figure 267724DEST_PATH_IMAGE092
And
Figure 218887DEST_PATH_IMAGE093
and the state matrix, the control matrix, the observation matrix and the feedforward matrix respectively represent the longitudinal motion state linearization model.
The aircraft lateral motion state linearization model is shown in the formula (36) (37):
Figure 341564DEST_PATH_IMAGE094
（36）
Figure 707955DEST_PATH_IMAGE095
（37）
wherein the content of the first and second substances,
Figure 625095DEST_PATH_IMAGE096
is the lateral motion state quantity of the aircraft,
Figure 935991DEST_PATH_IMAGE097
is the derivative of the lateral state of motion quantity of the aircraft,
Figure 986992DEST_PATH_IMAGE151
for the input of lateral motion, the output of lateral motion of the model
Figure 66944DEST_PATH_IMAGE098
，
Figure 295931DEST_PATH_IMAGE099
、
Figure 94122DEST_PATH_IMAGE100
、
Figure 948815DEST_PATH_IMAGE101
And
Figure 883273DEST_PATH_IMAGE102
and the state matrix, the control matrix, the observation matrix and the feedforward matrix respectively represent the lateral motion state linearization model.
The characteristic analysis and display module is configured to perform aircraft model characteristic analysis based on a pneumatic stability derivative, a full-state linearized model zero-pole, an aircraft longitudinal/lateral motion state linearized model zero-pole, and a corresponding transfer function and stability margin between any one of the manipulation input quantity and the flight state output quantity, and output a model analysis result.
In this embodiment, the aircraft model characteristic analysis is performed based on the aerodynamic stability derivative, the full-state linearized model zero-pole, the aircraft longitudinal/lateral motion state linearized model zero-pole, and the corresponding transfer function and stability margin between any one of the manipulated input quantities and the flight state output quantity, and data and curves of various aerodynamic stability derivative curves, a zero-pole point diagram, a bode diagram, stability margin and other characteristic model characteristics are output, so that the solution and analysis of the aircraft motion model are realized. The derivative of aerodynamic stability comprises an aerodynamic coefficient of one dimension along the three axes of the aircraft body coordinate system
Figure 407795DEST_PATH_IMAGE001
、
Figure 303070DEST_PATH_IMAGE002
、
Figure 836819DEST_PATH_IMAGE003
And aerodynamic moment coefficient of dimension one in three axial directions around the coordinate system of the aircraft body
Figure 15997DEST_PATH_IMAGE004
、
Figure 711420DEST_PATH_IMAGE005
、
Figure 359571DEST_PATH_IMAGE006
。
The analysis result comprises pneumatic stability derivative analysis, aircraft all-state linearized model pole-zero analysis, aircraft longitudinal motion state linearized model pole-zero analysis, aircraft lateral motion state linearized model pole-zero analysis, and corresponding transfer function and stability margin analysis between any one of the control input quantity and the flight state output quantity.
It should be noted that, the aircraft modeling and model characteristic analysis system provided in the foregoing embodiment is only exemplified by the division of the functional modules, and in practical applications, the above functions may be allocated to different functional modules according to needs, that is, the modules or steps in the embodiment of the present invention are further decomposed or combined, for example, the modules in the foregoing embodiment may be combined into one module, or may be further split into multiple sub-modules, so as to complete all or part of the functions described above. The names of the modules and steps involved in the embodiments of the present invention are only for distinguishing the modules or steps, and are not to be construed as unduly limiting the present invention.
An aircraft modeling and model characteristic analysis method according to a second embodiment of the present invention, as shown in fig. 2 and 3, includes the following steps:
step S100, initializing and setting a system; the initialization setting comprises aircraft structure parameter setting, reference motion condition setting and transfer function input/output quantity setting;
step S200, constructing an aerodynamic force and aerodynamic moment model, and calculating aerodynamic force acting on the mass center of the aircraft and aerodynamic moment acting on the mass center of the aircraft;
step S300, combining the aerodynamic force and the aerodynamic moment to construct a six-degree-of-freedom equation set of the aircraft;
step S400, solving a full-dimensional motion state of the aircraft under a reference motion condition;
step S500, linearizing an aircraft motion equation at a reference motion state to generate an aircraft all-state linearization model in a state space form;
s600, extracting a longitudinal/lateral motion equation set based on the aircraft all-state linearized model, and generating an aircraft longitudinal/lateral motion state linearized model in a state space form;
step S700, performing aircraft model characteristic analysis based on a pneumatic stability derivative, a full-state linearized model zero-pole, an aircraft longitudinal/lateral motion state linearized model zero-pole, a corresponding transfer function and a stability margin between any one of the control input quantity and the flight state output quantity, and outputting a model analysis result;
step S800, whether aircraft modeling and model characteristic analysis are finished or not is evaluated: if the task is completed, executing step S900, otherwise, executing step S100;
and S900, storing the aircraft state information and the model characteristic analysis result to the system.
It can be clearly understood by those skilled in the art that, for convenience and brevity of description, the specific working process and related description of the method described above may refer to the corresponding process in the foregoing system embodiment, and are not described herein again.
A storage device according to a third embodiment of the invention stores a plurality of programs adapted to be loaded by a processor and to implement the aircraft modeling and model characterization method described above.
A processing apparatus according to a fourth embodiment of the present invention includes a processor, a storage device; a processor adapted to execute various programs; a storage device adapted to store a plurality of programs; the program is adapted to be loaded and executed by a processor to implement the aircraft modeling and model characterization methods described above.
It can be clearly understood by those skilled in the art that, for convenience and brevity of description, the specific working processes and related descriptions of the storage device and the processing device described above may refer to the corresponding processes in the foregoing method examples, and are not described herein again.
Referring now to FIG. 4, there is illustrated a block diagram of a computer system suitable for use as a server in implementing embodiments of the method, system, and apparatus of the present application. The server shown in fig. 4 is only an example, and should not bring any limitation to the functions and the scope of use of the embodiments of the present application.
As shown in fig. 4, the computer system includes a Central Processing Unit (CPU)401 that can perform various appropriate actions and processes according to a program stored in a Read Only Memory (ROM) 402 or a program loaded from a storage section 408 into a Random Access Memory (RAM) 403. In the RAM 403, various programs and data necessary for system operation are also stored. The CPU 401, ROM 402, and RAM 403 are connected to each other via a bus 404. An Input/Output (I/O) interface 405 is also connected to the bus 404.
The following components are connected to the I/O interface 405: an input portion 306 including a keyboard, a mouse, and the like; an output section 407 including a Display such as a Cathode Ray Tube (CRT), a Liquid Crystal Display (LCD), and a speaker; a storage section 408 including a hard disk and the like; and a communication section 409 including a Network interface card such as a LAN (Local Area Network) card, a modem, or the like. The communication section 409 performs communication processing via a network such as the internet. A driver 410 is also connected to the I/O interface 405 as needed. A removable medium 411 such as a magnetic disk, an optical disk, a magneto-optical disk, a semiconductor memory, or the like is mounted on the drive 410 as necessary, so that a computer program read out therefrom is mounted into the storage section 408 as necessary.
In particular, according to an embodiment of the present disclosure, the processes described above with reference to the flowcharts may be implemented as computer software programs. For example, embodiments of the present disclosure include a computer program product comprising a computer program embodied on a computer readable medium, the computer program comprising program code for performing the method illustrated in the flow chart. In such an embodiment, the computer program may be downloaded and installed from a network through the communication section 409, and/or installed from the removable medium 411. The computer program performs the above-described functions defined in the method of the present application when executed by a Central Processing Unit (CPU) 401. It should be noted that the computer readable medium mentioned above in the present application may be a computer readable signal medium or a computer readable storage medium or any combination of the two. A computer readable storage medium may be, for example, but not limited to, an electronic, magnetic, optical, electromagnetic, infrared, or semiconductor system, apparatus, or device, or any combination of the foregoing. More specific examples of the computer readable storage medium may include, but are not limited to: an electrical connection having one or more wires, a portable computer diskette, a hard disk, a Random Access Memory (RAM), a read-only memory (ROM), an erasable programmable read-only memory (EPROM or flash memory), an optical fiber, a portable compact disc read-only memory (CD-ROM), an optical storage device, a magnetic storage device, or any suitable combination of the foregoing. In the present application, a computer readable storage medium may be any tangible medium that can contain, or store a program for use by or in connection with an instruction execution system, apparatus, or device. In this application, however, a computer readable signal medium may include a propagated data signal with computer readable program code embodied therein, for example, in baseband or as part of a carrier wave. Such a propagated data signal may take many forms, including, but not limited to, electro-magnetic, optical, or any suitable combination thereof. A computer readable signal medium may also be any computer readable medium that is not a computer readable storage medium and that can communicate, propagate, or transport a program for use by or in connection with an instruction execution system, apparatus, or device. Program code embodied on a computer readable medium may be transmitted using any appropriate medium, including but not limited to: wireless, wire, fiber optic cable, RF, etc., or any suitable combination of the foregoing.
Computer program code for carrying out operations for aspects of the present application may be written in any combination of one or more programming languages, including an object oriented programming language such as Java, Smalltalk, C + + or the like and conventional procedural programming languages, such as the "C" programming language or similar programming languages. The program code may execute entirely on the user's computer, partly on the user's computer, as a stand-alone software package, partly on the user's computer and partly on a remote computer or entirely on the remote computer or server. In the case of a remote computer, the remote computer may be connected to the user's computer through any type of network, including a Local Area Network (LAN) or a Wide Area Network (WAN), or the connection may be made to an external computer (for example, through the Internet using an Internet service provider).
The flowchart and block diagrams in the figures illustrate the architecture, functionality, and operation of possible implementations of systems, methods and computer program products according to various embodiments of the present application. In this regard, each block in the flowchart or block diagrams may represent a module, segment, or portion of code, which comprises one or more executable instructions for implementing the specified logical function(s). It should also be noted that, in some alternative implementations, the functions noted in the block may occur out of the order noted in the figures. For example, two blocks shown in succession may, in fact, be executed substantially concurrently, or the blocks may sometimes be executed in the reverse order, depending upon the functionality involved. It will also be noted that each block of the block diagrams and/or flowchart illustration, and combinations of blocks in the block diagrams and/or flowchart illustration, can be implemented by special purpose hardware-based systems which perform the specified functions or acts, or combinations of special purpose hardware and computer instructions.
The terms "first," "second," and the like are used for distinguishing between similar elements and not necessarily for describing or implying a particular order or sequence.
The terms "comprises," "comprising," or any other similar term 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.
So far, the technical solutions of the present invention have been described in connection with the preferred embodiments shown in the drawings, but it is easily understood by those skilled in the art that the scope of the present invention is obviously not limited to these specific embodiments. Equivalent changes or substitutions of related technical features can be made by those skilled in the art without departing from the principle of the invention, and the technical scheme after the changes or substitutions can fall into the protection scope of the invention.

Claims (8)

1. An aircraft modeling and model characterization system, the system comprising: the system comprises an aerodynamic module, an aerodynamic moment module, a six-degree-of-freedom motion equation module, a reference motion state solving module, a motion equation linearization module, a longitudinal/lateral motion module, a characteristic analysis and display module and a system management module;
the system management module is configured to perform initialization setting on the system; further configured to store aircraft state information and model characteristics; the initialization setting comprises aircraft structure parameter setting, reference motion condition setting and transfer function input/output quantity setting;
the aerodynamic module is configured to calculate aerodynamic forces acting on the center of mass of the aircraft;
the aerodynamic moment module is configured to calculate an aerodynamic moment acting around a center of mass of the aircraft;
the six-degree-of-freedom motion equation module is configured to combine the aerodynamic force and the aerodynamic moment to construct an aircraft full-dimensional motion state equation set; the full-dimensional motion state equation set comprises a moment equation set, an angular displacement equation set, a force equation set, a linear displacement equation set, an actuating mechanism model and an atmospheric model;
the reference motion state solving module is configured to solve the full-dimensional motion state of the aircraft under the reference motion condition;
the motion equation linearization module is configured to linearize the aircraft motion equation at the reference motion state to generate an aircraft all-state linearization model in a state space form;
the aircraft full-state linearization model is as follows:
Figure 735529DEST_PATH_IMAGE002
Figure 616897DEST_PATH_IMAGE004
wherein the content of the first and second substances,
Figure DEST_PATH_IMAGE005
is the full state quantity of the aircraft,
Figure 972792DEST_PATH_IMAGE006
、
Figure DEST_PATH_IMAGE007
、
Figure 233003DEST_PATH_IMAGE008
representing the longitudinal displacement, lateral displacement and altitude of the aircraft in a ground coordinate system,
Figure DEST_PATH_IMAGE009
、
Figure 750572DEST_PATH_IMAGE010
、
Figure DEST_PATH_IMAGE011
representing the roll, pitch and yaw of the aircraft,
Figure 853657DEST_PATH_IMAGE012
the speed of flight is indicated as a function of,
Figure DEST_PATH_IMAGE013
and
Figure 731352DEST_PATH_IMAGE014
respectively a flight attack angle and a sideslip angle,
Figure DEST_PATH_IMAGE015
、
Figure 360917DEST_PATH_IMAGE016
、
Figure DEST_PATH_IMAGE017
indicating flightRoll rate, pitch rate and yaw rate of the machine in a body coordinate system,
Figure 800120DEST_PATH_IMAGE018
the derivative of the full state quantity of the aircraft is represented,
Figure DEST_PATH_IMAGE019
as input quantities, output quantities of the model
Figure 515135DEST_PATH_IMAGE020
In the state quantity
Figure DEST_PATH_IMAGE021
On the basis of the three axial directions of the aircraft
Figure 88198DEST_PATH_IMAGE022
Acceleration, flight mach number and dynamic pressure value of the aircraft,
Figure DEST_PATH_IMAGE023
，
Figure 290379DEST_PATH_IMAGE024
respectively represent the flight mach number and dynamic pressure value,
Figure DEST_PATH_IMAGE025
、
Figure 415329DEST_PATH_IMAGE026
、
Figure DEST_PATH_IMAGE027
and
Figure 102794DEST_PATH_IMAGE028
respectively representing a state matrix, a control matrix, an observation matrix and a feedforward matrix of the model;
the longitudinal/lateral motion module is configured to extract a longitudinal/lateral motion equation set based on the aircraft all-state linearized model and generate an aircraft longitudinal/lateral motion state linearized model in a state space form;
the characteristic analysis and display module is configured to perform aircraft model characteristic analysis based on a pneumatic stability derivative, a full-state linearized model zero-pole, an aircraft longitudinal/lateral motion state linearized model zero-pole, and a corresponding transfer function and stability margin between any one of the control input quantity and the flight state output quantity, and output a model analysis result;
the derivative of aerodynamic stability comprises an aerodynamic coefficient of one dimension along the three axes of the aircraft body coordinate system
Figure DEST_PATH_IMAGE029
、
Figure 338603DEST_PATH_IMAGE030
、
Figure DEST_PATH_IMAGE031
And aerodynamic moment coefficient of dimension one in three axial directions around the coordinate system of the aircraft body
Figure 926448DEST_PATH_IMAGE032
、
Figure DEST_PATH_IMAGE033
、
Figure 97666DEST_PATH_IMAGE034
。
2. The system for modeling and modeling aircraft according to claim 1, wherein said aerodynamic module calculates aerodynamic forces acting on the center of mass of the aircraft by: based on the real-time state of the aircraft and aerodynamic force data, solving the aerodynamic force of the aircraft at the current moment through an interpolation algorithm; the method comprises the following specific steps:
Figure 787274DEST_PATH_IMAGE036
Figure 967719DEST_PATH_IMAGE038
Figure 442694DEST_PATH_IMAGE040
Figure 50393DEST_PATH_IMAGE042
Figure 368242DEST_PATH_IMAGE044
Figure 211433DEST_PATH_IMAGE046
wherein the content of the first and second substances,
Figure 931127DEST_PATH_IMAGE029
、
Figure 101207DEST_PATH_IMAGE030
、
Figure 640773DEST_PATH_IMAGE031
respectively are aerodynamic coefficients with the dimension of one along the three axial directions of the aircraft body coordinate system,
Figure 428600DEST_PATH_IMAGE021
、
Figure 861855DEST_PATH_IMAGE020
、
Figure DEST_PATH_IMAGE047
representing the aerodynamic force component in the aircraft body coordinate system,
Figure 421144DEST_PATH_IMAGE048
in order to generate a dynamic pressure,
Figure DEST_PATH_IMAGE049
for the purpose of reference area, the area of the reference,
Figure 916847DEST_PATH_IMAGE050
、
Figure DEST_PATH_IMAGE051
and
Figure 101841DEST_PATH_IMAGE052
respectively an aileron, an elevator and a rudder deflection of the aircraft,
Figure 638870DEST_PATH_IMAGE013
and
Figure 24852DEST_PATH_IMAGE014
respectively a flight attack angle and a sideslip angle,
Figure DEST_PATH_IMAGE053
caused by deflection of elevators
Figure 398065DEST_PATH_IMAGE021
The coefficient of variation of the axial aerodynamic force,
Figure 262116DEST_PATH_IMAGE054
generated when pitch rate is not zero
Figure 545330DEST_PATH_IMAGE021
Axial dynamic gasThe coefficient of the power variation is,
Figure DEST_PATH_IMAGE055
caused by deflection of ailerons and rudder
Figure 712000DEST_PATH_IMAGE020
The coefficient of variation of the axial aerodynamic force,
Figure 713454DEST_PATH_IMAGE056
generated when the yaw rate is not zero
Figure 505829DEST_PATH_IMAGE020
The coefficient of variation of the axial dynamic aerodynamic force,
Figure DEST_PATH_IMAGE057
generated when the roll rate is not zero
Figure 377970DEST_PATH_IMAGE020
The coefficient of variation of the axial dynamic aerodynamic force,
Figure 479656DEST_PATH_IMAGE058
caused by deflection of elevators
Figure 702827DEST_PATH_IMAGE047
The coefficient of variation of the axial aerodynamic force,
Figure DEST_PATH_IMAGE059
generated when pitch rate is not zero
Figure 767735DEST_PATH_IMAGE047
The coefficient of variation of the axial dynamic aerodynamic force,
Figure 635328DEST_PATH_IMAGE060
is a reference length.
3. The system for modeling and modeling aircraft as claimed in claim 2, wherein said aerodynamic moment module calculates the aerodynamic moment about the aircraft center of mass by: solving the aerodynamic moment of the aircraft at the current moment by an interpolation algorithm based on the real-time state of the aircraft and the aerodynamic moment data; the method comprises the following specific steps:
Figure 799593DEST_PATH_IMAGE062
Figure 244481DEST_PATH_IMAGE064
Figure 378659DEST_PATH_IMAGE066
Figure 490972DEST_PATH_IMAGE068
Figure 668881DEST_PATH_IMAGE070
Figure 132223DEST_PATH_IMAGE072
wherein the content of the first and second substances,
Figure DEST_PATH_IMAGE073
、
Figure 273355DEST_PATH_IMAGE074
and
Figure DEST_PATH_IMAGE075
representing the aerodynamic moment components in three axes around the aircraft body coordinate system,
Figure 849961DEST_PATH_IMAGE060
for the purpose of reference to the length of the strip,
Figure 90449DEST_PATH_IMAGE032
、
Figure 900142DEST_PATH_IMAGE033
、
Figure 251489DEST_PATH_IMAGE034
respectively are aerodynamic moment coefficients with the dimension of one around the three axial directions of the coordinate system of the aircraft body,
Figure 338394DEST_PATH_IMAGE076
caused by deflection of ailerons and rudder
Figure 129544DEST_PATH_IMAGE021
The coefficient of variation of the axial aerodynamic moment,
Figure DEST_PATH_IMAGE077
generated when the yaw rate is not zero
Figure 895375DEST_PATH_IMAGE021
The axial dynamic pneumatic moment variation coefficient,
Figure 50413DEST_PATH_IMAGE078
generated when the roll rate is not zero
Figure 991824DEST_PATH_IMAGE021
The axial dynamic pneumatic moment variation coefficient,
Figure DEST_PATH_IMAGE079
for elevatorsCaused by deflection
Figure 652744DEST_PATH_IMAGE020
The coefficient of variation of the axial aerodynamic moment,
Figure 171450DEST_PATH_IMAGE080
generated when pitch rate is not zero
Figure 130178DEST_PATH_IMAGE020
The axial dynamic pneumatic moment variation coefficient,
Figure DEST_PATH_IMAGE081
caused by deflection of ailerons and rudder
Figure 768839DEST_PATH_IMAGE047
The coefficient of variation of the axial aerodynamic moment,
Figure 646665DEST_PATH_IMAGE082
generated when the yaw rate is not zero
Figure 528034DEST_PATH_IMAGE047
The axial dynamic pneumatic moment variation coefficient,
Figure DEST_PATH_IMAGE083
generated when the roll rate is not zero
Figure 369082DEST_PATH_IMAGE047
And (4) axial dynamic aerodynamic moment variation coefficient.
4. The system for modeling and analyzing model characteristics of an aircraft according to claim 1, wherein said module for solving the reference motion state "solves the full-dimensional motion state of the aircraft under the reference motion condition" comprises:
carrying out square weighted sum on the derivative values of the full-dimensional motion state of the aircraft to construct a cost function;
and solving the full-dimensional motion state of the aircraft with the minimum cost function based on a preset reference motion condition.
5. The aircraft modeling and model characterization system of claim 4, wherein said reference motion conditions are:
Figure 19506DEST_PATH_IMAGE084
，
Figure DEST_PATH_IMAGE085
，
Figure 802654DEST_PATH_IMAGE086
wherein the content of the first and second substances,
Figure DEST_PATH_IMAGE087
and
Figure 14062DEST_PATH_IMAGE088
respectively the set flight speed value and the flight altitude value,
Figure 173647DEST_PATH_IMAGE014
、
Figure 678578DEST_PATH_IMAGE009
representing the sideslip angle, roll angle of the aircraft,
Figure 648939DEST_PATH_IMAGE015
、
Figure 770479DEST_PATH_IMAGE016
、
Figure 937018DEST_PATH_IMAGE017
indicating aircraft is seated on the bodyRoll rate, pitch rate and yaw rate in the system,
Figure 562035DEST_PATH_IMAGE012
the speed of flight is indicated as a function of,
Figure 467411DEST_PATH_IMAGE008
representing the altitude of the aircraft in the ground coordinate system,
Figure DEST_PATH_IMAGE089
the derivative of the flight speed is represented as,
Figure 138564DEST_PATH_IMAGE090
representing the angle of attack derivative and the sideslip angle derivative of the aircraft,
Figure DEST_PATH_IMAGE091
the roll rate derivative, pitch rate derivative, and yaw rate derivative of the aircraft are represented.
6. The aircraft modeling and model characteristic analysis system of claim 5, wherein said cost function model is:
Figure DEST_PATH_IMAGE093
Figure DEST_PATH_IMAGE095
wherein the content of the first and second substances,
Figure 328368DEST_PATH_IMAGE096
in order to be the weight coefficient,
Figure DEST_PATH_IMAGE097
the derivative of the flying height is indicated.
7. The aircraft modeling and model characteristic analysis system of claim 6 wherein said aircraft longitudinal motion state linearization model is:
Figure DEST_PATH_IMAGE099
Figure DEST_PATH_IMAGE101
wherein the content of the first and second substances,
Figure 113616DEST_PATH_IMAGE102
is the longitudinal motion state quantity of the aircraft,
Figure DEST_PATH_IMAGE103
is the derivative of the longitudinal state of motion quantity of the aircraft,
Figure 143889DEST_PATH_IMAGE104
for the longitudinal motion input, the longitudinal motion output of the model
Figure DEST_PATH_IMAGE105
，
Figure 286026DEST_PATH_IMAGE106
、
Figure DEST_PATH_IMAGE107
、
Figure 794368DEST_PATH_IMAGE108
And
Figure DEST_PATH_IMAGE109
state matrix, control matrix, observation matrix and feedforward matrix respectively representing state linearized model of longitudinal motion。
8. The aircraft modeling and model characteristic analysis system of claim 6 wherein said aircraft lateral motion state linearization model is:
Figure DEST_PATH_IMAGE111
Figure DEST_PATH_IMAGE113
wherein the content of the first and second substances,
Figure 66081DEST_PATH_IMAGE114
is the lateral motion state quantity of the aircraft,
Figure DEST_PATH_IMAGE115
is the derivative of the lateral state of motion quantity of the aircraft,
Figure 782102DEST_PATH_IMAGE116
for the input of lateral motion, the output of lateral motion of the model
Figure DEST_PATH_IMAGE117
，
Figure 427846DEST_PATH_IMAGE118
、
Figure DEST_PATH_IMAGE119
、
Figure 490612DEST_PATH_IMAGE120
And
Figure DEST_PATH_IMAGE121
state matrix, control, respectively representing a linearized model of the state of lateral motionA matrix, an observation matrix, and a feed forward matrix.
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