CN106096194A - Fixed Wing AirVehicle rapid modeling method for designing based on plug type interface - Google Patents

Abstract

The present invention proposes fixed wing airplane aerodynamic parameter edit specification based on plug type interface, designs plug type interface, is peeled off by aerodynamic parameter and abstract, it is achieved aerodynamic parameter plug and play from kinetic model.Use VC2010 developing instrument, complete the manipulation input of fixed wing airplane, kinetics equation and kinematical equation Development of Module.Design aerodynamic parameter XML file analysis program, combines the aerodynamic parameter of acquisition with dynamics module, it is achieved the quick exploitation of flying quality phantom.User is by replacing the aerodynamic parameter of different aircraft in XML file, on the premise of engine type and model are fixing, only need to revise the aerodynamic parameter in XML file and just can complete simulation model design during to different Modeling of Vehicle.Can quickly set up the Aerodynamics Model for different fixed wing airplane types, shorten the model development cycle.Reduce the repeated labor in aircraft process of mathematical modeling, improve the motility of Modeling of Vehicle, rapidity, and versatility.

Description

Fixed Wing AirVehicle rapid modeling method for designing based on plug type interface

Technical field

Present invention relates particularly to a kind of Fixed Wing AirVehicle rapid modeling method for designing based on plug type interface, belong to Computer aircraft performance emulation field.

Background technology

XML is extensible markup language or extensible, is a kind of markup language.The innovative point of the present invention is that application can The aerodynamic parameter of aircraft, due to the XML document of aerodynamic parameter, is detached out from Aviate equation and writes by extending mark language XML document, is resolved in Aviate equation can be carried out aircraft emulating when again from document by aerodynamic parameter Performance simulation.Use XML document labelling aerodynamic parameter so that flight side can be shared when different aircraft are emulated Journey, it is only necessary to the aerodynamic parameter in amendment document, so can improve development efficiency, it is achieved the multiplexing of code.XML language has There is the succinct feature effectively, efficiently expanded, and use XML can exchange information between different computer systems.XML supports Multiplexing document snippet, user can invent and use the label of oneself, and scalability is big, can effectively carry out XML file Expansion, aircraft platforms difference when can by amendment expand XML file be rapidly performed by aerial vehicle simulation model.

Summary of the invention

The present invention proposes a kind of Fixed Wing AirVehicle rapid modeling method for designing based on plug type interface, proposes base In the fixed wing airplane aerodynamic parameter edit specification of XML, design plug type interface, aerodynamic parameter is shelled from kinetic model From also abstract, it is achieved aerodynamic parameter plug and play.Use VC2010 developing instrument, complete fixed wing airplane manipulation input, Kinetics equation and kinematical equation Development of Module.Design aerodynamic parameter XML file analysis program, by same for the aerodynamic parameter obtained Dynamics module combines, it is achieved the quick exploitation of flying quality phantom.User flies by replacing difference in XML file The aerodynamic parameter of row device, on the premise of engine type and model are fixing, only need to revise XML during to different Modeling of Vehicle Aerodynamic parameter in file just can complete simulation model design.Can quickly set up the air for different fixed wing airplane types to move Mechanical model, shortens the model development cycle.Reduce the repeated labor in aircraft process of mathematical modeling, improve aircraft The motility of modeling, rapidity, and versatility.

A kind of Fixed Wing AirVehicle rapid modeling method for designing based on plug type interface, mainly comprises herein below:

Step one: definition XML document, peels off the aerodynamic parameter of aircraft from Aviate equation and writes in file, real Existing Aviate equation therefrom reads aerodynamic parameter when initializing；

Step 2: according to the Kinematic Decomposition of aircraft, is moved and is divided into horizontal lateral movement and lengthwise movement, write respectively Kinestate resolves equation；

Step 3: design XML analysis program, is loaded into aerodynamic parameter in Aerodynamics Model, it is achieved kinestate Resolve the initialization of equation；

Step 4: under conditions of engine type and model are fixing, can fly so that multiplexing is same for different aircraft Row equation, only need to revise the aerodynamic parameter in XML file and the most again be resolved in kinetic model；

Defined in described step one, the realization approach of XML document is:

(1). first created XML document, definition document root node by xml editor；

(2). definition ground floor node, totally 8 nodes, labelling ground floor nodes, lift resolve relevant pneumatic ginseng respectively Number, side force resolve aerodynamic parameter, pitching moment resolves aerodynamic parameter, rolling moment resolves aerodynamic parameter, yawing resolves gas Dynamic parameter, damping force resolve aerodynamic parameter, angular acceleration resolves aerodynamic parameter；

(3). for each node creation node of ground floor, create child node according to the aerodynamic parameter number of corresponding node, Relate to a how many aerodynamic parameter and be created that several child node, wherein the child node number of first sub-vertex ticks present node；

(4). one aerodynamic parameter of a sub-node on behalf, is series of discrete point owing to being stored in the aerodynamic parameter of document, because of This needs to create a series of nexine nodes for child node.Each child node extends 21 nexine nodes, wherein first nexine Vertex ticks aerodynamic parameter discrete point number, remaining nexine vertex ticks aerodynamic parameter centrifugal pump；

(5). Save and Close XML document, it is achieved the labelling of aerodynamic parameter；

In described step 3, XML document parsing realization approach is:

(1). load XML document, proceed by aerodynamic parameter and resolve；

(2). text pointer is navigated to root node, obtains the nodes of ground floor node；

(3). text pointer is navigated to successively each node of ground floor, begins stepping through the child node of each node；

(4). the nexine node of traversal child node extension, by the aerodynamic parameter value of labelling in nexine node, read corresponding gas In aerodynamic parameter array corresponding in dynamic resolving module；

(5). repeat aforesaid operations, until the nexine node in the child node of all nodes all reads；

(6). Save and Close XML document, start-up parameter is loaded in calculator memory；

Aerodynamic force, moment that described ground floor node on behalf is corresponding resolve module；The child node of node represents this pneumatic solution Calculate the aerodynamic parameter that module relates to；The dimensional table data amount check of this aerodynamic parameter labelling of nexine node on behalf of child node；

Need write XML file aerodynamic parameter table:

For just understanding, the XML document tag format of attached following airfoil lift coefficient.

Implement:

1 defined variable:

Variable module includes that abstract variable that each aerodynamic parameter of aircraft is corresponding, kinestate variable, mechanical state become Amount and control input.Wherein aerodynamic parameter variable is for the resolving of aircraft aerodynamic model；Kinestate variable is used for Attitude information that record cast calculates, positional information etc.；Pitching moment that mechanical state variable calculates for record cast, Rolling moment, yawing, lateral moment and the status information such as resistance, lift；Control input and include [δ_T δ_e δ_a δ_r], right Answer the input of throttle push rod, elevator drift angle, aileron movement angle, rudder.

2 definition XML document and XML analysis programs:

The aerodynamic parameter of vehicle aerodynamics model is write in XML file.By gas from Aerodynamics Model Dynamic parameter is stripped out and abstract, replaces with variable in each resolving module, by pneumatic by file of XML analysis program Parameter analysis of electrochemical is in corresponding variable, it is achieved the initialization of Aerodynamics Model.The definition of XML document is by the gas in table 1 Dynamic parameter recorded in document with tree-like hierarchy structure successively according to resolving equation, and all aerodynamic parameters of such aircraft are with regard to structure Having become to have the tree of multilamellar bifurcation structure, each aerodynamic parameter corresponds to an element of one of them branch.Analysis program Effect be exactly parsing tree, by carry out tree traversal just each element can be operated, be resolved to aircraft flight In equation, thus Aviate equation is embodied.The method creating XML document is shown in Figure of description 1, Fig. 2, XML document parsing side Method is shown in Figure of description 3.

3 states resolving modules:

State resolves module and includes the power equation group of aircraft, movement difference equations, momental equation group, navigation equation group four Point, write when state resolves module and corresponding equation group discretization is converted into computer program.Determining state vectorWith control input [δ_T δ_e δ_a δ_rRelation between] After, at the relevant aerodynamic parameter of known aircraft, characteristic parameter, very according to flying height h, Mach number M_aAnd state of flight, just May determine that power (F_x F_y F_z) and momentApplication state resolves module, and in office when just can solve aircraft The kinestate carved.Wherein (u, v w) are three velocity components of body axis system；For attitude angle, respectively rolling Angle, the angle of pitch, course angle；(p, q, r) be body angular velocity component, respectively body angular velocity in roll, body rate of pitch And body yaw rate；(x_g,y_g,z_g) it is geographical coordinates；δ_T,δ_e,δ_a,δ_rBe respectively accelerator open degree input, elevator inclined Input, aileron rudder partially inputs and course rudder inputs partially.

1). according to controlling input [δ_T δ_e δ_a δ_r] resolve power, moment.

Power mainly includes lift L, side force Y, motor power T.Concrete resolving is as follows:

Lift resolving equation:

Wherein have

Q: dynamic pressure；

ρ: current gas pressure level air density；

V: aircraft horizontal flight speed；

C_LW: airfoil lift coefficient；

C_Lb: fuselage lift coefficient；

S_b: fuselage cross-section is amassed；

S_W: wing area；

C_Lt: horizontal tail lift coefficient；

S_t: horizontal tail area；

Side force resolving equation:

ρ: current gas pressure level air density；

V: aircraft horizontal flight speed；

S_W: wing area；

C_Yβ: the side force derivative that yaw angle causes；

β: yaw angle；

: rudder side force derivative；

δ_r: rudder inputs；

: angular velocity in roll side force derivative；

: withFor the angular velocity in roll of dimension, b is wing length；

: yaw rate side force derivative；

: withFor the yaw rate of dimension, b is wing length；

Pitching moment resolving equation:

$M_A=\left(\frac{1}{2}{\mathrm{\ρV}}^2\right)(C_{m,\mathrm{\α}=0}+C_{m\mathrm{\α}}\mathrm{\α}+C_{{\mathrm{m\δ}}_e}{\mathrm{\δ}}_e+C_{m\stackrel{\‾}{q}}\stackrel{\‾}{q}+C_{m\stackrel{\‾}{\stackrel{\·}{\mathrm{\α}}}}\stackrel{\‾}{\stackrel{\·}{\mathrm{\α}}}+C_{m{\stackrel{\‾}{\stackrel{\·}{\mathrm{\δ}}}}_e}{\stackrel{\‾}{\stackrel{\·}{\mathrm{\δ}}}}_e)---\left(3\right)$

ρ: current gas pressure level air density；

V: aircraft horizontal flight speed；

C_{M, α=0}: static indeterminacy pitching moment；

C_mα: angle of attack pitching moment derivative；

α: the angle of attack；

: elevator pitching moment derivative；

δ_e: elevator drift angle；

: the additional pitching moment coefficient of horizontal tail；

: withFor the rate of pitch of dimension, b is wing length；

: wash time difference damping torque derivative under horizontal tail；

: withAngle of attack speed for dimension

: elevator drift angle speed pitching moment derivative；

: withElevator drift angle speed for dimension；

c_A: wing mean geometric of airfoil；

Rolling moment resolving equation:

C_lβ: roll static-stability derivative；

V: aircraft horizontal flight speed；

B:b is wing length；

ρ: current gas pressure level air density；

: roll guidance derivative；

δ_a: aileron movement angle；

: rudder control cross derivative；

: roll damping derivative；

: withFor the angular velocity in roll of dimension, b is wing length；

: intersection dynamic derivative；

: withFor the yaw rate of dimension, b is wing length；

Yawing resolving equation:

ρ: current gas pressure level air density；

V: aircraft horizontal flight speed；

C_nβ: driftage static-stability derivative；

: aileron control cross derivative；

δ_a: aileron movement angle；

: directional control derivative；

δ_r: course angle of rudder reflection；

: intersection dynamic derivative；

: withFor the angular velocity in roll of dimension, b is wing length；

: course damping derivative；

: withFor the yaw rate of dimension, b is wing length；

Damping force resolves equation: D_k=C_DV_k (6)

C_D: damped coefficient；

Aircraft can be calculated along body axis system force vector by equation (1)～(6):

$\left[\begin{array}{c}{F}_{kx}\\ F_{ky}\\ F_{kz}\end{array}\right]=\left[\begin{array}{c}{T}_k\\ 0\\ 0\end{array}\right]+\left[\begin{array}{c}0\\ Y_k\\ -L_k\end{array}\right]+\left[\begin{array}{c}-D_k-m_k{g}_k\sin \mathrm{\θ}\\ m_k{g}_k\cos \mathrm{\θ}\sin \mathrm{\φ}\\ m_k{g}_k\cos \mathrm{\θ}\cos \mathrm{\φ}\end{array}\right]$

Wherein k is that kth resolves the cycle；D_kFor flight resistance；m_kThe quality of cycle aircraft is resolved for kth；g_kFor kth The acceleration of gravity in individual resolving cycle.

(7) moment vector can be decomposed along body axis system by calculating aircraft by equation (1)～(6):

$\left[\begin{array}{c}{L}_k\\ M_k\\ N_k\end{array}\right]=\left[\begin{array}{c}0\\ T_k{l}_z\\ -T_k{l}_y\end{array}\right]+\left[\begin{array}{c}{\stackrel{\‾}{L}}_{kA}\\ M_{kA}\\ N_{kA}\end{array}\right]---\left(8\right)$

WhereinM_kA, N_kAIt is respectively kth and resolves cycle roll guidance moment, pitch control moment, yaw control power Square；T_kThe motor power in cycle is resolved for kth；l_y,l_zIt is respectively the rotary inertia of y-axis, the rotary inertia of z-axis；L_k,M_k, N_kIt is respectively rolling moment, pitching moment, yawing.

2). by [F_kx F_ky F_kz]^T、[u_k ν_k ω_k]^T、[φ_k θ_k ψ_k]^T、[p_k q_k r_k]^TCalculate current state to accelerate Degree component.

Component of acceleration resolving equation:

Current state component of acceleration is calculated by equation (9)

3). by [L_k M_k N_k]^T、[p_k q_k r_k]^TCalculate current state component of angular acceleration.

Component of angular acceleration resolving equation:

Wherein: I_xzFor the product of inertia；

$c_1=\frac{(I_y-I_z)I_z-I_{xz}^2}{I_x{I}_z-I_{xz}^2},c_2=\frac{(I_x-I_y+I_z)I_{xz}}{I_x{I}_z-I_{xz}^2},c_3=\frac{I_z}{I_x{I}_z-I_{xz}^2};$

$c_4=\frac{I_{xz}}{I_x{I}_z-I_{xz}^2},c_5=\frac{I_z-I_x}{I_y},c_6=\frac{I_{xz}}{I_y};$

$c_7=\frac{1}{I_y},c_8=\frac{(I_x-I_y)I_x+I_{xz}^2}{I_x{I}_z-I_{xz}^2},c_9=\frac{I_x}{I_x{I}_z-I_{xz}^2};$

Current state component of angular acceleration is calculated by equation (10)

4). by [u_k ν_k ω_k]^T、Calculate the velocity component [u of NextState_k+1 ν_k+1 ω_k+1]^T

The velocity component resolving equation of NextState:

Wherein Δ τ is the resolving cycle.

5). by [p_k q_k r_k]^T、Calculate the angular velocity component [p of NextState_k+1 q_k+1 r_k+1]^T

The angular velocity component resolving equation of NextState:

$\left[\begin{array}{c}{p}_{k+1}\\ q_{k+1}\\ r_{k+1}\end{array}\right]=\left[\begin{array}{c}{p}_k+{\stackrel{\·}{p}}_k\mathrm{\Δ}\mathrm{\τ}\\ q_k+{\stackrel{\·}{q}}_k\mathrm{\Δ}\mathrm{\τ}\\ r_k+{\stackrel{\·}{r}}_k\mathrm{\Δ}\mathrm{\τ}\end{array}\right]---\left(12\right)$

Wherein Δ τ is the resolving cycle.

6). by [φ_k θ_k ψ_k]^T、[p_k+1 q_k+1 r_k+1]^TCalculate current state attitude angular rate

Current state attitude angular rate resolving equation:

$\left\{\begin{array}{c}\stackrel{\·}{\mathrm{\φ}}=p+(rcos\mathrm{\φ}+q\sin \mathrm{\φ})tan\mathrm{\θ}\\ \stackrel{\·}{\mathrm{\θ}}=q\cos \mathrm{\φ}-rsin\mathrm{\φ}\\ \stackrel{\·}{\mathrm{\ψ}}=\frac{1}{\cos \mathrm{\θ}}(rcos\mathrm{\φ}+q\sin \mathrm{\φ})\end{array}\right.---\left(13\right)$

Current state attitude angular rate can be calculated by equation (13)

7). by[φ_k θ_k ψ_k]^TCalculate the attitude angle [φ of NextState_k+1 θ_k+1 ψ_k+1]^T

NextState solving of attitude equation:

$\left[\begin{array}{c}{\mathrm{\φ}}_{k+1}\\ {\mathrm{\θ}}_{k+1}\\ {\mathrm{\ψ}}_{k+1}\end{array}\right]=\left[\begin{array}{c}{\mathrm{\φ}}_k+{\stackrel{\·}{\mathrm{\φ}}}_k\mathrm{\Δ}\mathrm{\τ}\\ {\mathrm{\θ}}_k+{\stackrel{\·}{\mathrm{\θ}}}_k\mathrm{\Δ}\mathrm{\τ}\\ {\mathrm{\ψ}}_k+{\stackrel{\·}{\mathrm{\ψ}}}_k\mathrm{\Δ}\mathrm{\τ}\end{array}\right]---\left(14\right)$

Wherein Δ τ is the resolving cycle.

8). by [u_k+1 ν_k+1 ω_k+1]^T、[φ_k+1 θ_k+1 ψ_k+1]^TCalculate the location status information change rate of aircraft

Location status information change rate resolving equation:

$\left\{\begin{array}{c}{\stackrel{\·}{x}}_g=u\cos \mathrm{\θ}\cos \mathrm{\ψ}+v(\sin \mathrm{\φ}\sin \mathrm{\θ}\cos \mathrm{\ψ}-\cos \mathrm{\φ}\sin \mathrm{\ψ})+\mathrm{\ω}(\sin \mathrm{\φ}\sin \mathrm{\ψ}+\cos \mathrm{\φ}\sin \mathrm{\θ}\cos \mathrm{\ψ})\\ {\stackrel{\·}{y}}_g=u\cos \mathrm{\θ}\cos \mathrm{\ψ}+v(\sin \mathrm{\φ}\sin \mathrm{\θ}\sin \mathrm{\ψ}+\cos \mathrm{\φ}\cos \mathrm{\ψ})+\mathrm{\ω}(-\sin \mathrm{\φ}\cos \mathrm{\ψ}+\cos \mathrm{\φ}\sin \mathrm{\θ}\sin \mathrm{\ψ})\\ {\stackrel{\·}{z}}_g=u\sin \mathrm{\θ}-v\sin \mathrm{\φ}\cos \mathrm{\θ}-\mathrm{\ω}\cos \mathrm{\φ}\cos \mathrm{\θ}\end{array}\right.---\left(15\right)$

Current position state rate of change can be calculated by equation (15)

9). by[x_(k)g y_(k)g z_(k)g]^TResolve the next position information [x_(k+1)g y_(k+1)g z_(k+1)g]^T

NextState positional information resolving equation:

$\left[\begin{array}{c}{x}_{(k+1)g}\\ y_{(k+1)g}\\ z_{(k+1)g}\end{array}\right]=\left[\begin{array}{c}{x}_{\left(k\right)g}+{\stackrel{\·}{x}}_{\left(k\right)g}\mathrm{\Δ}\mathrm{\τ}\\ y_{\left(k\right)g}+{\stackrel{\·}{y}}_{\left(k\right)g}\mathrm{\Δ}\mathrm{\τ}\\ z_{\left(k\right)g}+{\stackrel{\·}{z}}_{\left(k\right)g}\mathrm{\Δ}\mathrm{\τ}\end{array}\right]---\left(16\right)$

Equation (1)～(16) are carried out discretization and can write computer program, set up Aviate equation, pass through Aviate equation Calculate the rate of change of power in aircraft each moment, moment variations rate, attitude angle rate of change, percentage speed variation can be obtained by The status information in each moment of aircraft and positional information.Resolved by state and just achieve dummy vehicle emulation, from And the performance parameter of aircraft can be simulated.

Compared with prior art, the invention have the advantages that:

Invention defines the aerodynamic parameter xml configuration file of aircraft dynamics model, can be by amendment configuration file The rapid modeling of different type of machines kinetic model can be realized.The feature of this Dynamics Model is with dynamic by aerodynamic parameter Mechanical equation is separated, it is achieved that the modularized design of kinetic model, is then loaded into by aerodynamic parameter before resolving state In model, resolve aerodynamic parameter according to current flight state and realize the resolving of airplane motion state.Use this modeling pattern can To realize the flexible combination of Dynamics Model, improve development efficiency, shorten the construction cycle, and engineering can be reduced Personnel's repeated labor in modeling process, improves the motility of Modeling of Vehicle, rapidity and versatility.

Accompanying drawing explanation

Fig. 1 is method for designing schematic flow sheet of the present invention；

Fig. 2 is that the definition of aerodynamic parameter XML document realizes framework 1；

Fig. 3 is that the definition of aerodynamic parameter XML document realizes framework 2；

Fig. 4 aerodynamic parameter XML document flow chart；

Fig. 5 is Aviate equation model realization structure chart；

Detailed description of the invention

The present invention proposes a kind of Fixed Wing AirVehicle rapid modeling method for designing based on plug type interface, proposes base In the fixed wing airplane aerodynamic parameter edit specification of XML, design plug type interface, aerodynamic parameter is shelled from kinetic model From also abstract, it is achieved aerodynamic parameter plug and play.Use VC2010 developing instrument, complete fixed wing airplane manipulation input, Kinetics equation and kinematical equation Development of Module.Design aerodynamic parameter XML file analysis program, by same for the aerodynamic parameter obtained Dynamics module combines, it is achieved the quick exploitation of flying quality phantom.User flies by replacing difference in XML file The aerodynamic parameter of row device, on the premise of engine type and model are fixing, only need to revise XML during to different Modeling of Vehicle Aerodynamic parameter in file just can complete simulation model design.Can quickly set up the air for different fixed wing airplane types to move Mechanical model, shortens the model development cycle.Reduce the repeated labor in aircraft process of mathematical modeling, improve aircraft The motility of modeling, rapidity, and versatility.

A kind of Fixed Wing AirVehicle rapid modeling method for designing based on plug type interface, mainly comprises herein below:

Step one: definition XML document, peels off the aerodynamic parameter of aircraft from Aviate equation and writes in file, real Existing Aviate equation therefrom reads aerodynamic parameter when initializing；

Step 2: according to the Kinematic Decomposition of aircraft, is moved and is divided into horizontal lateral movement and lengthwise movement, write respectively Kinestate resolves equation；

Step 3: design XML analysis program, is loaded into aerodynamic parameter in Aerodynamics Model, it is achieved kinestate Resolve the initialization of equation；

Step 4: under conditions of engine type and model are fixing, can fly so that multiplexing is same for different aircraft Row equation, only need to revise the aerodynamic parameter in XML file and the most again be resolved in kinetic model；

Defined in described step one, the realization approach of XML document is:

(1). first created XML document, definition document root node by xml editor；

(2). definition ground floor node, totally 8 nodes, labelling ground floor nodes, lift resolve relevant pneumatic ginseng respectively Number, side force resolve aerodynamic parameter, pitching moment resolves aerodynamic parameter, rolling moment resolves aerodynamic parameter, yawing resolves gas Dynamic parameter, damping force resolve aerodynamic parameter, angular acceleration resolves aerodynamic parameter；

(3). for each node creation node of ground floor, create child node according to the aerodynamic parameter number of corresponding node, Relate to a how many aerodynamic parameter and be created that several child node, wherein the child node number of first sub-vertex ticks present node；

(4). one aerodynamic parameter of a sub-node on behalf, is series of discrete point owing to being stored in the aerodynamic parameter of document, because of This needs to create a series of nexine nodes for child node.Each child node extends 21 nexine nodes, wherein first nexine Vertex ticks aerodynamic parameter discrete point number, remaining nexine vertex ticks aerodynamic parameter centrifugal pump；

(5). Save and Close XML document, it is achieved the labelling of aerodynamic parameter；

In described step 3, XML document parsing realization approach is:

(1). load XML document, proceed by aerodynamic parameter and resolve；

(2). text pointer is navigated to root node, obtains the nodes of ground floor node；

(3). text pointer is navigated to successively each node of ground floor, begins stepping through the child node of each node；

(4). the nexine node of traversal child node extension, by the aerodynamic parameter value of labelling in nexine node, read corresponding gas In aerodynamic parameter array corresponding in dynamic resolving module；

(5). repeat aforesaid operations, until the nexine node in the child node of all nodes all reads；

(6). Save and Close XML document, start-up parameter is loaded in calculator memory；

Aerodynamic force, moment that described ground floor node on behalf is corresponding resolve module；The child node of node represents this pneumatic solution Calculate the aerodynamic parameter that module relates to；The dimensional table data amount check of this aerodynamic parameter labelling of nexine node on behalf of child node；

Need write XML file aerodynamic parameter table:

For just understanding, the XML document tag format of attached following airfoil lift coefficient.

Implement:

1 defined variable:

Variable module includes that abstract variable that each aerodynamic parameter of aircraft is corresponding, kinestate variable, mechanical state become Amount and control input.Wherein aerodynamic parameter variable is for the resolving of aircraft aerodynamic model；Kinestate variable is used for Attitude information that record cast calculates, positional information etc.；Pitching moment that mechanical state variable calculates for record cast, Rolling moment, yawing, lateral moment and the status information such as resistance, lift；Control input and include [δ_T δ_e δ_a δ_r], right Answer the input of throttle push rod, elevator drift angle, aileron movement angle, rudder.

2 definition XML document and XML analysis programs:

The aerodynamic parameter of vehicle aerodynamics model is write in XML file.By gas from Aerodynamics Model Dynamic parameter is stripped out and abstract, replaces with variable in each resolving module, by pneumatic by file of XML analysis program Parameter analysis of electrochemical is in corresponding variable, it is achieved the initialization of Aerodynamics Model.The definition of XML document is by the gas in table 1 Dynamic parameter recorded in document with tree-like hierarchy structure successively according to resolving equation, and all aerodynamic parameters of such aircraft are with regard to structure Having become to have the tree of multilamellar bifurcation structure, each aerodynamic parameter corresponds to an element of one of them branch.Analysis program Effect be exactly parsing tree, by carry out tree traversal just each element can be operated, be resolved to aircraft flight In equation, thus Aviate equation is embodied.The method creating XML document is shown in Figure of description 1, Fig. 2, XML document parsing side Method is shown in Figure of description 3.

3 states resolving modules:

State resolves module and includes the power equation group of aircraft, movement difference equations, momental equation group, navigation equation group four Point, write when state resolves module and corresponding equation group discretization is converted into computer program.Determining state vectorWith control input [δ_T δ_e δ_a δ_rRelation between] After, at the relevant aerodynamic parameter of known aircraft, characteristic parameter, very according to flying height h, Mach number M_aAnd state of flight, just May determine that power (F_x F_y F_z) and momentApplication state resolves module, and in office when just can solve aircraft The kinestate carved.Wherein (u, v w) are three velocity components of body axis system；For attitude angle, respectively rolling Angle, the angle of pitch, course angle；(p, q, r) be body angular velocity component, respectively body angular velocity in roll, body rate of pitch And body yaw rate；(x_g,y_g,z_g) it is geographical coordinates；δ_T,δ_e,δ_a,δ_rBe respectively accelerator open degree input, elevator inclined Input, aileron rudder partially inputs and course rudder inputs partially.

3). according to controlling input [δ_T δ_e δ_a δ_r] resolve power, moment.

Power mainly includes lift L, side force Y, motor power T.Concrete resolving is as follows:

Lift resolving equation:

Wherein have

Q: dynamic pressure；

ρ: current gas pressure level air density；

V: aircraft horizontal flight speed；

C_LW: airfoil lift coefficient；

C_Lb: fuselage lift coefficient；

S_b: fuselage cross-section is amassed；

S_W: wing area；

C_Lt: horizontal tail lift coefficient；

S_t: horizontal tail area；

Side force resolving equation:

ρ: current gas pressure level air density；

V: aircraft horizontal flight speed；

S_W: wing area；

C_Yβ: the side force derivative that yaw angle causes；

β: yaw angle；

: rudder side force derivative；

δ_r: rudder inputs；

: angular velocity in roll side force derivative；

: withFor the angular velocity in roll of dimension, b is wing length；

: yaw rate side force derivative；

: withFor the yaw rate of dimension, b is wing length；

Pitching moment resolving equation:

$M_A=\left(\frac{1}{2}{\mathrm{\ρV}}^2\right)(C_{m,\mathrm{\α}=0}+C_{m\mathrm{\α}}\mathrm{\α}+C_{{\mathrm{m\δ}}_e}{\mathrm{\δ}}_e+C_{m\stackrel{\‾}{q}}\stackrel{\‾}{q}+C_{m\stackrel{\‾}{\stackrel{\·}{\mathrm{\α}}}}\stackrel{\‾}{\stackrel{\·}{\mathrm{\α}}}+C_{m{\stackrel{\‾}{\stackrel{\·}{\mathrm{\δ}}}}_e}{\stackrel{\‾}{\stackrel{\·}{\mathrm{\δ}}}}_e)---\left(3\right)$

ρ: current gas pressure level air density；

V: aircraft horizontal flight speed；

C_{M, α=0}: static indeterminacy pitching moment；

C_mα: angle of attack pitching moment derivative；

α: the angle of attack；

: elevator pitching moment derivative；

δ_e: elevator drift angle；

: the additional pitching moment coefficient of horizontal tail；

: withFor the rate of pitch of dimension, b is wing length；

: wash time difference damping torque derivative under horizontal tail；

: withAngle of attack speed for dimension

: elevator drift angle speed pitching moment derivative；

: withElevator drift angle speed for dimension；

c_A: wing mean geometric of airfoil；

Rolling moment resolving equation:

C_lβ: roll static-stability derivative；

V: aircraft horizontal flight speed；

B:b is wing length；

ρ: current gas pressure level air density；

: roll guidance derivative；

δ_a: aileron movement angle；

: rudder control cross derivative；

: roll damping derivative；

: withFor the angular velocity in roll of dimension, b is wing length；

: intersection dynamic derivative；

: withFor the yaw rate of dimension, b is wing length；

Yawing resolving equation:

ρ: current gas pressure level air density；

V: aircraft horizontal flight speed；

C_nβ: driftage static-stability derivative；

: aileron control cross derivative；

δ_a: aileron movement angle；

: directional control derivative；

δ_r: course angle of rudder reflection；

: intersection dynamic derivative；

: withFor the angular velocity in roll of dimension, b is wing length；

: course damping derivative；

: withFor the yaw rate of dimension, b is wing length；

Damping force resolves equation: D_k=C_DV_k (6)

C_D: damped coefficient；

Aircraft can be calculated along body axis system force vector by equation (1)～(6):

$\left[\begin{array}{c}{F}_{kx}\\ F_{ky}\\ F_{kz}\end{array}\right]=\left[\begin{array}{c}{T}_k\\ 0\\ 0\end{array}\right]+\left[\begin{array}{c}0\\ Y_k\\ -L_k\end{array}\right]+\left[\begin{array}{c}-D_k-m_k{g}_{k}sin\mathrm{\θ}\\ m_k{g}_k\cos \mathrm{\θ}\sin \mathrm{\φ}\\ m_k{g}_k\cos \mathrm{\θ}\cos \mathrm{\φ}\end{array}\right]$

Wherein k is that kth resolves the cycle；D_kFor flight resistance；m_kThe quality of cycle aircraft is resolved for kth；g_kFor kth The acceleration of gravity in individual resolving cycle.

(7) moment vector can be decomposed along body axis system by calculating aircraft by equation (1)～(6):

$\left[\begin{array}{c}{L}_k\\ M_k\\ N_k\end{array}\right]=\left[\begin{array}{c}0\\ T_k{l}_z\\ -T_k{l}_y\end{array}\right]+\left[\begin{array}{c}{\stackrel{\‾}{L}}_{kA}\\ M_{kA}\\ N_{kA}\end{array}\right]---\left(8\right)$

WhereinM_kA, N_kAIt is respectively kth and resolves cycle roll guidance moment, pitch control moment, yaw control power Square；T_kThe motor power in cycle is resolved for kth；l_y,l_zIt is respectively the rotary inertia of y-axis, the rotary inertia of z-axis；L_k,M_k, N_kIt is respectively rolling moment, pitching moment, yawing.

4). by [F_kx F_ky F_kz]^T、[u_k ν_k ω_k]^T、[φ_k θ_k ψ_k]^T、[p_k q_k r_k]^TCalculate current state to accelerate Degree component.

Component of acceleration resolving equation:

Current state component of acceleration is calculated by equation (9)

3). by [L_k M_k N_k]^T、[p_k q_k r_k]^TCalculate current state component of angular acceleration.

Component of angular acceleration resolving equation:

Wherein: I_xzFor the product of inertia；

$c_1=\frac{(I_y-I_z)I_z-I_{xz}^2}{I_x{I}_z-I_{xz}^2},c_2=\frac{(I_x-I_y+I_z)I_{xz}}{I_x{I}_z-I_{xz}^2},c_3=\frac{I_z}{I_x{I}_z-I_{xz}^2};$

$c_4=\frac{I_{xz}}{I_x{I}_z-I_{xz}^2},c_5=\frac{I_z-I_x}{I_y},c_6=\frac{I_{xz}}{I_y};$

$c_7=\frac{1}{I_y},c_8=\frac{(I_x-I_y)I_x+I_{xz}^2}{I_x{I}_z-I_{xz}^2},c_9=\frac{I_x}{I_x{I}_z-I_{xz}^2};$

Current state component of angular acceleration is calculated by equation (10)

4). by [u_k ν_k ω_k]^T、Calculate the velocity component [u of NextState_k+1 ν_k+1 ω_k+1]^T

The velocity component resolving equation of NextState:

Wherein Δ τ is the resolving cycle.

5). by [p_k q_k r_k]^T、Calculate the angular velocity component [p of NextState_k+1 q_k+1 r_k+1]^T

The angular velocity component resolving equation of NextState:

$\left[\begin{array}{c}{p}_{k+1}\\ q_{k+1}\\ r_{k+1}\end{array}\right]=\left[\begin{array}{c}{p}_k+{\stackrel{\·}{p}}_k\mathrm{\Δ}\mathrm{\τ}\\ q_k+{\stackrel{\·}{q}}_k\mathrm{\Δ}\mathrm{\τ}\\ r_k+{\stackrel{\·}{r}}_k\mathrm{\Δ}\mathrm{\τ}\end{array}\right]---\left(12\right)$

Wherein Δ τ is the resolving cycle.

6). by [φ_k θ_k ψ_k]^T、[p_k+1 q_k+1 r_k+1]^TCalculate current state attitude angular rate

Current state attitude angular rate resolving equation:

$\left\{\begin{array}{c}\stackrel{\·}{\mathrm{\φ}}=p+(rcos\mathrm{\φ}+qsin\mathrm{\φ})tan\mathrm{\θ}\\ \stackrel{\·}{\mathrm{\θ}}=q\cos \mathrm{\φ}-rsin\mathrm{\φ}\\ \stackrel{\·}{\mathrm{\ψ}}=\frac{1}{\cos \mathrm{\θ}}(rcos\mathrm{\φ}+q\sin \mathrm{\φ})\end{array}\right.---\left(13\right)$

Current state attitude angular rate can be calculated by equation (13)

7). by[φ_k θ_k ψ_k]^TCalculate the attitude angle [φ of NextState_k+1 θ_k+1 ψ_k+1]^T

NextState solving of attitude equation:

$\left[\begin{array}{c}{\mathrm{\φ}}_{k+1}\\ {\mathrm{\θ}}_{k+1}\\ {\mathrm{\ψ}}_{k+1}\end{array}\right]=\left[\begin{array}{c}{\mathrm{\φ}}_k+{\stackrel{\·}{\mathrm{\φ}}}_k\mathrm{\Δ}\mathrm{\τ}\\ {\mathrm{\θ}}_k+{\stackrel{\·}{\mathrm{\θ}}}_k\mathrm{\Δ}\mathrm{\τ}\\ {\mathrm{\ψ}}_k+{\stackrel{\·}{\mathrm{\ψ}}}_k\mathrm{\Δ}\mathrm{\τ}\end{array}\right]---\left(14\right)$

Wherein Δ τ is the resolving cycle.

8). by [u_k+1 ν_k+1 ω_k+1]^T、[φ_k+1 θ_k+1 ψ_k+1]^TCalculate the location status information change rate of aircraft

Location status information change rate resolving equation:

$\left\{\begin{array}{c}{\stackrel{\·}{x}}_g=u\cos \mathrm{\θ}\cos \mathrm{\ψ}+v(\sin \mathrm{\φ}\sin \mathrm{\θ}\cos \mathrm{\ψ}-\cos \mathrm{\φ}\sin \mathrm{\ψ})+\mathrm{\ω}(\sin \mathrm{\φ}\sin \mathrm{\ψ}+\cos \mathrm{\φ}\sin \mathrm{\θ}\cos \mathrm{\θ})\\ {\stackrel{\·}{y}}_g=u\cos \mathrm{\θ}\cos \mathrm{\ψ}+v(\sin \mathrm{\φ}\sin \mathrm{\θ}\sin \mathrm{\ψ}+\cos \mathrm{\φ}\cos \mathrm{\ψ})+\mathrm{\ω}(-\sin \mathrm{\φ}\cos \mathrm{\ψ}+\cos \mathrm{\φ}\sin \mathrm{\θ}\sin \mathrm{\ψ})\\ {\stackrel{\·}{z}}_g=u\sin \mathrm{\θ}-v\sin \mathrm{\φ}\cos \mathrm{\θ}-\mathrm{\ω}\cos \mathrm{\φ}\cos \mathrm{\θ}\end{array}\right.---\left(15\right)$

Current position state rate of change can be calculated by equation (15)

9). by[x_(k)g y_(k)g z_(k)g]^TResolve the next position information [x_(k+1)g y_(k+1)g z_(k+1)g]^T

NextState positional information resolving equation:

$\left[\begin{array}{c}{x}_{(k+1)g}\\ y_{(k+1)g}\\ z_{(k+1)g}\end{array}\right]=\left[\begin{array}{c}{x}_{\left(k\right)g}+{\stackrel{\·}{x}}_{\left(k\right)g}\mathrm{\Δ}\mathrm{\τ}\\ y_{\left(k\right)g}+{\stackrel{\·}{y}}_{\left(k\right)g}\mathrm{\Δ}\mathrm{\τ}\\ z_{\left(k\right)g}+{\stackrel{\·}{z}}_{\left(k\right)g}\mathrm{\Δ}\mathrm{\τ}\end{array}\right]---\left(16\right)$

Equation (1)～(16) are carried out discretization and can write computer program, set up Aviate equation, pass through Aviate equation Calculate the rate of change of power in aircraft each moment, moment variations rate, attitude angle rate of change, percentage speed variation can be obtained by The status information in each moment of aircraft and positional information.Resolved by state and just achieve dummy vehicle emulation, from And the performance parameter of aircraft can be simulated.

Table 1 kinetic model aerodynamic parameter

Claims (4)

1. a Fixed Wing AirVehicle rapid modeling method for designing based on plug type interface, it is characterised in that its comprise with Lower step:

Step one, defines XML document, is peeled off by the aerodynamic parameter of aircraft and write in file, it is achieved fly from Aviate equation Row equation therefrom reads aerodynamic parameter when initializing；

Step 2, according to the Kinematic Decomposition of aircraft, is moved and is divided into horizontal lateral movement and lengthwise movement, write motion respectively State resolves equation；

Step 3, designs XML analysis program, is loaded in Aerodynamics Model by aerodynamic parameter, it is achieved kinestate resolves The initialization of equation；

Step 4, under conditions of engine type and model are fixing, can be with the same flight side of multiplexing for different aircraft Journey, only need to revise the aerodynamic parameter in XML file and the most again be resolved in kinetic model.

A kind of Fixed Wing AirVehicle rapid modeling method for designing based on plug type interface the most according to claim 1, It is characterized in that, defined in described step one, XML document specifically comprises the following steps that

(1). first created XML document, definition document root node by xml editor；

(2). definition ground floor node, totally 8 nodes, labelling ground floor nodes, lift resolve relevant aerodynamic parameter, side respectively Power resolves aerodynamic parameter, pitching moment resolves aerodynamic parameter, rolling moment resolves aerodynamic parameter, yawing resolves pneumatic ginseng Number, damping force resolve aerodynamic parameter, angular acceleration resolves aerodynamic parameter；

(3). for each node creation node of ground floor, create child node according to the aerodynamic parameter number of corresponding node, relate to A how many aerodynamic parameter are created that several child node, wherein the child node number of first sub-vertex ticks present node；

(4). one aerodynamic parameter of a sub-node on behalf, it is series of discrete point owing to being stored in the aerodynamic parameter of document, therefore needs Child node to be creates a series of nexine nodes；Each child node extends 21 nexine nodes, wherein first nexine node Labelling aerodynamic parameter discrete point number, remaining nexine vertex ticks aerodynamic parameter centrifugal pump；

(5). Save and Close XML document, it is achieved the labelling of aerodynamic parameter.

A kind of Fixed Wing AirVehicle rapid modeling method for designing based on plug type interface the most according to claim 1, It is characterized in that, what in described step 3, XML document resolved specifically comprises the following steps that

(1). load XML document, proceed by aerodynamic parameter and resolve；

(2). text pointer is navigated to root node, obtains the nodes of ground floor node；

(3). text pointer is navigated to successively each node of ground floor, begins stepping through the child node of each node；

(4). the nexine node of traversal child node extension, by the aerodynamic parameter value of labelling in nexine node, read corresponding pneumatic solution Calculate in aerodynamic parameter array corresponding in module；

(5). repeat aforesaid operations, until the nexine node in the child node of all nodes all reads；

(6). Save and Close XML document, start-up parameter is loaded in calculator memory.

A kind of Fixed Wing AirVehicle rapid modeling method for designing based on plug type interface the most according to claim 1, It is characterized in that, in described step 3, aircraft kinestate resolves equation, including herein below:

Described state resolves equation and includes the power equation group of aircraft, movement difference equations, momental equation group, navigation equation group four Point, write when state resolves module and corresponding equation group discretization is converted into computer program；Determining state vectorWith control input [δ_T δ_e δ_a δ_rRelation between], Aerodynamic parameter that known aircraft is relevant, characteristic parameter, very according to flying height h, Mach number M_aAnd state of flight, it is possible to really Determine power (F_x F_y F_z) and momentApplication state resolves module just can solve aircraft fortune at any time Dynamic state；Wherein (u, v w) are three velocity components of body axis system；For attitude angle, respectively roll angle, bow The elevation angle, course angle；(p, q, r) be body angular velocity component, respectively body angular velocity in roll, body rate of pitch and machine Body yaw rate；(x_g,y_g,z_g) it is geographical coordinates；δ_T,δ_e,δ_a,δ_rBe respectively accelerator open degree input, elevator inputs partially, pair Wing rudder inputs partially and course rudder inputs partially；

1). according to controlling input [δ_T δ_e δ_a δ_r] resolve power, moment；

Power mainly includes lift L, side force Y, motor power T；Concrete resolving is as follows:

Lift resolving equation: Wherein have

Q: dynamic pressure；

ρ: current gas pressure level air density；

V: aircraft horizontal flight speed；

C_LW: airfoil lift coefficient；

C_Lb: fuselage lift coefficient；

S_b: fuselage cross-section is amassed；

S_W: wing area；

C_Lt: horizontal tail lift coefficient；

S_t: horizontal tail area；

Side force resolving equation:

ρ: current gas pressure level air density；

V: aircraft horizontal flight speed；

S_W: wing area；

C_Yβ: the side force derivative that yaw angle causes；

β: yaw angle；

Rudder side force derivative；

δ_r: rudder inputs；

Angular velocity in roll side force derivative；

WithFor the angular velocity in roll of dimension, b is wing length；

Yaw rate side force derivative；

WithFor the yaw rate of dimension, b is wing length；

Pitching moment resolving equation:

ρ: current gas pressure level air density；

V: aircraft horizontal flight speed；

C_{M, α=0}: static indeterminacy pitching moment；

C_mα: angle of attack pitching moment derivative；

α: the angle of attack；

Elevator pitching moment derivative；

δ_e: elevator drift angle；

The additional pitching moment coefficient of horizontal tail；

WithFor the rate of pitch of dimension, b is wing length；

Time difference damping torque derivative is washed under horizontal tail；

WithAngle of attack speed for dimension

Elevator drift angle speed pitching moment derivative；

WithElevator drift angle speed for dimension；

c_A: wing mean geometric of airfoil；

Rolling moment resolving equation:

C_lβ: roll static-stability derivative；

V: aircraft horizontal flight speed；

B:b is wing length；

ρ: current gas pressure level air density；

Roll guidance derivative；

δ_a: aileron movement angle；

Rudder control cross derivative；

Roll damping derivative；

WithFor the angular velocity in roll of dimension, b is wing length；

Intersection dynamic derivative；

WithFor the yaw rate of dimension, b is wing length；

Yawing resolving equation:

ρ: current gas pressure level air density；

V: aircraft horizontal flight speed；

C_nβ: driftage static-stability derivative；

Aileron control cross derivative；

δ_a: aileron movement angle；

Directional control derivative；

δ_r: course angle of rudder reflection；

Intersection dynamic derivative；

WithFor the angular velocity in roll of dimension, b is wing length；

Course damping derivative；

WithFor the yaw rate of dimension, b is wing length；

Damping force resolves equation: D_k=C_DV_k (6)

C_D: damped coefficient；

Aircraft can be calculated along body axis system force vector by equation (1)～(6):

$\left[\begin{array}{c}{F}_{kx}\\ F_{ky}\\ F_{kz}\end{array}\right]=\left[\begin{array}{c}{T}_k\\ 0\\ 0\end{array}\right]+\left[\begin{array}{c}0\\ Y_k\\ -L_k\end{array}\right]+\left[\begin{array}{c}-D_k-m_k{g}_{k}sin\mathrm{\θ}\\ m_k{g}_k\cos \mathrm{\θ}\sin \mathrm{\φ}\\ m_k{g}_k\cos \mathrm{\θ}\cos \mathrm{\φ}\end{array}\right]---\left(7\right)$

Wherein k is that kth resolves the cycle；D_kFor flight resistance；m_kThe quality of cycle aircraft is resolved for kth；g_kFor kth solution The acceleration of gravity in calculation cycle；

Moment vector can be decomposed along body axis system by calculating aircraft by equation (1)～(6):

$\left[\begin{array}{c}{L}_k\\ M_k\\ N_k\end{array}\right]=\left[\begin{array}{c}0\\ T_k{l}_z\\ -T_k{l}_y\end{array}\right]+\left[\begin{array}{c}{\stackrel{\‾}{L}}_{kA}\\ M_{kA}\\ N_{kA}\end{array}\right]---\left(8\right)$

WhereinM_kA, N_kAIt is respectively kth and resolves cycle roll guidance moment, pitch control moment, yaw control moment；T_k The motor power in cycle is resolved for kth；l_y,l_zIt is respectively the rotary inertia of y-axis, the rotary inertia of z-axis；L_k,M_k,N_kRespectively For rolling moment, pitching moment, yawing；

2). by [F_kx F_ky F_kz]^T、[u_k ν_k ω_k]^T、[φ_k θ_k ψ_k]^T、[p_k q_k r_k]^TCalculate current state acceleration to divide Amount；

Component of acceleration resolving equation:

Current state component of acceleration is calculated by equation (9)

3). by [L_k M_k N_k]^T、[p_k q_k r_k]^TCalculate current state component of angular acceleration；

Component of angular acceleration resolving equation:

Wherein: I_xzFor the product of inertia；

$c_1=\frac{\left(I_y-I_z\right)I_z-I_{xz}^2}{I_x{I}_z-I_{xz}^2},c_2=\frac{\left(I_x-I_y+I_z\right)I_{xz}}{I_x{I}_z-I_{xz}^2},c_3=\frac{I_z}{I_x{I}_z-I_{xz}^2};$

$c_4=\frac{I_{xz}}{I_x{I}_z-I_{xz}^2},c_5=\frac{I_z-I_x}{I_y},c_6=\frac{I_{xz}}{I_y};$

$c_7=\frac{1}{I_y},c_8=\frac{(I_x-I_y)I_x+I_{xz}^2}{I_x{I}_z-I_{xz}^2},c_9=\frac{I_x}{I_x{I}_z-I_{xz}^2};$

Current state component of angular acceleration is calculated by equation (10)

4). by [u_k ν_k ω_k]^T、Calculate the velocity component [u of NextState_k+1 ν_k+1 ω_k+1]^TNext shape The velocity component resolving equation of state:

Wherein Δ τ is the resolving cycle；

5). by [p_k q_k r_k]^T、Calculate the angular velocity component [p of NextState_k+1 q_k+1 r_k+1]^TNext shape The angular velocity component resolving equation of state:

$\left[\begin{array}{c}{p}_{k+1}\\ q_{k+1}\\ r_{k+1}\end{array}\right]=\left[\begin{array}{c}{p}_k+{\stackrel{\·}{p}}_k\mathrm{\Δ}\mathrm{\τ}\\ q_k+{\stackrel{\·}{q}}_k\mathrm{\Δ}\mathrm{\τ}\\ r_k+{\stackrel{\·}{r}}_k\mathrm{\Δ}\mathrm{\τ}\end{array}\right]---\left(12\right)$

Wherein Δ τ is the resolving cycle；

6). by [φ_k θ_k ψ_k]^T、[p_k+1 q_k+1 r_k+1]^TCalculate current state attitude angular rate current state attitude angular rate Resolving equation:

$\left\{\begin{array}{c}\stackrel{\·}{\mathrm{\φ}}=p+(rcos\mathrm{\φ}+qsin\mathrm{\φ})tan\mathrm{\θ}\\ \stackrel{\·}{\mathrm{\θ}}=q\cos \mathrm{\φ}-rsin\mathrm{\φ}\\ \stackrel{\·}{\mathrm{\ψ}}=\frac{1}{\cos \mathrm{\θ}}(rcos\mathrm{\φ}+q\sin \mathrm{\φ})\end{array}\right.---\left(13\right)$

Current state attitude angular rate can be calculated by equation (13)

7). by[φ_k θ_k ψ_k]^TCalculate the attitude angle [φ of NextState_k+1 θ_k+1 ψ_k+1]^TNextState Solving of attitude equation:

$\left[\begin{array}{c}{\mathrm{\φ}}_{k+1}\\ {\mathrm{\θ}}_{k+1}\\ {\mathrm{\ψ}}_{k+1}\end{array}\right]=\left[\begin{array}{c}{\mathrm{\φ}}_k+{\stackrel{\·}{\mathrm{\φ}}}_k\mathrm{\Δ}\mathrm{\τ}\\ {\mathrm{\θ}}_k+{\stackrel{\·}{\mathrm{\θ}}}_k\mathrm{\Δ}\mathrm{\τ}\\ {\mathrm{\ψ}}_k+{\stackrel{\·}{\mathrm{\ψ}}}_k\mathrm{\Δ}\mathrm{\τ}\end{array}\right]---\left(14\right)$

Wherein Δ τ is the resolving cycle；

8). by [u_k+1 ν_k+1 ω_k+1]^T、[φ_k+1 θ_k+1 ψ_k+1]^TCalculate the location status information change rate position shape of aircraft State information change rate resolving equation:

$\left\{\begin{array}{c}{\stackrel{\·}{x}}_g=u\cos \mathrm{\θ}\cos \mathrm{\ψ}+v\left(\sin \mathrm{\φ}\sin \mathrm{\θ}\cos \mathrm{\ψ}-\cos \mathrm{\φ}\sin \mathrm{\ψ}\right)+\mathrm{\ω}\left(\sin \mathrm{\φ}\sin \mathrm{\ψ}+\cos \mathrm{\φ}\sin \mathrm{\θ}\cos \mathrm{\ψ}\right)\\ {\stackrel{\·}{y}}_g=u\cos \mathrm{\θ}\cos \mathrm{\ψ}+v\left(\sin \mathrm{\φ}\sin \mathrm{\θ}\sin \mathrm{\ψ}+\cos \mathrm{\φ}\cos \mathrm{\ψ}\right)+\mathrm{\ω}\left(-\sin \mathrm{\φ}\cos \mathrm{\ψ}+\cos \mathrm{\φ}\sin \mathrm{\θ}\sin \mathrm{\ψ}\right)\\ {\stackrel{\·}{z}}_g=u\sin \mathrm{\θ}-v\sin \mathrm{\φ}\cos \mathrm{\θ}-\mathrm{\ω}\cos \mathrm{\φ}\cos \mathrm{\θ}\end{array}\right.---\left(15\right)$

Current position state rate of change can be calculated by equation (15)

9). byResolve the next position information [x_(k+1)g y_(k+1)g z_(k+1)g]^TNextState positional information resolving equation:

$\left[\begin{array}{c}{x}_{(k+1)g}\\ y_{(k+1)g}\\ z_{(k+1)g}\end{array}\right]=\left[\begin{array}{c}{x}_{\left(k\right)g}+{\stackrel{\·}{x}}_{\left(k\right)g}\mathrm{\Δ}\mathrm{\τ}\\ y_{\left(k\right)g}+{\stackrel{\·}{y}}_{\left(k\right)g}\mathrm{\Δ}\mathrm{\τ}\\ z_{\left(k\right)g}+{\stackrel{\·}{z}}_{\left(k\right)g}\mathrm{\Δ}\mathrm{\τ}\end{array}\right]---\left(16\right)$

Equation (1)～(16) are carried out discretization and can write computer program, set up Aviate equation, resolved by Aviate equation Go out the rate of change of power in aircraft each moment, moment variations rate, attitude angle rate of change, percentage speed variation can be obtained by flight The status information in each moment of device and positional information.Resolved by state and just achieve dummy vehicle emulation, thus can To simulate the performance parameter of aircraft.