CN112180728B - Method for constructing carrier-based aircraft landing longitudinal motion linear model - Google Patents
Method for constructing carrier-based aircraft landing longitudinal motion linear model Download PDFInfo
- Publication number
- CN112180728B CN112180728B CN202011064970.XA CN202011064970A CN112180728B CN 112180728 B CN112180728 B CN 112180728B CN 202011064970 A CN202011064970 A CN 202011064970A CN 112180728 B CN112180728 B CN 112180728B
- Authority
- CN
- China
- Prior art keywords
- carrier
- trim
- based aircraft
- formula
- equation
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Active
Links
Images
Classifications
-
- G—PHYSICS
- G05—CONTROLLING; REGULATING
- G05B—CONTROL OR REGULATING SYSTEMS IN GENERAL; FUNCTIONAL ELEMENTS OF SUCH SYSTEMS; MONITORING OR TESTING ARRANGEMENTS FOR SUCH SYSTEMS OR ELEMENTS
- G05B13/00—Adaptive control systems, i.e. systems automatically adjusting themselves to have a performance which is optimum according to some preassigned criterion
- G05B13/02—Adaptive control systems, i.e. systems automatically adjusting themselves to have a performance which is optimum according to some preassigned criterion electric
- G05B13/04—Adaptive control systems, i.e. systems automatically adjusting themselves to have a performance which is optimum according to some preassigned criterion electric involving the use of models or simulators
- G05B13/042—Adaptive control systems, i.e. systems automatically adjusting themselves to have a performance which is optimum according to some preassigned criterion electric involving the use of models or simulators in which a parameter or coefficient is automatically adjusted to optimise the performance
Landscapes
- Engineering & Computer Science (AREA)
- Health & Medical Sciences (AREA)
- Artificial Intelligence (AREA)
- Computer Vision & Pattern Recognition (AREA)
- Evolutionary Computation (AREA)
- Medical Informatics (AREA)
- Software Systems (AREA)
- Physics & Mathematics (AREA)
- General Physics & Mathematics (AREA)
- Automation & Control Theory (AREA)
- Aerodynamic Tests, Hydrodynamic Tests, Wind Tunnels, And Water Tanks (AREA)
Abstract
The invention discloses a method for constructing a longitudinal motion linear model of carrier-based aircraft landing, which comprises the following steps of firstly establishing a force balance equation and a moment balance equation according to a set reference motion state of the carrier-based aircraft; constructing a Simulink model according to a calling rule of a trim function of Matlab software and a force balance equation and a moment balance equation; calling a trim function of Matlab software to realize the motion trim of carrier-based aircraft carrier landing; then establishing a nonlinear equation of the longitudinal motion of the carrier-based aircraft; constructing another Simulink model according to a calling rule of a linmod function of Matlab software and a nonlinear equation of longitudinal motion; and calling a linmod function of Matlab software to obtain a coefficient matrix of a state space equation, and finally obtaining a carrier aircraft carrier landing longitudinal motion linear model. The model constructed by the method has high precision and can better describe the response of the carrier-based aircraft to instructions and control input in the stage of approach and landing.
Description
Technical Field
The invention relates to the technical field of carrier aircraft landing control, in particular to a method for constructing a carrier aircraft landing longitudinal motion linear model.
Background
The trim and linearization of the nonlinear dynamical model are the basis for aircraft control design and flight simulation: after a stable working point (balance point) of the system is obtained through trimming, a linear model corresponding to the point is obtained through a linearization method, and the stability and the maneuverability of the aircraft near the point can be researched by using a relatively complete linear system analysis means and used as a design guide of a flight control system.
At present, a Newton iteration method or a quasi-Newton iteration method is generally adopted at home and abroad to trim and linearize an aircraft motion model. Considering that the balancing and linearization operations are in great demand in system analysis and design, some computing software on the market provides corresponding tool functions to facilitate users to use, such as a trim function and a linmod function of MATLAB software, which can respectively balance and linearize a dynamic system; moreover, if the system is linear or only comprises relatively simple non-linear links, they can be generally called directly, which has application examples in the help interface of MATLAB and some control theory MATLAB courses. However, for linear modeling of a complex nonlinear system, it is generally difficult to implement by directly calling a trim function and a linmod function, for example, the take-off and landing process of an aircraft has the characteristics of strong coupling, nonlinearity and the like, and in addition, due to the special and complex degree of performance parameters of the aircraft body, a linearized model of the take-off and landing motion of the aircraft is not easy to obtain, but an effective solution is lacking in the prior art.
Disclosure of Invention
The invention aims to provide a method for constructing a carrier-based aircraft landing longitudinal motion linear model, the method for constructing the linear model is simple and convenient to operate and high in accuracy, and the response of the carrier-based aircraft to instruction and control input in an approach landing stage can be better described.
The purpose of the invention is realized by the following technical scheme:
a method for constructing a carrier-based aircraft landing longitudinal motion linear model comprises the following steps:
and 7, obtaining an approximate ratio of the thrust of the engine to the opening degree of an accelerator according to the working characteristics of the engine in the carrier landing stage of the carrier-based aircraft, and multiplying the 3 rd row elements of the state space equation coefficient matrix B obtained in the step 6 by the approximate ratio to obtain a carrier-based aircraft landing longitudinal motion linear model.
The technical scheme provided by the invention shows that the model constructed by the method has high precision, can better describe the response of the shipboard aircraft to instructions and control input in the stage of approach and landing, does not need to compile balancing and linearization algorithm codes, has simple and convenient operation steps and has strong practicability.
Drawings
In order to more clearly illustrate the technical solutions of the embodiments of the present invention, the drawings to be used in the description of the embodiments will be briefly introduced below, and it is apparent that the drawings in the description below are only some embodiments of the present invention, and it is obvious for those skilled in the art that other drawings may be obtained based on these drawings without creative efforts.
Fig. 1 is a schematic flow chart of a method for constructing a carrier-based aircraft landing longitudinal motion linear model according to an embodiment of the invention;
FIG. 2 is a schematic diagram of a Simulink model called by a trim function according to an exemplary embodiment of the present invention;
FIG. 3 is a schematic diagram of a Simulink model called by a linmod function according to an exemplary embodiment of the present invention;
FIG. 4 is a schematic diagram showing comparison of open-loop responses of a carrier-based aircraft landing linear model and a nonlinear model to an elevator step command, which are obtained by balancing and linearization according to an example of the invention;
FIG. 5 is a schematic diagram showing comparison of open-loop responses of a carrier-based aircraft landing linear model and a nonlinear model to an accelerator step command, which are obtained through balancing and linearization according to an embodiment of the invention;
fig. 6 is a schematic diagram showing comparison of step responses of a carrier-based aircraft landing linear model and a nonlinear model obtained by trimming and linearization according to an embodiment of the invention under the action of the same control system.
FIG. 7 is a schematic diagram of the control law configuration of the longitudinal inner loop for verifying the accuracy of the linear model according to the exemplary embodiment of the present invention.
Detailed Description
The technical solutions in the embodiments of the present invention are clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all embodiments. All other embodiments, which can be derived by a person skilled in the art from the embodiments of the present invention without making any creative effort, shall fall within the protection scope of the present invention.
The embodiment of the present invention will be further described in detail with reference to the accompanying drawings, and as shown in fig. 1, a schematic flow chart of a method for constructing a carrier-based aircraft landing longitudinal motion linear model provided in the embodiment of the present invention is shown, where the method includes:
in the step, the set reference motion state of the carrier-based aircraft, namely the steady-state flight state, is a steady-state linear flight state of the carrier-based aircraft in a calm atmosphere environment at a track angle of-3.5 degrees, an attack angle of 8.1 degrees and a constant speed of 70m/s in a vertical plane.
The force balance and moment balance equations established are expressed as:
wherein, the first formula in the formula (1) represents that the ship-based aircraft is stressed and balanced in the thrust direction of the engine thereof; the second expression represents that the stress of the carrier-based aircraft in the motion plane of the carrier-based aircraft is balanced in the direction vertical to the thrust direction of the engine; the third formula shows that the pitching moment borne by the carrier-based aircraft is balanced; wherein m is the mass of the carrier-based aircraft, and 21280kg is taken; g is gravity acceleration, and is 9.81m/s2(ii) a Alpha, theta and q respectively represent the attack angle, the pitch angle and the pitch angle speed of the shipboard aircraft; l, D, M respectively representing the aerodynamic lift, the resistance and the pitching moment borne by the carrier-based aircraft, wherein the specific expression is shown in formula (2); p represents the thrust of the engine of the carrier-based aircraft, the acting direction of the thrust is along the longitudinal axis of the aircraft body, and the mounting angle and the eccentricity of the thrust are both zero; i isyRepresenting the pitch moment of inertia, 368908kg/m2(ii) a Subscript denotes a reference value (nominal value) of the variable;
in the formula, ρ represents the air density and is 1.21kg · m-3(ii) a U represents the size of the airspeed of the carrier-based aircraft (the same as the ground speed of the carrier-based aircraft in the absence of wind); s represents the reference area of the wing, and 62m is taken2；CL、CD、CmThe lift coefficient, the drag coefficient and the pitching moment coefficient are respectively, and the specific expressions are shown as formula (3):
in the formula, deltaeAnd deltacRespectively representing a rudder deflection angle and a canard deflection angle; cAThe average pneumatic chord length is expressed, and 4m is taken; cL1、CL2、CL3、CD1、CD2、CD3、Cm1、Cm2、Cm3、Cmq、
Are all aerodynamic coefficients, expressed in the model as α and δeThe data in the table are obtained through wind tunnel tests or flight tests.
in this step, the output quantity of the constructed Simulink model comprises two parts of out1 and out2, and the expression is as follows:
the Simulink model is constructed according to formula (1) to formula (4), and fig. 2 is a schematic diagram of the Simulink model called by the trim function according to the illustrated example of the present invention, and formula (2) and formula (3) are included in the "air dynamics and momentums" subsystem in fig. 2. When the trim function calls the Simulink model, each integral or differential link in the Simulink model is processed to generate a state quantity, so that q is used as the only state quantity of the Simulink model; the input quantities, i.e. the terminals labeled In1 and In2 In FIG. 2, are selected to be δeAnd deltac；
In the above formula (4), 5 or 10 is selected5And 104These two larger values are to improve the trim accuracy because calling the trim function does not necessarily achieve the ideal balance point of the system. the mechanism of action of the trim function is: the absolute value of the difference between the output quantity and the reference value is made as small as possible by searching the value of the input quantity; obviously, a difference of 0 is the most ideal case, but cannot be guaranteed to be achieved. Therefore, to improve the trimming accuracy, q and (L · cos α) are not added*+D·sinα*-mg·cosθ*) Directly as the output of the Simulink model.
specifically, the call format of the trim function is [ x _ trim, u _ trim, y _ trim ] ═ trim ('sys 1', x _ ref, u _ ref, y _ ref); wherein sys1 is the name of the Simulink model constructed in step 2; x _ ref represents a reference value of the state quantity of the Simulink model, namely a reference value of q, and is 0; y _ ref represents a reference value of the model output quantity, namely reference values of out1 and out2 in formula (4), which are both 0; u _ ref is asserted, so it is taken as [ ];
after the call of the trim function, u _ trim ═ 1.0978, -20.1847 is obtained](corresponding to each of. delta.)eAnd deltacValue of (d), the state quantity trim error x _ trim becomes 4.0369 × 10-5And an output quantity trim error y _ trim ═ 20.1847,20.1847]；
Because y _ ref is a 0 vector, y _ trim gives values of the trimmed output vectors out1 and out2, as shown in formula (4), the trimming precision is high, and according to formula (2) of formula (1), the carrier thrust trim value at the approximate balance point is about 28208.04N.
in this step, the longitudinal motion equation is established under the condition that only the longitudinal motion of the carrier-based aircraft is considered and no atmospheric disturbance is considered, and the specific expression is as follows:
in the formula, vIGamma and h respectively represent the ground speed, the track angle and the height of the carrier-based aircraft; the first two formulas in the formula (5) are ship-borne aircraft centroid translation kinetic equations established under a track coordinate system; the formula 3 is an Euler kinetic equation; formulas 4 and 5 are respectively a ship-borne aircraft centroid rotation equation and a translational kinematic equation; the final equation gives the mathematical relationship between α, θ and γ.
similarly, when a linmod function calls the Simulink model, each of the modelsThe integration or differentiation element is processed to generate a state quantity. Therefore, the Simulink model should be built based on the first order differential equations of these variables given in equation (5). In a specific implementation, as shown in fig. 3, a schematic diagram of a Simulink model called by linmod function according to an embodiment of the present invention is shown, where the model input quantity is δe、δcAnd P, corresponding to the terminals labeled In1, In2, and P, respectively, In fig. 3; the state quantity is x ═ vI,α,q,θ,h]TAlso, the output y corresponds to the output terminals denoted by Vk, AOA, q, Theta, and h in fig. 3.
The "aerodyne processes and Moments" subsystem in fig. 3 is also established according to formula (2) and formula (3), the expressions reflected by the series of modules connected to the output terminals labeled q, Theta, h and Vk in fig. 3 are respectively formula 3, formula 4, formula 5 and formula 1 in formula (5), and the expressions reflected by the series of modules connected to the output terminals labeled AOA in fig. 3 are derived from formula 2, formula 4 and formula 6 in formula (5).
In addition, reference values (nominal values) of the partial variables are referred to in fig. 3, including: v. ofI*＝70m/s，γ*＝-3.5°，θ*＝4.6°，α*The initial values of the integrals q and h are 0 and 200, respectively, at 8.1 °.
specifically, the call format of the linmod function is [ A, B, C, D]Linmod ('sys 2', x _ trim, u _ trim); wherein sys2 is the name of another Simulink model constructed in step 5; x _ trim is the value of the state quantity of the Simulink model in the reference motion state of the shipboard aircraft, namely
Taking the balancing result obtained in the step 3 as u _ trim, namely u _ trim ═ 1.0978, -20.1847,28208.04]；
After a linmod function is called, coefficient matrixes A-D of the state space equation are obtained, wherein the matrix C is a 5-order unit matrix, the matrix D is a zero matrix with 5 multiplied by 3 dimensions, and the matrix A and the matrix B are as follows:
and 7, obtaining an approximate ratio of the thrust of the engine to the opening degree of an accelerator according to the working characteristics of the engine in the carrier landing stage of the carrier-based aircraft, and multiplying the 3 rd row elements of the state space equation coefficient matrix B obtained in the step 6 by the approximate ratio to obtain a carrier-based aircraft landing longitudinal motion linear model.
In the step, the finally obtained carrier-based aircraft landing longitudinal motion linear model is represented as:
in the formula, Δ represents the deviation of the variable from its reference value (nominal value); column vectors x and y are respectively the state quantity and the output quantity of the other Simulink model constructed in the step 5; u represents the linear model input, u ═ δe,δc,δp]TWherein δpIndicating an effective throttle opening; coefficient matrix A, B1C, D is a coefficient matrix of the state space equation, where the matrix B1The matrix B is obtained by multiplying the 3 rd row elements of the matrix B obtained in the step 6 by a coefficient 2314, and the coefficient reflects the ship-borne model of the example in the ship landing stage deltapAnd P.
According to the method, the Simulink model is constructed by reasonably selecting and simplifying the carrier aircraft landing motion equation according to the calling rules of the trim function and the linmod function of the MATLAB, so that the accuracy of the result can be ensured after the two functions are called, and the workload of compiling the balancing and linearization algorithm is reduced.
The accuracy of the obtained carrier-based aircraft landing longitudinal motion linear model is verified by an example, and in the example, the accuracy verification is mainly realized by two parts of simulation tests:
firstly, response conditions of a linear model and a nonlinear model to a control surface and an accelerator open-loop instruction are compared, as shown in fig. 4, a comparison schematic diagram of an open-loop response of a carrier-based aircraft landing linear model and a nonlinear model obtained by balancing and linearizing in an example provided by the invention to an elevator step instruction is shown, as shown in fig. 5, a comparison schematic diagram of an open-loop response of a carrier-based aircraft landing linear model and a nonlinear model obtained by balancing and linearizing in an example provided by the invention to an accelerator step instruction is shown, and as for local accuracy of small-disturbance linearization and organism instability of a carrier-based aircraft are considered, as shown in fig. 4 and fig. 5: although the open-loop command input quickly brings both models away from equilibrium, the linear model more accurately describes the (divergent) response of the nonlinear model to the open-loop command during divergence. Fig. 4 shows that, for the motion deviation caused by the elevator from the reference motion state, in the process that the speed deviation of the non-linear model reaches 0.5m/s, the attack angle deviation reaches 5.5 °, the pitch rate deviation reaches 6.8 °/s, the pitch attitude deviation reaches 7.4 °, and the altitude change rate deviation reaches 2.3m/s, the change trends of the corresponding states of the linear model are completely consistent, and the maximum errors are respectively about: the speed is 0.3m/s, the attack angle is 0.4 degrees, the pitching speed is 0.9 degrees/s, the pitching attitude is 0.7 degrees, and the height change rate is 0.4 m/s. Fig. 5 shows that, for the motion deviation caused by the throttle relative to the reference motion state, in the process that the speed deviation of the non-linear model reaches 0.3m/s, the attack angle deviation reaches 0.15 °, the pitch rate deviation reaches 0.1 °/s, the pitch attitude deviation reaches 0.075 °, and the altitude change rate deviation reaches 0.074m/s, the change trends of each corresponding state of the linear model are completely consistent, and the maximum errors are respectively about: the speed is 0.09m/s, the attack angle is 0.06 degrees, the pitching speed is 0.05 degrees/s, the pitching attitude is 0.04 degrees, and the height change rate is 0.025 m/s.
Secondly, response conditions of the linear model and the nonlinear model under the action of the same closed-loop control System are compared, as shown in fig. 6, a step response comparison schematic diagram of the Carrier aircraft Landing linear model and the nonlinear model obtained through trimming and linearization by the example of the invention under the action of the same control System is shown, the closed-loop control System selected here has a longitudinal inner loop control law configuration of an Automatic Carrier Landing System (ACLS) of an F/a-18A "hornet" fighter, as shown in fig. 7, a longitudinal inner loop control law configuration schematic diagram of the example of the invention for verifying accuracy of the linear model is shown, and includes two parts, namely an autopilot and an Approach Power Compensator (APC), and a control law of the APC is shown in a formula (7).
In the formula, deltaplDenotes a throttle lever, nzRepresenting normal overload, the mathematical relationship with other variables in the linear model can be represented by equation (8).
Δnz＝(4×10-4ΔvI+0.0432Δα-1.35×10-6Δh+1.65×10-4Δδe)×vI* (8)
The parameters of the closed-loop control system are corrected by taking the obtained linear model as a controlled object and referring to the design performance of the F/A-18A aircraft ACLS, and are finally determined as follows:
KQ＝0.6281 KQC＝350 T＝0.5342 β＝1.5357
as can be seen from fig. 6: except for overshoot and adjusting time, the control effect of the closed-loop control system on the nonlinear model is slightly inferior to that of the linear model, the closed-loop response curves of the two models are high in goodness of fit on the whole, and namely the linear model can accurately describe the response of the nonlinear model on the closed-loop control system.
Therefore, the accuracy of the carrier-based aircraft landing longitudinal motion linear model constructed by the method provided by the embodiment of the invention is higher.
It is noted that those skilled in the art will recognize that embodiments of the present invention are not described in detail herein.
The above description is only for the preferred embodiment of the present invention, but the scope of the present invention is not limited thereto, and any changes or substitutions that can be easily conceived by those skilled in the art within the technical scope of the present invention are included in the scope of the present invention. Therefore, the protection scope of the present invention shall be subject to the protection scope of the claims.
Claims (6)
1. A method for constructing a longitudinal motion linear model of carrier-based aircraft landing is characterized by comprising the following steps:
step 1, firstly, establishing a force balance equation and a moment balance equation according to a set reference motion state of a carrier-based aircraft by taking an equiangular descent stage of the fixed-wing carrier-based aircraft before landing as a research object;
step 2, constructing a Simulink model according to a calling rule of a trim function of Matlab software and the force balance equation and the moment balance equation established in the step 1, and specifying the input quantity, the output quantity and the state quantity of the Simulink model;
step 3, aiming at the Simulink model constructed in the step 2, calling a trim function of Matlab software to achieve trim of carrier-based aircraft landing motion, and obtaining balance points of the carrier-based aircraft landing motion;
step 4, establishing an equation capable of describing longitudinal motion of the carrier-based aircraft in the equal-angle descent stage before landing;
in step 4, the longitudinal motion equation is established under the condition that only the longitudinal motion of the carrier-based aircraft is considered and atmospheric disturbance is not considered, and the specific expression is as follows:
in the formula, vIGamma and h respectively represent the ground speed, the track angle and the height of the carrier-based aircraft; m is the carrier-based aircraft mass; g is the acceleration of gravity; alpha, theta and q respectively represent the attack angle, the pitch angle and the pitch angle speed of the shipboard aircraft; l, D, M respectively representing aerodynamic lift, drag and pitch experienced by a carrier-based aircraftMoment of force; p represents the thrust of the engine of the carrier-based aircraft, the acting direction of the thrust is along the longitudinal axis of the aircraft body, and the mounting angle and the eccentricity of the thrust are both zero; i isyRepresents the pitch moment of inertia;
the first two formulas of the formula (5) are ship-borne aircraft centroid translation kinetic equations established under a track coordinate system; the formula 3 is an Euler kinetic equation; formulas 4 and 5 are respectively a ship-borne aircraft centroid rotation equation and a translational kinematic equation; the final equation gives the mathematical relationship between α, θ and γ;
step 5, constructing another Simulink model according to the calling rule of the Matlab software linmod function and the longitudinal motion equation of the carrier-based aircraft established in the step 4;
the other Simulink model in step 5 is established according to formula (5), that is:
the state quantity is x ═ vI,α,q,θ,h]TAnd is also the output y; input quantity is deltae、δcAnd P; in addition, the reference values for the partial variables used in the other Simulink model include: v. ofI*＝70m/s，γ*＝-3.5°，θ*＝4.6°，α*8.1 °; the integral initial values of q and h are respectively 0 and 200;
step 6, calling a linmod function of Matlab software aiming at the other Simulink model constructed in the step 5 by using the balancing result obtained in the step 3 to obtain a coefficient matrix of the state space equation;
the process of calling the linmod function of the Matlab software in the step 6 is as follows:
the calling format of the linmod function is as follows: [ a, B, C, D ] ═ linmod ('sys 2', x _ trim, u _ trim); wherein sys2 is the name of another Simulink model constructed in step 5; and x _ trim is a value of the state quantity of the other Simulink model in the reference motion state of the carrier-based aircraft, namely:
taking the trim result obtained in the step 3 as u _ trim, namely u _ trim is [1.0978, -20.1847,28208.04 ];
after a linmod function is called, coefficient matrixes A-D of the state space equation are obtained, wherein the matrix C is a 5-order unit matrix, the matrix D is a zero matrix with 5 multiplied by 3 dimensions, and the matrix A and the matrix B are specifically as follows:
and 7, obtaining a ratio of the thrust of the engine to the opening degree of an accelerator according to the working characteristics of the engine in the carrier landing stage of the carrier-based aircraft, and multiplying the 3 rd row elements of the state space equation coefficient matrix B obtained in the step 6 by the ratio to obtain a carrier landing longitudinal motion linear model of the carrier-based aircraft.
2. The method for constructing the carrier-based aircraft landing longitudinal motion linear model according to claim 1, wherein in the step 1, the set carrier-based aircraft reference motion state is a steady-state linear flight state of the carrier-based aircraft in a calm atmosphere environment at a-3.5-degree track angle, an 8.1-degree attack angle and a constant speed of 70m/s in a vertical plane.
3. The method for constructing the carrier-based aircraft landing longitudinal motion linear model according to claim 1, wherein in the step 1, the established force balance and moment balance equations are expressed as:
wherein, the first formula in the formula (1) represents that the ship-based aircraft is stressed and balanced in the thrust direction of the engine thereof; the second expression represents that the carrier-based aircraft is transported in the second expressionThe force in the dynamic plane is balanced in the direction vertical to the thrust direction of the engine; the third formula shows that the pitching moment borne by the carrier-based aircraft is balanced; in the formula, m is the mass of the carrier-based aircraft; g is the acceleration of gravity; alpha, theta and q respectively represent the attack angle, the pitch angle and the pitch angle speed of the shipboard aircraft; l, D, M respectively representing the aerodynamic lift force, the resistance force and the pitching moment born by the carrier-based aircraft, and the specific expression is shown in formula (2); p represents the thrust of the engine of the carrier-based aircraft, the acting direction of the thrust is along the longitudinal axis of the aircraft body, and the mounting angle and the eccentricity of the thrust are both zero; i isyRepresents the pitch moment of inertia; subscript denotes a reference value of the variable;
where ρ represents the air density; u represents the airspeed of the carrier-based aircraft and is the same as the ground speed of the carrier-based aircraft under the windless condition; s represents a wing reference area; cL、CD、CmThe lift coefficient, the drag coefficient and the pitching moment coefficient are respectively, and the specific expressions are shown as formula (3):
wherein, deltaeAnd deltacRespectively representing a rudder deflection angle and a canard deflection angle; cAThe average pneumatic chord length is expressed, and 4m is taken; cL1、CL2、CL3、CD1、CD2、CD3、Cm1、Cm2、Cm3、Cmq、
Are all aerodynamic coefficients, expressed in the model as α and δeThe data in the table is obtained through wind tunnel test or flight test.
4. The method for constructing the longitudinal motion linear model of the carrier-based aircraft landing according to claim 3, wherein in the step 2, the input quantity, the output quantity and the state quantity of the Simulink model are specified as follows:
the Simulink model is specified with a state quantity q and an input quantity deltaeAnd deltacThe output quantity comprises two elements of out1 and out2, and the expression is as follows:
in the formula, take 5.105And 104These two larger values are to improve the trim accuracy because calling the trim function does not necessarily achieve the ideal balance point of the system, which is determined by the mechanism of action of the trim function.
5. The method for constructing the carrier-based aircraft landing longitudinal motion linear model according to claim 4, wherein in the step 3, the process of calling the trim function of Matlab software to realize the carrier-based aircraft landing motion trim comprises the following steps:
the call format of the trim function is [ x _ trim, u _ trim, y _ trim ] ═ trim ('sys 1', x _ ref, u _ ref, y _ ref); wherein sys1 is the name of the Simulink model constructed in step 2; x _ ref represents a reference value of the state quantity of the Simulink model, namely a reference value of q, and is 0; y _ ref represents a reference value of the output quantity of the model, namely, reference values of out1 and out2 in formula (4), which are both 0; u _ ref is an undetermined quantity, and is taken as [ ];
after the call of the trim function, u _ trim ═ 1.0978, -20.1847 is obtained]The state quantity trim error x _ trim is 4.0369 × 10-5And an output quantity trim error y _ trim ═ 20.1847,20.1847]；
Because y _ ref is a 0 vector, y _ trim gives values of the trim output quantities out1 and out2, as can be seen from formula (4), the trim precision is high, and according to formula (2) of formula (1), the carrier thrust trim value at the balance point is 28208.04N.
6. The method for constructing the carrier-based aircraft landing longitudinal motion linear model according to claim 3, wherein in the step 7, the finally obtained carrier-based aircraft landing longitudinal motion linear model is represented as:
in the formula, Δ represents the deviation of a variable from its reference value; column vectors x and y are respectively the state quantity and the output quantity of the other Simulink model constructed in the step 5; u represents the linear model input, u ═ δe,δc,δp]TWherein δpIndicating an effective throttle opening; coefficient matrix A, B1C, D is a coefficient matrix of the state space equation, where matrix B1The matrix B is obtained by multiplying the 3 rd row elements of the matrix B obtained in the step 6 by a coefficient 2314, and the coefficient reflects the ship-borne model of the example in the ship landing stage deltapAnd P.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202011064970.XA CN112180728B (en) | 2020-09-30 | 2020-09-30 | Method for constructing carrier-based aircraft landing longitudinal motion linear model |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202011064970.XA CN112180728B (en) | 2020-09-30 | 2020-09-30 | Method for constructing carrier-based aircraft landing longitudinal motion linear model |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN112180728A CN112180728A (en) | 2021-01-05 |
| CN112180728B true CN112180728B (en) | 2022-06-14 |
Family
ID=73948184
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202011064970.XA Active CN112180728B (en) | 2020-09-30 | 2020-09-30 | Method for constructing carrier-based aircraft landing longitudinal motion linear model |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN112180728B (en) |
Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN105676672A (en) * | 2016-01-06 | 2016-06-15 | 西北工业大学 | Simulation modeling method for composite flight strategy of flapping wing air vehicle |
| CN105759609A (en) * | 2016-02-29 | 2016-07-13 | 北京航空航天大学 | Carrier-based aircraft autonomous landing method based on explicitly nonlinear model predictive control |
| CN107220419A (en) * | 2017-05-16 | 2017-09-29 | 中国人民解放军海军总医院 | A kind of modeling and simulation method of carrier-borne aircraft seat harness constrained system |
Family Cites Families (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US10339266B2 (en) * | 2016-02-16 | 2019-07-02 | Board Of Regents Of The University Of Texas Systems | Mechanisms for constructing spline surfaces to provide inter-surface continuity |
-
2020
- 2020-09-30 CN CN202011064970.XA patent/CN112180728B/en active Active
Patent Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN105676672A (en) * | 2016-01-06 | 2016-06-15 | 西北工业大学 | Simulation modeling method for composite flight strategy of flapping wing air vehicle |
| CN105759609A (en) * | 2016-02-29 | 2016-07-13 | 北京航空航天大学 | Carrier-based aircraft autonomous landing method based on explicitly nonlinear model predictive control |
| CN107220419A (en) * | 2017-05-16 | 2017-09-29 | 中国人民解放军海军总医院 | A kind of modeling and simulation method of carrier-borne aircraft seat harness constrained system |
Non-Patent Citations (1)
| Title |
|---|
| 夏桂华 等."考虑扰流的舰载机终端进场线性模型".《航空学报》.2016,第37卷(第3期),第970-983页. * |
Also Published As
| Publication number | Publication date |
|---|---|
| CN112180728A (en) | 2021-01-05 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| Liu et al. | Robust attitude stabilization for nonlinear quadrotor systems with uncertainties and delays | |
| Heidlauf et al. | Verification Challenges in F-16 Ground Collision Avoidance and Other Automated Maneuvers. | |
| Raptis et al. | A novel nonlinear backstepping controller design for helicopters using the rotation matrix | |
| Yan et al. | Extended state observer‐based sliding mode fault‐tolerant control for unmanned autonomous helicopter with wind gusts | |
| Falcón et al. | Comparative analysis of continuous sliding-modes control strategies for quad-rotor robust tracking | |
| Zheng et al. | Planar path following control for stratospheric airship | |
| Chen et al. | Design of Flight Control System for a Novel Tilt‐Rotor UAV | |
| Lu et al. | Real-time simulation system for UAV based on Matlab/Simulink | |
| CN113296525B (en) | Data-driven-based non-linear control method for tilting three-rotor unmanned aerial vehicle | |
| Cheng et al. | Neural-networks control for hover to high-speed-level-flight transition of ducted fan uav with provable stability | |
| Millidere et al. | Newton-raphson methods in aircraft trim: A comparative study | |
| CN113156985B (en) | Fixed-wing unmanned aerial vehicle obstacle avoidance robust anti-interference flight control method based on preset performance | |
| Benaddy et al. | Comparison of the different control strategies for Quadrotor Unmanned Aerial Vehicle | |
| CN112180728B (en) | Method for constructing carrier-based aircraft landing longitudinal motion linear model | |
| Sharma et al. | Simulation of maritime helicopter dynamics during approach to landing with time-accurate wind-over-deck | |
| Wahid et al. | Comparative assesment using LQR and fuzzy logic controller for a pitch control system | |
| Tanner | Modeling, identification, and control of autonomous helicopters | |
| Xu et al. | Modelling and hovering control of a novel multi-tandem ducted fan vehicle | |
| Liu et al. | Robust attitude control for tail‐sitter unmanned aerial vehicles in flight mode transitions | |
| CN114035599A (en) | Aircraft attitude control method and device and electronic equipment | |
| Millidere et al. | Further investigations on Newton-Raphson methods in aircraft trim | |
| Sun et al. | Flexible modelling and altitude control for powered parafoil system based on active disturbance rejection control | |
| Sobron | Design and Testing of a Flight Control System for Unstable Subscale Aircraft | |
| CN112182889B (en) | Method for constructing carrier-based aircraft landing longitudinal motion linear model considering wind interference | |
| Cunha et al. | A path following controller for model-scale helicopters |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| GR01 | Patent grant | ||
| GR01 | Patent grant |