CN108594653A - The performance boundary analysis system of large envelope flight control rule design - Google Patents
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Abstract
The invention belongs to automatic control technology fields, and in particular to a kind of performance boundary analysis system of large envelope flight control rule design.Aim to solve the problem that the problem of prior art can not analyze big envelope curve aircraft in performance boundary.The present invention provides a kind of the performance boundary analysis method and device of the design of large envelope flight control rule, including the first computing unit, for converting the dummy vehicle built in advance to master pattern using expansion linearization technique；Second computing unit calculates the frequency domain performance robustness of master pattern for master pattern to be carried out nonlinear coordinate transformation；Third computing unit, the stability criterion for calculating master pattern according to linear time varying system consistent asymptotic stability criterion；4th computing unit is used for the performance boundary of analytical standard model.The present invention can restrain design verification for large envelope flight control and Aircraft Conceptual Design provides foundation.
Description
Technical field
The invention belongs to automatic control technology fields, and in particular to a kind of performance boundary of large envelope flight control rule design
Analysis system.
Background technology
As aircraft task mission range is constantly expanded, flight envelope is increasing.Usual aircraft is in atmosphere
Interior (in 2 myriametres) and atmosphere edge close to space vehicle (2 myriametres to 10 myriametres), the range of flying speed are located at Gao Ya
To between 5 Mach or more of hypersonic speed, flying speed changes greatly velocity of sound, and the kinetic characteristics variation of aircraft is also larger,
Therefore in its flight course, the performance parameter of aircraft, such as structural elasticity cannot be ignored, this sets for Flight Control Law
Meter proposes severe challenge.The control law of existing big envelope curve aircraft relates generally to control stability, performance robustness and Shandong
Stick stability lacks the method analyzed for control law performance boundary.And the control law performance boundary analysis of big envelope curve aircraft
Method can disclose the inherent mechanism of nonlinear system limited performance, provide the performance boundary of more levels, it is intended to will be linear
The performance boundary theory of system is expanded into nonlinear system, the performance boundary analysis theories of nonlinear system is enriched, from engineering
Angle discloses the inherent mechanism of aircraft portrait track following limited performance, and design verification, Yi Jifei are restrained for large envelope flight control
Row device master-plan provides foundation.
Therefore, how to propose it is a kind of meet large envelope flight control rule inherent flight quality calibration scheme be this field
The current problem to be solved of technical staff.
Invention content
It, can not be to big envelope curve aircraft in property in order to solve the prior art in order to solve the above problem in the prior art
The problem of being analyzed when energy limit, the present invention provide a kind of performance boundary analysis system of large envelope flight control rule design,
The system comprises：
First computing unit, first computing unit are configured as winged by what is built in advance using linearization technique is expanded
Row device model conversation is master pattern；
Second computing unit, second computing unit are configured as the master pattern carrying out non-linear scale change
It changes, calculates the frequency domain performance robustness of the master pattern；
Third computing unit, the third computing unit are configured as according to linear time varying system consistent asymptotic stability criterion
Calculate the stability criterion of the master pattern；
4th computing unit, the 4th computing unit are configured as the frequency domain performance robustness of second computing unit
The performance boundary of the master pattern is analyzed with the stability criterion of the third computing unit.
In the optimal technical scheme of the above method, first computing unit is additionally operable to：
The dummy vehicle is linearized using extended linearization method, divides the Unstable Zero of the dummy vehicle
Dynamically；
The time-domain response criterion of the PD feature structures and the dummy vehicle of the dummy vehicle is calculated, described in judgement
Whether dummy vehicle is controllable；
If the dummy vehicle is controllable, nonlinear coordinate transformation is carried out to the dummy vehicle.
In the optimal technical scheme of the above method, " judging whether the dummy vehicle is controllable ", method is：
The PD feature structures of the dummy vehicle and zero initial value time domain response of the dummy vehicle are calculated, judges institute
State dummy vehicle whether with standard type system equivalence, if with the standard type system equivalence, the dummy vehicle is controllable；
Wherein, the PD feature structures of the dummy vehicle are calculated, shown in the following formula of method：
pi(t)、For PD characteristic values ρi(t) right, left PD feature vectors, ck(t),bj(t) it is output input matrix
Correlated components；
Wherein, standard type system is：
In the optimal technical scheme of the above method, " the frequency domain performance robustness for calculating the master pattern ", method
For：
By scalar polynomial differential algorithm calculate the master pattern in the case that Parameter uncertainties PD spectrum and ranks
The perturbation situation of PD feature vectors；
According to the perturbation situation of PD spectrum and ranks PD feature vectors institute is analyzed in conjunction with the time domain response of the master pattern
The norm for stating the time domain performance parameter vector perturbation range of master pattern, calculates the frequency domain performance robustness of the master pattern.
Compared with the immediate prior art, the present invention provides a kind of performance boundary point of large envelope flight control rule design
Analysis system, including the first computing unit, first computing unit are configured as to build in advance using expansion linearization technique
Dummy vehicle be converted into master pattern；Second computing unit, second computing unit are configured as the master die
Type carries out nonlinear coordinate transformation, calculates the frequency domain performance robustness of the master pattern；Third computing unit, the third meter
Unit is calculated to be configured as calculating the stability criterion of the master pattern according to linear time varying system consistent asymptotic stability criterion；4th
Computing unit, the 4th computing unit are configured as the frequency domain performance robustness of second computing unit and the third meter
The stability criterion for calculating unit analyzes the performance boundary of the master pattern.
Above-mentioned technical proposal at least has the advantages that：Technical scheme of the present invention can disclose nonlinear system
Inherent mechanism that can be limited, provides the performance boundary of more levels, it is intended to by the performance boundary theory of linear system to non-linear
It is expanded in system, enriches the performance boundary analysis theories of nonlinear system, from engineering viewpoint, disclose aircraft portrait track following
The inherent mechanism that can be limited, large envelope flight control restrains design verification and Aircraft Conceptual Design provides foundation.
Description of the drawings
Fig. 1 is the structural representation of the performance boundary analysis system of an embodiment of the present invention large envelope flight control rule design
Figure.
Specific implementation mode
In order to make the object, technical scheme and advantages of the embodiment of the invention clearer, below in conjunction with the embodiment of the present invention
In attached drawing, technical scheme in the embodiment of the invention is clearly and completely described, it is clear that described embodiment is
A part of the embodiment of the present invention, instead of all the embodiments.Based on the embodiments of the present invention, those of ordinary skill in the art
The every other embodiment obtained without making creative work, shall fall within the protection scope of the present invention
The preferred embodiment of the present invention described with reference to the accompanying drawings.It will be apparent to a skilled person that this
A little embodiments are used only for explaining the technical principle of the present invention, it is not intended that limit the scope of the invention.
Refering to attached drawing 1, Fig. 1 illustratively gives the performance boundary point of large envelope flight control rule design in the present embodiment
The flow diagram of analysis system.As shown in Figure 1, the performance boundary analysis system that large envelope flight control rule designs in the present embodiment
Including the following contents：
The method of the present invention is converted using nonlinear system to linear system, is carried out using the theory of linear time varying system,
By nonlinear coordinate transformation, system is become into canonical form, gives time-domain and frequency-domain limited performance mechanism and performance boundary meter
It calculates, the limited mechanism of performance robustness calculates, the non-minimum systems compliant asymptotically stability criterion of a quasi-nonlinear calculates, Stable Robust
Property and its limited mechanism calculate implementation method.
Big envelope curve aircraft elastomer vehicle dynamics are modeled, Longitudinal Dynamic Model can be as follows shown in formula：
The model altogether include 11 state of flights, 5 rigid body states V, γ, h, α, Q indicate respectively speed, flight path angle,
Highly, the angle of attack, pitch rate；6 elastomeric phasesFirst three rank Elastic mode and its differential are indicated respectively.
Wherein, ωiFor the natural frequency of Elastic mode, ξiFor damping ratio, m, g, IyyQuality, acceleration of gravity are indicated respectively and around Y-axis
Rotary inertia；L,D,T,M,NiLift, resistance, thrust, pitching moment and broad sense elastic force are indicated respectively.Above-mentioned model with fly
There is complicated non-linear relation in row state, control input, can be indicated with high-precision model of fit, public affairs specific as follows
Shown in formula (7)：
Wherein,Indicate Elastic mode vector, δ=[δc,δe]TIndicate angle of rudder reflection vector, δc,
δeThe angle of rudder reflection of canard and elevator is indicated respectively,Indicate dynamic pressure, ρ is atmospheric density.
The Aerodynamic Coefficient of big envelope curve aircraft can be as shown in formula (8)：
Wherein, CT,φIt is thrust to the thrust coefficient of roll angle,Respectively CT,φTo the angle of attack
3 ranks, 2 ranks, 1 order derivative and deviant；CTIt is thrust to the thrust coefficient of the angle of attack,Respectively CTTo the angle of attack
3 ranks, 2 ranks, 1 order derivative and deviant；
CLFor lift coefficient, with the angle of attack(Elastic mode vector), δ=[δc,δe]T(rudder
Drift angle vector) related, δc,δeThe angle of rudder reflection of canard and elevator is indicated respectively,It is lift to the lift coefficient of the angle of attack,
It is lift to δeLift coefficient,It is lift to δcLift coefficient,For zero lift coefficient,It is lift to springform
The lift coefficient of state vector η.
CDFor resistance coefficient, with the angle of attack(Elastic mode vector), δ=[δc,δe]T(rudder
Drift angle vector) related, δc,δeThe angle of rudder reflection of canard and elevator is indicated respectively,Angle of attack quadratic sum is attacked for resistance
The resistance coefficient at angle,It is resistance to δeQuadratic sum δeResistance coefficient,It is resistance to δcQuadratic sum
δcResistance coefficient,It is zero resistance coefficient,It is resistance to the resistance coefficient of Elastic mode vector η.
CMFor pitching moment coefficient, with the angle of attack(Elastic mode vector), δ=[δc,
δe]T(angle of rudder reflection vector) related, δc,δeThe angle of rudder reflection of canard and elevator is indicated respectively,Attack against each other for pitching moment
The pitching moment coefficient coefficient of the angle quadratic sum angle of attack,It is pitching moment to δePitching moment coefficient,For pitching moment
To δcPitching moment coefficient,For zero pitching moment coefficient,Pitching for pitching moment coefficient to Elastic mode vector η
Torque coefficient.
L,D,T,M,NiLift, resistance, thrust, pitching moment and broad sense elastic force are indicated respectively.For thrust T, lift
L, M pairs of resistance D, pitching momentThrust coefficient, lift coefficient, the resistance of (Elastic mode vector)
Coefficient, pitching moment coefficient.For broad sense elastic force pairThe broad sense bullet of (Elastic mode vector)
Property force coefficient.
The canard of big envelope curve aircraft is connected by hinge with elevator, to eliminate non-minimum phase characteristic, therefore is flown
The control input to be designed of row device can be u=[δe,φ]T, φ is engine fuel equivalent proportion.Control output can be y=
[V,h]T, that is, control speed and height.
Aircraft mathematical model is typical multivariable, strong nonlinearity, a strong-coupling model, due to aircraft flight speed
Degree is fast, flight envelope is big, the complicated mechanisms such as scramjet engine burning, and lacks enough test flight data supports, control
There are serious uncertainty in simulation, belonging to typically has the probabilistic system of labyrinth.And it is uncertain right
Model characteristics have significant impact.By taking ablated configuration as an example, process is reentered with quick speed and height change, causes to move
PressureFast time variant and uncertain effect, the variation of dynamic pressure there is significant impact to flight characteristics.
Specifically, the performance boundary analysis method of large envelope flight control of the invention rule design includes the following steps：
Step S1：Calculate time-domain and frequency-domain limited performance mechanism and performance boundary；
Step S11：Using extended linearization method by the method for Differential Geometry, nonlinear system is led at equalization point
The method linearisation for crossing differomorphism, Unstable zero-dynamics are split, and zero dy namics are a concepts in nonlinear theory,
The zero being equivalent in transmission function, Unstable zero-dynamics system are also referred to as non-minimum phase system.To analog signal system
Speech, it is non-minimum phase system that all system functions have the system of one or more zeros in the right half plane of S planes, non-
In the control of minimum phase system, need to inhibit to bear the regulating time for adjusting and simultaneously shortening system caused by unstable zero,
Original system is set to be converted into linear time varying systemWherein, x is state variable, and A (t) is state equation, B
(t) it is input matrix, u is inputted in order to control, and t is the time, completes the standardization of system model；
Step S12：Using linear time varying system differential algebra spectral theory, item is wanted using fully controllable the filling of systems compliant
Part, i.e., with standard type systemEquivalence, z are the state variable of standard type system, the dimension of z and original system
The dimension of system is consistent, Az(t) be standard type system sytem matrix, Bz(t) be standard type system input matrix, t is time, u
Indicate control input, wherein
By judging to judge that can system consistent fully controllable, and then study unstable with standard type system equivalence
The confinement mechanism that zero dy namics configure frequency-domain index PD (Parallel D-Eigenvalue) characteristic values and PD feature vectors is
System cannot be fully controllable, then PD Characteristic Structure Configurations are necessarily limited.Wherein, PD characteristic values and PD feature vectors are nonlinear systems
In concept, it is consistent with the concept of characteristic value and feature vector in linear system.When judging that systems compliant is fully controllable,
Next step can be entered.
Step S13：According to PD feature structures and system zero initial value time domain response
Wherein, pi(t)、Respectively PD characteristic values ρi(t) right, left PD feature vectors, ck(t),bj(t) it is respectively
The row k vector of output matrix c, j-th of vector of input matrix b, uj(τ) indicates that control j-th of row vector of variable, K indicate
The line number of output matrix c, N indicate characteristic value PD characteristic values ρi(t) columns of feature vector, the number that m is inputted in order to control, t0
Indicate initial time.Relationship of the time-domain response criterion (rise time, overshoot, negative tune) with PD feature structures is calculated, analysis is unstable
Determine confinement mechanism of the zero dy namics to system time-domain response criterion：Single unstable zero z and single shakiness existing for system are provided simultaneously
When determining pole p, system overshoot, the negative limit adjusted are respectively：
Step S14：For the system of flight dynamics, it is based on time-scale separation thought, by aircraft manufacturing technology system point
For the attitude angle circuit changed slowly and the angular speed circuit changed soon, fast circuit is generally minimum phase system, corresponding PD
Characteristic value and PD feature vectors can be set as definite value
The step of method described in conjunction with the examples disclosed in this document or algorithm, can use hardware, processor to execute
The combination of software module or the two is implemented.Software module can be placed in random access memory (RAM), memory, read-only memory
(ROM), electrically programmable ROM, electrically erasable ROM, register, hard disk, moveable magnetic disc, CD-ROM or technical field
In any other form of storage medium well known to interior.
Those skilled in the art should be able to recognize that, side described in conjunction with the examples disclosed in the embodiments of the present disclosure
Method step, can be realized with electronic hardware, computer software, or a combination of the two, in order to clearly demonstrate electronic hardware and
The interchangeability of software generally describes each exemplary composition and step according to function in the above description.These
Function is executed with electronic hardware or software mode actually, depends on the specific application and design constraint of technical solution.
Those skilled in the art can use different methods to achieve the described function each specific application, but this reality
Now it should not be considered as beyond the scope of the present invention.
So far, it has been combined preferred embodiment shown in the drawings and describes technical scheme of the present invention, still, this field
Technical staff is it is easily understood that protection scope of the present invention is expressly not limited to these specific implementation modes.Without departing from this
Under the premise of the principle of invention, those skilled in the art can make the relevant technologies feature equivalent change or replacement, these
Technical solution after change or replacement is fallen within protection scope of the present invention.
Claims (4)
1. a kind of performance boundary analysis system of large envelope flight control rule design, which is characterized in that the system comprises：
First computing unit, first computing unit are configured as the aircraft that will be built in advance using linearization technique is expanded
Model conversation is master pattern；
Second computing unit, second computing unit are configured as the master pattern carrying out nonlinear coordinate transformation, meter
Calculate the frequency domain performance robustness of the master pattern；
Third computing unit, the third computing unit are configured as being calculated according to linear time varying system consistent asymptotic stability criterion
The stability criterion of the master pattern；
4th computing unit, the 4th computing unit are configured as frequency domain performance robustness and the institute of second computing unit
The stability criterion for stating third computing unit analyzes the performance boundary of the master pattern.
2. system according to claim 1, which is characterized in that first computing unit is additionally operable to：
The dummy vehicle is linearized using extended linearization method, the Unstable Zero for dividing the dummy vehicle is dynamic
State；
The time-domain response criterion for calculating the PD feature structures and the dummy vehicle of the dummy vehicle, judges the flight
Whether device model is controllable；
If the dummy vehicle is controllable, nonlinear coordinate transformation is carried out to the dummy vehicle.
3. system according to claim 2, which is characterized in that " judging whether the dummy vehicle is controllable ", method
For：
The PD feature structures of the dummy vehicle and zero initial value time domain response of the dummy vehicle are calculated, judges described fly
Row device model whether with standard type system equivalence, if with the standard type system equivalence, the dummy vehicle is controllable；
Wherein, the PD feature structures of the dummy vehicle are calculated, shown in the following formula of method：
Wherein, K indicates that the line number of output matrix c, N indicate characteristic value PD, characteristic value ρi(t) columns of feature vector, m are control
Make the number of input, t0Indicate initial time, pi(t)、For PD characteristic values ρi(t) right, left PD feature vectors, ck(t),
bj(t) be respectively output matrix c row k vector, j-th of input matrix b vector, uj(τ) indicates control j-th of row of variable
Vector；
Wherein, standard type system is：
4. system according to claim 3, which is characterized in that " the frequency domain performance robustness for calculating the master pattern ",
Its method is：
It is special in the PD spectrums and ranks PD of Parameter uncertainties that the master pattern is calculated by scalar polynomial differential algorithm
Levy the perturbation situation of vector；
According to the perturbation situation of PD spectrums and ranks PD feature vectors the mark is analyzed in conjunction with the time domain response of the master pattern
The norm of the time domain performance parameter vector perturbation range of quasi-mode type, calculates the frequency domain performance robustness of the master pattern.
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| CN113253616A (en) * | 2021-06-29 | 2021-08-13 | 中国科学院自动化研究所 | Flight control method and device for large envelope of fast time-varying aircraft |
| CN115128966A (en) * | 2022-04-11 | 2022-09-30 | 厦门大学 | Design method and simulation method of turbofan engine full-envelope controller |
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| CN115128966A (en) * | 2022-04-11 | 2022-09-30 | 厦门大学 | Design method and simulation method of turbofan engine full-envelope controller |
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