KR102093744B1 - Method of parameter optimization for ensuring longitudinal axis stability and flying quality for aircraft - Google Patents

Abstract

The present invention is to design a target dynamics configured to remove dynamic characteristics of an aircraft and to calculate a dynamic demand input, and calculate a mathematical nonlinear model of an aircraft by inputting the difference between the angular acceleration measured from the angular acceleration sensor and the dynamic demand input as input. It relates to a parameter optimization method for satisfying the stability and flightability of the vertical axis of the aircraft, including the step of optimizing the parameters of the controller of the target dynamics and designing an inverse motion model represented by angular acceleration by removing it with (Inversion).
The parameter optimization method for satisfying the stability and flightability of the vertical axis of the aircraft according to the present invention can satisfy the flightability margin based on the incremental dynamic reverse conversion control method using the angular acceleration sensor, and improve the phase margin and equivalent time delay characteristics It has the effect of satisfying the stability at the same time, making it easy to obtain airworthiness certification.

Description

How to optimize parameters to meet the stability and flightability of the vertical axis of the aircraft {METHOD OF PARAMETER OPTIMIZATION FOR ENSURING LONGITUDINAL AXIS STABILITY AND FLYING QUALITY FOR AIRCRAFT}

The present invention relates to a method for optimizing the parameters for the stability and flightability of the vertical axis of the aircraft, and more specifically, an incremental nonlinear dynamic inversion (INDI) control method using angular acceleration information measured from a sensor. It is about how to design and optimize.

Modern high-performance fighter planes are designed with the concept of Relaxed Static Stability (RSS) to ensure high maneuverability. Therefore, a flight control system (FCS) is designed to provide stability and flying qualities suitable for missions to unstable aircraft.

Furthermore, advanced control theory that complements the classic control or the classic control that has been verified due to limitations such as computational throughput of the flight control computer and memory storage space has been applied to mass-produced aircraft aiming for flight safety in addition to maneuvering performance. .

In particular, when the physical properties are changed during the operation of the aircraft, there is a problem in that robust control cannot be achieved in response to changes in the physical properties that are changed in a classical manner. In order to solve this problem, a technique for minimizing the change in the center of gravity is disclosed in Korean Patent No. 1,445,221.

However, as a conventional technique, it is difficult to actively change the physical characteristics of an aircraft during operation, and even if a model simulated for the flight characteristics of a vehicle is selected through wind tunnel tests, it is different from the characteristics of actual flight. There is a problem that the Anseongjeong is lowered due to.

Republic of Korea Registered Patent No. 1,445,221

An object of the present invention is to provide a method of optimizing control parameters that satisfy flightability and stability of a vertical axis in a sensor-based nonlinear dynamic model inverse transformation control method of an aircraft in response to a changing center of gravity of a conventional aircraft.

As a solution to the above problems, a step of designing a target dynamics configured to remove dynamic characteristics of an aircraft and to calculate a dynamic demand input, and inputting a difference between an angular acceleration measured by an angular acceleration sensor and a dynamic demand input as an input to develop a mathematical nonlinear model of the aircraft. A parameter optimization method for satisfying stability and flightability of the vertical axis of the aircraft may be provided, including the step of designing and removing an inversion to optimize the parameters of the controller of the target dynamics and designing an inverse motion model represented by angular acceleration.

Here, the controller of the target dynamics may be configured to calculate a pitch angular acceleration command by a pilot's vertical acceleration command, an aircraft's vertical acceleration measured by a sensor, and a pitch angular velocity feedback input.

On the other hand, the parameters of the controller of the target dynamics may be determined by setting the stability margin according to a predetermined criterion as the highest constraint (Hard Constraint) and the flightability design goal as the lower constraint (Soft Constraint).

In addition, the controller of the target dynamics,

는 조종사의 수직가속도 명령이며,

는 센서로부터 측정되는 항공기의 수직가속도이며,

는목표동역학으로부터 산출되는 목표로 하는 피치 각가속도 명령이며,

는 피치 각속도 궤환이며,

,

,

및

는 제어이득 파라미터일 수 있다.Designed as, where:

Is the pilot's vertical acceleration command,

Is the vertical acceleration of the aircraft measured from the sensor,

Is a target angular acceleration command calculated from target dynamics,

Is the pitch angular velocity feedback,

,

,

And

May be a control gain parameter.

In addition, the target dynamics may be designed to further include a lead compensator in the pilot's vertical acceleration command stage.

Meanwhile, the stability margin may include gain margin and phase margin.

In addition, the flight design target is short-period, and in some cases, may include a design target for the long-period mode.

,

,

로 결정될 수 있다.Meanwhile, the initial value of the parameter

,

,

Can be determined as.

The parameter optimization method for satisfying the stability and flightability of the vertical axis of the aircraft according to the present invention can satisfy the flightability margin based on the incremental dynamic reverse conversion control method using the angular acceleration sensor, and improve the phase margin and equivalent time delay characteristics It has the effect of satisfying the stability at the same time, making it easy to obtain airworthiness certification.

1 is a flow chart of a parameter optimization method for satisfying the stability and flightability of an aircraft in accordance with the present invention.
FIG. 2 is a detailed flowchart of the parameter optimization step of FIG. 1.
3 is a block diagram of an aircraft control system according to the present invention.
FIG. 4 is a block diagram of target dynamics for deriving the dynamic control gain of FIG. 1.
5 is a table showing an example of the flight parameter optimization results.
6 is a table showing flightability and stability margin according to case.
7 is a view showing a Bode diagram according to a design and optimization method.
8 is a graph showing response characteristics according to time at the time of unit input.

Hereinafter, a method for optimizing parameters for satisfying the stability and flightability of an aircraft in an embodiment of the present invention will be described in detail with reference to the accompanying drawings. In the following description of the embodiments, the names of each component may be referred to as other names in the art. However, if they have functional similarity and identity, even if a modified embodiment is employed, it can be regarded as an even configuration. In addition, symbols added to each component are described for convenience of description. However, the illustrated contents on the drawings in which these codes are described do not limit each component to the range in the drawing. Similarly, even if an embodiment in which some modifications of the configuration on the drawing is employed, it can be regarded as an equivalent configuration if there is functional similarity and identity. In addition, in light of the general technical level of the technical field, if it is recognized as a component that should be included, a description thereof will be omitted.

1 is a flowchart of a parameter optimization method for satisfying the stability and flightability of an aircraft according to the present invention, FIG. 2 is a detailed flowchart of the parameter optimization step of FIG. 1, and FIG. 3 is a block diagram of an aircraft control system according to the present invention And FIG. 4 is a block diagram of target dynamics for deriving the dynamic control gain of FIG. 1.

As shown in the figure, the parameter optimization method for meeting the vertical axis stability and flightability of the aircraft according to the present invention allows the output of the aircraft by the control input generated through the target dynamics (1) and the dynamic inversion (2) to be designed independently. It can be designed to meet performance stability.

The parameter optimization method for satisfying the stability and flightability of the vertical axis of the aircraft according to the present invention may include a target dynamics design step (S100), a reverse motion model design step (S200), and a parameter optimization step (S300).

The target dynamics design step (S100), as shown in FIG. 4, the controller may be designed in a proportional-integral manner with respect to the longitudinal axis. The controller consists of flightability parameters, sets initial values based on the flightability design goals, and configures the parameters to satisfy the design goals in a linear control system that includes aircraft models, sensors, and drivers through a subsequent optimization step. do.

The controller of the target dynamics 1 can be designed with the following control laws.

는 조종사의 수직가속도 명령이며,

는 센서로부터 측정되는 항공기의 수직가속도이며,

는목표동역학으로부터 산출되는 목표로 하는 피치 각가속도 명령이며,

는 피치 각속도 궤환이며,

,

,

및

는 제어이득 파라미터를 나타낸다.here,

Is the pilot's vertical acceleration command,

Is the vertical acceleration of the aircraft measured from the sensor,

Is a target angular acceleration command calculated from target dynamics,

Is the pitch angular velocity feedback,

,

,

And

Indicates the control gain parameter.

From the relationship between the vertical acceleration of the center of gravity and the linear system, the angle of attack can be expressed as the pitch angular velocity as follows.

,

,

,

,

는 항공기 속도,

는 중력가속도 및

는 피치자세각 시정수(pitch attitude time constant)로서 항공기 모델로부터 획득할 수 있다.here

Aircraft speed,

Is the gravitational acceleration and

Is the pitch attitude time constant, which can be obtained from the aircraft model.

When the relational expression is applied to the relational expression of the target dynamics by laplace transform, it can be expressed as follows.

,

,

,

이다.here

,

,

,

to be.

Meanwhile, each parameter can be determined as follows.

,

,

,

,

When the determined parameter is applied to the transfer function, it can be expressed as follows.

에 의한 효과를 상쇄하기 위하여

를 다음의 식으로 정의할 수 있다.And, molecular protest

To offset the effect of

Can be defined by the following equation.

In addition, the controller of the target dynamics 1 may be configured to include a lead compensator 11 in the vertical axis control command stage to improve the control characteristics. Here, except in special cases, the zero of the system is not affected by the feedback system and is determined by the aerodynamic characteristics of the aircraft itself. However, since the position of the zero point affects the steering characteristics, it must be designed to have appropriate values in order to obtain the desired steering characteristics. When the pole-zero cancelation is applied in the transfer function of the true compensator 11, the following can be obtained.

,

의 파라미터로 설정한다.From this, it is possible to satisfy the design goal by improving the steering characteristics using the compensator 11 for any feedback system. here,

,

Parameter.

The inverse motion model design step S200 is a step of calculating and removing a model of a nonlinear aircraft. The inverse motion model is expressed by angular acceleration. The inverse motion model is configured to obtain and control angular acceleration information directly from the angular acceleration sensor mounted on the aircraft.

Specifically, calculating the nonlinear dynamic model inverse transform control input can be expressed as follows.

는 상태 백터(state vector)이고,

는 제어입력벡터(control input vector)이다. 일반적으로 항공기 내부루프의 상태벡터 변수로는 3축의 가속도를 사용한다. 따라서 3축 가속도를 제어변수(control variable)로 선정하여 상태벡터를 구성할 수 있다.here

Is a state vector,

Is a control input vector. In general, three-axis acceleration is used as the state vector variable of the aircraft's inner loop. Therefore, a state vector can be constructed by selecting 3-axis acceleration as a control variable.

및 제어입력

에 대하여 Taylor 시리즈로 전개하고, 간략화하면 다음과 같이 표현이 가능하다.The nonlinear motion equation described above is the current state variable

And control input

For Taylor series, it can be expressed as follows.

이며,

이다. 상기 식을 제어면 변위에 대하여 표현하면 다음과 같다.here

And

to be. If the above expression is expressed with respect to the control plane displacement,

는

에서의 각가속도가 되며,

에서의 각가속도 (

)에 대한 함수로 표현가능하다. 여기서 B의 역행렬이 존재하면 제어입력

는 다름과 같이 표현될 수 있다.here

The

Is the angular acceleration from,

Angular acceleration at

). Here, if the inverse matrix of B exists, control input

Can be expressed as

는 항공기 설계 요구조건을 반영한 동적 특성으로 목표동역학이 되며, 이는 항공기 설계 요구조건에 따라 다양한 형태로 설계된다.here

Is a dynamic characteristic that reflects the aircraft design requirements and becomes the target dynamics, which is designed in various forms according to the aircraft design requirements.

When an inverse matrix is present, substituting the control input into the nonlinear motion equation is as follows.

Therefore, the dynamic characteristics of the existing aircraft are removed, and the aircraft follows characteristics according to the designer or aircraft requirements. After all, if the inverse matrix of B exists, it is possible to design the target dynamics part independently of the dynamic characteristics of the aircraft.

Here, since 3-axis angular acceleration information is obtained directly from the sensor, uncertain information that cannot be expressed mathematically accurately for changes in the center of gravity caused by fuel consumption or armed dropping can be used, so it is used as a very robust control method for model uncertainty. Can be.

The parameter optimization step (S300) is configured to optimize the parameter value that satisfies both the stability requirement and the flightability requirement.

2, the parameter optimization step (S300) is a flight condition setting step (S310), stability and flightability demand setting step (S320), initial value setting step (S330), stability satisfaction determination step (S340), It may be configured to include a flight performance satisfactory determination step (S350), a parameter update step (S360).

The flight condition setting step (S310) is a step of setting flight conditions for finding the parameter setting value of the aircraft. A specific flight condition may be selected, and, for example, conditions for speeds of Mach number 0.8 and 5000 ft may be set.

The stability and flightability demand setting step (S320) corresponds to a step of setting a reference value of performance to be satisfied. Stability and flightability requirements can be set to the requirements presented by MIL-HDBK-513C, the airworthiness certification.

The initial value setting step (S330) corresponds to a step of setting the initial value in order to derive a rapid optimization result among a number of parameter setting values.

In the above transfer function, the characteristic equation of the closed loop can be expressed as follows.

는 단주기 모드 감쇄비(damping ratio)이며,

는 고유진동수(natural frequency)를 나타낸다.here

Is the short cycle mode damping ratio,

Denotes the natural frequency.

From the transfer function and characteristic equation, the initial value of the flightability parameter can be determined as described above.

,

,

,

,

In the step of determining whether the stability and flight performance is satisfied (S340), the stability margin is set to Hard Constraint, and the flight design criteria are set to Soft constraint to evaluate whether the design criteria are satisfied.

Whether or not stability is satisfied corresponds to a step of determining whether or not the above-described stability requirement is satisfied when performing control of the aircraft by applying a set parameter value.

For example, whether stability is satisfied may be determined by gain margin, phase margin, eigenvalue, and Nichols Margin.

Whether or not flight performance is satisfied may be performed when stability is satisfied when the currently set parameter value is applied by performing stability satisfaction. The determination of whether or not flight performance is satisfied may be performed by determining whether or not the flight performance requirement is satisfied when performing control of the aircraft by applying a set parameter value.

Examples of flight design criteria can be Tier # 1 short-cycle mode attenuation ratio, frequency, control anticipation parameter (CAP), equivalent time delay, and dropback. In some cases, design goals for long-cycle mode It may include.

The parameter update step (S350) corresponds to a step of updating the parameter to satisfy both stability satisfaction and flight satisfaction, and increases the update direction, that is, the parameter in consideration of stability and flight tendency according to the change of parameters. The decreasing direction can be determined and updated.

On the other hand, the stability and flight performance satisfaction determination step (S340) to the update step (S350) may be repeatedly performed when the parameter values satisfying both stability and flightability requirements are determined.

Hereinafter, results according to optimization of parameters according to the present invention will be described with reference to FIGS. 5 to 8.

Data for the three cases will be described below, case 1 is a control group when the short-term mode flightability design target is set to Soft Constraint in the optimization step (S300), case 2 and case 3 are the present invention. This is an example according to the embodiment. In case 2, the stability margin and the Nichols Margin design standard are set to Hard Constraint, and the flightability design target is selected as Soft Constraint, case 3 designs the fact compensation 11 based on the flightability parameters calculated from case 1. The evaluation of the case (C2) when the parameters of the true phase compensator 11 is optimized by the optimization method of 2 is shown.

5 is a table showing an example of the flight parameter optimization results. Here, in case 2, the pitch angular velocity feedback flightability parameter was slightly reduced compared to case 1, but the flightability and phase margin were satisfied. Case 3 is the result of designing a leading edge compensation filter and optimizing the coefficients to improve unsatisfactory stability margins while maintaining the combination of flight parameters of case 1.

6 is a table showing flightability and stability margin according to case. As shown, case 1 satisfies flight level 1, but the phase margin is 39.3 ㅀ, which does not satisfy 45 ㅀ or higher. Therefore, as in case 2, if the stability margin and flightability are the design objectives and the flightability parameters are optimized, it can be seen that the results satisfying the flightability design objectives and stability margins can be obtained. In case 3, it is possible to secure stability by designing the true compensator 11 so as to secure stability in case 1 that satisfies flightability.

7 is a view showing a Bode diagram according to a design and optimization method. As shown, it can be seen that the phase margin is improved by pulling the phase while maintaining the gain crossing frequency of the control system, and the phase margin is improved from the viewpoint of the overall control system, while short-cycle attenuation within the flightability level 1 It can be seen that the ratio decreases. In addition, the phase margin and gain margin are inversely related, so if the phase margin increases, the gain margin exhibits a relatively decreasing characteristic, so flight and stability can be interpreted as inverse relations. .

8 is a graph showing response characteristics according to time at the time of unit input. As shown, it can be seen that in case 2 and case 3, there is a change in the response characteristic in the transient response characteristic and the stabilization step due to the reduction in the attenuation ratio. However, in the stabilization stage, it can be confirmed that it satisfies the level of flightability by stabilizing immediately with one overshoot response.

As described above, the parameter optimization method for satisfying the stability and flightability of the aircraft's vertical axis according to the present invention applies a nonlinear dynamic model inverse transformation control method, and when designing a controller, it is required through optimization of parameters according to stability and flightability requirements. Since it can satisfy the characteristics of the aircraft, it can exert the effect of satisfying the airworthiness certification.

1: Target dynamics
11: True Compensation
2: Inverse motion model
S100: Target dynamics design phase
S200: Design phase of inverse motion model
S300: parameter optimization step
S310: Flight condition setting step
S320: Stability and flight demand setting stage
S330: initial value setting step
S340: Stability and flight performance satisfaction determination step
S350: parameter update step

Claims (7)

Designing a target dynamics configured to remove dynamic characteristics of the aircraft and to calculate dynamic demand inputs;
Calculating a mathematical nonlinear model of the aircraft by taking the difference between the angular acceleration measured from the angular acceleration sensor and the dynamic demand input as an input, and removing it by inversion to design an inverse motion model represented by angular acceleration; And
Optimizing the parameters of the controller of the target dynamics,
The controller of the target dynamics,
Pilot's vertical acceleration command;
Vertical acceleration of the aircraft measured from a sensor;
It is configured to calculate the pitch angular acceleration command by inputting the pitch angular velocity feedback.
The parameters of the controller of the target dynamics,
The stability margin according to the predetermined standard is set as the highest constraint (Hard Constraint),
A method of optimizing the parameters to meet the stability and flight characteristics of the vertical axis of the aircraft, which is determined by designing the flight objective as a soft constraint. delete According to claim 1,
The controller of the target dynamics,

Designed as
here,

Is the vertical acceleration command of the pilot,

Is the vertical acceleration of the aircraft measured from the sensor,

Is a target angular acceleration command calculated from the target dynamics,

Is the pitch angular velocity feedback,
remind

,

,

And

Is a control gain parameter, a parameter optimization method for meeting the stability and flightability of the vertical axis of the aircraft. According to claim 1,
The target dynamics is a parameter optimization method for satisfying the stability and flightability of the vertical axis of the aircraft, characterized in that the pilot further comprises a lead compensator in the vertical acceleration command stage. According to claim 4,
The stability margin is a parameter optimization method for meeting the stability and flight characteristics of the vertical axis of the aircraft, characterized in that it includes a gain margin and a phase margin. According to claim 4,
The flightability design goal is a short-period (short-period) mode, the method of optimizing the parameters for the stability and flight performance of the vertical axis of the aircraft, characterized in that it comprises a target. According to claim 3,
The initial value of the parameter

,

,

Characterized in that the method of optimizing the parameters to meet the stability and flightability of the aircraft vertical axis.

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