KR102093745B1 - Control method for improving flying qualities and stability corresponding to longitudinal cg travel in flight operation - Google Patents

Abstract

Between the first mode control step of generating the predicted angular acceleration of the aircraft as the output, the second mode control step of generating the angular acceleration measured value as the output, and the predicted angular acceleration of the first mode control step and the measured angular acceleration of the second mode control step The angular acceleration error calculation step of calculating the error, the additional control gain design step of determining the additional augmented gain and applying it to the error between the angular accelerations, the step of limiting the amount of augmentation after applying the additional gain, and the error between the angular accelerations of which the amount of augmentation is limited A control method for securing flightability of an aircraft that copes with a changing center of gravity during flight may be provided, which includes adding to the 1 mode control step.
The control method for securing the stability of the aircraft to cope with the changing center of gravity according to the present invention can determine the augmented gain according to the estimated angular acceleration and the angular acceleration error measured from the actual sensor by reflecting the position information of the center of gravity. Even if a failure of the error and the center of gravity sensor occurs, it is effective in securing flightability and stability.

Description

CONTROL METHOD FOR IMPROVING FLYING QUALITIES AND STABILITY CORRESPONDING TO LONGITUDINAL CG TRAVEL IN FLIGHT OPERATION}

The present invention relates to a control method for improving flightability and stability of an aircraft that copes with a changing center of gravity, and more specifically, a control method for reflecting an additional augmentation algorithm to control in response to a position of the center of gravity of a changing aircraft. It is about.

During operation, the position of the center of gravity changes due to various factors such as consumption of fuel and use of armament. This change in the center of gravity degrades flightability and causes a difference between the maneuvering of the pilot and the maneuvering of the actual aircraft.

As a prior art, a technique of controlling fuel transfer using software is applied to minimize changes in the center of gravity of the aircraft. The technology for controlling the fuel transfer is a technology that minimizes the movement of the center of gravity by appropriately controlling the amount of fuel in the fuel tank that is distributed inside the aircraft. Korean Patent No. 1,445,221 is disclosed as a technique for minimizing the change in the existing center of gravity.

However, the prior art cannot provide cracked flightability by coping with the shifting of the center of gravity during flight, and ensures flightability and stability when the position of the center of gravity is out of the limited range due to malfunction / misoperation of the fuel system. There was a problem that could not be done.

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

The present invention provides a uniform flightability of an aircraft and satisfies stability by coping with a case where the position of the center of gravity deviates from a limited range due to a change in the center of gravity within the limited range during flight, failure / misoperation of the fuel system, etc. The purpose is to provide a control method.

)를 출력으로 발생시키는 제1 모드 제어 단계, 각가속도 센서 또는 항공기의 상태 정보로부터 직접 측정된 각가속도(

)를 출력으로 발생시키는 제2 모드 제어 단계, 두 각가속도 정보의 오차를 산출하고 평가하는 단계(

), 추가적인 증강 알고리즘의 제어 이득 설계 단계(

), 비행제어시스템 안정성을 고려하여 증강의 양을 제한하는 제한 단계, 계산된 증강 알고리즘(

)을 제1 모드 제어 단계에 합산하는 단계(

)로, 비행 중 변화하는 무게중심에 대처하여 항공기에 균일한 비행성을 제공하고 안정성을 확보를 위한 제어방법이 제공될 수 있다.As a solution to the above problems, the angular acceleration predicted based on the motion model of the aircraft (

), The first mode control step of generating the output, the angular acceleration measured directly from the angular acceleration sensor or the aircraft's status information (

), The second mode control step of generating the output, calculating and evaluating the error of the two angular acceleration information (

), Additional gain algorithm control gain design step (

), A limiting step to limit the amount of augmentation in consideration of flight control system stability, and a calculated augmentation algorithm (

) To the first mode control step (

), It is possible to provide a uniform flightability to the aircraft by coping with the changing center of gravity during flight and a control method for securing stability.

Here, the control gain design and the limiting step of the additional augmentation algorithm can be variably designed according to the magnitude of the error of the two angular acceleration information to adjust the amount of augmentation to cope with the normal and defective situation of the center of gravity sensor.

The control method for securing the flightability and stability of the aircraft in response to the changing center of gravity according to the present invention calculates the error between the calculated angular acceleration and the measured angular acceleration and designs and compensates additional augmentation algorithms to compensate for the center of gravity error during flight and Even if a failure of the center of gravity sensor occurs, it is effective in securing flightability and stability.

1 is a view showing the concept of the impact of the static stability (Static Margin, SM) of the aircraft on the change in the center of gravity during flight of the aircraft.
2 is a conceptual diagram of a design that maintains a center of gravity within a certain range through cause and fuel management that affects a change in the center of gravity of an aircraft.
3 is a flowchart of the first embodiment according to the present invention.
4 is a block diagram of a control system according to the present invention.
5 is a block diagram of a control system according to another embodiment of the present invention.

Hereinafter, a control method for improving flightability and stability of an aircraft that copes with a changing center of gravity 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.

Hereinafter, with reference to FIGS. 1 and 2, the impact on the change in the center of gravity of the aircraft will be described.

FIG. 1 is a concept of the effect of an aircraft's static margin (SM) on a change in the center of gravity during flight of the aircraft, and FIG. 2 is a schedule of the center of gravity through fuel management and causes affecting the change in the center of gravity of the aircraft. It is a conceptual diagram of the design to keep within the scope.

Changes in the weight of a fighter during flight include fuel inflation, external armed fire, external fuel tank separation, and fuel inclination and air due to high inertia movement such as high angle of attack (HAOA) roll motion. And changes in aerodynamics, center of gravity and inertia of the aircraft caused by refueling. In the design of the aircraft, the change in the center of gravity according to the operation is defined as the center of gravity area diagram (CG envelope), and in the center of gravity area diagram, the position of the front and rear centers of gravity and the increase in movement of the center of gravity in normal flight without defects are defined. The margin is reflected. This includes uncertainties such as weight distribution, stability levels, fluctuations in manufacturing, addition of planned equipment, payload, flight posture, fuel density, high angle of attack, roll and acceleration, and uncertainties such as fuel skew and future of operating fixtures. This is to ensure the possibility of volatility. In addition, the maximum achievable excursions of the system or major components, such as a significant heavy store, are determined.

On the other hand, the position of the center of gravity can be estimated based on data obtained from multiple sensors measured in flight. This can be achieved by using accurate and reliable center of gravity information through the VMS (Vehicle Management System) and FMS (Fuel Management System) when operating the aircraft.

Hereinafter, a control method for improving flightability and stability of an aircraft that copes with a changing center of gravity, which is an embodiment according to the present invention, will be described in detail with reference to FIGS. 3 to 4.

3 is a flowchart of a first embodiment according to the present invention, and FIG. 4 is a block diagram of a control system according to the present invention.

As shown, the control method for improving the flightability and stability of the aircraft to cope with the changing center of gravity according to the present invention is a first mode control step (S100), a second mode control step (S200), angular acceleration error calculation step (S300), a control gain design step of additional augmentation algorithm (S400), a limiting step (S500) of limiting the amount of augmentation, and a step of adding the calculated augmentation algorithm to the first mode control step (S600). You can.

The first mode control step (S100) is configured to generate a control input of the driver of the aircraft as an output according to the pilot's control input. The first mode control step (S100) is performed based on the On Board Model (OBM), and the OBM may include a plurality of sub-control models set before the flight of the aircraft.

가 될 수 있다. 한편, 다수의 서브 제어 모델은 다양한 조건에 의해 선택될 수 있도록 구성될 수 있으므로 이하에서는 무게중심의 위치에 따른 선택에 대하여 설명하고, 다른 조건, 예를들어 비행영역과 같은 조건에 대한 선택사항에 대한 설명은 생략하도록 한다. The first mode control step (S100) may be configured to select a sub control model based on a plurality of flight data acquired during flight of the aircraft. In the first mode control step (S100), the angular acceleration is predicted and generated as an output based on the selected aircraft motion model. The angular acceleration predicted here is

Can be On the other hand, a plurality of sub-control models can be configured to be selected by various conditions, so the following describes the selection according to the location of the center of gravity, and the other conditions, for example, options for conditions such as the flight area. The description will be omitted.

Referring to FIG. 4 again, the sub-control model selected in the first mode control step (S100) is configured to exert maneuvering of the aircraft according to the pilot's steering input according to the current state of the aircraft. Specifically, the operation amount of the driver provided in the aircraft is varied to perform a desired maneuvering according to the location of a plurality of centers of gravity, and this operation amount is determined and generated as a control input.

가 된다. The second mode control step S200 is configured to measure the angular acceleration. In the second mode control step S200, the angular acceleration measured directly from the angular acceleration sensor or the state information of the aircraft is generated as an output. Where the angular acceleration as output

Becomes

)가 소정범위를 벗어나는 경우 현재 제1 모드 제어 단계(S100)에서 판단하고 있는 무게중심의 위치에 오차가 발생하는 것으로 추정될 수 있다. 무게중심의 오차가 발생되는 것으로 추정되는 경우에는 후술할 증강 알고리즘의 제어 이득 설계 단계(S400) 및 증강의 양을 제한하는 제한 단계(S500)가 수행될 수 있다.The angular acceleration error calculation step S300 is configured to calculate the difference between the angular accelerations obtained from the first mode control step S100 and the second mode control step S200. Here, the error between the predicted angular acceleration and the measured angular acceleration (

) May be estimated that an error occurs in the position of the center of gravity currently determined in the first mode control step (S100). When it is estimated that an error of the center of gravity is generated, a control gain design step (S400) of the augmentation algorithm to be described later and a limiting step (S500) of limiting the amount of augmentation may be performed.

)을 결정하는 단계에 해당한다. 증강 알고리즘의 제어 이득 설계 단계(S400)는 각가속도 오차의 범위에 따라 추가적인 제어 이득을 설계하게 된다. 결국 잘못 판단된 무게중심의 위치가 제어입력에 미치는 영향을 최소화 할 수 있도록 측정된 각가속도 값과 예측된 각가속도 값의 오차에 따라 각가속도 궤환 값의 이득값이 조절된다.The control gain design step (S400) of the augmentation algorithm is controlled by the augmentation algorithm applied to the angular acceleration feedback signal to maintain flightability uniformly when it is estimated that an error occurs at the position of the center of gravity from the angular acceleration error calculation step (S300). benefit(

). In the control gain design step (S400) of the augmentation algorithm, additional control gains are designed according to the range of the angular acceleration error. As a result, the gain value of the angular acceleration feedback value is adjusted according to the error between the measured angular acceleration value and the predicted angular acceleration value to minimize the influence of the position of the wrongly determined center of gravity on the control input.

The limiting step (S500) of limiting the amount of augmentation is a step of limiting the amount of augmentation gain in consideration of the stability of the flight control system. The aircraft has stability to be secured under each flight condition. For example, in a certain flight condition, the gain margin and the phase margin for the frequency response of the aircraft must be greater than or equal to a predetermined value, and a feedback signal is provided within the limits to satisfy this stability. The final value is limited so that it can be augmented. The amount limiting the augmented gain may be determined within a limit where flight performance can be maintained in consideration of stability that must be satisfied in each flight condition and flight condition, and may be selected from preset values in consideration of stability requirements.

) 값에 증강된 각가속도(

)를 합산하는 단계에 해당한다. 본 단계에서는 무게중심의 측정 또는 추정되는 위치에 오차가 발생하는 경우 제2 모드 제어 단계(S200)에서 출력으로 얻어지는 측정된 각가속도와의 오차를 반영하여 합산하므로 오차를 최소화 할 수 있다. Step S600 of adding the calculated augmentation algorithm to the first mode control step is the estimated angular acceleration that is the output of the first mode control step S100 (

Angular acceleration augmented to the value (

). In this step, when an error occurs in the measurement or the estimated position of the center of gravity, the error can be minimized because it reflects and sums the error with the measured angular acceleration obtained as an output in the second mode control step (S200).

Hereinafter, a second embodiment according to the present invention will be described in detail with reference to FIG. 5.

5 is a block diagram of a control system according to another embodiment of the present invention.

As shown in Figure 5, the control method for securing the stability of the aircraft to cope with the changing center of gravity according to the present invention is a first mode control step, a weight center measurement or estimation step, a weight center information reliability determination step, a switching step And a second mode control step.

The first mode control step is configured to generate a control input of the driver of the aircraft as an output according to the pilot's control input. The first mode control step is performed using an OBM (On Board Model), and may be configured to include a plurality of sub-control models set before the flight of the aircraft.

The first mode control step may be configured to select a sub control model based on a plurality of flight data acquired during flight of the aircraft. Since multiple sub-control models can be configured to be selected by various conditions, the following describes the selection according to the location of the center of gravity, and describes the options for other conditions, for example, conditions such as the flight area. Is omitted.

Referring back to FIG. 5, the sub-control model selected in the first mode control step is configured to exert maneuvering of the aircraft according to the pilot's steering input according to the current state of the aircraft. Specifically, the operation amount of the driver provided in the aircraft is varied to perform a desired maneuvering according to the location of a plurality of centers of gravity, and this operation amount is determined and generated as a control input.

The center of gravity measurement or estimation step is configured to estimate the position of the center of gravity using the center of gravity DB measured from the center of gravity sensor or based on flight data. When the position of the center of gravity is determined, the maneuvering and the state of the aircraft according to the pilot's steering input can be predicted, and in particular, the angular acceleration can be predicted from the response pattern of the aircraft according to the steering input.

The reliability determination step of the center of gravity information corresponds to the step of evaluating the reliability of the position of the center of gravity by comparing the current predicted angular acceleration with the angular acceleration value measured from the sensor. When the error between the predicted angular acceleration and the measured angular acceleration occurs in the weight center information reliability determination step, it is evaluated as low reliability. Also, if the position of the center of gravity is measured from the center of gravity sensor, it is determined that the center of gravity sensor has failed. can do.

The switching step corresponds to a step of switching the control model because flight performance and stability cannot be secured when it is used for control because the reliability of the measured or estimated center of gravity is low. Looking at the block diagram of Figure 4, the function of the switching step in the Fault Detector Block can be performed. The switching step is configured to switch the control from the first mode control step to the second mode control step when it is determined that the reliability of the center of gravity information is low in the reliability determination step of the center of gravity information.

S100: first mode control step
S200: second mode control step
S300: Step of calculating the angular acceleration error
S400: Control gain design step of additional augmentation algorithm
S500: limiting step to limit the amount of augmentation
S600: adding the calculated augmentation algorithm to the first mode control step

Claims (6)

A first mode control step of generating a predicted angular acceleration of the aircraft according to the pilot's control input as an output;
A second mode control step of generating a measurement angular acceleration measured according to the pilot's control input as an output;
An angular acceleration error calculation step of calculating an error between the predicted angular acceleration of the first mode control step and the measured angular acceleration of the second mode control step;
An additional control gain design step of applying an additional augmented gain determined according to the calculated angular acceleration error to the measured angular acceleration and applying the error between the angular accelerations;
Limiting the amount of augmentation of the measured angular acceleration applied to the additional augmentation gain to satisfy flight stability; And
And adding an error between angular accelerations in which the amount of enhancement is limited to the first mode control step,
The step of designing the additional control gain and limiting the amount of augmentation and summing the error between the angular acceleration where the amount of augmentation is limited to the first mode control step,
Control in order to secure the flightability of the aircraft to cope with the changing center of gravity during flight, characterized in that it is performed when it is determined that the error occurs in the position of the center of gravity outside the predetermined range in the angular acceleration error calculation step. Way. According to claim 1,
In the first mode control step, the angular acceleration estimated according to the current flight state by the on-board model (OBM) is used as a predicted angular acceleration output. Control method. According to claim 2,
The second mode control step is a control method for securing flightability of an aircraft that copes with a changing center of gravity during flight, characterized in that the angular acceleration measured from the angular acceleration sensor is generated as an output. According to claim 1,
The additional control gain design step determines the additional gain according to the magnitude of the error between the predicted angular acceleration and the measured angular acceleration. A control method for securing flightability of an aircraft coping with a changing center of gravity during flight . According to claim 4,
The additional control gain design step is to determine the augmented gain variably according to the magnitude of the error between the predicted angular acceleration and the measured angular acceleration for securing flightability of the aircraft coping with the changing center of gravity during flight. Control method. According to claim 1,
The step of limiting the amount of augmentation is selected from preset values so as to maintain flight performance that satisfies stability according to flight conditions and flight conditions, thereby ensuring flightability of the aircraft to cope with the changing center of gravity during flight. Control method for.

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