KR101440508B1 - Method for designing flight control law of rotor craft - Google Patents

Abstract

The purpose of the present invention is to provide a flight control law designing method of a rotor craft capable of securing the controllability in a hover area in which the visibility of the rotor craft is degraded. To achieve the purpose, the light control law designing method of the rotor craft according to the present invention comprises: a step (S100) of designating a response type for compensating to follow an orientation angle for reducing an operational stress of a pilot caused by a degraded visual field by using a control stick of the rotor craft; a step (S200) of setting a model inverse transformation control system so that a command model follows the response type designated; and a step (S300) of verifying a control gain regarding the model inverse transformation control system. The model inverse transformation control system comprises: a command model unit which copies and produce a desire response type; a feedback compensating unit which compensates a difference between the response type of a real rotor craft and the copied command model; an inverse motion model unit which copies motion characteristics inversely in the hover of the rotor craft; and a motion model unit which copies the motion characteristics inversely in the hover of the rotor craft. The feedback compensation for compensating a difference between a real response of the rotor craft and the command model is executed between the feedback compensating unit, the inverse motion model unit, and the motion model unit.

Description

METHOD FOR DESIGNING FLIGHT CONTROL LAW OF ROTOR CRAFT}

The present invention relates to a method for designing a flight control law of a rotorcraft, and more particularly, to a method for designing a flight control law of a rotorcraft for ensuring safety and flightability of a helicopter when flying in place in a flight environment where a watch is not secured.

Conventional mechanical helicopters have limited automation of flight control and are difficult to maintain and maintain due to complicated mechanical devices. Moreover, in a flight environment in which a watch is not secured, when flying in place, deterioration of flying quality may occur, thereby deteriorating the safety and mission capability of the helicopter. Advanced airlines apply digital fly-by-wire flight control systems by highly developed digital control technology to solve these problems and ensure helicopter stability and flying qualities. Doing. Helicopter electric flight control system technology has been commercialized by applying U.S. and French Eurocopters to military helicopters such as NH-90, UH-60M, and CH-47, and S-92F helicopter developed by U.S. Sikolski is a civilian helicopter. For the first time, we are conducting a certified flight test. In particular, the United States is pursuing a project to replace mechanical systems with FBW systems for the majority of military helicopters in operation.

On the other hand, the helicopter flight control law developed a basic model following control (MFC) method through the ADOCS (Advanced Digital Optical Control System) program of the United States National Aeronautics and Space Administration (NASA) in the 1980s, and the modified UH- It was applied to a 60A Black Hawk. The model tracking control method developed by ADCOS was applied to V-22 Osprey and RAH-66 Comanche, and further utilized in various helicopter FBW flight control systems such as AH-64 Apach, UH-60M, CH-47 Chinook and S-92. have.

An object of the present invention is to provide a method for designing a flight control law of a rotorcraft that can secure maneuverability in an in-flight flying zone in which the field of view of the rotorcraft is degraded by applying a model inversion control method.

In order to achieve the above object, a method for designing a flight control law of a rotorcraft according to the present invention is performed in a flying area in which the field of view of the rotorcraft is lowered, according to a reduced field of view by using a steering wheel of the rotorcraft. Step (S100) of designating a response form for compensating to follow a posture angle for reducing the pilot's control burden, and setting a model inverse transformation control system so that a command model can follow the designated response form (S200) And, comprising the step of verifying the control gain for the set model inverse transformation control system (S300), wherein the model inverse transformation control system, a command model unit for simulating and generating the desired response form, and the actual rotorcraft Response form, a feedback compensation unit for correcting an error between the simulated command models, an inverse motion model unit for inversely simulating motion characteristics when in flight of the rotorcraft, and motion characteristics when in flight of the rotorcraft It includes a motion model unit for simulating, feedback compensation for correcting an error between the actual response of the rotorcraft and the command model between the feedback compensation unit, the inverse motion model unit, and the motion model unit It is characterized by.

In addition, the method for designing a flight control law of a rotorcraft according to the present invention is characterized in that the feedback compensation delays a signal for the command model.

In addition, in the method of designing a flight control law of a rotorcraft according to the present invention, in step S100, the response form is a posture command response form and a posture angle maintenance response form on a vertical axis and a horizontal axis, and a change rate command response on a direction axis It is characterized by designing a form and direction angle holding response form and a change rate command response form and an altitude holding response form on the elevation axis.

In addition, the method for designing a flight control law of a rotorcraft according to the present invention, in the response form, applies a PH (Position Hold) and a Translational Rate Command (TRC) to improve the maneuverability in a flight environment in which a watch is deteriorated. It is characterized by.

According to the present invention, by applying the Model Inversion Control (Model Inversion Control) method, there is an effect that can secure the maneuverability in the in-flight flight zone where the field of view of the rotorcraft is deteriorated.

1 is a floor chart showing the flow of a method for designing a flight control law of a rotorcraft according to the present invention.
3 is a block diagram showing the concept of model inverse transformation control of a method for designing a flight control law of a rotorcraft according to the present invention.

The present invention can be applied to various transformations and may have various embodiments, and specific embodiments will be illustrated in the drawings and described in detail in the detailed description. However, this is not intended to limit the present invention to specific embodiments, and should be understood to include all conversions, equivalents, and substitutes included in the spirit and scope of the present invention. In the description of the present invention, when it is determined that detailed descriptions of related known technologies may obscure the subject matter of the present invention, detailed descriptions thereof will be omitted.

Terms such as first and second may be used to describe various components, but the components should not be limited by the terms. The terms are used only for the purpose of distinguishing one component from other components.

The terminology used herein is only used to describe a specific embodiment, and is not intended to limit the present invention. Singular expressions include plural expressions unless the context clearly indicates otherwise. In this specification, terms such as “include” or “have” are intended to indicate that a feature, number, step, operation, component, part, or combination thereof described in the specification exists, and that one or more other features are present. It should be understood that the existence or addition possibilities of fields or numbers, steps, operations, components, parts or combinations thereof are not excluded in advance.

In addition, terms such as “…unit”, “…module” described in the specification mean a unit that processes at least one function or operation, and may be implemented by hardware or software, or a combination of hardware and software.

In addition, in the description of the present invention, when it is determined that detailed descriptions of related known technologies may unnecessarily obscure the subject matter of the present invention, detailed descriptions thereof will be omitted.

Hereinafter, embodiments of the present invention will be described in detail with reference to related drawings.

1 is a floor chart showing a flow of a method for designing a flight control law of a rotorcraft according to the present invention.

Referring to FIG. 1, a method for designing a flight control law for a rotorcraft according to the present invention is a method for designing a flight control law for a rotorcraft, in a hovering flight region in which the field of view of the rotorcraft is deteriorated, the steering wheel of the rotorcraft is controlled. A step (S100) of designating a response form for compensating for following a posture angle for reducing a pilot's steering burden according to the deteriorated field of view, and a model inverse transformation control system for the command model to follow the designated response form. A step (S200) of setting and a step of verifying a control gain for the set model inverse transform control system (S300), wherein the model inverse transform control system, a command model unit for simulating and generating a desired response form, and a real rotorcraft The feedback compensation unit that corrects the error between the response type of the aircraft and the simulated command model, the inverse motion model unit that inversely simulates the motion characteristics of the rotorcraft in flight, and the motion characteristics of the rotorcraft in flight It includes a simulated motion model unit, it characterized in that the feedback compensation for correcting the error between the actual response of the rotorcraft and the command model between the feedback compensation unit, the inverse motion model unit, and the motion model unit.

First, the response type of the rotorcraft will be described.

The response type of most helicopters adopts the rate response type that follows the angular velocity or vertical velocity. In such a control system, in order to follow the desired attitude angle of the pilot, it is necessary to continuously compensate using the steering wheel. Moreover, when flying in place near the ground in a degraded visual environment where the watch is not secured, the pilot's maneuvering burden is increased, leading to a decrease in maneuverability, which can lead to dangerous situations. To solve this problem, for example, the S-33E-PRF requires a response form according to UCE (Usable Cue Environment).

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Next, with reference to Figure 3 will be described in more detail with respect to the flight control law design method of a rotorcraft according to the present invention.

3 is a configuration diagram showing the concept of model inverse transformation control of a method for designing a flight control law of a rotorcraft according to the present invention.

Referring to FIG. 3, the model inverse transformation control system compensates feedback to correct an error between a command model unit 100 that simulates and generates a desired response form, and a response form of an actual rotorcraft and a simulated command model. Includes an inverse 200, an inverse motion model (300) for simulating the motion characteristics when flying in situ of a rotorcraft aircraft, and an exercise model (400) for simulating motion characteristics when flying in situ of a rotorcraft, feedback compensation unit Wow, a feedback compensation is performed between the inverse motion model and the motion model to correct errors between the actual response of the rotorcraft and the command model.

This model inverse transformation control system aims to enable a helicopter to follow a command model whose response characteristics are lowered. As shown in FIG. 3, the model inverse transformation control system includes a lowered command model (M(s)) for generating a desired command, and feedback compensation for suppressing disturbances and stabilizing the air vehicle. , E(s)), an inverse plant model (P-1(s)) and a dynamics model (P(s)). The control system of this structure allows the helicopter response to follow the command model without error if the inverse motion model can be accurately represented. However, it is difficult to perfectly cancel low frequency characteristics such as high frequency dynamics associated with helicopter rotors and actuators, and effects of dynamic trim, weight, and center of gravity.

Accordingly, feedback compensation is provided to correct an error between the helicopter response and the command model caused by the above-described effect. In addition, the command model signal is delayed in the feedback loop so that model following is acceptable in the high frequency band.

First, the instruction model unit 100 will be described.

Hereinafter, the variables used in the described formulas will be referred to Table 1.

The command model is an algorithm that generates the desired dynamic response desired by the user, and is constructed by low-ordering as in Equation (1). The command model applied to the vertical axis/horizontal axis is composed of the secondary model, and the direction axis and the elevation axis are composed of the primary model. In addition, the design coefficient of the command model is selected as shown in Table 2 so as to satisfy the bandwidth requirements defined in the ADS-33E-PRF.

... 식 (1)

... Expression (1)

Next, the feedback compensation unit 200 will be described.

The model inverse transform control system can follow the command model without error if the inverse motion model can be accurately represented. However, it is difficult to completely cancel the high and low frequency characteristics of the helicopter model. Therefore, it is necessary to design a feedback compensation unit 200 to correct an error between an aircraft response and a command. The feedback compensation unit 200 applies a Proportional-Integral-Derivatives (PID) controller as shown in Equation (2).

... 식 (2)

... Expression (2)

Next, the inverse motion model unit 300 will be described.

Model inverse transform control is a system that theoretically does not require control gain scheduling if accurate aerodynamic data is provided. The inverse motion model used in the present invention is configured as follows by converting a nonlinear system into a linear system.

The relationship between the angular velocity and the Euler change rate for the body frame is shown in Eq. (3).

... 식 (3)

... Expression (3)

Summarizing Eq. (3) gives Eq. (4) as follows:

... 식 (4)

... Expression (4)

Using equation (4) and a helicopter dynamics model, the inverse motion model for the control input of a helicopter with a constant rotor speed can be obtained as shown in equation (5).

... 식 (5)

... Expression (5)

Next, the evaluation of the control law will be described.

In a complex control system, designing the control gain through the trial and error of the designer increases the development cost by increasing the development period. Therefore, in the present invention, control designer's Unified Interface (CONDUIT), which is a control law development software, optimizes and verifies the control gain. The specifications provided by this tool can be divided into fixed-wing specifications based on MIL-F-8785C, rotating-wheel specifications based on ADS-33E-PRF, and general specifications commonly applied to controller design. Each standard includes Hard(H) defining stability criteria, Soft(S) defining handling quality criteria, and Objective(O) constraints defining performance and actuator usage criteria ( constraint). In the CONDUIT optimization process, control gain optimization that satisfies the specifications set in phases (Phase 1 to 3) is performed by constraints set by the designer.

Applicable standards

In the present invention, the design coefficient provided by CONDUIT is applied to optimize the design coefficient. Table 3 shows the specifications applied to control gain optimization and analysis. The stability margin and eigenvalue standard related to the stability of the helicopter were set as hard constraints, and specifications such as damping ratio, bandwidth, and robustness against disturbance related to flight performance were soft. Set it as a constraint. In addition, the specification for the crossover frequency and the use of the actuator is the objective constraint, the specification for agility and cross coupling of each axis is set as the check constraint to check whether the specification is satisfied.

Optimization results

Table 3 shows the results of CONDUIT optimization of design factors. As a result of initially setting the frequency of the horizontal axis command model to 2.0 and performing optimization, it is impossible to satisfy the roll quickness specification only by optimizing the control gain. As a result, optimization is performed by increasing the frequency of the roll axis command model to 3.0, so that the roll speed specification can be satisfied. In addition, when optimization is performed without considering robustness against disturbance, specifications such as stability margin, attenuation, and agility can be satisfied, but control gains vulnerable to disturbance are generated. Therefore, when designing a control gain using CONDUIT, it is possible to optimize the design coefficient by introducing a specification for helicopter response due to disturbance from the viewpoint of stability. In addition, the advantage of applying the inverse transformation of the entire linearization matrix by applying the model inverse transformation technique can remove most of the linkage effects without designing a separate controller. In the case of the model tracking control law, the error compensator is located at the rear of the controller and is compensated for the result of the inverse transform model. Therefore, it is necessary to design a decoupler algorithm to remove the linkage effect between each axis.

Advanced aviation companies are applying electric flight control systems by highly developed digital control technology to ensure the stability of unstable helicopters and ensure optimal flightability. In the present invention, a model inversion control method is applied to design and verify the control law in the in-situ flight region. The optimization of the design variables is based on the commercial software CONDUIT, designing a control gain that can satisfy the specifications presented in the ADS-33E-PRF. And, as a result of evaluating the defined standard using CONDUIT, it can be confirmed that the handling quality rating (HQR) Level 1 specified in the ADS-33E-PRF is satisfied in the in-flight flight area.

In the above, the embodiment of the present invention has been described as an example, but various changes are possible at the level of those skilled in the art. Therefore, it is obvious that the present invention should not be interpreted as being limited to the above-described examples, but should be interpreted by the claims set forth below.

100: command model unit
200: feedback compensation unit
300: inverse motion model
400: exercise model unit

Claims (4)

As a method of designing a flight control law for a rotorcraft,
In a flight area where the field of view of the rotorcraft is deteriorated, a response form for compensating to follow a posture angle for reducing a pilot's maneuvering load according to the deteriorated field of view is determined by using the steering wheel of the rotorcraft. Step (S100), and
Setting a model inverse transformation control system so that a command model can follow the designated response form (S200);
Comprising the step of verifying the control gain for the set model inverse transformation control system (S300),
The model inverse transformation control system,
A command model unit that simulates and generates the desired response form;
A feedback compensation unit for correcting an error between the response form of the rotorcraft and the simulated command model;
An inverse motion model unit that inversely simulates motion characteristics of the rotorcraft in flight,
It includes an exercise model unit that simulates the motion characteristics when flying in place of the rotorcraft,
Flight control of the rotorcraft, characterized in that feedback compensation for correcting errors between the actual response of the rotorcraft and the command model is made between the feedback compensation unit, the inverse movement model unit, and the movement model unit. How to design the law. According to claim 1,
The feedback compensation is a method for designing a flight control law of a rotorcraft, characterized in that it delays the signal for the command model. According to claim 1,
In the step (S100),
The response form is
On the vertical axis and horizontal axis, the posture command response form and posture angle maintenance response form,
On the direction axis, change rate command response type and direction angle maintenance response type,
A method of designing a flight control law of a rotorcraft, characterized by designing a change rate command response type and an altitude maintenance response type on the high axis. The method of claim 3,
In the above response form,
A flight control law design method for a rotary wing aircraft, characterized by applying a PH (Position Hold) and a TRC (Translational Rate Command) to improve maneuverability in a deteriorated flight environment.

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