KR101803450B1 - Flight control system of helicopter using sas actuator - Google Patents

Flight control system of helicopter using sas actuator Download PDF

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Publication number
KR101803450B1
KR101803450B1 KR1020150178885A KR20150178885A KR101803450B1 KR 101803450 B1 KR101803450 B1 KR 101803450B1 KR 1020150178885 A KR1020150178885 A KR 1020150178885A KR 20150178885 A KR20150178885 A KR 20150178885A KR 101803450 B1 KR101803450 B1 KR 101803450B1
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South Korea
Prior art keywords
displacement
actuator
main
flight control
stability enhancing
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KR1020150178885A
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Korean (ko)
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KR20170071033A (en
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김응태
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한국항공우주연구원
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    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C27/00—Rotorcraft; Rotors peculiar thereto
    • B64C27/54—Mechanisms for controlling blade adjustment or movement relative to rotor head, e.g. lag-lead movement
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C27/00—Rotorcraft; Rotors peculiar thereto
    • B64C27/54—Mechanisms for controlling blade adjustment or movement relative to rotor head, e.g. lag-lead movement
    • B64C27/56—Mechanisms for controlling blade adjustment or movement relative to rotor head, e.g. lag-lead movement characterised by the control initiating means, e.g. manually actuated
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C27/00—Rotorcraft; Rotors peculiar thereto
    • B64C27/54—Mechanisms for controlling blade adjustment or movement relative to rotor head, e.g. lag-lead movement
    • B64C27/58—Transmitting means, e.g. interrelated with initiating means or means acting on blades
    • B64C2700/6284—

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Control Of Position, Course, Altitude, Or Attitude Of Moving Bodies (AREA)
  • Toys (AREA)

Abstract

The present invention relates to a helicopter comprising a spindle motor for adjusting a pitch angle of a rotor blade of a helicopter and a mechanical linkage configured to move a spindle synchronous valve to generate a displacement of a spindle motor, A stability enhancing actuator for increasing or decreasing the movement of the valve in the main synchronous mode; and a flight control computer connected to the main pilot control and applying a control signal to the stability enhancing actuator, wherein the flight control computer comprises: And a second mode in which displacement of the main drive synchronism occurs only by displacement of the mechanical linkage due to the displacement of the pilot manipulation steering wheel, is switched between the first mode in which the displacement of the main drive synchronism occurs only by the displacement of the stability- Wherein the helicopter It discloses a flight control system.

Description

TECHNICAL FIELD [0001] The present invention relates to a flight control system for a helicopter,

The present invention relates to a flight control system for a helicopter for attitude control of a helicopter.

The attitude control of the helicopter is accomplished by adjusting the pitch angle of the rotor blades, and FIG. 1 illustrates a typical helicopter flight control system.

1 is a mechanical control system in which a valve of a master cylinder 30 for generating a driving force for adjusting a pitch angle of a rotor blade is connected to a pilot control rod 10 by a mechanical linkage 20. When the pilot manipulates the control section 10, the mechanical linkage 20 operates to cause the valve of the spindle motor 30 to move, thereby causing displacement of the spindle motor 30, thereby changing the pitch angle of the rotor blade.

A stability augmentation system (SAS) is applied to improve the maneuverability and stability of the helicopter. The stability augmentation system (SAS) is applied to the flight control computer 40 and the stability enhancing actuator 50 (SAS actuator). The flight control computer 40 is connected to various sensors 70 for sensing various states of the helicopter and the control station 10 and the stability boosting actuator 50 is connected to the mechanical linkage 20 to control the valve To or from the motion. The flight control computer 40 controls the operation of the stability enhancing actuator 50 based on command signals from the control station 10 and information sensed by various sensors 70. [ A trim actuator 60 connected to the flight control computer 40 is connected to the control station 10 so that the operation of the mechanical linkage 20 is performed by the operation of the trim actuator 60. The trim actuator 60 is commonly used in an automatic flight mode that maintains speed, altitude, and the like.

FIG. 2 shows a helicopter flight control system using a flywheel (FBW) system developed to replace a conventional mechanical flight control system.

According to this, the pilot control unit 10 is connected to the flight control computer 40 directly connected to the main driving motor 30 without the mechanical linkage. The flight control computer 40 receives the displacement of the control station 10 as an electrical signal, performs an operation on the control signal, and then controls the operation of the FBW-dedicated spindle motor 30. Such an FBW type flight control system has an advantage of excellent maneuverability and stability because it can implement an attitude command (AC: Attitude Command) or a TRC (Translational Rate Command) response type control.

However, in spite of these advantages, the FBW type flight control system requires not only the expensive FBW dedicated spindle motor 30 but also a failure in the flight control computer 40 or the sensor 70 due to lack of mechanical linkage This is a disadvantage. In addition, there is a disadvantage that it is difficult to acquire the certification and the development cost is high because the reliability must be very high.

In this regard, instead of using the expensive FBW exclusive spur gear 30, a stabilization actuator mounted on a conventional mechanical flight control system may be used to control the attitude command or the limited speed command response type (Limited Authority) control system has been proposed. According to this, since the mechanical linkage mounted on the conventional mechanical system is used as it is, in order to move the spindle synchronous valve according to the calculation result in the flight control computer, the stability enhancing actuator must cancel the displacement of the mechanical linkage moved by the control rod. In other words, the stability enhancing actuator must move the displacement of the flight control computer combined with the displacement by the steering. Since the maximum operating range of the stability enhancing field synchronizer is only 10% to 20% of the main actuator valve operating range, There is a problem that the stability enhancing actuator is saturated.

Japanese Patent Application Laid-Open No. 10-2015-0130822 (2015.11.24)

SUMMARY OF THE INVENTION It is an object of the present invention to provide a flight control system capable of realizing an FBW helicopter control law using a stability enhancing actuator used in a conventional mechanical flight control system.

It is to be understood that both the foregoing general description and the following detailed description are exemplary and explanatory and are not intended to limit the invention to the precise forms disclosed. Other objects, which will be apparent to those skilled in the art, It will be possible.

According to an aspect of the present invention, there is provided a helicopter comprising a main rotor for adjusting a pitch angle of a rotor blade of a helicopter, a mechanical linkage configured to generate a displacement of a main rotor by moving the main rotor, A stability enhancing actuator connected to the mechanical linkage and for increasing or decreasing the movement of the valve of the main synchronous motor; and a flight control computer connected to the main pilot control and applying a control signal to the stability enhancing actuator, A first mode in which displacement of the main drive synchronism occurs only by the displacement of the stability enhancing actuator according to the displacement of the main pilot control and a second mode in which displacement of the main drive synchronizer is caused only by displacement of the mechanical linkage according to displacement of the pilot manipulator A configuration in which the second mode that occurs can be switched Proposes a flight control system of the helicopter, characterized in that the.

According to the flight control system of the helicopter of the present invention, the flight control computer operates in the first mode, and when saturation occurs in the stability enhancement actuator, the pilot mode is switched to the second mode so that the pilot is transferred from the main pilot to the pilot.

According to another aspect of the present invention, there is provided a helicopter comprising: a main linkage mechanism for controlling a pitch angle of a rotor blade of a helicopter; a mechanical linkage configured to move a main synchronous valve to generate a displacement of a main synchronous path, And a trim actuator connected to the pilot pilot and operative to generate a displacement of the mechanical linkage; a stabilization actuator connected to the pilot pilot and connected to the stability enhancer and the trim, And a flight control computer for applying a control signal to the actuator, wherein the flight control computer has a first mode in which displacement of the main motive occurs only by displacement of the stability enhancing actuator according to displacement of the main pilot, The reinforcement and trim actuators work together And a second mode in which the displacement of the spark plug is generated by the sum of the displacements of the mechanical linkage and the stability enhancing actuator.

According to the flight control system of the helicopter of the present invention, when the flight control computer operates in the first mode and the displacement command value for the output of the stability enhancing actuator or the spark plug is larger than a preset value, The trim actuator can be operated to switch to the second mode.

According to the flight control system of the helicopter of the present invention, the flight control computer calculates a value obtained by subtracting the displacement value of the mechanical linkage from the displacement command value for the spark plug in the second mode, .

According to the present invention, it is possible to provide a flight control system capable of implementing the FBW helicopter control law using the stability enhancing actuator used in the conventional mechanical flight control system, and preventing saturation of the stability enhancing actuator. .

Further, according to the present invention, it is possible to implement a flight control system even in a low-cost single system instead of an expensive multiplexing system of the FBW control system by providing a backup system using a pilot pilot control system, .

1 is a block diagram showing a flight control system of a general helicopter.
Figure 2 is a block diagram of a FBW helicopter flight control system.
3 is a block diagram illustrating a flight control system for a helicopter according to an embodiment of the present invention.
4 is a block diagram illustrating a flight control system of a helicopter according to another embodiment of the present invention.
5 and 6 are diagrams showing signal processing processes in the first mode and the second mode in the flight control system of the helicopter shown in FIG.

Hereinafter, a flight control system for a helicopter according to the present invention will be described in detail with reference to the drawings.

3 is a block diagram illustrating a flight control system for a helicopter according to an embodiment of the present invention.

The flight control system of the helicopter according to the present embodiment includes a main motive 110, a mechanical linkage 120, a stability enhancing actuator 130 or an SAS actuator, and a flight control computer 150.

The main driving motor 110 generates driving force for adjusting the pitch angle of the rotor blades of the helicopter.

The mechanical linkage 120 is configured to mechanically drive the valve of the main motive 110 to generate displacement of the main motive 110. Unlike the general flight control system shown in FIG. 1, the mechanical linkage 120 of the present embodiment is connected to the pilot pilot control station 102 separately provided from the pilot pilot pilot 101. That is, the valve of the main motive 110 is connected to the copilot pilot 102 by the mechanical rackie 120 and the mechanical linkage 120 is operated by the displacement of the pilot pilot control 102. [

The stability enhancing actuator 130 is connected to the mechanical linkage 120 and functions to increase or decrease the valve movement of the main motive 110.

The flight control computer 150 is directly connected to the main pilot control 101 and is configured to apply a control signal to the stability enhancing actuator 130. [ The flight control computer 150 is connected to various sensors 160 mounted on the helicopter and controls the stability enhancing actuator 130 based on command signals from the main pilot control 101 and information sensed by various sensors 70. [ .

According to the present embodiment, the flight control computer 150 controls the first mode in which the displacement of the main motive 110 is caused only by the displacement of the stability enhancing actuator 130 according to the displacement of the main pilot control 101, And the second mode in which the displacement of the main spindle 110 is generated only by the displacement of the mechanical linkage 120 according to the displacement of the main spindle 10.

As an example, the flight control computer 150 may operate in a first mode, and may switch to a second mode when saturation occurs in the stability enhancing actuator to operate so as to transfer the steering right from the main pilot to the pilot.

According to this, in the first mode, the electric signal of the displacement measuring device is transmitted to the flight control computer 150 according to the displacement of the main pilot controllable 101, and the flight control computer 150 performs calculation thereon, 130). The operation of the stability enhancing actuator 130 causes the valve of the spindle motor 120 to move so that the displacement of the spindle motor 120 occurs and the pitch angle of the rotor blade changes. In this case, unlike the limited rights control system, there is no need to offset the stability enhancing actuator by displacement of the mechanical linkage moved by the steering wheel, so that the stabilization enhancing actuator 130 has a relatively small saturation.

As soon as saturation of the stability enhancing actuator 130 occurs, the pilot mode of the pilot is automatically transferred from the main pilot to the copilot, so that the operation of the main motive 110 is controlled by the pilot's manipulation. That is, the mechanical linkage 120 is operated according to the displacement of the copilot controllable block 102, so that the valve of the main block 120 is moved. As a result, the displacement of the main block 120 is generated, do.

In this way, the control method of the attitude command (AC) or the horizontal speed command (TRC) response type which can be implemented by the FBW control law can greatly improve the maneuverability and stability in the low speed flight which is difficult to steer. In low-speed flight, it is expected that the stabilization actuator 130 is not likely to saturate because the pilot moves finely. Therefore, it is possible to fly effectively by controlling the FBW system at low speed and switching from high speed to mechanical linkage system will be.

4 is a block diagram illustrating a flight control system of a helicopter according to another embodiment of the present invention.

The helicopter flight control system according to the present embodiment includes a main motive 110, a mechanical linkage 120, a stability enhancing actuator 130, a trim actuator 140, and a flight control computer 150.

The main assembly motors 110, the mechanical linkages 120, and the stability enhancing actuators 130 have the same configuration as that of the previous embodiment, and therefore, the description thereof will be omitted.

The trim actuator 140 is connected to the copilot controllable 102 and is operated to generate displacement of the mechanical linkage 120.

The flight control computer 150 is directly connected to the main pilot control 101 and is configured to apply control signals to the stability enhancing actuator 130 and the trim actuator 140. [ The flight control computer 150 is connected to various sensors 160 mounted on the helicopter and controls the stability enhancing actuator 130 based on command signals from the main pilot control 101 and information sensed by various sensors 70. [ And the trim actuator (140).

The flight control computer 150 is operated in the first mode in which the displacement of the main motive 110 is generated only by the displacement of the stability enhancing actuator 130 according to the displacement of the main pilot control 101 and the second mode in which the stability enhancing actuator 130 and the trim The actuators 140 are operated together to switch between the second mode in which the displacement of the master motive 110 occurs in the sum of the displacements of the mechanical linkage 120 and the stability enhancing actuator 130. [

The flight control computer 150 operates in the first mode and activates the trim actuator 140 when the output of the stability enhancing actuator 130 or the displacement command value for the main motive 110 is greater than a predetermined value, .

The main displacement motive 110 valve displacement in the second mode is dependent on the sum of the displacement of the stability enhancing actuator 130 and the displacement of the trim actuator 140 so that the displacement of the stability enhancing actuator 130 can be reduced. It is possible to prevent the saturation phenomenon of the stability enhancing actuator 130 because the trim actuator 140 can move the valve of the main synchronizer 110 to the maximum displacement even if the main pilot control rod 101 is moved to the maximum. However, since the operation speed of the trim actuator 140 is slower than the operation speed of the stability enhancer 130, the stability enhancer 130 may be saturated when the main pilot controllable 101 is moved quickly, The stability enhancing actuator 130 can be released from the saturation state after a long time depending on the operation of the actuator 140. [

An example of the control logic of the flight control computer 150 will be described below with reference to FIGS. 5 and 6. FIG. Fig. 5 schematically shows the signal processing in the first mode, and Fig. 6 schematically shows the signal processing in the second mode.

First, in the flight control computer 150, a master motion (130) displacement command (y_act_comm) for the desired aircraft control is calculated.

If the magnitude of the displacement command value y_act_comm for the main spindle motor 130 is equal to or smaller than the preset value y_trim_min, the stability increasing actuator 130 is operated with the value y_act_comm as shown in FIG. (Main actuator command = y_sas_comm = y_act_comm)

However, if the magnitude of the displacement command value y_act_comm with respect to the main action synchronizer 130 becomes larger than the preset value y_trim_min, the trim actuator 140 starts to operate with the value y_act_comm as shown in FIG.

The displacement of the trim actuator 140 becomes the y_link_out of the mechanical linkage 120 at the "A" position of FIG. 4 via the mechanical linkage 120. (Generally y_link_out = gain * y_trim_out)

The flight control computer 150 applies the value obtained by subtracting the displacement value y_link_out of the mechanical linkage 120 from the displacement command value y_act_comm for the main motive 130 to the stability increasing actuator 130 as an operation command. (y_sas_comm = y_act_comm - y_link_out)

Since the final displacement that actuates the main driving motive 110 is a sum of the displacement y_sas_out of the stability enhancing actuator 130 and the displacement y_link_out of the mechanical linkage 120 and the y_sas_out value is equal to y_sas_comm, It can be confirmed that the displacement command value (y_act_comm) is transferred as it is to the main motive 110 in the main motive 110 calculated in the flight control computer 150. [ (Main actuator command = y_link_out + y_sas_comm = y_act_comm)

When the magnitude of the displacement command value y_act_comm becomes equal to or smaller than the preset value y_trim_min in the second mode, the mode is switched again to the first mode.

In the case of the flight control system of the present embodiment, the steering mechanism using the pilot pilot control station 102 has a function as a backup system. That is, when an abnormality occurs in the FBW control system, the co-pilot can transfer the control right and continue steering. In this case, steering using the mechanical linkage 120 is performed.

The above-described helicopter flight control system is not limited to the configuration and the method of the embodiment described above, but various modifications can be made by those skilled in the art within the scope of the technical idea of the present invention.

101: Main Pilot Pilot 102: Pilot Pilot
110: Master Motive 120: Mechanical Linkage
130: Stability enhancing actuator 140: Trim actuator
150: flight control computer 160: sensor

Claims (5)

delete delete A master motive for controlling the pitch angle of the rotor blade of a helicopter;
A mechanical linkage configured to generate a displacement of a spindle motor to move the valve of the spindle motor, the mechanical linkage being operated by a pilot pilot control rod;
A stability enhancing actuator configured to move the main motive valve to generate a displacement of the main motive, the stability enhancing actuator being operatively connected to the main actuator valve by a mechanical linkage;
A trim actuator coupled to the copilot control and operative to generate a displacement of the mechanical linkage; And
And a flight control computer connected to the main pilot control and applying a control signal to the stability boost actuator and the trim actuator,
The flight control computer,
A first mode in which displacement of the main drive synchronism occurs only by displacement of the stability enhancing actuator according to displacement of the main pilot control,
Wherein the stability enhancing actuator and the trim actuator are operated together to switch between the second mode in which displacement of the main drive synchronism is generated by the sum of the displacements of the mechanical linkage and the stability enhancing actuator,
Wherein the flight control computer operates in a first mode and operates the trim actuator when the output of the stability enhancing actuator or the displacement command value for the spark plug is greater than a preset value and saturation of the stability enhancing actuator is expected, Mode of the helicopter is changed to reduce the displacement of the active stability enhancing actuator. delete The method of claim 3,
The flight control computer,
Wherein a value obtained by subtracting a displacement value of the mechanical linkage from a displacement command value for the spark plug in the second mode is applied to the stability enhancing actuator as an operation command.

KR1020150178885A 2015-12-15 2015-12-15 Flight control system of helicopter using sas actuator KR101803450B1 (en)

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Application Number Priority Date Filing Date Title
KR1020150178885A KR101803450B1 (en) 2015-12-15 2015-12-15 Flight control system of helicopter using sas actuator

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KR101803450B1 true KR101803450B1 (en) 2017-11-30

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