JPH01169601A - Multiplex channel control device for aircraft - Google Patents

Abstract

PURPOSE:To exchange only a computer to generate a defect by equipping with a majority processing means at respective interface devices provided on respective control channels. CONSTITUTION:A computer 45 for monitoring constituted in the same way as actuator control computers 15, 25 and 35 is used in order to execute the judgement to apply for majority principles. Consequently, rudder wing driving members 18, 28 and 38 of respective control channels C1-C3 are driven by rudder wing driving signals S1'-S3' outputted from actuator control computers 15, 25 and 35 provided on respective control channels C1-C3 and not driven by the output signal of a computer 45 for monitoring. The majority processing means in respective control channels C1-C3 judges whether or not respective control channels C1-C3 are normal. Thus, when the defect occurs at any of respective computers, only the computer to generate the defect may be exchanged.

Description

DETAILED DESCRIPTION OF THE INVENTION A1 Object of the Invention (1) Industrial Application Field The present invention relates to a multi-channel control device for an aircraft, and particularly to a computer that outputs a rudder blade drive signal and a rudder operated by the rudder blade drive signal. The present invention relates to a multichannel control device for an aircraft having three or more control channels each having a wing drive member.

(2) Prior Art Conventionally, various multichannel control devices for aircraft have been proposed, and for example, a multichannel control device for aircraft as shown in FIG. 2 is known.

This aircraft multi-channel control device has three channels, and the amount of displacement of the control rod A operated by the pilot is detected by three sensors 010, 020.030 using steering command signals α1. α2. Detected as α3. The steering command signal α1. detected by each of the sensors 010, 020°030. α2 and α3 are flight control computers 011 and 021, respectively. and 031.

Said flight control computer 1011゜02.
1,031 includes various state detection signals β1. β2. β3 is input.

Said Fly I control computer 011°02
1.031 are the steering command signals α1, α2 .
α3 and state detection signal βhβ2. rudder blade drive signal γ1°γ2 corresponding to the optimum rudder blade drive speed from β3. Calculate T3. These three calculated rudder blade drive signals γ1, γ
2. The values of γ3 are outputted to the other fly 1 to control computers 011, 021, 031 and compared there. And each rudder blade drive signal? '+, rz, T
If the differences between the respective values of the three flight control computers 011, 021, 031 are all within the predetermined range, the rudder blade drive signals of the three flight control computers 011, 021, 031 are set to the rudder blade drive members 013, 023. ．．
033. However, the three rudder blade drive signals T+
+Tz+r3, one rudder blade drive signal (for example, r+
) has a difference of more than a predetermined value compared to the other two rudder blade drive signals (for example, γ2゜γ3), the other two rudder blade drive signals rz+γ3 are determined to be normal signals based on the principle of majority voting. , a small number of rudder blade drive signals γ1 are determined to be abnormal signals. The rudder blade drive signal γ1 determined to be abnormal is not input to the rudder blade drive member 013, and the rudder blade drive signal γ2 determined to be normal. Only γ3 is the rudder blade drive member 023,0
33.

The rudder blade drive member 013, 023, or 033, etc. to which the rudder blade drive signal determined to be normal is input is the secondary actuator P. Displace. This secondary actuator = 5-tuator P. With the displacement of the power actuator P
is activated. The position of this power actuator P is detected by position sensors Oll, 024 and 034. This detected position signal θ,.

θ2. θ3 is a signal corresponding to the actual steering angle, that is, an actual steering angle signal. This actual steering angle signal θ1. θ2, θ3
is the flight control computer 011,0
21. and is fed back to 031.

As the power actuator P operates, rudder blades S such as elevators and rudders are driven, and the motion of the aircraft changes according to the displacement of the driven rudder blades S.

(3) Problems to be Solved by the Invention By the way, in the above-mentioned three-channel aircraft multi-channel control device, if a defect occurs in one channel, the correct operating channel is determined based on the principle of majority voting. However, if two channels are defective, the principle of majority voting cannot be applied, and it becomes impossible to determine which channel is operating correctly.

Therefore, we prepare a spare channel equipped with a spare fly controller, roll computer, rudder blade drive member, etc., and if one channel becomes defective, it can be used in place of the defective channel. There is a method of using the above-mentioned spare channel, or by preparing only a spare Fly 1 to control computer, and when a defect occurs in the flight control computer of one channel, the flight control computer of that channel can be replaced. Instead, methods of using the spare flight control computer are being considered.

However, in both the methods using the spare channel and the method using the spare computer, an expensive flight control computer is not required. If a defect occurs in the flight control computer and it is difficult to repair the defect, the entire expensive flight control computer must be replaced, resulting in high costs.

On the other hand, conventionally, a method is known in which an actuator control computer is disposed near a rudder blade drive member and the actuator control computer is made to share a part of the functions of the flight control computer. In this method of using the actuator control computer, the functions assigned to the actuator control computer can be removed from the flight control computer, so that the configuration of the flight control computer can be made more compact. Also, the method of using the actuator control computer is
Since the actuator control computer and the actuator control computer are configured separately, it is advantageous that when a defect occurs in these computers and it becomes necessary to replace them, they can be replaced independently.

Further, the actuator control computer disposed near the rudder blade drive member receives a rudder angle command signal from the flight control computer, and also feeds back a rudder angle signal corresponding to the actual rudder angle. It has become. According to the method using this actuator control computer, the actuator control computer controls the rudder blade drive member disposed nearby at an optimum speed based on the rudder angle command signal and the actual rudder angle signal. Since the rudder only operates to drive the rudder and does not perform any other operation, it has the characteristic of improving the response speed of the rudder blade.

The present invention has been made in view of the various circumstances described above, and includes three or more control channels each having a computer and a rudder blade drive member, and a spare computer having the same input/output characteristics as the computer in the control channel. An object of the present invention is to provide a new method of using the spare computer for each control channel in an aircraft multi-channel control device having a multi-channel controller, and to reduce maintenance costs.

B. Arrangement of the Invention (1) Means for Solving the Problems In order to achieve the above object, the multi-channel control device for an aircraft of the present invention provides a multi-channel control device for an aircraft in which the rudder angle command signal and the actual rudder angle signal from the flight control computer are controlled manually. It is equipped with three or more control channels each including an actuator computer 1 to a roll computer which output a rudder blade drive signal calculated from each of the input signals, and a rudder blade drive member operated by the rudder blade drive signal, and each control The rudder blade drive signals of the channels are compared, and if the rudder blade drive signals do not match, the majority rule is applied to determine that a large number of matching rudder blade drive signals are normal and a small number of rudder blade drive signals are abnormal. In an aircraft multi-channel control device, the aircraft multi-channel control device is equipped with a majority decision processing means for outputting a drive output control signal that prevents output of the rudder blade drive signal determined to be abnormal from being output to the rudder blade drive member. The actuator computer 1 is configured in the same manner as the roll computer, receives a rudder angle command signal and an actual rudder angle signal from the flight control computer, and outputs a rudder blade drive signal calculated from the individual saw signals. A monitoring channel equipped with a monitoring computer is provided, and each of the control channels is provided with an interface device through which rudder blade drive signals output from the roll computer and the monitoring computer are input to each of the actuator controllers 1, and each of the control channels Each interface device provided in the channel is characterized in that each interface device is provided with the majority voting processing means.

(2) Effect The aircraft multi-channel control device of the present invention having the above-mentioned configuration has a computer divided into a flight control computer and an actuator control computer, so the computer for the monitoring channel is also divided. Can be configured. Then, when a defect occurs in a computer, inspection can be performed separately for each computer to discover which part of the computer has the defect. Furthermore, when replacing a defective computer, each computer can be replaced separately.

Furthermore, a monitoring computer configured similarly to the actuator control computer is used only to make decisions based on the majority rule. Therefore, the rudder blade drive member of each control channel is driven by the rudder blade drive signal output from the actuator control computer provided in each control channel, and not by the output signal of the monitoring computer.

Each control channel is provided with an interface device having majority decision processing means, and whether or not each control channel is normal is determined by the majority decision processing means within each control channel.

(3) Embodiment An embodiment of an aircraft multi-channel control device according to the present invention will be described below with reference to the drawings.

In FIG. 1, the aircraft multi-channel control device U has three channels C, , C2, and C3, and the amount of displacement of the control rod A operated by the pilot is detected by three sensors 10, 20, . and by 30 the steering command signal α
1. Detected as C2 and C3. Each sensor 1
0,20. and the steering command signal α0. detected by 30. C2 and C3 are input to flight control computers 11.21 and 31, respectively.

Further, the flight control computer 11.2
1 and 31, various state detection signals β1. β2 and β3 are input.

Said flight control computer 11゜2],
3] are the steering command signals α5. α2, α3 and state detection signal β1. β2. From β3, etc., the optimum steering angle command signal δ3 at that moment. δ2. Calculate δ3.

These flight control computers 11.21
and the steering angle command signal δ5 calculated by 31. δ2
and δ3 is Hassline! . are input to the other flight control computers 11, 21.31, respectively, and their values are compared with each other. Then, those rudder angle command signals δ1. Any one signal (for example, δl) among δ2 and δ3 is connected to the other two signals (for example, δ2
．． δ3), the rudder angle command signal δ1 is determined to be abnormal within the flight control computer 11 based on the principle of majority voting.

In this case, the flight control computer 11
From then on, either the steering angle command signal δ2 or δ3, which is determined to be normal, is output to the bus controller 13 instead of the steering angle command signal δ1.

However, each of the steering angle command signals δ, , δ2 and δ
3 is determined to be a normal signal if the difference between those values is within a predetermined range, and the HAS controllers 13 and 2
3 and 33.

The bus controllers 13.23 and 33 are devices for receiving and passing signals between the flight control computers 11.21 and 31 and the interface devices 14, 24 and 34, which will be described later. Accordingly, said flight control computers 11.21 and 3
The steering angle command signal δ1.1 outputted from the steering angle command signal δ1. δ2 and δ
3 are input to the interface devices ff14.24 and 34 via the Hascon 1 rollers 13.23 and 33, respectively.

Said interface devices 14.24 and 34 respectively connect path lines j2+, j2z and ! 3 etc. to other interface devices 14, 24 and 34, respectively.
It is configured to be able to send and receive signals to and from the monitor, and also to be able to send and receive signals to and from another monitoring interface device 44.

Each said interface device 14.24 and 34 comprises:
Actuator control computer 15 respectively
．． 25 and 35 in a transmissible manner, and the monitoring interface device 44 is connected to a monitoring computer 45 in a transmissible manner. The interfice devices 14, 24, and 34 are connected to the selection circuits 16, 26, and 36, and are adapted to output driving output control signals of +, kz, and 3 to them.

The interface device 44 connects the hash controller (13, 23 or 33) and the selection circuit (16, 2
The interface device 14 is configured similarly to the interface devices 24 and 34, except that it is not connected to the actuator controller 1-role computer 15, 25.
．． It is configured similarly to No. 35. The interface device 44 and the monitoring computer 45 constitute a monitoring channel Ca'.

the interface device 14, 24, 34° 44, the actuator control computer 15, 25;
．． 35 and the monitoring computer 45 will be detailed later.

The actuator computer 1 to roll computer 15.
25 and 35 are connected to motors 18, 28 and 38 as rudder blade drive members via motor drive circuits 17, 27 and 37 having power source B, respectively.

The motor drive circuits 17, 27, 37 receive rudder blade drive signals δ, 2, which will be described later, output from the actuator control computer 15, 25, 35.

This is a circuit for amplifying δ2' and δ3', and its output signal is input to the motors 18, 28, and 38 as the rudder blade drive members.

Whether or not the motor drive circuit 17, 27.37 outputs an output signal is determined by the drive output control signal k+, k2゜k3 input to the selection circuit 16, 26.36. It looks like this. and,
When the output signal of the motor drive circuit 17, 27.37 is input to the motor 1.8, 28.38, the motor 18,
28.38 is configured to operate. Then, as the motors 18, 28, and 38 operate, the secondary actuator P. is displaced, and this secondary actuator P. The power actuator P is configured to operate in accordance with the displacement. The rudder S is configured to operate in accordance with the operation of the power actuator P, thereby causing a change in the body motion of the aircraft.

Further, the position of the power actuator P is detected by position sensors 19, 29, and 39. The position signals of the power actuator P detected by the position sensors 19, 29 and 39 are signals corresponding to the actual steering angle, that is, the actual steering angle signal θ1. θ2. θ3.

This actual steering angle signal θ7. θ2. θ3 are fed back to the interface devices 14, 24, and 34, respectively.

Next, the interface device 1.4, 24°3'4
, 44, the actuator control computers 15, 25, 35, and the monitoring computer 45 will be described in detail.

Said actuator control computer 15.2
5. 35 and the monitoring computer 45 all have clocks 15a, 25a, 35a and 4
Actuator control computer 15.25 and 3 with 5a, but normally in good working order.
All systems operate in synchronization with one of the clocks 15a, for example, 15a.

The interface devices 14, 24. ,． 34 are the respective input steering angle command signals δ1. δ2゜δ3
are configured to be input to the other interface devices 14, 24, 34, 44 via the path lines fi, , ffi, , ffi.

Each of the interface devices 14, 24, 34°44 receives the steering angle command signal δ1. δ2. δ3 are compared, and their steering angle command signals δ0. δ2. If any one signal (for example, δ) among δ3 has a difference of more than a predetermined value from the other two signals (for example, δ2, δ3), the steering angle command signal δ is normal based on the majority rule. What is not (
It is determined that the steering angle command signals δ2 and δ3 are normal.

In this way, the flight control computer 1
1, 21.31, the normal steering angle command signal δ1. was all within the predetermined value range. δ2. δ3 is
If only the steering angle command signal δ1 falls outside the predetermined value range when compared within the interface devices 14, 24, 34 and 44, it is presumed that the bus controller 13 is defective. In this case, actuator control computers 25 and 35 connected to said interface devices 24 and 34, respectively.
The steering angle command signals δ2 and δ3, which are respectively determined to be normal, are input to the actuator control computer 15 and the monitoring computer 45 connected to the interface devices 14 and 44, and the normal steering angle command signal δ2 is input to the or δ3 (
For example, the configuration is such that δ2) is input.

However, each of the rudder angle command signals δ5. δ2. If the difference between these values is within a predetermined range, δ3 is determined to be a normal signal, and is input to the actuator control computers 15, 25.35, respectively, and the monitoring computer 45 is sent a normal signal. Each steering angle command signal δ0. It is configured such that one of the signals δ2 and δ3 (for example, δ1) is input.

The interface devices 14, 24.34 receive the actual steering angle signals θ4. Also for θ2 and δ3, the above-mentioned steering angle command signal δ4. δ2. It is configured to perform the same processing as for δ3.

Said actuator control computer 15.2
5.35 and the monitoring computer 45 receive the rudder angle command signals 6 and δ2, respectively. δ3 and actual steering angle signal θ3. θ2. From θ3, the rudder blade drive signals δ1', δ2', δ3', and δ4' at each moment are calculated. These rudder blade drive signals δ,'', δ2'', δ3'' and δ4'' are output to the interface devices 14, 24, 34 and 44, respectively. Note that the rudder blade drive signals δ1°, δ
2゛, δ3゛ are the motor drive circuits 17, 27.3''
/ is also output.

the interface device 14, 24.34° 44;
The rudder blade drive signals δ, ゛, δ2'' inputted to each
, 63'', δ4゛, the path line !1゜22.23
via other interface devices 14°24.34.
44, respectively.

The interface device 14, 24, 34° 44 is
Initially, each rudder blade drive signal δ1゛, δ2 input to them
'', 63', and the steering angle command signals δ5, δ2 .
If the difference in the values of δ3 is within a predetermined range, all signals are determined to be normal. At this time, the interface device 14
, 24.34, a signal for operating the motor drive circuit 17°27.37 is outputted to the selection circuit 16, 26.36. When the motor drive circuits 17, 27.37 operate, they receive the rudder blade drive signal δ1 input from the actuator control computer 15, 25.35.
It is configured to output amplified signals of 5 δ2 and δ3 to the motors 18, 28, and 38 serving as the rudder blade drive members.

However, the interface device W14°24.
34 and 44 first compare the respective rudder blade drive signals δ1'', δ2', δ3'' inputted thereto, and select one of the rudder blade drive signals δ1'', δ2'', δ, ''. One signal (for example, δ1゛) is connected to another two signals (for example, 6□
If the difference is greater than a predetermined value between the rudder blade drive signals δ2 and δ3, it is determined that the rudder blade drive signal δ1 is not normal based on the majority vote principle, and the rudder blade drive signals δ2 and δ3 are determined to be normal. do.

In this way, the interface device 14°24.3
4 to actuator control computer 15,
25. The steering angle command signal δ1 when inputted at 35. δ2
．． Even though δ3 was within the predetermined value range,
Actuator control computer 15, 25.
If the rudder blade drive signals δ1, δ2, and δ3 calculated 26 by 35 are outside the predetermined value range, it is presumed that the actuator control computer 15 is defective. In this case, the driving output control signal outputted from the interface device 14 to the selection circuit 16 is set to 1.
makes the motor drive circuit 17 inactive, and the drive output control signals kz, ks output from the interface device 24.34 to the selection circuit 26.36 actuate the motor drive circuit 27.37.

That is, actuator control computer 1
5.25 or 35, the defective actuator control computer is configured to be disconnected from operation of the aircraft multichannel control system.

As described above, if any one of the actuator control computers 15, 25 and 35 is defective, the rudder blade drive signal δ1' of the defective actuator control computer 15 is replaced by the monitoring The rudder blade drive signal δ4 of the computer 45
The principle of majority voting is applied to this rudder blade drive signal δ4' and the rudder blade drive signals δ2' and δ3' determined to be normal. When a difference of more than a predetermined value occurs between the rudder blade drive signals δ2'' and δ3', the principle of majority voting is applied including the rudder blade drive signal δ4'', and the rudder blade drive signal δ2'' or 63' is selected. It is determined which one of the rudder blade drive signals δ2゛ or δ3 is normal.
The signal is amplified by a motor drive circuit 27 or 37 and then inputted to the motor 28 or 38.

In the embodiment of the aircraft multi-channel control device U according to the invention with the above-mentioned configuration, the three control channels C++Cz+c+ are normally operated simultaneously. and,
The flight control computers 11, 2L31 of each control channel CI, CZ, C3 mutually transmit their output signals, that is, the rudder angle command signal δ1. δ2. By mutually monitoring δ3, a normal steering angle command signal and an abnormal steering angle command signal are distinguished based on the principle of majority voting. The path controllers 13, 23, and 33 receive a normally normal steering angle command signal δ1. δ2 or δ is input.

The steering angle command signal δ1. input to the path controller 13, 23.33. δ2. δ3 is first input to the interface device 14.24.34 and then to the other interface devices 14°24.34 and 44, respectively. Further, the actual steering angle signal θ1. θ2. θ
3 and the rudder blade drive signals δ1, δ2, and δ3 are also the rudder angle command signal δ19. δ2. Similar to δ3, it is first input manually to the interface device 14.24.34 and then to the other interface devices 14.24.34 and 44.

Then, the interface devices 14, 24, 34, and 44 use the majority vote processing means provided therein to process the steering angle command signal δ1. δ2. δ3 is mainly compared to determine whether the path controllers 13, 23, 33 are normal or abnormal, and the actual steering angle signals θ1. θ2. θ, to determine whether the rudder angle sensors 19, 29, 39 are normal or abnormal, and further, by comparing the rudder blade drive signals δ1, δ2, δ, and δ, the actuator control computer 15.25.35 is determined to be normal or abnormal.

Furthermore, by comparing the signals, it is determined whether each interface device 14, 24, 34, and 44 itself is normal or abnormal.

said interface devices 14, 24, 34 and 44;
control channels CI, C2 by each majority processing means of
, if it is determined that an abnormality has occurred in any of C3,
The interface devices 14, 24, 34 belonging to the control channel in which the abnormality occurred are the selection circuits 16, 26, .
1. to the drive output control signal output to 36. kz, 'ni'
a disables the morph drive circuit 17, 27.37.

In this way, the control channels CI, C where the abnormality occurred
2, C3 is rendered inactive, and after that, the monitoring channel is used instead of the control channel where the abnormality occurred, and a majority decision is made based on this monitoring channel and the remaining normal control channels. Apply.

Although the embodiments of the aircraft multi-channel control device according to the present invention have been described in detail above, the present invention is not limited to the above-described embodiments, and there is no deviation from the scope of the present invention as set forth in the claims. However, various design changes can be made.

For example, the number of control channels can be any number greater than three instead of three, and the computer portion of each control channel can be
Instead of having two separate flight control computers and actuator control computers, three
It is also possible to divide it into more parts. Additionally, monitoring channels can be provided not only for the actuator control computer but also for the flight control computer, and the number of monitoring channels for each computer can be two or more. It is. Furthermore, the monitoring channel can be configured only by the monitoring computer without the interface device. Furthermore, the interface device can be configured with a wired logic circuit, but it can also be configured with a microcomputer. It is also possible to configure the interface device and the actuator control computer with one microcomputer.

C9 Effects of the Invention According to the aircraft multi-channel control device of the present invention described above, when an abnormality occurs in a control channel during operation of the aircraft multi-channel control device, the monitoring channel is switched on instead of the control channel in which the abnormality has occurred. A majority decision can be made using .

In addition, the control channel computer is divided into a flight control computer and an actuator control computer, and the divided computers (actuator control combination j, -
By configuring the monitoring computer in accordance with the above, each computer can be configured compactly, so if a defect occurs in one of the computers, only the defective computer can be replaced. Bayoshi)
Therefore, costs can be reduced.

[Brief explanation of the drawing]

FIG. 1 is an explanatory diagram of an embodiment of an aircraft multi-channel control device according to the present invention, and FIG. 2 is an explanatory diagram of a conventional aircraft multi-channel control device. CI, C2, C3...Control channel, C4...Monitoring channel, k+, kz, kz...Drive output control signal, δhδ2. δ3... Rudder angle command signal, θ1. θ2
．． θ3... Rudder angle signal, δ1゛, δ2', δ3゛...
Rudder blade drive signal, 10°20.30...sensor, 11,
2], 31... Flight control computer, 14, 24, 34. 44... Interface device, 15, 25° 35... Actuator control combination 54-Y, 45... Monitoring computer, 1
8.28.38... Rudder blade drive member (motor) Low Figure 2〃L0 port] Total - Column 7''11

Claims (1)

[Scope of Claims] An actuator control computer that receives a rudder angle command signal and an actual rudder angle signal from a flight control computer and outputs a rudder blade drive signal calculated from each of the input signals, and the rudder blade drive signal. At least three systems of control channels each having an actuated rudder blade drive member are provided, and each rudder blade drive signal of each control channel is compared, and when the rudder blade drive signals do not match, the majority rule is applied to A driving device that determines that matching rudder blade drive signals are normal and a small number of rudder blade drive signals are abnormal, and prevents output of the rudder blade drive signals determined to be abnormal to the rudder blade drive member. A multi-channel control device for an aircraft equipped with a majority vote processing means for outputting an output control signal, configured similarly to the actuator control computer, and receiving a rudder angle command signal and an actual rudder angle signal from the flight control computer. and a monitoring channel equipped with a monitoring computer that outputs a rudder blade drive signal calculated from each of the input signals; A multi-channel control device for an aircraft, comprising an interface device into which a signal is input, and each of the interface devices provided for each of the control channels is provided with the majority voting processing means.