JPH0358959B2 - - Google Patents

Description

DETAILED DESCRIPTION OF THE INVENTION [Industrial Field of Application] The present invention relates to an aircraft control system, and more particularly to an aircraft control system that enables control in a completely new manner.

[Prior Art] In both fixed-wing aircraft and rotary-wing aircraft (helicopters), pilots use the control stick, control levers,
Using control means such as control wheels and foot pedals,
Control the position of the aircraft's control surfaces to control the aircraft's altitude,
It controls posture, speed, etc. In the simplest systems, the flight controls are connected to the control surfaces by cables. For example, in fixed wing light aircraft, foot pedals are connected to the rudder by cables. More complex systems use steering devices with mechanical connections or the like that are augmented by hydraulic servos.

As aircraft systems become increasingly complex, a large number of instruments, switches, etc. are disposed within the cockpit, and these instruments, switches, etc. occupy much of the space around the cockpit. Therefore, it has become very difficult to arrange the control device and other devices within the cockpit.

In a typical aircraft, control wheels on the control stick control the roll and pitch of the aircraft, foot pedals control the rudder, and a throttle console controls engine thrust. On the other hand, in a typical helicopter, a cyclic pitch stick controls pitch and roll of the aircraft, a foot pedal controls yaw, and a collective pitch stick controls vertical lift. These control devices and mechanical connections or servomechanisms to the control surfaces occupy a large portion of the cockpit space. For example, if a control wheel or control stick is placed in front of the pilot's seat, the control wheel or control stick must be moved to various positions within the space, and these wheels or control sticks may obscure the pilot's field of view. This makes it difficult to place electronic displays in front of the pilot. Additionally, foot pedals may obstruct the pilot's forward and downward visibility during helicopter logging operations, construction work, and the like. Also, when a passenger is seated in one of the pilot seats, inadvertent contact of the passenger with the controls may result in unintended operation. Moreover, these controls become a nuisance when approaching and leaving the pilot's seat.

In systems where control is performed by a pilot and a copilot, the control systems are synchronized in position with each other so that handover from one to the other can occur without sudden inputs to the control system. This is essential. To this end, each pilot's controls are typically mechanically coupled to a corresponding co-pilot's controls. The hydraulic or electrical sensors and actuators required to eliminate the mechanical linkage are too slow and cumbersome to be practical for this purpose.

In order to overcome some of the above-mentioned drawbacks, attempts have been made to provide so-called "side arm" controls that can be operated by the pilot with his hands resting on the seat arm. This side-arm type control device allows the pilot to operate the controls from within the seat because the pilot's body must be supported by cushions on the seat in high-speed aircraft or spacecraft where the pilot must withstand high acceleration. It is considered to be effective because it can be obtained. For this reason, several sidearm control devices have come into use.

Typical sidearm control devices that have been used with some success in the past are limited to control in two axes, generally in the pitch and roll axes. Therefore, the throttle or collective
The pitch stick and foot pedal remain as they were, so to operate them the pilot must reach outside the seat and assume a position that allows him to press the foot pedal. Furthermore, it is not sufficient to eliminate the complexity of the seating room equipment.

Attempts have also been made to create sidearm controls that are operable in three or more axes. These axes include, for example, pitch, roll and yaw axes, or pitch, roll and collective pitch axes (or throttle axis in fixed wing aircraft). However, sidearm control systems designed to operate in three or more axes suffer from common drawbacks due to cross-coupling between each control axis. That is, when performing pitch direction control and yaw direction control by forward and backward and right and left movements,
It is also impossible to control the collective pitch of a helicopter by the same up-and-down movement of the control stick. This is because attempting to move the control stick forward and backward will result in the control stick also moving upward and backward to some extent (and vice versa). This is an inherent problem because the human hand is connected to the forearm using the wrist joint as a pivot point, and the natural movement of the human wrist joint causes cross-coupling between the different axes of the control stick. It is thought that. The same applies to torsional movements combined with forward-backward and right-left movements.

Mechanical interconnections within the aircraft are avoided to reduce aircraft weight, provide redundancy in systems to improve reliability and performance, and to take advantage of emerging technologies such as microprocessors. There has also been some consideration of "fly-by-wire" systems characterized by electrically or optically (or both) connecting sensors and actuators.
In such cases, the usual mechanical linkages driving booster servos for control of the aircraft's control surfaces are replaced by sensors in electrical positions, thereby controlling electro/hydraulic actuators. However, it has been difficult to implement fly-by-wire systems that can provide synchronization between pilot and co-pilot controls without increasing complexity and seat-mounted equipment. It was hot.
Therefore, in fly-by-wire systems currently configured for use in aircraft with general control means, the pilot's control means and the co-pilot's control means are mechanically connected, and the single It is commonly proposed to couple an electrical transducer to one mechanical link. This means that when a handover occurs between a pilot and a co-pilot, the position of the control device (for example, the control stick of a helicopter or the control wheels of a fixed-wing aircraft) is the same for the pilot and co-pilot. However, this synchronization of motion or position cannot easily be achieved other than by mechanical interconnections, due to the inherent difficulty in implementing high-speed tracking systems that do not take up excessive space. Ta.

An object of the present invention is to solve the above-mentioned problems in conventional aircraft control systems, and to provide multiple control functions such as pitch, roll, and yaw attitude control, lift control, and airspeed control using a single control device. It is an object of the present invention to provide a flight control system that can perform control and reduce the burden on the pilot in flight operations.

[Means for Solving the Problems] According to the first configuration of the present invention, for manually controlling at least one of attitude control, lift control, and speed control in pitch direction, roll direction, and yaw direction of an aircraft. A control device comprising: a control surface for controlling one of the attitude, lift, and speed of the aircraft in pitch, roll, and yaw directions; and a control surface for operating the control surface. At the same time, it operates with minute strokes in each operation direction to generate an operation input signal indicating the magnitude of the operation force and the operation direction according to the operation force, and as long as the input of the operation force continues, the operation is performed. a single manual input device that continues to output an input signal and stops outputting the operation input signal when the operation force is released; and a control surface that operates the control surface in accordance with a control value of the control signal. A servo mechanism for surface drive; and a signal processing means for generating a control signal according to the signal value of the operation input signal inputted by the manual operation input device, and the signal processing means includes: generating a proportional component of the control signal proportional to the signal value of the operation input signal and an integral component of the control signal by integrating the signal value of the input signal, and adjusting the control surface according to the proportional component and the integral component. determining a control value of the control signal that commands a control position, and supplying a control signal having the determined control value to the servo mechanism to control the control surface according to the signal value and input duration of the operation input signal. While there is no operation input signal, the integral component at the time of extinction of the operation input signal is held as the floating correction point corresponding to the position of the control surface, and the control surface is adjusted to the operation input. An aircraft control system is provided that operates to maintain a position at the time of signal disappearance.

Further, according to the second configuration of the present invention, attitude control in the pitch direction, roll direction, and yaw direction of the aircraft;
A control device for manually controlling at least two of lift control and speed control, wherein the control device manually controls at least one of the attitude, lift, and speed of an aircraft in pitch direction, roll direction, and yaw direction. first and second control surfaces for controlling two control surfaces; and operated for inputting operational inputs for controlling said first and second control surfaces;
Generates an operation input signal indicating the magnitude and direction of the operation force in accordance with the applied operation force by operating in each operation direction with a minute operation stroke that the pilot does not perceive the stroke, and inputs the operation force. a single manual input device that continues outputting the operating input signal while the operating force continues and stops outputting the operating input signal when the operating force is released; and a servo mechanism for driving a control surface that operates a wing surface; and while the control input in the same control direction input from the manual control input device continues, the same control input by the manual control input device continues. signal processing means for generating a corresponding control signal for the first or second control surface according to the signal value of the directional operation input signal, the signal processing means comprising: a proportional component of the control signal proportional to the signal value of the control signal, and an integral component of the control signal by integrating the signal value of the input signal, and control the control position of the control surface according to the proportional component and the integral component. determining a control value of the control signal to be commanded, supplying a control signal having the determined control value to the servo mechanism, and controlling the position of the operating blade surface according to the signal value and input duration of the operation input signal; control, and while the operation input signal is absent, the integral component at the time when the operation input signal disappears is held as the floating correction point corresponding to the position of the control surface, and the control surface is adjusted to the point at which the operation input signal disappears. An aircraft control system is provided that operates to maintain the aircraft in the position of the aircraft.

[Function] The present invention is configured in the first and second configurations as described above, and allows the pilot to perform a pilot operation using a manual operating device by setting the operating stroke of the operating device to be very small. The control value of the control signal is determined by the magnitude, direction, and input duration of the operating force applied to the control device, and the control position of the corresponding control surface is determined. In this manner, the operating direction, operating speed, and operating amount of the control surface to be controlled can be controlled. That is, the direction of the operating force acting on the control device corresponds to the direction of movement of the control surface,
The magnitude of the operating force corresponds to the operating speed of the control surface,
The integral value of the operating force and input duration corresponds to the amount of movement of the control surface. In addition, the control value of the control signal is fixed to the control value at the time when the operation force is released when the operation input to the control device is released, and the control value is fixed to the control value at the time when the operation input by the next control operation is input. The signal value of the signal is updated using the previously fixed control value as a reference value.

Therefore, the pilot can control pitch direction, roll direction,
In attitude control in the yaw direction, lift control, and airspeed control, when the behavior of the aircraft reaches a desired behavior in one control, the operation input to the flight control device can be canceled. Thereby, in particular, a plurality of control operations can be performed with a single control device, and a combined control operation of two or more of the above-mentioned controls can be performed easily and with a light burden.

[Example] The above and other objects, features and advantages of the present invention are as follows:
This will become clearer as preferred embodiments are described in detail below with reference to the drawings.

Referring to FIG. 1, a sidearm control device 10 according to the present invention includes a detection transducer 13;
It has a control stick 12 coupled to the transducer. The control stick 12 and transducer 13 constitute a manual control input device. The unit formed by the transducer 13 and the control stick 12 is arranged above the armrest 14 of the cockpit 16. In addition, the armrest 14 can be attached to the cockpit so as to be rotatable around the part marked with reference numeral 18 in order to facilitate seating in and leaving the cockpit 10 as needed. . As shown in FIG. 1, the side arm control device 10 can be operated in the front-rear direction, right-left direction, up-down direction, and rotational direction about the up-down axis. In this embodiment, the pitch direction of the aircraft is controlled by operating the control stick 12 in the longitudinal direction.
Thus, in a helicopter, the longitudinal cyclic pitch channel or the elevator of a fixed-wing aircraft is controlled. In addition, the control stick 12 is configured to control the roll direction of the aircraft by operating the control stick 12 in the left and right directions, and the lateral cyclic pitch channel of a helicopter or the ailerons of a fixed-wing aircraft are controlled. . Control stick 12 of control device 10
By manipulating the rotor in a rotational direction about a vertical axis, yaw control of the aircraft is achieved, which in turn controls the tail rotor pitch channel of a helicopter or the rudder of a fixed wing aircraft. Furthermore, in the illustrated embodiment, the lift or speed can be controlled by vertically operating the control stick 12 of the control device 10, which is useful in the collective pitch channel of a helicopter or in a fixed-wing aircraft. engine throttle and/or propeller blade pitch are controlled.

In the side arm control device 10 according to the preferred embodiment of the present invention, the operating stroke of the control stick 12 in the control operation in all or some of the four operation directions described above is the same as that of the transducer 13.
The transducer 13 is set to the minimum stroke that can be detected by the transducer 13.
detects not the operating stroke of the control stick 12 but the operating force applied to the control stick, and generates an output signal according to this operating force.

By making the motion stroke of the control stick minute as described above, it is possible to reduce the number of maneuvers in multiple directions that occur due to the natural movement of the wrist joints during pilot operations. This makes it possible to prevent cross-coupling (for example, a phenomenon in which a backward stroke motion of the control stick is caused due to an upward steering operation). Therefore, the steering operations in the four directions described above can be performed independently. This type of force-sensitive control stick is easily available on the market, and one example is the Model 404 manufactured by Measurement Systems, Inc., Norwalk, Connecticut, USA. -There is G517. Other force-sensitive control sticks are also available. Note that the necessary conditions for use in the control system of the present invention are that the control stick has sufficient rigidity in all operating directions, and that the control stick has an appropriate range of operation (for example, on the order of 0 to 18 kg in each operating direction). It has the sensitivity to produce the required output signal in response to the input. Furthermore, a strain gauge or the like can be used as a transducer for detecting such operating force.
According to the configuration of this embodiment described above, the pilot hardly senses the movement of the control stick during piloting operations. Here, motion that is not perceivable by the pilot means that the displacement caused by the operating force is so small that it does not give the pilot a sense of movement;
This refers to a degree of movement in which the cross-coupling between axes that accompanies hand movement does not become a problem.

As described above, according to this embodiment, the multi-axis operating force type control stick is sensitive only to the operating force input through the control stick, and the operating stroke of the control stick is so small that it cannot be detected by the pilot. The use of this eliminates cross-coupling between different operating directions in three- or four-axis maneuvers.

On the other hand, when using this type of control device 10,
Continuing to apply a nearly constant force can make the pilot feel fatigued. It is clear that if the pilot continues to perform maneuvers in a plurality of directions, the pilot will become more fatigued by continuing to apply a constant maneuvering force.

Furthermore, it is difficult to single-handedly maneuver rapid flight movements in multiple maneuvering directions, such as a 180° turn of a helicopter in hover under gusty wind conditions. While not all ergonomic factors, including the function of the hands themselves and the pilot's response to the aircraft's responses, are not fully understood, the above-mentioned difficulties arise when maneuvering in more than one direction during such complex flight movements. This is thought to be due to the need to give commands with mutual correlation between the pilot operations. The force-sensitive control stick used in the present invention differs from conventional control sticks that give commands according to the stroke position of the control stick, and the control stick itself does not provide visual or visual feedback or the position of the hand relative to other body support parts (e.g., knees). You cannot steer by feeling. In addition, in conventional control devices, control operations are assigned to various body parts, and the only thing that needs to be controlled with one hand in relation to each other is the control stick or control wheel that controls the pitch and roll directions. In contrast, the steering system of the present invention requires one-handed control in all directions.

In order to solve the above-mentioned problems associated with the use of a multi-axis operating force-sensitive control stick, the control system according to this embodiment uses a signal processing circuit that uses proportional components and integral components (PI control components). The control starting point for each direction is floated. That is, a new control starting point (or reference point) for each corresponding control direction is updated by all operation inputs given by the pilot. Therefore, according to the present invention, the pilot can perform maneuvers to correct the current position of the control surface while knowing the attitude, speed, altitude, and changes thereof of the aircraft by visual observation or instruments. .

FIG. 2 shows a block diagram of a signal processing circuit of a control system according to the present invention using the multi-axis operating force sensitive control stick described in FIG. The control device 10 has a plurality of output ends 20 to 23, each of which corresponds to the operating force input to the control stick 12 in the vertical direction, front-rear direction, left-right direction, or rotational direction. A voltage signal having a voltage value that is a predetermined function is generated. The control stick 12 described in FIG. 1 has bidirectionality in each operating direction. That is,
Vertical operations include upward and downward operations, forward and backward operations include forward and backward operations, and left and right operations include right and left operations. There are several operations, and the operations in the rotational direction include clockwise and counterclockwise operations. The control device 10 of this embodiment is configured to generate voltages with different polarities depending on the operating direction of the control stick 12 in each control direction. In the above-mentioned operating force sensitive control device, the voltage has a substantially linear relationship with the operating force, and therefore, in this embodiment, the absolute value of the output voltage generated at the output terminals 20 to 23 is: It is linearly proportional to the operating force. However, it is not essential in implementing the present invention that the output voltages of the output terminals 20 to 23 be linearly proportional to the operating force.
Even when the relationship between the output voltage and the operating force is non-linear, signal matching circuits 24 to 27 are provided for each of the output terminals 20 to 23, and the signal matching circuits 24 to 27 adjust the operating force. This is because voltage signals having a desired functional relationship to the voltage signals can be generated on the conductors 28-31, and the output voltages of the signal matching circuits 24-27 can be the actual signal inputs to the control system.

An example of signal matching performed by circuit 26 is shown in FIG. The horizontal axis represents the lateral operating force to the left or right, and the vertical axis represents the output voltage of the circuit 20 on the conducting wire 30. This signal matching is of course a voltage-matching depending on the operating force-voltage relationship of the signal on conductor 22. However, in this example, the desired functional relationship is such that the control system is not activated when the control column is slightly deflected from its lateral zero due to inadvertent contact or other reasons. Approximately 0.5lb (0.23Kg) on the left side
A dead zone is provided. This dead zone is necessary, as mentioned above, to prevent small unintended signals from being integrated over long periods of time. For forces from 0.5lb (0.23Kg) to 4lb (1.8Kg) in each direction, the output voltage increases linearly from 0V to 0.8V. Furthermore, the output voltage increases non-linearly for forces of 4 lb (1.8 kg) or more in each direction, and the sensitivity is set high in the range where the input operating force is large. In Figure 3, the voltage-operating force relationship is shown as a non-linear relationship whose slope increases as the force increases, but other elements of the control system, such as the characteristics of the fluid pressure servo, the flight characteristics of the aircraft, the desired response, etc. Depending on the equation, any functional relationship suitable for implementing the invention can be selected.

The signal matching circuit used to perform the signal matching operation shown in FIG. 3 can be easily realized by an appropriate combination of a biased amplifier and a limited amplifier, as illustrated in FIG. The signal matching circuit 26 in FIG. 4 includes six amplifiers 26a.
~26f. Amplifiers 26a and 2
6b is a biased amplifier, 0.5lb (0.24Kg)
The gain is 0 up to the input voltage representing the force of
Above that, the gain is 1. These amplifiers can easily provide a deadband of ±0.5lb (±0.24Kg). Amplifiers 26c and 26d are limited amplifiers, with a gain of 0 for input voltages of one polarity, but a gain of 0.2V/1lb (0.24Kg) for input voltages of the other polarity. However, after the input voltage reaches the 4lb (1.8Kg) limit, it saturates at an output voltage of 0.8V. These amplifiers form a region of low sensitivity. Furthermore, the amplifier 26
e and 26f are biased amplifiers,
For input voltages below 4V, the gain is 0, but for input voltages exceeding 4V, the gains are set to be greater than the gains of the amplifiers 26c and 26d. These amplifiers form a sensitive area. The output voltages of the amplifiers 26c to 26f for the low sensitivity region and the high sensitivity region are at the addition point 2.
6g will be added. The summing point may be constructed from a dedicated summing amplifier, or may be constructed as an input circuit to a circuit with proportional-integral characteristics shown in FIG. 2, which will be described later.

The characteristics of the pitch and yaw channel signal matching circuits are not limited to those described with reference to FIG. 3, and any circuit that performs similar operations can be used. In the embodiment of the present invention, the characteristics of the pitch channel signal matching circuit are the same as those of the roll channel signal matching circuit shown in FIG. 3, and the characteristics of the yaw channel signal matching circuit are the same as the roll channel signal matching circuit characteristics shown in FIG.・Channel characteristics and gain
They are the same except that they are 0.225V/1 inch·lb (0.01 m·Kg) and the dead zone is ±0.27 inch·lb (±0.003 m·Kg).

On the other hand, the collective pitch channel or vertical channel signal matching circuit differs in characteristics from the other three channel signal matching circuits in that the slope decreases as the force increases. As in the example shown in Figure 5, a vertical channel requires a maximum control stick input of 40 lb (18.2 Kg) of force (instead of 20 lb (9.1 Kg) as in the left/right and up/down directions). shall be. A dead band of ±1lb (±0.45Kg) is not provided, and the gain in the linear part is
On the order of 0.19V/1lb (0.45Kg), below
It is on the order of 0.8V/1lb (0.45Kg). moreover,
Figure 5 shows that the slope decreases (instead of increasing as in pitch, roll and yaw channels) with increasing force to accommodate the droop characteristics between collective pitch and airspeed. ing. The voltage-operating force relationship in any operating direction is realized by a combination of several biased amplifiers and limiting amplifiers as shown in Figure 4, and the dead band and gain are independent on the positive and negative sides. Can be adjusted. If a suitable digital computer is used, Figures 3 and 5 can be used.
The characteristics shown in the figures can also be obtained using lookup tables based on the magnitude of the voltages on conductors 20-23, and such methods are described in a patent application herein assigned to the same assignee. August 1978
No. 938,583, entitled "Full-Operational, Full-Safe Multicomputer Control System," filed on June 31, 2003. Referring to FIG. 2, the signals on matched conductors 28-31 are provided to a plurality of amplifiers 32-39. Among them, amplifiers 32 to 35 are proportional amplifiers, and amplifier 36
-39 is an integrating amplifier. Therefore, the amplifier 32~
39 provides the proportional/integral gain of the control input to the aircraft control surfaces. The pairs of outputs produced by each of the amplifiers on corresponding leads 40-47 are summed together with the negative feedback signals on corresponding leads 54-57 at summing points 50-53, respectively. The output of each summing point on the corresponding conductor 60-63 is a position error signal which, via appropriate amplifiers 64-67, controls a solenoid valve 70-73 of a hydraulic servo 74-77. The mechanical outputs 80-82 of the three servos 74-76 are sent to the switch via a mixer 84.
Controls mechanical inputs 86-88 to plate 90. Thereby, the pitch of the blades of the main rotor 92 is controlled. Yaw servo 77 mechanical output 9
4 through the pitch beam 96 to the tail rotor 98
control the pitch of the plate.

In addition, the proportional amplifiers 32 to 35 and the integral amplifier 36
~39 and additional points 50 to 53 are for longitudinal cyclic pitch channel, transverse cyclic pitch channel, tail rotor pitch channel of an aircraft, and collective pitch channel of a helicopter or engine throttle of a fixed wing aircraft. and a signal processing means for generating a control signal for controlling the position of each control surface in each control channel of the propeller blade pitch channel in accordance with the signal value of the manual input signal inputted via the side arm control device. Configure. In the illustrated embodiment, these signal processing means are shown as having an analog circuit configuration, but it is of course possible to configure them as a digital circuit using a digital computer or the like.

Each servo 74-77 has a position sensor 100-10
3, mechanical output of each servo 80~8
An electrical signal indicating the position of 2,94 is emitted on the corresponding conductor wires 104-107. These signals are provided to feedback leads 54-57 via amplifiers 108-111, respectively, for appropriate scaling and isolation purposes. Therefore, whenever the actual position of the mechanical output of each servo does not match the target position, the electromagnetic valves 70 to 73 are driven via the amplifiers 64 to 67 by the deviation signals on the conductors 60 to 63, and the valves are activated. The servo brought into an unbalanced state is transferred from the working fluid source 113 to the conduit 11.
Pressurized working fluid applied via 2 moves the piston to modify the actual position of a mechanical output coupled to the piston. Thus, the actual position of each servo's mechanical output always tracks its target position. The above servos and various devices 6
All of 4 to 113 can be constructed from publicly known materials. However, the servos 74 to 77 are
It must be an electrically controlled servo at high speed and up to a specified limit, rather than an electrically modified mechanical booster servo as is conventionally used to control control surfaces. Servos suitable for use with the present invention are readily available.

A description of the operation of one of the control systems shown in FIG. 2 for one control direction is sufficient to demonstrate the novelty of the flight control according to the invention. For example, when the pilot wishes to increase the collective pitch, he applies an upward force to the control stick, and an electrical signal that is a function of the force applied appears at the vertical axis output 20. This signal is converted by the signal matching circuit 24 into a signal that follows the functional relationship illustrated in FIG. 5, and is applied to the conductor 28 as a maneuver command signal. Proportional amplifier 32 immediately amplifies the signal on lead 28 and provides it via lead 40 to summing point 50 as its one input signal. As a result, the output of summing point 50 becomes unbalanced. This is because the position sensor 100 of the servo 74 provides a signal indicating the current position of the mechanical linkage 80 via the conductor 54 to the feedback input of the summing point 50. Therefore, additional points 50
A deviation signal appears on the output lead 60 of the amplifier 64 which, upon amplification, drives the solenoid valve 70 from its balanced condition to its unbalanced condition, thereby actuating the servo 74 to cause the mechanical linkage 80 to move as desired. Drive in the direction of. The servos 74 to 77 are capable of driving the control surface over its entire specified limit range in a very short time on the order of one second. Signal matching circuit 24 and amplifiers 32, 6
With the gain of 4, if the operating force applied by the pilot to the control stick 12 of the operating force-sensitive control device 10 is greater than a certain limit, a signal of sufficient magnitude is given to the solenoid valve 70. Thus, servo 74 applies maximum fluid pressure to its piston to provide maximum acceleration force to mechanical linkage 80. On the other hand,
If the force applied by the pilot is small, first proportional amplifier 32, summing point 50 and amplifier 6
The signal applied to the solenoid valve 70 through the servo 74 is so small that it does not actually move the piston within the servo 74.

In order to achieve complete steering in a system having only a proportional gain as described above, it is necessary to provide a signal on lead 40 that is parallel to the signal fed back onto lead 54 at the desired location of the mechanical output, for example. The necessary force must be continuously applied by the pilot to the control column of the force-sensitive control system for the period of time to maintain its mechanical output position. It is obvious that if this period lasts for a long time, for example, several tens of minutes, the pilot will become fatigued. Several operating directions (if the invention is used in a steering system for controlling four control directions: roll, pitch, yaw and vertical (lift) or fore-and-aft (velocity)) Having to continue to apply force in all directions at the same time further fatigues the pilot.

In order to solve the above-mentioned fatigue problem, in a control system using conventional position-sensitive control means, that is, a control system in which the position of the control stick etc. and the position of the control surface correspond to each other, the control surface The position of the control means is restrained by, for example, a detent spring until the position needs to be corrected, and when correction is made, the restraint of the position of the control means is released by resisting the spring force of the detent spring. Then, a method is adopted in which the control means is moved to a new correction point and the position of the control means is restrained again there. However, it is impossible to adopt such a method in a control system in which control operations in three or four control directions are performed using a control stick of a one-handed control force-sensitive control device, and which has only a proportional gain. be. There are several reasons for this, but firstly, if you try to restrain the control stick in one operating direction using the button attached to the control stick itself of a one-handed operation force-sensitive control device, your hand will not touch the button. A slight movement when touching can cause errors in command input in that operating direction or in other operating directions. Secondly, even when the restraint is released, the command input is distorted, so that it is almost impossible to restore it to the original state in each operating direction, even if a force indicator is provided. Third, providing special servomechanisms for holding position or force in each of the four operating directions is complicated and negates the benefits of a sidearm control stick. Furthermore, if it is desired to change the control surface over a certain period of time, the control stick must be operated in accordance with the lapse of time, so the above method does not reduce the burden on the pilot. For these reasons, a control system having only a proportional gain is insufficient in order to take advantage of the above-mentioned advantages of employing a multi-axis operating force-sensitive control device.

Therefore, another feature of the present invention is that a signal processing circuit with proportional/integral characteristics is used in combination with a multi-axis operating force sensitive control device. The feature obtained thereby is the manner in which the motion of the control surface in each control direction follows the control operation in each control direction in response to a command given by the multi-axis force-sensitive control device.
In an embodiment of the invention, the position of the control surface in each control direction is determined by a proportional component obtained by proportional amplifiers 32-35 and a predetermined feed forward integral gain provided in parallel with proportional amplifiers 32-35. It is determined by the integral components obtained by the integral amplifiers 36 to 39 having the following values.

That is, in the above operation example, when the pilot gives a signal indicating a desired change in the position of the linkage 80 using the control stick of the force-sensitive control device, the servo 74 is instantaneously operated as described above. This is done in accordance with a command based on the proportional component output from the proportional amplifier 32 to the conductor 40. but,
Before servo 74 reaches a position where it feeds back a signal equal to the proportional component of lead 40 onto lead 54, an integral component signal from integrator amplifier 36 having the same polarity as the proportional component signal of lead 40 is applied to lead 40.
given to 4. The time constants of the integral amplifiers 36 to 39 are set so that operational inputs in all control directions can be input in a time width (for example, on the order of seconds) that corresponds to the pilot's reaction to the behavior of the aircraft. Thus, typically, if a pilot wishes to modify a control surface by a certain amount, very small inputs that cause the pilot to immediately revert will achieve the desired result.
This is because servo 74 initially responds to the proportional signal on lead 40 and then quickly reaches a steady state where the feedback signal on lead 54 and the integrating amplifier output signal on lead 44 are balanced. If you want to change the position of the control surface largely and slowly,
The pilot only needs to continuously give small inputs. In this case, if the signal value of the conductor 20 corresponding to the operating force input through the control stick is small, the output signal of the integrating amplifier 36 will also be correspondingly small, and the
By continuously applying a signal with a small signal value to 0, the output signal of the integrating amplifier 36 increases over time to far exceed the proportional component signal applied to the conductor 40 (as described below). (up to a maximum value limited to ), the servo 74
The position of linkage 80 continues to be rotated until the feedback signal above balances the sum of the proportional signal on lead 40 and the integral signal on lead 44.

In fact, the operation of the servo is controlled by the operation force-sensitive control device of this embodiment whose operation stroke is so small that it is not sensed by the pilot, and by the control signal containing the proportional component and the integral component of this control device. In combination with a controlling proportional/integral signal processing channel, the pilot continues to apply force on the control stick until the desired aircraft behavior is detected, and the integrated control stick output signal is in equilibrium with the feedback signal. By operating the control stick to return the force to zero, the aircraft can be easily controlled. Therefore, regarding the four control directions shown in FIG. 2, the multi-axis operation force sensitive control device according to this embodiment has a floating correction point in each operation direction, and each servo mechanism 74 to 77 has a corresponding one. The locations of mechanical linkages 80-82 and 94 that
The feedback signals on the conductors 40 to 57 are controlled to be in equilibrium with the integral signals on the corresponding conductors 40 to 47, thereby controlling the position of the corresponding control surfaces. Each time an aircraft is operated, the pilot applies the time and operating force to the control column of the force-sensitive control system to change the floating correction point in either operating direction by the desired magnitude and at the desired rate of change. This is done by adding in the direction. The pilot is
Only when you want to change the position of the control surface do you apply force to the control stick to change its floating correction point.
If the force applied to the control stick is zero, the float correction point will not change. In reality, in order to avoid large commands being generated due to the unintended integration of very small forces, a dead zone is provided in the signal processing circuit, so the force applied to the control stick is within the range corresponding to this dead zone. If the force is very small, the floating correction point will not change. Therefore, according to the present invention, the pilot can fly with his hands lightly touching the control stick, and with his hands off the control stick during steady flight.

In FIG. 2, the yaw channel conductor 31 is connected to an additional integrating amplifier 117. This integral amplifier provides a signal that integrates the operating force in the rotational direction to the wheel operating mechanism 119 via a conductor 118. This is not an essential part of the invention, but if the foot pedals are omitted (to give the pilot a better view of the ground around his feet and to reduce the weight of the control system) , added to indicate the fact that the control column can be used for both ground wheel operations and flight movements.

A conductor 114 in the upper left corner of FIG. 2 provides a signal indicating that the aircraft has landed, ie, that the wheels or skids have touched the ground. Such signals may be provided by "squat switches" or derived by some other means from the aircraft's wheels or skid support mechanisms. Such signals are commonly used in many aircraft for various purposes, such as locking the operation of autopilot stabilizers on the ground. The signal on conductor 114 is provided to each of integrating amplifiers 36-39 to act as an integral hold signal. In one example, this signal opens an electronic switch connected to the feedback circuit of the integrating amplifier, breaking the connection between the integrating capacitor and the amplifier input. Thus, when the aircraft contacts the ground, the float correction point is held constant at its instantaneous value and the pilot then completes the flight motion through the proportional circuit only. When the aircraft is shut down, all floating correction points are electrically reduced to zero by a reset switch or other known means. Afterwards, when the aircraft resumes operation, conductor 11
The signal on 4 holds all of the integrating amplifiers at their initial values, ie, zero. Therefore, there is no risk of command integration due to stray inputs to the control system during shutdown or retraction. In this way, unintended control inputs cannot be present at the beginning of takeoff, since all control surface correction points are guaranteed to be in a neutral position during takeoff. Therefore, takeoff is performed by the pilot through a proportional loop only. The signal on conductor 114 is applied via inverter 116 to integrating amplifier 117 as a complementary signal. The output of the integrating amplifier 117 is used to operate the wheels on the ground, if any.

Referring now to FIG. 6, there is shown a device for indicating margins to specified limits as may be required in the system of the present invention. In conventional systems, mechanical linkages that are actually activated by the pilot via the control stick, control levers, control wheels or foot pedals are used to alert that a specified limit has been reached in a given wheel. It includes position responsive means. As an alternative to such position responsive means, the system of the invention may be provided with electronic means as shown in FIG. For example, the proportional output and the integral output are added together in an adder circuit 50a that does not receive the position feedback signal from the conductor 54, and the result is provided on the conductor 60a as a position command signal. This signal is applied at summing point 120 to a constant voltage source 1, for example, which indicates the 100% defined limit for a given channel.
22 and the difference is provided on conductor 124 as a signal indicating margin to specified limits. This signal is provided to an indicator 126 and to a level detection circuit 128 in order to always indicate to the pilot the margin to the specified limit. The signal provided on lead 130 from this level detector indicates, for example, that 90% of the specified limit is currently being commanded on the axis in question. This signal is OR circuit 13
2 and with similar signals provided by leads 134 for the other axes, an alarm signal appears on leads 136 when the command in either axis reaches 90% of the specified limit. This signal gives a warning to the pilot by lighting the warning lamp 138, vibrating the control stick shaker 140, etc. The combination of the control column shaker, warning lamp and margin indicator replaces the position responsive means previously used to alert the pilot that a specified limit has been reached. In fact, the shaker causes the sensitive control stick to vibrate as a particular axis approaches a defined limit.
This is preferable to an alarm provided by a position responsive means upon reaching a specified limit.

The control system of the present invention, which is configured as shown in FIG. 2 by combining the operational force sensitive control device described in FIG. 1 with a signal processing circuit, has been used in a light helicopter with good results. In this embodiment, the signal matching circuit 2
4-27 have the characteristics previously explained in FIGS. 3-5. The gains of amplifiers 32-39 were adjusted to allow the control surface position to change to specified limits in 0.5 to 2 seconds when maximum force input is applied to the control column. For example, the constant Kc of the integrating amplifier 36 is
1.25, the minimum time for a full stroke of servo 74 in either direction when maximum force input is applied perpendicularly to sensitive control stick 10 and maximum voltage appears on conductor 20. It was heated in 1.5 seconds.
The constant Kp of the amplifier 37 is selected to be 0.5, and the servo 7
The minimum time for the entire stroke of 5 was about 2 seconds in either direction. The constant K _R of amplifier 38 was chosen to be 1.0, and the minimum time for a full stroke of servo 76 was about 1 second in either direction. Also, amplifier 3
The constant Ky of 9 was chosen to be 1.25, and the minimum time for a full stroke of servo 77 was about 0.8 seconds. These gains are relative to the gains of the corresponding proportional channels. However, each of these gains is adjusted in a manner well known to those skilled in the art in relation to the gain relationships provided by the signal matching circuits 24-27 and the characteristics provided in other parts of the system (e.g., servomechanism gain). .

The signal processing channel of the steering system described above is of the analog type, using amplifiers with appropriate gain, limit and integration characteristics and analog voltage summers to drive the servo valves. However, the present invention can also be effectively implemented as a digital system, that is, a system in which signal matching, integration, addition, etc. are all performed by one or two digital computers. An example is the dual computer system described in the aforementioned US patent application. When implementing the present invention in an aircraft using such a computer, the operating force sensitive control device 1
The voltage output of control stick 12 at 0 is applied to an analog-to-digital converter via a multiplexer as described in FIG. 1 of said U.S. patent application;
Further, the solenoid valves 70 to 73 are described in the first patent application of the above US patent application.
It is driven in the manner shown in FIG. It is clear that if two computers are used, both computers are connected to each of the steering axes.
On the other hand, it is of course possible to configure the system using one computer.

Signal processing can be performed by table lookup alone, as briefly suggested above, or by selection of coefficients by table lookup and calculation using the coefficients. All of the digital techniques necessary to perform the functions described in FIG. It may be similar to that in the control system.

The present invention is also suitable for combination with an autopilot system, such as an autopilot system that automatically controls altitude, speed, and heading, a stability augmentation system that compensates for disturbances to the aircraft attitude due to gusts, etc. There is. The coupling of the autopilot system to the pilot system according to the invention is quite simple. This is because the present invention has already implemented a method of transitioning from a floating correction point to an updated correction point, which is corrected by the autopilot system as a function of the gyro output, and by the stability enhancement system at a rate - This is because it only needs to be stabilized as a function of the gyro output. For example, the output signal of the autopilot system can be applied to the inputs of the integrating amplifiers 36-39, and the output signal of the stability multiplier can be applied to the inputs of the proportional amplifiers 32-35 or to the summing points 50-53. In some cases, the output signals of both systems may be added at summing points 50 to 5.
3 can be easily added. In either case, the electrical signals from the autopilot system and the stability augmentation device must be properly matched to take into account the differences between the device of the present invention and the device that will normally be used subsequently. For example, the stability increase signal should be kept small in magnitude, on the order of 5% of the specified limit, and the autopilot signal should be limited in its rate of change, although over the full range of the specified limit. . Also, when the stability increase signal is applied to the floating correction point of the present invention after the integration circuit, the center of the change width of the correction point for stability enhancement is set to It should be updated continuously by the signal. As above,
Known systems used for autopilot and stability enhancement can be combined with the steering system of the present invention.

Above, an example in which the present invention is mainly applied to a rotary wing aircraft (helicopter) has been described. However, the principles of the invention are equally applicable to fixed wing aircraft control systems. For fixed wing aircraft, the vertical axis controls the elevator, the horizontal axis controls the ailerons, and the torsion axis controls direction. The vertical axis is used to control speed and/or lift (ie, engine thrust or propeller pitch) as desired. Of course, the time constants and signal alignment characteristics are selected based on knowledge of servo control for fixed wing aircraft control surfaces. There are no other special considerations that must be taken into account when using the control system of the present invention on a fixed wing aircraft.

If desired, the function of mechanical mixer 84 may be replaced by an electrical signal mixing circuit in a "fly-by-wire" system incorporating the present invention. In this case, the mixed signal directly drives a main electromagnetic servo (not shown) in the switch plate 90, which is electromagnetic rather than mechanical. Furthermore, the four axes of the control stick of the operating force sensitive control device do not necessarily have to correspond one-to-one with specific servos.
Importantly, proportional/integral control of the aircraft's control surfaces is provided in response to force inputs to each axis of the control column, which has at least three axes.

Although the present invention has been illustrated and described with respect to its preferred embodiments, those skilled in the art will recognize that these and various other changes, omissions, and additions may be made without departing from the scope of the invention.

[Effects] As described above, the present invention allows the pilot to respond to changes in the aircraft's attitude, altitude, speed, nose direction, etc. by using a multi-axis sensitive flight control system and a signal processing circuit with proportional/integral characteristics. It becomes possible to control the aircraft by applying control input only when the position of the control surface is changed. This involves a completely new concept of aircraft control (floating correction point method).

According to the invention, a single control stick (e.g. a sidearm control stick) can be used for three or four control directions without requiring any cross-coupling between the control directions.
can be used. Unlike conventional control systems that operate based on the position of the control sticks, the control system of the present invention allows the pilot to take a comfortable posture and does not require large movements, thereby significantly reducing pilot fatigue. . Since the flight control system of the present invention continually updates correction points on each of the control axes, synchronization between the pilot's stick and the co-pilot's stick is unnecessary. According to the invention,
The conventional large control stick, pedals, etc. that made instruments and the outside world difficult to see and occupied a large space are omitted. The invention makes it possible to operate an aircraft with only one hand, without using the feet. Furthermore, the present invention allows for the implementation of highly capable aircraft control systems at a cost lower than that of conventional control stick and foot pedal systems. The steering system of the invention can be implemented using devices and techniques that are individually familiar to those skilled in the art.

[Brief explanation of drawings]

FIG. 1 is a perspective view of a side arm control device according to a preferred embodiment of the present invention. FIG. 2 is a block diagram of a helicopter control system using a side arm control device according to an embodiment of the present invention. FIG. 3 is a graph showing the input/output characteristics of one signal matching circuit used in the maneuvering system of FIG. 2. Fourth
This figure is a block diagram of a signal matching circuit for obtaining the characteristics shown in FIG. 3 in the system shown in FIG. 2. FIG. 5 is a graph showing the input/output characteristics of another signal matching circuit. FIG. 6 is a block diagram of a portion of the system shown in FIG. 2 that has been partially modified to indicate margins for specified limits. 10... Side arm control device, 12... Control stick, 13... Transducer assembly, 14...
Arm, 16... Pilot seat, 18... Swing axis,
24-27... Signal matching circuit, 32-35... Proportional amplifier, 36-39... Integrating amplifier, 50-5
3...Adder, 64-67...Amplifier, 70-7
6... Solenoid valve, 74-77... Fluid pressure servo, 84
...Mixer, 90...Swish plate,
92...Main rotor, 96...Pitch beam, 9
8... Tail rotor, 100-103... Position sensor, 108-111... Amplifier, 113... Working fluid source, 117... Integrating amplifier, 119... Wheel operation mechanism, 120... Adder, 122... Constant voltage source, 126...Indicator, 128...Level detection circuit, 132...OR circuit, 138...Warning lamp, 148...Control stick shaker.

Claims (1)

[Scope of Claims] 1. A control device for manually controlling at least one of attitude control, lift control, and speed control in pitch direction, roll direction, and yaw direction of an aircraft, the control device controlling a control surface for controlling one of the attitude, lift, and speed of the aircraft in pitch, roll, and yaw directions; a first operating direction for operating the control surface in a first control direction; The surface is operated in a second operating direction different from the first operating direction, and the surface is operated in a second operating direction different from the first operating direction, and the surface is operated with a minute operation stroke in each operating direction. generates an operation input signal indicating the magnitude and direction of the operation force in accordance with the force, continues outputting the operation input signal while the input of the operation force continues, and outputs the operation input signal when the operation force is released; a single manual input device that stops outputting the operation input signal; a servo mechanism for driving the control surface that operates the control surface in a first or second control direction according to a control value of the control signal; signal processing means for generating a control signal in accordance with the signal value of the operation input signal inputted by the manual operation input device, the signal processing means generating a control signal proportional to the signal value of the operation input signal; The control signal generates a proportional component of a control signal to be inputted, and an integral component of a control signal by integrating the signal value of the input signal, and commands a control position of the control surface according to the proportional component and the integral component. and supplying a control signal having the determined control value to the servo mechanism to control the position of the control surface according to the signal value and input duration of the operation input signal, and While there is no operational input signal, the integral component at the time when the operational input signal disappears is held as the floating correction point corresponding to the position of the control surface, and the control surface is held at the position at the time when the operational input signal disappears. , an aircraft control system characterized in that it operates as follows. 2. A control device for manually controlling at least two of pitch, roll, and yaw attitude control, lift control, and speed control of an aircraft, wherein the control device controls at least pitch and roll direction control. , first and second control surfaces that control at least two of the attitude, lift, and speed of the aircraft in the yaw direction; and inputting operational inputs for controlling the first and second control surfaces. a first control direction for controlling the first control surface, and a second control direction intersecting the first control direction for controlling the second control surface; a first operating direction for operating a corresponding first or second control surface in a first control direction in each of the first and second control directions; operating one or the second control surface in a second control direction opposite to the first control direction; Generates an operation input signal indicating the magnitude and direction of the operation force in accordance with the applied operation force by operating in each operation direction with a minute operation stroke that the pilot does not perceive the stroke, and inputs the operation force. a single manual input device that continues outputting the operation input signal while the operation force continues and stops outputting the operation input signal when the operation force is released; and a servo mechanism for driving a control surface that operates the wing surface in a first or second control direction in a first or second control direction in which the control force is input; While the operation input in the control direction continues, a control signal is sent to control the corresponding first or second control surface according to the signal value of the operation input signal in the same control direction input by the manual operation input device. and a signal processing means that integrates a proportional component of a control signal that is proportional to the signal value of the operation input signal, and a signal value of the input signal to integrate the control signal. determining a control value of the control signal that commands the control position of the control surface according to the proportional component and the integral component, and transmitting a control signal having the determined control value to the servo mechanism. control the position of the control surface according to the signal value and input duration of the control input signal, and calculate the integral component of the control surface at the point of disappearance of the control input signal while there is no control input signal. An aircraft control system characterized in that it operates as follows: holding the control surface as the floating correction point corresponding to the position and holding the control surface at the position at the time of disappearance of the control input signal.