RU2380280C2 - Automatic speed control system for aircraft - Google Patents

Abstract

FIELD: physics; control.

SUBSTANCE: invention relates to automatic control systems. Aircraft, method and system for controlling flight of the aircraft assume a selected value of a first parametre which is either airspeed or inertial speed of the aircraft. A primary feedback loop generates a primary error signal which is proportional to difference between the selected value and the measured value of the first parametre. A secondary feedback loop generates a secondary error signal which is proportional to difference between the selected value of the first parametre and the measured value of the second flight parametre which is the remaining parametre: either airspeed or inertial speed. The primary and secondary error signals are summed to generate a speed error signal, and this speed error signal and the integrated value of the primary error signal are summed to generate an actuation control signal, which is then used to actuate aircraft devices to control the first parametre in order to minimise the primary error signal.

EFFECT: minimisation of undesirable accelerations acting on aircraft passengers.

16 cl, 24 dwg

Description

Technical field

The present invention generally relates to the field of flight control systems for an aircraft, and more particularly, to an aircraft speed automatic control system.

State of the art

Many modern aircraft have flight control systems to keep selected flight parameters accurate or close to selected values. These parameters may include flight altitude, heading, position in space and / or airspeed, and the control system supports each parameter through the operation of aircraft flight control systems. For example, altitude can be controlled by using flight control surfaces, such as elevators, or by using a throttle to control the airspeed of an aircraft. These flight control systems are usually closed loop control systems that allow the input from the control system to respond to changes in the value of the controlled parameter.

Typical closed loop control systems control aircraft speed using either airspeed or inertial speed. Air speed is defined as the forward speed of the aircraft relative to the air masses in which the aircraft is flying, while inertial speed is defined as the forward speed of the aircraft relative to the surface of the earth over which the aircraft is flying. The flight control system compares the target speed (airspeed or inertial speed) with the measured speed, and the difference between the target speed and the measured speed is considered a speed error. When the speed error is not zero, the control system enters an adjustment command to one or more aircraft systems, such as throttles for an aircraft with fixed wing geometry or the angle of inclination of the propeller blade in a helicopter, in order to increase or decrease the measured speed in order to achieve zero speed error. Usually the correction command is proportional to the speed error.

A schematic view of a known airspeed control system is shown in FIG. The system 11 comprises a command input device 13 for sending commands to the power drives 15 of the aircraft, and the airspeed of the aircraft is measured by a sensor 17 in the feedback loop 19. The airspeed command from the device 13 and the value opposite to the measured airspeed output value from the sensor 17 are summed in the node 21 producing the airspeed error signal, which is sent to the power drives 15. The system 11 controls the operation of the power drives 15 to reduce this air error signal speeds to zero.

In the absence of wind, typical closed loop closed loop systems work pretty well to control airspeed. However, an aircraft flying in an environment representing turbulent air will move from air masses moving in one direction to air masses moving in the other direction. The effects of this turbulence will cause positive and negative longitudinal accelerating forces on the aircraft. These accelerations alter the airspeed and inertial speed of the aircraft, which creates a speed error that the control system is trying to eliminate. In an aircraft with unchanged wing geometry, the control system will issue a command to change the throttle position, which changes the engine power and produces additional accelerations. In helicopters or other rotary-wing aircraft, such as convertiplanes, the control system can command the throttle position, the position of the engine nacelle and / or the input signals of the tilt of the blades, which can also cause a change in the tilt position of the aircraft. Changes in engine power and tilt position are transmitted to the aircraft cabin, producing undesirable effects of acceleration and movement on passengers.

An example illustrates the effects of turbulent air on the operation of a flight control system, such as system 11, which instructs to maintain a selected airspeed. 2A-2E are diagrams of the input signal and response versus time for a long oncoming gust of wind when using the prior art system shown in FIG. 1, and FIGS. 3A-3E are similar diagrams showing the input signal and response for short-term headwind.

In an aircraft flying through air masses that do not have speed (still air), the control system measures a small speed error or does not record it at all, while the accelerations caused by negligible changes in the throttle input signal are not felt by passengers. However, when the aircraft encounters air masses that are moving in the opposite direction with respect to the aircraft, the airspeed sensor will detect increased airspeed. For example, Diagram 2A shows the results of a long oncoming gust of wind at a speed of 9.1 m / s, seen at a time point of 5 s on a time line, which increases linearly to its maximum value over a period of approximately 1 s. A gust causes the measured air speed shown in FIG. 2B to increase from a given air speed of 370 km / h (200 knots) to about 383 km / h (207 knots) at a time of about 7.5 s. This is also the reason for the decrease in ground speed, as shown in figs. In response to increased airspeed, control system 11 issues a throttle position command to lower engine power in order to achieve initial airspeed. The position of the throttle as a function of time is shown in FIG. 2D, and this position decreases from about 36 degrees just before meeting with a gust of wind to about 12 degrees after it at a time of 8 s, lowering engine power. The aircraft, accordingly, slows down to even lower ground speed, achieving a total decrease in ground speed of 55.6 km / h (30 knots) at a point in time of about 14 s.

After reaching a maximum value of 383 km / h (207 knots), airspeed begins to decrease due to lower engine power and falls below 370 km / h (200 knots) at a time of about 11 s. At the same time, the throttle position increases to increase engine power to achieve and maintain a given airspeed, but the control system 11 causes the throttle position to go beyond the desired value and it will not stabilize until about 35 s. In addition to the longitudinal speeds, the vertical speed of the aircraft is affected, as shown in FIG. 2E, with a maximum value of +2.44 m / s (8 ft / s) and a minimum value of -2.74 m / s (9 ft / from).

When the aircraft returns to stationary air masses (zero wind speed), the measured airspeed will be less than the set airspeed. Then the control system issues a command to change the throttle position to increase engine power, leading the aircraft to accelerate back to the original airspeed and the original ground speed.

Similar effects occur in the case of short-term oncoming gusts of wind. FIGS. 3B-3E show the results of a headwind at a speed of 9.1 m / s (30 ft / s), which occurs for a period of 5 s, as shown in FIG. 3A. As shown in FIG. 3B, a gust of wind causes the measured airspeed to increase to 389 km / h (210 knots) at a time of about 7 s, while the ground speed decreases, as shown in FIG. 3C. In response to the increased airspeed, the control system 11 issues a throttle position command to lower engine power in order to achieve the original airspeed. 3D position of the throttle as a function of time is shown in FIG. 3D, and this position decreases from about 36 degrees just before meeting with a gust of wind, to about 22 degrees after it at a point in time of about 7 s, reducing engine power. The aircraft, accordingly, slows down to even lower ground speed, achieving a total decrease in ground speed of 42.6 km / h (23 knots) at a time point of about 11 s.

After reaching a maximum of 389 km / h (210 knots), airspeed begins to decrease due to lower engine power, and airspeed drops below 370 km / h (200 knots) at a time of about 9.5 s. At the same time, the throttle position increases to increase engine power to achieve and maintain a given airspeed, but the control system 11 causes the throttle position to go beyond the required value, and this throttle position does not stabilize until about 35 s. Longitudinal acceleration is shown in the form of a diagram in FIG. 3E with an initial maximum deceleration of 2.4 m / s ² (8 ft / s ² ) and a subsequent maximum acceleration of 2.1 m / s ² (7 ft / s ² ).

The combination of positive and negative accelerations due to the behavior of the system 11 leads to undesirable effects on the passengers of the aircraft. The initial deceleration caused by a prolonged or short-term gust of wind is worsened due to accelerations arising from a large shortfall in the throttle position to the required regulation value or transition through it.

SUMMARY OF THE INVENTION

There is a need for an automatic control system for controlling aircraft airspeed that minimizes unwanted accelerations affecting aircraft passengers.

Consequently, an object of the present invention is to provide an automatic control system for controlling the airspeed of an aircraft, which minimizes undesired accelerations affecting aircraft passengers.

The flight control system for the aircraft takes the selected value of the first parameter, which is either the airspeed or the inertial speed of the aircraft. The primary feedback loop generates a primary error signal that is proportional to the difference between the selected value and the measured value of the first parameter. The secondary feedback loop generates a secondary error signal that is proportional to the difference between the selected value of the first parameter and the measured value of the second flight parameter, which is either airspeed or inertial speed, but different from the first parameter. The primary and secondary error signals are summed to produce a speed error signal, and the speed error signal and the integrated value of the primary error signal are summed to produce a power drive control signal. The power drive control signal is then used for aircraft operating devices to control the first parameter in order to minimize the primary error signal.

The present invention provides several advantages, including:

1) reduction of undesirable longitudinal acceleration caused by automatic responses to oncoming wind gusts and air turbulence;

2) a decrease in automatic changes in engine power caused as a response to air turbulence;

3) increasing the stability of the flight control system, thereby reducing the transition beyond the required regulation value and the lack of access to it, caused by turbulence and predetermined changes;

4) improving the efficiency of the aircraft by reducing the acceleration caused by air turbulence.

Brief Description of the Drawings

For a more complete understanding of the present invention, including its characteristic features and advantages, we turn to the detailed description of the invention, taken in combination with the accompanying drawings, in which the same parts are indicated by the same reference numbers, and where

figure 1 shows a schematic view of the components of a flight control system of the prior art;

on figa-2E shows a diagram of the dependence of input and response from time to time oncoming gust of wind when using the prior art system shown in figure 1;

on figa-3E shows a diagram of the dependence of input and response on time for a short oncoming gust of wind when using the prior art system shown in figure 1;

4 is a schematic view of components of a preferred embodiment of a flight control system in accordance with the present invention;

on figa-5E shows a diagram of the dependence of input and response from time to time oncoming gust of wind when using the system depicted in figure 4;

on figa-6E shows a diagram of the dependence of input and response on time for a short oncoming gust of wind when using the system depicted in figure 4;

FIG. 7 is a perspective view of an aircraft comprising the flight control system shown in FIG. 4;

on Fig shows an alternative embodiment of a flight control system in accordance with the present invention.

Description of the preferred embodiment

The present invention is directed to an airspeed control system configured to automatically control the airspeed of an aircraft and reduce longitudinal accelerations due to air turbulence encountered during flight. When a gust of wind having a longitudinal component is detected, the system according to the invention uses a combination of an airspeed signal and an inertial speed signal (longitudinal component of ground speed) as a speed feedback signal for the control system. With still air, steady-state airspeed and inertial speed are one and the same value.

Turning to the drawings, we note that Fig. 4 shows a schematic view of a preferred embodiment of a control system in accordance with the invention in which the selected airspeed is set by the operator or pilot. System 23 is a closed loop feedback control system that simultaneously uses airspeed and inertial speed (ground speed) to determine an adequate throttle response to changes in airspeed. In the system shown, the selected airspeed signal is the output from the command device 25, which may be an onboard docking device used by a pilot or control system such as an autopilot. Alternatively, the command device 25 may dock with a receiver that receives commands transmitted from a location remote from the aircraft, such as an unmanned or remotely piloted vehicle. The airspeed control signal is added at node 27 to the output from the airspeed feedback loop 29, which is the primary feedback loop. The airspeed sensor 31 transmits data to the airspeed feedback loop 29 to obtain a signal representing the measured airspeed of the aircraft, and a value opposite to the measured airspeed is added to the set airspeed in node 27 to calculate the airspeed error signal. Similarly, the inertial speed or ground speed feedback loop 33 provides a signal representing the value of the inertial speed measured by the inertial speed sensor 35 transmitting data to the feedback loop 33. In this embodiment, the inertial speed feedback loop 33 is a secondary feedback loop. The value opposite to the inertial speed measured by the sensor 35 is added to the set airspeed at node 37 to calculate the inertial speed error.

The airspeed error calculated at node 27 is used in two subsequent calculations. The inertial velocity error (calculated in node 37) is added to the positive value of the airspeed error in node 39 to calculate the velocity error. The integral value of the speed error is calculated using the integrator 41 and the positive value of this integral value is summed with the positive value of the speed error in the node 43. The output signal from the node 43 is the drive control signal used by the power drives or other devices represented by a rectangle 45, to adjust the airspeed of the aircraft so as to minimize airspeed.

By using a combination of the airspeed signal and the inertial speed signal as the speed feedback signal, a dynamic combination of these two signals will reduce the amplitude of the changes issued as a command by the system 23 caused by air turbulence when only the airspeed sensor 31 is used. The sensors 31, 35 show speed errors in opposite directions, but since the proportional speed error is calculated from the combination of these two signals, the unwanted acceleration is much less due to the effect of mutual compensation of these two signals. However, the low frequency or stationary speed error used for the integral value of the speed error is detected only by the airspeed sensor 31, so that the stationary airspeed is not affected by the inertial speed signal. An improved response can be seen in FIGS. 5A-5E and FIGS. 6A-6E, which show input and improved response diagrams for headwind gusts of the same speed and duration as those shown for the prior art system 11 2A-2E and FIGS. 3A-3E, respectively.

For example, the diagram in FIG. 5A shows that a long oncoming gust of wind at a speed of 9.1 m / s (30 ft / s) occurs at a point in time of 5 s on the time line and rapidly increases linearly to its maximum value in about 1 from. A gust of wind causes the measured air speed shown in FIG. 5B to increase from a given air speed of 370 km / h (200 knots) to about 383 km / h (207 knots) at a time of about 7.5 s. On figs shows that the ground speed is also reduced, as expected. In response to increased airspeed, control system 23 issues a change command to a power drive or other device in order to affect airspeed. In this example, the throttle position is used to control engine power, with the throttle position first being lowered in order to reach the original airspeed. However, the throttle position, as shown in FIG. 5P, decreases from about 36 degrees just before meeting with a gust of wind to about 30 degrees after a gust of wind at a time of about 7 s. The throttle position then increases uniformly up to approximately 62 degrees, while airspeed and ground speed uniformly stabilize at new values. The system stabilizes in about 15 seconds from the start of a gust of wind. As shown in the diagram in FIG. 5E, the vertical accelerations and motions are also reduced.

When comparing with the responses of the prior art system 11, it should be noted that the diagrams in FIGS. 5B-5D have no shortfall to the required control values and transitions through them found in the prior art system. When the system gently stabilizes to new values without these fluctuations, passenger comfort increases.

The same improvements are also visible in the responses for a brief gust of wind, as seen in FIGS. 6A-6E. A headwind at a speed of 9.1 m / s (30 ft / s) occurs at a time of 5 s and lasts for 5 s. FIG. 6B shows that the measured airspeed reaches a maximum of 210 knots (389 km / h) at a time of 7 s and does not reach the desired control value at a value of about 194 knots (359 km / h) at a time of about 12 s . The ground speed shown in FIG. 6C has a maximum decrease of about 27.8 km / h (15 knots) at a time of about 10 s, but the ground speed is restored after a gust of wind without crossing the desired control value. Referring to FIG. 6D, note that the throttle position changes from the initial setting of 36 degrees to about 26 degrees in response to a gust of wind, and then increases to a value close to 60 degrees to increase air speed after the gust of wind has ended . The position of the throttle is then stabilized in the opposite direction to a value of approximately 36 degrees without a loss to the desired control value. The response of the system stabilizes at approximately 15 s from the start of a gust of wind.

Comparing the response of the system in accordance with the present invention with the responses shown in FIGS. 3B-3E for the prior art system, it should be noted that the proposed system lowers the maximum deviations from the conditions that existed before the gust of wind, without income, to the desired regulation and transition through him observed in the responses of the prior art system. In addition, the system stabilizes rather than the prior art system, with the longitudinal accelerations shown in the diagram in FIG. 6E lasting a shorter time. All this contributes to improving the comfort of travel of aircraft passengers.

Aircraft devices used to control airspeed can be of various types, depending on the type of aircraft. For example, FIG. 7 shows a tiltrotor 47 with an airspeed control system in accordance with the present invention, having two rotary screws 49 with multiple blades 51, each screw 49 rotating with a torque provided by an engine located in a corresponding nacelle 53. Each the nacelle 53 is rotatably attached to the outer edge of the wing 55 of the aircraft 47, which allows each nacelle to rotate between a horizontal position, as shown in the drawing, and a vertical position. Each engine has means (not shown) for controlling the output power and / or speed of the engine, and these means in this document are collectively referred to as a “throttle”.

While a tiltrotor is shown, it should be understood that the proposed airspeed control system 23 can be applied to all types of aircraft, including aircraft with fixed wing geometry and helicopters. In addition, although the engines of the aircraft 47 are turbine engines, the proposed system 23 can also be applied to other types of aircraft engines, including piston engines. Also, although throttles are mainly used to control the output of engines on an aircraft 47, the control system 23 can be used to control other devices to control the magnitude or direction of thrust produced by the screws 49. For example, the control system 23 can be used to control the angular position of the nacelles 53 or the inclination of the blades 51. In other types of aircraft, the control system 23 can be used to control air speed by using the device shaft, specifying the direction of thrust, such as those used to direct the turbine exhaust.

On Fig shows a schematic view of an alternative embodiment of the proposed control system 23. The control system 57 is designed to maintain a given inertial speed, or ground speed, and not to maintain a given airspeed, as the system 23 in figure 4, shown above.

System 57 is a closed loop feedback control system that uses both airspeed and inertial speed (ground speed) to determine the corresponding throttle response to inertial speed changes. In the system shown, the selected inertial speed signal is the output from the command device 59, which may be an onboard docking device used by the pilot or control system, such as an autopilot system. Alternatively, the command device 59 may be connected to a receiver that receives commands transmitted from a location remote from the aircraft. The predetermined inertial speed control signal is summed in the node 61 with the output signal from the inertial speed feedback loop 63, which in this embodiment is the primary feedback loop. The inertial speed sensor 65 communicates with the inertial speed feedback loop 63 to obtain a signal displaying the measured inertial speed of the aircraft, and the value opposite to the measured inertial speed is added to the given inertial speed in node 61 to calculate the inertial speed error signal. Similarly, the airspeed feedback loop 67, which in this embodiment is the secondary feedback loop, provides a signal representing the airspeed value measured by the airspeed sensor 69, which is communicating with the feedback loop 67. The value opposite to the airspeed measured by the sensor 69 is summed with a predetermined inertial speed in the node 71 to calculate the airspeed error.

The inertial velocity error calculated at node 61 is used in two subsequent calculations. The airspeed error (calculated in node 71) is added to the positive value of the inertial velocity error in node 73 to calculate the speed error. The integral value of the inertial speed error is calculated using the integrator 75, and the positive value of this integral value is summed with the positive value of the speed error in the node 77. The output signal from the node 77 is the drive control signal used by the power drives or other devices represented by a rectangle 79, to control the airspeed of the aircraft so that the inertial speed error is minimized.

The combination of the airspeed signal and the inertial speed signal as the speed feedback signal reduces the amplitude of the changes defined by system 57 caused by air turbulence. When a gust of wind occurs, sensors 65, 69 detect changes in speed in opposite directions. The proportional error of the speed is calculated using these two signals, so that unwanted power or a sharp rise in traction is much less due to the effect of mutual compensation. However, the low-frequency, or stationary, inertial velocity error used for the integral value of the velocity error is determined only by the inertial velocity sensor, so that the steady-state speed is not affected by the airspeed signal.

For example, an aircraft using an inertial speed control system may encounter a stream of air that moves in the opposite direction with respect to the direction of movement of the aircraft. When this happens, the inertial speed sensor will detect a decrease in inertial speed due to increased aerodynamic drag. The inertial speed control system issues a command to maintain a constant inertial speed, and the system will control the devices on the aircraft so as to achieve and maintain the initial inertial speed.

The present invention provides several advantages, among which

1) reduction of undesirable longitudinal acceleration caused by automatic responses to oncoming wind gusts and air turbulence;

2) a decrease in automatic changes in engine power caused by a response to air turbulence;

3) an increase in the stability of the flight control system and, as a consequence, a decrease in transitions beyond the required regulation value and non-profit to it, caused by turbulence and predetermined changes;

4) improving the efficiency of the aircraft by reducing the acceleration caused by air turbulence.

While the proposed invention is described with reference to illustrative embodiments, this description is not intended to be construed in a limiting sense. Various modifications and combinations of illustrative embodiments, as well as other embodiments of the invention, will be apparent to those skilled in the art when referring to this description.

Claims (16)

1. The aircraft flight control system containing means for receiving an input signal representing a selected value of the first parameter, which is either the airspeed of the aircraft or the inertial speed of the aircraft, a primary feedback loop designed to generate a primary error signal that is proportional to the difference between the selected value of the first parameter and the measured value of the first parameter, a secondary feedback loop for generating a second the initial error signal, which is proportional to the difference between the selected value of the first parameter and the measured value of the second flight parameter, which is either the airspeed of the aircraft or the inertial speed of the aircraft, but different from the first parameter, while in this system the primary error signal and the secondary signal errors are summed to produce a speed error signal, the speed error signal and the integrated value of the primary error signal are summed to produce a control signal The pressure actuator and actuator control signal is used to activate the operating aircraft devices to control the first parameter of the aircraft so as to minimize the primary error signal. 2. The control system according to claim 1, in which the means for receiving the input signal are configured to receive an input signal generated directly on board the aircraft. 3. The control system according to claim 1, in which the means for receiving the input signal are configured to receive an input signal generated remotely with respect to the aircraft. 4. The control system according to claim 1, in which the first parameter is the airspeed of the aircraft, and the second parameter is the inertial speed of the aircraft. 5. The control system according to claim 1, in which the first parameter is the inertial speed of the aircraft, and the second parameter is the air speed of the aircraft. 6. The control system according to claim 1, in which the control signal of the power drive is made for use by working devices selected from the group consisting of throttles, controls for the screw system, as well as controls for the position of the nacelle. 7. An aircraft containing power devices for moving the aircraft, at least one device for controlling the thrust output of power devices, a flight control system comprising means for receiving an input signal representing a selected value of a first parameter, which is either the air speed of the air the vessel, or the inertial speed of the aircraft, the primary feedback loop, designed to generate a primary error signal, which is proportionally different between the selected value of the first parameter and the measured value of the first parameter, a secondary feedback loop designed to generate a secondary error signal that is proportional to the difference between the selected value of the first parameter and the measured value of the second flight parameter, the second parameter being either the airspeed of the aircraft or the inertial speed of the aircraft, but different from the first parameter, while the primary error signal and the secondary error signal are summed for the production of the speed error signal, the speed error signal and the integrated value of the primary error signal are summed to produce the power drive control signal and the power drive control signal is used to actuate the aircraft operating devices to control the first parameter of the aircraft so as to minimize the primary error signal . 8. The aircraft according to claim 7, in which the specified at least one device contains at least one throttle. 9. The aircraft according to claim 7, in which the specified at least one device contains at least one power drive to specify the direction of thrust. 10. The aircraft according to claim 7, in which the means for receiving the input signal are configured to receive an input signal generated directly on board the aircraft. 11. The aircraft according to claim 7, in which the means for receiving an input signal is configured to receive an input signal generated remotely with respect to the aircraft. 12. The aircraft according to claim 7, in which the first parameter is the airspeed of the aircraft, and the second parameter is the inertial speed of the aircraft. 13. The aircraft according to claim 7, in which the first parameter is the inertial speed of the aircraft, and the second parameter is the air speed of the aircraft. 14. A method of automatically controlling the flight of an aircraft, including a) introducing a signal representing a selected value of the first parameter, which is either the airspeed of the aircraft or the inertial speed of the aircraft, b) generating a primary error signal by calculating the difference between the selected value of the first parameter and the measured value of the first parameter, c) the generation of the secondary error signal by calculating the difference between the selected value of the first parameter and the measured the value of the second parameter, the second parameter being either the airspeed of the aircraft or the inertial speed of the aircraft, but different from the first parameter, d) generating a speed error signal by summing the primary error signal and the secondary error signal, e) generating the power drive control signal by summing the speed error signal and the integrated value of the primary error signal and then f) actuating the aircraft devices to control the first air a stuffy vessel so as to minimize the initial signal of error. 15. The method according to 14, in which the first parameter is the airspeed of the aircraft, and the second parameter is the inertial speed of the aircraft. 16. The method of claim 14, wherein the first parameter is the inertial speed of the aircraft, and the second parameter is the airspeed of the aircraft.

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