MX2008003435A - Automatic velocity control system for aircraft - Google Patents

Abstract

A flight control system for an aircraft receives a selected value of a first parameter, which is either the airspeed or inertial velocity of the aircraft. A primary feedback loop generates a primary error signal that is proportional to the difference between the selected value and a measured value of the first parameter. A secondary feedback loop generates a secondary error signal that is proportional to the difference between the selected value of the first parameter and a measured value of a second flight parameter, which is the other of the airspeed and inertial velocity. The primary and secondary error signals are summed to produce a velocity error signal, and the velocity error signal and an integrated value of the primary error signal are summed to produce an actuator command signal. The actuator command signal is then used for operating aircraft devices to control the first parameter to minimize the primary error signal.

Description

AUTOMATIC SPEED CONTROL SYSTEM FOR AIRCRAFT TECHNICAL FIELD The present invention relates generally to the field of flight control systems for aircraft and particularly relates to a system for automatically controlling the speed of an aircraft.

DESCRIPTION OF THE PREVIOUS TECHNIQUE Many modern aircraft have flight control systems to maintain the selected flight parameters at or near the selected values. These parameters may include altitude, heading, position, and / or flight speed, and the control system maintains each parameter operating the flight control systems of the aircraft. For example, altitude can be controlled through the use of flight control surfaces, such as elevators, or through the use of the accelerator to control the flight speed of the aircraft. These flight control systems are normally cyclic program feedback control systems, which allow entry from the control system to respond to changes in the controlled parameter. Typical cyclic program systems control the speed of the aircraft using either flight speed or inertial velocity. The speed of flight is defined as the speed of advance of the aircraft with respect to the mass of air in which the aircraft is flying, while the inertial velocity is defined as the speed of advance of the aircraft with respect to the earth above. which the aircraft is flying. The flight control system compares the ordered speed (flight speed or inertial velocity) to the measured speed, and the difference between the ordered speed and the measured speed is the speed error. When the speed error is not zero, the control system enters a corrective command signal to one or more systems of the aircraft, such as accelerators in a fixed-wing aircraft or pitch of the rotor blades in a helicopter, to increase or decrease the measured speed in order to achieve a zero speed error. Typically, the corrective command signal is proportional to the speed error. A schematic view of a typical flight speed control system of the prior art is shown in the Figural. The system 11 comprises a control signal input device 13 for sending control signals to the actuators of the aircraft 15, and the flight speed of the aircraft is measured by a sensor 17 in the feedback circuit 19. The signal of The flight speed control of the device 13 and the negative of the flight speed output measured from the sensor 17 are added to the node 21, producing an error signal of the flight speed sent to the actuators 15. The system 11 operates the actuators 15 to reduce this error signal from the flight speed to zero. In quiet wind, typical cyclic program feedback systems operate to control flight speed quite well. However, an aircraft that flies in a turbulent wind environment will move from a mass of air moving in one direction to a mass of air moving in another direction. The effects of this turbulence will cause negative and positive longitudinal acceleration forces in the aircraft. These accelerations change the flight speed and the inertial velocity of the aircraft, which creates a speed error that the control system tries to eliminate. In a fixed-wing aircraft, the control system will send a change in throttle position, which changes the engine power and produces additional accelerations. In helicopters or other rotary wing aircraft, such as incunabula rotors, the control system may order a change in throttle position, the engine's nacelle position, and / or the pitching pitch of the blades, which it can also cause a change in the position of the aircraft's pitch. Changes in engine power and pitch position are transmitted to the cabin of the aircraft, producing undesirable acceleration and movement effects in passengers. An example will illustrate the effects of turbulent wind on the operation of a flight control system, such as system 11, which is commanded to maintain a selected flight speed. Figures 2A to 2E are graphs about the time of entry and response to a continuous front burst using the prior art system of Figure 1, and Figures 3A to 3E are similar graphs showing the input and response for a front burst. transient In an aircraft that flies through the wind that has no speed (calm wind), the control system measures little or no speed error, and accelerations caused by insignificant changes in the throttle input are not felt by the passengers However, when the aircraft encounters wind that is moving in the opposite direction to the aircraft, the flight speed sensor will detect the increased flight speed. For example, graph 2A shows the results of a continuous frontal burst of 30 ft / sec located at 5 seconds on the time scale and which rises to its maximum value in about 1 second. The burst causes the measured flight speed, shown in Figure 2B, to increase from the ordered flight speed from 200 knots per second (kts) to about 207kts in about 7.5 seconds. This also causes a decrease in velocity with respect to earth, as shown in Figure 2C. In response to the increased flight speed, the control system 11 commands a change in throttle position to reduce the power of the engine to reach the original flight speed. The position of the throttle against time is shown in Figure 2D, and the position is decreased by approximately 36 degrees just before the burst is about 12 degrees shortly after 8 seconds, reducing the power of the engine. The aircraft is in this way decelerated at a speed with respect to earth even slower, reaching a total decrease of the speed with respect to earth of 30 kts approximately in 14 seconds. After reaching the maximum at 207 kts, the flight speed begins to decrease due to the reduction in engine power, and the flight speed drops below 200 kts in approximately 1 1 seconds. Simultaneously, the position of the accelerator is raised to increase the power of the engine to achieve and maintain the orderly flight speed, but the control system 11 causes the throttle position to exceed that is not established until approximately 35 seconds. In addition to the longitudinal speeds, the vertical speed of the aircraft is affected, as shown in Figure 2E, with a maximum of + 8ft / sec and a minimum of -9ft / sec. When the aircraft moves back to a stationary air mass (zero wind speed), the measured flight speed will be less than the orderly flight speed. The control system then orders a change in throttle position to increase engine power, causing the aircraft to accelerate back to the original flight speed and to the original ground speed. Similar effects occur in the case of a transient frontal burst. Figures 3B to 3E show the results of a transient frontal burst of 30 feet / seconds that is found for 5 seconds, as shown in Figure 3A. As shown in Figure 3B, the burst causes the measured flight speed to rise to 210 kts in about 7 seconds as the speed relative to ground decreases, as shown in Figure 3C. In response to the increased flight speed, the control system 11 orders a change in throttle position to reduce the power of the engine in order to achieve the original flight speed. The position of the accelerator against time is shown in Figure 3D, and the position is decreased by approximately 36 degrees just before the burst is at approximately 22 degrees shortly after approximately 7 seconds, reducing the power of the engine. The aircraft is thus decelerated at an even lower ground speed, reaching a total decrease in ground speed of 23 kts in approximately 11 seconds. After reaching the maximum at 210 kts, the flight speed begins to decrease due to the reduction of the power in the engine, and the flight speed falls below the 200kts in approximately 9.5 seconds. Simultaneously, the position of the accelerator is raised to increase the power of the engine to achieve and maintain the orderly flight speed, but the control system 1 1 causes the position of the accelerator to exceed that not established for approximately 35 seconds. The longitudinal acceleration is represented graphically in Figure 3E, with an initial maximum deceleration of 8ft / sec / sec followed by a maximum acceleration of 7ft / sec / sec. The combination of positive and negative accelerations due to the behavior of the system 11 causes undesirable effects on the passengers of the aircraft. The initial deceleration caused by a continuous or transient burst is worsened by the accelerations due to a large undercorrection and overcorrection of the position of the accelerator.

BRIEF DESCRIPTION OF THE INVENTION There is a need for an automatic control system to control the flight speed of the aircraft that minimizes the undesirable accelerations encountered by passengers on the aircraft. Therefore, it is an object of the present invention to provide an automatic control system for controlling the speed of flight of the aircraft that minimizes the undesirable accelerations encountered by the passengers in the aircraft. A flight control system for an aircraft receives a selected value from a first parameter, which is either the flight or inertial velocity of the aircraft. A primary feedback circuit generates a primary error signal that is proportional to the difference between the selected value and a measured value of the first parameter. A secondary feedback circuit generates a secondary error signal that is proportional to the difference between the selected value of the first parameter and a measured value of a second flight parameter, which is the other of the flight speed and inertial velocity. The first and second error signals are summed to produce a speed error signal, and the speed error signal and an integrated value of the primary error signal are summed to produce an actuator command signal. The actuator control signal is then used to operate the aircraft devices to control the first parameter to minimize the primary error signal. The present invention provides several advantages, including; (1) the reduction of unwanted longitudinal acceleration caused by automatic responses to frontal gusts and wind turbulence; (2) reduction of automatic changes in engine power caused in response to wind turbulence; (3) increased stability for a flight control system, thus reducing overcorrections and undercorrections caused by turbulence and orderly changes; and (4) improving the efficiency of the aircraft by reducing the accelerations caused by wind turbulence.

BRIEF DESCRIPTION OF THE DRAWINGS For a more complete understanding of the present invention, including its features and advantages, reference is now now made to the detailed description of the invention taken in conjunction with the accompanying drawings in which similar numbers identify similar parts, and in which : Figure 1 is a schematic view of the components of a flight control system of the prior art; Figures 2A to 2E are graphs through the time of the input and response for a continuous front burst using the prior art system of Figure 1. Figures 3A to 3E are graphs over time of the input and response of a transient frontal burst using the prior art system of Figure 1; Figure 4 is a schematic view of the components of a preferred embodiment of a flight control system according to the current invention; Figures 5A to 5E are graphs over time of the input and response of a frontal burst supported using the system of Figure 4; Figures 6A to 6E are graphs over time of the input and response of a transient frontal burst using the system of Figure 4. Figure 7 is a perspective view of an aircraft comprising a flight control system of the Figure 4; Figure 8 is an alternative embodiment of the flight control system of the present invention.

DESCRIPTION OF THE PREFERENTIAL MODALITY The present invention is directed to a flight speed control system configured to automatically control the flight speed of an aircraft and reduce longitudinal accelerations due to the wind turbulence encountered during flight. When a gust of wind that has a longitudinal component is detected, the system of the invention uses the combination of a signal of the flight speed and an inertial velocity signal (longitudinal velocity to earth) as the velocity feedback signal for the control system. During a calm wind, the speed of flight and the steady state inertial velocity have the same value. Referring to the Figures, Figure 4 shows a schematic view of a preferred embodiment of the control system of the invention in which a selected flight speed is commanded by the operator or pilot. System 23 is a cyclic program feedback control system that uses both flight speed and inertial velocity (ground speed) to determine the appropriate response of the accelerator to changes in flight speed. In the system shown, a selected signal of the flight speed is produced from a device of command 25, which may be an on-board interface used by the pilot or a control system, such as an autopilot system. Alternatively, the control device 25 may be interconnected with a receiver that receives control signals transmitted from a distant site of the aircraft, such as an unmanned or remotely piloted vehicle. The flight speed control signal is added at node 27 with a signal output from the flight speed feedback circuit 29, which is the primary feedback loop. A flight speed sensor 31 is in data communication with the flight speed feedback circuit 29, to provide a signal representing the measured flight speed of the aircraft, and the negative value of the measured flight speed is added with the flight speed ordered at node 27 to calculate a signal of the error in the flight speed. In the same way, an inertial velocity, or a velocity with respect to earth, a feedback circuit 33 provides a signal representing a value of the inertial velocity measured by an inertial velocity sensor 35 that is in data communication with the feedback 33. In this mode, the inertial velocity feedback circuit 33 is the secondary feedback loop. The negative value of the inertial velocity measured by a sensor 35 is added to the ordered flight velocity at the node 37 to calculate the error of the inertial velocity. The flight speed error calculated at node 27 is used in two subsequent calculations. The error of the nervous speed (calculated at node 37) is added to the positive value of the flight speed error at node 39 to calculate a speed error. The integral value of the flight speed error is calculated using the integrator 41, and the positive value of this integral value is added to the positive value of the speed error in the node 43. The output signal of the node 43 represents the signal from control of the actuator used by the actuators or other devices represented by box 45 to control the flight speed of the aircraft so that the flight speed is reduced to a minimum. Using the combination of a flight velocity signal and an inertial velocity signal as the velocity feedback signal, the dynamic combination of these two signals will reduce the amplitude of the changes ordered by the system 23 caused by a turbulence of wind that only the flight speed sensor 31 was used. The sensors 31, 35 indicate the speed errors in opposite directions, but because the proportional error of the speed is computed from the combination of these two signals , the undesirable acceleration is significantly lower due to the cancellation effect of these two signals. However, the error of the low frequency speed, or steady state, used for the integral of the speed error is determined by the flight speed sensor 31 only, such that the constant flight speed is not affected by the signal of inertial velocity. The improved response can be observed in the Figures 5A through 5E and Figures 6A through 6E, which are graphs showing the input and improved response for frontal bursts at the same speed and duration as those plotted for the prior art control system 11 in Figures 2A to 2E and in Figures 3A to 3E, respectively. For example, the graph in Figure 5A shows that there is a frontal burst of 30ft / sec sustained at 5 seconds on the time scale and rises to its maximum value in about 1 second. The burst causes the measured flight speed, shown in Figure 5B, to rise from the orderly flight speed from 200 kts to about 207 kts in about 7.5 seconds. Figure 5C shows that the speed with respect to earth also decreases, as expected. In In response to the increased flight speed, the control system 23 orders a change in an actuator or other device to affect the flight speed. In this example, the position of the accelerator is used to control the power of the engine, and the position of the accelerator is initially reduced in order to achieve the original flight speed. However, the position of the accelerator, as shown in Figure 5D, is decreased by 36 degrees just before the burst is found at approximately 30 degrees shortly after in approximately 7 seconds. The position of the accelerator then gently rises to approximately 62 degrees while the speed of flight and the speed with respect to earth are established without problems at the new values. The system is set to approximately 15 seconds from the beginning of the burst. As shown in the graph of Figure 5E, a reduction is also carried out for accelerations and vertical movements. When compared with the responses of the prior art system 11, it should be noted that the graphs in FIGS. 5B to 5D lack undercorrection and overcorrection found in the response of the prior art system.

When the system carefully establishes new values without these oscillations, passengers increase comfort. The same improvements are also observed in response to a momentary wind gust, as shown in Figures 6A to 6E. A frontal burst of 30 feet / sec is in the tempo = 5 seconds, and the burst lasts approximately 5 seconds. Figure 6B shows the flight velocity peaks measured at 210kts for approximately 7 seconds and the subcorrections at approximately 194 kts to around 12 seconds. The speed with respect to ground, shown in Figure 5C, has a maximum decrease of about 15kts in about 10 seconds, but the speed with respect to ground is recovered after the burst without a overcorrection. Referring now to Figure 6D, the position of the accelerator changes from the initial setting of 36 degrees to approximately 26 degrees in response to the burst, then increases to about 60 degrees to increase the flight speed after the burst has finished. The throttle position is then adjusted back to approximately 36 degrees without an undercorrection. The system responds to the settings in approximately 15 seconds from the start of the burst. By comparing the response of the system of the present invention to the responses shown in FIGS. 3B to 3E for the prior art system, it should be noted that the system of the present invention reduces derivations to the maximum from the conditions of a prior art. burst without undercorrection and overcorrection observed in the responses of the prior art system. Also, the system quickly adjusts that the prior art system, and the longitudinal accelerations, plotted in Figure 6E last a short time. All this contributes to improve the comfort of the passengers in the aircraft. The devices of the aircraft used to control the flight speed can be of various types depending on the type of aircraft. For example, Figure 7 shows an aircraft with incunable rotor 47 has a flight speed control system in accordance with the present invention. The aircraft 47 has two rotors 49 that have multiple vanes 51, and each rotor 49 is rotated with a torque movement provided from a motor carried in an associated nacelle 53. Each nacelle 53 is pivotally mounted to an outer end of a wing 55 of the aircraft 47, allowing each pod 53 to rotate between a horizontal position, as shown in the figure, and a vertical position. Each motor has means (not shown) for controlling the power output and / or speed of the motor, and these means are collectively referred to herein as an "accelerator".

Although shown as an aircraft with incunable rotor, it should be understood that the flight speed control system 23 of the present invention is applicable to all types of aircraft, including fixed-wing aircraft and helicopters. In addition, although the engines of aircraft 47 are turbine engines, system 23 of the invention is also applicable to other types of aircraft engines, including reciprocating engines. Also, although the accelerators are mainly used to control the output of motors in an aircraft 47, the control system 23 can be used to control other devices to control the amount or direction of thrust produced by the rotors 49, For example, the system 23 can be used to control the rotating position of the nozzles 53 or the pitch of the blades 51. In other types of aircraft, the control system 23 can be used to control the flight speed through the use of the devices of thrust vector, such as those used to direct the exhaust of the turbine. Figure 8 is a schematic view of an alternative embodiment of the control system of the present invention. The control system 57 is configured to maintain an ordered inertial speed, or velocity with respect to ground, instead of maintaining an orderly flight speed, as was the system 23 of Figure 4 above. System 57 is a cyclic programming feedback control system that uses both a flight speed and an inertial velocity (velocity to ground) to determine the appropriate response of the accelerator to change at an inertial velocity. In the system shown, a selected inertial velocity signal is output from a control device 59, which may be in the on-board interface used by a pilot or a control system, such as an autopilot system. Alternatively, the control device 59 can be connected to a receiver that receives the commands transmitted from a remote location of the aircraft. The inertial velocity of the ordinate signal is added to the node 61 with a signal output from the inertial velocity feedback circuit 63, which is the primary feedback circuit in this mode. An inertial velocity sensor 65 is in data communication with the inertial velocity feedback circuit 63 to provide a signal representing the measured inertial velocity of an aircraft, and the negative value of the measured inertial velocity is added with the inertial velocity. commanded at node 61 to calculate an inertial velocity error signal. Similarly, a feedback loop of the flight speed 67, which is the secondary feedback circuit in this mode, provides a signal representing the value of the flight speed measured by a flight speed sensor 69. in data communication with the feedback circuit 67. The negative value of the flight speed measured by the sensor 69 is added with the inertial speed commanded in the node 71 to calculate the flight speed error. The error of the inertial velocity calculated in node 61 is used in two subsequent calculations. The flight velocity error (calculated at node 71) is added to the positive value of the inertial velocity error at node 73 to calculate a velocity error. The integral value of the inertial velocity error is calculated using an integrator 75, and the positive value of this integral value is summed with the positive value of the velocity error in the node 77. The output signal of the node 77 represents the signal of control of the actuator used by the actuators or other devices represented by the rectangle 79 to control the speed of navigation of the aircraft so that the error of the inertial speed is minimized. The combination of the flight speed signal and an inertial velocity signal as the feedback velocity signal reduces the amplitude of the changes commanded by the system 57 caused by an air turbulence. When a gust of wind is encountered, sensors 65, 69 detect changes in speed in opposite directions. The proportional speed error is computed using these two signals, so that the undesirable power or push that arises is significantly less due to cancellation effects. However, the low frequency, or steady-state, inertial velocity error used for the integral velocity error is determined by the inertial velocity sensor only, so that the fixed velocity is not affected by the velocity signal. Of flight. For example, an aircraft that uses an inertial speed control system may find wind moving in the opposite direction of the aircraft.

When this happens, the inertial speed sensor will detect a decrease in inertial velocity due to the increase in aerodynamic drag. The inertial speed control system is commanded to maintain a constant inertial speed, and the system will operate aircraft devices to obtain and maintain the original inertial speed. The present invention provides several advantages, including: (1) the reduction of an undesired longitudinal acceleration caused by automatic responses for frontal bursts and air turbulence; (2) reduction of automatic power changes of the engine caused in response to air turbulence; (3) increased stability for a flight control system, reducing overcorrections and subcorrections caused by turbulence and command changes; and (4) improvements in the efficiency of the aircraft by reducing the accelerations caused by air turbulence. While this invention has been described with reference to exemplary embodiments, this description is not intended to be construed as limiting.

Various modifications and combinations of the illustrative embodiments, as well as other embodiments of the invention, will be apparent to those skilled in the art above the reference of the description.

Claims (16)

1. A flight control system for an aircraft, the system comprises: means for receiving an input signal representing a selected value of a first parameter, the first parameter being one of the flight speed of the aircraft and the speed of inertia of the aircraft. The aircraft; a primary feedback circuit for generating a primary error signal, the primary error signal being proportional to the difference between the selected value of the first parameter and the measured value of the first parameter; and a secondary feedback circuit for generating a secondary error signal, the secondary error signal being proportional to the difference between the selected value of the first parameter and a measured value of the second flight parameter, the second parameter being the other one of the speed of flight of the aircraft and the inertial speed of the aircraft; characterized in that the primary error signal and the secondary error signal are summed to produce a speed error signal; wherein the speed error signal and an integrated value of the primary error signal are summed to produce an actuator control signal, and wherein the actuator control signal is adapted to be used to operate the aircraft devices to control the first parameter of the aircraft, in such a way that the primary error signal is minimized.

2. The control system according to claim 1, further characterized in that the means for receiving the input signal is configured to receive an input signal generated on board the aircraft.

3. The control system according to claim 1, further characterized in that the means for receiving the input signal is configured to receive an input signal generated remotely from the aircraft.

4. The control system according to claim 1, further characterized in that the first parameter is the flight speed of the aircraft and the second parameter is the inertial velocity of the aircraft.

5. The control system according to claim 1, further characterized in that the first parameter is the inertial velocity of the aircraft and the second parameter is the flight speed of the aircraft. The control system according to claim 1, further characterized in that the control signal of the actuator is adapted to be used to operate the devices selected from the group consisting of accelerators, controls of the rotor system, and controls of the position of the car. 7. An aircraft comprises: means of propulsion to propel the aircraft; at least one device configured to control a thrust outlet of the propulsion means; and a flight control system, comprising: means for receiving an output signal representing the selected value of a first parameter, the first parameter being one of the flight speed of the aircraft and the inertial velocity of the aircraft; a primary feedback circuit for generating a primary error signal, the primary error signal being proportional to the difference between the selected value of the first parameter and the measured value of the first parameter; Y a secondary feedback circuit for generating a secondary error signal, the secondary error signal being proportional to the difference between the selected value of the first parameter and the measured value of a second flight parameter, the second parameter being the other of the speed of flight of the aircraft and the inertial speed of the aircraft wing; characterized in that the primary error signal and the secondary error signal are summed to produce a velocity error signal; wherein the speed error signal and the integrated value of the primary error signal are summed to produce an actuator control signal, and wherein the control signal of the actuator is used to operate at least one control device of the first parameter of the aircraft, in such a way that the primary error signal is minimized. The aircraft according to claim 7, further characterized in that at least one device comprises at least one accelerator. 9. The aircraft according to claim 7, further characterized in that at least one device comprises at least one actuator to vector the thrust. 10. The aircraft according to claim 7, further characterized in that the means for receiving the input signal is configured to receive an input signal generated on board the aircraft. The aircraft according to claim 7, further characterized in that the means for receiving the input signal is configured to receive an input signal generated remotely from the aircraft. 12. The aircraft according to claim 7, further characterized in that the first parameter is the flight speed of the aircraft and the second parameter is the inertial velocity of the aircraft. 13. The aircraft according to claim 7, further characterized in that the first parameter is the inertial velocity of the aircraft and the second parameter is the flight speed of the aircraft. 14. A method for automatically controlling the flight of an aircraft, the method characterized in that it comprises: a) entering a signal representing a selected value of a first parameter, the first parameter being one of the flight speed of the aircraft and the speed inertial of the aircraft; b) generating a primary error signal by calculating the difference between the selected value of the first parameter and a measured value of the first parameter; c) generating a secondary error signal by calculating the difference between the selected value of the first parameter and a measured value of a second parameter, the second parameter being the other of the flight speed of the aircraft and the inertial velocity of the aircraft; d) generating a speed error signal by summing the primary error signal and the secondary error signal; e) generating an actuator control signal by summing the speed error signal and an integrated value of the primary error signal; and f) operating the devices on the aircraft to control the first parameter of the aircraft, such that the primary error signal is minimized. 15. The method according to claim 12, further characterized in that the first parameter is the flight speed of the aircraft and the second parameter is the inertial velocity of the aircraft. 1

6. The method according to claim 14, characterized also because the first parameter is the inertial velocity of the aircraft and the second parameter is the flight speed of the aircraft.