Aircraft automatic control system, effectiveness evaluation method and measuring device for the same

Abstract

To provide an automatic control technique capable of reducing oscillations of an aircraft body when an aircraft encounters turbulence, even in the presence of a minor measurement error in forecast information.SOLUTION: An aircraft automatic control system comprises: a measurement section which measures a difference between a wind speed that an aircraft currently experiences and a wind speed that the same will experience in the future as forecast information and outputs measurement information with an estimated error in a measurement value or information to determine validity of the measurement information added to the measurement information; a ladder to control lift force, drag force or posture of the aircraft or a device to control propulsion force of the aircraft; and a control calculation section which calculates an angle of the ladder or the propulsion force to mitigate an influence of the wind speed on the aircraft body based on a wind speed value in a planned flight direction of the aircraft measured with the measurement section.SELECTED DRAWING: Figure 1

Description

[Technical field]

[0001] The present invention relates to technology relating to an automatic control system for an aircraft, and more particularly to a gust response alleviation system and method used to reduce the vibration of an aircraft when the aircraft flies through turbulence, for example, or a gust load alleviation system and method used to reduce the load applied to the aircraft. [Background technology]

[0002] Turbulence is a particularly important cause of passenger aircraft accidents, and technology related to Doppler LIDAR using laser light has been researched and developed as a device to be installed on aircraft to detect turbulence in advance (see, for example, Non-Patent Document 1). Since then, as research and development has progressed, the performance limits of Doppler LIDAR have become clear, and the present inventors have proposed a method of use that is thought to be highly effective (see, for example, Non-Patent Document 2).

[0003] To use Doppler LIDAR to prevent aircraft turbulence accidents, there are two methods: one is to transmit turbulence information ahead of the flight direction to the pilot, who can then take measures such as evasive maneuvering or turning on the seatbelt sign; the other is to transmit the airflow information to an onboard computer and automatically control the rudder to reduce the shaking of the aircraft when it enters turbulence (see, for example, Patent Document 1).

[0004] In order to control the rudder, it is generally necessary to obtain a vertical airflow vector. In Patent Document 2, the present inventors proposed a technology to obtain a vertical airflow vector by geometrically converting the measured values ​​by two sets of Doppler LIDARs (remote airflow measurement devices).

[0005] Furthermore, in Patent Document 3, the inventors proposed a remote airflow measurement device, a remote airflow measurement method, and a program that can improve the estimation accuracy of secondary airflow vectors, including vertical airflow vectors, and also widen the airflow estimation range.

[0006] However, when automatically controlling the rudder using an airflow vector as preview information for vibration reduction control, the preview information must be extremely reliable, and there is a possibility that vibration may increase if automatic control is performed based on erroneous information. Therefore, the present inventors proposed a technology for adding reliability information to observation signals in Patent Document 4. Then, how to effectively utilize the observation information was an issue.

[0007] Meanwhile, the present inventors have proposed a technology for simply reducing the vertical oscillation of an airplane in Patent Document 5. However, the effect of this technology is considered to be limited, and the target for reduction is only the vertical oscillation of the airplane. [Prior art documents] [Patent documents]

[0008] [Patent Document 1] Patent No. 5771893 [Patent Document 2] Patent No. 5398001 [Patent Document 3] Patent No. 6583677 [Patent Document 4] Patent application 2017-234165 [Patent Document 5] Patent application 2018-72705 [Non-patent literature]

[0009] [Non-Patent Document 1] H.Inokuchi, H.Tanaka, and T.Ando, ​​"Development of an Onboard Doppler LIDAR for Flight Safety," Journal of Aircraft,Vo1.46, No.4, PP.1411-1415, AIAA, July-August, 2009. [Non-Patent Document 2] H.Inokuchi, T.Akiyama, "Performance Evaluation of an Airborne Coherent Doppler Lidar and Investigation of its Practical Application" Transactions of JSASS, Vol.65, No.2, PP.47-55, March 2022. [Non-Patent Document 3] Hamada, "Discrete-Time Preview Feedforward Compensation and Its Application to Gust Response Mitigation Control", 57th Joint Conference on Automatic Control, 2014. [Non-Patent Document 4] Y. Hamada, "New lmi-based conditions for preview feedforward synthesis," Control Engineering Practice, Vol.90, PP.19-26, 2019. Summary of the Invention [Problem to be solved by the invention]

[0010] In conventional feedback control to reduce aircraft sway, the rudder angle is usually controlled based on the output of an acceleration sensor attached to the aircraft. In this case, there is a delay due to the inertial force of the aircraft from when it encounters a gust of wind until the aircraft starts to move. Furthermore, the aircraft's movement is measured by an acceleration sensor, and the appropriate rudder angle is calculated, and then a rudder angle command is sent to the rudder actuator, but there is a delay even before the aerodynamic force of the rudder is changed. Therefore, there is a possibility that the initial movement of the sway cannot be responded to, or that it may actually cause vibration.

[0011] For example, when airflow information is used for automatic control of control surfaces as described above, inappropriate control of control surfaces due to erroneous signals or observation errors is not acceptable for the safety of aircraft operation. Nevertheless, in the past, there was a possibility that inappropriate control would occur due to the influence of erroneous signals that occur rarely. In other words, in the past, it was not possible to utilize reliability information on the control input signal, and therefore, depending on the conditions, safety could be reduced compared to the case where no control was performed. Since erroneous signals are generated by noise that is always present, they cannot be completely eliminated. In addition, there was a drawback in that the accuracy of control would decrease if data with large observation errors were mixed in.

[0012] To estimate the vertical and front-rear airflow of a gust using a Doppler LIDAR, optical axes with two or more lines of sight are required. In the case of the Doppler LIDAR prototyped so far, the measurement accuracy of the optical axis line of sight direction is 0.2 to 0.3 m / s, based on the results of Monte Carlo simulations and flight tests. When this is converted to a vertical airflow vector, the estimation accuracy is 0.6 to 0.9 m / s when the angle between the optical axes is 20 degrees.

[0013] As long as the above accuracy is always guaranteed, flight simulations have shown that the vibration of the aircraft can be appropriately reduced by half, but in actual measurements, the measurement error may temporarily increase due to the influence of noise, or the signal may become a singular value due to inability to measure properly. If such low-quality measurement information is used as is for control, the vibration of the aircraft may become greater than if no control was performed.

[0014] In addition, automatic control of control surfaces requires a high sampling rate, but when measuring the optical axis directions of two or more lines of sight, each local flow must be measured, and if the measurement time is short, it becomes difficult to estimate the overall flow due to the effects of minute turbulence.

[0015] In view of the above circumstances, an object of the present invention is to provide an automatic control technique that can reduce the shaking of an aircraft when it enters turbulence, or can reduce the load applied to the aircraft, even if there is a slight measurement error in the preview information. [Means for solving the problem]

[0016] An automatic control system for an aircraft according to one aspect of the present invention comprises: A measurement unit that measures the difference between the wind speed that the aircraft is currently experiencing and the wind speed that the aircraft will encounter in the future as prediction information, and adds information that determines the estimated error or validity of the measurement value to the measurement information and outputs the result; A device for controlling the rudder or thrust that controls the lift, drag, or attitude of an aircraft; a control and calculation unit that calculates the rudder angle or thrust so as to reduce the effect of the wind speed on the aircraft, based on the wind speed value in the planned flight direction of the aircraft measured by the measurement unit; Equipped with:

[0017] The above-described automatic control system can reduce the adverse effects of measurement errors in the preview information by selecting and utilizing only high-quality preview information to automatically control the aircraft.

[0018] The measurement unit may be configured to emit electromagnetic waves in the planned flight direction of the aircraft, receive the scattered waves in the atmosphere, and measure the remote wind speed in the radiation axis direction of the electromagnetic waves based on the Doppler shift amount of the scattered electromagnetic waves relative to the radiated electromagnetic waves.

[0019] The measurement unit may obtain a two-dimensional or three-dimensional vector of wind speed by setting or scanning the radiation axis of the electromagnetic wave from two or more lines of sight, thereby making it possible to add a difference between the respective measurement values ​​obtained from the radiation axis of the electromagnetic wave from two or more lines of sight to the measurement information as information for determining whether the measurement information is valid or invalid.

[0020] The control and calculation unit may be configured to calculate the wind speed value based on a moving average value of the spectrum integral of the received signal from which the received signal with invalid information added has been removed. By moving the time range for the spectrum integration over time, it is possible to use high data rate measurements as input for automatic control.

[0021] The control calculation unit may be configured to execute automatic control of the rudder based on the output of an acceleration sensor that detects acceleration acting on the aircraft when the estimation error is equal to or greater than a set value, or when preview information to which invalid information has been added is received.

[0022] The control calculation unit may be configured to divide the estimation error by a set value, subtract the result from 1, define the result as an authority, and multiply the control command by either 0 or the authority, whichever is larger.

[0023] The control calculation unit may be configured to generate the control signal by regarding, as the wind speed value, a value obtained by subtracting the estimation error from the measured value or 0, whichever is larger.

[0024] The control calculation unit may be configured to calculate the rudder angle using measurement values ​​in range bins before and after a measurement error in a certain range bin is larger than a predetermined absolute value or invalid information is added to the measurement value.

[0025] The control calculation unit may be configured to use a control gain that offsets the influence of a bias-like measurement error that is a constant value added to the measurement value of each range bin.

[0026] The method for evaluating the effectiveness of an automatic control system according to one aspect of the present invention includes the steps of: Conducting flight tests or flight simulations of the aircraft under predetermined wind and flow conditions, including vertical wind speeds; Plotting the results of the flight test or flight simulation with the amount of change in acceleration of the aircraft when the automatic control based on the control command is not performed as the horizontal axis and the amount of change in acceleration of the aircraft when the automatic control is performed as the vertical axis; The downward deviation of the plot from a preset broken line is regarded as the effect of the automatic control in reducing the gust response.

[0027] A measurement device according to one embodiment of the present invention is a measurement device for an automatic control system of an aircraft, which measures a difference between a wind speed currently being experienced by an aircraft and a wind speed that the aircraft will encounter in the future as prediction information, and The device is provided with a signal processing unit which adds the difference between the respective measurement values ​​obtained from the electromagnetic wave radiation axes of two or more lines of sight to the measurement information as information for determining whether the measurement is valid or invalid, and outputs the result. Effect of the Invention

[0028] According to the present invention, in automatic control for reducing the vibration and load of an aircraft when the aircraft enters turbulence, it is possible to reduce the adverse effects of measurement errors in preview information. [Brief description of the drawings]

[0029] [Figure 1] 1 is a block diagram showing a configuration of an automatic control system for an aircraft according to an embodiment of the present invention. FIG. [Diagram 2] FIG. 2 is a conceptual diagram showing a method for measuring the airflow ahead of the aircraft in the above-mentioned automatic control system. [Diagram 3] 4 is a flowchart showing an operation of a control calculation unit in the automatic control system. [Figure 4] 5 is an explanatory diagram showing a method of calculating a moving average value executed by the control calculation unit. FIG. [Diagram 5] FIG. 4 is an explanatory diagram showing a method of using an estimated measurement error in the above automatic control system. [Figure 6] FIG. 11 is an explanatory diagram showing a method of estimating the measurement value of a range bin from front and rear measurement points in the automatic control system to calculate the rudder angle. [Figure 7] FIG. 2 is an explanatory diagram showing a configuration of a control input in the automatic control system. [Figure 8] FIG. 2 is an explanatory diagram showing a method for reducing fluctuations in roll angle by observing four lines of sight in the automatic control system. [Figure 9] FIG. 13 is a diagram showing the distribution of vertical wind speed used in a flight simulation for evaluating the effectiveness of the above-mentioned automatic control system. [Figure 10]FIG. 11 is a diagram showing the measured and true wind speeds ahead of the aircraft at a certain time in an example flight simulation for evaluating the effectiveness of the above-mentioned automatic control system. [Figure 11] FIG. 11 is a diagram showing changes over time in elevator steering angle and vertical acceleration in an example flight simulation for evaluating the effectiveness of the above-mentioned automatic control system. [Figure 12] FIG. 13 is a diagram of an example flight simulation to illustrate a method for evaluating the effectiveness of the above-mentioned automatic control system. [Figure 13] 11 is a diagram for explaining a method for simply calculating reliability information of wind speed measurement. FIG. DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

[0030] Hereinafter, an embodiment of the present invention will be described with reference to the drawings.

[0031] [Gust response or gust load reduction system configuration] 1 is a block diagram showing the configuration of an automatic control system 1 for an aircraft according to one embodiment of the present invention. In this embodiment, the automatic control system 1 is configured as a Doppler LIDAR type optical gust response or gust load alleviation system mounted on an aircraft.

[0032] As shown in FIG. 1, the automatic control system 1 of this embodiment includes a measurement unit 10, a control and calculation unit 20, and a rudder 30.

[0033] (Measuring section) The measurement unit 10 is a device that measures the difference between the wind speed that the aircraft body 200 is currently experiencing and the wind speed that it will encounter in the future as prediction information, and adds information that determines whether the measurement value is an estimated error or valid or invalid to the measurement information and outputs it.

[0034] In this embodiment, the measurement unit 10 emits pulsed laser light in two directions in the atmosphere, receives the reflected light, and measures the wind speed in the axial direction of each optical axis based on the amount of Doppler shift in frequency between the emitted laser light and the reflected light. The measurement unit 10 has an optical transceiver 12, a switch 13, a first telescope 15, a second telescope 14, and a signal processing unit 11.

[0035] The optical transceiver 12 generates the emitted laser light and converts the received laser light into an electrical signal. The switch 13 sequentially selects between the first telescope 14 and the second telescope 15. The first telescope 14 emits and receives laser light below the axis through a window 40. The second telescope 15 emits and receives laser light above the axis through a window 40. Here, the angles made by the optical axes of the upper and lower laser lights with respect to the axis are set to θ.

[0036] Here, an example is described in which the first telescope 14 is a telescope for long distances and the second telescope 15 is a telescope for close distances, but this is not limited to this, and the same telescope (e.g., a telescope for close distances or a telescope for long distances) may be used for the first telescope 14 and the second telescope.

[0037] The optical transceiver 12 generates and amplifies a laser beam of a single wavelength, for example 1.55 μm, and receives the scattered light to measure the frequency change (wavelength change) caused by the Doppler effect, thereby measuring wind speed. This is generally called a Doppler LIDAR, and LIDAR is an abbreviation of "Light Detection And Ranging," a remote observation method that uses light.

[0038] The amount of frequency change due to the Doppler effect can be found by comparing the frequency (wavelength) of the received light (scattered light) received via the first telescope 15 or the second telescope 14 with the frequency (wavelength) of the transmitted light. Using this principle, the measurement unit 10 measures preview information, which is the difference between the wind speed the aircraft is currently experiencing and the wind speed it will encounter in the future.

[0039] In this embodiment, the laser light, which is the transmission signal, is a continuous pulse train emitted into the atmosphere, so the reception signal is also a pulse train. The amount of frequency change due to the Doppler effect, that is, the signal train representing the difference between the frequency of the reception signal that receives the reflection signal of the transmission signal and the frequency of the transmission signal, is also a pulse train.

[0040] The signal processor 11 calculates the wind speed and reliability information of the wind speed measurement. The reliability information means information for determining the estimation error or validity / invalidity of the measurement value, which is prediction information.

[0041] The estimated error refers to the difference between the wind speed measured using the first telescope 14 (wind speed at optical axis L1) and the wind speed measured using the second telescope 15 (wind speed at optical axis L2). The estimated error is in the same units as the measured value, and the range estimated to include the true value is expressed as the absolute value of the deviation centered on the measured value.

[0042] The measurement value is judged to be valid or invalid if the difference between the upper and lower wind speeds (the difference between the measurement value on optical axis L1 and the measurement value on optical axis L2) is less than a first set value, and the measurement value is judged to be invalid if the difference exceeds the first set value. Note that the reliability information of the wind speed measurement may be calculated using, for example, the technology of Patent Document 4.

[0043] The first set value is set as an absolute value based on the difference between the wind speed the aircraft is currently experiencing and the wind speed it will encounter in the future, taking into account the reliability of the entire system. The first set value is not a fixed value such as 1 m / s, but a variable obtained by multiplying the up and down wind measurement value by a coefficient of 1 or less can be used.

[0044] As shown in FIG. 2, for example, the measurement unit 10 emits laser light in each direction (optical axes L1, L2) that form angles θ1 and θ2 above and below the aircraft axis X of the aircraft 200, and receives the scattered light as spectrum data for the distance along the optical axis direction of the laser light. For example, 20 measurement values ​​are obtained for each range bin that has a distance width (for example, 25 m) corresponding to the sampling interval from a measurement range of within 500 m. In other words, the measurement unit 10 measures the wind speed at 20 positions for each range bin of each optical axis L1, L2 by emitting laser light in each direction (optical axes L1, L2). This measured wind speed data is also called a lidar measurement value.

[0045] Since the wind speed data may contain erroneous measurement values, the measurement unit 10 adds independent reliability information to the wind speed data of each range bin. In other words, reliability information about the wind speed data (estimated error, valid / invalid flag that is the result of determining whether the wind speed data is valid / invalid) is added to the wind speed data in each range bin. Since the wind speed data is the wind speed component in the optical axis direction, it can be calculated by determining the up / down wind and wind speed vector using the technology of Patent Document 3, for example.

[0046] (Control and calculation section) The control and calculation unit 20 calculates the angle of the rudder 30 or the thrust force that reduces the effect of the wind speed on the aircraft 200, based on the wind speed value (wind speed data) in the planned flight direction of the aircraft measured by the measurement unit 10. The calculated angle of the rudder 30 or the thrust force generates input information (control command) for automatically controlling the rudder 30 or a thrust generator (not shown).

[0047] In addition, when the control and calculation unit 20 detects, based on reliability information about the wind speed data in the planned flight direction of the aircraft measured by the measurement unit 10, that a predetermined amount of wind speed data with an invalid flag added thereto is included (in this embodiment, when an invalid flag is added to the wind speed data in all range bins), the control and calculation unit 20 executes automatic control of the rudder angle based on the output of the acceleration sensor 50 (see Figure 1) attached to the aircraft 200.

[0048] The acceleration sensor is configured to detect, for example, acceleration in one or two axial directions perpendicular to the aircraft axis. For automatic control of the rudder angle based on the output of the acceleration sensor 50, for example, a feedback control technique of the rudder angle is adopted so that the acceleration acting on the aircraft 200 is kept below a predetermined value.

[0049] Specifically, the rudder 30 as a control object corresponds to a moving surface or an auxiliary wing for controlling the lift, drag or attitude of an aircraft, such as an elevator, rudder, flaperon, throttle, spoiler, DLC (Direct Lift Control) flap, aileron, etc. If the aircraft 200 is a propeller aircraft, the pitch angle of the propeller may also be a control object. Note that the control object is not limited to the rudder 30, and may include devices such as a jet engine or a propeller thruster that generate thrust for the aircraft 200.

[0050] FIG. 3 is a flowchart showing the operation of the control and calculation unit 20.

[0051] The control calculation unit 20 acquires wind speed data from the measurement unit 10, and proceeds with processing according to the reliability information included in the wind speed data (step 101). First, if an invalid flag is added to all range bins as reliability information (No in step 102), the control according to the present invention is not performed while this state continues, and only feedback control based on the output of the acceleration sensor 50 described above is used (step 103).

[0052] Although the wind speed data can be used as is, using the average value over a certain period of time will provide more reliable data that eliminates minute fluctuations. In this case, if the wind speed data to be averaged contains invalid information, the error in the average value will be large, so it is preferable to calculate the average value excluding data with invalid flags attached, as shown in Figure 4.

[0053] In this embodiment, the control and calculation unit 20 removes the received signal with the invalid flag added, then performs spectrum integration on the received signal, and calculates the wind speed data from the moving average of the integral value. By moving the time range for spectrum integration over time, it is possible to use the measurement value with a high data rate as the input for automatic control.

[0054] The average value can be calculated by averaging the wind speed values, but in that case, no improvement in the S / N ratio can be expected. Therefore, the wind speed spectrum data is integrated to simultaneously improve the S / N ratio. An improved S / N ratio is expected to improve reliability and measurement accuracy. However, if the average value for a certain period is used as input for automatic control every time it is obtained, the data rate will be low and the discontinuity of control will be large. For this reason, a moving average value is calculated by parallel calculation to prevent deterioration of the data rate. Delays caused by averaging are compensated for by setting the range of information used to a distant location.

[0055] When an invalid flag is added to some range bins as reliability information, or when an invalid flag is not added to all range bins, the control calculation unit 20 divides the estimated measurement error as reliability information by a second set value, subtracts the result from 1, and calculates the authority for each range bin (step 104).

[0056] The second set value uses, for example, a variable obtained by multiplying the vertical wind speed measurement value (wind speed data on optical axes L1, L2) by a coefficient of 1 or less (for example, 0.5 to 1). The vertical wind speed measurement value is expressed as a relative value to the vertical wind that the aircraft 200 is currently encountering. When the reliability of the entire system is high, setting the coefficient to a value close to 1 increases the effectiveness of control. On the other hand, when the reliability of the entire system is low, setting a small value reduces the risk of vibration due to control.

[0057] Next, the control calculation unit 20 calculates a rudder angle that reduces the effect of the wind speed on the aircraft (step 105). In the rudder angle calculation, a rudder angle that reduces the shaking of the aircraft due to changes in wind speed is calculated. At this time, the rudder angle command is sent to the actuator of the rudder 30 by multiplying the calculated rudder angle change amount by the larger value of 0 or the authority and sending out the rudder angle command (step 106).

[0058] At this time, if an invalid flag is added to some range bins as reliability information, or if an invalid flag is not added to all range bins, the control calculation unit 20 calculates a value by subtracting an estimated measurement error from the wind speed measurement value for each range bin, and regards the greater of this value or 0 as the correct wind speed measurement value and uses it as the control input.

[0059] 5, when the measured value of the wind speed is used as the control input as is, for example, the function W0 shown by the dashed line is used as the control input. In contrast, in one embodiment of the present invention, the function W1 obtained by subtracting an estimated measurement error from the wind speed data as shown by the solid line is used as the control input. As a result, the risk of the measured value of the wind speed being inflated by an error, resulting in vibration due to excessive control even when the wind speed is actually small, is reduced.

[0060] When the measurement error of a certain range bin is larger than a predetermined absolute value, or when an invalid flag is added to only some of the range bins as reliability information, the control calculation unit 20 may use a value obtained by interpolating the two closest values ​​instead of the value of that range bin as the control input.

[0061] In Figure 6, an invalid flag is added to range B, so the wind speed value for this range bin is W B is not used, and the wind speed values ​​W in the adjacent range bins A and B are used. A and W C The interpolated value W BWhen there are consecutive range bins with invalid flags, such as ranges E and F, the wind speed values ​​W D and W G The interpolated value W E ' and W F The control input is calculated using these wind speed values ​​and the technology of, for example, Patent Document 1, Non-Patent Document 3, or Non-Patent Document 4. As a result, the risk of vibration being generated by excessive control even when the wind speed is actually small due to an inflated wind speed measurement value caused by an error is reduced.

[0062] When a bias-like measurement error, which is a constant value, is added to the measurement value of each range bin, the control calculation unit 20 may calculate a control input that is robust against the measurement error by using a control gain that offsets the effect of the error. In this case, the control input is first calculated based on the following formula.

[0063]

number

[0064] The control law is given as a discrete-time system, and u(k) represents the current value of the control input (elevator command). x(k) is a vector representing the state quantity (specifically, velocity, angular velocity, attitude angle, etc.) of the aircraft 200. a (k) represents the vertical wind speed currently encountered by the aircraft 200. a (k+1) represents the vertical wind speed that the aircraft 200 will encounter one discrete time step from the present. a (k+n) represents the vertical wind speed that the aircraft 200 encounters after n discrete time steps. These wind speeds are replaced by lidar measurements. For example, if the sampling period is T seconds and the speed of the aircraft 200 is V, then w a For (k+n), the measurement value V×T×n (m) ahead of the aircraft 200 is used. If this position does not match the range bin, the measurement value in the nearby range bin is used instead. K B is a constant gain matrix multiplied by the state quantity, and can be designed by any method. Also, k0, k1, ..., kn is a gain multiplied by the wind speed, and can be calculated using the technique described in Non-Patent Document 4. The structure of this control input is shown in Figure 7.

[0065] The control gains that offset the bias error are k0, k1, ..., k n can be obtained by solving the following constraints: k n =-k0-k1-…-k n-1

[0066] The design method in Non-Patent Document 4 solves an optimization problem with linear constraints on these gains, so this design method can be applied even when the above constraints are imposed. When a bias error δ is added to each vertical wind speed value, the control input is calculated using the gains obtained under this constraint, and the result is k0+k1+...k n Since =0,

[0067]

number

[0068] Here, when the objective is to reduce the vertical oscillation of the aircraft 200, the rudder 30 to be controlled is assumed to be, for example, a flaperon that changes the left and right ailerons in phase, but is not limited to this and may be a spoiler or DLC flap, or a combination of two or more of these. Alternatively, the attitude may be changed by the elevator to indirectly change the lift.

[0069] In addition, when the purpose is to reduce the change in airspeed of the aircraft 200 due to wind shear, the rudder to be controlled is assumed to be a throttle or spoiler, but the present invention is not limited to this, and either one of them may be used, or both may be used in combination. Alternatively, in the case of a propeller aircraft, the pitch angle may be used.

[0070] In general, for jet aircraft, it is appropriate to adjust the airspeed with the throttle, but because of the large time constant, a delay of several tens of seconds occurs. Deploying spoilers increases drag, which reduces fuel efficiency. For this reason, when decelerating, it is appropriate to deploy the throttle and spoilers simultaneously, and gradually retract the spoilers as the airspeed decreases. When accelerating, only the throttle is used. However, it is difficult to imagine a situation in which acceleration control is required during cruising, and it is assumed that this is to prevent stalls during takeoff and landing, but at low altitudes, the observation range of the Doppler LIDAR is long, so throttle control alone is considered to be effective. In this case, it is not necessary to have two optical axes; for example, one line of sight can be directed in the direction of flight, and the number of range bins can be 80, with the range bin distance width set to, for example, 100 m, resulting in a measurement range of 8 km.

[0071] Furthermore, if the objective is to reduce lateral oscillation of the aircraft 200, a rudder is assumed as the rudder to be controlled. Ailerons may also be used in combination to apply steady lateral G. In this case, unlike the above example, it is necessary to measure the crosswind by emitting laser light forward to the left and right.

[0072] If the objective is to reduce the fluctuation in the roll angle of the aircraft 200, the aileron is assumed to be the rudder to be controlled. In this case, for example, four lines of sight A to D are measured as shown in FIG. 8. The measured values ​​of lines of sight A and C are vector-converted to obtain the vertical wind W that the right wing encounters. ZR The vertical wind W that the left wing encounters is calculated by vector conversion of the measured values ​​of line of sight B and line of sight D. ZL Calculate W ZR and W ZL Based on the difference between the line of sight and the ailerons, the ailerons can be controlled to reduce the fluctuation in roll angle. 4 By measuring the line of sight, it is possible to reduce vertical oscillation, airspeed changes due to wind shear, and lateral oscillation.

[0073] By repeatedly executing the above process, automatic control of the steering angle based on the preview information (hereinafter also referred to as preview control based on remote airflow) is performed. Here, in the case of conventional feedback control for reducing aircraft sway (feedback control of rudder 30 based on the output of acceleration sensor 500), there was a possibility that the initial vibration could not be responded to due to control delays, or that vibration could be exacerbated. In contrast, in the case of preview control based on remote airflow as in this embodiment, the rudder can be controlled in advance, anticipating delays, so there is no effect of average delay, and only slight effects such as observation errors of remote airflow and rudder angle errors remain.

[0074] As described above, according to this embodiment, the estimation error of the measurement or the information for determining whether the measurement is valid or invalid is appropriately used in the control, thereby improving the accuracy and reliability of the control. Furthermore, according to this embodiment, even if there is a slight measurement error in the preview information, it is possible to reduce the shaking of the aircraft when the aircraft enters turbulence, or to reduce the load applied to the aircraft.

[0075] There have been many accidents caused by turbulence, such as the accident of American Airlines Flight 280 on December 16, 2014. The aircraft unexpectedly encountered clear air turbulence that could not be detected by radar, causing the aircraft to shake violently, and passengers and crew were seriously injured. It is believed that such accidents can be prevented by applying the present invention to appropriately automatically control lift and reduce aircraft shaking when encountering turbulence.

[0076] In addition, in the case of the accident of Japan Airlines Flight 356 that occurred on October 21, 2002, the aircraft was shaken by inappropriate control in response to an increase in airspeed due to wind shear, and passengers and crew members were seriously injured. It is believed that such accidents can be prevented by applying the present invention to appropriately automatically control thrust to offset the increase in airspeed due to wind shear.

[0077] Furthermore, in the case of the accident involving All Nippon Airways Flight 569 that occurred on September 27, 2002, passengers were seriously injured when the aircraft swayed sideways due to a sudden change in crosswind. It is believed that such accidents could be prevented by applying the present invention to appropriately automatically control the rudder to reduce aircraft shaking caused by crosswind changes just before landing.

[0078] [Application example of gust response mitigation system] An example of the application of the present invention as a gust response mitigation system is shown below. As in Non-Patent Document 3, consider the minute vertical motion obtained by linear approximation from a horizontal steady flight state of a large passenger aircraft equipped with four jet engines. The conditions such as altitude and speed in the horizontal steady flight state are shown in Table 1 below.

[0079] [Table 1]

[0080] For the mathematical model representing the motion, the technology described in Non-Patent Document 4 is used to design a control law that generates a control input for reducing the gust response. The elevator is used as the control surface. In this case, the control law is expressed by the following formula.

number

[0081] The control law is given as a discrete-time system, where u(k) represents the current value of the control input (elevator command). x(k) is a vector representing the state of the aircraft (specifically, velocity, angular velocity, attitude angle, etc.). a (k) represents the vertical wind speed currently being encountered by the aircraft. a (k+1) represents the vertical wind speed that the aircraft will encounter one discrete time step from the present. w a (k+n) represents the vertical wind speed encountered after n discrete time steps. These wind speeds are replaced by measurements by the lidar. For example, if the sampling period is T seconds and the aircraft speed is V, then w a(k+n) is the measurement value V×T×n(m) ahead of the aircraft position. If this position does not match the range bin, the measurement value in the nearby range bin is used instead. K B is a constant gain matrix multiplied by the state quantity, and can be designed by any method. Also, k0, k1, ..., k n is a gain multiplied by the wind speed, and can be calculated using the technique described in Non-Patent Document 4.

[0082] A simulation was performed in which the aircraft was subjected to a vertical gust of wind. The aircraft moves along the X-axis from X=0, at a speed of approximately 220 m / s. The measurement unit obtains vertical wind speed measurements up to 600 m ahead on the X-axis every 0.1 seconds, with a range bin distance width of 25 m. The vertical wind speed encountered by the aircraft is assumed to be a uniform flow (wind speed flow with the same distribution regardless of altitude) with the distribution shown in Figure 9. The vertical flow has a peak value of 10 m / s at 4,676 m on the X-axis. Spike noise is assumed to occur with a 1% probability in the measurements for each range bin. Figure 10 shows an example of measured values ​​and true values.

[0083] The circles in Fig. 10 represent the measurement values ​​at each range bin 20.2 seconds after the start of the simulation, with the measurement values ​​near 4,800 m containing large spike noise, resulting in outliers. Figure 11 shows the results of a simulation in which control was performed using these measurement values. The dashed lines represent the elevator steering angle and vertical acceleration when directly affected by spike noise, and the solid lines represent when spike noise is nullified using a nullification flag. In the case of the dashed lines, an excessive elevator steering angle command is generated due to the influence of spike noise, resulting in large vertical acceleration. In contrast, in the case of the solid lines, the spike noise is nullified by the present invention, and the behavior shown by the dashed lines does not occur.

[0084] This simulation shows the effect of the present invention as a gust response alleviation system, but the present invention is also applicable to a gust load alleviation system. This simulation was performed using the technique described in Non-Patent Document 4 to minimize the vertical acceleration K B and K. FHowever, this technology can also calculate a gain that minimizes the load, and the present invention can be applied to a gust load reduction system configured using such a gain.

[0085] Furthermore, at low altitudes during takeoff and landing, the concentration of aerosol particles in the atmosphere is high and the observation range of the Doppler LIDAR is long, passengers and crew are wearing seat belts, the flight route can be changed at the pilot's sole discretion, and there is a risk of crashing due to a decrease in altitude if a malfunction occurs in the control of the present invention. Therefore, at flight altitudes of, for example, 500m or less, it is more desirable to simply provide the pilot with advance information about turbulence rather than performing the vertical control of the present invention.

[0086] Regardless of the application of the present invention, it is necessary to evaluate the effectiveness of the automatic control system in order to apply it. FIG. 12 shows an example of a flight simulation, in which the horizontal axis indicates the vertical acceleration fluctuation without control, and the vertical axis indicates the vertical acceleration fluctuation when the present invention is applied. According to the automatic control system of this embodiment, a certain reduction effect was confirmed with respect to the shaking of the aircraft caused by wind speed, as shown by the plot of the result being displayed below the dashed line in part A. In this example, most of the plots are below the dashed line, and it can be seen that the shaking is reduced to almost half by the control, but some plots are above the dashed line. In proceeding with practical application, it is sufficient to determine whether the probability and the deviation amount are acceptable. Regarding the dashed line in part B, it includes a part where the shaking increases due to the control, but it is sufficient to set an acceleration fluctuation amount that has little adverse effect even if it increases.

[0087] The above examples are flight simulation results based on wind data that an aircraft actually encountered, but flight test results may also be displayed. Also, although this explanation is about acceleration changes, the effectiveness of control can be evaluated by a similar method for load changes as well.

[0088] The reliability information of the wind speed measurement output from the measurement unit 10 of the automatic control system 1 of the aircraft is calculated by hardware processing, as in the technique of Patent Document 4, for example, but this has the drawback of increasing costs and size. For this reason, a simple calculation method using software will be described with reference to FIG.

[0089] In FIG. 13, when the angles of the optical axes L1 and L2 with respect to the axial direction are θ1 and θ2, respectively, θ1 and θ2 are assumed to be 5 to 10 degrees, and in this case, it can be considered that the observations are in almost the same direction. Therefore, the difference between the wind speeds V1 and V2 of range bins at the same distance is the measurement error. However, when there is turbulence, it may be superimposed on the measurement error, so the measurement error may be multiplied by a coefficient less than 1. When estimating the airflow vector from each measurement value obtained from each observation axis, the airflow vector is estimated assuming that the measurement error is included in the measurement value, and the estimation error is calculated at the same time. At this time, when the measurement error or estimation error is extremely large compared to normal turbulence that can occur in the natural world, for example, when the wind speed difference between each observation axis exceeds 10 m / s, the reliability of the airflow vector estimation decreases, so the measurement value is judged to be invalid and an invalid flag is applied to the output data string. The threshold value for this judgment may be different depending on the range.

[0090] [others] More than half of all passenger aircraft accidents are caused by turbulence, so reducing turbulence-related accidents is an urgent issue. For this reason, passenger aircraft are required to be equipped with weather radar, which can detect cumulonimbus clouds that cause turbulence, but cannot detect turbulence that occurs in clear weather conditions.

[0091] In contrast, while Doppler LIDAR has the advantage of being able to observe remote air currents in clear weather, pilots at airline operators have pointed out that its effective observation range is short, which has been an obstacle to its practical use. However, even with a short observation range of around 500 m, the gust response mitigation system of the present invention can be suitably applied as a means of reducing aircraft shaking when an aircraft encounters turbulence.

[0092] The present invention is not limited to the above-described embodiment, but can be implemented in various modified forms, and the scope of such modifications falls within the scope of the technical idea of ​​the present invention.

[0093] For example, in the above embodiment, the explanation is based on the assumption that a Doppler lidar that uses light waves, which are a type of electromagnetic wave, is used as the measurement unit 10, but the present invention can also be applied to a Doppler radar that uses radio waves. In this case, the measurement unit emits radio waves in the planned flight direction of the aircraft, receives the scattered waves in the atmosphere, and measures the remote wind speed in the radiation axis direction of the radio waves based on the Doppler shift amount of the scattered radio waves with respect to the emitted radio waves.

[0094] In the above embodiment, a case has been described in which the optical axis of the laser light is directed in two directions, up and down, forward. However, the present invention can also be applied to a case in which the optical axis of the laser light is directed in two directions, left and right, forward, or in three or more directions (four directions in the example of Figure 8).

[0095] Furthermore, although the above embodiment is premised on application to an aircraft as the airframe 200, it can also be applied to, for example, a case where the motion of a ship is reduced by observing waves and water currents in the traveling direction of the ship. Even in the case of a vehicle, it is expected to be effective in preventing accidents caused by gusts of wind, for example, at the exit of a tunnel. [Explanation of symbols]

[0096] 1...Gust response or gust load reduction systems 10…Measuring section 11...Signal processing unit 12...Optical transmitter / receiver 13...Switch 14...Second telescope 15…First Telescope 20...Control calculation section 30…Rudder 40…Window 200…Aircraft

Claims

1. a measurement unit that measures the difference between the wind speed that the aircraft is currently experiencing and the wind speed that the aircraft will encounter in the future as prediction information, and adds information that determines the estimated error or validity of the measurement value to the measurement information and outputs the measurement information; A device for controlling the rudder or thrust that controls the lift, drag, or attitude of an aircraft; a control calculation unit that calculates the rudder angle or thrust that reduces the effect of the wind speed on the aircraft based on the wind speed value in the planned flight direction of the aircraft measured by the measurement unit, and generates a control command to be output to the rudder or device based on the calculated value; An automatic control system for an aircraft comprising:

2. 2. The automatic control system for an aircraft according to claim 1, The measurement unit emits electromagnetic waves in a planned flight direction of the aircraft, receives the scattered waves in the atmosphere, and measures a remote wind speed in a radiation axis direction of the electromagnetic waves based on a Doppler shift amount of the scattered electromagnetic waves with respect to the emitted electromagnetic waves. Automatic control system of aircraft.

3. 3. The automatic control system for an aircraft according to claim 2, The measurement unit obtains a two-dimensional or three-dimensional vector of wind speed by setting or scanning two or more lines of sight along the radiation axis of the electromagnetic wave. Automatic control system of aircraft.

4. An automatic control system for an aircraft according to any one of claims 1 to 3, The control and calculation unit calculates the wind speed value based on a moving average value of a spectrum integral of a received signal from which a received signal to which invalid information has been added has been removed. Automatic control system of aircraft.

5. An automatic control system for an aircraft according to any one of claims 1 to 3, further comprising an acceleration sensor for detecting an acceleration acting on the aircraft; The control and calculation unit executes automatic control of the rudder based on the output of the acceleration sensor when the estimation error is equal to or greater than a set value, or when the control and calculation unit receives preview information to which invalid information is added. Automatic control system of aircraft.

6. An automatic control system for an aircraft according to any one of claims 1 to 3, The control calculation unit divides the estimated error by a set value, subtracts the result from 1, defines the result as an authority, and multiplies the control command by either 0 or the authority, whichever is larger. Automatic control system of aircraft.

7. An automatic control system for an aircraft according to any one of claims 1 to 3, The control calculation unit generates the control signal by regarding the wind speed value as the larger value of the measured value minus the estimated error or 0. Automatic control system of aircraft.

8. An automatic control system for an aircraft according to any one of claims 1 to 3, When the measurement error of a certain range bin is larger than a predetermined absolute value or when invalid information is added to the measurement value, the control calculation unit calculates the rudder angle using the measurement values ​​in the range bins before and after the certain range bin. Automatic control system of aircraft.

9. An automatic control system for an aircraft according to any one of claims 1 to 3, The control calculation unit uses a control gain that offsets the effect of a bias-like measurement error that is a constant value added to the measurement value of each range bin. Automatic control system of aircraft.

10. 2. A method for evaluating the effectiveness of an automatic control system of an aircraft according to claim 1, comprising: Conducting flight tests or flight simulations of the aircraft under predetermined wind and flow conditions, including vertical wind speeds; Plotting the results of the flight test or flight simulation with the amount of change in acceleration of the aircraft when the automatic control based on the control command is not performed as the horizontal axis and the amount of change in acceleration of the aircraft when the automatic control is performed as the vertical axis; The deviation of the plot downward from the preset broken line is regarded as the effect of the automatic control in reducing the gust response. A method for evaluating the effectiveness of automatic aircraft control systems.

11. A measuring device for an automatic control system of an aircraft, which measures a difference between a wind speed currently being experienced by an aircraft and a wind speed that the aircraft will encounter in the future as predictive information, A signal processing unit that adds the difference between the respective measurement values ​​obtained from the radiation axis of electromagnetic waves of two or more lines of sight to the measurement information as information for determining whether it is valid or invalid, and outputs it. A measuring device comprising:

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