Patent Description:
Conventionally, a Doppler navigation device with an added function to output position information with excellent accuracy by reducing the effect of velocity errors due to Doppler frequency shift caused by the beam length change velocity that occurs when the aircraft attitude changes has been known (see, for example, Patent Literature <NUM>). In this Doppler navigation device, a correction quantity is determined to compensate for the velocity error caused by the beam length change.

<CIT> and <CIT> discloses an automatic landing system provided with: an imaging device mounted on a vertical take-off and landing aircraft; a relative-position acquisition unit that performs image processing on an image, of a marker provided to a target landing point, captured by the imaging device, and that acquires a relative position between the vertical take-off and landing aircraft and the target landing point; a relative-altitude acquisition unit for acquiring a relative altitude between the vertical take-off and landing aircraft and the target landing point; and a control unit for controlling the vertical take-off and landing aircraft in a plurality of control modes so that the relative position becomes zero, wherein the control modes include a hovering mode and a landing mode. In the hovering mode, the relative altitude of the vertical take-off and landing aircraft is lowered to a predetermined relative altitude when the relative position is within a first threshold value. The transition to the landing mode takes place upon satisfaction of predetermined conditions including a condition in which the relative position is within a predetermined threshold value which is less than the first threshold value. In the landing mode, the relative altitude of the vertical take-off and landing aircraft is further lowered so as to make a landing on the target landing point.

<CIT> discloses a navigation apparatus with image recognition, including: an imaging section for obtaining a stereo image of a target spot; an inertial information detecting section for measuring an attitude angle of a body and an acceleration of the body; an image process calculating section for calculating a relative position of the body with respect to the target spot based on the stereo image and the attitude angle; and a navigation calculating section for calculating navigation information based on the attitude angle, the acceleration and the relative position.

<CIT> discloses a device including: an acquisition unit that acquires information indicating a position estimation system selected from among a plurality of position estimation systems for estimating a position of a flight vehicle; and a position estimation unit that estimates the position of the flight vehicle from first information generated by using an inertial sensor of the flight vehicle and second information generated through the position estimation system based on a parameter for the position estimation system.

In the paper titled "Autonomous Landing of a Multirotor Micro Air Vehicle on a High Velocity Ground Vehicle", Alexandre Borowczyk et al. disclose the problem of the automated landing of a quadcopter on a ground vehicle moving at relatively high speed and a system architecture, including the structure of a Kalman filter for the estimation of the relative position and velocity between the quadcopter and the landing pad, as well as a controller design for the full rendezvous and landing maneuvers.

In Patent Literature <NUM>, a Doppler navigation device that transmits radio waves from an aircraft toward a ground surface is used. The Doppler navigation device measures the velocity of the aircraft by correcting the velocity error. The velocity of the aircraft measured by the Doppler navigation is used to control the
position of the aircraft in flight. If the corrected velocity of the aircraft can be used in the flight control, it is possible to make the position less likely to deviate from a target landing point.

On the other hand, it has been known to measure the acceleration of the aircraft and perform flight control based on the measured acceleration in order to control the position of the aircraft in flight.

When an aircraft is to hover stably at target coordinates, the flight control of the aircraft is performed using the acceleration of the aircraft measured by an accelerometer as input information. In this case, the output value of the acceleration acquired by the accelerometer may include a drift component due to a drift phenomenon. If the aircraft is made to hover using output values of the acceleration that include the drift component, it is difficult to obtain accurate acceleration, and the position will deviate from the original target landing point, making it difficult to hover with high accuracy over the target landing point.

In view of this, it is an object of the present invention to provide an aircraft position control system, an aircraft, and an aircraft position control method that can stably keep the position of an aircraft in an inertial space.

An aircraft position control system according to the present disclosure keeps an aircraft at target coordinates in an inertial space with respect to a target landing point that fluctuates. The aircraft position control system includes: an acceleration correction processing unit that, based on acceleration of the aircraft and an attitude of the aircraft, outputs first attitude correction acceleration in which the acceleration of the aircraft is corrected; a complementary filter that, based on the first attitude correction acceleration and inertial velocity of the aircraft, outputs second attitude correction acceleration in which a drift component included in the first attitude correction acceleration is removed; and a smoothing processing unit that, based on the second attitude correction acceleration and relative coordinates between the aircraft and the target landing point, outputs smoothed relative coordinates obtained by smoothing the relative coordinates.

An aircraft according to the present disclosure includes: an acceleration acquisition unit that acquires acceleration of the aircraft; an attitude acquisition unit that acquires an attitude of the aircraft; an inertial velocity acquisition unit that acquires inertial velocity of the aircraft; a relative coordinate acquisition unit that acquires relative coordinates of the aircraft; and the above-described aircraft position control system.

An aircraft position control method according to the present disclosure is for keeping an aircraft at predetermined coordinates in an inertial space with respect to a target landing point that moves. The aircraft position control method includes the steps of: outputting, based on acceleration of the aircraft and an attitude of the aircraft, first attitude correction acceleration in which the acceleration of the aircraft is corrected; outputting, based on the first attitude correction acceleration and inertial velocity of the aircraft, second attitude correction acceleration in which a drift component included in the first attitude correction acceleration is removed; and outputting, based on the second attitude correction acceleration and relative coordinates between the aircraft and the target landing point, smoothed relative coordinates obtained by smoothing the relative coordinates.

According to the present disclosure, even if the target landing point moves, the average position of the motion of the target landing point can be kept stably with only the information available on the aircraft side without being affected by the motion.

An embodiment of the present disclosure will hereinafter be described based on the drawings. This invention is not limited by the embodiment. In addition, the components in the following embodiment include those that can be replaced and easily conceived by those skilled in the art, or those that are substantially the same. Furthermore, the components described below can be combined as appropriate, and if there is more than one embodiment, the embodiments can be combined.

<FIG> is a schematic structure diagram illustrating one example of a position control system of an aircraft according to the present embodiment, and <FIG> is an explanatory diagram illustrating a state in which the aircraft according to the present embodiment heads to a target landing point.

As illustrated in <FIG>, an aircraft <NUM> is a flight vehicle as a rotorcraft (for example, helicopter, drone, etc.). In this embodiment, the aircraft <NUM> is an unmanned vehicle. The aircraft <NUM> may be any flight vehicle that can move forward, backward, sideways, swirl, and hover, and may be a manned vehicle. The aircraft <NUM> includes a position control system <NUM>, and its flight is controlled by the position control system <NUM> so that the aircraft <NUM> lands at a target landing point <NUM> illustrated in <FIG>.

In this embodiment, the target landing point <NUM> is provided on a vessel <NUM> as illustrated in <FIG>. Therefore, the aircraft <NUM> lands on (lands on a deck of) the vessel <NUM> as a moving vehicle moving on the water. The vessel <NUM> includes a restraint device to restrain the aircraft <NUM> when the aircraft <NUM> lands at the target landing point <NUM>, which is not illustrated in the drawing. However, the target landing point <NUM> is not limited to the vessel <NUM>, and may alternatively be provided on a vehicle or the like as a moving object moving on the ground, or on non-moving equipment or on the ground.

The target landing point <NUM> includes a marker <NUM> for the aircraft <NUM> to capture the position of the target landing point <NUM>. <FIG> is an explanatory diagram illustrating one example of the marker provided at the target landing point. As illustrated in the drawing, the marker <NUM> is a square-shaped AR marker with two colors, black and white, for example. The marker <NUM> is not limited to the AR marker and may alternatively be any marker that can capture the position of the target landing point <NUM> by image processing, such as an H mark or an R mark indicating the landing point of a heliport. As the marker <NUM>, a plurality of markers with different shapes may be provided on the vessel <NUM>, and the aircraft <NUM> may be guided to the target landing point <NUM> corresponding to any of the different markers <NUM>. In this embodiment, the marker <NUM> is provided on the vessel <NUM> to capture the position of the target landing point <NUM>, but the configuration is not limited in particular as long as the position of the target landing point <NUM> can be acquired.

The position control system <NUM> for the aircraft according to this embodiment is a system for controlling the position of the aircraft <NUM> so that the aircraft <NUM> in flight lands at the target landing point <NUM>. The position control system <NUM> is mounted on the aircraft <NUM>. The position control system <NUM> includes a camera <NUM>, a navigation device <NUM>, and a control unit <NUM> as illustrated in <FIG>.

The camera <NUM> is a photographing device mounted on the aircraft <NUM> via a gimbal that is not illustrated. The camera <NUM> may be a monocular camera, a compound-eye camera, an infrared camera, or the like, as long as the marker <NUM> can be photographed. The camera <NUM> is provided to photograph the marker <NUM> at the target landing point <NUM> from the aircraft <NUM>. The camera <NUM> is able to adjust a photographing direction via the gimbal that is not illustrated. In this embodiment, the camera <NUM> is controlled by the control unit <NUM> so that its photographing range B (see <FIG>) faces right downward in the vertical direction as one example. The camera <NUM> may be controlled by the control unit <NUM> so that the photographing range B faces forward at an angle to the vertical direction. The camera <NUM> may omit the gimbal and be fixed right under the body of the aircraft <NUM> so that the photographing direction faces downward in the vertical direction.

The navigation device <NUM> is, for example, an inertial navigation system (INS). In this embodiment, the navigation device <NUM> will be described in application to an inertial navigation system; however, any navigation device <NUM> may be used without particular limitations. The navigation device <NUM> may alternatively be an inertial navigation device that includes a global positioning system (GPS) to improve the accuracy of position measurement. In this embodiment, the application to the inertial navigation device including the GPS is described; however, it is not limited to the GPS and any position measurement unit that can measure the position with high accuracy may be used. For example, a quasi-zenith satellite system may be used. If the position can be measured with high accuracy only with the navigation device <NUM>, the GPS or other position measurement unit may be omitted.

The navigation device <NUM> including the GPS acquires, for example, attitude angles of the aircraft <NUM> in a roll direction, a yaw direction, and a pitch direction, and aircraft velocity, inertial velocity, aircraft acceleration, a nose azimuth, and position coordinates in the earth coordinate system of the aircraft <NUM>. The navigation device <NUM> may have an attitude angle sensor to detect the attitude angle of the aircraft <NUM>, a velocity sensor to detect the aircraft velocity of the aircraft <NUM>, an acceleration sensor to detect the aircraft acceleration of the aircraft <NUM>, and a sensor to detect the nose azimuth of the aircraft <NUM>. The navigation device <NUM> outputs the acquired attitude angle, aircraft velocity, inertial velocity, aircraft acceleration, nose azimuth, and position coordinates of the aircraft <NUM> to the control unit <NUM>. Thus, the navigation device <NUM> functions as an acceleration acquisition unit to acquire the acceleration of the aircraft <NUM>, an attitude acquisition unit to acquire the attitude of the aircraft <NUM>, and an inertial velocity acquisition unit to acquire the inertial velocity of the aircraft <NUM>.

The position control system <NUM> also includes an altitude sensor <NUM> that detects the altitude of the aircraft <NUM> from the ground or a water surface as illustrated in <FIG>. The altitude sensor <NUM> is, for example, a laser altimeter, which measures relative altitude Δh (see <FIG>) from the aircraft <NUM> to the target landing point <NUM>. The altitude sensor <NUM> may be either a radio altimeter or a barometric altimeter. These altimeters may be used in combination as appropriate in accordance with the environment in which they are used, i.e., to measure the altitude from the ground surface and the altitude from the sea level. The altitude sensor <NUM> outputs the detected relative altitude Δh of the aircraft <NUM> to the control unit <NUM>. The altitude sensor <NUM> may measure the altitude of the aircraft <NUM> and output the measured altitude to the control unit <NUM>. The control unit <NUM> may cause a guidance calculation unit <NUM> described below to calculate the relative altitude Δh to the target landing point <NUM>, based on the altitude of the aircraft <NUM>. The position control system <NUM> may cause, instead of the altitude sensor <NUM>, an image processing unit <NUM>, which is described below, to calculate the relative altitude Δh between the aircraft <NUM> and the vessel <NUM> by applying image processing to an image including the marker <NUM> photographed by the camera <NUM>.

The control unit <NUM> includes the image processing unit <NUM>, the guidance calculation unit <NUM>, and a flight control unit <NUM>. The control unit <NUM> includes a photographing control unit, which is not illustrated, that controls the photographing direction of the camera <NUM> via a gimbal, which is not illustrated, that is installed in the aircraft <NUM>. In this embodiment, the camera <NUM> is adjusted so that the photographing range B of the camera <NUM> faces right downward in the vertical direction as described above.

The image processing unit <NUM> applies image processing to an image photographed by the camera <NUM> to calculate a center (Cx, Cy) (see <FIG>) of the marker <NUM>, that is, the target landing point <NUM>. The center (Cx, Cy) here is a point of coordinates in a camera fixed coordinate system whose origin is the center of the image photographed by the camera <NUM>, and can be calculated based on the number of pixels from the center of the image. Specifically, as illustrated in <FIG>, the image processing unit <NUM> specifies two diagonal lines Ld that extend between the corners of the marker <NUM> by image processing, and sets the intersection of the two specified diagonal lines Ld as the center (Cx, Cy) of the marker <NUM>. The target landing point <NUM> is not limited to the center of the marker <NUM> (Cx, Cy), and may be any of the four corners of the marker <NUM> or offset from the center of the marker <NUM>. The image processing unit <NUM> outputs the calculated center (Cx, Cy) of the marker <NUM> to the guidance calculation unit <NUM>.

The image processing unit <NUM> may also calculate a bow azimuth of the vessel <NUM> by applying image processing to the image including the marker <NUM> photographed by the camera <NUM> to specify the direction of the marker <NUM> and mapping the direction to the nose azimuth of the aircraft <NUM>, which is acquired by the navigation device <NUM>. The image processing unit <NUM> may calculate the relative altitude Δh between the aircraft <NUM> and the vessel <NUM> by applying image processing to the image including the marker <NUM> photographed by the camera <NUM> as described above.

The guidance calculation unit <NUM> calculates the control quantity of the aircraft <NUM> to guide the aircraft <NUM> to the target landing point <NUM>. The control quantity is the control quantity for adjusting the aircraft velocity, the attitude angle, the rate of change of the attitude angle, etc. of the aircraft <NUM>. The guidance calculation unit <NUM> calculates relative coordinates between the aircraft <NUM> and the target landing point <NUM> in order to calculate the control quantity. Specifically, the guidance calculation unit <NUM> calculates the relative position (X, Y) between the aircraft <NUM> and the target landing point <NUM> as the relative coordinates, the relative altitude Δh between the aircraft <NUM> and the target landing point <NUM>, the relative velocity between the aircraft <NUM> and the target landing point <NUM>, and the like. The relative position (X, Y) is the distance between the aircraft <NUM> and the target landing point <NUM> in the horizontal direction. The relative altitude Δh is the distance between the aircraft <NUM> and the target landing point <NUM> in the vertical direction.

The guidance calculation unit <NUM> calculates the relative position (X, Y) between the aircraft <NUM> and the target landing point <NUM>, based on the center (Cx, Cy) of the marker <NUM> calculated by the image processing unit <NUM>, the azimuth of the camera <NUM>, i.e., the nose azimuth of the aircraft <NUM>, and the altitude of the aircraft <NUM> (relative altitude Δh to the target landing point <NUM>). In this embodiment, the azimuth of the camera <NUM> is aligned with the nose azimuth of the aircraft <NUM>, but the embodiment is not limited to this example and the azimuth of the camera <NUM> does not have to be aligned with the nose azimuth of the aircraft <NUM>. Thus, the image processing unit <NUM> and the guidance calculation unit <NUM> function as a relative position acquisition unit (relative coordinate acquisition unit) to acquire the relative position between the aircraft <NUM> and the target landing point <NUM>.

The guidance calculation unit <NUM> calculates the relative altitude Δh to the target landing point <NUM>, based on the altitude of the aircraft <NUM> detected by the altitude sensor <NUM>. Therefore, the altitude sensor <NUM> and the guidance calculation unit <NUM> function as a relative altitude acquisition unit (relative coordinate acquisition unit) to acquire the relative altitude Δh between the aircraft <NUM> and the target landing point <NUM>. When the relative altitude Δh between the aircraft <NUM> and the vessel <NUM> is calculated by applying image processing to the image including the marker <NUM> photographed by the camera <NUM> in the image processing unit <NUM>, the image processing unit <NUM> serves as the relative altitude acquisition unit.

The guidance calculation unit <NUM> also calculates the relative velocity between the aircraft <NUM> and the target landing point <NUM>. Therefore, the guidance calculation unit <NUM> functions as the relative velocity acquisition unit to acquire the relative velocity between the aircraft <NUM> and the target landing point <NUM>. More specifically, the guidance calculation unit <NUM> executes a relative velocity calculation process of calculating relative velocity (ΔVx, ΔVy) between the aircraft <NUM> and the target landing point <NUM> on the basis of the relative position (X, Y) and the aircraft velocity (Vx, Vy). Therefore, the guidance calculation unit <NUM> functions as the relative velocity acquisition unit to acquire the relative velocity (ΔVx, ΔVy) between the aircraft <NUM> and the target landing point <NUM>.

The guidance calculation unit <NUM> then calculates the control quantity by feedback control (for example, PID control), based on the relative position (X, Y), the relative altitude Δh, the relative velocity (ΔVx, ΔVy), and the aircraft acceleration. The feedback control is not limited to PID control, and may be P control, PI control, PD control, or the like. The guidance calculation unit <NUM> outputs a calculated control quantity C' (see <FIG> and <FIG>) to the flight control unit <NUM>.

The flight control unit <NUM> controls each component of the aircraft <NUM> to fly the aircraft <NUM> according to the control quantity calculated by the guidance calculation unit <NUM>. The flight control unit <NUM> controls the blade pitch angle, rotation velocity, etc. of each rotor blade according to the control quantity, so as to adjust the aircraft velocity, attitude angle, rate of change of the attitude angle, etc. of the aircraft <NUM>. The aircraft <NUM> is thereby guided to the target landing point <NUM>. Although the image processing unit <NUM> and the guidance calculation unit <NUM> are described in this embodiment as functional units separate from the flight control unit <NUM>, the flight control unit <NUM>, the image processing unit <NUM>, and the guidance calculation unit <NUM> may be an integral functional unit. In other words, the process in the image processing unit <NUM> and the guidance calculation unit <NUM> may be performed in the flight control unit <NUM>.

Next, the position control of the aircraft <NUM> according to this embodiment is described with reference to <FIG> and <FIG>. In the position control of the aircraft <NUM>, spatial stable hovering, which executes stable hovering in the space, is performed. With reference to <FIG>, the spatial stable hovering is described here. <FIG> is an explanatory diagram illustrating one example about the position control of the aircraft according to the present embodiment. A line L1 in <FIG> indicates a position of the target landing point <NUM>, which changes over time due to motion such as waves. As illustrated in <FIG>, the spatial stable hovering is the flight control to perform the hovering so that the relative altitude Δh and the relative position (X, Y) between the aircraft <NUM> and the target landing point <NUM> in which the displacement quantity due to the motion is smoothed (averaged) coincide with the target relative altitude and the target relative position even if the relative altitude Δh and the relative position (X, Y) between the aircraft <NUM> and the target landing point <NUM> change over time due to motion such as waves. When the relative position (X, Y) and the relative altitude Δh are measured using the altitude sensor <NUM>, etc., their values are instantaneous values that include mutual motion, so that changes due to motion or the like are also included. Therefore, a smoothing process is required to remove the changes due to motion or the like, and control needs to be performed on the smoothed position/altitude. In other words, in the spatial stable hovering, the flight of the aircraft <NUM> is controlled so that the difference between the smoothed (average) relative altitude Δh and the target relative altitude is zero, and additionally, the flight of the aircraft <NUM> is controlled so that the difference between the smoothed (average) relative position (X, Y) and the target relative position is zero.

<FIG> is a block diagram illustrating one example of calculating the smoothed relative altitude of the aircraft. <FIG> is a block diagram illustrating one example of calculating the smoothed relative position of the aircraft. In the position control of the aircraft <NUM>, the guidance calculation unit <NUM> performs the position control about the spatial stable hovering, which keeps the aircraft at the target coordinates in the space with respect to the target landing point <NUM>, based on the block diagrams in <FIG> and <FIG>. For this reason, the guidance calculation unit <NUM> calculates the control quantity of the aircraft <NUM> to execute the spatial stable hovering.

In <FIG>, the position control about the spatial stable hovering is performed so that the difference between the smoothed relative altitude and the target relative altitude is zero. The guidance calculation unit <NUM> includes an acceleration correction processing unit <NUM>, a complementary filter <NUM>, a smoothing processing unit <NUM>, and a feedback control unit <NUM>. The acceleration correction processing unit <NUM>, the complementary filter <NUM>, and the smoothing processing unit <NUM> may be implemented by the guidance calculation unit <NUM>, by a processing unit separate from the guidance calculation unit <NUM>, or by a combination thereof, without particular limitations.

The acceleration correction processing unit <NUM> outputs first attitude correction acceleration in which the acceleration of the aircraft <NUM> is corrected, based on the acceleration of the aircraft <NUM> and the attitude of the aircraft <NUM>. The attitude correction acceleration is the acceleration in an aircraft axis coordinate system converted to acceleration in an inertial space coordinate system by a coordinate conversion on the basis of the attitude angle of the aircraft <NUM>. Specifically, to the acceleration correction processing unit <NUM>, the acceleration of the aircraft <NUM> acquired in the navigation device <NUM> is input, and the attitude angle of the aircraft <NUM> acquired in the navigation device <NUM> is also input. The acceleration to be input includes longitudinal (front-rear direction in the aircraft coordinate system), lateral (left-right direction in the aircraft coordinate system), and vertical (up-down direction in the aircraft coordinate system) acceleration. The attitude angles to be input include the attitude angles in the pitch axis, the roll axis, and the yaw axis. Upon the input of the acceleration of the aircraft <NUM> and the attitude angle of the aircraft <NUM>, the acceleration correction processing unit <NUM> calculates the first attitude correction acceleration in which the acceleration of the aircraft <NUM> in the vertical direction is corrected. The acceleration correction processing unit <NUM> outputs the calculated first attitude correction acceleration to the complementary filter <NUM>.

The complementary filter <NUM> outputs second attitude correction acceleration in which the drift component included in the first attitude correction acceleration is removed, based on the first attitude correction acceleration and the inertial velocity of the aircraft <NUM>. Specifically, to the complementary filter <NUM>, the first attitude correction acceleration output from the acceleration correction processing unit <NUM> is input, and the inertial velocity of the aircraft <NUM> in the vertical direction acquired in the navigation device <NUM> is also input. The inertial velocity to be input is the velocity in the aircraft inertial system with the aircraft <NUM> as the origin. Upon the input of the first attitude correction acceleration and the inertial velocity of the aircraft <NUM>, the complementary filter <NUM> combines a high-frequency component in which a low-frequency component included in the output value of the first attitude correction acceleration is removed, and a low-frequency component included in a derivative of the inertial velocity, and outputs the second attitude correction acceleration in the vertical direction.

<FIG> is a block diagram illustrating one example of the process in the complementary filter. As illustrated in <FIG>, the complementary filter <NUM> includes a pseudo-differential filter <NUM>, a low-pass filter <NUM>, and a high-pass filter <NUM>. The pseudo-differential filter <NUM> is a filter that applies pseudo-differentiation to the inertial velocity to obtain an acceleration component. A transfer function G1(S) of the pseudo-differential filter <NUM> is expressed by the following Equation (<NUM>). In Equation (<NUM>), "s" is the operator and "τ1" is the time constant.

The low-pass filter <NUM> is a filter that attenuates frequencies of a predetermined cutoff frequency or more for the derivative of the inertial velocity input from the pseudo-differential filter <NUM>. A transfer function G2(s) of the low-pass filter <NUM> is expressed by Equation (<NUM>). In Equation (<NUM>), "s" is the operator and "τ2" is the time constant. Therefore, the predetermined cutoff frequency is "<NUM>/τ2". Thus, by passing the derivative of the inertial velocity through the low-pass filter <NUM>, the low-frequency component included in the derivative of the inertial velocity can be obtained.

The high-pass filter <NUM> is a filter that attenuates the frequencies below the predetermined cutoff frequency for the first attitude correction acceleration. A transfer function G3(s) of the high-pass filter <NUM> is expressed by Equation (<NUM>). In Equation (<NUM>), "s" is the operator and "τ2" is the time constant common to the low-pass filter <NUM>. Therefore, the predetermined cutoff frequency is "<NUM>/τ2" in the high-pass filter <NUM> as well. Thus, by passing the first attitude correction acceleration in the vertical direction through the high-pass filter <NUM>, the high-frequency component in the first attitude correction acceleration can be obtained.

The complementary filter <NUM> then combines the low-frequency component included in the derivative of the inertial velocity output from the low-pass filter <NUM> and the high-frequency component included in the first attitude correction acceleration output from the high-pass filter <NUM> to calculate the second attitude correction acceleration. In this way, the complementary filter <NUM> removes the low-frequency component as the drift component included in the first attitude correction acceleration. After this, the complementary filter <NUM> outputs the calculated second attitude correction acceleration to the smoothing processing unit <NUM>.

The smoothing processing unit <NUM> performs the process for the flight control so that the relative altitude Δh between the aircraft <NUM> and the target landing point <NUM>, in which the displacement quantity due to motion is smoothed (averaged), coincides with the target relative altitude, even when the target landing point <NUM> changes due to motion. The smoothing processing unit <NUM> outputs the smoothed relative altitude Δh in which the relative altitude Δh is smoothed, based on the second attitude correction acceleration and the relative altitude Δh. Specifically, to the smoothing processing unit <NUM>, the second attitude correction acceleration output from the complementary filter <NUM> is input and the relative altitude Δh calculated by the guidance calculation unit <NUM> is also input. Upon the input of the second attitude correction acceleration and the relative altitude Δh, the smoothing processing unit <NUM> calculates the smoothed relative altitude Δh to make the relative altitude Δh between the aircraft <NUM> and the target landing point <NUM> in which the displacement quantity due to the motion is smoothed (averaged) coincide with the target relative altitude. The smoothing processing unit <NUM> then outputs the calculated smoothed relative altitude Δh to a subtraction circuit unit <NUM>. The subtraction circuit unit acquires the target relative altitude and calculates the difference between the acquired target relative altitude and the smoothed relative altitude Δh input from the smoothing processing unit <NUM>. The subtraction circuit unit <NUM> then outputs the calculated difference to the feedback control unit <NUM>.

The feedback control unit <NUM> acquires the inertial velocity of the aircraft <NUM> and calculates the control quantity C' on the basis of the acquired inertial velocity of the aircraft <NUM> and the difference between the target relative altitude and the smoothed relative altitude Δh input from the subtraction circuit unit <NUM>. The feedback control unit <NUM> then outputs the calculated control quantity C' to the flight control unit <NUM>. The flight control unit <NUM> executes the flight control on the basis of the control quantity C'. As an example of the flight control by the flight control unit <NUM>, when the aircraft <NUM> is a helicopter, the flight control is performed to change the angle of the collective pitch angle of the rotating blades on the helicopter.

Here, with reference to <FIG>, the signal processing related to the relative altitude in the smoothing processing unit <NUM> is described. <FIG> is a Bode diagram from the relative altitude to the smoothed relative altitude of the smoothing processing unit. In <FIG>, its horizontal axis expresses frequency, and its vertical axis expresses a gain (dB) and a phase (deg). In <FIG>, the frequency band where motion occurs is a motion frequency band f, and the frequency band that includes the motion frequency band f and is higher than or equal to the motion frequency band f is a frequency band f1 (first frequency band) and the frequency band that is lower than the motion frequency band f1 is a frequency band f2 (second frequency band).

The smoothing processing unit <NUM> amplifies the output signal by applying a gain to the input signal. The input signal is the relative altitude Δh and the output signal is the smoothed relative altitude Δh. As illustrated in <FIG>, the frequency band f1 has a sufficiently low gain compared to the frequency band f2, which reduces the effect of the motion in the frequency band f1. The frequency band f2 has a large damping coefficient (ζ > <NUM>) so that the low-frequency disturbance response is not oscillatory. Overshoot is suppressed as the value of the damping factor ζ is increased, and desirable results are obtained if ζ > <NUM>.

In <FIG>, the position control about the spatial stable hovering is performed so that the difference between the smoothed relative position and the target relative position is zero. The guidance calculation unit <NUM> includes the acceleration correction processing unit <NUM>, the complementary filter <NUM>, the smoothing processing unit <NUM>, the feedback control unit <NUM>, a Kalman filter <NUM>, and a relative velocity estimation processing unit <NUM>. The acceleration correction processing unit <NUM>, the complementary filter <NUM>, the smoothing processing unit <NUM>, the Kalman filter <NUM>, and the relative velocity estimation processing unit <NUM> may be implemented by the guidance calculation unit <NUM>, by a processing unit separate from the guidance calculation unit <NUM>, or by a combination thereof, without particular limitations. In <FIG>, the component in the X direction, which is the direction of the pitch axis, and the component in the Y direction, which is the direction of the roll axis, are expressed together, and the control quantity for each component is calculated by the guidance calculation unit <NUM>.

In a manner similar to the above, the acceleration correction processing unit <NUM> outputs the first attitude correction acceleration corresponding to the corrected acceleration of the aircraft <NUM> on the basis of the acceleration of the aircraft <NUM> and the attitude of the aircraft <NUM>. Upon the input of the acceleration of the aircraft <NUM> and the attitude angle of the aircraft <NUM>, the acceleration correction processing unit <NUM> calculates the first attitude correction acceleration corresponding to the corrected acceleration of the aircraft <NUM> in the longitudinal direction and the lateral direction. The acceleration correction processing unit <NUM> outputs the calculated first attitude correction acceleration to the complementary filter <NUM>.

In a manner similar to the above, the complementary filter <NUM> outputs the second attitude correction acceleration in which the drift component included in the first attitude correction acceleration is removed, based on the first attitude correction acceleration and the inertial velocity of the aircraft <NUM>. Specifically, to the complementary filter <NUM>, the first attitude correction acceleration output from the acceleration correction processing unit <NUM> is input, and the inertial velocity of the aircraft <NUM> in the longitudinal direction and the lateral direction acquired in the navigation device <NUM> is also input. Upon the input of the first attitude correction acceleration and the inertial velocity of the aircraft <NUM>, the complementary filter <NUM> combines the high-frequency component in which the low-frequency component included in the output value of the first attitude correction acceleration is removed, and the low-frequency component included in the derivative of the inertial velocity, and outputs the second attitude correction acceleration in the longitudinal direction and the lateral direction. The complementary filter <NUM> outputs the second attitude correction acceleration to the smoothing processing unit <NUM> and the feedback control unit <NUM>. The complementary filter <NUM> is similar to that in <FIG>.

The Kalman filter <NUM> performs an estimation based on the relative position (X, Y) and outputs the estimated relative position (X, Y) after the estimation. Specifically, the relative position (X, Y) calculated by the guidance calculation unit <NUM> is input to the Kalman filter <NUM>. Upon the input of the relative position (X, Y), the Kalman filter <NUM> calculates the estimated relative position (X, Y) by estimating the change of the relative position (X, Y) over time. The Kalman filter <NUM> outputs the calculated estimated relative position (X, Y) to the smoothing processing unit <NUM>.

The smoothing processing unit <NUM> outputs a smoothed relative position (X, Y), in which the estimated relative position (X, Y) is smoothed, on the basis of the second attitude correction acceleration and the estimated relative position (X, Y). Specifically, to the smoothing processing unit <NUM>, the second attitude correction acceleration output from the complementary filter <NUM> is input, and the estimated relative position (X, Y) calculated by the Kalman filter <NUM> is also input. Upon the input of the second attitude correction acceleration and the estimated relative position (X, Y), the smoothing processing unit <NUM> calculates the smoothed relative position (X, Y) to make the relative position between the aircraft <NUM> and the target landing point <NUM> in which the displacement quantity due to motion is smoothed (averaged) coincide with the target relative position. The smoothing processing unit <NUM> then outputs the calculated smoothed relative position (X, Y) to the feedback control unit <NUM>.

The relative velocity estimation processing unit <NUM> outputs the estimated relative velocity on the basis of the relative position (X, Y) and the aircraft velocity of the aircraft <NUM>. Specifically, to the relative velocity estimation processing unit <NUM>, the relative position (X, Y) calculated by the guidance calculation unit <NUM> is input, and the aircraft velocity of the aircraft <NUM> in the longitudinal direction and the lateral direction acquired by the navigation device <NUM> is also input. The relative velocity estimation processing unit <NUM> estimates the relative velocity from the relative position (X, Y) and the aircraft velocity that are input, and outputs the estimated relative velocity to the feedback control unit <NUM>.

The feedback control unit <NUM> calculates the control quantity C' on the basis of the second attitude correction acceleration input from the complementary filter <NUM>, the smoothed relative position (X, Y) input from the smoothing processing unit <NUM>, and the estimated relative velocity input from the relative velocity estimation processing unit <NUM>. The feedback control unit <NUM> then outputs the calculated control quantity C' to the flight control unit <NUM>. The flight control unit <NUM> executes the flight control on the basis of the control quantity C'. As an example of the flight control by the flight control unit <NUM>, when the aircraft <NUM> is a helicopter, the flight control is performed by tilting the helicopter's main rotor in the longitudinal direction and the lateral direction.

With reference to <FIG>, the signal processing related to the relative position in the smoothing processing unit <NUM> is described. <FIG> is a Bode diagram from the relative position to the smoothed relative position of the smoothing processing unit. In <FIG>, its horizontal axis expresses frequency, and its vertical axis expresses a gain (dB) and a phase (deg). In <FIG>, in a manner similar to <FIG>, the frequency band where motion occurs is the motion frequency band f, and the frequency band that includes the motion frequency band f and is higher than or equal to the motion frequency band f is the frequency band f1 (first frequency band) and the frequency band that is lower than the motion frequency band f1 is the frequency band f2 (second frequency band).

The smoothing processing unit <NUM> amplifies the output signal by applying a gain to the input signal. The input signal is the estimated relative position (X, Y) and the output signal is the smoothed relative position (X, Y). As illustrated in <FIG>, the frequency band f1 has a sufficiently low gain compared to the frequency band f2, similar to <FIG>, which reduces the effect of the motion in the frequency band f1. The frequency band f2 has a large damping coefficient (ζ > <NUM>), similar to the relative altitude, so that the low-frequency disturbance response is not oscillatory. Overshoot is suppressed as the value of the damping factor ζ is increased, and desirable results are obtained if ζ > <NUM>.

Next, with reference to <FIG>, a position control method using the position control system <NUM> for the aircraft <NUM> according to the present embodiment, specifically, the smoothing process until the smoothed relative altitude Δh and the smoothed relative position (X, Y) are calculated in the smoothing processing unit <NUM> is described. <FIG> is a flowchart expressing one example about the position control method for the aircraft according to the present embodiment. <FIG> illustrates the smoothing process about the relative altitude and the relative position, and also the smoothing process until the smoothed relative altitude Δh and the smoothed relative position (X, Y) are output from the smoothing processing unit <NUM>.

In the smoothing process, the acceleration correction processing unit <NUM> first outputs the first attitude correction acceleration on the basis of the acceleration of the aircraft <NUM> and the attitude of the aircraft <NUM> that are input (step S1). At step S1, for the position control about the relative altitude Δh, the acceleration correction processing unit <NUM> outputs the first attitude correction acceleration in the vertical direction. At step S1, in the case of the position control regarding the relative position (X, Y), the acceleration correction processing unit <NUM> outputs the first attitude correction acceleration in the longitudinal direction and the lateral direction.

Then, in the position control method for the aircraft <NUM>, the complementary filter <NUM> outputs the second attitude correction acceleration on the basis of the first attitude correction acceleration and the inertial velocity of the aircraft <NUM> (step S2). At step S2, the drift component included in the first attitude correction acceleration is removed by removing the low-frequency component of the first attitude correction acceleration. At step S2, in the case of the position control about the relative altitude Δh, the complementary filter <NUM> outputs the second attitude correction acceleration in the vertical direction. At step S2, in the case of the position control about the relative position (X, Y), the complementary filter <NUM> outputs the second attitude correction acceleration in the longitudinal direction and the lateral direction.

After that, in the position control method for the aircraft <NUM>, the smoothing processing unit <NUM> outputs the smoothed relative altitude Δh and the smoothed relative position (X, Y) on the basis of the second attitude correction acceleration in the vertical direction and the relative altitude Δh and the second attitude correction acceleration in the longitudinal direction and the lateral direction and the estimated relative position (X, Y) (step S3).

As described above, the position control system <NUM> for the aircraft <NUM>, the aircraft <NUM>, and the position control method for the aircraft <NUM> according to the embodiment are understood as follows, for example.

The position control system <NUM> for the aircraft <NUM> according to a first aspect is the position control system <NUM> for the aircraft <NUM> that keeps the aircraft <NUM> at target coordinates in an inertial space with respect to the target landing point <NUM> that moves, and includes: the acceleration correction processing unit <NUM> that, based on acceleration of the aircraft <NUM> and an attitude of the aircraft <NUM>, outputs first attitude correction acceleration for correcting the acceleration of the aircraft <NUM>; the complementary filter <NUM> that, based on the first attitude correction acceleration and inertial velocity of the aircraft <NUM>, outputs second attitude correction acceleration in which a drift component included in the first attitude correction acceleration is removed; and the smoothing processing unit <NUM> that, based on the second attitude correction acceleration and relative coordinates between the aircraft <NUM> and the target landing point <NUM>, outputs smoothed relative coordinates obtained by smoothing the relative coordinates.

With this configuration, the drift component included in the first attitude correction acceleration output based on the acquired acceleration of the aircraft <NUM> can be removed in the complementary filter <NUM>. Therefore, since the smoothed relative coordinates can be acquired using the second attitude correction acceleration from which the drift component has been removed, the position of the aircraft <NUM> in the space can be stably kept even when the target landing point <NUM> moves.

In a second aspect, the complementary filter <NUM> combines a high-frequency component in which a low-frequency component included in an output value of the first attitude correction acceleration is removed, and the low-frequency component included in a derivative of the inertial velocity, and outputs the second attitude correction acceleration.

With this configuration, the drift component can be suitably removed by removing the low-frequency component included in the output value of the first attitude correction acceleration. The removed low-frequency component can be suitably interpolated by the inertial velocity.

In a third aspect, the relative coordinates include at least one of relative altitude Δh between the aircraft and the target landing point in the vertical direction and a relative position (X, Y) between the aircraft and the target landing point in the horizontal plane.

With this configuration, the position of the aircraft <NUM> can be kept stable at the relative altitude Δh and the relative position (X, Y).

In a fourth aspect, when the relative coordinates include the relative position (X, Y), the smoothing processing unit <NUM> outputs a smoothed relative position as the smoothed relative coordinates, the position control system <NUM> further includes the relative velocity estimation processing unit <NUM> that, based on aircraft velocity of the aircraft <NUM> and the relative position (X, Y), outputs estimated relative velocity obtained by estimating relative velocity between the aircraft <NUM> and the target landing point <NUM>, and the flight of the aircraft <NUM> is controlled based on the estimated relative velocity and the smoothed relative position.

With this configuration, the position control for the target landing point <NUM> can be performed with high accuracy at the relative position (X, Y).

In a fifth aspect, the position control system <NUM> further includes the Kalman filter <NUM> that, when the relative coordinates include the relative position (X, Y), outputs an estimated relative position estimated based on the relative position (X, Y), and the smoothing processing unit <NUM> outputs the smoothed relative position as the smoothed relative coordinates, based on the second attitude correction acceleration and the estimated relative position.

With this configuration, the smoothed relative position with high accuracy can be output at the relative position (X, Y). Thus, the position of the aircraft <NUM> can be kept stable with high accuracy at the relative position (X, Y).

In a sixth aspect, the smoothing processing unit outputs the smoothed relative coordinates after applying a gain to the relative coordinates, when a frequency band that is higher than or equal to a motion frequency band in which the motion occurs is a first frequency band and a frequency band that is lower than the motion frequency band is a second frequency band, the gain in the first frequency band is lower than the gain in the second frequency band, and a damping coefficient ζ in the second frequency band satisfies ζ > <NUM>.

With this configuration, the first frequency band f1 has a sufficiently low gain compared to the second frequency band, which reduces the effect of the motion in the first frequency band f1. The second frequency band f2 has a larger damping coefficient (ζ > <NUM>), so that the low-frequency disturbance response is not oscillatory.

The aircraft <NUM> according to a seventh aspect includes an acceleration acquisition unit that acquires the acceleration of the aircraft, an attitude acquisition unit that acquires an attitude of the aircraft, an inertial velocity acquisition unit that acquires inertial velocity of the aircraft, a relative coordinate acquisition unit that acquires relative coordinates of the aircraft, and the aforementioned position control system <NUM> for the aircraft <NUM>.

With this configuration, it is possible to provide the aircraft <NUM> with the flight controlled so that its position with respect to the target landing point <NUM> is stable, even when the target landing point <NUM> moves.

A position control method for the aircraft <NUM> according to an eighth aspect is the position control method for the aircraft to keep the aircraft at predetermined coordinates in an inertial space with respect to a target landing point that moves, and includes: a step of, based on acceleration of the aircraft and an attitude of the aircraft, outputting first attitude correction acceleration for correcting the acceleration of the aircraft; a step of, based on the first attitude correction acceleration and inertial velocity of the aircraft, outputting second attitude correction acceleration in which a drift component included in the first attitude correction acceleration is removed; and a step of, based on the second attitude correction acceleration and relative coordinates between the aircraft and the target landing point, outputting smoothed relative coordinates obtained by smoothing the relative coordinates.

Claim 1:
An aircraft position control system (<NUM>) configured to keep an aircraft (<NUM>) at target coordinates in an inertial space with respect to a target landing point (<NUM>) that moves, the aircraft position control system (<NUM>) comprising:
an acceleration correction processing unit (<NUM>) configured to, based on acceleration of the aircraft (<NUM>) and an attitude of the aircraft (<NUM>), output first attitude correction acceleration in which the acceleration in an aircraft axis coordinate system is converted to acceleration in an inertial space coordinate system;
a complementary filter (<NUM>) configured to, based on the first attitude correction acceleration and inertial velocity of the aircraft (<NUM>), which is the velocity in an aircraft coordinate system with the aircraft (<NUM>) as the origin, output second attitude correction acceleration in which a drift component included in the first attitude correction acceleration is removed;
a smoothing processing unit (<NUM>) configured to, based on the second attitude correction acceleration and relative coordinates between the aircraft (<NUM>) and the target landing point (<NUM>), output smoothed relative coordinates obtained by smoothing the relative coordinates;
a feedback control unit (<NUM>) configured to calculate a control quantity of the aircraft (<NUM>) on the basis of the smoothed relative coordinates; and
a flight control unit (<NUM>) configured to execute the position control of the aircraft (<NUM>) on the basis of the control quantity.