Patent Description:
In the related art, there are known techniques for guiding an aircraft to a target point. For example, Patent Literature <NUM> discloses a navigation system that performs image processing on image information of a target point acquired by a stereo camera to determine a distance to the target point, calculates a relative position of the aircraft with respect to the target point based on the distance and the attitude angle of the aircraft, and generates navigation information using the relative position.

Patent Literature <NUM>: <CIT>
<CIT> is directed to methods of determining a vessel-relative off-deck waypoint (VRODW) location comprising the steps of providing an aircraft in flight; determining vessel range and vessel bearing relative the aircraft; and determining the VRODW location using the range and bearing measurements of the vessel. <NPL>), relates to a system architecture, including the structure of the Kalman filter for the estimation of the relative position and velocity between the quadcopter and the landing pad, as well as the controller design for the full rendezvous and landing maneuvers. <CIT> discloses a control device and method for integrated unmanned aerial vehicle (UAV) flight. The control device comprises a data link communication unit, an acquisition unit, a processing unit and a control unit. The data link communication unit is used for interacting with a ground station, acquiring a flight task of a UAV and monitoring the flight state of the UAV; the acquisition unit is used for acquiring current flight acceleration and angular speed of the UAV, static air pressure on the UAV, motion speed of the UAV relative to surrounding air, magnetic field intensity of geomagnetic field on the UAV and positioning information of the UAV; the processing unit is used for processing data acquired by the acquisition unit to acquire the flight attitude of the UAV, and comparing the flight attitude with the flight task received by the data link communication unit so as to give flight height, flight speed and flight direction which need to be adjusted by the UAV; and the control unit is used for controlling a steering engine controller of the UAV to adjust according to the flight height, flight speed and flight direction which need to be adjusted by the UAV and are given by the processing unit, so that the UAV meets the requirement of the flight task. The device and the method can effectively and quickly control the flight of the UAV.

The navigation system described in Patent Literature <NUM> described above causes the aircraft to descend while controlling the position, altitude, rate of descent or the like of the aircraft based on the navigation information generated using the relative position, so as to cause the aircraft to land on the landing point. Here, the landing point may move in a case where the landing point is a movable body such as a marine vessel, for example. In this case, the navigation system described in Patent Literature <NUM> generates navigation information using the relative position, whereby it can grasp the relative position between the aircraft and the landing point, as well as grasping the relative velocity with respect to a fixed landing point. However, Patent Literature <NUM> does not allow for appropriate control of the aircraft based on the navigation information in a case where the landing point is moving, and may fail to accurately and quickly control the position of the aircraft such as, for example, the aircraft taking an excessive time to approach the target landing point, or passing over the target landing point.

It is an object of the present invention, which has been made in consideration of the foregoing description, to move the aircraft more accurately and faster toward a target landing point moving relatively thereto.

Further preferred embodiments may be derived from the dependent claims.

In order to solve the problems described above and achieve the object, an aircraft position control system according to the present invention includes: a relative position acquisition unit configured to acquire a relative position between an aircraft and a target landing point; a relative velocity acquisition unit configured to acquire relative velocity of the aircraft with respect to the target landing point; and a control unit configured to control the aircraft, wherein the control unit has: a feedback control unit configured to calculate a feedback manipulated variable of the aircraft by feedback control so that the aircraft heads toward the target landing point, based on at least the relative position and the relative velocity; a multi-value control unit configured to set, by referring to a switching line preliminarily provided in a manner passing through an origin of a coordinate plane whose orthogonal axes represent the relative position and the relative velocity and separating an acceleration region in which the relative velocity is increased and a deceleration region in which the relative velocity is decreased, an addition value that tends to increase the relative velocity when a coordinate point of the current relative position and the current relative velocity is located in the acceleration region, or set an addition value that tends to decrease the relative velocity when the coordinate point is located in the deceleration region with respect to the switching line; and an addition circuit configured to calculate a manipulated variable of the aircraft by adding the addition value to the feedback manipulated variable.

In order to solve the problems described above and achieve the object, an aircraft position control system according to the present invention includes: a relative position acquisition unit configured to acquire a relative position between an aircraft and a target landing point; a relative velocity acquisition unit configured to acquire relative velocity between the aircraft and the target landing point; and a control unit configured to control the aircraft to head toward the target landing point, based on at least the relative position and the relative velocity, wherein the relative velocity acquisition unit calculates the relative velocity by adding a value acquired by applying, to a differential value of the relative position, a low-pass filter that attenuates frequencies equal to or higher than a cutoff frequency being predetermined, and a value acquired by applying, to the velocity of the aircraft, a high-pass filter that attenuates frequencies below the cutoff frequency.

In order to solve the problems described above and achieve the object, an aircraft according to the present invention includes the aircraft position control system.

In order to solve the problems described above and achieve an object, an aircraft position control method according to the present invention comprising: acquiring a relative position between an aircraft and a target landing point; acquiring relative velocity of the aircraft with respect to the target landing point; calculating a feedback manipulated variable of the aircraft by feedback control so that the aircraft heads toward the target landing point, based on at least the relative position and the relative velocity; setting, by referring to a switching line preliminarily provided in a manner passing through an origin of a coordinate plane whose orthogonal axes represent the relative position and the relative velocity and separating an acceleration region in which the relative velocity is increased and a deceleration region in which the relative velocity is decreased, an addition value that tends to increase the relative velocity when a coordinate point of the current relative position and the current relative velocity is located in the acceleration region, or setting an addition value that tends to decrease the relative velocity when the coordinate point is located in the deceleration region with respect to the switching line; and calculating a manipulated variable of the aircraft by adding the addition value to the feedback manipulated variable.

The aircraft position control system, the aircraft, and the aircraft position control method according to the present invention exhibits an effect that allows for moving the aircraft more accurately and faster toward a relatively moving target landing point.

In the following, detailed description of embodiments of an aircraft position control system, an aircraft, and an aircraft position control method according to the present invention will be provided, based on drawings. Note that, the invention is not limited to the embodiments.

<FIG> is a schematic configuration diagram illustrating an example of an aircraft position control system according to a first embodiment, and <FIG> is an explanatory diagram illustrating an aircraft according to the first embodiment heading toward a target landing point. An aircraft <NUM> according to the first embodiment is a flying object as a rotorcraft (e.g., helicopter, drone, etc.). In the present embodiment, the aircraft <NUM> is an unmanned aerial vehicle.

Here, the aircraft <NUM> may be any flying object that can fly forward, fly backward, turn, fly sideways, or hover, and may be a manned aerial vehicle. The aircraft <NUM> equipped with a position control system <NUM> is flight-controlled by the position control system <NUM> so as to land on a target landing point <NUM> illustrated in <FIG>.

In the present embodiment, the target landing point <NUM> is provided on a marine vessel <NUM>, as illustrated in <FIG>. Accordingly, the aircraft <NUM> lands on (touches down) the marine vessel <NUM>, which is a movable body on the water. Here, although not illustrated, the marine vessel <NUM> has an arresting gear provided thereon for arresting the aircraft <NUM> when the aircraft <NUM> lands on the target landing point <NUM>. However, the target landing point <NUM> is not limited to the marine vessel <NUM> and may be provided on a vehicle or the like, which is a movable body on the ground, or may be provided on a fixed facility or the ground.

The target landing point <NUM> has a marker <NUM> provided thereon that allows the aircraft <NUM> to recognize the position of the target landing point <NUM>. <FIG> is an explanatory diagram illustrating an example of a marker provided on the target landing point. As illustrated, the marker <NUM> is a square AR marker color-coded with two colors, black and white, for example. Here, the marker <NUM> is not limited to an AR marker and may be any marker that can recognize the position of the target landing point <NUM> via image processing, such as, for example, an H mark or an R mark indicating a landing point in a heliport. In addition, the marine vessel <NUM> may have a plurality of markers of different shapes provided thereon as the marker <NUM>, and the aircraft <NUM> may be guided toward the target landing point <NUM> corresponding to any one of the different markers <NUM>.

An aircraft position control system <NUM> according to the first embodiment is a system that controls the position of the aircraft <NUM> in order to cause the aircraft <NUM> in flight to land on the target landing point <NUM>. The position control system <NUM> is mounted on the aircraft <NUM>. The position control system <NUM> includes, as illustrated in <FIG>, a camera <NUM>, a navigation system <NUM>, and a control unit <NUM>.

The camera <NUM> is an imaging device mounted on the aircraft <NUM> with a gimbal (not illustrated) interposed therebetween. The camera <NUM> may be a monocular camera, a compound-eye camera, an infrared camera or the like, provided that it can capture an image of the marker <NUM>. The camera <NUM> is provided in order to capture, from the aircraft <NUM>, an image of the marker <NUM> provided on the target landing point <NUM>. It is assumed that the camera <NUM> can adjust the shooting direction via a gimbal (not illustrated). In the present embodiment, the camera <NUM> is controlled by the control unit <NUM> so that its shooting range B (see <FIG> and <FIG>) faces directly downward in the vertical direction, for example. Here, the camera <NUM> may be controlled by the control unit <NUM> so that its shooting range B faces diagonally forward with respect to the vertical direction. In addition, the camera <NUM> may be devoid of a gimbal and fixed directly below the body of the aircraft <NUM> so that the shooting direction faces downward in the vertical direction.

The navigation system <NUM> is an inertial navigation system (INS), for example. Here, although the present embodiment is described applying an inertial navigation system as the navigation system <NUM>, it is not particularly limited thereto and any type of navigation system may be used as the navigation system <NUM>. In addition, the navigation system <NUM> serves as an inertial navigation system including the global positioning system (GPS) in order to improve the accuracy of position measurement. Although the present embodiment is described applying an inertial navigation system including the GPS, it is not particularly limited to the GPS and any position measurement unit that can accurately measure the position may be used, such as those using the Quasi-Zenith Satellite System, for example, and there may also be a configuration with the position measurement unit such as the GPS being omitted provided that the position can be accurately measured using only the navigation system <NUM>. The navigation system <NUM> including the GPS acquires attitude angles of the aircraft <NUM> in the roll, yaw and pitch directions, and aircraft velocity Vh (see <FIG>), aircraft acceleration a (see <FIG>), an aircraft heading ψh (see <FIG>), and position coordinates of the aircraft <NUM>. Here, the navigation system <NUM> may include an attitude angle sensor that detects the attitude angle of the aircraft <NUM>, a velocity detection sensor that detects the aircraft velocity Vh of the aircraft <NUM>, an acceleration detection sensor that detects the aircraft acceleration a of the aircraft <NUM>, and a sensor that detects the aircraft heading ψh of the aircraft <NUM>. The navigation system <NUM> outputs, to the control unit <NUM>, the attitude angle, the aircraft velocity Vh, the aircraft acceleration a, and the position coordinates of the aircraft <NUM>, all of which have been acquired.

In addition, the position control system <NUM> includes, as illustrated in <FIG>, an altitude sensor <NUM> that detects the altitude of the aircraft <NUM> from the ground level or the water surface. The altitude sensor <NUM> may use any of a laser altimeter, a radio altimeter, or a barometric altimeter, for example. In addition, these altimeters may also be applied in combination as appropriate, depending on the use environment, i.e., measurement of the altitude from the ground level or the altitude from the sea level. The altitude sensor <NUM> outputs a detected altitude of the aircraft <NUM> to the control unit <NUM>. However, the position control system <NUM> is not limited to the altitude sensor <NUM>, and may be one that calculates a relative altitude between the aircraft <NUM> and the marine vessel <NUM> by performing, in an image processing unit <NUM> described below, image processing on an image including the marker <NUM> captured by the camera <NUM>.

The control unit <NUM> includes the image processing unit <NUM>, a guidance calculation unit <NUM>, and a flight control unit <NUM>. Here, the control unit <NUM> includes an imaging control unit (not illustrated) that controls the shooting direction of the camera <NUM> via a gimbal (not illustrated) provided on the aircraft <NUM>. In the present embodiment, as described above, the camera <NUM> is adjusted so that the shooting range B faces directly downward in the vertical direction.

The image processing unit <NUM> performs image processing on an image captured by the camera <NUM> and calculates a center (Cx, Cy) of the marker <NUM>, i.e., the target landing point <NUM> (see <FIG>). Here, the center (Cx, Cy) is a coordinate point in a fixed camera coordinate system, with the center of the image captured by the camera <NUM> being an origin Oc (see <FIG>), and can be calculated from the number of pixels from the center of the image. The calculation method for the center (Cx, Cy) will be described below. The image processing unit <NUM> outputs the calculated center (Cx, Cy) of the marker <NUM> to the guidance calculation unit <NUM>. Here, the target landing point <NUM> is not limited to the center (Cx, Cy) of the marker <NUM>, and may be any of the four corners of the marker <NUM>, or may be a position offset from the center of the marker <NUM>.

In addition, the image processing unit <NUM> determines the orientation of the marker <NUM> by performing image processing on the image including the marker <NUM> captured by the camera <NUM>, and calculates a ship heading ψs (see <FIG>) of the marine vessel <NUM> by associating the orientation of the marker <NUM> with the aircraft heading ψh of the aircraft <NUM> acquired by the navigation system <NUM>. Here, as described above, the image processing unit <NUM> may calculate the relative altitude between the aircraft <NUM> and the marine vessel <NUM> by performing image processing on an image including the marker <NUM> captured by the camera <NUM>.

The guidance calculation unit <NUM> calculates a manipulated variable C' (see <FIG>) of the aircraft <NUM> for guiding the aircraft <NUM> toward the target landing point <NUM>. The manipulated variable C' is a manipulated variable for adjusting the aircraft velocity Vh, attitude angle, attitude rate, or the like, of the aircraft <NUM>.

Specifically, the guidance calculation unit <NUM> performs a relative position calculation process of calculating a relative position (Xhg, Yhg) (see <FIG>) 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 aircraft heading ψh, and the altitude of the aircraft <NUM>. Accordingly, the image processing unit <NUM> and the guidance calculation unit <NUM> function as a relative position acquisition unit that acquires the relative position (Xhg, Yhg) between the aircraft <NUM> and the target landing point <NUM>.

In addition, the guidance calculation unit <NUM> performs a relative velocity calculation process that calculates relative velocity (ΔVx, ΔVy) (see <FIG>) between the aircraft <NUM> and the target landing point <NUM> based on the relative position (Xhg, Yhg) and aircraft velocity (Vx, Vy). Accordingly, the guidance calculation unit <NUM> functions as a relative velocity acquisition unit that acquires the relative velocity (ΔVx, ΔVy) between the aircraft <NUM> and the target landing point <NUM>.

In addition, as described above, the guidance calculation unit <NUM> calculates the relative altitude with respect to the target landing point <NUM>, based on the altitude of the aircraft <NUM>. Therefore, the altitude sensor <NUM> and the guidance calculation unit <NUM> function as a relative altitude acquisition unit that acquires the relative altitude between the aircraft <NUM> and the target landing point <NUM>. Here, the image processing unit <NUM> serves as the relative altitude acquisition unit in a case where the relative altitude between the aircraft <NUM> and the marine vessel <NUM> is calculated in the image processing unit <NUM> by performing image processing on the image including the marker <NUM> captured by the camera <NUM>.

Subsequently, the guidance calculation unit <NUM>, while calculating a feedback manipulated variable C by feedback control (PID control), based on the relative position (Xhg, Yhg), the relative velocity (ΔVx, ΔVy), and aircraft acceleration (ax, ay), calculates the manipulated variable C' (see <FIG>) by adding an addition value D to the feedback manipulated variable C via multi-value control which will be described below. Here, feedback control is not limited to PID control and may be P control, PI control, PD control, or the like. Details of the calculation process in the aforementioned guidance calculation unit <NUM> will be described below. The guidance calculation unit <NUM> outputs the calculated manipulated variable C' to the flight control unit <NUM>.

The flight control unit <NUM> controls respective components of the aircraft <NUM> in accordance with the manipulated variable C' calculated by the guidance calculation unit <NUM> described below, and assists the aircraft <NUM> in flight. The flight control unit <NUM> controls the blade pitch angle, rotational speed or the like of respective rotary blades in accordance with the manipulated variable, and adjusts the aircraft velocity Vh, attitude angle, and attitude rate of the aircraft <NUM>. Accordingly, the aircraft <NUM> is guided toward the target landing point <NUM>. Note that, although the image processing unit <NUM> and the guidance calculation unit <NUM> are described as separate functional units from the flight control unit <NUM> in the present embodiment, the flight control unit <NUM>, the image processing unit <NUM>, and the guidance calculation unit <NUM> may be integrated as a single functional unit. In other words, the flight control unit <NUM> may substitutionally perform processes that have been supposed to be performed by the image processing unit <NUM> and the guidance calculation unit <NUM>.

Next, a procedure of calculating the manipulated variable C' of the aircraft <NUM> by the control unit <NUM> will be described as an aircraft position control method according to embodiments. <FIG> is a block diagram illustrating an example of a configuration of calculating the manipulated variable of the aircraft by the guidance calculation unit. According to the block diagram illustrated in <FIG>, the control unit <NUM> calculates the manipulated variable of the aircraft <NUM> for guiding the aircraft <NUM> toward the target landing point <NUM>. Here, <FIG> illustrates both the component in the X direction that serves as the direction of the pitch axis, and the component in the Y direction that serves as the direction of the roll axis, with the manipulated variable of each component being calculated by the guidance calculation unit <NUM>.

The control unit <NUM> performs a relative position calculation process of calculating the relative position (Xhg, Yhg) between the aircraft <NUM> and the target landing point <NUM>. The relative position calculation process is performed by the image processing unit <NUM> and the guidance calculation unit <NUM> according to the procedure illustrated in <FIG> is a flowchart illustrating an example of the relative position calculation process performed by the image processing unit and the guidance calculation unit. The flowchart illustrated in <FIG> is repeatedly performed by the image processing unit <NUM> and the guidance calculation unit <NUM> at a predetermined time interval. In addition, <FIG> is an explanatory diagram illustrating an aircraft being guided toward a target landing point. The following explanation describes calculation of the relative position between the aircraft <NUM> and the target landing point <NUM> in the horizontal direction, with the camera <NUM> having successfully captured the marker <NUM>. The relative altitude of the aircraft <NUM> with respect to the target landing point <NUM> is calculated based on the altitude of the aircraft <NUM> detected by the altitude sensor <NUM>, and is appropriately controlled in accordance with the relative position or the like between the aircraft <NUM> and the target landing point <NUM>. Additionally, in a case where the aircraft <NUM> and the marine vessel <NUM> are separated to an extent that the camera <NUM> cannot capture the marker <NUM>, the aircraft <NUM> is assisted to fly toward the marine vessel <NUM> using, for example, GPS-based position information of each other.

The control unit <NUM> acquires, in the image processing unit <NUM>, the image captured by the camera <NUM> (step S1). Next, the control unit <NUM> calculates, in the image processing unit <NUM>, the center (Cx, Cy) of the marker <NUM> in the fixed camera coordinate system (step S2). Specifically, as illustrated in <FIG>, the image processing unit <NUM> determines, by image processing, two diagonal lines Ld extending between corners of the marker <NUM>, and defines the intersection of the two determined diagonal lines Ld as the center (Cx, Cy) of the marker <NUM>. Here, the image processing unit <NUM> may determine only one diagonal line Ld, and may define the center position of the length of the determined diagonal line Ld as the center (Cx, Cy) of the marker <NUM>. In addition, the image processing unit <NUM> may determine two or more diagonal lines Ld, and define a position that is the average of the center position of the determined diagonal lines Ld as the center (Cx, Cy) of the marker <NUM>. Furthermore, the image processing unit <NUM> may, when performing keystone correction of the marker <NUM> having a square shape using a projective transformation-based function, calculate the center (Cx, Cy) of the square based on the function. On this occasion, keystone correction may be performed using coordinate points of the four corners of the marker <NUM>, or coordinate points of respective points along the black and white color-coded boundary of the marker <NUM>, and other coordinate points may be calculated by interpolation.

Next, the control unit <NUM> calculates, in the guidance calculation unit <NUM>, the relative position (Xhg, Yhg) between the aircraft <NUM> and the target landing point <NUM> based on the center (Cx, Cy) of the marker <NUM>, orientation of the camera <NUM>, i.e., aircraft heading ψh of the aircraft <NUM>, and altitude of the aircraft <NUM> (relative altitude with respect to the target landing point <NUM>) (step S3). The relative position (Xhg, Yhg) serves as the distance between the aircraft <NUM> and the target landing point <NUM> in the horizontal direction. Note that the process at step S3 may be performed by the image processing unit <NUM>. Specifically, the guidance calculation unit <NUM> first transforms the coordinates of the center (Cx, Cy) of the marker <NUM> calculated by the image processing unit <NUM> into a target coordinate point (Ximg, Yimg) in the fixed camera coordinate system.

Next, the guidance calculation unit <NUM> calculates a relative position (Xsg, Ysg) (see <FIG>) between the aircraft <NUM> and the target landing point <NUM> in a ship inertial reference frame SG (see <FIG>), based on the following Equations (<NUM>) and (<NUM>). The ship inertial reference frame SG is a coordinate system with the target landing point <NUM> being an origin Osg (<NUM>, <NUM>), the direction along the ship heading ψs of the marine vessel <NUM> being the X-axis, the direction orthogonal to the ship heading ψs in the horizontal direction being the Y axis, and the vertical direction being the Z-axis. <NUM>] <MAT>
[Math. <NUM>] <MAT>.

Next, the guidance calculation unit <NUM> calculates the relative position (Xhg, Yhg) (see <FIG>) between the aircraft <NUM> and the target landing point <NUM> in an aircraft inertial reference frame HG (see <FIG>), based on the following Equations (<NUM>) and (<NUM>). The aircraft inertial reference frame HG is a coordinate system with the aircraft <NUM> being the origin Ohg (<NUM>, <NUM>), the direction along the aircraft heading ψh of the aircraft <NUM> being the X-axis, the direction orthogonal to the aircraft heading ψh in the horizontal direction being the Y-axis, and the vertical direction being the Z-axis. Accordingly, the relative position (Xhg, Yhg) between the aircraft <NUM> and the target landing point <NUM> in the aircraft inertial system is calculated in the horizontal direction. The relative position (Xhg, Yhg) is the distance from the aircraft <NUM> to the target landing point <NUM>. <NUM>] <MAT>
[Math. <NUM>] <MAT>.

In addition, the control unit <NUM> performs, in the guidance calculation unit <NUM>, a relative velocity calculation process of calculating the relative velocity (ΔVx, ΔVy) of the aircraft <NUM> with respect to the target landing point <NUM>, i.e., the marine vessel <NUM>, in accordance with the block diagram illustrated in <FIG> is a block diagram illustrating an example of a relative velocity calculation process performed by the guidance calculation unit. Here, <FIG> also illustrates both the component in the X direction that serves as the direction of the pitch axis, and the component in the Y direction that serves as the direction of the roll axis. Specifically, the guidance calculation unit <NUM> calculates the relative velocity (ΔVx, ΔVy) based on the relative position (Xhg, Yhg) between the aircraft <NUM> and the target landing point <NUM>, and the aircraft velocity (Vx, Vy) of the aircraft <NUM> detected by the navigation system <NUM>.

The guidance calculation unit <NUM> first calculates, as illustrated in <FIG>, the relative velocity (ΔV1x, ΔV1y) between the aircraft <NUM> and the target landing point <NUM>, by differentiating the relative position (Xhg, Yhg) between the aircraft <NUM> and the target landing point <NUM>. In the present embodiment, the relative position (Xhg, Yhg) is subject to pseudo-differential operation using a pseudo-differential filter <NUM>. A transfer function G1(s) of the pseudo-differential filter is represented by the following Equation (<NUM>). In Equation (<NUM>), the symbol "s" is an operator, and "τ1" is a time constant.

Using the relative velocity (ΔV1x, ΔV1y) calculated by the pseudo-differential filter <NUM>, such as that of Equation (<NUM>), for subsequent control may result in decreased controllability due to primary delay. Therefore, the guidance calculation unit <NUM> uses a complementary filter <NUM> to calculate the relative velocity (ΔVx, ΔVy), as illustrated in <FIG>. The complementary filter <NUM> includes a low-pass filter <NUM> and a high-pass filter <NUM>.

The guidance calculation unit <NUM> applies a low-pass filter <NUM> to the relative velocity (ΔV1x, ΔV1y) calculated by the pseudo-differential filter <NUM>, and calculates relative velocity (ΔV2x, ΔV2y) having attenuated frequencies equal to or higher than a predetermined cutoff frequency. A transfer function G2(s) of the low-pass filter <NUM> is represented by Equation (<NUM>). In Equation (<NUM>), the symbol "s" is an operator, and "τ2" is a time constant. The predetermined cutoff frequency turns out to be "<NUM>/i2". Accordingly, it is possible to acquire the relative velocity (ΔV2x, ΔV2y) reflecting relatively reliable and moderate variation of the relative velocity (ΔV1x, ΔV1y), i.e., having a value in a low-frequency range, which is equal to or lower than the predetermined cutoff frequency.

In addition, the guidance calculation unit <NUM> applies the high-pass filter <NUM> to the aircraft velocity (Vx, Vy) of the aircraft <NUM> detected by the navigation system <NUM>, and calculates relative velocity (ΔV3x, ΔV3y) having attenuated frequencies below a predetermined cutoff frequency. A transfer function G3(s) of the high-pass filter <NUM> is represented by Equation (<NUM>). In Equation (<NUM>), the symbol "s" is an operator, and "τ2" is a time constant in common with the low-pass filter <NUM>. Therefore, the predetermined cutoff frequency turns out to be "<NUM>/τ2" also in the high-pass filter <NUM>. In other words, variation of the short-term relative velocity (ΔVx, ΔVy) is estimated to have been generated by variation of the aircraft velocity (Vx, Vy) itself of the aircraft <NUM>, and the value acquired by applying the high-pass filter <NUM> to the aircraft velocity (Vx, Vy) is estimated to be the value of the relative velocity (ΔVx, ΔVy) in the high-frequency range. Subsequently, the guidance calculation unit <NUM> calculates the value acquired by adding the relative velocity (ΔV2x, ΔV2y) and the relative velocity (ΔV3x, ΔV3y) as the relative velocity (ΔVx, ΔVy). Accordingly, it becomes possible to accurately calculate the relative velocity (ΔVx, ΔVy) by adding, to the relative velocity (ΔV2x, ΔV2y) with values in a high-frequency range having been cut-off to achieve an increased reliability, the relative velocity (ΔV3x, ΔV3y) estimated to be a value in the high-frequency range.

Let us return to explanation of <FIG>. The guidance calculation unit <NUM>, having calculated the relative position (Xhg, Yhg) and the relative velocity (ΔVx, ΔVy) in a manner described above, applies a Kalman filter <NUM> to the calculated relative position (Xhg, Yhg) and the relative velocity (ΔVx, ΔVy), and calculates a relative position (Xhgf, Yhgf) with noise removed and error reduced. Here, the Kalman filter <NUM> may be omitted. Subsequently, the guidance calculation unit <NUM> performs PID control by a PID control unit <NUM> (feedback control unit) using the relative position (Xhgf, Yhgf), the relative velocity (ΔVx, ΔVy), and the aircraft acceleration (ax, ay) and calculates a feedback manipulated variable C. Specifically, the guidance calculation unit <NUM> calculates the feedback manipulated variable C of the aircraft <NUM> by PID control so that the relative position (Xhgf, Yhgf) corresponding to the distance in the horizontal direction between the aircraft <NUM> and the target landing point <NUM> becomes <NUM>. Accordingly, the feedback manipulated variable C can be determined so that the aircraft <NUM> is guided toward the target landing point <NUM> and the aircraft <NUM> becomes relatively stationary with respect to the marine vessel <NUM> directly above the target landing point <NUM>. In addition, the guidance calculation unit <NUM> calculates the feedback manipulated variable C of the aircraft <NUM> by PID control so that values of the relative velocity (ΔVx, ΔVy) and the aircraft acceleration (ax, ay) become <NUM>. Accordingly, the accuracy with regard to guidance of the aircraft <NUM> toward the target landing point <NUM> can be improved. Note that, although it is assumed that the PID control uses the aircraft acceleration (ax, ay) in the present embodiment, it suffices that the PID control is performed based on at least the relative position and the relative velocity between the aircraft <NUM> and the target landing point <NUM>.

In addition, the guidance calculation unit <NUM> may skip the integration operation of PID control when the relative velocity (ΔVx, ΔVy) is equal to or higher than a predetermined value. In the present embodiment, a relatively high gain of the integration operation of PID control is set when the aircraft <NUM> is relatively stationary with respect to the target landing point <NUM>, in consideration of canceling the force that the aircraft <NUM> receives from surrounding wind. Here, when the force of wind being received by the aircraft <NUM> has weakened, the aircraft <NUM> may temporarily transit from the relatively stationary state and, being balanced with the force of wind at a certain position, return to the stationary state again. On this occasion, the feedback manipulated variable C for causing the aircraft <NUM> to return to the target landing point <NUM> is calculated by the integration operation of PID control. Accordingly, although the integration operation of PID control is a necessary component, a relatively high gain is set as described above, and therefore the value calculated by the integration operation may be too large when the relative velocity (ΔVx, ΔVy) is high, which may result in occurrence of overshooting that causes the aircraft <NUM> to pass through the target landing point <NUM>. Therefore, skipping the integration operation of PID control when the relative velocity (ΔVx, ΔVy) is equal to or higher than a predetermined value allows for suppressing overshooting even when the integration gain is set to be relatively large.

Furthermore, the guidance calculation unit <NUM> performs, by a multi-value control unit <NUM> in parallel with PID control, a multi-value control of setting the addition value D to be added to the feedback manipulated variable C. The multi-value control unit <NUM> receives the relative position (Xhgf, Yhgf) and the relative velocity (ΔVx, ΔVy) input thereto. The multi-value control unit <NUM> calculates the addition value D, based on the input relative position (Xhgf, Yhgf) and the relative velocity (ΔVx, ΔVy).

<FIG> is an explanatory diagram illustrating a coordinate plane with the relative position and the relative velocity being orthogonal axes. The horizontal axis represents the relative position (Xhgf, Yhgf) and the vertical axis represents the relative velocity (ΔVx, ΔVy). Here, <FIG> also illustrates both the component in the X direction that serves as the direction of the pitch axis, and the component in the Y direction that serves as the direction of the roll axis. The relative position (Xhgf, Yhgf) here corresponds to the distance from the aircraft <NUM> to the target landing point <NUM>, which is assumed to be positive (right direction) when the aircraft <NUM> is ahead of the target landing point <NUM> in the forward direction of the aircraft <NUM>, or negative (left direction) when the aircraft <NUM> is behind the target landing point <NUM> in the forward direction of the aircraft <NUM>. In addition, the relative velocity (ΔVx, ΔVy) here is assumed to be positive (upward direction) when the aircraft velocity (Vx, Vy) of the aircraft <NUM> is higher than the velocity of the marine vessel <NUM>, and negative (downward direction) when the aircraft velocity (Vx, Vy) of the aircraft <NUM> is lower than the velocity of the marine vessel <NUM>. Accordingly, the origin of the coordinate plane indicates a coordinate point at which the aircraft <NUM> is directly above the target landing point <NUM> and relatively stationary with respect to the target landing point <NUM>. Note that the positive and negative signs may be reversed with respect to the relative position (Xhgf, Yhgf) and the relative velocity (ΔVx, ΔVy). <FIG> illustrates an example of a coordinate point P of the current relative position (Xhgf, Yhgf) and the current relative velocity (ΔVx, ΔVy).

The coordinate plane has preliminarily set thereon a switching line L1 passing through the origin and separating the plane into an acceleration region A1 (the range indicated by diagonal lines in <FIG>) in which the relative velocity (ΔVx, ΔVy) is increased and a deceleration region A2 (the range without diagonal lines in <FIG>) in which the relative velocity (ΔVx, ΔVy) is decreased. The switching line L1 is a straight line passing through the origin of the coordinate plane and extending between a quadrant in which the aircraft velocity (Vx, Vy) of the aircraft <NUM> is higher than the velocity of the marine vessel <NUM> (velocity of target landing point <NUM>) and also the aircraft <NUM> is behind the target landing point <NUM> in the forward direction, and a quadrant in which the aircraft velocity (Vx, Vy) of the aircraft <NUM> is lower than the velocity of the marine vessel <NUM>, and also the aircraft <NUM> is ahead of the target landing point <NUM> in the forward direction. Therefore, in this embodiment, the switching line L1 passes through the origin and extends between the second quadrant and the fourth quadrant, as illustrated in <FIG>. The angle of the switching line L1 is not limited to that illustrated in <FIG>, and may be set by the user as appropriate. For example, although increasing the angle of the switching line L1 (tilting toward the relative velocity side) increases the aircraft velocity (Vx, Vy) of the aircraft <NUM> returning to the target landing point <NUM>, it also increases the likelihood of overshooting and therefore the angle is set to an appropriate inclination.

<FIG> is an explanatory diagram illustrating an example of a map defining an addition value to be set in multi-value control. In <FIG>, the horizontal axis represents the distance from the switching line L1 to the current coordinate point P on the coordinate plane of <FIG>, and the vertical axis represents the addition value D to be set. The distance from the switching line L1 to the current coordinate point P is positive when the current coordinate point P is located in the acceleration region A1, and negative when the current coordinate point P is located in the deceleration region A2. In the present embodiment, the addition value D, when taking a positive value, is set to a trend that increases the relative velocity (ΔVx, ΔVy), i.e., a trend that increases the velocity of the aircraft <NUM> in the forward direction. Additionally, in the present embodiment, the addition value D, when taking a negative value, is set to a trend that reduces the relative velocity (ΔVx, ΔVy), i.e., a trend that reduces the velocity of the aircraft <NUM> in the forward direction or increases the same in the backward direction. In addition, as indicated by the solid line in the drawing, the addition value D is set in a stepwise manner so that its absolute value becomes larger for a further distance between the coordinate point P and the switching line L1, or smaller for a closer distance between the coordinate point P and the switching line L1.

The multi-value control unit <NUM> sets the addition value D illustrated in <FIG> in accordance with the distance between the switching line L1 and the coordinate point P of the relative position (Xhgf, Yhgf) and relative velocity (ΔVx, ΔVy), and adds, in an addition circuit <NUM>, the addition value D to the feedback manipulated variable C calculated by the PID control unit <NUM>, as illustrated in <FIG>. Subsequently, the guidance calculation unit <NUM> outputs, to the flight control unit <NUM>, the manipulated variable C' to which the addition value D has been added.

Accordingly, when the coordinate point P is located in the acceleration region A1, the flight of the aircraft <NUM> is controlled by the manipulated variable C' having added thereto the addition value D that tends to increase the relative velocity (ΔVx, ΔVy). Consequently, as illustrated by the solid line in <FIG>, the relative velocity (ΔVx, ΔVy) becomes larger than the case where the aircraft <NUM> flies according to the feedback manipulated variable C without the addition value D illustrated by the dashed line in <FIG> added thereto. In other words, the aircraft velocity (Vx, Vy) of the aircraft <NUM> rapidly increases, reducing the time required to travel a same distance. When, on the other hand, the coordinate point P crosses the switching line L1 and reaches the deceleration region A2, the flight of the aircraft <NUM> is controlled by the manipulated variable C' having added thereto the addition value D that tends to decrease the relative velocity (ΔVx, ΔVy). Consequently, as illustrated by the solid line in <FIG>, although the relative velocity (ΔVx, ΔVy) indicated by the solid line becomes higher than the relative velocity indicated by the dashed line immediately after the coordinate point P moved from the acceleration region A1 to the deceleration region A2 across the switching line L1, the relative velocity indicated by the solid line becomes lower than the relative velocity indicated by the dashed line as the aircraft <NUM> approaches the target landing point <NUM>. In other words, the aircraft velocity (Vx, Vy) of the aircraft <NUM> rapidly decelerates, thereby suppressing overshooting that causes the aircraft <NUM> to pass through the origin of the coordinate plane.

In addition, a plurality of addition values D may be set in accordance with the altitude of the aircraft <NUM> (relative altitude with respect to the target landing point <NUM>). For example, when the aircraft <NUM> is located directly above the target landing point <NUM> at a predetermined altitude (e.g., in a range of <NUM> or more and <NUM> or less with respect to the target landing point <NUM>) (e.g., in a low-altitude hovering mode described below), the absolute value of the addition value D may be set larger than the value in the normal time indicated by the solid line (e.g., in a high-altitude hovering mode described below), as illustrated by the dashed line in <FIG>. On this occasion, setting of the addition value D may be gradually changed in accordance with the value of the altitude. Accordingly, the manipulated variable C' of the aircraft <NUM> can be more accurately calculated in the vicinity of the target landing point <NUM> where the aircraft <NUM> lands on, thereby facilitating accurate landing on the target landing point <NUM>.

Similarly, a plurality of the addition values D may be set in accordance with the control mode of the aircraft <NUM>. For example, let us consider a case where there are set, as the control mode for the aircraft <NUM>, a high-altitude hovering mode that maintains the relative altitude with respect to the target landing point <NUM> at a first relative altitude (e.g., <NUM>) directly above the target landing point <NUM>, and a low-altitude hovering mode that lowers the relative altitude with respect to the target landing point <NUM> from the high-altitude hovering mode to a second relative altitude (e.g., <NUM>) directly above the target landing point <NUM>. Here, transition from the high-altitude hovering mode to the low-altitude hovering mode can be performed under a condition that the relative position (Xhgf, Yhgf) is equal to or lower than a predetermined threshold value, and also the operator has instructed mode transition. In addition, transition from the high-altitude hovering mode to the low-altitude hovering mode may be performed under a condition that, in place of instructing mode transition by the operator, the attitude rate, the attitude angle, the relative velocity (ΔVx, ΔVy) of the aircraft <NUM>, the angle of the target landing point <NUM> in the horizontal direction, the altitude of the aircraft <NUM> with respect to the target landing point <NUM>, the relative altitude with respect to the target landing point <NUM>, or the like are equal to or lower than a predetermined value.

In a case where such a high-altitude hovering mode and a low-altitude hovering mode have been set, setting of the addition value D is switched after a first predetermined time period (e.g., <NUM> seconds) has elapsed since transition from the high-altitude hovering mode to the low-altitude hovering mode started. In other words, the addition value D is switched from the value indicated by the solid line in <FIG> to the value indicated by the dashed line in <FIG>. On this occasion, the addition value D may be set so as to gradually change during a second predetermined time period (e.g., <NUM> seconds). Accordingly, the manipulated variable C' of the aircraft <NUM> can be more accurately calculated in the vicinity of the target landing point <NUM> where the aircraft <NUM> lands on, thereby facilitating accurate landing on the target landing point <NUM>.

As has been described above, the aircraft position control system <NUM> according to the first embodiment includes: the image processing unit <NUM> and the guidance calculation unit <NUM> (relative position acquisition unit) that acquires the relative position (Xhgf, Yhgf) between the aircraft <NUM> and the target landing point <NUM>; a guidance calculation unit <NUM> (relative velocity acquisition unit) that acquires the relative velocity (ΔVx, ΔVy) of the aircraft <NUM> with respect to the target landing point <NUM>; and the control unit <NUM> that controls the aircraft <NUM>. The control unit <NUM> includes, in the guidance calculation unit <NUM>: the PID control unit <NUM> (feedback control unit) that calculates the feedback manipulated variable C of the aircraft <NUM> by feedback control so that the aircraft <NUM> is directed toward the target landing point <NUM> based on at least the relative position (Xhgf, Yhgf) and the relative velocity (ΔVx, ΔVy); the multi-value control unit <NUM> that sets, by referring to a switching line L1 preliminarily provided in a manner passing through an origin (<NUM>) of a coordinate plane whose orthogonal axes represent the relative position (Xhgf, Yhgf) and the relative velocity (ΔVx, ΔVy) and separating an acceleration region A1 in which the relative velocity (ΔVx, ΔVy) is increased and a deceleration region A2 in which the relative velocity is decreased (ΔVx, ΔVy), the addition value D that tends to increase at least the relative velocity (ΔVx, ΔVy) when the coordinate point P of the current relative position (Xhgf, Yhgf) and the current relative velocity (ΔVx, ΔVy) is located in the acceleration region A1, or sets the addition value D that tends to decrease at least the relative velocity (ΔVx, ΔVy) when the coordinate point P is located in the deceleration region A2 with respect to the switching line L1; and the addition circuit <NUM> that calculates the manipulated variable C' of the aircraft <NUM> by adding the addition value D to the feedback manipulated variable C.

According to the aforementioned configuration, the flight of the aircraft <NUM> is controlled by the manipulated variable C' having added thereto the addition value D that tends to increase the relative velocity (ΔVx, ΔVy), whereby the relative velocity ΔV increases, when the coordinate point P of the current relative position (Xhgf, Yhgf) and the current relative velocity (ΔVx, ΔVy) is located in the acceleration region A1 separated by the switching line L1 on a coordinate plane whose orthogonal axes represent the relative position (Xhgf, Yhgf) and the relative velocity (ΔVx, ΔVy). As a result, the aircraft <NUM> can be moved fast. When, on the other hand, the current coordinate point P is located in the deceleration region A2 separated by the switching line L1, the flight of the aircraft <NUM> is controlled by the manipulated variable C' having added thereto the addition value D, which tends to decrease the relative velocity (ΔVx, ΔVy), whereby the relative velocity ΔV decreases. As a result, overshooting that causes the aircraft <NUM> to pass through the target landing point <NUM> can be suppressed. Therefore, according to the aircraft <NUM>, the aircraft position control system <NUM>, and the position control method according to the first embodiment, the aircraft <NUM> can be moved more accurately and faster toward the target landing point <NUM>. Accordingly, accurately controlling the position of the aircraft <NUM> with respect to the target landing point <NUM> allows for suppressing interference between the aircraft <NUM> and a device or a structure provided in the vicinity of the target landing point <NUM>.

In addition, the switching line L1 is a straight line extending over the coordinate plane between a quadrant in which the aircraft velocity (Vx, Vy) of the aircraft <NUM> is higher than the velocity of target landing point <NUM> (velocity of the marine vessel <NUM>) and also the aircraft <NUM> is behind the target landing point <NUM> in the forward direction, and a quadrant in which the aircraft velocity (Vx, Vy) of the aircraft <NUM> is lower than the velocity of the target landing point <NUM>, and also the aircraft <NUM> is ahead of the target landing point <NUM> in the forward direction. According to the aforementioned configuration, it is possible to appropriately set the acceleration region A1 and the deceleration region A2, and set the addition value D to an appropriate value.

In addition, the addition value D is set in a stepwise manner so that its absolute value becomes larger for a larger distance on the coordinate plane between the switching line L1 and the coordinate point P of the current relative position (Xhgf, Yhgf) and the current relative velocity (ΔVx, ΔVy), or smaller for a closer distance.

According to the aforementioned configuration, the addition value D is set larger for a further distance between the switching line L1 and the coordinate point P of the current relative position (Xhgf, Yhgf) and the current relative velocity (ΔVx, ΔVy) so that the aircraft <NUM> moves fast, or smaller for a closer distance between the current coordinate point P and the switching line L1 so that the aircraft <NUM> can be finely manipulated. However, the manner of setting the addition value D is not limited to the foregoing. <FIG> is an explanatory diagram illustrating another example of a map defining an addition value to be set in multi-value control. As illustrated, the addition value D may be a positive constant value in the acceleration region A1, or a negative constant value in the deceleration region A2, regardless of the distance between the current coordinate point P and the switching line L1.

In addition, the PID control unit <NUM> (feedback control unit), having calculated the feedback manipulated variable by PID control, skips the integration operation of PID control when the relative velocity (ΔVx, ΔVy) is equal to or higher than a predetermined value. According to the aforementioned configuration, in case where a relatively large integrated gain of PID control has been set, it is possible to suppress reduction of controllability of the PID control, or more specifically, occurrence of overshooting that causes the aircraft <NUM> to pass through the target landing point <NUM> due to the relative velocity (ΔVx, ΔVy) between the aircraft <NUM> and the target landing point <NUM> being large.

In addition, a plurality of the addition values D may be set in accordance with the altitude or control mode of the aircraft <NUM>. The aforementioned configuration allows for increasing the degree of freedom of setting the addition value D. Accordingly, for example, in a case where the aircraft <NUM> is located in the vicinity of the target landing point <NUM>, the position of the aircraft <NUM> can be controlled more accurately by setting the addition value D to a large value. Here, a plurality of gains of the PID control may also be set in accordance with the altitude or control mode of the aircraft <NUM>.

In addition, in the first embodiment, the guidance calculation unit <NUM> (the relative velocity acquisition unit) calculates the relative velocity (ΔVx, ΔVy) by adding a value (relative velocity (ΔV2x, ΔV2y)) acquired by applying, to a differential value (relative velocity (ΔV1x, ΔV1y)) of the relative position (Xhgf, Yhgf), a low-pass filter <NUM> that attenuates frequencies equal to or higher than a predetermined cutoff frequency, and a value (relative velocity (ΔV3x, ΔV3y)) acquired by applying, to the aircraft velocity (Vx, Vy) of the aircraft <NUM>, a high-pass filter <NUM> that attenuates frequencies below the cutoff frequency.

According to the aforementioned configuration, acquisition of the relative position (Xhgf, Yhgf) between the aircraft <NUM> and the target landing point <NUM> and the aircraft velocity (Vx, Vy) of the aircraft <NUM> allows for calculating the relative velocity (ΔVx, ΔVy) between the aircraft <NUM> and the target landing point <NUM> without having to acquire the velocity of the marine vessel <NUM> side on which the target landing point <NUM> is provided. Accordingly, it is not necessary to perform communication between the aircraft <NUM> and the marine vessel <NUM> side on which the target landing point <NUM> is provided, thereby preventing reduction of control accuracy and response speed due to communication delay. Furthermore, the relative velocity (ΔVx, ΔVy) can be accurately calculated by applying the high-pass filter <NUM> to the aircraft velocity (Vx, Vy) of the aircraft <NUM> and adding thereto the relative velocity (ΔV3x, ΔV3y) in the high-frequency range due to variation of the aircraft velocity (Vx, Vy) of the aircraft <NUM>, while using, as the relative velocity (ΔV2x, ΔV2y), a highly reliable value in a low-frequency range, which has been acquired by applying the low-pass filter <NUM> to the differential value of the relative position (Xhgf, Yhgf). Accordingly, it becomes possible to move the aircraft <NUM> more accurately and faster toward the target landing point <NUM>. In addition, data communication is not required, and therefore the system can be simplified.

In addition, the position control system <NUM> further includes the camera <NUM> (imaging device) mounted on the aircraft <NUM>, and the image processing unit <NUM> and the guidance calculation unit <NUM> (relative position acquisition unit) calculate the relative position (Xhgf, Yhgf) by image processing using images, captured by the camera <NUM>, of the marker <NUM> provided at the target landing point <NUM>. According to the aforementioned configuration, the relative position (Xhgf, Yhgf) between the aircraft <NUM> and the target landing point <NUM> can be acquired by image processing, whereby it becomes unnecessary to acquire position information from the marine vessel <NUM> side on which the target landing point <NUM> is provided. Accordingly, it is unnecessary to perform communication between the aircraft <NUM> and the marine vessel <NUM> side on which the target landing point <NUM> is provided, and therefore the system can be simplified.

Next, there will be described an aircraft position control system <NUM> and a position control method according to a second embodiment. <FIG> is a schematic configuration diagram illustrating an example of a position control system according to the second embodiment. The position control system <NUM> according to the second embodiment includes, as illustrated in <FIG>, a data transmission device <NUM> in addition to the configuration of the position control system <NUM> according to the first embodiment. In addition, the position control system <NUM> includes a guidance calculation unit 34A, in place of the guidance calculation unit <NUM> of the first embodiment. Other configuration of the position control system <NUM> is similar to that of the position control system <NUM> and therefore description thereof is omitted, and like components are indicated by like reference signs.

Additionally, in the second embodiment, the marine vessel <NUM> having the target attaching point <NUM> provided thereon includes a navigation system <NUM>, as illustrated in <FIG>. Similarly to the first embodiment, the navigation system <NUM> is an inertial navigation system, for example, and the marine vessel <NUM> acquires, via the inertial navigation system, a ship velocity (Vsx, Vsy), as well as the acceleration, the ship heading ψs (see <FIG>), or the like of the marine vessel <NUM>. The inertial navigation system may include the GPS or may include sensors to acquire various data. In addition, the marine vessel <NUM> includes a data transmission device <NUM>.

The data transmission device <NUM> and the data transmission device <NUM> communicate with each other to exchange information between the aircraft <NUM> and the marine vessel <NUM>. Specifically, in the second embodiment, the data transmission device <NUM> of the marine vessel <NUM> transmits the ship velocity (Vsx, Vsy) of the marine vessel <NUM> acquired by the navigation system <NUM> to the data transmission device <NUM> of the aircraft <NUM>. The data transmission device <NUM> of the aircraft <NUM> transmits the received ship velocity (Vsx, Vsy) of the marine vessel <NUM> to the guidance calculation unit 34A.

<FIG> is a block diagram illustrating an example of a configuration for calculating a manipulated variable of an aircraft by the guidance calculation unit of the second embodiment. In the second embodiment, the guidance calculation unit 34A skips the relative velocity calculation process of the first embodiment. The guidance calculation unit 34A calculates, as illustrated in <FIG>, a difference between the aircraft velocity (Vx, Vy) of the aircraft <NUM> acquired by the navigation system <NUM> and the ship velocity (Vsx, Vsy) of the marine vessel <NUM>, and sets the calculated difference as the relative velocity (ΔVx, ΔVy) of the aircraft <NUM> with respect to the target landing point <NUM>. On this occasion, the ship velocity (Vsx, Vsy) of the marine vessel <NUM> is the coordinates of the ship inertial reference frame SG, and therefore subjected to coordinate conversion into the aircraft inertial reference frame HG when calculating the difference. Therefore, the guidance calculation unit 34A functions, also in the second embodiment, as a relative velocity acquisition unit that acquires the relative velocity (ΔVx, ΔVy) between the aircraft <NUM> and the target landing point <NUM>. The guidance calculation unit 34A calculates the manipulated variable C' by feedback control and multi-value control similar to those in the first embodiment, using the calculated relative velocity (ΔVx, ΔVy). Here, although the Kalman filter <NUM> is omitted in <FIG>, the relative position (Xhgf, Yhgf) calculated by applying the Kalman filter <NUM> to the relative position (Xhg, Yhg) and the relative velocity (ΔVx, ΔVy) may be used, similarly to the first embodiment.

According to the aforementioned configuration, the aircraft <NUM>, the aircraft position control system <NUM>, and the position control method of the second embodiment eliminates the necessity of calculating the pseudo-differential values of the relative position (Xhg, Yhg), or performing a calculation process using the low-pass filter <NUM> and the high-pass filter <NUM>, as in the first embodiment, when calculating the relative velocity (ΔVx, ΔVy) of the aircraft <NUM> with respect to the target landing point <NUM>. Therefore, the calculation required for calculating the relative velocity (ΔVx, ΔVy) can be simplified. Here, a filter such as a low-pass filter and a high-pass filter may be applied to the ship velocity (Vsx, Vsy) of the marine vessel <NUM> which have been subjected to coordinate conversion.

It is assumed in the first and the second embodiments that an image of the marker <NUM> is captured by the camera <NUM> and image processing is performed on the captured image, and that the relative position between the aircraft <NUM> and the target landing point <NUM> is calculated by the relative position calculation process described above. However, the method for acquiring the relative position between the aircraft <NUM> and the target landing point <NUM> is not limited to the foregoing. For example, a laser irradiation device may be mounted on the aircraft <NUM> to acquire a relative position by irradiating a laser beam toward the target landing point <NUM> on the marine vessel <NUM> and receiving reflected waves by the aircraft <NUM>.

In the first and the second embodiments, the guidance calculation units <NUM> and 34A may measure or estimate the wind velocity around the aircraft <NUM> or the airspeed, which is the relative velocity between the aircraft <NUM> and the atmosphere. Subsequently, the guidance calculation units <NUM> and 34A may adjust each gain of feedback control in accordance with the acquired wind velocity or airspeed. Similarly, the guidance calculation units <NUM> and 34A may change the setting of the addition value of multi-value control in accordance with the acquired wind velocity or airspeed. Accordingly, it is possible to output a manipulated variable in accordance with the strength of wind around the aircraft <NUM>, and guide the aircraft <NUM> toward the target landing point <NUM> more accurately and faster.

Claim 1:
An aircraft position control system comprising:
a relative position acquisition unit (<NUM>) configured to acquire a relative position (Xhgf, Yhgf) between an aircraft (<NUM>) and a target landing point (<NUM>);
a relative velocity acquisition unit (<NUM>) configured to acquire relative velocity (ΔVx, ΔVy) between the aircraft (<NUM>) and the target landing point (<NUM>); and
a control unit (<NUM>) configured to control the aircraft (<NUM>) to head toward the target landing point (<NUM>), based on at least the relative position (Xhgf, Yhgf) and the relative velocity (ΔVx, ΔVy), wherein
the relative velocity acquisition unit (<NUM>) calculates the relative velocity (ΔVx, ΔVy) by adding a value (ΔV2x, ΔV2y) acquired by applying, to a differential value (ΔV1x, ΔV1y) of the relative position (Xhgf, Yhgf), a low-pass filter (<NUM>) that attenuates frequencies equal to or higher than a cutoff frequency (<NUM>/τ2) being predetermined, and a value (ΔV3x, ΔV3y) acquired by applying, to the velocity of the aircraft (<NUM>), a high-pass filter (<NUM>) that attenuates frequencies below the cutoff frequency (<NUM>/τ2).