JP5127284B2 - Flying object and apparatus for detecting rotational position of flying object - Google Patents

Description

  The present invention relates to a flying object that is launched from a launching device while rotating around a central axis, and a rotational position detection device for the flying object.

  Unlike a projectile with a propulsive force such as a rocket or missile, a projectile that does not have a propelling force such as a cannonball basically has a landing point depending on the direction and launch speed when fired from a launching device such as a gun. As a result, the deviation from the target point is caused by the error in the launch direction and the launch speed and the disturbance during the flight such as wind. Various improvements have been made to reduce the error between the landing point and target point of a projectile that does not have propulsive force, but these mainly improve the accuracy of the launch direction and launch speed and reduce the effects of disturbance. Met. Such a flying object is described in Patent Document 1, for example.

  On the other hand, in recent years, a new flying object that has a flying direction control function for a flying object that does not have a propulsive force to land at a target point more accurately has been proposed.

  FIG. 1 is an external view of the new flying object 1. As shown in the figure, the flying object 1 includes a main body 2 having a cylindrical shape with a sharp tip, a tail 3, and a control wing 4. The tail wing 3 and the control wing 4 are accommodated in the main body 2 before being launched by the launching device, and project from the main body 2 after being launched into the air. Each control wing 4 can change the amount of protrusion independently, and adjusts the amount of protrusion and adjusts a flight direction. The flying object 1 further includes a GPS device 5, a movement control device 6, and an inertial navigation device 7 in the main body 2. The inertial navigation device 7 includes a gyro device that detects rotation about the central axis (x-axis) of the flying object 1, the y-axis and the z-axis that are perpendicular to the x-axis and perpendicular to each other, and acceleration in these three-axis directions. And an accelerometer to be detected. The flight direction is calculated based on the output of the gyro device and the accelerometer, and the error of the flight position and direction is corrected as needed by the output of the GPS device 5. Here, the x axis, the y axis, and the z axis are referred to as a first axis, a second axis, and a third axis, respectively.

  FIG. 2 is a diagram for explaining the trajectory and operation from the launch of the flying object 1 of FIG. The flying object 1 is launched toward the target point by the launching device 8. As is well known, a flying object having no propulsive force is launched from the launching device 8 while rotating around the central axis (x-axis) at high speed so as to fly stably along the trajectory. The flying object 1 is premised on using a conventional launching device as it is, and is launched while rotating around the x-axis at a high speed as in the prior art.

  Since the tail wing 3 and the control wing 4 are projected, they cannot be fired, so they are accommodated in the main body 2 before being fired by the apparatus.

  The launched flying object 1 flies on the trajectory 9 of FIG. 2, and a battery (not shown) mounted on the main body 2 is turned on at the point A, and the GPS device 5, the movement control device 6 and the inertia are correspondingly turned on. The movement control device 6 causes the tail blade 3 and the control blade 4 to protrude so that the navigation device 7 starts to operate and decelerates rotation about the x axis. Then, it is detected that the rotation has stopped at point B.

  When it is detected that the rotation has stopped, the inertial navigation device 7 and the GPS device 5 detect the flight position and flight direction of the flying object 1. The movement control device 6 can control the flight of the flying object 1 at the point C. Thereafter, the movement control device 6 detects the flying position of the flying object 1 detected by the inertial navigation device 7 and the GPS device 5 and The control wing 4 is controlled based on the flight direction so that the flying object 1 is landed on the target point.

JP 2007-24450 A

  In order to perform the above-described control in the proposed flying object, the attitude angle of the flying object 1 (main body part 2) needs to be known.

  The GPS device 5 can detect the flight position, speed, and flight direction of the flying object 1, but cannot determine the attitude angle of the flying object 1 (main body 2).

  The conventional inertial navigation system detects the earth's gravitational acceleration component proportional to the inclination with the accelerometer of the gyroscope when it is stopped on the ground or flying horizontally in the air. (Pitch angle and rotation (roll) angle) are calculated. However, since the flying object 1 shown in FIG. 1 does not have a propulsive force, it does not fly horizontally. In this manner, that is, by using an accelerometer provided in an inertial navigation device, The attitude angle of the main body 1 cannot be determined by the technique.

  An object of the present invention is to detect the attitude angle of the new type of flying object shown in FIG. 1, and particularly to detect it at an early stage.

  In order to achieve the above object, the flying object of the present invention detects the rotational position around the center axis of the flying object from the output of the gyro device of the inertial navigation device while the flying object is rotating at low speed, Used for control.

  That is, the flying object of the present invention is a flying object that is launched while rotating from the launching device, and is detected by the variable wing protruding after launching, the inertial navigation device that detects the flight position, and the inertial navigation device. A flying body comprising a movement control device that controls the variable wing so as to reach a target based on a flying position, wherein the inertial navigation device uses the center axis of the flying body as a first axis, and the first An axis perpendicular to each other and perpendicular to each other is used as a second and third axes, and a gyro device that detects rotation about the first to third axes is provided. The inertial navigation device includes the movement control device that performs the flight When the number of rotations of the flying object around the first axis becomes smaller than a first predetermined value by projecting the variable wing after the body is launched, the output of the gyro device is used to rotate around the first axis of the flying object. Detects the rotation position of And wherein the Rukoto.

  The inertial navigation device calculates the pitch angle from the velocity signal of the GPS device, and determines the attitude angle of the flying object from the rotational position and pitch angle around the first axis.

  The movement control device controls the variable wing so that the flying object has a predetermined attitude angle based on the detected attitude angle of the flying object.

  The inertial navigation device includes an accelerometer that detects acceleration in the directions of the first to third axes, and a GPS device, and after the rotational speed of the flying object around the first axis becomes smaller than a second predetermined value, the accelerometer The flying position of the flying object is detected from the output of the GPS device.

  According to the present invention, after launching, the flying object that rotates at high speed is made to rotate at a low speed by projecting the variable wing, and in this state, the rotational position around the first axis of the flying object is detected from the output of the gyro device. To do. This makes it possible not only to detect the rotational position of the flying object, but also to detect the rotational position even when the flying object is rotating at a low speed, so that the rotation is controlled to stop in a desired posture. It is possible to detect the flight position and the flight direction in a short time from the launch and control to reach the target point, and the accuracy during the life can be improved.

  The inertial navigation device detects the rotational position around the first axis from the time when the rotation signals of the second and third axes of the gyro device pass through the center level when the flying object rotates at a low speed. Specifically, the inertial navigation device calculates a moving average of the rotation signals of the second and third axes of the gyro device, calculates an approximate sine wave from the calculated moving average, and the calculated approximate sine wave has the center level. The rotational position around the first axis is detected from the passing time. This method is a kind of sample-like detection method, and the error is not sufficiently small for guidance control. Therefore, the inertial navigation device uses the calculated rotation position around the first axis as the initial rotation position, calculates the ballistic pitch rate from the rotation rates around the second and third axes of the gyro device, and sets the gyro device at the initial rotation position. The pseudo rotation position is calculated by adding the integral value of the rotation rate around the first axis, the pseudo pitch rate (corresponding to the pseudo rotation rate around the second axis) is calculated from the pseudo rotation position and the ballistic pitch rate, and the pseudo pitch rate is calculated. The difference in rotation rate around the second axis of the gyro device is calculated, the peak value is calculated from the average of the calculated change in half cycle, and the rotational position error is calculated from the calculated peak value and ballistic pitch rate. The initial rotational position is newly calculated from the rotational position error, and the process is repeated so that the rotational position error converges.

  The flying object rotation position detection device of the present invention is a flying object rotation position detection device that is fired while rotating around the central axis from the launching device, wherein the center axis of the flying object is the first axis. An axis perpendicular to the first axis and perpendicular to each other as the second and third axes, and a gyro device that detects rotation about the first to third axes, and the second and third gyro devices The rotation position around the first axis is detected from the time when the rotation signal of the third axis passes the center level.

  According to the present invention, it becomes possible to detect the rotational position and posture angle of a flying object launched while rotating from a launching apparatus that has been difficult so far, and to detect the rotational position and posture angle in a short time after launch. Become.

  In addition, since the rotational position and attitude angle of the flying object can be detected in a short time after launching, the accuracy with which the flying object having variable wings reaches the target point can be improved.

  The flying object of the embodiment of the present invention has a configuration as shown in FIG. 1 and the configurations of the inertial navigation device 7 and the movement control device 6 are different.

  FIG. 3 is a diagram showing a configuration of the flying object inertial navigation device 7 of the embodiment, and shows the GPS device 5, the movement control device 6, and the variable wing drive unit 9 together. The movement control device 6 performs control so that not only the control blade 4 but also the tail blade 3 protrudes through the variable blade drive unit 9. As shown in the figure, the inertial navigation device 7 includes an x-axis gyro 11 that detects a rotation rate (angular velocity) around the central axis (x-axis) of the flying object 1, and a y-axis and a z-axis that are perpendicular to the x-axis and perpendicular to each other. A y-axis gyro 12 and a z-axis gyro 13 for detecting the rotational rate of rotation, an x-axis accelerometer 14 for detecting acceleration in the x-axis direction, a y-axis accelerometer 15 for detecting acceleration in the y-axis direction, and z The z-axis accelerometer 16 for detecting the acceleration in the axial direction, the x-axis to z-axis gyros 11 to 13, the x-axis to z-axis accelerometers 14 to 16 and the output of the GPS device 5 and the attitude angle of the flying object, the flight position, and And an inertial navigation calculation device 17 for calculating the flight direction. The x-axis to z-axis gyros 11 to 13 and the x-axis to z-axis accelerometers 14 to 16 are well known and will not be described in detail. It is desirable to be of a mold type. The inertial navigation calculation device 17 is realized by a computer. The variable blade drive unit 9 is an actuator that allows the variable blades (tail and control blades) to protrude and the amount of protrusion to be varied. An actuator for projecting the tail 3 after launch is also provided, but the illustration is omitted here.

  Here, the rotation rate around the x axis detected by the x axis gyro 11 is the p rate, the rotation rate around the y axis detected by the y axis gyro 12 is the q rate, and the rotation rate around the z axis detected by the z axis gyro 13 is here. The rotation rate is called r rate.

FIG. 4 is a diagram showing a coordinate system representing the flight of a flying object with respect to the earth, with the launch point as the origin, the east direction as the _VEG axis, the north direction as the _VNG axis, and the reverse direction of gravity as the _VZG axis. . The component projected onto the plane formed by the _VEG axis and the _VNG axis of the trajectory is the azimuth angle, and the angle formed by this plane of the trajectory is represented by the pitch angle θ. The pitch angle θ changes from the initial launch angle to the landing, and the rate of change of the pitch angle θ is called the ballistic pitch rate. The trajectory pitch rate is expressed by a derivative of θ. Conversely, when the trajectory pitch rate is integrated and added to the initial detection angle, the pitch angle θ is obtained. The flying object launched from the launching device rotates at high speed around the x axis.

  FIG. 5 is a flowchart showing processing operations after launch of the movement control device 6 and the inertial navigation device 7. The processing operation will be described below with reference to the flowchart.

  The flying object is launched at a very high speed from the launching device so as to fly, for example, 50 km. When launched, the projectile is rotating around the x axis at high speed.

  In step 101, the movement control device 6 controls the tail blade 3 and the control blade 4 to protrude immediately after launch. By projecting the tail wing 3 and the control wing 4, the rotation of the flying object around the x-axis is decelerated, and is decelerated to a region where the rotation rate can be measured by the x-axis to z-axis gyros 11 to 13. The output signals of the x-axis to z-axis gyros 11 to 13 are processed by an inertial navigation calculation device 17 constituted by a computer. When the flying object is rotating at a high speed, the output signal of the x-axis gyro 11 changes at a high speed and cannot be processed by the inertial navigation calculation device 17, so it is necessary to decelerate until the processing area is reached. is there. As will be described later, the measurable rotational speed region is, for example, 100 deg / sec to 200 deg / sec. Although it can be determined from the output signals of the x-axis to z-axis gyros 11 to 13 that the rotation rate has been reduced to a measurable region, it may be determined by other methods.

  In steps 102 and 103, a pseudo roll angle is calculated. Hereinafter, processing for calculating the pseudo roll angle will be described.

  Since the flying object rolls (spins) around the x-axis, the orthogonal y-axis gyro 12 and z-axis gyro 13 have a ballistic pitch rate proportional to the change in the pitch angle θ of the ballistic flight and a yaw in the yaw direction.・ Detect the rate. As represented by the following equation (1), the q rate and the r rate are represented by a ballistic pitch rate, a ballistic yaw rate, and a roll angle.

  As can be seen from equation (1), the y-axis gyro 12 detects the ballistic pitch rate when the roll angle is zero degrees, and the z-axis gyro 13 detects the ballistic pitch rate of the reverse polarity when the roll angle is 90 degrees. .

  Since the launched projectile almost flies along a straight trajectory, it can be said that the rate in the yaw direction is almost zero. Therefore, the equation (1) can be simplified as the following equations (2) and (3).

  As a result, the output signals (q rate and r rate) of the y-axis gyro 12 and the z-axis gyro 13 change sinusoidally as shown in FIG. 6, and the maximum value (amplitude) is the ballistic pitch rate. The position where the q rate becomes zero is the rotational position where the roll angle is zero degrees. The horizontal axis in FIG. 6 is an arbitrary time, and the cycle of the sine wave is determined by the rotational speed.

  However, the output signals (q rate and r rate) of the y-axis gyro 12 and the z-axis gyro 13 have fluctuations in rotation and noise of the gyro sensor itself. There are many signals. In this state, the position that passes through the amplitude and zero level cannot be detected.

  Therefore, in step 102, the inertial navigation calculation unit 17 calculates the moving average of the digital signal obtained by A / D converting the output signal of the y-axis gyro 12, and uses the moving average sample points to perform the least-squares calculation. The optimal sinusoidal envelope as shown in FIG.

  In step 103, a position where this envelope becomes zero is obtained. This position is an angle corresponding to a roll angle of 90 degrees, and the roll angle is obtained. Here, the envelope of the output signal of the y-axis gyro 12 is sinusoidal, has little change near the peak, is flat, and has a large error when judging the apex. Determine the roll angle at the zero cross point. Thereby, the roll angle can be determined with a small error.

  However, it is a sample value calculated from a signal including noise as described above, and the accuracy is insufficient. In this embodiment, the roll angle is set as a pseudo roll angle, and processing for improving accuracy is performed so that the guidance control of the flying object can be performed reliably and smoothly.

  Steps 104 to 108 are processes for further improving the accuracy by using the roll angle as a pseudo roll angle. Here, the pseudo roll angle is an angle obtained by adding a roll angle error to the true roll angle. When the true roll angle is φ and the roll angle error is Δφ, the pseudo roll angle is represented by φ + Δφ.

  In step 104, the ballistic pitch rate is calculated from the output signals (q rate and r rate) of the y-axis gyro 12 and the z-axis gyro 13.

  Since the q rate and the r rate are expressed as in equations (2) and (3), the average ballistic pitch rate is calculated when the square root of these square sums is calculated as in the following equation (4). be able to.

  In step 105, a pseudo pitch rate (pseudo q rate) is calculated by multiplying the cosine of the pseudo roll angle by the ballistic pitch rate.

  In step 106, a difference Δρ between the pseudo pitch rate (pseudo q rate) and the raw q rate actually detected by the y-axis gyro 12 is calculated. The time when the pseudo q rate matches the raw q rate generated by the y-axis gyro 12 is the minimum roll angle error. In order to obtain this roll angle error, as shown in the following equation (5), a difference Δρ between the calculated pseudo q rate and the actual q rate actually detected by the y-axis gyro 12 is calculated.

The pseudo roll angle φ + Δφ is expressed as an angle obtained by adding a change angle around the x axis over time to the initial roll angle φ ₉₀ at the zero cross point of the envelope calculated by the above least square calculation, and the change angle around the x axis is This is a value obtained by integrating the output signal (p rate) of the x-axis gyro 11 and is expressed by the following equation.

  As shown in FIG. 8, the difference Δρ changes according to the change in the roll angle φ. However, if there is an error in the positive direction (phase advance direction), the q-rate becomes a sine wave Δρ1 in phase with the raw r rate. Becomes q1, and when there is an error in the negative direction (phase lag direction), it becomes a sine wave Δρ2 opposite in phase to the r rate and the q rate becomes q2.

  If Δθ is a small angle, cos Δθ is approximately 1, so Δρ can be approximated by the following equation.

  Furthermore, if the peak value is taken, φ is approximately 90 degrees and sin φ is approximately 1, so Δρ can be approximated as the following equation.

  From this, the roll error angle Δφ can be obtained by the following equation (6).

In step 107, a roll error angle Δφ is obtained. Specifically, in order to obtain a rate component Δρ based on the roll error angle Δφ, a Δρ signal calculated in synchronization with the r rate phase is detected, and an average for a half cycle (180 °) is obtained for noise reduction. Since it is a sinusoidal signal, the peak value ΔP _PEAK can be obtained from this average value. The roll error angle Δφ is obtained from the arc sine of the ratio between the peak value and the ballistic pitch rate. Thus, Δφ can be obtained by a simple arithmetic expression.

  In step 108, it is determined whether the roll error angle Δφ is smaller than a predetermined threshold value.

  If the roll error angle Δφ is larger than the predetermined threshold value, the process proceeds to step 109, the pseudo roll angle is determined so that the roll error angle Δφ becomes smaller, and the process returns to step 104, where the roll error angle Δφ is larger than the predetermined threshold value. The above processing is repeated until it becomes smaller. The new pseudo roll angle is determined by estimation using, for example, a Kalman filter. Note that the threshold value can be determined by the dispersion value convergence of the Kalman filter.

When the roll error angle Δφ becomes smaller than a predetermined threshold value, the routine proceeds to step 110 where the attitude is set in the inertial navigation calculation device 17. Since the pitch angle θ is estimated to be substantially correlated with the elevation angle, the pitch angle θ can be estimated by calculating the attack angle from the velocity (V _NG , V _EG , V _ZG ) of the flying object detected by the GPS device 5. It can be calculated by equation (7).

  As described above, since the pitch angle θ and the rotational position (roll angle) φ around the central axis (x axis) can be calculated, these are set in the inertial navigation calculation device 17. Thereby, since the attitude | position of the flying body was set, the inertial navigation by a gyro apparatus can be started.

  If it is necessary for the flying object to have a substantially predetermined posture in order to control the flight direction, the movement control device 6 controls the control wing 4 to perform posture control in step 111.

  Thereafter, the inertial navigation calculation device 17 initializes the flight position and direction based on the flight position and direction in the earth coordinates detected by the GPS device 5, and thereafter determines the flight position and direction based on the output signal of the gyro device. The calculated flight information is output to the movement control device 6. The inertial navigation calculation device 17 calculates the flight position and direction based on the output signal of the gyro device, but corrects the accumulated error as needed based on the flight position and direction in the earth coordinates detected by the GPS device 5.

  The movement control device 6 controls the control wing 4 based on the flight information from the inertial navigation calculation device 17 so that the flying object reaches the target point.

  According to the result of the error simulation of the feedback process using the Kalman filter for the process of estimating the pseudo roll angle so that the roll error angle Δφ is small, the roll angle error is 1 ° or less if the process is repeated about 5 times. The result that can be made.

  Further, because of the ability of the computer constituting the inertial navigation calculation device 17, when the calculation cycle is 200 Hz (5 msec) and sampling is performed with a resolution of 0.5 degrees, 0.5 × 200 = 100 deg / sec, and one rotation 3.6 seconds and half rotation is 1.8 seconds, so that the rotation speed at which the calculation can be performed is about 100 deg / sec to 200 deg / sec. Therefore, in step 101, control is performed so that the rotational speed is reduced to this level. Of course, if a computer capable of high-speed calculation is used, the calculation can be performed even at a higher rotational speed.

  As mentioned above, although the Example of this invention was described, it cannot be overemphasized that various modifications are possible. For example, a flying object of any shape and configuration may be used as long as it rotates around the central axis.

  The present invention is applicable to any flying object that does not have a propulsive force and flies according to a straight trajectory when it is launched, and that is rotated and launched at the time of launch.

It is a figure which shows the shape and structure of a flying body to which this invention is applied. It is a figure explaining the operation | movement in each point on the trajectory and the trajectory of the flying body of FIG. It is a figure which shows the structure of the inertial navigation apparatus of the flying body of the Example of this invention combining a GPS apparatus, a movement control apparatus, and a variable wing drive part. It is a figure which shows the coordinate system showing the flight of the flying body with respect to the topography. It is a flowchart which shows the processing operation after discharge of a movement control apparatus and an inertial navigation apparatus. It is a figure explaining the change of the output signal of a y-axis and a z-axis gyro when a flying body flies while rotating. It is a figure which shows the actual output signal containing the noise of the y-axis and z-axis gyro when a flying body flies while rotating. It is a figure which shows the output signal and error signal of a y-axis and z-axis gyro in case there exists a roll angle error.

Explanation of symbols

DESCRIPTION OF SYMBOLS 1 Flying object 2 Main part 3 Tail 4 Control wing 5 GPS device 6 Movement control device 7 Inertial navigation device 11 X-axis gyro 12 Y-axis gyro 13 Z-axis gyro 14 X-axis accelerometer 15 Y-axis accelerometer 16 z-axis accelerometer 17 Inertial navigation arithmetic unit

Claims (7)

Without a propulsion fired while rotating from the launcher, a projectile you flying accordance trajectory by direction and rate of fire when it is fired,
Variable wings protruding after launch,
A GPS device;
An inertial navigation device for detecting the flight position;
In a flying object comprising a movement control device that controls the variable wing so as to reach a target based on a flight position detected by the GPS device and the inertial navigation device,
The inertial navigation device has a center axis of the flying object as a first axis, axes perpendicular to the first axis and perpendicular to each other as second and third axes, and rotation about the first to third axes. Equipped with a gyro device for detecting
The inertial navigation device is configured such that when the movement control device causes the variable wing to project after the flying object is launched, the rotational speed of the flying object around the first axis becomes smaller than a first predetermined value. The roll angle which is the rotational position around the first axis of the flying object is detected from the output of the device, the pitch angle is calculated from the speed signal of the GPS device, and the roll angle which is the rotational position around the first axis And a posture angle of the flying object from the pitch angle . The flying object according to claim 1 , wherein the movement control device controls the variable wing so that the flying object has a predetermined attitude angle based on the detected attitude angle of the flying object. The inertial navigation system includes an accelerometer for detecting acceleration in the direction of the third axis from said first, after the rotational speed of the first axis of the flying object is smaller than a second predetermined value, the acceleration The flying object according to claim 1, wherein a flying position of the flying object is detected from a total and an output of the GPS device. The inertial navigation device detects a rotational position around the first axis from a position where a rotational signal of the second and third axes of the gyro device that is a vibration wave passes a center level of the rotational signal. Item 1. A flying object according to item 1. The inertial navigation system, the moving average of the second and the rotation signal of the third shaft to calculate the gyro device, and calculates an approximate sine wave from the moving average calculated, passes the calculated approximate sine wave with said center level The flying object according to claim 4 , wherein a rotational position around the first axis is detected from a position to be operated. The inertial navigation device
The calculated rotational position around the first axis is the initial rotational position,
Calculating a ballistic pitch rate from the rotation rates of the gyro device around the second and third axes;
A pseudo rotation position is calculated by adding an integral value of a rotation rate around the first axis of the gyro device to the initial rotation position;
Calculate the pseudo pitch rate from the pseudo rotation position and the ballistic pitch rate,
Calculating a change in the difference between the pseudo pitch rate and the rotation rate of the gyro device around the second axis;
Calculate the peak value from the average of the calculated change in half cycle,
Calculate the rotational position error from the calculated peak value and the trajectory pitch rate,
The initial rotation position is newly set to reduce the rotation position error,
The flying object according to claim 4 , wherein the process is repeated using the new initial rotational position so that the rotational position error converges. The flying object according to claim 6 , wherein the inertial navigation device newly sets the initial rotational position so as to reduce the rotational position error by Kalman filter processing.

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