JP2018169368A - Flight vehicle-purpose navigation device, and flight vehicle-purpose navigation - Google Patents

Abstract

To provide a flight vehicle-purpose navigation device that can properly correct posture errors even when using a general-purpose inertial measurement device of low accuracy.SOLUTION: A flight vehicle-purpose navigation device, which is mounted to a fight vehicle, comprises: a detection device that detects inertia force acting on the flight vehicle; a first estimation unit that estimates a first error serving as at least a part of error ingredients between an initial setting posture of the detection device and a mounting posture relative to the flight vehicle on the basis of a result of a correction by gravity acceleration with respect to acceleration detected by the detection device when the flight vehicle stands still; and a positioning unit that derives a position and velocity of the flight vehicle on the basis of the initial setting posture in which the first error estimated by the first estimation unit is removed.SELECTED DRAWING: Figure 5

Description

  The present invention relates to an aircraft navigation apparatus and an aircraft navigation.

  In recent years, research on navigation sensors for measuring the position and speed of a flying object such as a rocket with high accuracy has been underway (see Patent Documents 1 and 2). This navigation sensor includes an inertial sensor as a sensor source, and is attached to the flying object in a preset posture for calculating position and velocity. However, if there is an error (hereinafter referred to as “attitude error”) between the attitude set in advance for calculating the position and speed and the actual mounting attitude of the navigation sensor, the position and speed of the aircraft will be It cannot be measured accurately.

  In order to eliminate the influence of the above-mentioned attitude error, a method for estimating the attitude error of the navigation sensor and correcting it is known. For example, in a navigation sensor for a rocket, the attitude error is separated into three components in the axial direction of the roll axis, pitch axis, and yaw axis of the flying object. There is known a method of correcting using an acceleration direction vector and correcting an attitude error around the roll axis using an earth rotation angular velocity (earth rate).

U.S. Pat. No. 8,868,258 US Pat. No. 5,739,787

  When correction using the direction vector of the gravitational acceleration is performed for the posture error about the pitch axis and the yaw axis, the accuracy of the correction depends on the accuracy of the inertial sensor. If the accuracy of the inertial sensor is poor, an error occurs in the direction vector itself of the gravitational acceleration detected by the inertial sensor. For this reason, when correction is performed using the direction vector of gravitational acceleration including an error, the posture error cannot be appropriately removed. The accuracy of the inertial sensor depends on the accuracy of the built-in accelerometer and gyroscope, and the higher the accuracy, the higher the cost. Because of the low cost, relatively low-precision consumer parts are often used as inertial sensors, so correction using the gravitational acceleration direction vector cannot accurately measure the position and velocity of the flying object. .

  Also, when correcting the attitude error around the roll axis using the earth rate, an expensive inertial sensor that can be used in the flight environment and can detect the earth rate with the accuracy required for the correction is required. Become. A low-priced inertial sensor is difficult to achieve both the ability to be used in severe flight environments and the ability to detect the earth rate with the accuracy required for correction. Therefore, there is a need for a method for correcting an attitude error that does not use an earth rate.

  The present invention has been made in consideration of such circumstances, and is a navigation system for an aircraft that can appropriately correct an attitude error even when a general-purpose low-precision inertial measurement device is used. Another object is to provide navigation for a vehicle.

  One aspect of the present invention is a navigation apparatus for an aircraft mounted on an aircraft, the detection device detecting an inertial force acting on the aircraft, and detected by the detection device when the aircraft is stationary. A first estimation unit that estimates a first error that is at least a part of an error component between an initial setting posture and an attachment posture of the detection device with respect to the flying object based on a result of correction by acceleration due to gravitational acceleration; A positioning unit that derives the position and velocity of the flying object based on the initial setting posture from which the first error estimated by the first estimating unit is removed.

  The present invention can appropriately correct the posture error even when a general-purpose low-precision inertial measurement device is used.

1 is a schematic configuration diagram of a rocket 1. FIG. It is sectional drawing in the A-A 'direction of the rocket 1 shown by FIG. 1 is a configuration diagram of an aircraft navigation system 100. FIG. 2 is a configuration diagram of a radio navigation positioning module 110. FIG. 3 is a configuration diagram of a navigation calculation module 130. FIG. 5 is a flowchart illustrating an example of leveling processing in a navigation calculation module 130. It is a figure which shows pitch angle error (epsilon) (theta) _P in an ENU coordinate system. It is a figure showing yaw angle error εθ _Y in the ENU coordinate system. 5 is a flowchart illustrating an example of azimuth angle estimation processing in a navigation calculation module 130. It is a figure showing roll angle error εθ _R in the ENU coordinate system. It is a figure which shows the azimuth angle error (theta) _D in an ENU coordinate system. It is a figure which shows the simulation result of a leveling process.

  DESCRIPTION OF EMBODIMENTS Hereinafter, embodiments of an aircraft navigation apparatus and an aircraft navigation according to the present invention will be described with reference to the drawings. The flying object navigation apparatus is an apparatus mounted on a flying object such as a rocket, an artificial satellite, or a space probe. In the following description, the flying object is described as a rocket as an example, but may be an artificial satellite or a space probe as described above. Hereinafter, the configuration and function of such an apparatus will be disclosed step by step.

[rocket]
FIG. 1 is a schematic configuration diagram of the rocket 1. The rocket 1 is, for example, a multi-stage rocket. A satellite is stored in the uppermost stage, and the lower stages are separated from the uppermost stage and then land on the sea or land on the ground. The rocket 1 includes, for example, a first stage rocket 10, a second stage rocket 12, and an uppermost third stage rocket 14. Immediately after being launched, the rocket 1 flies with the propulsive force of the first-stage rocket 10, and then the first-stage rocket 10 is separated and then flies with the propulsive force of the second-stage rocket 12. 14 makes the payload 20 such as the artificial satellite reach the outer space. The flying object navigation apparatus 100 is stored in the second stage rocket 12, for example.

[Aircraft navigation system]
The inertial measurement device included in the flying object navigation device 100 needs to be accurately mounted in the initial setting posture set in advance in the second stage rocket 12 in order to measure the position and speed of the rocket 1 with high accuracy. There is. However, there may be a posture error between the actual mounting posture of the inertial measurement device and the initial setting posture set in advance for calculating the position and speed of the flying object navigation device 100. The flying object navigation apparatus 100 has a function for correcting this attitude error.

  FIG. 2 is a cross-sectional view of the rocket 1 shown in FIG. 1 in the A-A ′ direction. As shown in FIG. 2, the navigation system 100 for an aircraft detects attitude errors in terms of a roll axis that is an axis in the traveling direction of the rocket 1, a pitch axis that is a horizontal axis of the rocket 1, and the vertical direction of the rocket 1. The correction is performed by separating into three axial components of the yaw axis, which is the first axis. For example, the aircraft navigation apparatus 100 launches the rocket 1 with respect to attitude errors around the pitch axis (hereinafter referred to as “pitch angle errors”) and attitude errors around the yaw axis (hereinafter referred to as “yaw angle errors”). Correction is performed by the leveling process performed at the previous stationary time. On the other hand, the attitude error around the roll axis (hereinafter referred to as “roll angle error”) is corrected by azimuth angle estimation processing performed during the flight after the launch of the rocket 1. Details of the leveling process and the azimuth angle estimation process will be described later.

  FIG. 3 is a configuration diagram of the aircraft navigation apparatus 100. The aircraft navigation apparatus 100 is connected to one or more antennas AT and the communication apparatus TM. The aircraft navigation apparatus 100 includes, for example, a radio navigation positioning module 110, an IMU (Inertial Measurement Unit) 120, a navigation calculation module 130, a power supply module 140, and a storage unit 150.

  The antenna AT receives radio waves from the outside and outputs a signal corresponding to the received radio waves (hereinafter referred to as satellite positioning signal) to an LNA (Low Noise Amplifier). The LNA amplifies the satellite positioning signal input from the antenna AT.

  The radio navigation positioning module 110 derives the position and velocity of the rocket 1 based on the satellite positioning signal amplified by the LNA, and outputs it to the navigation calculation module 130.

  The IMU 120 includes, for example, a triaxial acceleration sensor configured by MEMS (Micro Electro Mechanical Systems) or an optical fiber, and a triaxial gyro sensor. The IMU 120 detects the inertial force that acts on the rocket 1. The IMU 120 outputs values detected by these sensors (horizontal, vertical, and depth accelerations, pitch, roll, and yaw rates, etc.) to the navigation calculation module 130. The IMU 120 is an example of a “detection device”.

  The navigation calculation module 130 corrects a posture error between the actual mounting posture of the IMU 120 and the initial setting posture. The navigation calculation module 130 derives the position and speed of the rocket 1 based on the detection value (detection result) detected by the IMU 120. The navigation calculation module 130 integrates the position and speed derived by the radio navigation positioning module 110 and the position and speed derived using the detection result by the IMU 120 to derive the position and speed of the rocket 1.

  The power supply module 140 is connected to a power supply device (not shown). The power supply module 140 includes, for example, a protection circuit and a DC-DC converter, and supplies power to each part of the aircraft navigation apparatus 100.

  The storage unit 150 stores initial setting attitude data of the IMU 120 of the flying object navigation apparatus 100 used for calculation when deriving the position and velocity of the rocket 1. This initial setting posture predefines the mounting posture of the IMU 120 with respect to the rocket 1. When the navigation calculation module 130 derives the position and speed of the rocket 1 based on the detection values detected by the IMU 120, the navigation calculation module 130 calculates the position and speed on the assumption that the IMU 120 is attached to the rocket 1 in the initial setting posture. To do.

  The storage unit 150 also includes a satellite positioning signal input to the radio navigation positioning module 110, detection values detected by the IMU 120, various calculation results derived by the radio navigation positioning module 110 and the navigation calculation module 130 (of the rocket 1). (Position or speed) etc. The storage unit 150 is realized by, for example, a ROM (Read Only Memory), a RAM, an EEPROM (Electrically Erasable Programmable Read Only Memory), and the like. The storage unit 150 may be realized by an HDD (Hard Disc Drive), a flash memory, or the like.

  The communication device TM transmits information to a ground monitoring device (not shown) using, for example, a telemeter line. For example, the ground monitoring device acquires the position and speed of the rocket 1 after the rocket 1 is launched (takes off) from the rocket 1 by telemetry communication. Then, the ground monitoring device repeatedly estimates a drop point when the rocket 1 falls, and transmits an instruction signal to the rocket 1 so as to avoid a danger due to the fall.

[Radio navigation positioning module]
FIG. 4 is a configuration diagram of the radio navigation positioning module 110. The radio navigation positioning module 110 includes a positioning side interface 112 and a radio navigation positioning unit 114. The satellite positioning signal received by the antenna AT is amplified by the LNA and input to the positioning-side interface 112. The positioning side interface 112 is, for example, an RF (Radio Frequency) interface.

  The radio navigation positioning unit 114 performs processing on the satellite positioning signal input to the positioning side interface 112 to derive the position and velocity of the rocket 1 and outputs the position and velocity to the navigation calculation module 130. The radio navigation positioning unit 114 obtains a relative distance (shoot range) to the artificial satellite based on a satellite positioning signal corresponding to each of radio waves received from four or more artificial satellites by the antenna AT. The position of is calculated.

[Navigation calculation module]
FIG. 5 is a configuration diagram of the navigation calculation module 130. The navigation calculation module 130 includes, for example, a leveling unit 132, an azimuth angle estimation unit 134, and an inertial navigation positioning unit 136.

  The leveling unit 132 estimates and removes the pitch angle error and the yaw angle error in the attitude error of the aircraft navigation device 100 when the aircraft is stationary before the launch of the rocket 1. The leveling unit 132, based on the result of correction by the gravitational acceleration with respect to the acceleration detected by the IMU 120 when the rocket 1 is stationary, components (pitch, yaw, roll of the pitch) constituting the error between the initial setting posture and the mounting posture of the IMU 120. An error (an example of a first error) around the pitch axis and the yaw axis, which is at least a part of the error components around each axis), is estimated.

  When an attitude error has occurred, a correction residual occurs in the acceleration that should be “0” after correction by gravitational acceleration. This is because a gravity correction residual is generated by performing gravity correction in the wrong vector direction due to the posture error. Therefore, the leveling unit 132 integrates the correction residual of the acceleration corrected by the gravitational acceleration, and obtains the attitude angle at which the calculated speed becomes “0”, so that the pitch angle of the aircraft navigation apparatus 100 is obtained. Estimate and remove errors and yaw angle errors. For example, the leveling unit 132 constructs a linear approximation expression using an attitude error as a parameter on the IMU 120 coordinate system (ENU (East North Up) coordinate system), and applies a Kalman filter to the correction result by gravity acceleration. Pitch angle error and yaw angle error are estimated and removed. The leveling unit 132 is an example of a “first estimation unit”.

  The azimuth angle estimation unit 134 estimates and removes the roll angle error in the attitude error of the flying object navigation apparatus 100 during the flight after the launch of the rocket 1. The azimuth estimation unit 134 is a roll that is at least a part of components that constitute an error between the initial setting posture and the mounting posture based on the detection result of the IMU 120 and the radio wave received from the satellite during the flight of the rocket 1. An error around the axis (an example of a second error) is estimated. The azimuth angle estimation unit 134 uses the fact that a horizontal component with respect to the ground surface is generated in the velocity vector in the traveling direction of the aircraft at the timing when the rocket 1 gradually changes its attitude in the horizontal direction during the flight after launch. The velocity derived by the radio navigation positioning module 110 based on the satellite positioning signal is not affected by the azimuth error (roll angle error). Therefore, the azimuth estimation unit 134 refers to the velocity vector derived by the radio navigation positioning module 110 as positive data, and compares this velocity vector with the velocity vector calculated based on the detection value of the IMU 120. Thus, the roll angle error is estimated.

  For example, the azimuth angle estimation unit 134 includes the horizontal velocity vector of the rocket 1 body obtained by integrating the detection values of the three-axis acceleration sensor of the IMU 120, and the velocity vector derived by the radio navigation positioning module 110. Are respectively passed through a second-order digital low-pass filter, and their inner product and outer product are obtained. Then, the azimuth angle estimation unit 134 estimates the magnitude of the error angle from the inner product, and estimates and removes the posture error by estimating the error direction from the outer product. The azimuth angle estimation unit 134 is an example of a “second estimation unit”.

  The angle formed by these two velocity vectors includes only the roll angle error when the pitch angle error and the yaw angle error are removed by the leveling unit 132 before the launch of the rocket 1. Since the IMU 120 accumulates a bias error with time, the azimuth angle estimation unit 134 performs the above azimuth angle estimation process at a timing at which the attitude of the rocket 1 gradually changes in the horizontal direction.

  The inertial navigation positioning unit 136 detects the rocket based on the detected value detected by the IMU 120, the attitude error estimated by the leveling unit 132 and / or the initial setting attitude from which the attitude error estimated by the azimuth angle estimating unit 134 is removed. The position and velocity of 1 are derived. The inertial navigation positioning unit 136 calculates the speed by integrating the value input from the triaxial acceleration sensor of the IMU 120, and further calculates the displacement (position) by integrating the speed. The inertial navigation positioning unit 136 calculates a more accurate posture by combining displacement and direction vectors based on values input from the triaxial acceleration sensor and the triaxial gyro sensor of the IMU 120. The inertial navigation positioning unit 136 is an example of a “positioning unit”.

  The inertial navigation positioning unit 136 integrates the position and speed of the rocket 1 derived based on the detection value detected by the IMU 120 and the position and speed of the rocket 1 derived by the radio navigation positioning module 110, The position and speed of the rocket 1 are derived. The inertial navigation positioning unit 136 integrates both by applying a Kalman filter, for example, and derives the position and velocity of the rocket 1. In a situation where the radio navigation positioning module 110 cannot receive the satellite radio waves with the antenna AT due to the influence of multipath, the inertial navigation positioning unit 136 determines the position and velocity of the rocket 1 derived by the radio navigation positioning module 110 until immediately before. Is supplemented by the position and speed of the rocket 1 derived based on the detection value detected by the IMU 120.

  Each of the radio navigation positioning module 110 and the navigation calculation module 130 is configured by, for example, an MPU (Micro Processing Unit). The radio navigation positioning module 110 implements the radio navigation positioning unit 114 by executing a program stored in a program memory (not shown), for example. Moreover, the navigation calculation module 130 implement | achieves the leveling part 132, the azimuth angle estimation part 134, and the inertial navigation positioning part 136 by running the program stored in the program memory (not shown), for example. Some or all of these functional units may be realized by hardware such as LSI (Large Scale Integration) or ASIC (Application Specific Integrated Circuit), or the software and hardware cooperate with each other. It may be realized.

[Leveling process]
Hereinafter, the leveling process performed by the leveling unit 132 of the navigation calculation module 130 will be described. When an attitude error occurs between the mounting attitude of the IMU 120 and the preset initial setting attitude with respect to the E axis and the N axis of the ENU coordinate system, in the calculation in the navigation calculation module 130, after gravity correction, A correction residual is generated in the acceleration that should be 0 ”. By obtaining an attitude angle such that the speed generated by the integration of the acceleration correction residual after the gravity correction is “0”, the true attachment attitude with respect to the ENU coordinate system can be obtained.

FIG. 6 is a flowchart illustrating an example of leveling processing in the navigation calculation module 130. First, the leveling unit 132 calculates a correction residual of acceleration after gravity correction caused by the attitude error of the ENU coordinate system based on the detection value detected by the IMU 120 (step S101). For example, the leveling unit 132 calculates a correction residual of acceleration after gravity correction caused by an attitude error of the ENU coordinate system by the following equations (1) to (3). In Expressions (1) to (3), εa _E , εa _N , and εa _U are components of the E axis, N axis, and U axis of the gravity-corrected acceleration of the ENU coordinate system, respectively, and εθ _P and εθ _Y are respectively The pitch angle error and yaw angle error of the rocket 1 body axis coordinate system. The roll direction is the U-axis direction.

FIG. 7 is a diagram showing the pitch angle error εθ _P in the ENU coordinate system. FIG. 8 is a diagram showing the yaw angle error εθ _Y in the ENU coordinate system. 7 and 8, the actual mounting posture of the IMU 120 is indicated by a broken line (IMU 120A), and the calculated initial setting posture of the IMU 120 is indicated by a solid line (IMU 120B).

Here, in the above formulas (1) to (3), if εθ _P and εθ _Y are very small, the above formulas (1) and (2) can be approximated by the following formulas (4) and (5). .

Next, when εθ _P and εθ _Y are substantially the same value and are very small, the above equation (3) can be approximated by the following equation (6).

Here, if εθ _P and εθ _Y are almost the same value, the above equation (6) can be approximated by the following equation (7).

Furthermore, when α _p and α _Y are posture estimation coefficients, the above equation (7) is expressed as the following equation (8). For α _p and α _Y , target posture accuracy (target posture angle estimation error) of the pitch axis and yaw axis is set.

Next, the leveling unit 132 calculates a speed error εv _L using the calculated acceleration correction residual (step S103). The speed error εv _L is expressed by the following equation (9).

Next, the leveling unit 132 applies a Kalman filter to the calculated speed error, and estimates and removes the posture error (step S105). The leveling unit 132 estimates the attitude error with the Kalman filter using the above equation (9) and the condition that the velocity of the ENU coordinate system when the aircraft is stationary before launching the rocket 1 is “0”, with the speed increment as an observed value. To do. The update cycle of the Kalman filter is set to, for example, the update cycle 100 Hz of the flying object navigation device 100. The state quantity x _L , the transition matrix A _L , and the observation matrix C _L are expressed by the following equation (10).

Regarding the process noise Q _L , the dispersion of the acceleration noise of the IMU 120 and the observation noise R _L may be determined by the fluctuation value of the acceleration due to the motion of the aircraft when the rocket 1 is stationary detected by the IMU 120. Also, the initial value of the covariance value P _L may be determined from the attitude error or the like is assumed. By updating the quaternion of the body axis posture with the estimated pitch angle error εθ _P and yaw angle error εθ _Y , the posture is corrected over time and finally converges to a constant value. . Thereby, the leveling unit 132 estimates and removes the pitch angle error and the yaw angle error.

  Next, the leveling unit 132 causes the storage unit 150 to store the data from which the pitch angle error and the yaw angle error have been removed as data of the true mounting posture (step S107). Thus, the process of this flowchart is completed.

[Azimuth estimation processing]
Hereinafter, the azimuth angle estimation process performed by the azimuth angle estimation unit 134 of the navigation calculation module 130 will be described. The azimuth angle estimation uses velocity matching, and in the ENU coordinate system with the launch point of the rocket 1 as the origin, the velocity vector of the rocket 1 derived by the radio navigation positioning module 110 and the detected value detected by the IMU 120 Based on the inner product and outer product of the velocity vector of the rocket 1 derived based on the above, the angle formed by these two vectors is calculated as an azimuth error. The angle formed by these two vectors includes only the roll angle error when the pitch angle error and the yaw angle error are removed by the leveling unit 132 before the launch of the rocket 1.

  FIG. 9 is a flowchart illustrating an example of the azimuth angle estimation process in the navigation calculation module 130. First, the azimuth estimation unit 134 determines whether or not a horizontal component with respect to the ground surface is generated in the velocity vector of the rocket 1 based on the detection value detected by the IMU 120 during the flight after the launch of the rocket 1 ( Step S201). If the azimuth angle estimation unit 134 determines that the horizontal component of the velocity vector is not generated with respect to the ground surface, the azimuth angle estimation unit 134 continues monitoring the velocity vector.

  On the other hand, when the azimuth angle estimation unit 134 determines that a horizontal component with respect to the ground surface is generated in the velocity vector, that is, when it is determined that it is time to gradually change the attitude of the rocket 1 in the horizontal direction, the IMU 120 The horizontal velocity vector (hereinafter referred to as “first velocity vector”) of the rocket 1 obtained by integrating the output of the three-axis accelerometer and the rocket 1 derived by the radio navigation positioning module 110 The two velocity vectors, the horizontal velocity vector of the aircraft (hereinafter referred to as “second velocity vector”), are individually filtered by a second-order digital low-pass filter to remove noise (step S203). Thereby, random noise included in the detection value of the IMU 120, thermal noise included in the satellite positioning signal, and the like can be removed.

FIG. 10 is a diagram showing the roll angle error εθ _R in the ENU coordinate system. In FIG. 10, the actual mounting posture of the IMU 120 is indicated by a broken line (IMU 120A), and the initial setting posture for calculation of the IMU 120 is indicated by a solid line (IMU 120B).

  Next, the azimuth angle estimation unit 134 calculates the inner product of the first speed vector and the second speed vector from which noise has been removed (step S205). The azimuth angle estimation unit 134 estimates the angle between the first speed vector and the second speed vector, that is, the magnitude of the error angle, based on the calculated inner product.

  Next, the azimuth angle estimation unit 134 calculates the outer product of the first speed vector and the second speed vector from which noise has been removed (step S207). The azimuth angle estimation unit 134 estimates the error direction between the first speed vector and the second speed vector, that is, in which direction the roll angle error is generated, based on the calculated outer product.

Figure 11 is a diagram showing an azimuth angle error theta _D in ENU coordinate system. As shown in FIG. 11, a first velocity vector, the angle formed between the second velocity vector becomes azimuth error theta _D.

  Next, the azimuth angle estimation unit 134 estimates a roll angle error based on the estimated error angle size and error direction (step S209).

  Next, the leveling unit 132 removes the roll angle error from the calculation initial setting posture data (the mounting posture data from which the pitch angle error and the yaw angle error have been removed by the leveling unit 132) stored in the storage unit 150. Then, the data after the removal is stored in the storage unit 150 as data of the true mounting posture (step S211). Thus, the process of this flowchart is completed.

[Simulation test]
The applicant of the present application conducted a simulation test described below. FIG. 12 shows a simulation result of the leveling process when an error angle of 10 degrees is given to each of the pitch angle direction and the yaw angle direction of the airframe. From the simulation results, it was found that the error angle converged to 0 degrees within 20 seconds.

If the posture error is larger than the assumed value, it may take time to converge. In this case, the parameters of the Kalman filter including α _P and α _Y are made variable, and the parameters are set so that rough posture estimation is performed at the initial stage of posture estimation in order to shorten the time required for convergence. Then, by confirming the convergence state and changing the parameters of the Kalman filter step by step so as to improve the estimation accuracy, the time required for convergence can be shortened while maintaining the accuracy of posture estimation.

  According to the flying object navigation apparatus 100 and the flying object navigation of the embodiment described above, even when a general-purpose relatively low-precision inertial measurement apparatus is used by performing the leveling process as described above. The pitch angle error and the yaw angle error can be appropriately removed. Thereby, the navigation system 100 for a vehicle can measure a highly accurate position and speed required for the flight safety control of the rocket 1. Furthermore, the acceleration fluctuation due to the movement of the airframe when the rocket 1 is standing up and the acceleration noise of the inertial measurement device can also be reduced, and the accuracy of the position and speed measurement values can be further improved.

  Further, by performing the azimuth angle estimation process as described above, it is possible to appropriately remove the roll angle error even when a general-purpose relatively low-precision inertial measurement device that cannot detect the earth rate is used. . Furthermore, the velocity vector of the rocket 1 derived based on the detection value detected by the inertial measurement device and the velocity vector of the rocket 1 derived by the radio navigation positioning module 110 are individually filtered by a secondary digital low-pass filter. Therefore, it is possible to prevent a speed error due to random noise included in the detection value of the inertial measurement device and a speed error due to thermal noise included in the satellite positioning signal.

  As mentioned above, although the form for implementing this invention was demonstrated using embodiment, this invention is not limited to such embodiment at all, In the range which does not deviate from the summary of this invention, various deformation | transformation and substitution Can be added.

AT ... Antenna, 1 ... Rocket, 100 ... Aircraft navigation device, 110 ... Radio navigation positioning module, 112 ... Positioning side interface, 114 ... Radio navigation positioning unit, 120 ... IMU, 130 ... Navigation calculation module, 132 ... Leveling unit , 134 ... azimuth angle estimation unit, 136 ... inertial navigation positioning unit, 140 ... power supply module, 150 ... storage unit, TM ... communication device

Claims (9)

A navigation device for an aircraft mounted on an aircraft,
A detection device for detecting an inertial force acting on the flying object;
Based on the result of correction by gravitational acceleration with respect to the acceleration detected by the detection device when the flying object is stationary, at least a part of error components between the initial setting posture and the mounting posture of the detection device with respect to the flying object. A first estimation unit for estimating a certain first error;
A flying object navigation apparatus comprising: a positioning unit that derives a position and a velocity of the flying object based on the initial setting posture from which the first error estimated by the first estimating unit is removed. A second error that is at least a part of an error component between the initial setting posture and the mounting posture is estimated based on a detection result of the detection device and a radio wave received from a satellite during the flight of the flying object. 2 further includes an estimation unit,
The positioning unit is configured to determine the position of the flying object based on the initial setting posture from which the first error estimated by the first estimation unit and the second error estimated by the second estimation unit are removed, and Deriving speed,
The aircraft navigation apparatus according to claim 1. The first error is an error around the pitch axis and the yaw axis of the aircraft.
The aircraft navigation system according to claim 1 or 2. The first estimation unit estimates the first error by applying a Kalman filter to the correction result by the gravitational acceleration,
The flying object navigation apparatus according to any one of claims 1 to 3. The second error is an error around the roll axis of the aircraft.
The navigation apparatus for an aircraft according to claim 2. The second estimation unit includes an inner product and an outer product of a first velocity vector of the flying object calculated from a detection result of the detection device and a second velocity vector of the flying object calculated from a radio wave received from the satellite. Estimating the second error based on:
The navigation apparatus for an aircraft according to claim 5. The first velocity vector and the second velocity vector are velocity vectors in the horizontal direction with respect to the ground surface.
The aircraft navigation apparatus according to claim 6. A vehicle navigation system using a vehicle navigation system equipped with a detection device that is mounted on a vehicle and that detects an inertial force acting on the vehicle.
Estimating a first error which is at least a part of an error component between the detection device initial setting posture and the mounting posture with respect to the flying object based on a result of correction by gravitational acceleration with respect to the acceleration detected by the detection device. When,
Deriving the position and velocity of the flying object based on the initial setting posture from which the estimated first error is removed;
Aircraft navigation. Estimating a second error that is at least a part of an error component between the initial setting posture and the mounting posture based on a detection result of the detection device and a radio wave received from a satellite during the flight of the flying object. When,
Deriving the position and velocity of the flying object based on the initial setting posture from which the estimated first error and the second error are removed;
The vehicle navigation according to claim 8, further comprising:

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