JP2843903B2 - Inertial navigation device - Google Patents

Description

DETAILED DESCRIPTION OF THE INVENTION

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to an inertial navigation system (hereinafter referred to as INS) for detecting an attitude angle of an aircraft, a vehicle, or the like.
In particular, the present invention relates to a highly accurate inertial navigation device that corrects a detection error due to a temporal deformation of a cushion interposed between a gyro / accelerometer package and an INS chassis, thereby improving accuracy.

[0002]

2. Description of the Related Art A conventional attitude angle detection INS will be described with reference to FIGS. A gyro / accelerometer package 3 to which a gyro 1 and an accelerometer 2 are attached is fixed in parallel on a bottom plate 6a of a chassis 6 of an INS 5 via a rubber cushion 4, and the chassis 6 is fixed to a body 7 such as an aircraft or a vehicle. . Although the accelerometer 2 is not damaged by vibration and impact, the gyro 1 is structurally weak, so that the gyro / accelerometer package 3 cannot be attached to the chassis 6 without the rubber cushion 4.

[0003] In the gyro 1, the body 7 of an aircraft or a vehicle is used.
The rotational angular velocity ωθ around the pitch axis (the reference axis from left to right orthogonal to the longitudinal direction of the aircraft on the horizontal plane) and the roll axis (the reference axis orthogonal to the pitch axis and traveling forward from the rear of the aircraft on the horizontal plane) The rotation angular velocity ωφ around is detected and input to the first attitude angle calculation unit 8. As shown in FIG.
In the accelerometer 2, when the body 7 before starting is stationary, i.e. components of the pitch angle theta ₁ direction of the gravitational acceleration G a ₁ and the roll angle phi ₁ direction component b ₁ is detected in a navigation preparation, the It is input to the two attitude angle calculator 9.

The second attitude angle calculator 9 uses the gravitational acceleration G to calculate the pitch angle θ ₁ (initial value of the pitch angle) and the roll angle φ ₁ (initial value of the roll angle) at rest before departure. It is calculated by the formula, input to the first attitude angle calculator 8, and stored in the memory. θ ₁ = sin ^-1 (a ₁ / G) (1) φ ₁ = sin ^-1 (b ₁ / G) (2)

[0005] The pitch angle and the roll angle are collectively called attitude angles. In the first posture angle calculating portion 8 is integral rotational angular velocity ωθ and ωφ than the starting point, the respective integrated value added to the pitch angle initial value theta ₁ or roll angle initial value phi _1, constantly changing current And the roll angle φ θ = ∫ωθdt + θ ₁ (3) φ = ∫ωφdt + φ ₁ (4) is calculated and output to the outside. The first and second posture angle calculators 8 and 9 constitute an operation unit 10.

[0006]

By the way, the conventional IN
In S, gyro / accelerometer package 3 and chassis 6
The rubber cushion 4 interposed between the bottom plate 6a and the bottom plate 6a gradually changes with time and gradually becomes harder and slightly deformed, so that the parallelism of the bottom surface 3a of the package 3 to the bottom plate 6a is deviated. For this reason, errors occur in the components a ₁ and b ₁ of the gravity G in the attitude angles θ ₁ and φ ₁ detected by the accelerometer 2, and thus errors occur in the initial attitude angles θ ₁ and φ ₁ . Eventually, an error occurs in the attitude angles θ and φ output from the INS, and the measurement accuracy has not been improved beyond 0.05 °.

An object of the present invention is to solve such a conventional drawback and to provide an inertial navigation apparatus with improved attitude angle measurement accuracy.

[0008]

In order to achieve the above object, an inertial navigation device according to the present invention comprises a gyro / accelerometer package, a correction accelerometer, and an arithmetic unit, wherein the arithmetic unit comprises: The gyro / accelerometer package includes first, second and third attitude angle calculators, a correction value calculator, and an attitude angle corrector, and accommodates the gyro and the accelerometer, and is fixed to the body. The gyro detects rotation angular velocities ωθ and ωφ around the pitch axis and the roll axis of the airframe and inputs the gyro to the first attitude angle calculation unit. , and the said accelerometer of said gyro accelerometer package detects the pitch angle and roll angle direction component of the gravitational acceleration G in the stationary state of the machine body, each of the detected values a ₁ and b ₁ The correction attitude accelerometer is directly attached to the chassis, and detects a pitch angle and a roll angle direction component of the gravitational acceleration G in a stationary state of the body, The detected values a ₂ and b ₂ are input to the third attitude angle calculation unit, and the second attitude angle calculation unit calculates the pitch angle θ ₁ = sin ^{−1 of the} aircraft in a stationary state of the aircraft. (A
₁ / G) and the roll angle φ ₁ = sin ⁻¹ (b ₁ / G), which are input to the first attitude angle calculation section and the correction value calculation section, and the third attitude angle calculation section Is the pitch angle θ ₂ = sin ⁻¹ (a ₂ /
G) and the roll angle φ ₂ = sin ⁻¹ (b ₂ / G),
The first attitude angle calculation unit inputs the correction value calculation unit to the correction value calculation unit.
t + θ ₁ and roll angle φ = ∫ωφdt + φ ₁ are calculated and input to the attitude angle correction unit. The correction value calculation unit calculates the error Δθ = θ ₁ −θ ₂ and Δφ = φ _{1 of the} attitude angle. −
are those to calculate the phi ₂ to enter into the attitude angle correcting unit, the posture angle correction unit includes a structure which calculates and outputs the pitch angle theta-[Delta] [theta] and the roll angle phi-[Delta] [phi has been corrected the error component Has become.

[0009]

In the inertial navigation device according to the present invention, the gyro in the gyro / accelerometer package detects the rotational angular velocities ωθ and ωφ around the pitch axis and the roll axis of the body to which the chassis of the inertial navigation device is fixed. To the first attitude angle calculator. The accelerometer in the gyro / accelerometer package detects the pitch angle and roll angle direction components of the gravitational acceleration G in a stationary state before the aircraft departs, and calculates the respective detected values a ₁ and b ₁ as a second. Input to attitude angle calculator. The accelerometer for correction directly attached to the chassis, when the aircraft is stationary before departure,
The pitch angle and roll angle direction components of the gravitational acceleration G are detected, and the detected values a ₂ and b ₂ are input to the third attitude angle calculation unit. The second attitude angle calculation unit calculates a pitch angle θ ₁ = sin ⁻¹ (a ₁ / G) and a roll angle φ ₁ = sin ⁻¹ (b ₁ / G) of the aircraft at rest before departure. And the first
Input to the attitude angle calculator and the correction value calculator. The third attitude angle calculation unit calculates the pitch angle θ ₂ = sin ⁻¹ (a ₂ / G) and the roll angle φ ₂ = sin ⁻¹ (b ₂ /
G) is calculated and input to the correction value calculator. The first
The attitude angle calculation unit calculates a pitch angle θ = 機 ω of the moving aircraft.
θdt + θ ₁ and inputs to calculate the roll angle φ = ∫ωφdt + φ ₁ to the posture angle correction unit. Then, the correction value calculation unit calculates the error Δθ = θ ₁ −θ ₂ and Δφ = φ ₁
The -.phi ₂ calculated input to the attitude angle correcting unit, the attitude angle correction unit outputs to the external computing a pitch angle theta-[Delta] [theta] and the roll angle phi-[Delta] [phi has been corrected the error component. By doing so, it is possible to correct the detection error due to the influence of the temporal deformation of the cushion interposed between the gyro / accelerometer package and the chassis of the inertial navigation device, thereby achieving higher accuracy.

[0010]

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS An embodiment of an inertial navigation system according to the present invention will be described below with reference to FIGS. In FIGS. 1 to 3, the same reference numerals are given to the portions corresponding to FIGS. 4 to 6, and the overlapping description will be omitted. In the present invention, a correction accelerometer 11 is added to the conventional configuration of FIG. 4 and is directly mounted on the bottom plate 6 a of the chassis 6. Third attitude angle calculation unit 1
2. A correction value calculation unit 13, an attitude angle correction unit 14, and a filter 15 are added. The first, second, and third posture angle calculation units 8, 9, 12, the correction value calculation unit 13, the posture angle correction unit 14, and the filter unit 15 constitute the calculation unit 10.

In the correction accelerometer 11, as shown in FIG. 3, similarly to the accelerometer 2 described in the conventional example, a component a ₂ of the gravitational acceleration G in the direction of the pitch angle θ _{2 and} a component of the gravitational acceleration G in the direction of the roll angle φ ₂ are used. The component b ₂ is detected and input to the third posture angle calculator 12. Since a rubber cushion is not used for mounting the correction accelerometer 11, the components a ₂ ,
error due to the rubber cushion as acceleration in b ₂ meter 2 detects output a _1, b ₁ is not included. Calculation unit 1
In the second example, similarly to the second attitude angle calculation unit 9 described in the conventional example,
The attitude angles (initial attitude angles) θ ₂ and φ ₂ of the aircraft at rest before departure shown by the following equations are detected and input to the correction value calculation unit 13. θ ₂ = sin ^-1 (a ₂ / G) (5) φ ₂ = sin ^-1 (b ₂ / G) (6)

On the other hand, initial posture angles θ ₁ and φ ₁ including errors caused by the rubber cushion 4 are input from the second posture angle calculation unit 9 to the first posture angle calculation unit 8 and the correction value calculation unit 13.
In the correction value calculation unit 13, the initial attitude angle theta _1, phi ₁ from the initial attitude angle theta _2, phi ₂ is subtracted respectively, theta _1, the error component contained in phi ₁ (correction amount) Δθ = θ ₁ -θ ₂ (7) Δφ = φ ₁ −φ ₂ (8) is calculated, and the posture angle correction unit 14
(However, the filter unit 15 may be omitted.) In the attitude angle correction unit 14, the first attitude angle calculation unit 8
The errors Δθ and Δφ are subtracted from the input posture angles θ and φ (the same as the conventional example in FIG. 4), and the posture angle θ ′ = θ−Δθ = ∫ not including the error due to the rubber cushion 4. ωθdt + θ ₁ −Δθ = ∫ωθdt + θ ₂ (9) φ ′ = φ−Δφ = ∫ωφdt + φ ₁ −Δφ = ∫ωφdt + φ ₂ (10) is calculated and output to the outside.

Since the detection values a ₂ and b ₂ of the correction accelerometer 11 include noise components due to the wind, the rotation of the engine of an aircraft, a vehicle or the like, and the vibration of the body 7 caused by the getting on and off of a person, the initial attitude angle Noise components are superimposed on θ ₂ , φ ₂ , and therefore on the correction amounts Δθ, Δφ. The filter section 15 is for removing this noise component.

Although the embodiments of the present invention have been described above, it is obvious to those skilled in the art that the present invention is not limited to the embodiments and various modifications and changes can be made within the scope of the claims. Would.

[0015]

As described above, according to the inertial navigation system of the present invention, the errors Δθ and Δφ of the initial attitude angle caused by the aging of the cushion used to mount the gyro / accelerometer package on the chassis are reduced. The detected attitude angles θ ′ and φ ′ can be measured more accurately by correcting the detected errors. Therefore, the conventional measurement accuracy of 0.05 ° was the limit, but according to the present invention, 0.02 ° can be achieved.

[Brief description of the drawings]

FIG. 1 is a block diagram showing an embodiment of an inertial navigation device according to the present invention.

FIG. 2 is a cross-sectional view showing a main structure of the embodiment.

FIG. 3 is a cross-sectional view for explaining a component in a posture angle direction of a gravitational acceleration G detected by a correction accelerometer used in the embodiment.

FIG. 4 is a block diagram of a conventional inertial navigation device for detecting an attitude angle.

FIG. 5 is a cross-sectional view showing a main structure of the conventional inertial navigation device of FIG.

6 is a cross-sectional view for explaining a component in a posture angle direction of a gravitational acceleration G detected by an accelerometer used in the conventional inertial navigation device of FIG.

[Explanation of symbols]

 Reference Signs List 1 gyro 2 accelerometer 3 gyro / accelerometer package 5 INS 8 first attitude angle calculator 9 second attitude angle calculator 10 calculator 11 correction accelerometer 12 third attitude angle calculator 13 correction value calculator 14 attitude angle Correction unit 15 Filter unit

──────────────────────────────────────────────────続 き Continued on front page (58) Field surveyed (Int.Cl. ⁶ , DB name) G01C 21/00 G01C 21/16

Claims (1)

(57) [Claims] 1. An inertial navigation device comprising a gyro / accelerometer package, a correction accelerometer, and a calculation unit, wherein the calculation unit includes first, second, and third attitude angle calculation units, and a correction value calculation unit. A gyro / accelerometer package, which accommodates the gyro and the accelerometer, and is attached to a chassis fixed to the body substantially in parallel via a cushion, The gyro detects the rotational angular velocities ωθ and ωφ around the pitch axis and the roll axis of the aircraft and inputs the detected angular velocity to the first attitude angle calculation unit.The accelerometer in the gyro / accelerometer package includes: to detect the pitch angle and roll angle direction component of the gravitational acceleration G in the stationary state of the body, which each detected value a ₁ and b ₁ input to the second posture angle calculating portion, the correction Accelerometer, attached directly to the chassis, the pitch angle of the gravitational acceleration G in the stationary state of the machine body and to detect the roll angle direction component, each of the detected values a ₂ and b ₂ the third posture angle calculator The second attitude angle calculation unit calculates a pitch angle θ ₁ = sin ⁻¹ (a ₁ / G) and a roll angle φ ₁ = sin ⁻¹ (b) of the aircraft in a stationary state of the aircraft. ₁ / G) and inputs it to the first attitude angle calculation section and the correction value calculation section. The third attitude angle calculation section calculates a pitch angle θ _{2 of the} aircraft in a stationary state of the aircraft. = Sin ^-1 (a ₂ / G) and the roll angle φ ₂ = sin ^-1 (b ₂ / G) are input to the correction value calculation unit, and the first attitude angle calculation unit is , The pitch angle θ of the aircraft being moved = 移動 ωθdt + θ ₁ and the roll angle φ = ∫ωφdt +
The correction value calculation unit calculates φ ₁ and inputs the error to the attitude angle correction unit. The correction value calculation unit calculates Δθ = θ ₁ −θ ₂ and Δφ = φ ₁ −φ ₂ for the error of the attitude angle. The attitude angle correction unit is configured to input the angle to the angle correction unit.
An inertial navigation device for calculating and outputting −Δθ and roll angle φ−Δφ.