JPS5866120A - Attitude controller for flying object - Google Patents

Abstract

PURPOSE:To lead a flying material to its destination accurately, by calculating the attitude angle of the flying object in an acceleration time on a basis of the angular velocity detected by a rate gyro and controlling the flying object so that this attitude angle becomes a set reference attitude angle. CONSTITUTION:Detection signals of angular velocities (p), (q), and (r) around (x), (y), and (z) axes of the machine body coordinate system which are detected by rate gyros 1, 2, and 3 are selected successively by a multiplexer 5 and are converted to digital quantities by an A/D converter 6. A microcomputer 7 performs the coordinate conversion from angular velocities (p), (q), and (r) to angular velocities of the inertia coordinate system and performs the time integration for an acceleration time (motor combustion time) of a rocket to obtain a roll angle, a pitch angle, and a yaw angle of the rocket in the inertia coordinate system. The microcomputer 7 compares this attitude angle of the rocket with a set reference attitude angle to issue a control signal to an attitude controller 9.

Description

[Detailed Description of the Invention] This invention calculates the attitude angle of a flying object during flight, particularly during the acceleration period, based on the angular velocity detected by the rate gyro; The present invention relates to an attitude control device that accurately guides a flying object to a destination by controlling the attitude angle so that the attitude angle is the same.

Conventionally, as a means for making a flying object (such as a rocket) reach its destination, there have been two methods: a free method in which only the initial attitude angle, that is, the launch angle, is set based on trajectory calculation, and no guidance is required during flight; There are known systems that perform highly precise guidance such as homing over the entire flight period. The former is
Although it is possible to reduce the cost and reduce the size of the eaves, it is difficult to accurately reach the destination because disturbances such as surface winds and nozzle misalignment easily cause deviations from the planned trajectory. The latter allows accurate guidance to a destination, but because it uses a rate-integrating gyro, the device is expensive, complicated, and large, making it difficult to equip a small flying object at a low cost.

This invention was made with attention to the above points, and it is inexpensive,
It is an object of the present invention to provide an attitude control device for a flying object, which has a lightweight and compact configuration and can make the flying object reach its destination with high precision.

As described above, the object of the present invention is achieved by controlling the attitude of the flying object during the period when the flying object is accelerated after being launched, that is, during motor combustion. This is because the main reason why a flying object deviates from its planned trajectory is that the mounting of the flying object on the launcher is affected by factors such as attitude and nozzle misalignment, which have a significant effect during the initial period of flight, especially during the acceleration period. , Therefore, this is based on the fact that if the attitude angle is properly controlled during motor combustion, deviation from the orbit can be minimized and the accuracy of the flying object's arrival can be greatly improved. In order to achieve this object, the attitude control device of the present invention includes a small and lightweight rate gyro that detects the angular velocity in the aircraft coordinate system set for the flying object, and an attitude control system that detects the attitude of the flying object based on the angular velocity detected by the rate gyro. It includes a computing unit that calculates the angle and compares it with a preset reference attitude angle and outputs a control signal, and an actuator that generates an attitude control force in response to control of the computing unit (C8),
The attitude angle of the flying object is controlled according to a preset reference attitude angle.

The present invention will be explained below based on the drawings.

Figures 1.2 and 3.4 are diagrams showing an embodiment of the present invention, Figure 1 is a block diagram of the attitude control device of this embodiment, and Figure 2 is a body coordinate system and an inertial coordinate system. shows the relationship between
Fig. 3 shows the entire calculated trajectory of the rocket, and Fig. 4 shows the program for the reference pitch angle θC. In Figure 1, (
1) (21 (3) indicates three rate gyros. These rate gyros + 11 (2 + (31) are 3 of the aircraft coordinate system installed on the rocket and set on the aircraft with the rocket's center of gravity as the origin O. two orthogonal coordinate axes (x shown in Figure 2)
, y, and z axes, where the Z axis is the axial direction of the rocket. )
Angular velocity of rotation p (roll rate), q (pitch rate)
, r (yaw rate), respectively. (4) indicates an arithmetic unit. The processor ngi(4) consists of an input section, an arithmetic section, and an output section.

The input section includes a multiplexer (5) that sequentially selects detection signals of angular velocities p and qlr around the aircraft coordinate system X, y, and z axes detected by the rate gyro (11 (21 (31)), and It has an A/D converter (6) that converts the signal from an analog quantity to a digital quantity.The arithmetic unit is composed of a microcomputer (7),
The color difference Up, ci, r is the inertial coordinate thread (x, y shown in Figure 2)
, z-axis, here the z-axis.

The Y axis is horizontal, and the Z axis is vertically downward. Note that the launch point (S) and the destination point (■) shown in FIG. 3 are on the X axis, and are assumed to be several tens of kilometers in size. ) angular velocity;
Coordinates are transformed to t,,b,small. In this example,
The inertial coordinate system (x, y, z axes) is transformed into the aircraft coordinate system (X, )',
As shown in Figure 2, the order of transfer to
, among the Z axes, first rotate Z@ by ψ, move the X and Y axes to Y' and Y', respectively, and then rotate the Y' axis by θ.

Let the Y' and Z axes be X' (x @ ) and 2, respectively.
", then rotate the X"←) axis by φ and change the zl axis to Z.
Since the order of transfer to siqu is adopted, the angular velocity p
, q, r are angular velocities; t=, a, coordinates are transformed (Euler transformation) using the following formula.

i=p+(qle nφ+r songφ) −〇 Mi-qenφ-r＆nφ Te = (q lesson φ + r plane φ) recitation θ And the microcomputer (7) has a, δ.

φ=fφdt θ=fθdt ψ=fψdt Attitude angle φ (roll angle) of the rocket in the inertial WrJ frame
, θ (pitch θ) and ψ (yaw angle) are determined. Next, the microcomputer (7) calculates the attitude angle θ of this rocket.
, ψ is the preset reference attitude angle θ. ,ψ.

Compared to these errors Δ0=θ−θ.

△ψtwoψ−ψ.

If is above a certain value, a control signal is output.

In this embodiment, the roll angle φ is not controlled and is left free, and the pitch angle θ changes in size over time during the control time (T) (equal to the motor combustion time), as shown in FIG. The yaw angle ψ is controlled according to a pitch angle program in which the angle changes from θC1 to θc2, and the yaw angle ψ is constant over the control time (T). It is controlled to 20.

The output section receives control signals from the driver (8) and the microcomputer (7) to control an attitude controller, which will be described later. The attitude controller (9) receives a control signal from the driver (8) and generates an attitude control force.

This posture control device (9) is provided with a plurality of tabs arranged in the circumferential direction at a predetermined distance from each other at the rear end of the nozzle so as to be able to enter and retract into the nozzle, and any tab can be inserted into the nozzle. A jet tab type that protrudes into the flowing combustion gas flow and thereby blocks the combustion gas flow to generate attitude control force, or a plurality of jet tabs arranged circumferentially at a predetermined angle from each other on the nozzle. A so-called secondary control force is provided, and by injecting control fluid into the nozzle from any fluid injection mechanism, the flow direction of the combustion gas flow flowing through the nozzle is changed to generate an attitude control force. These include injection type, and fin control type, which uses the wings to aerodynamically generate an attitude control device. Changes in the aircraft (10) caused by the attitude controller (9) are detected by rate gyros (1), (2), and (3).

The attitude control device configured as described above controls the attitude angle of the rocket as follows and guides the rocket to the destination point.

In order to give the rocket axis stability, several H2 during flight.
A degree of spin is given. As described above, the attitude control device of this embodiment does not control the spin, that is, the roll angle φ, and leaves it free, and detects the roll rate p with the rate gyro m to safely calculate the pitch angle θ and yaw angle ψ. It is used as a raw material. The reference pitch angle θC is given as a reference pitch angle program that changes from θc1 to θc2 as shown in FIG. 4, based on calculations for determining the trajectory shown in FIG. Therefore, the attitude control device controls the pitch angle θ according to this progera l, and when the rocket deviates from the reference pitch angle θC due to nozzle misalignment or the like, the attitude brake (9) corrects it. Reference yaw angle ψ
C is set to zero as mentioned above, and the rocket is at the launch point (
Control is performed so as not to deviate from the X-Z plane including Sl and the target point (2) due to nozzle misalignment or the like. The rocket is attitude controlled during its motor burn time (same as control time T). Therefore, the rocket is guided along the set trajectory shown in FIG. 3 during motor combustion. After the motor burns, that is, after the control time FT), the attitude of the rocket is not controlled.

There will be no guidance, but at this time it is in the deceleration region and the above-mentioned causes of orbit deviation have been corrected, so there will be no major deviation from the orbit, and there will be no difference in the sliding degree to the rocket's destination point (■). It has no effect on

Note that in the above embodiment, the yaw angle ψ is set to zero, but the yaw angle ψ can also be controlled.

As explained above, the present invention includes a rate gyro that detects the angular velocity in the body coordinate system set for the flying object, coordinate conversion of the angular velocity detected by the rate gyro to an inertial coordinate system, and , a calculation that calculates the attitude angle of the flying object by time-integrating it during the acceleration period of the rocket, compares this attitude angle with a preset reference attitude angle, and outputs a control signal when the error between them exceeds a predetermined value. and an attitude controller for controlling the attitude of the rocket in response to control signals from the arithmetic unit, and the attitude angle of the projectile at the initial stage of flight is controlled to the reference attitude angle.
A flying object can be guided to a destination with high precision using a simple configuration. Furthermore, since a rate gyro, which is smaller and lighter in weight than other sensors such as a rate integral gyro, is used as a sensor, the entire attitude control device can be made smaller and lighter. Therefore, it is possible to equip a small flying object at low cost.

[Brief explanation of drawings]

Fig. 1 is a block diagram showing a flying object attitude control device according to an embodiment of the present invention, Fig. 2 is a diagram showing the relationship between the body coordinate thread and the inertial coordinate thread, and Fig. 3 is based on rocket trajectory calculation. FIG. 4 is a diagram showing a ballistic trajectory, and shows a program for a reference pitch angle θC based on the ballistic trajectory of FIG. 3, with time t5 plotted on the horizontal axis and reference pitch angle θC plotted on the vertical axis. (1) (2) (3)...Rate gyro (4)...
Arithmetic unit (5)...Multiflexa (6)...A/
D converter (7)...Microcomputer (8)...Driver (9)...Actuator aα...Machine, body Patent applicant Nissan Motor Co., Ltd. Agent Patent attorney Yuga Military - Department

Claims (1)

[Claims] There is a rate gyro that detects the angular velocity in the aircraft coordinate system set for the flying object, and the angular velocity detected by this rate gyro is converted into the inertial coordinate system, and the attitude of the flying object is determined by integrating over time during the acceleration period hJ of the rocket. an arithmetic unit that calculates the attitude angle, compares this attitude angle with a preset reference attitude angle, and outputs a control signal 1 when the error between these angles exceeds a predetermined value; An attitude control device for a flying object, comprising: an attitude controller for controlling the attitude of a rocket; and controlling an attitude angle of the flying object at the initial stage of flight to the reference attitude angle.