JPH11201700A - Airframe - Google Patents

Abstract

PROBLEM TO BE SOLVED: To reduce the weight, cost, and size of an airframe by reducing the pitch axis and yaw axis of an auto-pilot device to one axis by spinning the airframe by setting the rear wing of the airframe at a fixed angle. SOLUTION: An auto-pilot section 2 calculates a steering angle command for pitch axis spinning system based on an acceleration command calculated by means of an inertial system proportional navigation command calculating section 1 and inputs the command to a pitch axis actuator 4. Then an auto-pilot gain is calculated by changing the steering angle by means of the actuator 4 and detecting the response from a flying body by means of an inertial device 6. The inertial device 6 sends the angular velocity and acceleration of a rolling flying body to a coordinate transforming section 8 and the section 8 transforms the angular velocity and acceleration into the angular velocity and acceleration of a non-rolling flying body. The airframe is constituted by feeding back the transformed angular velocity, acceleration, rolling angle, and the above- mentioned auto-pilot gain to the auto-pilot section 2.

Description

DETAILED DESCRIPTION OF THE INVENTION

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a flying object having a control device for controlling a flying object only by a pitch axis.

[0002]

2. Description of the Related Art FIG. 10A is a block diagram showing functions of a conventional flying object. FIG. 10B is a conceptual diagram of a conventional flying object. In order to perform flying control in the pitch direction and the yaw direction, four sheets are used as shown in FIG. A rudder is required, and the paired rudder must be controlled by two actuators or independently controlled by using four actuators. FIG.
In (a) of FIG. 1, 1 is an inertial coordinate system proportional navigation command calculation unit that calculates inertial coordinate system proportional navigation acceleration commands S3 and S26, 16 is an autopilot unit that obtains steering angle commands S4 and S27 from acceleration commands S3 and S26, 3 is the first rudder, the second rudder S, from the rudder angle commands S4 and S27.
Rudder angle command S for third rudder F3 and fourth rudder F4
5, S6, S28, S29, a fin mixing section, which converts the first rudder, the second rudder, the third rudder F3, and the fourth rudder F3;
Angle commands S5, S6, S28, S for the rudder F4
29, the rudder is turned accordingly, and the rudder angles S7, S
8, a steering device section that outputs S30, S31, and a body 5 that responds by the steering angle and generates a body acceleration and an angular velocity;
Reference numeral 6 denotes an acceleration sensor output S34, an angular velocity sensor output S33, and an acceleration sensor output S34, which are outputs of the airframe 5, and a velocity S10 obtained by integrating the angular velocity sensor output S33.
An inertial device section 7 for detecting the altitude S9 is a dynamic pressure compensation calculating section for calculating an autopilot gain S32 from the speed S10 detected by the inertial device section 6 and the altitude S9. The autopilot gain S32, the angular velocity sensor output S33, The acceleration sensor output S34 is fed back to the input of the auto pilot calculation unit 16.

Next, the operation will be described. Since the conventional flying object is configured as described above, the inertial coordinate system proportional navigation command calculation unit 1 uses the formula shown in Equation 1 as follows:
The inertial coordinate system proportional navigation acceleration commands S3 and S26 are calculated from the guidance signal S1 and the approach speed S2. Signals S3 and S input as inertial coordinate system proportional navigation acceleration commands
26 is converted into steering angle commands S4 and S27 for pitch and yaw by 16 autopilot units. The pitch and yaw rudder angle commands S4 and S27 are input to the fin mixing unit 3 and the first rudder, the second rudder, and the third
Angle commands S5, S to the rudder F3 and the fourth rudder F4
6, S28 and S29. The steering device 17 performs steering according to the steering angle commands S5, S6, S28, S29 to the first rudder, the second rudder, the third rudder F3 and the fourth rudder F4, and sets the steering angles S7, S8, S30, and S31 are output. In the fuselage 5, the steering angles S7, S8, S30, S
In response to 31, the acceleration and angular velocity of the aircraft are responded. Then, the aircraft acceleration S34, which is the output of this aircraft, the aircraft angular velocity S33, and the aircraft speed S1, which is the integral value of these, are obtained.
0 and the altitude S9 are detected by the inertial device unit 6 and passed to the dynamic pressure compensation calculation unit 7. The dynamic pressure compensation calculation unit 7 calculates the auto pilot gain S32 of the auto pilot unit from the altitude S9 and speed S10 of the airframe and the aerodynamic coefficient, mass, and moment of inertia stored in the memory of the dynamic pressure compensation calculation unit 7. The acceleration S34 and angular velocity S33 of the airframe and the autopilot gain S11 of the dynamic pressure compensation calculator 7 are fed back to the input of the autopilot calculator 16.

[0004]

(Equation 1)

[0005]

In the conventional flying object as described above, the autopilot section 16 must output the steering angle commands S4 and S27, and the two autopilots for the pitch axis and the yaw axis are required. Needed.
Also, two to four actuators are required in the steering device. This not only leads to an increase in the cost of the flying object, but also causes problems such as an increase in the mass of the flying object and an increase in the volume of the flying object.

SUMMARY OF THE INVENTION The present invention solves such a problem. An autopilot device for a pitch axis and a yaw axis is provided by spinning a flying object by giving a constant angle to a rear wing of the flying object. With a single axis, which enables weight reduction, cost reduction and miniaturization of the flying object.

[0007]

The flying object according to the first aspect of the present invention receives a command transmission from the proportional navigation command calculation unit mounted on the ground equipment, and cuts off a fixed steering angle with the front wing for controlling the flying object. By having the rear wing attached, the flying object can be spun and fly, and the guidance of the flying object can be controlled only by the pitch axis autopilot.

The flying object according to the second aspect of the present invention has a rear wing attached to the flying object at a fixed steering angle by receiving a command transmitted from a proportional navigation command calculation unit mounted on the ground apparatus. The flying object is made to fly by spinning, and guidance control can be performed using only an impulse thruster added around the flying object.

The flying object according to the third aspect of the present invention receives command transmission from a proportional navigation command calculation unit mounted on the ground equipment, and is mounted after being cut off at a fixed steering angle from a front wing for controlling the flying object. Guidance control is performed by combined steering with an impulse thruster provided around the flying object and equipped with wings, thereby shortening the response time.

The flying object according to the fourth aspect of the present invention receives a command transmitted from a proportional navigation command calculation section mounted on an airplane, and is attached to the front wing for controlling the flying object at a fixed steering angle to the rear wing. With this, the flying object can be spun and fly, and the guidance control of the flying object can be performed only by the auto pilot on the pitch axis. This makes it possible to construct a system with better mobility than the first, second and third inventions.

The flying object according to the fifth aspect of the present invention receives commands from a proportional navigation command calculation unit mounted on an airplane, and has a rear wing attached to the flying object at a fixed steering angle. The flying object is made to fly by spinning the flying object, and guidance control can be performed using only the impulse thruster added around the flying object. This makes it possible to construct a system with better mobility than the first, second and third inventions.

The flying object according to the sixth aspect of the present invention receives a command transmitted from a proportional navigation command calculation unit mounted on an airplane, and is attached to a front wing for controlling the flying object at a fixed steering angle to a rear wing. In addition, guidance control is performed by combined steering with an impulse thruster added around the flying object, and the response time can be reduced. This makes it possible to construct a system with better mobility than the first, second and third inventions.

The flying object according to the seventh aspect of the present invention receives a command transmitted from a proportional navigation command calculation unit mounted on an artificial satellite, and is attached to the front wing for controlling the flying object at a predetermined steering angle. By having the wings, the flying object can be spun and fly, and the guidance of the flying object can be controlled only by the pitch axis autopilot. This makes it possible to construct a system that excels in long-distance guidance as compared with the fourth, fifth, and sixth inventions.

The flying object according to the eighth aspect of the present invention has a rear wing attached to the flying object at a fixed steering angle by receiving a command transmitted from a proportional navigation command calculation unit mounted on an artificial satellite. The flying object is made to fly by spinning, and guidance control can be performed using only an impulse thruster added around the flying object. This makes it possible to construct a system that excels in long-distance guidance as compared with the fourth, fifth, and sixth inventions.

Further, the flying object according to the ninth invention receives command transmission from the proportional navigation command calculation unit mounted on the artificial satellite, and is attached to the front wing for controlling the flying object at a fixed steering angle. Guidance control is performed by combined steering with an impulse thruster provided around the flying object and equipped with wings, thereby shortening the response time. This makes it possible to construct a system that excels in long-distance guidance as compared with the fourth, fifth, and sixth inventions.

[0016]

DESCRIPTION OF THE PREFERRED EMBODIMENTS Embodiment 1 FIG. 1 is a block diagram showing Embodiment 1 of a flying object of the present invention.
3, 5, 6, and 7 are the same as the above-mentioned conventional. And
1 is an inertial coordinate system proportional navigation command calculation unit mounted on the ground equipment, 2 is an autopilot unit for the pitch axis spin method, 4
Denotes a pitch axis actuator unit.

In FIG. 1, the inertial coordinate system proportional navigation command calculation unit 1 generates a guidance signal S1 from a ground-based guidance computer.
The approach speed S2 is multiplied by a proportional navigation constant to generate an inertial coordinate system proportional navigation acceleration command S3. The signal S3 input as the inertial coordinate system proportional navigation acceleration command is converted into a pitch steering angle command S4 by the pitch axis spin type autopilot calculation unit 2. The steering angle command S4 of this pitch is input to the fin mixing unit 3, and the steering angle command S for the first rudder and the second rudder is calculated by the conversion formula shown in Expression 3.
5, S6. In the pitch axis actuator section 4, the steering angle commands S5 and S5 for the first rudder and the second rudder are provided.
6 and the steering angles S7 and S8 are output. The body 5 responds to the steering angles S7 and S8 with the acceleration and angular velocity of the body. And the output of this aircraft is the aircraft acceleration,
The acceleration sensor output S13, which is the angular velocity, the angular velocity sensor output S12, and the speed S10 of the body, which is the integral value thereof,
The altitude S9 is detected by the inertial device unit 6 and passed to the dynamic pressure compensation calculation unit 7. The dynamic pressure compensation calculator 7 calculates the altitude S of the aircraft.
9, the auto pilot gain S11 of the pitch axis spin type auto pilot unit from the speed S10 and the aerodynamic coefficient, mass, and moment of inertia stored in the memory for itself.
Is calculated. The coordinate conversion unit 8 converts the output of the roll body coordinate system acceleration sensor S13, the output of the roll body coordinate system angular velocity sensor S12, and the roll angle S14 into the output of the non-roll body coordinate system acceleration sensor S16 and the output of the non-roll body coordinate system angular velocity sensor S15. . Further, the acceleration S16 and the angular velocity S15 of the airframe, and the autopilot gain S of the dynamic pressure compensation calculation unit
Numeral 11 is fed back to the input of the pitch axis spin scheme autopilot calculation unit 2.

[0018]

(Equation 2)

Embodiment 2 FIG. 2 is a block diagram showing a flying object according to a second embodiment of the present invention. 1,5
6 and 8 are the same as those in FIG. Reference numeral 9 denotes a side thruster compensator, 10 denotes a side thruster command / mixing unit, and 11 denotes a thruster generator.

The side thruster compensator 9 outputs a thruster command S17 using the acceleration command S3 calculated by the inertial coordinate system proportional navigation command calculation unit 1 mounted on the ground apparatus. In the side thruster command / mixing unit 10, the thruster command S17 is transmitted to each port of the thruster generator by the thruster commands S18 and S1.
Convert to 9. Equation 4 shows this conversion formula. In accordance with the thruster commands S18 and S19, thruster output accelerations S20 and S21 are generated by a thruster generator 11 for generating thruster acceleration. Then, the body 5 outputs the acceleration and the angular velocity based on the thruster output acceleration by the thruster. Further, the inertial device unit 6 detects the acceleration and angular velocity of the body, and outputs a roll body coordinate system angular velocity sensor output S12 and a roll body coordinate system acceleration sensor output S13. Then, the coordinate conversion unit 8 converts the output of the roll machine coordinate system angular velocity sensor S12 and the output of the roll machine coordinate system acceleration sensor S13, which are the outputs of the inertial device unit, from the roll angle S14 to the output of the non-roll machine coordinate system angular velocity sensor S15 and the non-roll. The output is converted into the body coordinate system acceleration sensor output S16. The output of the non-rolled body coordinate system angular velocity sensor S15 and the output of the non-rolled body coordinate system acceleration sensor S16 are converted to the side thruster compensator 1.
Input to 0 to form a loop.

[0021]

(Equation 3)

Embodiment 3 FIG. FIG. 3 is a block diagram showing Embodiment 3 of the flying object of the present invention. 1,2,
3, 4, 5, 6 and 8 are the same as those in FIG. 10 and 11 are the same as those in FIG. And
Reference numerals 12 and 13 denote adder / subtracters, 14 an autopilot for side thrusters, and 15 a thruster output estimating device.

This is a combination of the first and second embodiments. In the block diagram, 1, 2, 3,
Up to 4, 5, and 6 are the same as the blocks configured in the first embodiment, and the adder / subtractor unit 12 outputs a roll body coordinate system acceleration sensor output S13 which is one of the outputs from the inertial device unit 6.
Then, a roll body coordinate system acceleration sensor output S13 that compensates for the thruster output is calculated from the thruster output estimation acceleration S24 output from the thruster output estimator 15, and is output.
The coordinate conversion unit 8 converts the output of the roll body coordinate system acceleration sensor S13 and the output of the roll body coordinate system angular velocity sensor S12, which are the body acceleration and the angular velocity, into the output of the non-roll body coordinate system acceleration sensor S16 and the output of the non-roll body coordinate system angular velocity sensor. Convert to S15. The output from the coordinate conversion unit 8 is input to a pitch axis spin type autopilot calculation unit 2. The adder / subtractor 13 calculates the acceleration for the side thruster from the acceleration command S3 calculated by the inertial coordinate system proportional navigation command calculation unit 1 and the non-rolled body coordinate system acceleration sensor output S16 which is a part of the output from the coordinate conversion unit 8. Calculate and output the command S22. Then, the thruster command S1 is input to the side thruster autopilot 14, and the thruster command S1
7 is output. Hereinafter, the operations of 10 and 11 are the same as those of the second embodiment, and the outputs S20 and S20 from the thruster generator 11
21 is input to the airframe 5. Also, a thruster command S, which is an output from the side thruster autopilot 14,
Reference numeral 17 is input to the thruster output estimator 15 and outputs a thruster output estimated acceleration S24. The thruster output estimated acceleration S24 is input to the adder / subtractor 12 to form a control loop.

Embodiment 4 FIG. 4 is a block diagram showing Embodiment 4 of the flying object of the present invention. 3,5
Reference numerals 6 and 7 are the same as the above-mentioned conventional ones. Reference numeral 1 denotes an inertial coordinate system proportional navigation command calculating unit mounted on an airplane, 2 denotes a pitch axis spin type autopilot unit, and 4 denotes a pitch axis actuator unit.

In FIG. 4, the inertial coordinate system proportional navigation command calculation unit 1 generates a guidance signal S from an aircraft guidance computer.
1. Multiply the approach speed S2 by a proportional navigation constant to generate an inertial coordinate system proportional navigation acceleration command S3. The signal S3 input as the inertial coordinate system proportional navigation acceleration command is converted into a pitch steering angle command S4 by the pitch axis spin type autopilot calculation unit 2. The steering angle command S4 of this pitch is input to the fin mixing unit 3 and is converted into steering angle commands S5 and S6 for the first rudder and the second rudder by the conversion formula shown in Expression 3. In the pitch axis actuator section 4, a steering angle command S for the first rudder and the second rudder is provided.
The steering is performed according to steps S5 and S6, and steering angles S7 and S8 are output. The body 5 responds to the steering angles S7 and S8 with the acceleration and angular velocity of the body. The acceleration of the body, the acceleration sensor output S13 and the angular velocity sensor output S12, which are the angular velocities, and the velocity S10 and the altitude S9 of the body, which are the integral values thereof, are detected by the inertial device unit 6 and the dynamic pressure compensation calculation unit Passed to 7. The dynamic pressure compensation calculation unit 7 calculates an autopilot gain S11 of the pitch axis spin type autopilot unit from the altitude S9, the speed S10 of the airframe, and the aerodynamic coefficient, mass, and moment of inertia stored in the memory of the aircraft. . The coordinate conversion unit 8 outputs a roll body coordinate system acceleration sensor output S13, which is a body acceleration and an angular velocity, a roll body coordinate system angular velocity sensor output S12, and a roll angle S.
14 is converted into a non-rolled machine coordinate system acceleration sensor output S16 and a non-rolled machine coordinate system angular velocity sensor output S15.
The acceleration S16 and angular velocity S15 of the airframe and the autopilot gain S11 of the dynamic pressure compensation calculation unit are fed back to the inputs of the pitch axis spin type autopilot calculation unit 2.

Embodiment 5 FIG. 5 is a block diagram showing Embodiment 5 of the flying object of the present invention. 1,5
6 and 8 are the same as those in FIG. Reference numeral 9 denotes a side thruster compensator, 10 denotes a side thruster command / mixing unit, and 11 denotes a thruster generator.

The side thruster compensator 9 outputs a thruster command S17 using the acceleration command S3 calculated by the inertial coordinate system proportional navigation command calculation unit 1 mounted on the airplane. Side thruster command / mixing unit 1
0, the thruster command S17 is changed to the thruster commands S18, S19 for each port of the thruster generator.
Convert to Equation 4 shows this conversion formula. The thruster output acceleration S is generated by the thruster generator 11 which generates thruster acceleration in accordance with the thruster commands S18 and S19.
20 and S21 are generated. Then, the body 5 outputs the acceleration and the angular velocity based on the thruster output acceleration by the thruster. Further, the inertial device unit 6 detects the acceleration and angular velocity of the body, and outputs a roll body coordinate system angular velocity sensor output S12 and a roll body coordinate system acceleration sensor output S13. Then, the coordinate conversion unit 8 converts the output of the roll machine coordinate system angular velocity sensor S12 and the output of the roll machine coordinate system acceleration sensor S13, which are the outputs of the inertial device unit, from the roll angle S14 to the output of the non-roll machine coordinate system angular velocity sensor S15 and the non-roll. The output is converted into the body coordinate system acceleration sensor output S16. The output of the non-rolled body coordinate system angular velocity sensor S15 and the output of the non-rolled body coordinate system acceleration sensor S16 are output to the side thruster compensator 10.
To form a loop.

Embodiment 6 FIG. FIG. 6 is a block diagram showing Embodiment 6 of the flying object of the present invention. 1,2,
3, 4, 5, 6, and 8 are the same as those in FIG. 10 and 11 are the same as those in FIG. And
Reference numerals 12 and 13 denote adder / subtracters, 14 an autopilot for side thrusters, and 15 a thruster output estimating device.

This is a combination of the fourth embodiment and the fifth embodiment. In the block diagram, 1, 2, 3,
Up to 4, 5, and 6 are the same as the blocks configured in the fourth embodiment, and the adder / subtractor unit 12 outputs a roll body coordinate system acceleration sensor output S13 which is one of the outputs from the inertial device unit 6.
Then, a roll body coordinate system acceleration sensor output S13 that compensates for the thruster output is calculated from the thruster output estimation acceleration S24 output from the thruster output estimator 15, and is output.
The coordinate conversion unit 8 converts the output of the roll body coordinate system acceleration sensor S13 and the output of the roll body coordinate system angular velocity sensor S12, which are the body acceleration and the angular velocity, into the output of the non-roll body coordinate system acceleration sensor S16 and the output of the non-roll body coordinate system angular velocity sensor. Convert to S15. The output from the coordinate conversion unit 8 is input to a pitch axis spin type autopilot calculation unit 2. The adder / subtractor 13 calculates the acceleration for the side thruster from the acceleration command S3 calculated by the inertial coordinate system proportional navigation command calculation unit 1 and the non-rolled body coordinate system acceleration sensor output S16 which is a part of the output from the coordinate conversion unit 8. Calculate and output the command S22. Then, the thruster command S1 is input to the auto-pilot 14 for the side thruster,
7 is output. Hereinafter, the operations of 10 and 11 are the same as those of the second embodiment, and the outputs S20 and S20 from the thruster generator 11
21 is input to the airframe 5. Also, a thruster command S, which is an output from the side thruster autopilot 14,
Reference numeral 17 is input to the thruster output estimator 15 and outputs a thruster output estimated acceleration S24. The thruster output estimated acceleration S24 is input to the adder / subtractor 12 to form a control loop.

Embodiment 7 FIG. 7 is a block diagram showing Embodiment 7 of the flying object of the present invention. 3,5
Reference numerals 6 and 7 are the same as the above-mentioned conventional ones. Reference numeral 1 denotes an inertial coordinate system proportional navigation command calculation unit mounted on a satellite, reference numeral 2 denotes a pitch axis spin type autopilot unit, and reference numeral 4 denotes a pitch axis actuator unit.

In FIG. 7, the inertial coordinate system proportional navigation command calculation unit 1 generates an inertial coordinate system proportional navigation acceleration command S3 by multiplying a guidance signal S1 and an approach speed S2 from a guidance computer mounted on an artificial satellite by a proportional navigation constant. The signal S3 input as the inertial coordinate system proportional navigation acceleration command is converted into a pitch steering angle command S4 by the pitch axis spin type autopilot calculation unit 2. The steering angle command S4 of this pitch is input to the fin mixing unit 3 and is converted into steering angle commands S5 and S6 for the first rudder and the second rudder by the conversion formula shown in Expression 3. The pitch axis actuator section 4 performs steering according to the steering angle commands S5 and S6 for the first rudder and the second rudder, and outputs steering angles S7 and S8. The body 5 responds to the steering angles S7 and S8 with the acceleration and angular velocity of the body. The acceleration of the body, the acceleration sensor output S13 and the angular velocity sensor output S12, which are the angular velocities, and the velocity S10 and the altitude S9 of the body, which are the integral values thereof, are detected by the inertial device unit 6 and the dynamic pressure compensation calculation unit Passed to 7. The dynamic pressure compensation calculation unit 7 calculates an autopilot gain S11 of the pitch axis spin type autopilot unit from the altitude S9, the speed S10 of the airframe, and the aerodynamic coefficient, mass, and moment of inertia stored in the memory of the aircraft. . The coordinate conversion unit 8 outputs a roll body coordinate system acceleration sensor output S13, which is a body acceleration and an angular velocity, a roll body coordinate system angular velocity sensor output S12, and a roll angle S.
14 is converted into a non-rolled machine coordinate system acceleration sensor output S16 and a non-rolled machine coordinate system angular velocity sensor output S15.
The acceleration S16 and angular velocity S15 of the airframe and the autopilot gain S11 of the dynamic pressure compensation calculation unit are fed back to the inputs of the pitch axis spin type autopilot calculation unit 2.

Embodiment 8 FIG. FIG. 8 is a block diagram showing Embodiment 8 of the flying object of the present invention. 1,5
6 and 8 are the same as those in FIG. Reference numeral 9 denotes a side thruster compensator, 10 denotes a side thruster command / mixing unit, and 11 denotes a thruster generator.

The side thruster compensator 9 outputs a thruster command S17 using the acceleration command S3 calculated by the inertial coordinate system proportional navigation command calculation unit 1 mounted on the satellite. In the side thruster command / mixing unit 10, the thruster command S17 is transmitted to each port of the thruster generator by the thruster commands S18 and S1.
Convert to 9. Equation 4 shows this conversion formula. In accordance with the thruster commands S18 and S19, thruster output accelerations S20 and S21 are generated by a thruster generator 11 for generating thruster acceleration. Then, the body 5 outputs the acceleration and the angular velocity based on the thruster output acceleration by the thruster. Further, the inertial device unit 6 detects the acceleration and angular velocity of the body, and outputs a roll body coordinate system angular velocity sensor output S12 and a roll body coordinate system acceleration sensor output S13. Then, the coordinate conversion unit 8 converts the output of the roll machine coordinate system angular velocity sensor S12 and the output of the roll machine coordinate system acceleration sensor S13, which are the outputs of the inertial device unit, from the roll angle S14 to the output of the non-roll machine coordinate system angular velocity sensor S15 and the non-roll. The output is converted into the body coordinate system acceleration sensor output S16. The output of the non-rolled body coordinate system angular velocity sensor S15 and the output of the non-rolled body coordinate system acceleration sensor S16 are converted to the side thruster compensator 1.
Input to 0 to form a loop.

Embodiment 9 FIG. 9 is a block diagram showing Embodiment 9 of the flying object of the present invention. 1,2,
3, 4, 5, 6 and 8 are the same as those in FIG. 10 and 11 are the same as those in FIG. And
Reference numerals 12 and 13 denote adder / subtracters, 14 an autopilot for side thrusters, and 15 a thruster output estimating device.

This is a combination of the seventh embodiment and the eighth embodiment. In the block diagram, 1, 2, 3,
Up to 4, 5, and 6 are the same as the blocks configured in the seventh embodiment, and the adder / subtractor unit 12 outputs a roll body coordinate system acceleration sensor output S13 which is one of the outputs from the inertial device unit 6.
Then, a roll body coordinate system acceleration sensor output S13 that compensates for the thruster output is calculated from the thruster output estimation acceleration S24 output from the thruster output estimator 15, and is output.
The coordinate conversion unit 8 converts the output of the roll body coordinate system acceleration sensor S13 and the output of the roll body coordinate system angular velocity sensor S12, which are the body acceleration and the angular velocity, into the output of the non-roll body coordinate system acceleration sensor S16 and the output of the non-roll body coordinate system angular velocity sensor. Convert to S15. The output from the coordinate conversion unit 8 is input to a pitch axis spin type autopilot calculation unit 2. The adder / subtractor 13 calculates the acceleration for the side thruster from the acceleration command S3 calculated by the inertial coordinate system proportional navigation command calculation unit 1 and the non-rolled body coordinate system acceleration sensor output S16 which is a part of the output from the coordinate conversion unit 8. Calculate and output the command S22. Then, the thruster command S1 is input to the auto-pilot 14 for the side thruster,
7 is output. Hereinafter, the operations of 10 and 11 are the same as those of the second embodiment, and the outputs S20 and S20 from the thruster generator 11
21 is input to the airframe 5. Also, a thruster command S, which is an output from the side thruster autopilot 14,
Reference numeral 17 is input to the thruster output estimator 15 and outputs a thruster output estimated acceleration S24. The thruster output estimated acceleration S24 is input to the adder / subtractor 12 to form a control loop.

[0036]

Since the present invention is configured as described above, it has the following effects.

According to the first aspect of the present invention, two autopilots for the pitch axis and the yaw axis are conventionally required. However, the guidance of the flying object can be controlled by one autopilot by spinning the airframe. As a result, an inexpensive system can be constructed.

Further, according to the second aspect of the present invention, it is possible to control the guidance of the flying object only by the impulse thruster added around the airframe by spinning the airframe, so that a system with a quick response can be constructed.

Further, according to the third aspect of the present invention, a system which is faster in response than Embodiments 1 and 2 is constructed by spinning the airframe and performing combined control of the flying body front wing and the impulse thruster. it can.

According to the fourth invention, two autopilots of the pitch axis and the yaw axis are conventionally required, but the guidance of the flying object is controlled by one autopilot by spinning the aircraft. System by receiving commands of the guidance signal S1 and the approach speed S2 from the inertial coordinate system proportional command calculation unit of the airplane, thereby making it possible to construct a system that is not easily affected by obstacles such as mountains and buildings. .

According to the fifth aspect of the present invention, the flying body can be guided and controlled only by the impulse thruster added around the aircraft by spinning the aircraft. Can be constructed.

According to the sixth aspect of the present invention, a system which is faster in response than Embodiments 4 and 5 is constructed by spinning the airframe and performing combined control of the flying body front wing and the impulse thruster. it can.

According to the seventh aspect, two autopilots of the pitch axis and the yaw axis are conventionally required, but the guidance of the flying object is controlled by one autopilot by spinning the aircraft. , A cheaper system can be constructed, the response is faster than in the fourth and fifth embodiments, and the command transmission of the guidance signal S1 and the approach speed S2 from the inertial coordinate system proportional command calculation unit mounted on the artificial satellite. Thus, long-distance guidance becomes possible as compared with the fourth, fifth, and sixth inventions.

According to the eighth aspect of the present invention, the guidance of the flying object can be controlled only by the impulse thruster added around the aircraft by spinning the aircraft. Can be constructed.

Further, according to the ninth aspect, a system which is faster in response than Embodiments 7 and 8 is constructed by performing combined control of the flying wing and the impulse thruster by spinning the airframe. it can.

[Brief description of the drawings]

FIG. 1 is a block diagram showing Embodiment 1 of a flying object according to the present invention.

FIG. 2 is a block diagram showing Embodiment 2 of a flying object according to the present invention.

FIG. 3 is a block diagram showing Embodiment 3 of a flying object according to the present invention.

FIG. 4 is a block diagram showing Embodiment 4 of a flying object according to the present invention.

FIG. 5 is a block diagram showing Embodiment 5 of a flying object according to the present invention.

FIG. 6 is a block diagram showing Embodiment 6 of a flying object according to the present invention.

FIG. 7 is a block diagram showing Embodiment 7 of a flying object according to the present invention.

FIG. 8 is a block diagram showing Embodiment 8 of a flying object according to the present invention.

FIG. 9 is a block diagram showing Embodiment 9 of a flying object according to the present invention.

FIG. 10 is a block diagram showing a conventional flying object.

[Explanation of symbols]

1 Inertial coordinate system proportional navigation command calculation unit, 2 pitch axis spin type auto pilot unit, 3 fin mixing unit, 4 pitch axis actuator, 5 airframe, 6 inertial device unit, 7 dynamic pressure compensation calculation unit, 8 coordinate conversion unit, 9
Side thruster compensator, 10 side thruster command / mixing unit, 11 thruster generator, 12 adder / subtractor, 13 adder / subtractor, 14 side thruster autopilot, 15 thruster output estimator, 16 autopilot unit, 17 steering unit, S1 Pitch guidance signal, S2 approach speed, S3 pitch axis inertial coordinate system proportional navigation acceleration command, S4 guidance signal, S5 steering angle command for first rudder, S6 steering angle command for second rudder, S7 first rudder steering Angle, S8 Rudder angle of second rudder, S
9 altitude, S10 flying body speed, S11 auto pilot gain, S12 roll body coordinate system angular velocity sensor output, S13 roll body coordinate system acceleration sensor output,
S14 Roll angle, S15 Non-roll aircraft coordinate system angular velocity sensor output, S16 Non-roll aircraft coordinate system acceleration sensor output, S17 thruster command, S18 thruster command for first thruster port, S19 thruster command for second thruster port, S20 first thruster Output acceleration, S21 second thruster output acceleration, S
22 Side thruster acceleration command, S23 Roller body coordinate system acceleration sensor output compensated for thruster output,
S24 thruster output estimated acceleration, S25 yaw guidance signal, S26 yaw axis inertial coordinate system proportional navigation acceleration command, S27 yaw axis steering angle command, S28 steering angle command for third rudder, S29 steering angle command for fourth rudder, S30 Third rudder steering angle, S31 Fourth rudder steering angle, S32 Autopilot gain, S33 Angular velocity sensor output, S34 Acceleration sensor output, A traveling direction, B lateral acceleration, F1 first rudder, F2 second Rudder,
F3 third rudder, F4 fourth rudder.

Claims (9)

[Claims] 1. Outputs of an inertial coordinate system proportional navigation command calculator and an inertial coordinate system proportional navigation command calculator mounted on a ground device for calculating an inertial coordinate system proportional navigation acceleration command using a guidance signal and an approach speed. An autopilot unit for a pitch axis spin method for converting an inertial coordinate system proportional navigation acceleration command into a steering angle command, and a steering angle command for the first rudder and the second rudder to match the steering angle command with the input of the steering device. A fin mixing unit for converting, a pitch axis actuator unit for steering by a steering angle command for the first rudder and the second rudder, and outputting a steering angle, and a steering angle output from the pitch axis actuator unit, acceleration and angular velocity. And the output of the roll body coordinate system angular velocity sensor and the output of the roll body coordinate system acceleration sensor that are the acceleration and angular velocity of the body. An inertial device unit for calculating the speed, altitude and roll angle of the aircraft by integrating the output of the roll aircraft coordinate system angular velocity sensor and the output of the roll aircraft coordinate system acceleration sensor, and the roll aircraft which is the output of the inertial device unit A coordinate conversion unit that converts the coordinate system angular velocity sensor output and the roll body coordinate system acceleration sensor output into a non-roll machine body coordinate system angular velocity sensor output and a non-roll machine body coordinate system acceleration sensor output based on the roll angle, and an inertia unit. A dynamic pressure compensation calculator for calculating an autopilot gain of the pitch axis spin type autopilot from the altitude, speed, aerodynamic coefficient, mass, and moment of inertia stored in a memory of the dynamic pressure compensation calculator; The output of the non-roll aircraft coordinate system angular velocity sensor output and the output of the non-roll aircraft coordinate system acceleration sensor are the pitch axis spin method Flying object, characterized by comprising a means for feedback to an input of over preparative pilot portion. 2. An inertial coordinate system proportional navigation command calculation unit mounted on a ground device for calculating an inertial coordinate system proportional navigation acceleration command using a guidance signal and an approach speed, and a side thruster compensator for outputting a thruster command from the acceleration command. And a side thruster command / mixing unit for converting a thruster command into a thruster command for each thruster port which is an input of the thruster generator, and a thruster output acceleration from a thruster command for each thruster port which is an output of the side thruster command / mixing unit. A thruster generating device that generates the thruster, an airframe that generates acceleration and angular velocity based on a thruster output acceleration that is an output of the thruster generating device, an output of a roll airframe coordinate system angular velocity sensor that is an acceleration and an angular speed of the airframe, and an output of a roll airframe coordinate system acceleration sensor Detect Along with the output of the roll body coordinate system angular velocity sensor, the inertia device unit that calculates the roll angle of the body by integrating the output of the roll body coordinate system acceleration sensor, and the output of the roll body coordinate system angular velocity sensor that is the output of the inertial device unit, A coordinate conversion unit that converts the output of the roll body coordinate system acceleration sensor into a non-roll body coordinate system angular velocity sensor output and a non-roll body coordinate system acceleration sensor output based on the roll angle, and only an impulse thruster added around the fuselage of the flying object A flying object, comprising: means for controlling the flying object by using 3. An output of an inertial coordinate system proportional navigation command calculation unit and an inertial coordinate system proportional navigation command calculation unit mounted on the ground device for calculating an inertial coordinate system proportional navigation acceleration command using the guidance signal and the approach speed. An autopilot unit for a pitch axis spin method for converting an inertial coordinate system proportional navigation acceleration command into a steering angle command, and a steering angle for the first rudder and the second rudder so that the steering angle command matches the input of the pitch axis actuator unit. A fin mixing unit that converts the command into a command, a pitch axis actuator that outputs a steering angle by steering according to a steering angle command for the first rudder and the second rudder, and an acceleration based on the steering angle output from the pitch axis actuator. And angular velocity, the output of the roll body coordinate system angular velocity sensor, which is the acceleration and angular velocity of the body, and the acceleration sensor of the roll body coordinate system The inertial device section that detects the output of the roll, integrates the output of the roll body coordinate system angular velocity sensor and the output of the roll body coordinate system acceleration sensor to calculate the roll angle of the body, and the roll body coordinates that are the outputs of the inertial device section. A coordinate conversion unit that converts the output of the system angular velocity sensor and the output of the roll body coordinate system acceleration sensor into the output of the non-rolled body coordinate system angular velocity sensor and the output of the non-rolled body coordinate system acceleration sensor based on the roll angle, and the non-roll from the coordinate conversion unit An adder / subtractor that generates a side thruster acceleration command from the inertial coordinate system proportional navigation acceleration command output from the body coordinate system acceleration sensor output and the inertial coordinate system proportional navigation command calculator, and the side thruster acceleration output from the adder / subtractor An autopilot for the side thruster that converts commands into thruster commands, and a thruster And side thruster command mixing unit for converting the command into thruster command for thruster ports thruster generator,
A thruster generator that generates a thruster output acceleration according to a thruster command for a thruster port, a thruster output estimator that converts a thruster command output from a side thruster autopilot into a thruster output estimated acceleration, and a thruster output estimator that is output from the thruster output estimator A flying body comprising: an adder / subtractor for calculating an output of a roll body coordinate system acceleration sensor from the estimated thruster output acceleration and the body acceleration output from the inertial device; and a means for performing combined steering by aerodynamic steering and an impulse thruster. . 4. An inertial coordinate system proportional navigation command calculator for calculating an inertial coordinate system proportional navigation acceleration command using a guidance signal and an approach speed, and an inertia output from the inertial coordinate system proportional navigation command calculator. A pitch axis spin type autopilot for converting the coordinate system proportional navigation acceleration command into a steering angle command, and converting the steering angle command into a steering angle command for the first rudder and the second rudder so as to match the input of the steering device. Fin mixing section, the first rudder and the second
A pitch axis actuator that outputs a steering angle by steering with a rudder angle command for a rudder, a body that generates acceleration and angular velocity from a steering angle output from the pitch axis actuator, and an acceleration and angular velocity of the body. Detects the output of the roll body coordinate system angular velocity sensor and the output of the roll body coordinate system acceleration sensor, and integrates the output of the roll body coordinate system angular velocity sensor and the output of the roll body coordinate system acceleration sensor to calculate the velocity, altitude, and roll angle of the body. The inertial device section, and the output of the roll body coordinate system angular velocity sensor and the output of the roll body coordinate system acceleration sensor, which are the outputs of the inertial apparatus section, are output from the non-roll body coordinate system angular velocity sensor output and the non-roll machine coordinate system acceleration sensor based on the roll angle. Coordinate conversion unit to convert to output, altitude, speed,
Aerodynamic coefficient, mass,
A dynamic pressure compensation calculation unit for calculating an autopilot gain of the pitch axis spin type autopilot unit from the moment of inertia; a non-roll body coordinate system angular velocity sensor output and a non-roll body coordinate system acceleration sensor output from the coordinate conversion unit; Means for feeding back an output to an input of an autopilot unit for the pitch axis spin method. 5. An inertial coordinate system proportional navigation command calculator for calculating an inertial coordinate system proportional navigation acceleration command using a guidance signal and an approach speed, and a side thruster compensator for outputting a thruster command from the acceleration command. ,
A side thruster command / mixing unit for converting a thruster command into a thruster command for each thruster port which is an input of the thruster generator, and a thruster output acceleration is generated from a thruster command for each thruster port which is an output of the side thruster command / mixing unit. A thruster generator, a body that generates the acceleration and angular velocity of the body based on the thruster output acceleration that is an output of the thruster generator, and an output of a roll body coordinate system angular velocity sensor and a roll body coordinate system acceleration sensor that are the acceleration and angular velocity of the body And an inertial device unit that calculates the roll angle of the body by integrating the output of the roll body coordinate system angular velocity sensor and the output of the roll body coordinate system acceleration sensor,
Coordinate transformation for converting the output of the roll body coordinate system angular velocity sensor and the output of the roll body coordinate system acceleration sensor, which are the outputs of the inertial device unit, into the output of the non-roll body coordinate system angular velocity sensor and the non-roll body coordinate system acceleration sensor according to the roll angle. And a means for controlling only using an impulse thruster added around the fuselage of the flying object. 6. An inertial coordinate system proportional navigation command calculator for calculating an inertial coordinate system proportional navigation acceleration command using a guidance signal and an approach speed, and an inertia output from the inertial coordinate system proportional navigation command calculator. A pitch axis spin type autopilot unit for converting a coordinate system proportional navigation acceleration command into a steering angle command, and a steering angle command for the first rudder and the second rudder so that the steering angle command matches the input of the pitch axis actuator unit. To a fin mixing unit, a pitch axis actuator unit that steers by a steering angle command for the first rudder and the second rudder, and outputs a steering angle, and an acceleration and a steering angle obtained from the steering angle output from the pitch axis actuator unit. Aircraft that generates angular velocity, output of angular velocity sensor of roll airframe coordinate system that is acceleration and angular velocity of the airframe, acceleration sensor of roll airframe coordinate system An inertial device unit that detects the output and integrates the output of the roll body coordinate system angular velocity sensor and the output of the roll machine coordinate system acceleration sensor to calculate the roll angle of the body; and the roll machine coordinate system that is the output of the inertial device unit. A coordinate conversion unit that converts the output of the angular velocity sensor and the output of the roll body coordinate system acceleration sensor into the output of the non-rolled body coordinate system angular velocity sensor and the output of the non-rolled body coordinate system acceleration sensor based on the roll angle, and the non-rolled body from the coordinate conversion unit An adder / subtractor for generating a side thruster acceleration command from the inertial coordinate system proportional navigation acceleration command output from the coordinate system acceleration sensor output and the inertial coordinate system proportional navigation command calculator, and a side thruster acceleration command output from the adder / subtractor To the thruster command And side thruster command mixing unit for converting the command into thruster command for thruster ports thruster generator,
A thruster generator that generates a thruster output acceleration according to a thruster command for a thruster port, a thruster output estimator that converts a thruster command output from a side thruster autopilot into a thruster output estimated acceleration, and a thruster output estimator that is output from the thruster output estimator A flying machine comprising: an adder / subtractor for calculating a roll body coordinate system acceleration sensor output from the estimated thruster output acceleration and the body acceleration output from the inertial device; body. 7. An output of an inertial coordinate system proportional navigation command calculator and an inertial coordinate system proportional navigation command calculator mounted on a satellite for calculating an inertial coordinate system proportional navigation acceleration command using the guidance signal and the approach speed. An autopilot unit for a pitch axis spin method for converting an inertial coordinate system proportional navigation acceleration command into a steering angle command, and a steering angle command for the first rudder and the second rudder to match the steering angle command with the input of the steering device. A fin mixing unit for converting, a pitch axis actuator unit for steering by a steering angle command for the first rudder and the second rudder, and outputting a steering angle, and an acceleration and an angular velocity based on the steering angle output from the pitch axis actuator unit. And the output of the roll body coordinate system angular velocity sensor, which is the acceleration and angular velocity of the body, and the output of the roll body coordinate system acceleration sensor And an inertial unit for calculating the speed, altitude and roll angle of the airframe by integrating the output of the roll airframe coordinate system angular velocity sensor and the output of the roll airframe coordinate system acceleration sensor, and the roll airframe coordinates which are the outputs of the inertia device unit. A coordinate conversion unit that converts the system angular velocity sensor output and the roll body coordinate system acceleration sensor output into a non-roll machine body coordinate system angular velocity sensor output and a non-roll machine body coordinate system acceleration sensor output based on the roll angle, and the altitude output by the inertial unit. A dynamic pressure compensation calculation unit that calculates an autopilot gain of the pitch axis spin type autopilot unit from the aerodynamic coefficient, mass, and moment of inertia stored in the memory of the speed and dynamic pressure compensation calculation unit; The output of the non-roll airframe coordinate system angular velocity sensor output and the output of the non-roll airframe coordinate system acceleration sensor are output for pitch axis spin method. Flying object, characterized by comprising a means for feedback to the input of the bets pilot portion. 8. An inertial coordinate system proportional navigation command calculator for calculating an inertial coordinate system proportional navigation acceleration command using a guidance signal and an approach speed, and a side thruster compensator for outputting a thruster command from the acceleration command. And a side thruster command / mixing unit for converting a thruster command into a thruster command for each thruster port which is an input of the thruster generator, and a thruster output acceleration from a thruster command for each thruster port which is an output of the side thruster command / mixing unit. A thruster generating device, a thruster output device which generates thruster output acceleration which is an output of the thruster generating device, a device which generates acceleration and angular speed of the device, a roll device coordinate system angular speed sensor output which is the acceleration and angular speed of the device, a roll device coordinate system acceleration Check the sensor output And an inertial unit for calculating the roll angle of the body by integrating the output of the roll body coordinate system angular velocity sensor and the output of the roll body coordinate system acceleration sensor, and a roll body coordinate system angular velocity sensor which is the output of the inertial unit. The output of the roll body coordinate system acceleration sensor output by the roll angle, the output of the non-roll body coordinate system angular velocity sensor,
A flying object, comprising: a coordinate conversion unit for converting the output into a non-rolled body coordinate system acceleration sensor output; and means for controlling using only an impulse thruster added around the fuselage of the flying object. 9. An output of an inertial coordinate system proportional navigation command calculation unit mounted on an artificial satellite for calculating an inertial coordinate system proportional navigation acceleration command using the guidance signal and the approach speed, and an output of the inertial coordinate system proportional navigation command calculation unit. An autopilot unit for a pitch axis spin method for converting an inertial coordinate system proportional navigation acceleration command into a steering angle command, and a steering angle for the first rudder and the second rudder so that the steering angle command matches the input of the pitch axis actuator unit. A fin mixing unit that converts the command into a command, a pitch axis actuator that outputs a steering angle by steering according to a steering angle command for the first rudder and the second rudder, and an acceleration based on the steering angle output from the pitch axis actuator. And angular velocity, the output of the roll body coordinate system angular velocity sensor, which is the acceleration and angular velocity of the body, and the acceleration sensor of the roll body coordinate system The inertial device section that detects the output of the roll, integrates the output of the roll body coordinate system angular velocity sensor and the output of the roll body coordinate system acceleration sensor to calculate the roll angle of the body, and the roll body coordinates that are the outputs of the inertial device section. A coordinate conversion unit that converts the output of the system angular velocity sensor and the output of the roll body coordinate system acceleration sensor into the output of the non-rolled body coordinate system angular velocity sensor and the output of the non-rolled body coordinate system acceleration sensor based on the roll angle, and the non-roll from the coordinate conversion unit An adder / subtractor that generates a side thruster acceleration command from the inertial coordinate system proportional navigation acceleration command output from the body coordinate system acceleration sensor output and the inertial coordinate system proportional navigation command calculator, and the side thruster acceleration output from the adder / subtractor An autopilot for the side thruster that converts commands into thruster commands, and a thruster And side thruster command mixing unit for converting the command into thruster command for thruster ports thruster generator,
A thruster generator that generates a thruster output acceleration according to a thruster command for a thruster port, a thruster output estimator that converts a thruster command output from a side thruster autopilot into a thruster output estimated acceleration, and a thruster output estimator that is output from the thruster output estimator A flying machine comprising: an adder / subtractor for calculating the output of a roll body coordinate system acceleration sensor from the estimated thruster output acceleration and the body acceleration output from the inertial device; and means for performing combined steering by aerodynamic steering and impulse thrusters. body.