JP2001174198A - Guided missile - Google Patents

Abstract

PROBLEM TO BE SOLVED: To solve a problem that a guided missile launched backward from a flying aircraft becomes aerodynamically unstable due to pitch-up moment for increasing attack angle while the missile flies backward, so that it is difficult to secure the attitude stability of the missile body. SOLUTION: Grating airfoils each formed of a plurality of plane airfoils intersecting one another are arranged at the rear portion of the missile body. While the guided missile separated from the aircraft flies backward, the grating airfoils is fixed in rudder angle range such that they become perpendicular to the missile body axis to reduce lift. The rudder angle range of the grating airfoils is turned by 90 degrees when the flying speed become almost zero, so that the grating airfoils generate large lift when the guided missile flies forward thereafter.

Description

DETAILED DESCRIPTION OF THE INVENTION

[0001]

TECHNICAL FIELD The present invention relates to a helicopter,
It relates to a guided flying vehicle that can be mounted on an aircraft such as an airplane, fired or dropped toward another aircraft located behind the aircraft, or a target such as a guided missile, and fly backward. , To elaborate,
In the initial stage of guidance in which the guidance flying vehicle receives airflow from behind the aircraft, and in the middle and late stages of guidance in which airflow is received from the front of the aircraft, it proposes a guidance flying vehicle that keeps the attitude of the aircraft stable. is there.

[0002]

2. Description of the Related Art Mounted on an aircraft (hereinafter referred to as "base machine"),
A conventional technique relating to a guided flying vehicle capable of rearward launch will be described with reference to the drawings. FIG. 7 is a conceptual diagram of a conventional guided flying object that is fired from a base machine toward a predetermined target body located behind the base machine. In the figure, 1 is a guided flying object that can be fired backward, 2 is a base unit, and 3 is a target object such as an aircraft or a guided bullet that is a threat behind the base unit 2. The guided flying object 1
The main body 2 is mounted on the main body 2 such that the front of the body faces rearward (to the right in the drawing) of the main body 2 in flight. When the target body 3 exists behind the base unit 2, the guided flying object 1 is given a firing command from the base unit 2, is separated from the base unit 2, and is separated from the base unit 3.
Let's fly to

FIG. 8 (a) is a view showing the components of a conventional guided flying vehicle 1. In FIG. 8, reference numeral 4 denotes the body of the guided flying vehicle 1 mounted on the base unit 2; (A right side in the figure), a guidance device having a seeker unit 6 such as a radio wave sensor or a light wave sensor in the front part, 7 is a dome mounted on the front part of the body 4, and 8 is a rear part of the body 4 (see FIG. To the left of
The rear wing 9 is a stable wing fixed to the steering wheel, 9 is a steering wing rotatably supported by the fuselage 4, 10 is a front wing fixed to the front part, 1
1 is a propulsion device that generates a propulsion force on the guided flying object 1,
Reference numeral 2 denotes a cover attached to the rear of the body 4 so as to cover the propulsion device 11. FIG. 8B shows the internal configuration of the propulsion device 11. In the figure, reference numeral 13 denotes a combustion gas that is provided inside the propulsion device 11 and ejects combustion gas in the rearward direction of the body 4 (to the left in the drawing). A thrust deflecting device for deflecting the thrust of the fuselage 4; and 15, a vane.

[0004] Each component of such a guided flying object 1 operates as follows. Rear wing 8, steering wing 9, front wing 10
Each of the rear wing 8, the steering wing 9, and the front wing 10 is mounted as a set at each position where the fuselage outer periphery of the fuselage 4 is divided into four equal parts when viewed from the machine axis direction. By the thrust deflection device 14, the attitude stability of the guidance flying object 1 is ensured. The dome 7 is formed of a material that transmits radio waves and light waves, and has functions of protecting the seeker 6 of the guidance device 5 and reducing the air resistance of the body 4. Also, the guided flying object 1 flies in the direction from the dome 7 toward the propulsion device 11 (rearward of the fuselage) with the propulsion device 11 at the head for a while after being fired, and covers until the propulsion device 11 is ignited. 12 is mounted. Also, the thrust deflection device 14
Is provided by providing a vane 15 around the ejection port of the nozzle 13, drives the vane 15 to a desired position in accordance with a thrust deflection angle command described later, and controls the propellant ejected from the nozzle 13. By deflecting the combustion gas, a desired rotational moment can be generated in the body 4.

[0005] FIG. 9 shows a conventional guided flying vehicle 1 as a target 3
FIG. 4 is a diagram illustrating a configuration of a control system that guides the vehicle to stabilize the attitude of the aircraft. In the figure, 16 is an inertial device, 1
Reference numeral 7 denotes a navigation calculation circuit, reference numeral 18 denotes a gain calculation circuit, reference numeral 19 denotes a steering angle and thrust deflection angle command calculation circuit, and reference numeral 20 denotes a steering blade driving device for rotating the steering blade 9.

Next, the operation of the control system will be described. The guided flying object 1 is moved from the base unit 2 to the target
Target information indicating the position, speed, etc. of the vehicle. After the launch, the guidance device 5 searches for the target body 3 based on the target information, and the seeker unit 6 supplements and tracks the target body 3. Further, the guidance device 5 estimates the rate of change of the line-of-sight angle formed between the guidance flying object 1 and the target body 3 and indicates the guiding direction and the target speed of the guidance flying object 1 to the target body 3. Generate a target guidance signal. In the inertial device 16, the angular velocity and acceleration of the body 4 are measured by an inertial sensor unit provided therein, and the measurement results are output to the navigation calculation circuit 17 and the steering angle and thrust deflection angle command calculation circuit 19 as an inertia information signal. You. The navigation calculation circuit 17 calculates an acceleration command and an angular velocity command necessary for guidance based on a target guidance signal from the guidance device 5 and an inertia information signal from the inertia device 16. Further, when the guided flying vehicle 1 is fired from the base unit 2, the inertial device 16 is given the initial altitude and initial speed of the guided flying unit 1 from the base unit 2. In the inertial device 16,
Based on the initial altitude and speed, and the angular velocity and acceleration of the aircraft 4 measured by the internal inertial sensor unit after the launch, the altitude and speed of the guided flying object 1 are calculated by a calculation unit provided inside. Further, the gain calculation circuit 18 calculates an autopilot gain according to the altitude and the speed calculated by the inertial device 16. The steering angle and thrust deflection angle command calculation circuit 19 calculates an acceleration deviation from the acceleration command given from the navigation calculation circuit 17 and the measurement data of the acceleration given from the inertial device 16, and the deviation is calculated by the gain calculation circuit 18. Multiplied by a multiplier of the gain of the autopilot system, and based on the result of the multiplication and the angular velocity given by the inertial device 16, a rudder for realizing a predetermined navigation until the guided flying vehicle 1 meets the target 3. Calculate the angle command and the thrust deflection angle command. This steering angle command is output to the steering wing drive device 20, and the steering wing 9 is steered to obtain a required steering angle in the guided flying object 1. The thrust deflection angle command from the steering angle and thrust deflection angle command calculation circuit 19 is input to the thrust deflection device 14, and the thrust of the guidance vehicle 1 is deflected in a required direction.

Next, the conventional guided flying object 1 is
The guidance process from launching to the target 3 will be described.

FIG. 10 is a diagram showing the behavior of the guided flying object 1 launched from the base unit 2 and flying backward. In the figure, reference numeral 21 denotes a stage in which the guided flying object 1 is fired rearward from the base unit 2 in flight at the speed V0, for example, and is flying at the speed Vb toward the rear of the aircraft.
Indicates a stage in which the propulsion device 11 is ignited and the guided flying vehicle 1 is flying at the speed Vc facing the rear of the aircraft, and 23 indicates a direction in which the guided flying vehicle 1 is directed from the propulsion device 11 to the dome 7 (front of the fuselage). Shows the stage of flying at the speed Va.

After the existence of the target body 3 which poses a threat to the base unit 2 is confirmed, the guided flying object 1 fired from the base unit 2 returns to the state in step 21 before the propulsion device 11 is ignited. The aircraft flies at a flying speed Vb heading rearward of the aircraft, which is almost the same as the speed V0 of the mother machine 2. Thereafter, when the propulsion device 11 is ignited, an initial stage 22 having a speed Vc toward the rear of the vehicle is passed. Thereafter, the guided flying object 1 is accelerated, and finally reaches the stage 23 having the flying speed Va heading forward of the aircraft, and is guided to the target body 3. During such a process, the guided flying vehicle 1 passes through a stage 22 having a velocity Vc toward the rear of the aircraft, which is an aerodynamically unstable velocity area, and a velocity area in which the aircraft velocity is near zero is also increased. Elapses.

FIG. 11 is a diagram showing the aerodynamic moment acting on the conventional guided flying vehicle 1. FIG. 11 (a) shows a case where the vehicle flies forward, and FIG. 11 (b) shows a case where it flies backward. Each case is shown below. In FIG. 11A, Va is the velocity vector of the guided flying vehicle 1 in the stage 23 of flying forward the aircraft, α is the angle of attack with respect to the airflow around the aircraft, L1 is the lift of the wing 10, and Xc1 is the center of gravity CG1.
, The distance from the center of gravity CG1 to the center of the aerodynamic force of the rear wing 8, Ma is around the center of gravity CG1 in the case of the stage 23 of flying forward. , Vair indicates the velocity vector (airspeed) of the airflow.

In the stage 23 in which the guided flying vehicle 1 has a speed Va toward the front of the fuselage, in order to secure aerodynamic static stability, for example, when the angle of attack α, the front wing 10 Lift L1, distance Xc from center of gravity CG1
1. Lift L2 of rear wing 8 and distance Xc from center of gravity CG1
In view of the relationship with 2, a configuration is made such that a head-down moment Ma for reducing the angle of attack α is generated on the dome 7 side. (Moment is positive.)

[0012]

(Equation 1)

That is, even if disturbance such as a change in the direction of the air flow due to the turbulence of the air flow around the airframe acts to generate the angle of attack α, a moment Ma for canceling the angle of attack α is generated, and the static stability of the airframe 4 with respect to the airflow is improved. Can be secured. At this time, the aerodynamic center of the guidance flying vehicle 1 is located downstream of the center of gravity CG1 with respect to the airflow.

On the other hand, immediately after the launch from the base unit 2, the guided flying object 1 enters a stage 21 having a velocity toward the rear of the aircraft. Here, assuming that the steering wing 9 and the thrust deflecting device 14 do not operate when the guidance vehicle 1 takes the angle of attack α, as in the case of the guidance vehicle 1 in step 23, FIG. As shown in FIG.
The lift L2 of the rear wing 8 is generated. In this case, if the positions of the centers of gravity are substantially the same, a moment is generated by the lift as in the case of the step 23, but the moment Mb is represented by the following equation (2). And acts in a direction to further increase the angle of attack α.

[0015]

(Equation 2)

As a result, it becomes difficult to maintain the attitude with respect to the airflow only by the front wing 10 and the rear wing 8, and the thrust deflecting device 14 and the steering wing 9 constantly generate a moment for canceling the moment generated by the angle of attack α. There is a need. At this time, the aerodynamic center of the guided flying object 1 is located upstream of the center of gravity CG1 with respect to the airflow.

[0017]

When the guided flying vehicle is fired backward from the base unit, the flying speed changes from the rear (negative speed) to the forward (positive speed) during the flying process. I do. At this time, the guiding flying object 1 using the thrust deflection device as shown in FIG. 8 shown in the first conventional example has the following problems.

During the flight of the guided flying vehicle 1, a head-lifting moment that increases the angle of attack α is generated due to the balance between the lifts acting on the front wing 10 and the rear wing 8 during the flight behind the fuselage. Is unstable, and the Aircraft 4
There is a problem that it is difficult to secure the posture stability of the person.

Further, in order to secure the attitude stability of the airframe 4, the thrust deflection device 14 and the steering wing 9 are used to deflect the thrust and the lift to generate a moment in a direction to cancel the head-lifting moment. Even if the attitude control of No. 4 is performed, there is a problem that the control becomes impossible if a disturbance exceeding the control force is applied.

The thrust deflection device 14 is configured such that the body 4 is
When the propulsion device 11 does not operate, for example, immediately after being fired from, it is impossible to deflect the thrust, so that it is impossible to generate a moment that cancels this head-up moment, and for some reason Propulsion device 11
There is a problem that control becomes impossible if is not activated.

Further, in the propulsion device 11 having the thrust deflection device 14 inside, the attitude of the head is always unstable due to the moment of lifting while flying at the speed toward the rear of the aircraft, so that the thrust in the axial direction of the aircraft is generated. In addition, it is necessary to constantly generate a thrust in a direction perpendicular to the machine axis for thrust deflection to maintain the attitude of the body 4 at all times. For this reason, it is necessary to mount a large-capacity propellant on the fuselage 4 and mount a high-output nozzle, and there is a problem that the fuselage 4 becomes large. In addition, since an aircraft on which the guidance flying vehicle 1 is mounted is generally limited in the mass and size of the load, such an increase in the size of the airframe 4 may cause a problem in the mounting base unit.

SUMMARY OF THE INVENTION The present invention has been made to solve the above-mentioned problems, and has been made in comparison with a conventional device using only a thrust deflection device or only an aerodynamic characteristic changing device, from the rear of the aircraft to the front. It is an object of the present invention to obtain a guided flying object capable of securing a more stable flying in the speed range of.

[0023]

According to a first aspect of the present invention, there is provided a guided flying object which is separated from an aircraft in flight and has a front portion flying toward the rear of the aircraft. A stabilizing wing disposed at the front of the fuselage, a lattice wing disposed at the rear of the fuselage and having a plurality of plane wings intersecting to form a lattice shape, and thrust deflecting means for deflecting the thrust of the fuselage A detecting means for detecting that the flying speed of the airframe is near zero, and a steering wing drive device for switching the steering angle range of the lattice wing in accordance with the detection by the detecting means for steering. It is.

[0024] The guided flying object according to the second invention is the first flying object.
In the invention, a fin is provided at a rear portion of the fuselage to reduce the resistance of the lattice wing when the mother machine is mounted, and to rotate around the fuselage axis immediately after launch to avoid obstructing deflection of lift due to steering of the lattice wing. Is provided.

[0025]

DESCRIPTION OF THE PREFERRED EMBODIMENTS Embodiment 1 First Embodiment A first embodiment according to the present invention will be described with reference to FIGS. 1, 2, 3, 4, and 5. FIG. FIG. 1 is a diagram showing a configuration of a control system of a guided flying object 24 in this embodiment. In FIG.
Is the same as the conventional example. FIG. 2 is a diagram showing components of the guidance flying object 24 in this embodiment. FIG.
(B) shows the case of flying forward of the aircraft.
In the drawing, reference numeral 26 denotes a lattice wing provided at the rear part (left side in the figure) of the fuselage 4 and having a plurality of plane wings intersecting to form a lattice shape. The lattice wings 26 are mounted as a set of four blades, one at each position, which divides the outer periphery of the fuselage of the fuselage 4 into four equal parts when viewed from the axial direction. FIG. 3 is a diagram showing the behavior of the guided flying object 24 fired or dropped from the base unit 2.

Next, the operation will be described. In FIG. 3, the stage at which the guided flying object 24 is mounted on the base unit 2 flying at the speed V0 (for example, in the supersonic range) is designated as 27. The guided flying object 24 fired or dropped from the base unit 2 has a speed Vb immediately after separation, as shown in step 28.
And fly in the direction of travel of the main unit 2. At this time,
As shown in FIG. 2A, the guided flying object 24 is
With the dome 7 at the rear and the cover 12 held by the fuselage 4
Fly to the rear of the aircraft with In this step 28, the speed Vb in the same direction as the base unit 2 is reduced by the action of the air resistance of the body 4. At this time, the provision of the cover 12 can prevent the airflow from flowing into the propulsion device 11.

After that, after being separated from the base unit 2, tr1
After a few seconds, the propulsion device 11 is ignited, and the holding member of the cover 12 is released and the cover 12 is discharged to the rear of the aircraft, as shown in step 29 of FIG. In this stage 29, the speed Vc (for example, a subsonic region) in the same direction as the base unit 2, that is, the rearward direction of the unit, is reduced by the propulsion force of the propulsion device 11 and the air resistance of the unit 4, and separated from the base unit 2. After tw1 seconds from the start, the vehicle has a direction Va opposite to the traveling direction of the base unit 2, that is, the speed Va toward the front of the aircraft, and the flying state shown in step 30 is reached.
Further, when the guided flying object 24 is separated from the base unit 2, information for setting the ignition time tr 1 is given from the base unit 2, and the ignition time tr 1 is set in advance in the propulsion device 11. The steering angle range of the lattice wing 26 is fixed to the range from A to B in FIG. 2A by the steering blade drive device 20 until Step 29, and in Step 30, the steering angle range from C to D in FIG. The steering angle range is fixed within the range.

Here, the aerodynamic characteristics of the lattice wing 26 will be described with reference to FIG. FIG. 4A shows the state of the aerodynamic characteristics of the flat wing. Numeral 31 indicates a state near an attack angle of zero, and numeral 32 indicates a state near an attack angle of 90 degrees. When the angle of attack is small, the flat wing has a large lift L5 compared to the resistance D5, but when the angle of attack is large, the flat wing stalls and the lift L6 becomes very small as compared with the resistance D6. On the other hand, as shown in FIG. 4 (b), the lattice wings 26 are formed by rows of the flat plate wings of FIG. 4 (a) arranged in parallel. 33
Indicates a state near an attack angle of zero, and 34 indicates a state near an attack angle of 90 degrees. The state 33 corresponds to the state where the steering angle range in FIG. 2B is from c to d, and the state 34 corresponds to the state from a to b in FIG. 2A. In the state 33, each flat blade generates a lift L5. However, in the state 34, each flat blade is in a row, so it is difficult to stall, and a lift L7 larger than the lift L6 of the single flat blade is generated. Therefore, the lattice wing 26 can secure a lift even at a high angle of attack at which a single flat wing stalls. However, the lift L7 generated in the state 34 is smaller than the lift L5 generated in the state 33.

Next, the rotation control of the lattice blade 26 and its principle will be described with reference to FIG. In transitional stages 29 and 30 in which the speed of the airframe 4 is reduced, the gain of a circuit required for controlling the airframe 4 and the like change significantly due to a change in the flying speed. Therefore, similarly to the conventional guidance flying vehicle 1 shown in FIG. 9, the guidance device 5 calculates the target guidance signal based on the target tracking signal. In the inertial device 16, the angular velocity and acceleration of the body 4 are measured by the inertial sensor unit and output to the inertial device 16 and the steering angle and thrust deflection angle command calculation circuit 19, and the flying height and speed etc. are calculated by the calculation unit. Is calculated and output to the gain calculation circuit 18, and the gain calculation circuit 1
In step 8, a multiplier of the autopilot system gain according to the altitude and the speed is calculated. Further, the navigation calculation circuit 17 calculates an acceleration command or an angular velocity command based on the target guidance signal from the guidance device 5 and the angular velocity and acceleration from the inertial device 16. The steering angle and thrust deflection angle command calculation circuit 19 uses a predetermined navigation based on the angular velocity command and acceleration command from the navigation calculation circuit 17, the angular velocity and acceleration observation data from the inertial device 16, and the multiplier from the gain calculation circuit 18. The steering angle and thrust deflection angle command that realizes
The steering angle command is output to the steering wing drive device 20, and the thrust deflection angle command is output to the thrust deflection device 14. Based on the output steering angle command and thrust deflection angle command,
The steering blade drive device 20 and the thrust deflection device 14 perform required attitude control of the airframe 4 moment by moment. For example, FIG.
In (a), when a steering angle command for tilting the fuselage 4 of the guidance flying object 24 in a direction to raise the tip of the cover 12 is output, the steering blade driving device 20 turns the lattice blade 26 counterclockwise in the figure ( (In the direction of arrow a in the figure), and
When the steering angle command for tilting the fuselage 4 of the guidance flying vehicle 24 in the direction of lowering the tip of the steering wheel is output, the steering wing drive device 20
The steering is performed by turning the lattice wing 26 clockwise as viewed in the figure (in the direction of arrow A in the figure). In addition, for example, in FIG. 2B, when a steering angle command for tilting the fuselage 4 of the guidance flying object 24 in a direction to raise the tip of the dome 7 is output, the steering blade driving device 20 Turn clockwise (in the direction of arrow d in the figure) toward
When the steering angle command for tilting the fuselage 4 of the guidance flying vehicle 24 in the direction of lowering the tip of the steering wheel is output, the steering wing drive device 20
The steering is performed by turning the lattice wing 26 counterclockwise (in the direction of the arrow C in the figure) toward the figure. Further, when a thrust deflection angle command for tilting the fuselage 4 of the guidance flying object 24 in the direction of raising the tip of the dome 7 is output, the thrust deflection device 14 moves the thrust in the downward direction in the figure (the direction of the arrow f in the figure). To guide the flying object 24 in the direction of lowering the tip of the dome 7.
If the thrust deflection angle command for tilting the aircraft 4 is output,
Thrust deflection is performed such that the thrust deflection device 14 applies a thrust in the upward direction in the figure (the direction of the arrow e in the figure).

Further, in this embodiment, the flying speed observed by the inertial device 16 is given to the lattice wing rotation determination circuit 25. The lattice wing rotation determination circuit 25 detects that the observed flying speed falls below a predetermined speed reference value (for example, 50 km / h) near zero speed, and that the flying speed reverses from the rear of the aircraft to the front. Then, a steering angle rotation command is generated. Steering blade drive 20
Then, a steering angle rotation command is given from the lattice blade rotation determination circuit 25 to rotate the lattice blade 26 by 90 degrees. As a result, as shown in FIG. 4B, the lift generated by the lattice wings 26 increases, and the static stability of the airframe 4 on the flight path is ensured.

Here, the relationship between the flying speed and the static stability will be described in detail. FIG. 5A shows the guided flying object 24 when the rudder angle range of the lattice wing 26 is limited to the range from c to d.
FIG. 5B shows the aerodynamic moment acting on the guided flying vehicle 24 when the steering angle range of the lattice wing 26 is limited to the range from A to B. It is.

In the case of step 28 in FIG. 3 in which the guidance vehicle 24 flies at the speed Vb toward the rear of the aircraft, assuming that the guidance vehicle 24 has an angle of attack α with respect to the airflow, FIG. ), A lift L4 is generated by the lattice wing 26. This lift L4 and the lift L generated by the front wing 10
1. The distance Xc between the aerodynamic center of the lattice wing 26 and the center of gravity CG2
3. The relationship of the moment balance shown in Equation 3 is established between the aerodynamic center of the front wing 10 and the distance Xc1 between the center of gravity CG2.

[0033]

(Equation 3)

In step 28, the range of elevation of the lattice wing 26 is limited to the range from A to B, so that the generated lift L4 can be reduced. Therefore, when the angle of attack α is taken,
, A head-down moment Md (Md <0) for reducing the angle of attack α on the cover 12 side can be generated around the center of gravity CG2. (Moment is head-up positive.)
Note that the speed Vc at which the guided flying object 24 moves toward the rear of the aircraft.
In the case of step 29 in which the flight is performed, the same moment as in the case of step 28 is generated. Therefore, guided flying object 2
When the aircraft 4 is flying toward the rear of the aircraft, the angle of attack α
Is generated, a moment Md for canceling the occurrence is generated, and the static stability of the body 4 against the airflow can be secured.

The flight speed of the guidance vehicle 24 is reversed by the action of the thrust and air resistance of the propulsion device 11, and the vehicle flies forward at the speed Va at step 30 in FIG.
, The angle of attack range of the lattice wing 26 is limited to the range from c to d. Here, assuming that the guidance flying vehicle 24 has the same angle of attack α with respect to the airflow as in the case of FIG. 5B, as shown in FIG.
Occurs. This lift L3 and the lift L generated by the front wing 10
1. The distance Xc between the aerodynamic center of the lattice wing 26 and the center of gravity CG2
3. The relationship of the moment balance shown in Formula 4 is established between the aerodynamic center of the front wing 10 and the distance Xc1 between the center of gravity CG2.

[0036]

(Equation 4)

In step 30, the angle of attack of the lattice wing 26 is limited to the range from c to d, so that the generated lift L3 can be increased. Therefore, a head-down moment M for reducing the angle of attack α on the dome 7 side around the center of gravity CG2 of the body 4
c (Mc <0) can be generated. As a result, even when the guidance flying body 24 flies toward the front of the aircraft, aerodynamic static stability is ensured, and the attitude of the aircraft 4 with respect to the airflow can be kept stable.

On the other hand, the guided flying object 24 passes through a speed region where the speed is near zero in a transition period in which the flying speed changes from the rear of the aircraft to the front of the aircraft. In this speed range, the lift acting on each wing becomes small, and an aerodynamically unstable state occurs for the guided flying object 24.

For this reason, in the guided flying object 24 of this embodiment, control is performed by the thrust deflection device 14 so that the attitude of the aircraft body 4 is kept stable. For example, in the steering angle and thrust deflection angle command calculation circuit 19 shown in FIG. 1, the magnitude of the flying speed observed by the inertial device 16 is smaller than a predetermined value (for example, when the magnitude of the speed is 30 m / s or less). became)
When this is detected, a thrust deflection command for giving a thrust in a direction perpendicular to the machine axis is generated based on the speed and acceleration observed by the inertial device 16 so as to spatially stabilize the machine body 4. In the thrust deflection device 14, the thrust deflection angle command is given from the steering angle and thrust deflection angle command calculation circuit 19, during which thrust deflection control is constantly performed by the thrust deflection device 14 to keep the attitude of the body 4 stable. It is.

Embodiment 2 FIG. 6 is a diagram showing components of the guided flying object 24 in this embodiment. In the figure, reference numeral 35 denotes a fin attached to the cover 12, and the other components are the same as those of the first embodiment according to the present invention. FIG. 6A shows a form in a state 27 in which the guided flying object 24 is mounted on the base unit 2 shown in FIG.
FIG. 6B shows the state in the state 28 immediately after the launch from the base unit 2, FIG.
(C) shows a configuration in a state 29 in which the propulsion device 11 is ignited.

As shown in FIG.
As shown in the figure, the fins 35 are mounted at the same position as the lattice wing 26 as a set of four fins 35 at each position where the fuselage outer periphery of the fuselage 4 is divided into four equal parts when viewed from the machine axis direction. , The effect of shortening the cruising distance due to an increase in resistance when the guide flying object 24 of the base unit 2 is mounted is reduced.

As shown in FIG. 6 (b), the guided flying object 24 is fired until the propulsion device 11 is ignited.
The fins 35 rotate 45 degrees around the body axis together with the cover 12 by hydraulic power or power of a motor or the like. In this way, in the state 28 in which the guided flying object 24 flies toward the rear of the aircraft, it is possible to prevent the lift deflection caused by the steering of the lattice wing 26 from being hindered.

Further, when the propulsion device 11 is ignited, FIG.
As shown in (c), the holding member of the cover 12 comes off,
The fins 35 are discharged to the rear of the fuselage together with the cover 12.

[0044]

Since the guided flying object according to the present invention is constituted as described above, the following effects can be obtained.

According to the first aspect of the present invention, in the guided flying vehicle separated from the aircraft and flying rearward, when the flying vehicle speed is at the rear of the aircraft, the grid wing is directed in a direction perpendicular to the aircraft. Steering to secure static stability, and when the flying body speed is in front of the aircraft, steer the lattice wing in a direction parallel to the aircraft to secure static stability, and head forward from behind the aircraft. In all speed ranges up to the above, it is possible to obtain a guided flying object that can secure a more stable flight.

According to the second aspect of the present invention, in the guided flying vehicle separated from the aircraft and flying rearward toward the rear of the aircraft, a fin rotatable in the machine axis direction is attached to the rear of the body of the lattice wing, so that the aircraft It is possible to reduce the resistance of the guided flying vehicle mounted on the vehicle, suppress the increase in the resistance of the aircraft, and avoid hindering the control by the steering of the lattice wing immediately after the launch.

[Brief description of the drawings]

FIG. 1 is a configuration diagram of a control system according to a first embodiment of a guided flying object of the present invention.

FIG. 2 is a configuration diagram showing Embodiment 1 of the guided flying object of the present invention.

FIG. 3 is a diagram showing the behavior of the embodiment of the guided flying object of the present invention.

FIG. 4 is a diagram showing characteristics of a lattice wing in Embodiment 1 of the guided flying object of the present invention.

FIG. 5 is a diagram showing an aerodynamic moment in the first embodiment of the guided flying object of the present invention.

FIG. 6 is a configuration diagram showing Embodiment 2 of the guided flying object of the present invention.

FIG. 7 is a diagram showing the concept of a guided flying object system.

FIG. 8 is a configuration diagram showing a conventional guided flying object.

FIG. 9 is a configuration diagram of a conventional control system for a guided flying object.

FIG. 10 is a diagram showing the behavior of a conventional guided flying object.

FIG. 11 is a diagram showing an aerodynamic moment in a conventional guided flying object.

[Explanation of symbols]

DESCRIPTION OF SYMBOLS 1 Conventional guidance flying object 2 Base unit 3 Target body 4 Airframe 5 Guidance device 6 Seeker part 7 Dome 8 Rear wing 9 Steering wing 10 Front wing 11 Propulsion device 12 Cover 13 Nozzle 14 Thrust changing device 15 Vane 16 Inertial device 17 Navigation calculation Circuit 18 Gain calculation circuit 19 Attack angle and thrust deflection angle command calculation circuit 20 Steering blade drive device 21 Stage in which guided flying vehicle flies at speed Vb toward rear of aircraft 22 Guided flying vehicle at speed Vc toward rear of aircraft Flying stage 23 Stage in which the guided flying vehicle flies at the speed Va heading in front of the aircraft 24 Guided flying vehicle 25 Lattice wing rotation determination circuit 26 Lattice wing 27 Stage in which the guided flying vehicle is mounted on the base unit 28 Guidance Stage where the flying object flies at the speed Vb toward the rear of the aircraft 29 Guided flying object goes toward the rear of the aircraft Stage of flying at a degree of Vc 30 Stage of flying at a speed Va of a guided flying vehicle toward the front of the aircraft 31 State of aerodynamic load near a plane of wing at zero angle of attack 32 State of aerodynamic load at about 90 degree of angle of attack of a plane wing State 33 Lattice wing is in the state of aerodynamic load near zero angle of attack 34 Lattice wing is in the state of aerodynamic load near 90 degree angle of attack 35 Fin

Claims (2)

[Claims] 1. A guided vehicle separated from an aircraft in flight and having a front portion flying toward the rear of the aircraft, a stabilizer wing disposed at a front portion of the aircraft of the guided flying device, A lattice wing disposed at the rear of the aircraft, a plurality of plane wings intersect to form a lattice shape, thrust deflecting means for deflecting the thrust of the aircraft, and detecting that the flying speed of the aircraft is near zero. And a steering wing drive device for switching the steering angle range of the lattice wing in accordance with the detection by the detection means for steering. 2. The device is disposed at the rear of the fuselage to reduce the resistance of the lattice wing when the mother machine is mounted, and to rotate around the fuselage axis immediately after firing to avoid hindering the deflection of lift due to the steering of the lattice wing. The guided flying object according to claim 1, further comprising a fin.