JP4647985B2 - Flying object propulsion control device - Google Patents

Description

  The present invention can efficiently perform control associated with propulsion of a flying object such as thrust control, attitude control, or trajectory control of a flying object such as a rocket, and achieve a reduction in size and weight by adopting a simple structure. The present invention relates to a flying object propulsion control device that can

  Conventionally, the following two methods are known as a flying object propulsion control apparatus using a rocket motor that is used for thrust control and attitude control of a flying object and contains a solid propellant. That is, the first is a method in which the solid propellant is burned and the flow rate and injection direction of the combustion gas is controlled by a valve or pintle (for example, see Patent Document 1), and the second is accommodated in the flying object. This is a system in which the combustion gas is injected from a number of small rocket motors from nozzles facing the outer periphery of the flying object (see, for example, Patent Document 2).

  This will be specifically described with reference to the drawings. FIG. 4 is an explanatory diagram showing a flying object propulsion control device that controls combustion gas using a pintle and a valve. In this example, a large combustion chamber 32 in which a solid propellant 31 is loaded in a flying body 30 is provided, and a gas flow path 33 having a transverse T-shaped cross section is connected to the rear end of the combustion chamber 32. Yes. A valve 34 and a pintle 35 are provided at an intermediate portion of the gas flow path 33, and sub nozzles 36 are provided at both ends of the gas flow path 33. A main nozzle 37 is provided at the rear end of the flying body 30. Then, the solid propellant 31 is always combusted, and the generated combustion gas is controlled by a valve 34 and a pintle 35 provided in the gas flow path 33 so that the combustion gas is ejected from the sub-nozzle 36, and thrust for an arbitrary time in the lateral direction The attitude of the flying object 30 is controlled by generating the same.

On the other hand, FIG. 5 is an explanatory view showing a flying object propulsion control apparatus of a type controlled using a number of conventional small rocket motors. (A) is a front view which fractures | ruptures and shows a part of flying object propulsion control apparatus, (b) is sectional drawing in the 5b-5b line. As shown in these drawings, a plurality of small rocket motors 38 are radially arranged in the flying body 30 so as to extend in the radial direction of the flying body 30, and sub nozzles are arranged at the outer peripheral side ends of the small rocket motors 38. 36 is provided. A main nozzle 37 is provided at the rear end of the flying body 30. The attitude of the flying object 30 is controlled by operating any one of the small rocket motors 38.
Japanese Patent No. 3510527 (first page, second page and FIG. 2) Japanese Patent No. 3075055 (first page, FIG. 3)

  However, in the flying object propulsion control device shown in FIG. 4, the solid propellant 31 is loaded together in one place in the flying object 30, so that it is necessary to generate thrust once the solid propellant 31 is burned. Even if there is no, the solid propellant 31 continues to burn, the utilization efficiency of the solid propellant 31 is poor, and the propulsion control efficiency of the flying object 30 is poor. In addition, control mechanical parts such as a rod for moving the opening / closing means of the valve 34 and the pintle 35 and its driving means are indispensable, and there is a problem that the structure of the flying object 30 becomes complicated and large. .

  On the other hand, in the flying object propulsion control apparatus shown in FIG. 5, one sub nozzle 36 is required for one small rocket motor 38. In order to obtain a high thrust with each small rocket motor 38, each large sub nozzle is required. 36 has to be mounted, which causes a problem in that the volume and mass increase. Also, if thrust is required again in the direction of the sub nozzle 36 of the small rocket motor 38 that has been burned, it is necessary to rotate the airframe or operate a plurality of small rocket motors 38, which is not efficient. There was also a problem.

  Furthermore, in order to extend the flight time, for example, a flying object propulsion control apparatus as shown in FIG. 6 can be considered. That is, a pair of front and rear rocket motors 39 a and 39 b are disposed in the flying body 30. Each rocket motor 39a, 39b includes a cylindrical solid propellant 31 and a front end ignition device 40, and a pressure-resistant partition wall 41 is provided between the rocket motors 39a, 39b.

  In such a flying object propulsion control apparatus, the solid propellant 31 and the ignition device 40 constituting the rocket motor 39a at the front position are protected from the combustion heat and combustion pressure by the rocket motor 39b at the rear position. The thrust is limited to a certain value because the structure is complicated and the solid propellants 31 of both rocket motors 39a and 39b cannot be ignited simultaneously. Moreover, it may be difficult to arrange three or more rocket motors side by side in the flying object 30 in terms of the arrangement of components, heat, pressure, and the like.

  Therefore, an object of the present invention is to make it possible to efficiently perform control associated with propulsion of a flying object such as thrust control and attitude control, and to achieve a small structure and light weight with a simple structure. The object is to provide a body propulsion control device.

To achieve the above object, the projectile propulsion control apparatus of the first invention in the present invention is provided in the flying body, it is formed in a pressure heat-resistant material in the ignition device and the front end to the rear end portion A plurality of rocket motors having a cylindrical solid propellant having an inner hole are loaded into a primary combustion chamber provided with a closure, and a single secondary combustion chamber communicates with the plurality of primary combustion chambers. A plurality of secondary combustion chambers configured as described above, and a nozzle having a throat having a smaller cross-sectional area than the throat portion of the main nozzle is installed between the primary combustion chamber and the secondary combustion chamber, and 2 A sub-nozzle having a throat portion is installed in the next combustion chamber. Further, the solid propellant combustion gas loaded in the rear part is jetted from the main nozzle to obtain thrust, and each solid propellant is supplied from the outside. Individual based on ignition signal The ignition device is operated to burn the solid propellant, and the sub-nozzles of the secondary combustion chamber are arranged so as to extend in a direction orthogonal to the axis of the flying object at the center of gravity of the flying object. It is characterized by being.

The flying object propulsion control device according to a second aspect of the present invention is the flying object propulsion control device according to the first aspect, wherein each of the rocket motors is configured such that the combustion of the subsequent rocket motor is started after the completion of the combustion of the previous rocket motor. It is what.

A flying object propulsion control device according to a third aspect of the invention is the first or second aspect of the invention, wherein the plurality of rocket motors are arranged to extend in a direction parallel to the axis of the flying object.

According to the present invention, the following effects can be exhibited.
Projectile propulsion control apparatus of the first aspect of the invention is loaded into the primary combustion chamber the closure formed by the pressure- and heat-resistant material in the ignition device and the front end to the rear end portion is provided, a cylinder having an inner bore a plurality of rocket motors with Jo of solid propellant, and a plurality of secondary combustion chamber which is constructed as a single secondary combustion chamber is communicated to said plurality of primary combustion chamber. Thus, the primary combustion chamber of a plurality of rocket motors having a solid propellant, by a combination of a plurality of secondary combustion chamber in which a single secondary combustion chamber is communicated to said plurality of primary combustion chamber, Control accompanying propulsion of the flying object such as thrust control and attitude control can be performed efficiently. In addition, a valve and a pintle are not required as in the prior art, and an extra nozzle for each rocket motor is not required, and a simple structure can reduce the size and weight.

Further, each rocket motor is configured so that the solid propellant is burned by individually operating the ignition device based on an external ignition signal. For this reason, each rocket motor can be ignited at an arbitrary timing, and thrust can be generated based on the ignition signal when necessary. In addition to the above effects, the average flight speed and the flight time can be increased. . In addition, since it is possible to ignite one or more arbitrary numbers of rocket motors at the same time, an arbitrary amount of thrust can be obtained by simultaneous ignition of a plurality of rocket motors when necessary. Performance can be improved. Furthermore, since it is possible to operate different rocket motors connected to the secondary combustion chamber, thrust can be generated in the resultant direction of the thrust vector, and the control characteristics accompanying the propulsion of the flying object are greatly improved. be able to.
In addition, since the sub-nozzles of the secondary combustion chamber are arranged so as to extend in a direction orthogonal to the axis of the flying object at the center of gravity of the flying object, the flying object can be moved in parallel, effectively controlling the attitude. The thrust generated in the lateral direction of the flying object can be used as a side thruster (side thrust body).
In the flying object propulsion control device of the second invention, each rocket motor is configured so that the combustion of the subsequent rocket motor is started after the combustion of the previous rocket motor is completed. As a result, the flight distance of the flying object can be extended, and the flight time can be extended while maintaining a predetermined flight speed.

In the flying object propulsion control device of the third invention, the plurality of rocket motors are arranged so as to extend in a direction parallel to the axis of the flying object. Therefore, a plurality of rocket motors can be arranged at high density in the flying object, and in addition to the effects according to the first or second invention, the flying object propulsion control device can be reduced in size and weight. In addition, since the solid propellant can be filled from one direction at a time into the primary combustion chambers of the rocket motors arranged in parallel at the time of manufacturing the solid propellant, the manufacture of the rocket motor becomes easy. This is also advantageous from the viewpoint of economy.

(First embodiment)
Hereinafter, a first embodiment in which the flying object propulsion control device of the present invention is embodied will be described in detail with reference to the drawings.

  FIGS. 1A to 1D are views showing the flying object propulsion control apparatus of the first embodiment, and FIG. 1A is a front view showing a main part of the flying object propulsion control apparatus in a cutaway state. 1B is an enlarged sectional view of the rocket motor, FIG. 1C is a sectional view taken along line 1c-1c in FIG. 1B, and FIG. 1D is taken along line 1d-1d in FIG. It is sectional drawing.

  As shown in FIG. 1A, the airframe 11 constituting the flying body 10 of the first embodiment is formed in a substantially cylindrical shape, and the front end portion is formed in a tapered shape (bullet shape) whose diameter is reduced toward the front end. At the rear end, a main nozzle 12 that opens rearward through a throat portion 12 a is disposed so as to extend in a direction parallel to the axis of the flying object 10, that is, the axis X of the airframe 11. A plurality of rocket motors 13 are disposed slightly behind the central portion in the body 11 so as to extend in a direction parallel to the axis X of the body 11. As shown in FIG. 1D, in the first embodiment, 19 rocket motors 13 are arranged so as to be packed most closely. These rocket motors 13 are arranged symmetrically with respect to the axis X of the body 11 and are also symmetrically arranged with respect to the horizontal line Y and the vertical line Z.

  As shown in FIG. 1B, each rocket motor 13 has a primary combustion chamber 14 inside, and an ignition device 15 is provided at the front end portion of the primary combustion chamber 14, and a pressure-resistant and heat-resistant property is provided at the rear end portion. A closure 16 made of a material is provided, and a solid propellant 17 is loaded between them. The ignition device 15 is configured by connecting an ignition part and a lead wire for sending an ignition signal to the ignition part. Each rocket motor 13 is individually ignited by an ignition device 15 based on an external ignition signal, and the solid propellant 17 is combusted. The closure 16 is ruptured by gas ejection due to combustion of the solid propellant 17, and preferably has a thickness of 1 to 2 mm. As shown in FIG. 1C, the solid propellant 17 has a cylindrical shape and has an inner hole 18 at the center. The solid propellant 17 is combusted from the inner hole 18 side by ignition of the ignition device 15, and the combustion spreads in the radial direction to reach the outer peripheral surface (internal combustion).

  As each rocket motor 13, two or more types of rocket motors 13 having different sizes (shapes and sizes) can be used. In this case, the loading rate in the airframe 11 can be increased. In addition, if the composition and burning rate of the solid propellant 17 suitable for each size are set, the generated thrust can be made equal even if different rocket motors 13 are used. The solid propellant 17 does not need to have the same burning rate and shape, and two or more kinds of solid propellants 17 having different burning rates and shapes may be used in combination.

As shown in FIG. 1 (a), the space inside the fuselage 11 between the primary combustion chamber 14 and the main nozzle 12 is a secondary combustion chamber 19. Therefore, the primary combustion chamber 14 and the secondary combustion chamber 19 in each rocket motor 13 are in communication . B installed a nozzle having a throat portion cross-sectional area smaller than the throat portion 12a of the main nozzle 12 between the primary combustion chamber 14 of the socket motor 13 and the secondary combustion chamber 19, combustion pressure in the primary combustion chamber 14 Is set to be several times higher than the pressure in the secondary combustion chamber 19 so that one or a plurality of rocket motors 13 are burned, and the combustion pressure of the rocket motor 13 is changed even if the pressure in the secondary combustion chamber 19 changes. Can be kept constant. Thus, secondary combustion chamber pressure in the 19 without changing the burning time of the rocket motor 13, that is, to increase the thrust generated.

  The flying object propulsion control device includes a plurality of rocket motors 13 each having a solid propellant 17 loaded in the primary combustion chamber 14 and a secondary combustion chamber 19 communicated with the primary combustion chamber 14. Yes.

  When the flying object 10 flies (flys) in the air, the speed of the airframe 11 gradually decreases due to air resistance. In the conventional flying object, the initial speed of the airframe is sufficiently increased to ensure the final speed by burning the solid rocket motor with a large thrust once. As a result, the higher the speed, the greater the air resistance, and the propulsion energy is wasted, and the initial acceleration becomes too high. This will be described with reference to FIGS. 2A is a velocity pattern diagram showing the relationship between the speed of the flying object 10 and time, FIG. 2B is a thrust pattern diagram showing the relationship between the thrust of the flying object 10 and time, and FIG. It is a flight distance pattern figure which shows the relationship between the flight distance of the flying body 10, and time.

  As shown in FIGS. 2A to 2C, the conventional flying object propulsion control device has a curve indicated by a broken line, whereas the flying object propulsion control device of the first embodiment has a curve indicated by a solid line. Become. That is, in the first embodiment, since the plurality of rocket motors 13 can individually burn based on the signal from the outside, the combustion of the next rocket motor 13 is completed after the combustion of the initial rocket motor 13 is completed. After the combustion is completed, the next rocket motor 13 can be burned. As a result, the flight distance of the flying object 10 can be extended, and the flight time can be extended while maintaining a predetermined flight speed.

  On the other hand, in the case of a structure in which two or more solid propellants 31 and an ignition device 40 are arranged in the flying body 30 as shown in FIG. In addition, the structure of the pressure-resistant partition 41 is complicated and the thrust is limited to a certain level. Moreover, it may be difficult to incorporate three or more solid propellants 31 and the ignition device 40 in terms of arrangement of components or protection of heat and pressure.

  However, in the first embodiment, since the primary combustion chamber 14 of each rocket motor 13 is partitioned from the secondary combustion chamber 19 by a pressure-resistant and heat-resistant closure 16, measures for heat and pressure are facilitated, and the rocket motor It is also structurally easy to increase the number of 13 shots and generate a plurality of thrusts. Moreover, since a plurality of rocket motors 13 can be ignited simultaneously, if necessary, all the rocket motors 13 can be burned in a short time and accelerated at once.

Next, the shape and arrangement of the rocket motor 13 and the type of the solid propellant 17 will be described.
Although the shape of the rocket motor 13 of the first embodiment is a cylindrical shape, a hexagonal cylindrical shape or a fan-shaped cylindrical shape may be used in order to arrange a plurality of rocket motors 13 in the airframe 11 with high density. When the primary combustion chamber 14 in the rocket motor 13 is used at a high pressure, a cylindrical shape is adopted in consideration of the pressure resistance, and when the filling rate is important, a hexagonal cylinder shape or a fan-shaped cylinder shape is used, and the combustion pressure is Can be reduced. As the solid propellant 17 loaded in the primary combustion chamber 14, a known propellant such as a double base propellant or a composite propellant is used, and is not particularly limited.

  Further, since the plurality of rocket motors 13 are arranged so as to extend in a direction parallel to the axis X of the airframe 11, the solid propellant is placed in all the primary combustion chambers 14 at once in the manufacturing process of the solid propellant 17. 17 can be loaded, and the manufacturing operation thereof is facilitated. Furthermore, the ignition device module, the propellant module, and the secondary combustion chamber 19 with a nozzle are arranged in order from the front to the rear of the airframe 11, and it is possible to manufacture and assemble each module in the manufacture of the flying object propulsion control device. Cost reduction can be achieved. As described above, by arranging a plurality of rocket motors 13 in the axial direction X of the airframe 11, the rocket motors 13 can be disposed in the airframe 11 at a high density, and the size and weight can be reduced. In addition, the flying object propulsion control device can be manufactured easily, and a great advantage can be obtained economically.

Next, the combustion method of the solid propellant 17 will be described.
Since each rocket motor 13 preferably employs an internal combustion method, each solid propellant 17 has a structure having an inner hole 18 therein. In the example of FIG. 1, the inner hole 18 has a round hole shape so that the thickness of the solid propellant 17 is uniform in any direction, but the inner periphery as used in a general solid propellant 17 is used. The shape of the part may be a star shape or a slot shape (a shape having a slit). It is preferable that the inner hole 18 has an appropriate shape according to the cross-sectional shape of the primary combustion chamber 14 and the required thrust pattern. However, in this embodiment, since a large number of rocket motors 13 are arranged adjacent to each other, it is necessary to suppress the inner wall surface of the primary combustion chamber 14 from being heated by the combustion flame as much as possible. A round hole or a star-shaped hole that expands in order is a preferable shape.

The effects exhibited by the above first embodiment will be summarized below.
· Projectile propulsion control apparatus of the first embodiment includes a plurality of rocket motors 13 having a solid propellant 17 loaded in the primary combustion chamber 14, a primary combustion chamber 14 Ru communicates with the secondary combustion chamber 19 It has. That is, the main nozzle 12 of the secondary combustion chamber 19 is shared with the primary combustion chamber 14 of each rocket motor 13. Thus, the primary combustion chamber 14 of a plurality of rocket motors 13 having a solid propellant 17, in combination with the secondary combustion chamber 19 Ru communicated, in particular control with the promotion of the projectile 10 of the thrust control, etc. Can be performed efficiently. In addition, a valve and a pintle are not required as in the prior art, and an extra nozzle is not required, and a simple structure can reduce the size and weight.

  Each rocket motor 13 is configured such that the solid propellant 17 is burned by individually operating the ignition device 15 based on an external ignition signal. For this reason, each rocket motor 13 can be ignited at an arbitrary timing, and thrust can be generated based on the ignition signal when necessary, so that the average flight speed and the flight time can be increased. Further, since it is possible to simultaneously ignite one or more arbitrary numbers of rocket motors 13, a thrust of an arbitrary magnitude can be obtained by simultaneous ignition of a plurality of rocket motors 13 when necessary. 10 flight performance can be improved.

  The solid propellant 17 is formed in a cylindrical shape, and is configured such that combustion proceeds from the inner side of the inner hole 18 to the outer side, that is, radially outward from the inner peripheral surface of the inner hole 18. Therefore, it is possible to suppress the combustion heat of the rocket motor 13 from being transmitted to the inner wall surface of the primary combustion chamber 14, and it is difficult to raise the temperature of the surrounding rocket motor 13, and the combustion performance can be stabilized.

  In addition, a plurality of rocket motors 13 are arranged so as to extend in a direction parallel to the axis X of the body 11. Therefore, a plurality of rocket motors 13 can be arranged with high density, and the flying body propulsion control device can be reduced in size and weight. In addition, since the solid propellant 17 can be filled from one direction at a time into the primary combustion chambers 14 of the rocket motors 13 arranged in parallel at the same time, the rocket motor 13 can be easily manufactured. This is also advantageous from the viewpoint of economics during production.

In addition, since the main nozzle 12 of the secondary combustion chamber 19 is disposed so as to extend in a direction parallel to the axis X of the flying object 10, thrust control of the flying object 10 can be particularly easily performed. .
(Second Embodiment)
Next, a second embodiment that embodies the flying object propulsion control device of the present invention will be described with reference to FIG. In the second embodiment, differences from the first embodiment will be mainly described. In addition, the same member number is attached | subjected and demonstrated about the same member as 1st Embodiment.

  FIGS. 3A to 3E are views showing a flying object propulsion control apparatus according to a second embodiment, and FIG. 3A is a front view showing a main part of the flying object propulsion control apparatus in a broken state. 3B is an enlarged cross-sectional view of the rocket motor, FIG. 3C is a cross-sectional view taken along line 3c-3c in FIG. 3B, and FIG. 3D is taken along line 3d-3d in FIG. Sectional drawing and FIG.3 (e) are sectional drawings in the 3e-3e line | wire of Fig.3 (a).

  As shown in FIG. 3 (a), the secondary combustion chamber 19 is provided in the central portion (center of gravity) of the airframe 11, and as shown in FIG. Each secondary combustion chamber 19a is formed in a substantially triangular prism shape. A sub-nozzle 21 having a throat portion 21 a is arranged on the outer peripheral portion of each secondary combustion chamber 19 a so as to extend in a direction perpendicular to the axis X of the body 11 (radial direction). Therefore, a total of four sub nozzles 21 are provided. The number of sub-nozzles 21 in the secondary combustion chamber 19a can be appropriately set according to the combustion performance of the solid propellant 17, the size of the secondary combustion chamber 19a, the size of the flying object 10, its weight, the control purpose, and the like. it can. Here, the orthogonal direction does not necessarily have to be perpendicular (90 degrees) to the axis X of the airframe 11, and means a direction close to a right angle.

  A plurality (32 in the second embodiment) of the rocket motors 13 are disposed at the rear position of the secondary combustion chamber 19a, so that each of the eight rocket motors 13 faces the four secondary combustion chambers 19a. The eight primary combustion chambers 14 communicate with one secondary combustion chamber 19a. As shown in FIG. 3B, the ignition device 15 is disposed at the rear end position of the primary combustion chamber 14, and the closure 16 is disposed at the front end position. Moreover, as shown in FIG.3 (c), the solid propellant 17 is formed in the cylindrical shape similarly to 1st Embodiment, and has the inner hole 18. As shown in FIG. In the flying body 10 of the second embodiment, although not shown, a solid propellant 17 for injecting combustion gas from the main nozzle 12 is loaded in the rear part of the body 11, and the axis X direction of the body 11 It is comprised so that the thrust to can be obtained.

  Therefore, according to the flying object propulsion control device of the second embodiment, the sub-nozzle 21 of the secondary combustion chamber 19 is arranged so as to extend in the direction orthogonal to the axis X of the flying object 10, so that the flying object 10 can be easily controlled, and thrust generated in the lateral direction of the flying object 10 can be used as a side thruster. The timing of this thrust is controlled by the timing of combustion of the solid propellant 17. Moreover, since it is possible to ignite different rocket motors 13 connected to the secondary combustion chamber 19a, thrust can be generated in the resultant direction of the thrust vector, and flight control characteristics can be greatly enhanced. .

  Further, since the space between the secondary combustion chambers 19a is divided into four by the partition wall 20, the number and timing of the solid propellants 17 to be burned at a time can be controlled, so that the number of solid propellants 17 to be loaded is minimized. It is possible to control the size and direction of the fine thrust. In addition, once the combustion gas is accumulated in the secondary combustion chamber 19a, the combustion reaction can be sufficiently advanced, and the sub-nozzle 21 having an appropriate opening ratio can be disposed with sufficient space. The combustion efficiency and the nozzle efficiency are improved as compared with the conventional method in which the combustion gas is directly ejected from the sub nozzle 36 provided in each small rocket motor 38.

  In addition, the primary combustion chamber 14 of the rocket motor 13 is communicated with four secondary combustion chambers 19a, but the flow path between the primary combustion chamber 14 and the secondary combustion chamber 19a is narrowed. Thus, even if the pressure in the secondary combustion chamber 19a changes, the combustion pressure in the primary combustion chamber 14 can be kept constant. Therefore, the generated thrust can be increased without changing the combustion time of the rocket motor 13.

It should be noted that the present embodiment can be implemented with the following modifications.
-In 1st Embodiment, it can also comprise so that the secondary combustion chamber 19 may be divided | segmented into plurality. In that case, the control mode of thrust control can be changed as in the second embodiment.

In the second embodiment, it is also possible to place as in the first embodiment and a positional relationship between the primary combustion chamber 14 and the secondary combustion chamber 19a in the opposite.

  The configuration of the second embodiment, that is, a plurality of rocket motors 13 and a plurality of secondary combustion chambers 19a may be provided at a position ahead of the rocket motor 13 of the first embodiment. In that case, the flying object 10 can be controlled by combining thrust control and attitude control.

The rocket motor 13 may be arranged so as not to be symmetric with respect to at least one of the axis X, the horizontal line Y, and the vertical line Z of the body 11.
It is also possible to configure the solid propellant 17 as a solid body that does not have the inner hole 18.

The flying object propulsion control device may be provided with a timer so that thrust control and attitude control of the flying object 10 are performed at predetermined intervals.
Further, the technical idea that can be grasped from the embodiment will be described below.

· Said plurality of rocket motors, projectile propulsion control apparatus according to any one of claims 1 to 3, which is arranged at a position which is a point symmetry with respect to the axis of the projectile. In such a configuration, in addition to the effect of the invention according to any one of claims 1 to 3, the control associated with the promotion of the projectile in a stable state, can be easily performed.

The flying object propulsion control device according to any one of claims 1 to 3 , wherein the plurality of rocket motors are arranged so as to be closest packed in the flying object. In such a configuration, in addition to the effects of the invention according to any one of claims 1 to 3 , the control associated with the propulsion of the flying object can be performed more efficiently, and the flying object can be reduced in size. Can do.

(A) is the front view which fractures | ruptures and shows the principal part of the flying object propulsion control apparatus of 1st Embodiment, (b) is an expanded sectional view of a rocket motor, (c) is the cross section in the 1c-1c line of (b). The figure, (d) is sectional drawing (a) in the 1d-1d line | wire of (a). (A) is a graph showing the relationship between time and the velocity of the flying object, (b) is a graph showing the relationship between time and the thrust of the flying object, and (c) shows the relationship between time and the flying distance of the flying object. Graph. (A) is the front view which fractures | ruptures and shows the principal part of the flying object propulsion control apparatus of 2nd Embodiment, (b) is an expanded sectional view of a rocket motor, (c) is the cross section in the 3c-3c line | wire of (b). The figure, (d) is sectional drawing in the 3d-3d line of (a), (e) is sectional drawing in the 3e-3e line of (a). The front view which fractures | ruptures and shows the principal part of the conventional flying object propulsion control apparatus. (A) is the front view which fractures | ruptures and shows the principal part of another conventional flying object propulsion control apparatus, (b) is sectional drawing in the 5b-5b line | wire of (a). The front view which fractures | ruptures and shows the principal part of the conventional flying object propulsion control apparatus which can be considered.

Explanation of symbols

  DESCRIPTION OF SYMBOLS 10 ... Flying object, 12 ... Main nozzle as nozzle, 13 ... Rocket motor, 14 ... Primary combustion chamber, 15 ... Ignition device, 17 ... Solid propellant, 18 ... Inner hole, 19, 19a ... Secondary combustion chamber, 21 ... Sub nozzle as a nozzle, X ... Axis center.

Claims (3)

Provided flying body, the closure formed in the ignition device and the front end portion in a pressure heat-resistant material to the rear end portion is loaded into the primary combustion chamber which is provided, a cylindrical solid propellant having an inner bore a plurality of the rocket motor, and a plurality of secondary combustion chamber which is constructed as a single secondary combustion chamber is communicated to said plurality of primary combustion chamber, said primary combustion chamber and the secondary combustion with A nozzle having a throat having a smaller cross-sectional area than the throat portion of the main nozzle is installed between the chamber and a sub- nozzle having a throat portion in the secondary combustion chamber, and a solid propellant loaded in the rear portion Combustion gas is ejected from the main nozzle to obtain thrust,
Each solid propellant is configured so that the solid propellant is burned by individually operating an ignition device based on an external ignition signal, and the sub-nozzle of the secondary combustion chamber is positioned at the center of gravity of the flying object. The flying object propulsion control apparatus is arranged so as to extend in a direction orthogonal to the axis of the flying object. 2. The flying object propulsion control device according to claim 1, wherein each of the rocket motors is configured such that combustion of a subsequent rocket motor is started after combustion of the preceding rocket motor is completed. The flying object propulsion control apparatus according to claim 1, wherein the plurality of rocket motors are arranged so as to extend in a direction parallel to the axis of the flying object.

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