CN218034681U - Small missile - Google Patents

Abstract

The application discloses small-size guided missile, it includes guide head cabin, guidance control cabin, rudder cabin, draws battle cabin, engine compartment and the fin that connects gradually from the warhead to the missile tail, wherein, controlgear on the guided missile is located respectively guide head cabin, guidance control cabin and rudder cabin, controlgear's connecting cable arranges the inside in guide head cabin, guidance control cabin, rudder cabin. The guidance control cabin integrates 90% of electrical equipment on the missile, and comprises a falling plug, a guidance control device, a thermal battery and a cabin body structure, and all control equipment of the small missile can be formed only by connecting the front end of the guidance control device with a guidance head and connecting the rear end of the guidance control device with a rudder cabin.

Description

Small missile

Technical Field

The invention relates to the field of aerospace weapons, in particular to a structural design of a small missile.

Background

Along with the development of missile weapon systems, unmanned system combat has become a new trend of the development of missile weapon equipment. Unmanned combat platforms such as unmanned aerial vehicles and unmanned vehicles can show hands in local wars, and the unmanned combat platforms carrying weapons inevitably become important directions for weapon research and development, can meet the mounting requirements of unmanned aerial vehicles and ground unmanned vehicles, and can attack miniaturized missiles on battlefields, which becomes a new demand point for current weapon system development. The unmanned launching platform has great use requirements on small-sized missiles, so that the development of the small-sized missiles with the characteristics of long range, low cost and convenient use has wide market prospect.

Due to size limitation, the small missile has a plurality of limiting factors on the structural layout and the cabin section design of the missile. The conventional missile is generally in a normal aerodynamic layout and is generally divided into a nose cabin, a control cabin, a war induction cabin, an engine cabin and a rudder cabin from a warhead to a missile tail, wherein the rudder cabin is positioned at the tail part of the missile.

Because the rudder cabin is at the back and the control cabin is at the front, the control cabin needs to communicate and supply power with the steering engine through the missile-mounted cable, and because the cable can not be threaded in the engine cabin section, a cable network needs to be arranged on the missile-mounted cable supply groove reserved on the outer wall of the missile for routing.

Therefore, the small missile needs to design a cabin section layout with compact structure and high integration level.

In addition, the electrical system of a missile generally comprises a missile cable network, a missile-borne computer, an inertia measurement unit, a power distribution system, a thermal battery, a steering engine system and a guidance system. Each system needs to be customized according to the functional layout on the missile and is distributed in each cabin section structure of the missile. The distributed design can lead to the dispersion of equipment layout, and the unit equipment number is more, and the connection cable connection between equipment is more, can't realize the demand that small-size guided missile highly integrates to control system.

Therefore, the electrical control part of the small missile needs to be designed to realize the above-mentioned compact and high-integration cabin section layout alone or in combination with the improvement of other missile structures.

Disclosure of Invention

One of the objects of the present invention is: according to the functional requirements of the small missile, the small missile with compact structure and higher integration level is designed.

For this reason, some embodiments of the present application provide a small missile, which is a guidance head cabin, a guidance control cabin, a rudder cabin, a guidance control cabin, an engine cabin and a tail wing in sequence from a warhead to a missile tail, wherein control devices on the missile are respectively located in the guidance head cabin, the guidance control cabin and the rudder cabin, and connection cables of the control devices are arranged inside the guidance head cabin, the guidance control cabin and the rudder cabin.

In some embodiments, the rudder nacelle comprises a nacelle shell and four X-shaped trailing rudder pieces coupled to the nacelle shell, the four trailing rudder pieces being equally angularly spaced on an outer contour of the nacelle shell. Each trailing edge rudder piece comprises a fixed surface and a control surface, and the control surface is connected with the steering engine to realize angle control of the control surface, so that the attitude control of the whole missile is realized.

In some embodiments, the securing surface is secured to the deck section housing by clips. The four steering engines are connected with a guidance control device of the control cabin through internal cables.

In some embodiments, the guidance control cabin comprises a cabin body, wherein the cabin body is a cylinder with the same diameter as the projectile body and is provided with a front sleeving section and a rear sleeving section which are positioned at two end parts, and an inner cavity is formed inside the cylinder; the front sleeving section is connected with the guide head cabin through a radial screw, and the rear sleeving section is connected with the rudder cabin through a radial screw; the shedding plug protrudes out of the circumference of the cabin body of the guidance control cabin, the shedding plug is fixed on a shedding plug mounting plate, the shedding plug mounting plate comprises a fixing part and a flange part, the fixing part is fixed in the inner cavity of the cabin body, the flange part and the fixing part are basically arranged perpendicularly and can be integrally formed, and the flange part extends out of the outer side of the side wall of the cabin body, for example, extends out of the side wall of the cabin body perpendicularly; the breakaway plug is detachably attached to the flange portion by a screw connection, snap fit, or other means.

In some embodiments, the guidance control device is comprised of a computer board, an interface board, and an inertial measurement unit; the computer mounting plate is radially fixed with a radially arranged shoulder structure close to a front sleeving section of the cabin body; the front side of the computer mounting plate is arranged and connected with the inertia measurement unit, and the rear side of the computer mounting plate is arranged with a computer plate and an interface plate which are axially fixed with the interface plate.

In some embodiments, a ZYNQ main control unit is configured on the computer board, the ZYNQ main control unit collects inertial measurement unit data through an SPI interface implemented by a first rigid-flexible board, and performs error compensation on the measurement unit data; and the ZYNQ main control unit integrates the control function of the steering engine in the rudder engine room at the same time.

In some embodiments, an interface conversion chip circuit can be further configured on the computer board, and the interface conversion chip circuit is connected with the interface board through a second rigid flexible board.

In some embodiments, the tail is a rotating tail structure comprising a sleeve and tail fins rotatably fitted on an outer wall of the sleeve; the sleeve is provided with an annular sleeve joint surface for connecting an engine compartment and a sleeve body formed by extending from the sleeve joint surface along the axial direction, and a flange is formed between the sleeve joint surface and the sleeve body; the tail fin piece cylindrical tail fin shell is fixed on the outer wall of the cylindrical tail fin shell at intervals of the same angle; each tail fin is fixed on the cylindrical tail shell through a clamping piece.

In some embodiments, the inner contour of the tubular tail shell is matched with the outer contour of the sleeve body of the sleeve through two bearings, and the two bearings are sleeved outside the sleeve. The tail wing can rotate freely without being controlled, when the guided missile is installed on the guided missile frame, the guided missile can collide with the guided missile frame without limiting the rotation of the guided missile frame, and for the condition, a stop ring can be arranged at the tail of the cylindrical tail wing shell to limit the axial movement of the cylindrical tail wing shell and prevent the cylindrical tail wing shell from interfering with the guided missile frame.

In some embodiments, the connecting frames of adjacent cabins are respectively processed into inner and outer surfaces of a cylinder, and are sleeved by matching surfaces of the inner and outer surfaces and then connected and fixed along the circumference from a radial row to form a whole.

In some embodiments, the control surface is a low aspect ratio trailing edge duck control surface; the tail fin is a high aspect ratio tail fin.

According to the small missile, all cabin sections are connected in a modularized manner, and different guide heads, different guidance control cabins, different warheads, different rudder cabins and different empennage structures can be assembled. Function integration of the guidance control cabin is achieved through function distribution of all cabin sections, and no control equipment is arranged at the tail part, so that a wiring groove does not need to be reserved outside the projectile body.

In some embodiments, the pneumatic scheme of the small missile adopts four rudder pieces in an X-shaped arrangement and four tail wing pieces in an X-shaped arrangement to form a duck-shaped arrangement, wherein a warhead is a seeker cabin section, a guidance control cabin section is arranged behind the seeker cabin section, a stabilizing plane and a small-aspect-ratio rear edge duck rudder are arranged on the helm cabin section, and a tail part selects a rotatable empennage with a large aspect ratio, so that rolling interference can be reduced, and sufficient lift force can be provided.

Small missiles according to some embodiments of the present application are in a duck-type configuration, the electrical system of the entire missile is entirely centralized in front of the missile warfare compartment, and the missile-borne cables are entirely routed through the inside of the missile. Considering the technical maturity and the bullet finishing cost, the guide head with the diameter slightly larger or smaller than the diameter of the bullet body can be selected as the guide head; the structure is provided with a section connecting section.

According to the small missile of some embodiments of the application, the integrated design of the guidance control device is adopted, and all control devices except a guidance head are concentrated in the cabin body of the guidance control device, so that the number of the control devices on the missile is reduced, the small-sized design of the small missile is facilitated, and the modularized reloading capacity of the missile is improved.

According to the small missile provided by some embodiments of the application, the guidance control device integrates a steering engine control function, only a steering engine and a cable are left in a steering engine cabin, and a separate steering engine controller is not required to be arranged.

This summary is not intended to identify key or essential features of the claimed subject matter, nor is it intended to be used alone to determine the scope of the claimed subject matter. The subject matter should be understood by reference to appropriate portions of the entire specification of the disclosure, any or all of the drawings, and each claim.

The foregoing and other features and examples will be described in more detail in the following specification, claims and drawings.

The terms and expressions which have been employed are used as terms of description and not of limitation, and there is no intention in the use of such terms and expressions of excluding any equivalents of the features shown and described or portions thereof. Nevertheless, it will be understood that various modifications may be made within the scope of the claimed system and method. Thus, while the present systems and methods have been specifically disclosed by way of examples and optional features, those skilled in the art will recognize modifications and variations of the concepts disclosed herein, and such modifications and variations are considered to be within the scope of the systems and methods as defined by the following claims.

Drawings

The features of the various embodiments described above, as well as other features and advantages of certain embodiments of the present invention, will become more apparent from the following detailed description when taken in conjunction with the accompanying drawings, wherein:

FIG. 1 is a schematic illustration of a missile structural layout from the side according to an embodiment of the application;

FIG. 2 is a side view schematic illustration of a rudder nacelle structure according to an embodiment of the present application;

FIG. 3 is a schematic cross-sectional structural view of a guidance control bay according to an embodiment of the present application;

FIG. 4 is a schematic diagram of the electrical connections of the guidance control bay according to an embodiment of the present application;

FIG. 5 is a schematic side view of a configuration of a guidance control bay according to an embodiment of the present application from the front side;

FIG. 6 is a schematic plan view of a configuration of a guidance control cabin according to an embodiment of the present application;

FIG. 7 is a schematic side view of a configuration of a guidance control bay according to an embodiment of the present application from the rear side;

FIG. 8 is a top view of a rotating tail structure according to an embodiment of the present application;

fig. 9 is an exploded view of the structure of the rotating tail according to an embodiment of the present application.

Detailed Description

In the following description, various examples of small missiles are described. For purposes of explanation, specific configurations and details are set forth in order to provide a thorough understanding of the embodiments. It will be apparent, however, to one skilled in the art that certain embodiments may be practiced or carried out without each of the disclosed details. Furthermore, well-known features may be omitted or simplified to help prevent any confusion over the novel features described herein.

The following high-level summary is intended to provide a basic understanding of some novel innovations depicted in the accompanying drawings and presented in the corresponding description provided below.

Referring to fig. 1, a small missile 100 according to an embodiment of the present application is composed of a guidance head cabin 5, a guidance control cabin 4, a rudder cabin 3, a guidance cabin 2, an engine cabin 1, a tail wing 6, and the like, which are connected in sequence; the pneumatic scheme adopts an X-shaped and X-shaped duck layout; by adopting a modular structure scheme, each module can be replaced and expanded.

Specifically, a solid rocket engine and an ignition head (not shown) are arranged in the engine compartment 1, ignition of the engine is controlled by adopting a guidance control device, and ignition of the engine is controlled by a cable embedded in the missile compartment 2.

The fuse warfare cabin 2 comprises a fuse, a warfare part and a fuse warfare cabin body structure, the fuse is connected with a guidance control device 44 in the control cabin 4 through a cable on the bullet, and the fuse cable needs to penetrate through the rudder cabin 3.

As shown in fig. 1 and 2, the rudder nacelle 3 includes a nacelle shell 31 and four X-shaped trailing rudder pieces 32 coupled to the nacelle shell 31, and the four trailing rudder pieces are equally angularly spaced on the outer contour of the nacelle shell 31. Only two trailing edge flaps can be shown in fig. 1, 2 because of the perspective. Each trailing edge rudder 32 comprises a fixed surface 321 and a control surface 322, the control surface 322 is connected with a steering engine (not shown) to realize control surface angle control, so as to realize attitude control of the whole missile, wherein the fixed surface 321 is fixed with the cabin shell 31 through a clamping piece 33. The four steering engines are connected with a guidance control device of the control cabin 4 through internal cables. The trailing edge rudder 32 can also be replaced by a full-motion rudder.

The front side of the guidance control cabin 4 is provided with control equipment comprising a falling plug 41 and a guidance control device 44, the rear side is provided with control equipment comprising a thermal battery 46 and a corresponding cabin body 42, the falling plug 41 protrudes out of the cabin body and is positioned right above the control cabin, and the falling plug automatically falls off after the guided missile is ignited.

As shown in fig. 3, 5, 6 and 7, the guidance control cabin 4 includes a cabin body 42, the cabin body 42 is a cylinder with the same diameter as the projectile body, such as a cylinder, and has a front socket section 421 and a rear socket section 422 at two ends, and an inner cavity 423 is formed inside the cylinder. The front socket joint section 421 is connected with the guide head cabin 5 through a radial screw, and the rear socket joint section 422 is connected with the rudder cabin 3 through a radial screw. The connection mode of the cabin segment pieces is exemplarily described, and the connection mode can also be applied to the connection of other adjacent cabin segments.

The disconnection plug 41 projects outside the circumference of the housing 42 of the guidance control pod, for example, directly above the circumference shown in fig. 3. The falling plug 41 is fixed on the falling plug mounting plate 411, and the falling plug mounting plate 411 includes a fixing portion 4111 and a flange portion 4112, the fixing portion 4111 is fixed on the inner cavity 423 of the cabin 42, the flange portion 4112 is substantially vertically arranged with the fixing portion 4111 and may be integrally formed, and the flange portion 4112 extends out of the outer side of the side wall 424 of the cabin 42, for example, extends out perpendicularly from the side wall 424 of the cabin 42. The drop-off plug 41 is detachably attached to the flange portion 4112 by screwing, snapping, or other means. The falling plug adopts a mechanical separation mode, and the falling plug 41 is separated by the thrust of an engine after the missile is ignited.

A conformal shield 43 is included in the missile-facing direction of travel of the breakaway plug 41 to reduce aerodynamic drag. Conformal shell 43 may be formed by extending from sidewall 424 of body 42 to the maximum outer diameter of the breakaway plug 41 to form a tapered surface.

As shown in fig. 3, the guidance control device 44 is disposed inside the cabin body of the guidance control cabin, and is constituted by a computer board 441, an interface board 442, and an inertia measurement unit 443. Wherein, on the front side of the guidance control cabin, i.e. the advancing direction side of the missile, a computer mounting plate 445 is arranged, and the computer mounting plate 445 is fixed with the internal structure of the cabin body 42, for example, a radially arranged shoulder structure 4211 close to the front socket section 421 of the cabin body 42 by radial screws; the inertia measurement unit 443 is disposed on the front side of the computer mounting plate 445, the computer plate 441 and the interface plate 442 are disposed on the rear side of the computer mounting plate 445, and the computer plate 441 and the interface plate 442 are axially fixed by screws, wherein the computer plate 441 and the interface plate 442 are connected and communicated through a rigid-flexible plate. The computer mounting plate 445, the computer plate 441, and the interface plate 442 are configured to have an arcuate, or bow-shaped, configuration, with the remaining space of the interior chamber 423 being a pop-up cable network routing channel, the cable network illustratively including a first cable 451, a second cable 452, and a third cable 453.

The interior functional connection schematic diagram of the guidance control cabin is shown in figure 4. A high-performance ZYNQ 7020 minimum system 4411 or ZYNQ main control unit 4411 can be configured on the computer board 441, the flight control function of a conventional missile-borne computer is completed, IMU data acquisition and calibration processing functions are integrated, the ZYNQ main control unit 4411 acquires the data of the inertia measurement unit 443 through an SPI interface realized by the first rigid-flexible board 4412, and error compensation is carried out on the data of the measurement unit; the ZYNQ main control unit 4411 simultaneously integrates the control function of the steering engine in the rudder engine room, for example, the steering engine control is performed on the rudder engine room 3 through PWM, and the ADC performs steering engine angle feedback sampling. The guidance control device 44 is responsible for carrying out unified power supply and distribution control on the guidance head, the inertia measurement unit, the steering engine PWM signal and other power utilization equipment in the guidance head cabin.

An interface conversion chip circuit 4413, such as a conversion circuit based on RS-232 or RS-422 protocol, may also be disposed on the computer board 441, and is coupled to the interface board 442 via a second rigid-flexible board 4414. The computer board 441 may be further configured with a third conversion circuit 4415 to receive power from the interface board 442 and convert the power into a ZYNQ main control unit 4411 and an interface conversion chip circuit 4413.

The interface board 442 can be arranged with a steering engine power driving circuit 4421, a timing distribution circuit 4422, and an external plug connector 4423, which is connected with a drop plug, a thermal battery, a seeker, a fuse, etc. The interface board may further be provided with a primary power supply combiner 4424, a secondary power supply conversion circuit 4425 for supplying power to the ZYNQ main control unit 411, and an ADC acquisition circuit 4426 for supplying a steering engine angle feedback signal to the ZYNQ main control unit 4411.

A heat shield plate 461 is arranged between the thermal battery 46 and the guidance control device 44 to prevent heat generated after the thermal battery 46 works from influencing the guidance control device 44. The thermal battery 46 is fixedly mounted to the housing 424 of the nacelle 42 by radial screws. In some embodiments, thermal battery 46 is designed to be arched to maximize cabin space utilization, the bottom of thermal battery 46 is the routing channel for pop-up cables 452, 453 that extend through the bottom of thermal battery 46 to the rear of the guidance control cabin.

The working flow of the small missile 100 in the embodiment of the application is as follows:

the small missile 100 automatically runs the flight control software of the missile after being electrified, the flight control software acquires gyro and acceleration information of an inertia measurement unit through the SPI, receives data of a seeker through a seeker serial port, acquires initial alignment information of an external launching platform through a fire control serial port, completes launching-leading missile self-checking, initial binding and initial alignment, simultaneously sends state information of the missile through a ground test serial port, and the missile controls a thermal battery to be activated and an engine to be ignited according to a fire control serial port instruction; after the missile is launched, the target position is obtained according to the capture information of the seeker, fuze relief is carried out, the steering engine is controlled to deflect, and the posture of the missile is changed to control the missile to fly to the target.

The guidance control cabin integrates 90% of electrical equipment on the missile, comprises a shedding plug, a guidance control device, a thermal battery and a cabin body structure, and only needs to be connected with a seeker at the front end and a rudder cabin at the rear end, so that all control equipment of the small missile is formed.

The guide head cabin 5 can realize modularized reloading, and can be provided with a laser guide head, an infrared guide head, a radar guide head or a visible light guide head, wherein different guide heads use the same mechanical and electrical interface, the flight control software is communicated with the guide heads after being electrified to identify different guide head types, and different terminal guide software is loaded according to different guide head types to carry out terminal guide information calculation.

The missile tail wing 6 adopts a rotary tail wing structure, and the whole structure is as shown in figures 8 and 9 and comprises a sleeve 61 and tail fins 62 which are rotatably matched on the outer wall of the sleeve 61. The sleeve 61 has a socket surface 611 for annular coupling with the engine compartment and a sleeve body 612 formed to extend axially from the socket surface 611, with a flange formed between the socket surface 611 and the sleeve body 612. Tail fin 62 a tubular tail housing 621, four tail fins fixed at equal angular intervals on the outer wall of the tubular tail housing 621, i.e., a first tail fin 622, a second tail fin 623, a third tail fin 624, and a fourth tail fin, which is not shown in the drawing because of the view angle. Each tail fin is secured to a cylindrical tail shell 621 by clips 627. The inner contour of the tubular tail housing 621 is matched with the outer contour of the sleeve body 612 of the sleeve 61 through two bearings 625A and 625B to realize rotatable matching, and the two bearings 625A and 625B are sleeved outside the sleeve. A stop ring 626 may be provided at the tail of the cylindrical tail housing 621 to limit axial movement of the cylindrical tail housing 621 to prevent interference with the hair guide. The stop ring 626 is fixed to the outer wall of the sleeve 61 by radial screws. The tail wing pieces 62 can freely rotate on the sleeve 61, and when the hair guiding frame is hung in a naked mode, the tail wing pieces are limited and fixed through the rear sliding blocks 73, so that the tail wing pieces 62 are prevented from rotating to damage the structure. After firing, tail flaps 62 are free to rotate. The limiting slide block is fixed right above the cylindrical empennage shell 621 through a radial screw.

In order to meet the launching requirements of the missile launcher, the small missile is provided with three sliding blocks which are divided into a front sliding block 71, a middle sliding block 72 and a rear sliding block 73. The front slider 71 is located on the outer contour of the missile bay 2, for example, directly above the outer contour of the missile bay 2 as shown in fig. 1, the middle slider 72 is fixed on the outer contour of the engine bay 1 through a strap, for example, directly above the engine bay 1 as shown in fig. 1, the rear slider 73 is located on the outer contour of the circular cylinder 621 of the rotating tail fin 6, for example, above the circular cylinder 621 of the rotating tail fin 6 as shown in fig. 8, and the rotating tail fin 6 is limited before launching.

In the embodiment shown in fig. 1, the connecting frames of the adjacent full-elastic cabin sections are respectively processed into the inner surface and the outer surface of a cylinder, are sleeved by utilizing the matching surfaces of the inner surface and the outer surface, and are connected and fixed by a single row of radial screws along the circumference to form a whole.

The small missile in the embodiment of the application adopts the X-shaped arrangement of four rudder blades and the X-shaped arrangement of four tail wing blades to form a duck-shaped layout, and all cabin sections are connected in a modularized manner, so that different guide heads, different guidance control cabins, different warheads, different rudder cabins and different tail wing structures can be assembled.

When the trailing edge rudder piece and the rotary tail wing are selected, the rudder cabin section is provided with the stabilizing surface and the trailing edge rudder with small aspect ratio, and the tail part is provided with the rotary tail wing with large aspect ratio, so that the rolling interference can be reduced, and the enough lift force can be provided. When the duck-type layout is adopted, all the electrical systems of the whole missile are concentrated at the front part of the missile fighting cabin, and all the cables on the missile are wired through the inside of the missile without reserving wiring grooves on the outer wall of the missile.

In some embodiments of the application, the control devices except the seeker are completely integrated in the cabin body of the guidance control device through the integrated design of the guidance control device, the number of the electric devices on the missile is reduced, the small-sized design of the small-sized missile is facilitated, and meanwhile the modularized reloading capacity of the missile is improved.

In some embodiments of the application, because the guidance control device integrates a steering engine control function, only a steering engine and a cable are left in the steering engine room, and an independent steering engine controller is not required to be arranged.

The trailing edge rudder blade is selected for the rudder nacelle in the embodiment shown in fig. 1 to 9, and in other embodiments, the trailing edge rudder blade can be replaced by a full-motion rudder blade.

The tail of the embodiment shown in fig. 1 to 9 is a rotating tail, but in other embodiments, a folding rotating tail or a non-rotating fixed wing may be used instead.

In the above embodiment, the front slider and the middle slider are configured as separate parts and then connected to the corresponding cabin sections after molding; in other embodiments, the front and middle sliders may also be configured in such a way that they are integrally formed with the respective cabin segment.

In the above embodiments, the breakaway plug 41 is disposed on the guidance control pod 4 at the head of the projectile, while in other embodiments, the breakaway plug 41 may be disposed at the tail of the projectile; additionally or alternatively, the drop-off plug 41 may be configured to be embedded within the guidance control compartment, for example, having a conformal configuration with the compartment in which it is located.

In an alternative embodiment, the main control unit on the computer board 441 may be replaced by DSP + FPGA, or ARM + FPGA or other main control chips;

in an alternative embodiment, the drop plug 41 can be detached and inserted by a small wire cutter, and the detachment and insertion are performed in a tangent mode;

in an alternative embodiment, the breakaway plug 41 may be a small rectangular connector that is embedded within the housing of the guidance control device;

the control unit of the guidance control device is integrated into a computer board, an interface board, an inertia measurement unit and a thermal battery, wherein the functions of the computer board and the interface board can be split into a flight control computer, a navigation computer, a power distribution system and a steering engine control board.

The previous description of the disclosed embodiments is provided to enable any person skilled in the art to make or use the present invention. Various modifications to these embodiments will be readily apparent to those skilled in the art, and the generic principles defined herein may be applied to other embodiments without departing from the spirit or scope of the invention. Thus, the present invention is not intended to be limited to the embodiments shown herein but is to be accorded the widest scope consistent with the principles and novel features disclosed herein.

Moreover, while advantages associated with certain embodiments of the technology have been described in the context of those embodiments, other embodiments may also exhibit such advantages, and not all embodiments need necessarily exhibit such advantages to fall within the scope of the technology. Accordingly, the present disclosure and associated techniques may include other embodiments not explicitly shown or described herein. Accordingly, the disclosure is to be limited only by the following claims.

Reference numerals

Small missile 100

Engine compartment 1

War-leading cabin 2

Rudder nacelle 3

Cabin shell 31

Trailing edge rudder blade 32

Fixing surface 321

Control surface 322

Clamping piece 33

Guidance control cabin 4

Fall-off plug 41

Fall-off plug mounting plate 411

Fixing portion 4111

Flange portion 4112

Cabin 42

Front socket joint section 421

Rear socket joint section 422

Lumen 423

Conformal cover 43

Outer shell 424 of cabin 42

Guidance control device 44

Computer board 441

ZYNQ Master control Unit 4411

First rigid and flexible plate 4412

Interface conversion chip circuit 4413

Second rigid and flexible plate 4414

Interface board 442

Steering engine power drive circuit 4421

Timing distribution circuit 4422

Butt external plug connector 4423

Primary power supply combiner 4424

Secondary power supply conversion circuit 4425

ADC acquisition circuit 4426

Inertial measurement unit 443

Computer mounting plate 445

First cable 451

Second cable 452

Third cable 453

Thermal battery 46

Thermal shield 461

Guide head cabin 5

Empennage 6

Sleeve 61

Tail fin 62

Annular socket surface 611

Sleeve barrel 612

Tubular tail fin outer casing 621

First tail vane 622

Second tail fin 623

Third tail vane 624

Clamping piece 627

Bearings 625A, 625B

Retainer ring 626

Front slider 71

Middle slider 72

Rear slider 73

Claims (10)

1. A small missile, characterized by: the missile control system comprises a guide head cabin, a guidance control cabin, a rudder cabin, a war-guiding cabin, an engine cabin and a tail wing which are sequentially connected from a warhead to a missile tail, wherein control equipment on the missile is respectively positioned in the guide head cabin, the guidance control cabin and the rudder cabin, and connecting cables of the control equipment are arranged in the guide head cabin, the guidance control cabin and the rudder cabin.

2. A small missile according to claim 1, wherein: the rudder cabin comprises a cabin shell and four X-shaped trailing edge rudder pieces coupled to the cabin shell, and the four trailing edge rudder pieces are arranged on the outer contour of the cabin shell at equal angular intervals; each trailing edge rudder piece comprises a fixed surface and a control surface, and each control surface is connected with a steering engine to realize control surface angle control, so that the attitude control of the whole missile is realized.

3. A small missile according to claim 2, wherein: the fixed surface is fixed with the cabin section shell through a clamping piece; the four steering engines are connected with a guidance control device of the guidance control cabin through internal cables.

4. A small missile according to claim 1, wherein: the guidance control cabin comprises a cabin body, the cabin body is a cylinder with the same diameter as the projectile body, the cabin body is provided with a front sleeving section and a rear sleeving section which are positioned at two end parts, and an inner cavity is formed inside the cylinder; the front sleeve joint section is radially connected with the guide head cabin, and the rear sleeve joint section is radially connected with the rudder cabin; the shedding plug protrudes out of the circumferential outer side of the cabin body of the guidance control cabin, the shedding plug is fixed on a shedding plug mounting plate, the shedding plug mounting plate comprises a fixing part and a flange part, the fixing part is fixed in the inner cavity of the cabin body, the flange part and the fixing part are basically vertically arranged and can be integrally formed, and the flange part extends out of the outer side of the side wall of the cabin body, for example, extends out of the side wall of the cabin body vertically; the shedding plug is detachably attached to the flange portion.

5. A small missile according to claim 1, wherein: the guidance control device consists of a computer board, an interface board and an inertia measurement unit; the computer mounting plate is radially fixed with a radially arranged shoulder structure close to a front sleeving section of the cabin body; the front side of the computer mounting plate is arranged and connected with the inertia measurement unit, and the rear side of the computer mounting plate is arranged with a computer plate and an interface plate which are axially fixed with the interface plate.

6. A small missile according to claim 5, wherein: a ZYNQ main control unit is configured on the computer board, and the ZYNQ main control unit acquires inertial measurement unit data through an SPI (serial peripheral interface) realized by a first rigid flexible board and performs error compensation on the measurement unit data; and the ZYNQ main control unit integrates the control function of the steering engine in the rudder engine room at the same time.

7. A small missile according to claim 5, wherein: the computer board can be also provided with an interface conversion chip circuit which is connected with the interface board through a second rigid flexible board.

8. A small missile according to claim 1, wherein: the tail wing is of a rotary tail wing structure and comprises a sleeve and tail wing pieces which can be rotatably matched on the outer wall of the sleeve; the sleeve is provided with an annular sleeve joint surface for connecting an engine compartment and a sleeve body formed by extending from the sleeve joint surface along the axial direction, and a flange is formed between the sleeve joint surface and the sleeve body; the tail fin piece cylindrical tail fin shell is fixed on the outer wall of the cylindrical tail fin shell at intervals of the same angle; each tail fin is fixed on the cylindrical tail shell through a clamping piece.

9. The small missile according to claim 8, wherein the control surface is a low aspect ratio trailing edge duck control surface;

the tail vane is a high aspect ratio tail vane.

10. A small missile according to claim 9, wherein: the inner contour of the cylindrical empennage shell is matched with the outer contour of the sleeve body of the sleeve through two bearings so as to realize rotatable matching, and the two bearings are sleeved outside the sleeve.

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