A detachable attitude control device

Abstract

A detachable attitude control device comprises a guided aircraft (15), characterized in that an attitude control device (40) is installed at the front end of the guided aircraft (15), wherein the attitude control device (40) comprises a cylindrical shell (2) and an end cover (1), the end cover (1) is sealed and installed at the upper end of the cylindrical shell (2), and a pulse engine assembly, a controller (11), an inertial measurement unit (12) and a power supply unit (13) are installed in the cylindrical shell (2); the controller (11), the inertial measurement unit (12) and the power supply unit (13) are all installed on the inner wall of the cylindrical shell (2); the attitude control device of the present invention is installed at the front end of the guided aircraft, and can change the direction of the guided aircraft at a large angle during the launch process of the guided aircraft, and the speed of changing the direction of the guided aircraft is unattainable by previous attitude control devices.

Description

Technical Field

[0001] The present invention relates to the field of aircraft engineering, and in particular to a detachable attitude control device for a guided power-propelled combat aircraft (hereinafter referred to as a guided aircraft) to make a rapid turn after take-off, which is widely applicable to anti-ship cruise guided aircraft and air defense guided aircraft with different launch modes. Background Art

[0002] The inclined launch method of guided aircraft mainly relies on rotating launchers or launch boxes for aiming. For ship-borne guided aircraft, there are various facilities on the ship that block the target, and there are certain shooting blind spots. The vertical launch method does not require directional aiming, and has the advantages of fast response time, large payload, and all-round combat. It is widely used in modern anti-ship cruise guided aircraft, air defense guided aircraft and other fields. For medium and low altitude defense, offshore and sea surface defense or attack, the key technology of modern anti-ship cruise guided aircraft and air defense guided aircraft is its all-round rapid turning in the target direction, requiring the guided aircraft to complete the program turning with the minimum turning radius in the shortest time and quickly enter the range interval in order to expand the minimum range and maximum range interval. Since the flight speed of the guided aircraft is low in the initial ascent, it is impossible to achieve it only by air rudders. In order to achieve rapid turning in the ascent stage with low dynamic pressure, gas rudders, gas generators or pulse small engines are generally used to generate turning torque to turn the missile axis to the predetermined spatial orientation.

[0003] For example, a certain guided aircraft uses a gas rudder. The advantages are simple structure, easy combination with aerodynamic control surfaces, and ability to achieve full attitude stability and control. The disadvantages are large thrust loss, harsh working environment of the gas rudder, and rudder efficiency that varies with the initial charge temperature and degree of erosion.

[0004] A certain guided aircraft is based on the combination of a gas generator and the tail nozzle connected to it and an air rudder. The tail nozzle is designed to be perpendicular to the rudder surface, so as to effectively utilize the thrust generated by the jet flow ejected from the nozzle and reduce energy loss. However, this solution still has some disadvantages. Since the gas generator nozzle is integrated into the air rudder, the rudder root becomes thicker, and the rudder is in a high temperature and high pressure harsh environment. Overall, the weight, processing difficulty and cost of the air rudder are increased. In addition, the gas generator and its accessories cannot be separated after completing the rapid turn, and they always accompany the subsequent flight of the guided aircraft, resulting in slow acceleration after the main engine of the guided aircraft is ignited (relative to throwing away the negative weight), reducing maneuverability and shortening the range. In addition, this solution is not completely universal, because this system needs to adopt different configurations on different types of guided aircraft according to the launch method of the guided aircraft, which undoubtedly pushes up the mass production cost, and the difficulty of using the system in technical positions and vehicles also increases accordingly. Summary of the invention

[0005] The purpose of the present invention is to provide a detachable attitude control device to directly generate a large thrust to act on a guided aircraft in a short time, so that the guided aircraft can achieve a large maneuver turn instantly after launch and quickly enter a kinematic trajectory.

[0006] A detachable attitude control device includes a guided aircraft 15, characterized in that an attitude control device 40 is installed at the front end of the guided aircraft 15, wherein the attitude control device 40 includes a cylindrical shell 2 and an end cover 1, the end cover 1 is sealed and installed at the upper end of the cylindrical shell 2, and a pulse engine assembly, a controller 11, an inertial measurement unit 12 and a power supply unit 13 are installed in the cylindrical shell 2; the controller 11, the inertial measurement unit 12 and the power supply unit 13 are all installed on the inner wall of the cylindrical shell 2; and the pulse engine assembly, the controller 11, the inertial measurement unit 12 and the power supply unit 13 are connected to each other;

[0007] The lower side of the cylindrical shell 2 is open, and the lower end surface of the cylindrical shell 2 is connected to the end of the guided aircraft 15. There is a space in the lower part of the cylindrical shell 2, and the end of the guided aircraft 15 is inserted from the lower end surface of the cylindrical shell 2, so that a part of the end of the guided aircraft 15 extends into the cylindrical shell 2, and the lower end of the cylindrical shell 2 is sleeved on the upper part of the guided aircraft 15, and they are sealed and installed together by shear screws or explosive screws 5;

[0008] The pulse engine assembly comprises at least a pair of yaw control engines 6, a pair of pitch control engines 7, four roll control engines 8 and two separation control engines 9; the pair of yaw control engines 6 are symmetrically arranged on the upper inner wall of the cylindrical shell 2, and the nozzles of the two yaw control engines 6 face oppositely, both facing the outside of the cylindrical shell 2 and perpendicular to the longitudinal axis of the guided aircraft 15; the two pitch control engines 7 are symmetrically arranged on the upper inner wall of the cylindrical shell 2, and the nozzles of the two pitch control engines 7 face oppositely, both facing the outside of the cylindrical shell 2; The two yaw control engines 6 and the two pitch control engines 7 are arranged in a cross shape when viewed from the longitudinal axis of the guided vehicle 15; the nozzles of the four roll control engines 8 are facing opposite directions, all facing the outside of the cylindrical shell 2, in a straight line, and distributed on the inner walls of the lower part of the cylindrical shell 2, and the two straight lines are parallel to each other; the nozzle of the separation control engine 9 is facing an acute angle with the longitudinal axis of the guided vehicle 15; the cylindrical shell 2 corresponding to all the engine nozzles of the pulse engine assembly has corresponding holes for the airflow of each engine nozzle of the pulse engine assembly to be ejected;

[0009] When in use, the guided aircraft 15 and the attitude control device 40 connected thereto are placed in the launch tube 14. The attitude control device 40 is at the exit of the launch tube 14. The upper end surface of the launch tube 14 and the lower surface of the upper edge of the attitude control device 40 are sealed with sealing material 3. The attitude control device 40 is connected to the inner wall of the launch tube 14 with a shear tape 4.

[0010] The end cover 1 is both the end cover of the attitude control device 40 and the launch tube cover.

[0011] A plurality of windows 10 are provided on the wall of the cylindrical housing 2 beside the controller 11, the inertial measurement unit 12 or the power supply unit 13. The plurality of windows 10 are all opened and closed by small door panels to facilitate ground debugging and maintenance.

[0012] The inertial measurement unit 12 measures the three-axis attitude angular rate and acceleration of the guided aircraft 15 in real time, and transmits the real-time measured data to the controller 11 for resolution.

[0013] The controller 11 is used to receive the attitude parameters of the guided aircraft and the given target position output by the inertial measurement unit 12, and according to the requirements of turning attitude control and separation, the received data is calculated and processed in real time to form control instructions, which are transmitted to the actuator pulse engine assembly to control the guided aircraft to complete single or combined actions of rolling, pitching, and yaw, and then realize autonomous separation of the device and the guided aircraft.

[0014] The data communication between the controller 11 and the ground is realized by WIFI wireless mode.

[0015] The power supply unit 13 supplies power to the controller 11 and can be powered by a lithium battery, a chemical battery, or a super capacitor.

[0016] The sealing material 3 is a double-pass sealing ring.

[0017] The four roll control engines 8 are multi-pulse solid engines, which can be started once or multiple times according to the actual combat situation to provide multi-pulse thrust and achieve precise attitude adjustment of the guided aircraft 15.

[0018] Beneficial effects of the present invention

[0019] 1. The attitude control device 40 of the present invention is installed at the front end of the guided aircraft 15. During the launch of the guided aircraft 15, it can change the direction of the guided aircraft 15 by a large angle. The speed of changing the direction of the guided aircraft 15 is unattainable by previous attitude control devices.

[0020] Second, the attitude control device 40 of the present invention is equipped with multiple pulse engines to form a pulse engine assembly, which includes at least a pair of yaw control engines 6, a pair of pitch control engines 7, four roll control engines 8 and two separation control engines 9; the present invention does not have the necessary components such as rudders used in conventional attitude control devices to change the direction of the guided aircraft. The present invention only uses pulse engines to change the direction of the guided aircraft.

[0021] Third, the attitude control device installed on the upper end of the guided aircraft of the present invention shares an end cover 1 with the launch tube 14, so that the launch tube 14 does not need to be opened, so that the launch of the guided aircraft is not affected by the launch tube cover, which is more efficient than the existing guided aircraft launch. Another advantage of sharing an end cover 1 is that the weight of the launch tube is reduced.

[0022] 4. In the present invention, the attitude control device is autonomously and controllably separated from the head of the guided aircraft after completing the rapid turn of the guided aircraft, thereby ensuring the safety of the launch platform, and discarding the passive weight reduces the weight and volume of the entire guided aircraft. The acceleration performance of the booster engine device (or main engine) of the guided aircraft in the air acceleration stage is greatly enhanced after ignition, which comprehensively improves the maneuverability and range of the guided aircraft, thereby improving the strike or defense capability of the guided aircraft.

[0023] 5. The present invention provides a complete set of ideas for the roll-pitch-yaw control and separation control of the guided aircraft in the ascent phase of launching the guided aircraft out of the tube, which can completely replace the gas generator + air rudder control method, effectively improve the harsh working environment of the air rudder, and after removing the gas generator and its accessories, the design complexity of the air rudder is reduced (for example, the internal flow channel of the rudder is removed, the thickness of the rudder root is thinned, etc.), and at the same time avoids the problem of large thrust loss of the gas generator + air rudder method; and according to the characteristics of different guided aircraft, functional modules are selectively installed to achieve specific functions (for example, only choose to install the pulse engine for separation).

[0024] 6. The attitude control device of the present invention adopts an integrated independent disposable module design and is installed externally on the head of the guided aircraft, which greatly reduces the complexity of the interface design of the guided aircraft. The present invention has a simple structure, is easy to install, and has wide adaptability.

[0025] 7. The purpose of the present invention is to create a detachable general attitude control device for guided aircraft turning, so that it can meet the different requirements of launching guided aircraft in vertical or inclined modes in most modern roadbeds, surface ships, underwater submarines and other different vehicles, and realize all-round rapid turning of the guided aircraft in the target direction after launch.

[0026] 8. The present invention can replace the traditional control method of gas generator + air rudder, and solve the problems brought by the traditional control method, such as complex structure, harsh working environment of air rudder, high complexity of rudder design, large thrust loss, high cost, etc.

[0027] 9. The attitude control device of the present invention is based on the direct force control method of the pulse engine, which directly generates a large thrust to act on the guided aircraft in a short time, so that the guided aircraft can achieve a large maneuver turn instantly after launch and quickly enter the kinematic trajectory. After completing the turn, it can autonomously separate from the guided aircraft, reducing the flight weight of the guided aircraft, thereby achieving the purpose of improving the maneuverability (rapid response) effectiveness and survivability of the guided aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0028] Figure 1 , is a schematic structural diagram of the attitude control device 40 of the present invention;

[0029] Figure 2 , is a schematic diagram of the positional relationship between the attitude control device 40 and the guided aircraft 15 during the launch process of the present invention; wherein

[0030] (a) is a schematic diagram of a state in which the attitude control device 40 of the present invention is connected to the guided aircraft 15 and installed in a launch tube;

[0031] (b) is a schematic diagram of the attitude control device 40 of the present invention being connected to the guided aircraft 15 and being launched from the launch tube;

[0032] (c) is a schematic diagram of the attitude control device 40 of the present invention and the guided aircraft 15 being separated from each other after being launched out of the tube;

[0033] Figure 3 , a schematic diagram of the distribution structure of the pulse engine assembly of the present invention in the attitude control device 40; wherein,

[0034] (a) Figure 3 AA section view of (b);

[0035] (b) is a cross-sectional view of the distribution structure of the pulse engine assembly of the present invention in the attitude control device 40;

[0036] (c) Figure 3 (b) BB cross-sectional view;

[0037] Figure 4 , a schematic diagram showing the connection relationship between the controller 11, the inertial measurement unit 12 and the power supply unit 13 of the present invention on the inner wall of the cylindrical housing 2;

[0038] Figure 5 , structural principle diagram of the attitude control device of the present invention.

[0039] Among them, 1 is the end cover, 2 is the cylindrical shell, 3 is the sealing material, 4 is the shear band, 5 is the explosive screw, 6 is the yaw control engine, 7 is the pitch control engine, 8 is the roll control engine, 9 is the separation control engine, 10 is the window, 11 is the controller, 12 is the inertial measurement unit, 13 is the power supply unit, 14 is the launch tube, 15 is the guided vehicle, and 40 is the attitude control device. DETAILED DESCRIPTION

[0040] A detachable attitude control device includes a guided aircraft 15, characterized in that an attitude control device 40 is installed at the front end of the guided aircraft 15, wherein the attitude control device 40 includes a cylindrical shell 2 and an end cover 1, the end cover 1 is sealed and installed at the upper end of the cylindrical shell 2, and a pulse engine assembly, a controller 11, an inertial measurement unit 12 and a power supply unit 13 are installed in the cylindrical shell 2; the controller 11, the inertial measurement unit 12 and the power supply unit 13 are all installed on the inner wall of the cylindrical shell 2; and the pulse engine assembly, the controller 11, the inertial measurement unit 12 and the power supply unit 13 are connected to each other;

[0041] The lower side of the cylindrical shell 2 is open, and the lower end surface of the cylindrical shell 2 is connected to the end of the guided aircraft 15. There is a space in the lower part of the cylindrical shell 2, and the end of the guided aircraft 15 is inserted from the lower end surface of the cylindrical shell 2, so that a part of the end of the guided aircraft 15 extends into the cylindrical shell 2, and the lower end of the cylindrical shell 2 is sleeved on the upper part of the guided aircraft 15, and they are sealed and installed together by shear screws or explosive screws 5;

[0042] The pulse engine assembly comprises at least a pair of yaw control engines 6, a pair of pitch control engines 7, four roll control engines 8 and two separation control engines 9; the pair of yaw control engines 6 are symmetrically arranged on the upper inner wall of the cylindrical shell 2, and the nozzles of the two yaw control engines 6 face oppositely, both facing the outside of the cylindrical shell 2 and perpendicular to the longitudinal axis of the guided aircraft 15; the two pitch control engines 7 are symmetrically arranged on the upper inner wall of the cylindrical shell 2, and the nozzles of the two pitch control engines 7 face oppositely, both facing the outside of the cylindrical shell 2; The two yaw control engines 6 and the two pitch control engines 7 are arranged in a cross shape when viewed from the longitudinal axis of the guided vehicle 15; the nozzles of the four roll control engines 8 are facing opposite directions, all facing the outside of the cylindrical shell 2, in a straight line, and distributed on the inner walls of the lower part of the cylindrical shell 2, and the two straight lines are parallel to each other; the nozzle of the separation control engine 9 is facing an acute angle with the longitudinal axis of the guided vehicle 15; the cylindrical shell 2 corresponding to all the engine nozzles of the pulse engine assembly has corresponding holes for the airflow of each engine nozzle of the pulse engine assembly to be ejected;

[0043] When in use, the guided aircraft 15 and the attitude control device 40 connected thereto are placed in the launch tube 14. The attitude control device 40 is at the exit of the launch tube 14. The upper end surface of the launch tube 14 and the lower surface of the upper edge of the attitude control device 40 are sealed with sealing material 3. The attitude control device 40 is connected to the inner wall of the launch tube 14 with a shear tape 4.

[0044] The end cover 1 is both the end cover of the attitude control device 40 and the launch tube cover.

[0045] A plurality of windows 10 are provided on the wall of the cylindrical housing 2 beside the controller 11, the inertial measurement unit 12 or the power supply unit 13. The plurality of windows 10 are all opened and closed by small door panels to facilitate ground debugging and maintenance.

[0046] The inertial measurement unit 12 measures the three-axis attitude angular rate and acceleration of the guided aircraft 15 in real time, and transmits the real-time measured data to the controller 11 for resolution.

[0047] The controller 11 is used to receive the attitude parameters of the guided aircraft and the given target position output by the inertial measurement unit 12, and according to the requirements of turning attitude control and separation, the received data is calculated and processed in real time to form control instructions, which are transmitted to the actuator pulse engine assembly to control the guided aircraft to complete single or combined actions of rolling, pitching, and yaw, and then realize autonomous separation of the device and the guided aircraft.

[0048] The data communication between the controller 11 and the ground is realized by WIFI wireless mode.

[0049] The power supply unit 13 supplies power to the controller 11 and can be powered by a lithium battery, a chemical battery, or a super capacitor.

[0050] The sealing material 3 is a double-pass sealing ring.

[0051] The four roll control engines 8 are multi-pulse solid engines, which can be started once or multiple times according to the actual combat situation to provide multi-pulse thrust and achieve precise attitude adjustment of the guided aircraft 15.

[0052] Example:

[0053] The attitude control device of the present invention is based on the direct force control method of the pulse engine, which directly generates a large thrust to act on the guided aircraft in a short time, so that the guided aircraft can achieve a large maneuver turn instantly after launch and quickly enter the kinematic trajectory. After completing the turn, it can autonomously separate from the guided aircraft, reducing the flight weight of the guided aircraft, thereby achieving the purpose of improving the maneuverability (rapid response) effectiveness and survivability of the guided aircraft.

[0054] The attitude control device is mainly composed of a set of pulse engine components, controller, inertial measurement unit, power supply unit, end cover and cylindrical shell, etc. Figure 1As shown, the main structure consists of an end cover 1 and a cylindrical shell 2. The cylindrical shell 2 can be formed by integrated processing or split processing and welding. The end cover 1 can be formed by integrated processing, and each component is installed on the cylindrical shell 2. The device is an integrated independent module design, which is installed on the front of the guided aircraft 15 by shear screws or explosive screws 5, and can be separated from the guided aircraft independently. The device is sealed between the launch tube 14 by the sealing material 3 and connected to the launch tube 14 by the shear band 4, as shown in FIG. Figure 2 shown.

[0055] The pulse engine assembly is mainly composed of an engine and an igniter, and is functionally independent as an actuator. A layout form of the engine is as follows Figure 3 As shown (the engine layout is not limited to this form). A pair of engines 6 responsible for yaw control, a pair of engines 7 for pitch control, four engines 8 for roll control, and a pair of engines 9 for separation control. Engine 6 is arranged symmetrically, with nozzles facing opposite directions and perpendicular to the longitudinal axis of the guided aircraft 8; the arrangement method and nozzle direction of engine 7 are similar to engine 6, and engines 6 and 7 are arranged in a cross shape when viewed from the longitudinal axis of the guided aircraft; the nozzle direction of engine 8 is tangent to the shell 2, and the nozzles of two engines are facing opposite directions to the nozzles of the other two engines; the nozzle direction of engine 9 is at an acute angle to the longitudinal axis of the guided aircraft 8. Among them, the four engines 8 responsible for roll control can use dual-pulse solid engines for secondary precise attitude adjustment. All pulse solid engines are installed on the wall of the cylindrical shell 2 through flanges at the nozzle position, and holes are opened at the corresponding cylinder wall to allow airflow to spray out.

[0056] The controller 11 is used to receive the attitude parameters of the guided aircraft output by the inertial measurement unit 12 and the given target position, and according to the requirements of turning attitude control, separation control, etc., the received data is calculated and processed in real time and a control instruction is formed, which is transmitted to the actuator-pulse engine to control the guided aircraft to complete the single or combined action of roll-pitch-yaw to realize the autonomous separation of the device and the guided aircraft. The controller 11 mainly includes hardware and software. The hardware provides a physical platform and foundation for the operation control of the software and the transmission control of the interface information, and the software realizes the turning program control, the separation program control, etc. The data communication between the controller 11 and the ground is realized by WIFI wireless mode.

[0057] The inertial measurement unit 12 is used to measure the three-axis attitude angle (or angular velocity) and acceleration of the guided aircraft in real time, and transmit the parameters to the controller 11 for solution.

[0058] The power supply unit 13 supplies power to the controller 11 , and can be powered by lithium batteries, chemical batteries, super capacitors, etc.

[0059] The controller 11, the inertial measurement unit 12 and the power supply unit 13 can be installed on the side wall of the cylindrical shell 2, and a window is provided on the shell, and a small door plate 10 is used to open and close the window to protect the window and facilitate ground debugging and maintenance. Figure 4 As shown. For a small guided aircraft, the installation space of the cylindrical shell 2 may be relatively small. In this case, the above components can be installed on the upper part of the cylindrical shell 2, and ground debugging and maintenance can be carried out by removing the end cover 1. The inertial measurement unit 12 and the power supply unit 13 can also be integrated into the controller 11 according to the size of the guided aircraft.

[0060] The end cover 1 is conformally designed with the launch tube cover, and is both an attitude control device end cover and a launch tube cover, thereby reducing the weight of the launch tube.

[0061] The housing 2 is also provided with a sealing material 3 for sealing the annular gap between the guided aircraft and the launch tube from the front end of the transport launch tube. The sealing material can be in the form of a double-pass sealing ring.

[0062] All components, single machines and structural shells are integrated into an integrated independent cabin module, which can be conveniently installed externally with different guided aircraft through shear screws or explosive screws. It has a simple interface, easy installation and wide adaptability.

[0063] The internal connection relationship of the attitude control device system is as follows Figure 5 As shown. The controller, as a central processor, is connected to the power supply unit, inertial measurement unit, and igniter respectively. The power supply unit supplies power to the controller, and the inertial measurement unit transmits the attitude parameters of the guided aircraft measured in real time to the controller. The controller performs calculation and comparison based on the given target position parameters and the attitude parameters of the guided aircraft collected in real time, and forms a control instruction to be transmitted to the engine igniter, starting the engine to control the guided aircraft to complete the program turn.

[0064] Directions:

[0065] After the guided aircraft leaves the water or tube and is recorded by the sensor, the attitude control device is activated to perform rolling, pitching or yaw and other actions to achieve a rapid turn toward the target. For example, the guided aircraft is first rolled, then pitched, and finally separated. The controller 11 checks the real-time attitude of the guided aircraft according to the estimated trajectory, issues a rolling command to start a pair of rolling engines 8, and generates a positive torque around the longitudinal axis of the guided aircraft under the action of the engine thrust. The angular velocity of the guided aircraft in the rolling channel increases and begins to accelerate. The inertial measurement unit 12 transmits the attitude parameters of the guided aircraft measured in real time to the controller 11, and the controller 11 calculates the attitude parameters of the guided aircraft according to the given target parameters and the attitude parameters of the guided aircraft measured by the inertial measurement unit. The number is used to calculate the start time of subsequent instructions. When the guided aircraft benchmark is about to align with the target shooting surface, an instruction is given to start another pair of rolling engines 8 to spray in the opposite direction to produce negative torque, suppress the rolling angular velocity, and make the guided aircraft quickly stable. Then a pitch engine 7 is started to produce a "head-down" torque under the engine thrust, and the angular velocity of the guided aircraft in the pitch channel is increased. The controller solves in real time and gives an instruction to start another pitch engine 7 to produce a "stop" torque to calm the pitch angular velocity when the guided aircraft turns into the required attitude relative to the horizon. Similarly, the guided aircraft also adopts a similar control method in the yaw channel. After the turn is completed, the controller 11 sends an instruction to start the separation engine 9, so that the attitude control device outputs impulse along the flight direction of the guided aircraft. The attitude control device separates from the front of the guided aircraft and falls in front of the launch platform along the parabolic trajectory. Then the guided aircraft starts its own engine and begins to accelerate sharply. When the flight speed is large enough, the attitude control of the guided aircraft is completed by the air rudder.

[0066] The controller 11 and the inertial measurement unit 12 mentioned above can be optionally installed according to the specific needs of the guided aircraft. For example, the control and measurement functions can be realized by the control system and inertial navigation of the guided aircraft itself. The pulse engine can also be selectively installed to realize specific functions, such as launching the guided aircraft at a small horizontal angle (10-30 degrees) on the water surface. The guided aircraft does not need an attitude control device to make large azimuth turns, and only chooses to install a pulse engine for separation, which further reduces costs.

[0067] By adopting different action sequences of the device of the present invention and the use method start-up program recommended above, the guided aircraft is given the ability to be launched vertically or obliquely from different vehicles and to turn quickly after launch.

[0068] In summary, the present invention can guide the development of a fully universal guided aircraft that can meet the different requirements of modern conventional sea-based and land-based weapons, be widely adapted to various types of vehicles, and fully reflect the advantages of universal weapons such as simple operation and low use cost.

[0069] Technical problems solved by the present invention:

[0070] 1. The purpose of the present invention is to create a detachable attitude control universal device for guiding aircraft turning, so that it can meet the different requirements of launching guided aircraft in vertical or inclined modes by most modern roadbeds, surface ships, underwater submarines and other different vehicles, and realize the all-round rapid turning of the guided aircraft towards the target direction after launch;

[0071] 2. The present invention can replace the traditional control method of gas generator + air rudder, and solve the problems brought by the traditional control method, such as complex structure, harsh working environment of air rudder, high complexity of rudder design, large thrust loss, high cost, etc.

[0072] 3. The present invention can realize the specific functions (effects) of different types of guided aircraft by only changing the working procedures and operation methods of the components in the system, thus solving the problems of high mass production cost and difficulty in using the system in technical positions and vehicles;

[0073] 4. Solve the problem of safe and reliable separation of the device from the guided aircraft after use, reduce the weight and volume of the guided aircraft, and improve the maneuverability and range of the guided aircraft;

[0074] 5. The device of the present invention adopts an integrated independent disposable module design and is installed externally with the guided aircraft, which solves the problems of complex design interface, low installation efficiency and low adaptability of the guided aircraft;

[0075] The key points of the present invention are:

[0076] 1. The working sequence, ignition timing and other principle and method innovations of the device of the present invention can guide the development of a fully universal guided aircraft, which can meet the different requirements of launching guided aircraft in vertical or inclined modes from most modern roadbeds, surface ships, underwater submarines and other different vehicles, and realize the rapid turning of guided aircraft;

[0077] 2. The device of the present invention adopts an integrated independent disposable module design and is installed externally with the guided aircraft, which greatly reduces the complexity of the interface design of the guided aircraft, has a simple structure, is easy to install, and has wide adaptability;

[0078] 3. The present invention can realize specific functions and effects of different types of guided aircraft by selectively installing functional units and changing the working procedures and operation methods of components according to different launch modes and different types of guided aircraft, thereby greatly reducing the mass production cost and the difficulty of using the system in technical positions and vehicles;

[0079] 4. The present invention has the ability of autonomous and controllable separation, and can separate from the guided aircraft after the mission is completed, thereby reducing the weight and volume of the entire guided aircraft while ensuring the safety of the launch platform. The acceleration performance of the booster engine (or main engine) of the guided aircraft in the air acceleration stage is greatly enhanced after ignition, which comprehensively improves the maneuverability and range of the guided aircraft, thereby improving the strike and defense capabilities of the guided aircraft;

[0080] 5. The present invention provides a complete set of ideas for the roll-pitch-yaw control and separation control of the guided aircraft launching out of the tube ascent phase, which can completely replace the gas generator + air rudder control mode, effectively improve the harsh working environment of the air rudder, and after removing the gas generator and its accessories, the design complexity of the air rudder is reduced (for example, the internal flow channel of the rudder is removed, the thickness of the rudder root is thinned, etc.), and at the same time avoids the problem of large thrust loss of the gas generator + air rudder mode; and according to the characteristics of different guided aircraft, functional modules are selectively installed to achieve specific functions, such as only installing the pulse engine for separation.

[0081] The benefits brought by the technical solution of the present invention are:

[0082] 1. The device of the present invention can guide the development of a fully universal guided aircraft, which can meet the different requirements of launching guided aircraft in vertical or inclined modes by most modern roadbeds, surface ships, underwater submarines and other different vehicles, and realize rapid turning of guided aircraft.

[0083] 2. The device of the present invention adopts an integrated independent disposable module design and is installed externally on the head of the guided aircraft, which greatly reduces the complexity of the design interface of the guided aircraft. It has a simple structure, is easy to install, and has wide adaptability.

[0084] 3. For different launch modes and different types of guided aircraft, the device of the present invention can realize the specific functions and effects of different types of guided aircraft by selecting or changing the working procedures and operating methods of the components in the system, which greatly reduces the mass production cost and the difficulty of using the system in technical positions and vehicles.

[0085] 4. In the present invention, the attitude control device is autonomously and controllably separated from the head of the guided aircraft after completing the rapid turn of the guided aircraft, thereby ensuring the safety of the launch platform, and discarding the passive weight reduces the weight and volume of the entire guided aircraft. The acceleration performance of the booster engine device or the main engine of the guided aircraft in the air acceleration stage is greatly enhanced after ignition, which comprehensively improves the maneuverability and range of the guided aircraft, thereby improving the strike or defense capability of the guided aircraft.

[0086] 5. The present invention provides a complete set of ideas for the roll-pitch-yaw control and separation control of the guided aircraft launching out of the tube ascent phase, which can completely replace the gas generator + air rudder control mode, effectively improve the harsh working environment of the air rudder, and after removing the gas generator and its accessories, the design complexity of the air rudder is reduced (for example, the internal flow channel of the rudder is removed, the thickness of the rudder root is thinned, etc.), and at the same time avoids the problem of large thrust loss of the gas generator + air rudder mode; and according to the characteristics of different guided aircraft, functional modules are selectively installed to achieve specific functions, such as only installing the pulse engine for separation.

Claims

1. A detachable attitude control device, comprising a guided aircraft (15), characterized in that: An attitude control device (40) is installed at the front end of the guided aircraft (15), wherein the attitude control device (40) comprises a cylindrical shell (2) and an end cover (1), the end cover (1) being sealed and installed at the upper end of the cylindrical shell (2), and a pulse engine assembly, a controller (11), an inertial measurement unit (12) and a power supply unit (13) being installed in the cylindrical shell (2); the controller (11), the inertial measurement unit (12) and the power supply unit (13) are all installed on the inner wall of the cylindrical shell (2); and the pulse engine assembly, the controller (11), the inertial measurement unit (12) and the power supply unit (13) are interconnected; The lower side of the cylindrical shell (2) is open, the lower end surface of the cylindrical shell (2) is connected to the end of the guided aircraft (15), the lower part of the cylindrical shell (2) has a space, the end of the guided aircraft (15) is inserted from the lower end surface of the cylindrical shell (2), so that a part of the end of the guided aircraft (15) extends into the cylindrical shell (2), the lower end of the cylindrical shell (2) is sleeved on the upper part of the guided aircraft (15), and they are sealed and installed together by shear screws or explosive screws (5); The pulse engine assembly comprises at least a pair of yaw control engines (6), a pair of pitch control engines (7), four roll control engines (8) and two separation control engines (9); the pair of yaw control engines (6) are symmetrically arranged on the upper inner wall of the cylindrical shell (2), the nozzles of the two yaw control engines (6) are oriented in opposite directions, both facing the outside of the cylindrical shell (2) and perpendicular to the longitudinal axis of the guided aircraft (15); the two pitch control engines (7) are symmetrically arranged on the upper inner wall of the cylindrical shell (2), the nozzles of the two pitch control engines (7) are oriented in opposite directions, both facing the cylindrical shell (2). outside; when viewed from the longitudinal axis direction of the guided aircraft (15), the two yaw control engines (6) and the two pitch control engines (7) are arranged in a cross shape; the nozzles of the four roll control engines (8) are oriented in opposite directions, all facing the outside of the cylindrical shell (2), in a straight line, and distributed on the inner walls of the lower part of the cylindrical shell (2) on both sides, and the two straight lines are parallel to each other; the nozzle of the separation control engine (9) is oriented at an acute angle to the longitudinal axis of the guided aircraft (15); the cylindrical shell (2) corresponding to all the engine nozzles of the pulse engine assembly has corresponding holes for the airflow of each engine nozzle of the pulse engine assembly to be ejected; When in use, the guided aircraft (15) and the attitude control device (40) connected thereto are placed in a launch tube (14), the attitude control device (40) is at the exit of the launch tube (14), the upper end surface of the launch tube (14) and the lower surface of the upper edge of the attitude control device (40) are sealed with a sealing material (3); the attitude control device (40) is connected to the inner wall of the launch tube (14) with a shear band (4); The end cover (1) is both an end cover of the attitude control device (40) and a launch tube cover; A plurality of windows (10) are provided on the wall of the cylindrical housing (2) next to the controller (11), the inertial measurement unit (12) and the power supply unit (13). The plurality of windows (10) are all opened and closed by small door panels to facilitate ground debugging and maintenance.

2. A detachable posture control device according to claim 1, characterized in that: The inertial measurement unit (12) measures the three-axis attitude angular rate and acceleration of the guided aircraft (15) in real time, and transmits the real-time measured data to the controller (11) for resolution.

3. A detachable posture control device according to claim 2, characterized in that: The controller (11) is used to receive the attitude parameters and given target position of the guided aircraft (15) output by the inertial measurement unit (12), and to perform real-time calculation and processing on the received data according to the requirements of turning attitude control and separation, and to form control instructions, which are transmitted to the actuator pulse engine assembly to control the guided aircraft (15) to complete single or combined actions of rolling, pitching, and yaw, thereby realizing autonomous separation of the device from the guided aircraft.

4. The detachable posture control device according to claim 2, characterized in that: The data communication between the controller (11) and the ground is achieved by wireless WIFI.

5. The detachable posture control device according to claim 1, characterized in that: The power supply unit (13) supplies power to the controller (11) by using a lithium battery, a chemical battery or a super capacitor.

6. The detachable posture control device according to claim 1, characterized in that: The sealing material (3) is a double-pass sealing ring.

7. The detachable posture control device according to claim 1, characterized in that: The four roll control engines (8) are multi-pulse solid engines, which are started once or multiple times according to actual situations to provide multi-pulse thrust and achieve precise attitude adjustment of the guided aircraft.

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