Solid rocket engine for sounding rocket based on thrust vector system control

Abstract

The invention relates to the technical field of solid rocket engines, in particular to a solid rocket engine for a sounding rocket based on thrust vector system control, which comprises a combustion chamber, a spray pipe, a heat insulation layer, an arc end face, a control support, a combustion chamber cover plate, a fixed support, a base plate, a fixed rod, a steering engine, a push rod, an attitude sensor, a single chip microcomputer and an upper computer, the upper and lower base plates and the combustion chamber cover plate are connected with the combustion chamber through threads; the heat insulation layer is limited and positioned through a limiting groove in the combustion chamber; the combustion chamber cover plate is in limited connection with the combustion chamber through threads; the push rod is matched with the steering engine and the control support through pins; the push rod can be intelligently controlled by a single-chip microcomputer control steering engine. According to the solid rocket, on one hand, the flight stability of the rocket in the task execution process can be improved by adjusting the attitude of the solid rocket, on the other hand, the solid rocket can be used repeatedly, and the product cost is reduced.

Description

Technical Field

[0001] The present invention relates to the technical field of solid rocket engines for sounding rockets, and particularly to a solid rocket engine controlled by a thrust vector system and with low cost. Technical Background

[0002] At present, solid rocket engines have the advantages of simple structure, high reliability, strong maneuverability, and low requirements for processing and materials. They have been widely used in various rockets, tactical missiles, and boosters of other various aircraft.

[0003] At present, in order to reduce the launch cost in China, most sounding rockets are uncontrolled rockets, and only rely on stable fins or the rotation of the rocket around the longitudinal axis to ensure flight stability. As a result, the flight trajectory of the sounding rocket is greatly affected by the wind, and after the traditional solid rocket engine is used once, the inner wall of the combustion chamber and the nozzle will be corroded on a large area, resulting in the inability to continuously carry out multiple launch missions. However, with the continuous advancement of China's space industry, sounding rockets need to reach designated heights and positions to deploy precision instruments. Therefore, how to achieve the reusable use of sounding rockets while controlling the sounding rocket to fly along a predetermined trajectory and reach the designated position is a difficult problem that needs to be solved urgently at present. Summary of the Invention

[0004] The purpose of the present invention is to provide a solid rocket engine for sounding rockets controlled by a thrust vector system in view of the above-mentioned deficiencies of the prior art.

[0005] The present invention adopts the following technical solutions to achieve the invention purpose:

[0006] A solid rocket engine for sounding rockets controlled by a thrust vector system includes a combustion chamber, a nozzle, an insulating layer, an arc end face, a control support, a combustion chamber cover plate, a fixed support, a backing plate, a fixed rod, a servo, a push rod, an attitude sensor, a single-chip microcomputer, a host computer, and a client.

[0007] The insulating layer is fixedly connected through a limit groove in the combustion chamber. The insulating layer uses high-temperature resistant graphite material, and the combustion chamber uses GH4169 high-temperature resistant alloy.

[0008] The grain is located in the combustion chamber, and an insulating layer is coated on its outer side, and the grain is loaded from the upper part of the combustion chamber.

[0009] The inner diameter of the upper end opening of the combustion chamber and the inner diameter of the insulating layer are both larger than the outer diameter of the grain, and the outer diameter of the grain is equal to the inner diameter of the upper part of the nozzle.

[0010] The nozzle is placed at the bottom of the combustion chamber and is fixedly connected to the bottom of the combustion chamber through a hexagonal nut. The nozzle uses high-temperature resistant graphite material.

[0011] The arc end face is placed under the nozzle and fixedly connected to the lower backing plate and the fixed support through a fixing rod, which is used to adjust the thrust direction. The adjustable angle is 0° - 15°. The arc end face is made of high-temperature-resistant graphite material, and the lower backing plate and the fixed support are made of ordinary aluminum alloy material.

[0012] The control support is fixedly connected to the arc end face through threads and fixedly connected to the push rod through a pin. The control support is made of GH4169 high-temperature-resistant alloy.

[0013] The lower part of the combustion chamber cover plate is fixedly connected to the combustion chamber through threads, and the upper part is fixedly connected to the fixed support through a fixing rod. The combustion chamber cover plate is made of high-temperature-resistant graphite material.

[0014] The servo motor is fixedly connected to the push rod through a pin, and the servo motor is limited and fixedly connected to the central support through a hexagon nut. The number of servo motors and push rods is three. If the thrust needs to be enhanced, the number of servo motors and push rods can be increased appropriately. The push rod is made of ordinary aluminum alloy material.

[0015] The attitude sensor is communicatively connected to the single-chip microcomputer. The accelerometer of the attitude sensor measures the acceleration of the rocket on three axes, and the magnetometer measures the direction and intensity of the earth's magnetic field, calculates the attitude information such as the tilt angle and pitch angle of the rocket, and sends the rocket attitude information to the single-chip microcomputer. The attitude sensor uses MPU6050, and the single-chip microcomputer uses STM32.

[0016] The single-chip microcomputer is communicatively connected to the servo motor. After receiving the information from the attitude sensor, the single-chip microcomputer analyzes and processes it, outputs a PWM control signal to the servo motor through the PID algorithm, and the servo motor changes the fire direction of the arc end face by controlling the push rod.

[0017] Beneficial effects:

[0018] Compared with the prior art, the beneficial effects of the present invention are as follows:

[0019] (1) Through the thrust vector control technology, the present invention can adjust the rocket attitude in real time, control the rocket to fly according to the set ballistic trajectory, and thus reach the predetermined position.

[0020] (2) The present invention adopts a modular design instead of an integral molding, which can reduce the processing difficulty, save the manufacturing cost, and can directly remove and replace some damaged parts without the need for overall replacement.

[0021] (3) For parts such as the thermal insulation layer, nozzle, and combustion cover of the present invention, high-temperature-resistant graphite thermal insulation materials are used, and for parts such as the combustion chamber and control support, GH4169 high-temperature-resistant alloy is used. After one launch, the corrosion area of the inner wall of the combustion chamber, the control support, and the nozzle can be greatly reduced, and the launch cost can be reduced.

[0022] (4) The present invention can transmit the rocket attitude information and detection information to the client in real time through the host computer, enabling the user to monitor the rocket flight state in real time and receive the detection information. Description of the Drawings

[0023] Figure 1 It is a schematic diagram of the overall structure of the present invention

[0024] Figure 2 It is an enlarged schematic diagram of the control part structure of the present invention

[0025] Figure 3 It is a schematic sectional view of the present invention

[0026] Figure 4 It is a flowchart of thrust vector control of the present invention

[0027] Explanation of component numbers: 1 - Combustion chamber, 2 - Nozzle, 3 - Insulation layer, 4 - Arc end face, 5 - Control support, 6 - Combustion chamber cover plate, 7 - Fixed support, 8 - Pad, 9 - Fixed rod, 10 - Servo, 11 - Push rod, 12 - Attitude sensor, 13 - Single-chip microcomputer, 14 - Host computer, 15 - Client Detailed Embodiment

[0028] Next, in combination with the drawings of the present invention, the present invention will be further explained through specific embodiments, but it does not constitute any limitation to the present invention.

[0029] Embodiment

[0030] Referring to the drawings, a solid rocket engine for a sounding rocket based on thrust vector system control includes: a combustion chamber (1), a nozzle (2), an insulation layer (3), an arc end face (4), a control support (5), a combustion chamber cover plate (6), a fixed support (7), a pad (8), a fixed rod (9), a servo (10), a push rod (11), an attitude sensor (12), a single-chip microcomputer (13), a host computer (14), and a client (15).

[0031] Specifically, when assembling the solid rocket engine, first place the nozzle (2) at the lower outlet of the combustion chamber (1), then place the insulation layer (3) at the limit groove inside the combustion chamber (1), place the propellant grain inside the insulation layer (3), fixedly connect the combustion chamber cover plate (6) to the combustion chamber (1) by threads, then fixedly connect the arc end face (4) to the lower outlet of the combustion chamber (1) through the lower pad (8), then use the fixed rod (9) to fixedly connect the lower pad (8) to the fixed support (7), then fixedly connect the control support (5) to the arc end face (4) by threads, and then pass the push rod (11) through the fixed support (7) and fixedly connect it to the control support (5) by a pin.

[0032] Specifically, the combustion chamber (1) is made of GH4169 high-temperature resistant alloy with a thickness of 5 - 8 mm. After testing, when the combustion chamber (1) is in a high-temperature environment of 1000 °C and under an internal pressure of 8 MPA, it will not undergo fission.

[0033] Specifically, the thermal insulation layer (3), the nozzle (2), and the combustion chamber cover plate (6) are made of high-temperature resistant graphite material. After testing, when the thermal insulation layer (3) is in an environment of 1900 °C and an internal pressure of 6 - 9 MPA, it will not undergo fission and has good heat insulation effect. The nozzle (2) and the combustion chamber cover plate (6) can still maintain their original shape at a high temperature of 1900 °C and will not deform.

[0034] Specifically, the working steps of the thrust vector control system are as follows:

[0035] Step S1: The accelerometer of the attitude sensor (12) measures the acceleration of the rocket on three axes, and the magnetometer measures the direction and intensity of the earth's magnetic field, and calculates attitude information such as the tilt angle and pitch angle of the rocket.

[0036] Step S2: The attitude sensor (12) sends the rocket attitude information to the single-chip microcomputer (13), and the single-chip microcomputer (13) analyzes and processes the data, and outputs a PWM control signal to the servo (10) through the PID algorithm.

[0037] Step S3: The servo (10) receives the PWM control signal, drives the push rod (11) of the rear-end actuator to move, and adjusts the flight attitude of the rocket.

[0038] Step S4: The attitude sensor (12) sends the rocket attitude information to the upper computer (14) in real time, and the upper computer (14) displays information such as the position, attitude, and flight altitude of the rocket in real time.

[0039] Specifically, a solid rocket engine for a sounding rocket based on thrust vector system control has an operating time of 4 s ± 0.5 s. The ground test runs are divided into three groups. The first group: The thermal insulation layer (3) is made of high-temperature resistant graphite material, and the combustion chamber (1) and the nozzle (2) are made of aluminum alloy material, without thrust vector control; the second group: Without a thermal insulation layer, the combustion chamber (1) is made of GH4169 high-temperature resistant alloy, and the nozzle (6) is made of high-temperature resistant graphite material, without thrust vector control; the third group: The thermal insulation layer (3) and the nozzle (6) are made of high-temperature resistant graphite material, and the combustion chamber (1) is made of GH4169 high-temperature resistant alloy.

[0040] Specifically, during the test run process, the first group of tests was successfully completed, but it was found that 60% of the inner wall area of the combustion chamber (1) was corroded and deformed and could not be reused. The nozzle (6) cracked, resulting in a change in the fuel injection direction, reducing the rocket thrust, and the thrust direction could not be adjusted during the engine operation.

[0041] Specifically, the second group of test runs failed, and after the rocket engine worked for 1.5 seconds, the combustion chamber (1) exploded.

[0042] Specifically, the third group of test runs was successful. It was found that only 15% of the inner wall area of the combustion chamber (1) was corroded and there was no deformation, so it could be reused. The nozzle (6) did not rupture during the operation of the engine, and in the third group of tests, the thrust vector system could quickly adjust the attitude of the engine and was in good working condition.

[0043] Specifically, through analysis, the situation of the first group of tests was determined to be that due to the inability of ordinary aluminum alloy materials to withstand large internal pressures in high-temperature environments, the combustion chamber (1) and the nozzle (6) would have obvious deformations; the situation of the second group of tests was determined to be that due to the large heat flux density of the fuel flowing through the combustion chamber shell during combustion, the temperature of the combustion chamber (1) rose rapidly without adiabatic conditions, resulting in a decrease in the strength of the metal shell and an inability to withstand the internal pressure during operation, leading to the explosion of the combustion chamber (1).

[0044] For those skilled in the art, it is obvious that the present invention is not limited to the details of the above exemplary embodiments, and the present invention can be implemented in other specific forms without departing from the spirit or basic characteristics of the present invention. Therefore, from any point of view, the embodiments should be regarded as exemplary and non-limiting. The scope of the present invention is defined by the appended claims rather than the above description. Therefore, all changes falling within the meaning and scope of the equivalent elements of the claims are intended to be encompassed by the present invention. Any reference signs in the claims should not be regarded as limiting the claims involved.

Claims

1. A solid rocket engine for a sounding rocket based on thrust vector system control, characterized in that: It comprises a combustion chamber (1), a nozzle (2), an insulation layer (3), an arc end surface (4), a control support (5), a combustion chamber cover plate (6), a fixed support (7), a pad (8), a fixed rod (9), a steering gear (10), a push rod (11), a posture sensor (12), a single chip computer (13), a host computer (14), and a client (15); The lower end of the heat insulating layer (3) is limited and positioned by a limiting groove in the combustion chamber (1), and the upper end is limited and positioned by a limiting groove in the combustion chamber cover plate (6); the material used for the combustion chamber (1) is GH4169 high temperature resistant alloy, and the material used for the heat insulating layer (3) is high temperature resistant graphite material; The upper end of the combustion chamber (1) is fixedly connected to the combustion chamber cover plate by means of threads, the lower end of the combustion chamber (1) is threadedly connected to the pad (8), and the lower pad (8) and the fixed support (7) are limitedly connected by means of a fixing rod (9), and the fixing rod (9) is fixedly connected to the lower pad (8) and the fixed support (4) by means of a hexagonal nut.

2. A solid rocket engine for a sounding rocket based on thrust vector system control according to claim 1, characterized in that: Four fixed rods (9) are evenly distributed along the fixed support (7), three push rods (11) are evenly distributed along the fixed support (7), and the push rods (11) are connected to the control support (5) via pins.

3. The solid rocket engine for a sounding rocket based on thrust vector system control according to claim 1, characterized in that: The charge is located in the combustion chamber (1), the outer side of which is covered with a heat insulating layer (3), and the charge is loaded from the upper part of the combustion chamber (1).

4. The solid rocket engine for a sounding rocket based on thrust vector system control according to claim 1, characterized in that: The inner diameter of the upper opening of the combustion chamber (1) and the inner diameter of the heat insulating layer (3) are both larger than the outer diameter of the charge, and the outer diameter of the charge is equal to the inner diameter of the upper portion of the nozzle (2).

5. The solid rocket engine for a sounding rocket based on thrust vector system control according to claim 1, characterized in that: The posture sensor (12) is MPU6050, and the single chip microcomputer (13) is STM32.

6. The solid rocket engine for a sounding rocket based on thrust vector system control according to claim 1, characterized in that: The number of the steering gear (10) and the push rod (11) used is three. If the thrust force needs to be enhanced, the number of the steering gear (10) and the push rod (11) can be increased according to the actual situation.

7. The solid rocket engine for a sounding rocket based on thrust vector system control according to claim 1, characterized in that: The maximum angle at which the control support (5) can adjust the arc end surface (4) is 15°.

8. The solid rocket engine for a sounding rocket based on thrust vector system control according to claim 1, characterized in that: include: S1, the attitude sensor (12) calculates the attitude information such as the inclination angle and pitch angle of the rocket by measuring the acceleration of the rocket in different directions and the direction and strength of the earth's magnetic field. S2, the attitude sensor (12) sends the rocket attitude information to the single-chip microcomputer (13), the single-chip microcomputer (13) analyzes and processes the data, and outputs a PWM control signal to the steering gear (10) through a PID algorithm. S3, the steering engine (10) receives the PWM control signal and controls the push rod (11) to move, and the push rod (11) then drives the control support (5) to adjust the flight attitude of the rocket. S4, the attitude sensor (12) sends the rocket attitude information to the host computer (14) in real time, and the host computer (14) displays the rocket's position, attitude, flight altitude and other information in real time.

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