A fiber optic rate gyro combination for civilian launch vehicle control

Abstract

The present invention relates to a fiber optic rate gyro combination for civilian launch vehicle control, which includes a complete machine with five major components, including a fiber optic gyro unit, an information acquisition and processing circuit module unit, a power module unit, a complete machine structure unit, and an interactive interface unit. Its main function is to measure the angular velocity of the pitch and heading of the rocket body through two orthogonal fiber optic gyros, thereby realizing the measurement of the pitch and yaw state of the rocket body, and at the same time collect the working state of the secondary power supply, and then send the measurement result to the information acquisition and processing circuit module, and send the attitude information to the overall system through the 1553B bus protocol to complete the corresponding attitude control. The device can work stably and reliably after high-frequency random vibration up to 13.1g root mean square value and shock up to 1500g. The rate gyro combination has a 1553B communication module integrated inside the complete machine, and has the technical characteristics of miniaturization, light weight, low cost, reliability and versatility, meeting the market demand of civilian launch vehicles.

Description

Technical Field

[0001] The invention relates to the technical field of aerospace attitude measurement, and specifically designs an optical fiber rate gyro combination for controlling a civilian launch vehicle. Background Art

[0002] During the space flight, the launch vehicle needs to adjust its flight attitude in time to ensure the controllability and safety of the flight. Therefore, the detection of its flight attitude (pitch, yaw) becomes particularly important. The fiber optic gyroscope has the ability to measure attitude. It is composed of all-solid-state devices and has the characteristics of strong impact resistance and large dynamic range. It can meet the application needs of launch vehicle attitude detection.

[0003] With the rise of the civil launch vehicle market, it is mainly aimed at the commercial consumer market, with the characteristics of large launch volume and low launch cost. Therefore, it puts forward small size, lightweight and low-cost application requirements for rocket launch. Whether it is the traditional aerospace market or the civil aerospace market, the rocket is irreparable once launched, so the reliability index requirements for the sensors on the rocket body have not been reduced. On the other hand, the large impact caused by the rocket launch and separation process will bring severe challenges to the stability of the rate gyro sensor itself. If the rate gyro is equipped with a vibration reduction system, the bandwidth index of the rate gyro combination will be affected. Therefore, the traditional solution of adding a vibration reduction system will not meet the needs of civil rocket systems in terms of volume, weight, and cost. Therefore, it is necessary to optimize the fiber optic rate gyro system to meet the large overload, large vibration, and large impact index requirements of the civil rocket field for fiber optic gyroscopes. Summary of the invention

[0004] The present invention aims at the technical problems existing in the prior art and provides a miniaturized, low-cost, lightweight, and highly reliable fiber optic rate gyroscope combination for controlling a civilian launch vehicle, so as to meet the index requirements of large overload, large vibration, and large impact of the fiber optic gyroscope in the field of civilian rockets.

[0005] The technical solution of the present invention to solve the above technical problems is as follows:

[0006] A fiber optic rate gyro assembly for controlling a civilian launch vehicle, comprising:

[0007] A rectangular cylindrical main frame structure, one end of which has a protruding edge at its opening, a bottom plate matching the size of the opening, and a gas valve assembly installed on one side of the bottom plate close to the inside of the main frame structure;

[0008] A fiber optic gyroscope unit, wherein the fiber optic gyroscope unit is fixedly installed inside the main frame structure;

[0009] An information collection and processing circuit module unit is fixedly mounted on an outer side wall of the main frame structure and is communicatively connected with the fiber optic gyroscope unit;

[0010] A power module unit is fixedly mounted on another outer side wall of the main frame structure opposite to the information acquisition and processing circuit module unit; the power module unit is electrically connected to the fiber optic gyroscope unit and the information acquisition and processing circuit module unit through a cable;

[0011] The shell is a cavity structure with one end open; the shell is sleeved outside the main frame structure and is sealed with the protruding edge; the shell is also provided with a power connector and a communication connector, the power connector connects the power module unit with an external power supply, and the communication connector connects the information acquisition and processing circuit module unit with an external device.

[0012] The beneficial effects of the present invention are as follows: the device can work stably and reliably after high-frequency random vibration up to 13.1g RMS and shock up to 1500g, and the rate gyro combination is internally integrated with a 1553B communication module, and has the technical characteristics of miniaturization, light weight, low cost, reliability and versatility, meeting the market demand for civil launch vehicles.

[0013] Furthermore, the fiber optic gyroscope unit includes a fiber optic gyroscope Y1 and a fiber optic gyroscope Z1, and the fiber optic gyroscope Y1 and the fiber optic gyroscope Z1 are orthogonally arranged and are used to measure the angular velocity of the pitch and heading of the rocket body, respectively.

[0014] Furthermore, the fiber optic gyroscope unit is fixedly installed away from the power module unit.

[0015] Furthermore, a wire board is installed at one end of the main frame structure away from the protruding edge.

[0016] Furthermore, a conductive rubber rope is installed between the opening edge of the shell and the protruding edge.

[0017] Furthermore, a conductive rubber sealing ring is installed between the bottom plate and the edge of the opening at one end of the main frame structure provided with a protruding edge.

[0018] Furthermore, the main frame structure is made of metal.

[0019] Furthermore, thermal conductive silicone grease is coated between the power module unit and the main frame structure. BRIEF DESCRIPTION OF THE DRAWINGS

[0020] Figure 1 It is a block diagram of the composition of the present invention;

[0021] Figure 2 It is a three-dimensional schematic diagram of the present invention (without the outer cover);

[0022] Figure 3 It is the main structural framework diagram of the present invention;

[0023] Figure 4 It is a schematic diagram of the cross-sectional structure of the whole machine of the present invention;

[0024] Figure 5 It is a schematic diagram of the overall appearance of the present invention.

[0025] In the accompanying drawings, the components represented by the reference numerals are listed as follows:

[0026] 1. Fiber optic gyro unit, 2. Information acquisition and processing circuit module unit, 3. Power module unit, 4. Whole machine structure unit, 5. Interactive interface unit

[0027] 101 Fiber Optic Gyroscope Y1, 102 Fiber Optic Gyroscope Z1,

[0028] 401 air valve assembly, 402 bottom plate, 403 screw I, 404 screw II, 405 housing, 406 main frame structure, 407 conductive rubber rope, 408 conductive rubber sealing ring, 409 countersunk screw, 410 locking screw I, 411 hexagon socket mounting screw, 412 screw III, 413 locking screw II, 414 wire plate, 415 locking screw III

[0029] 501 power connector, 502 communication connector DETAILED DESCRIPTION

[0030] The principles and features of the present invention are described below in conjunction with the accompanying drawings. The examples given are only used to explain the present invention and are not used to limit the scope of the present invention.

[0031] Example

[0032] The embodiment of the present invention provides a fiber optic rate gyro combination for civilian launch vehicle control, and the design principle schematic diagram is shown as follows: Figure 1 As shown, the whole machine specifically includes five major components, namely, a fiber optic gyroscope unit 1, an information acquisition and processing circuit module unit 2, a power module unit 3, a whole machine structure unit 4, and an interactive interface unit 5. Its main function is to measure the angular velocity of the pitch and heading of the rocket body through two orthogonal fiber optic gyroscopes Y1 and Z1, thereby realizing the measurement of the pitch and yaw state of the rocket body, and at the same time collecting the working state of the secondary power supply, and then sending the measurement result to the information acquisition and processing circuit module, and sending the attitude information to the overall system through the 1553B bus protocol to complete the corresponding attitude control.

[0033] Specifically, Figures 2 to 5 As shown, the overall structural unit 4 includes: a main frame structure 406 , a shell 405 and a bottom plate 402 .

[0034] The main frame structure 406 is a rectangular cylindrical structure made of metal as a whole, with a protruding edge at one end of the opening. The bottom plate 402 is matched with the size of the opening, and the side of the bottom plate 402 close to the inside of the main frame structure 406 is equipped with an air valve assembly 401.

[0035] The outer shell 405 is a cavity structure with one end open, and is also made of metal. The outer shell 405 is sleeved outside the main frame structure 406 and is sealed and connected to the protruding edge.

[0036] The fiber optic gyroscope unit 1 is fixedly installed inside the main frame structure 406; the fiber optic gyroscope unit 1 includes a fiber optic gyroscope Y1101 and a fiber optic gyroscope Z1102, and the fiber optic gyroscope Y1101 and the fiber optic gyroscope Z1102 are orthogonally arranged and are used to measure the angular velocity of the pitch and heading of the rocket body respectively.

[0037] The information collection and processing circuit module unit 2 is fixedly mounted on an outer side wall of the main frame structure 406 and is communicatively connected with the fiber optic gyro unit 1 .

[0038] The power module unit 3 is fixedly mounted on the other outer side wall of the main frame structure 406 opposite to the information acquisition and processing circuit module unit 2; the power module unit 3 is electrically connected to the fiber optic gyroscope unit 1 and the information acquisition and processing circuit module unit 2 through cables. When fixing the power module 3, considering that it generates more heat, it is kept away from the two fiber optic gyroscopes.

[0039] The housing 405 is also provided with a power connector 501 and a communication connector 502. The power connector 501 connects the power module unit 3 with an external power source, and the communication connector 502 connects the information collection and processing circuit module unit 2 with an external device.

[0040] Preferably, a wire board 414 is installed at one end of the main frame structure 406 away from the protruding edge.

[0041] A conductive rubber rope 407 is installed between the opening edge of the housing 405 and the protruding edge.

[0042] A conductive rubber sealing ring 408 is installed between the bottom plate 402 and the edge of the opening of one end of the main frame structure 406 with a protruding edge.

[0043] Thermal conductive silicone grease is coated between the power module unit 3 and the main frame structure 406 .

[0044] When implementing it, Figures 2 to 5As shown, the air tightness of the whole structure is tested first. The test method is to assemble the air valve assembly 401 on the bottom plate 402 and tighten it. The power connector 501 and the 1553B communication connector 502 are fixed to the housing 405 of the whole structure with screws I403 and screws II404 respectively, and the fixed position is sealed.

[0045] First, the bottom plate 402 with the gas valve 401 and the main frame structure 406 are assembled and fixed with the conductive rubber seal ring 408 using fourteen countersunk screws 409, and then the whole is fixed with the housing 405 equipped with the power connector 501 and the 1553B connector 502 and the conductive rubber rope 407 using twenty locking screws I410. When fixing, all screws are sealed with removable thread glue. Finally, nitrogen gas at 1 atmosphere pressure is injected through the gas valve hole and maintained for 1 hour. After one hour, the gas leakage rate is no more than 25%, and the subsequent assembly of the whole machine can be carried out.

[0046] During the specific implementation, each unit module is tested to work normally, and the air tightness of the structural parts meets the requirements before the whole machine is assembled. The specific assembly process is: first, the fiber optic gyroscope 101 and the fiber optic gyroscope 102 are fixed to the main frame structure 406 with four hexagon socket mounting screws 411 respectively, and the non-verticality of the two gyroscopes is less than 20" after compensation. The second is to fix the voltage-stabilized power supply module 3. When fixing and assembling, it is necessary not only to ensure the tightness of the assembly, but also to achieve thermal conductivity and soft fit between the main heat dissipation device DC / DC module and the main frame structure 406 to ensure contact heat dissipation of the heat-generating module. It is fixed to the main frame structure 406 with six locking screws III412. When fixing the power supply module 3, considering that its heat generation is more serious, it is kept away from the two fiber optic gyroscopes. Screw. Considering the center of gravity of the assembly, the information acquisition and processing circuit module unit 2 is assembled to the opposite end face of the power module unit 3 with six locking screws II413. Then, the electrical circuit is assembled and the connecting cables are fixed to the wire board 414 with a tie rope. The wire board 414 is fixed to the main frame structure 406 with four locking screws III415. Finally, the conductive rubber rope 407 is installed, and the outer shell 405 is fastened with twenty locking screws I410. After turning the whole machine upside down, a conductive rubber ring 408 is placed at the bottom of the main frame structure 406, and the bottom plate 402 equipped with the air valve 401 is fastened to the mounting surface at the bottom of the main frame structure 406 with fourteen countersunk screws 409, and the whole machine is assembled.

[0047] The size of the whole machine can be reduced to 130mm*142mm*95mm, the weight is less than 1.8kg, the hardware cost is less than 200,000, and it can work stably and reliably after high-frequency random vibration up to 13.1g RMS and shock up to 1500g. The rate gyro combination has a 1553B communication module integrated inside the whole machine, and has the technical characteristics of miniaturization, light weight, low cost, reliability and versatility, meeting the market demand for civil launch vehicles.

[0048] The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc. made within the spirit and principle of the present invention should be included in the protection scope of the present invention.

Claims

1. A fiber optic rate gyro assembly for civilian launch vehicle control, characterized in that: include: A rectangular cylindrical main frame structure (406) has an opening at one end thereof with a protruding edge extending outwardly, a bottom plate (402) matching the size of the opening is provided at the opening, and a gas valve assembly (401) is installed on one side of the bottom plate (402) close to the inside of the main frame structure (406); A fiber optic gyroscope unit (1), the fiber optic gyroscope unit (1) being fixedly mounted inside the main frame structure (406); An information collection and processing circuit module unit (2) is fixedly mounted on an outer side wall of the main frame structure (406) and is communicatively connected to the fiber optic gyroscope unit (1); A power module unit (3) is fixedly mounted on another outer side wall of the main frame structure (406) opposite to the information collection and processing circuit module unit (2); the power module unit (3) is electrically connected to the fiber optic gyroscope unit (1) and the information collection and processing circuit module unit (2) via a cable; The housing (405) is a cavity structure with one end open; the housing (405) is sleeved outside the main frame structure (406) and is sealed with the protruding edge; the housing (405) is also provided with a power connector (501) and a communication connector (502); the power connector (501) connects the power module unit (3) with an external power source, and the communication connector (502) connects the information collection and processing circuit module unit (2) with an external device; The fiber optic gyroscope unit (1) comprises a fiber optic gyroscope Y1 (101) and a fiber optic gyroscope Z1 (102), wherein the fiber optic gyroscope Y1 (101) and the fiber optic gyroscope Z1 (102) are arranged orthogonally and are used to measure the angular velocity of the pitch and heading of the rocket body respectively; A conductive rubber rope (407) is installed between the opening edge of the housing (405) and the protruding edge; A conductive rubber sealing ring (408) is installed between the bottom plate (402) and the edge of the opening at one end of the main frame structure (406) provided with a protruding edge; Thermal conductive silicone grease is coated between the power module unit (3) and the main frame structure (406).

2. The fiber optic rate gyro assembly according to claim 1, characterized in that: The optical fiber gyroscope unit (1) is fixedly installed away from the power module unit (3).

3. The fiber optic rate gyro assembly according to claim 1, characterized in that: A wire board (414) is installed at one end of the main frame structure (406) away from the protruding edge.

4. The fiber optic rate gyro assembly according to claim 1, characterized in that: The main frame structure (406) is made of metal.

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