WO2011140804A1 - 一种微型惯性测量系统 - Google Patents

一种微型惯性测量系统 Download PDF

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Publication number
WO2011140804A1
WO2011140804A1 PCT/CN2010/079483 CN2010079483W WO2011140804A1 WO 2011140804 A1 WO2011140804 A1 WO 2011140804A1 CN 2010079483 W CN2010079483 W CN 2010079483W WO 2011140804 A1 WO2011140804 A1 WO 2011140804A1
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WO
WIPO (PCT)
Prior art keywords
sensing
circuit board
bracket
measurement system
inertial measurement
Prior art date
Application number
PCT/CN2010/079483
Other languages
English (en)
French (fr)
Inventor
汪滔
Original Assignee
Wang Tao
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Wang Tao filed Critical Wang Tao
Priority to CN201080068428.XA priority Critical patent/CN103210280B/zh
Priority to US13/809,407 priority patent/US9213046B2/en
Priority to EP10851308.6A priority patent/EP2604974B1/en
Priority to JP2013523464A priority patent/JP6154324B2/ja
Publication of WO2011140804A1 publication Critical patent/WO2011140804A1/zh
Priority to US14/940,721 priority patent/US10132827B2/en
Priority to US16/189,782 priority patent/US10732200B2/en
Priority to US16/945,208 priority patent/US11215633B2/en
Priority to US17/646,635 priority patent/US20220120782A1/en

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Classifications

    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01P—MEASURING LINEAR OR ANGULAR SPEED, ACCELERATION, DECELERATION, OR SHOCK; INDICATING PRESENCE, ABSENCE, OR DIRECTION, OF MOVEMENT
    • G01P15/00—Measuring acceleration; Measuring deceleration; Measuring shock, i.e. sudden change of acceleration
    • G01P15/14—Measuring acceleration; Measuring deceleration; Measuring shock, i.e. sudden change of acceleration by making use of gyroscopes
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
    • G01C19/00—Gyroscopes; Turn-sensitive devices using vibrating masses; Turn-sensitive devices without moving masses; Measuring angular rate using gyroscopic effects
    • G01C19/56—Turn-sensitive devices using vibrating masses, e.g. vibratory angular rate sensors based on Coriolis forces
    • G01C19/5783—Mountings or housings not specific to any of the devices covered by groups G01C19/5607 - G01C19/5719
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F16—ENGINEERING ELEMENTS AND UNITS; GENERAL MEASURES FOR PRODUCING AND MAINTAINING EFFECTIVE FUNCTIONING OF MACHINES OR INSTALLATIONS; THERMAL INSULATION IN GENERAL
    • F16F—SPRINGS; SHOCK-ABSORBERS; MEANS FOR DAMPING VIBRATION
    • F16F15/00—Suppression of vibrations in systems; Means or arrangements for avoiding or reducing out-of-balance forces, e.g. due to motion
    • F16F15/02—Suppression of vibrations of non-rotating, e.g. reciprocating systems; Suppression of vibrations of rotating systems by use of members not moving with the rotating systems
    • F16F15/022—Suppression of vibrations of non-rotating, e.g. reciprocating systems; Suppression of vibrations of rotating systems by use of members not moving with the rotating systems using dampers and springs in combination
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F16—ENGINEERING ELEMENTS AND UNITS; GENERAL MEASURES FOR PRODUCING AND MAINTAINING EFFECTIVE FUNCTIONING OF MACHINES OR INSTALLATIONS; THERMAL INSULATION IN GENERAL
    • F16F—SPRINGS; SHOCK-ABSORBERS; MEANS FOR DAMPING VIBRATION
    • F16F15/00—Suppression of vibrations in systems; Means or arrangements for avoiding or reducing out-of-balance forces, e.g. due to motion
    • F16F15/02—Suppression of vibrations of non-rotating, e.g. reciprocating systems; Suppression of vibrations of rotating systems by use of members not moving with the rotating systems
    • F16F15/04—Suppression of vibrations of non-rotating, e.g. reciprocating systems; Suppression of vibrations of rotating systems by use of members not moving with the rotating systems using elastic means
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F16—ENGINEERING ELEMENTS AND UNITS; GENERAL MEASURES FOR PRODUCING AND MAINTAINING EFFECTIVE FUNCTIONING OF MACHINES OR INSTALLATIONS; THERMAL INSULATION IN GENERAL
    • F16F—SPRINGS; SHOCK-ABSORBERS; MEANS FOR DAMPING VIBRATION
    • F16F15/00—Suppression of vibrations in systems; Means or arrangements for avoiding or reducing out-of-balance forces, e.g. due to motion
    • F16F15/02—Suppression of vibrations of non-rotating, e.g. reciprocating systems; Suppression of vibrations of rotating systems by use of members not moving with the rotating systems
    • F16F15/04—Suppression of vibrations of non-rotating, e.g. reciprocating systems; Suppression of vibrations of rotating systems by use of members not moving with the rotating systems using elastic means
    • F16F15/08—Suppression of vibrations of non-rotating, e.g. reciprocating systems; Suppression of vibrations of rotating systems by use of members not moving with the rotating systems using elastic means with rubber springs ; with springs made of rubber and metal
    • F16F15/085—Use of both rubber and metal springs
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
    • G01C21/00—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00
    • G01C21/10—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration
    • G01C21/12—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration executed aboard the object being navigated; Dead reckoning
    • G01C21/16—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration executed aboard the object being navigated; Dead reckoning by integrating acceleration or speed, i.e. inertial navigation
    • G01C21/166—Mechanical, construction or arrangement details of inertial navigation systems
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01P—MEASURING LINEAR OR ANGULAR SPEED, ACCELERATION, DECELERATION, OR SHOCK; INDICATING PRESENCE, ABSENCE, OR DIRECTION, OF MOVEMENT
    • G01P1/00—Details of instruments
    • G01P1/02—Housings
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01P—MEASURING LINEAR OR ANGULAR SPEED, ACCELERATION, DECELERATION, OR SHOCK; INDICATING PRESENCE, ABSENCE, OR DIRECTION, OF MOVEMENT
    • G01P15/00—Measuring acceleration; Measuring deceleration; Measuring shock, i.e. sudden change of acceleration
    • G01P15/18—Measuring acceleration; Measuring deceleration; Measuring shock, i.e. sudden change of acceleration in two or more dimensions

Definitions

  • the present invention relates to strapdown inertial navigation technology for use in a carrier such as a drone, and more particularly to a miniature inertial measurement system for use in strapdown inertial navigation. Background technique
  • Strapdown inertial navigation is an advanced navigation technology that is rapidly developing today. It uses the inertial components such as gyroscopes and accelerometers directly attached to the carrier to measure the acceleration of the carrier relative to the inertial reference frame. The integral operation is performed according to the Newton's inertia principle to obtain the velocity, attitude angle and position in the navigation coordinate system. Information, guiding the carrier from the starting point to the destination. Strapdown inertial navigation technology uses the control computer to perform coordinate transformation on the data measured by the gyroscope and accelerometer, and solve mathematical operations such as differential equations, and extracts the attitude and heading data from the elements of the attitude matrix to realize the navigation task.
  • inertial components such as gyroscopes and accelerometers directly attached to the carrier to measure the acceleration of the carrier relative to the inertial reference frame.
  • the integral operation is performed according to the Newton's inertia principle to obtain the velocity, attitude angle and position in the navigation coordinate system. Information, guiding the carrier from the starting point to the
  • the Strapdown Inertial Navigation System uses a data such as the Strapdown Matrix that is updated at any time to create a "mathematical platform" that replaces the traditional electromechanical navigation platform, which greatly simplifies the system structure, greatly reduces the size and cost of the system, and facilitates the installation and maintenance of inertial components.
  • the strapdown inertial navigation system does not rely on external system support, obtains attitude, speed and position information autonomously, and does not radiate any information to the outside world. It has real-time autonomy, is free from interference, and is not limited by geographical, temporal and climatic conditions.
  • the output parameters are comprehensive and so on, and are widely used in various fields such as aviation, navigation, and transportation.
  • the strapdown inertial navigation system typically consists of an inertial measurement system, a control computer, a control display, and associated support components. Its core component inertial measurement system is equipped with a gyroscope and accelerometer.
  • the working principle of the inertial measurement system is: The gyroscope detects the three-axis angular velocity of the carrier, and the accelerometer detects the linear acceleration of the aircraft along the three-axis motion, and the control computer measures the gyroscope.
  • the angular rate signal is used for time integral calculation, and the navigational attitude information such as instantaneous heading and dip angle is calculated.
  • the acceleration signal measured by the accelerometer is used to calculate the instantaneous navigation speed information for the time integral calculation.
  • the second integration is used to calculate the time period. The distance and location of the voyage.
  • the inertial measurement system and its attitude calculation technology are the key technical links that affect the performance of the strapdown inertial navigation system. This is because inertial measurement and its attitude calculation are the premise for the trajectory control of the vehicle. Its accuracy and efficiency directly affect the timeliness and accuracy of the navigation. Second, the inertial measurement system should directly withstand the vibration in a harsh aerodynamic environment. Impact and angular motion, which cause many instability and error effects, become the weak link in the strapdown inertial navigation system. Third, the strapdown inertial navigation system faces the challenges of miniaturization and industrialization, especially with microelectronics. The development of technology requires the use of medium-precision or even low-precision MEMS inertial components to achieve the goal of low-cost mass production of strapdown inertial navigation products.
  • Inertial measurement systems must propose targeted technical measures in terms of mechanical structure, vibration reduction design, and micro-chemical technology to overcome the drawbacks of unstable navigation, reduced accuracy, and even shortened service life of electronic components.
  • the sensing bracket 11 is fastened inside the casing 12 by fastening screws, and the damping unit 13 is composed of four rubber pads, and the casing is fixed to the aircraft from the bottom.
  • the sensing bracket is composed of three mutually perpendicular gyro circuit boards 111, 112, 113 (see Fig. 2) on which three single-axis gyroscopes 111a, 112a, 113a are respectively mounted.
  • the horizontally placed gyro circuit board 1 11 is a combined gyro circuit board, which is equipped with a three-axis accelerometer in addition to the gyro 111a.
  • the three gyroscopes should be mounted on three orthogonal planes whose their sensitive axes are perpendicular to each other to form a measured orthogonal coordinate system.
  • the measuring axis of the three-axis accelerometer 111b on the combined gyro circuit board 111 and the gyro 111a on the circuit board The measuring axes are placed in parallel.
  • the combined gyro circuit board 111 is directly connected to the conditioning circuit board 114 and the main processor circuit board 115 through a connector.
  • Fig. 3 The equivalent analysis of the damping structure of the above inertial measurement system is shown in Fig. 3.
  • the mass M represents the inertial measurement system, and its center of mass is m;
  • the damping unit is represented by ⁇ , ⁇ , which represents the stiffness, represents the damping coefficient, and the subscript i Indicates the number of damping units included in the damper.
  • i l, 2, 3, 4;
  • B represents the voyage carrier;
  • P is the damper Elastic center.
  • the damping unit ⁇ , ⁇ absorbs and consumes the forced vibration energy from the carrier B, and takes the P point as the center to perform the upper and lower elastic motion, thereby reducing The impact of the small carrier B vibration on the inertial measurement system m.
  • the sensing bracket structure is three separate circuit boards, occupying a large space, and the three axial stiffnesses are significantly different;
  • the vibration damping unit is installed outside the inertial measurement system, which not only takes up extra space, but more importantly, when the inertial measurement unit is forced to vibrate, the mechanical structure is unreasonable due to unbalanced stiffness, and the inertial measurement system is prone to twist when subjected to vibration. Vibration
  • the ideal range of the damper is limited to the uniaxial direction, that is, the vibration from the vertical X direction can only be normally attenuated, and the vibration damping in other directions cannot be effectively suppressed, so that the line vibration and the angular vibration in different degrees of freedom are enabled. Coupling occurs, and the damping band is narrow.
  • the technical solution of the present invention is to construct a miniature inertial measurement system, including a housing, a sensing component, and a damper; wherein the sensing component includes a rigid sensing bracket and is mounted on a measurement and control circuit board on the surface of the sensing bracket, and an inertial sensor disposed on the measurement and control circuit board, the inertial sensor includes a gyroscope and an accelerometer; the sensing component is mounted in the housing; The vibrator is mounted in the housing and disposed in a gap between the sensing component and the inner wall of the housing.
  • the sensing bracket is a rectangular rigid bracket, and a groove is engraved on at least one surface thereof;
  • the measuring and controlling circuit board is a flexible measuring and controlling circuit board; and the flexible measuring and controlling circuit board has at least a part of the circuit The component is embedded in a recess of the at least one surface.
  • the six surfaces of the sensing bracket are engraved with grooves; the flexible measuring and controlling circuit board has six blocks and covers the six surfaces of the sensing bracket respectively;
  • the circuit components on the flexible measuring and controlling circuit board are respectively embedded in the grooves of the surface of the sensing bracket, so that the flexible measuring and controlling circuit board smoothly covers each surface of the sensing bracket.
  • the sensing component further includes an anti-aliasing circuit and an A/D conversion circuit disposed on the flexible measuring and controlling circuit board;
  • the inertial sensor includes three gyroscopes and an accelerometer;
  • the six circuit modules are respectively disposed on the six flexible measurement and control circuit boards.
  • the six flexible measuring and controlling circuit boards are of a unitary structure, and are bent 90° along the edges of the sensing bracket to completely cover the respective surfaces of the sensing bracket.
  • the damper includes at least two damper units disposed in a gap between one of the surfaces of the sensing assembly and the inner wall of the housing. Among them, it is preferable that the damper includes six damper units.
  • the sensing component is suspended by the six damping unit at a center of a cavity of the housing, and an elastic center point P of the damper and a center of mass of the sensing component m coincides.
  • the housing includes an upper casing having a lower opening, and is mounted on the The lower cover at the opening.
  • the present invention has the following advantages: (1) not only enhances the rigidity of the bracket, but also improves the mechanical structure of the system, and achieves three-way stiffness reduction, so that the noise resistance of the inertial measurement system is greatly improved; (2) Improve the vibration characteristics of the inertial measurement system, so that the natural frequency is away from the operating frequency of sensitive devices such as the gyro shaker, and the relative amplitude of the inertial sensor mounting surface is minimized; (3) The inertia measurement unit volume and weight are greatly reduced. , expanding the load space of the carrier.
  • FIG. 1 Schematic diagram of the existing small-sized UAV strap-down inertial measurement system.
  • FIG. 2 is a schematic structural view of a sensing bracket in the inertial measurement system shown in FIG. 1.
  • FIG. 3 is a schematic diagram of an equivalent model of a vibration damping system in the inertial measurement system shown in FIG.
  • Fig. 4 is a schematic view showing the distribution of the internal damping unit of the damper according to an embodiment of the present invention, wherein S is the upper and lower sides and the left and right inner walls of the housing.
  • Figure 5 is a schematic illustration of a sensing bracket in accordance with a preferred embodiment of the present invention.
  • Fig. 6 is a schematic view showing the outline and component arrangement of the flexible measuring and controlling circuit board cooperated with Fig. 5.
  • Figure 7 is a schematic view showing the construction of a sensing unit in a preferred embodiment of the present invention.
  • Figure 8 is a schematic view of the structure of the housing mated with Figure 7.
  • Figure 9 is a schematic illustration of the positional relationship of the internal damping unit and the sensing assembly employed in a preferred embodiment of the present invention.
  • Figure 10 is a schematic illustration of the complete assembly of a miniature inertial measurement system in accordance with a preferred embodiment of the present invention. detailed description
  • Vigorous random vibration is the main mechanical environment that the strapdown inertial navigation system faces during operation. Vibration causes system performance instability or electronic component damage, which has a great impact on system stability. In order to reduce the violent random vibration of the carrier, the electronic components are damaged or the inertial measurement unit is unstable.
  • the damper is used as a damping medium to elastically couple the inertial measurement unit to the carrier to obtain a satisfactory damping effect.
  • the selection of the damping mode not only affects the damping performance of the inertial navigation system, but also affects the measurement accuracy of the system. It has always been an important part of the structural design of the inertial navigation system.
  • the invention starts from two aspects of improving the design of the sensing bracket and rationalizing the vibration damping mechanical structure, and improving the performance of the micro inertial measurement system.
  • the sensing bracket is a key component for installing the gyroscope and the measuring and controlling circuit board and the connecting wire. It is subjected to various severe vibrations during operation.
  • the relative amplitude of the mounting surface of the gyroscope on the bracket is the largest, and the dynamic performance of the structure will affect the working of the gyroscope. Reliability and accuracy require a certain amount of static strength, vibration strength and fatigue life. In terms of process, the bracket is required to be easy to install and easy to manufacture.
  • the bracket structure is rationally designed to improve the stiffness and damping characteristics of the structure, so that the natural frequency of the structure must be kept away from the operating vibration frequency of the gyro shaker, so that the relative amplitude of the gyroscope mounting surface is minimized.
  • the improved bracket design can not be contrary to the traditional thinking.
  • the method of increasing the wall thickness is adopted to increase the stiffness and increase the natural frequency of the structure.
  • the structural rigidity and damping of the bracket should be improved by improving the structural design of the material, shape and joint surface. Moreover, it is necessary to proceed from the whole, to deal with the mutual restraint relationship between the bracket and the vibration damping device, and also to consider the installation position and the line direction of the measuring and controlling circuit board on the bracket.
  • the technical measures adopted by the present invention are: Starting from improving the mechanical structure of the inertial measurement system, providing a greatly reduced volume, three-way stiffness reduction
  • the miniature inertial measurement system of the vibrating structure overcomes the adverse effects of the three-way stiffness, the resonance excitation, and the generation of torsional vibrations on the strapdown inertial navigation system.
  • the miniature inertial measurement system includes a sensing component 12, a vibration damping unit, an upper casing 16, a lower cover 18, and the like, wherein: the sensing component 12 is transmitted.
  • the sensing bracket 121, the inertial sensor 122, and the flexible measuring and controlling circuit board 123 are composed of:
  • the sensing bracket 121 is a rectangular rigid bracket with grooves in each plane and meeting certain specific gravity and rigidity requirements.
  • the inertial sensor 122 includes a gyroscope and an accelerometer including three gyroscopes and an accelerometer that are coupled to the flexible measurement and control circuit board 123.
  • the flexible measurement and control circuit board 123 should include a sensor signal pre-processing function including at least an anti-aliasing circuit and an A/D conversion circuit; the circuit board base and the connecting wires are made of a flexible material to withstand a 90° bend; the flexible measurement and control circuit board The shape should be identical to the plane of the sensing bracket. When it is bent 90 ° along the edge of the sensing bracket, the entire flexible measuring and controlling circuit board can completely and smoothly cover each plane of the sensing bracket.
  • the anti-aliasing circuit, the A/D conversion circuit, the three gyroscopes, and one accelerometer are provided on six flexible measurement and control circuit boards.
  • the circuit components on each flexible measurement and control circuit board are respectively embedded in the grooves on the surface of the sensing bracket.
  • the shape of the inner cavity formed by the upper casing 16 and the lower cover 18 should be similar to the outer shape of the sensing component 12 and the space is slightly larger, so that substantially the same space is left between the inner walls of the casing and the corresponding planes of the sensing components for installation and reduction.
  • Vibration unit 14 The inner damper is composed of a plurality of internal damping unit constituent units ⁇ , ⁇ 14 having appropriate damping characteristics, which are installed between the inner wall S of the upper casing 16 and the six planes of the sensing assembly 12, which are determined according to different vibration characteristics of the carrier. The number can be up to six.
  • the sensing component is suspended in the center of the inner cavity of the housing, and the deformation force axes of the inner damping unit are orthogonal to each other, and the elastic center point P of the inner damper coincides with the centroid m of the sensing component to balance absorption and consume the forcing from the carrier. vibration.
  • the vibration damping unit is composed of an elastic material having a certain damping effect, and may be, but not limited to, a spring, a rubber mat, a silica gel, a sponge, or other vibration damping material.
  • the sensing bracket is made of a metal or non-metal material having a certain specific gravity and rigidity, and is integrally processed into a square sensing bracket 121. The overall processing rather than the assembly is to ensure that the bracket itself has sufficient Rigidity to reduce stiffness and anisotropy See Figure 5 for volume error;
  • Figure 6 is a plan view showing the planar development and component arrangement of the flexible measurement and control circuit board 123 in a preferred embodiment of the present invention.
  • the circuit board base and connecting wires of the flexible measuring and controlling circuit board 123 are made of a flexible material and can withstand a 90° bend; the shape is designed to be identical to the outer plane of the sensing bracket, and thus has six unfolding planes.
  • the sensor and other electronic components are attached to the appropriate positions on the front faces of the six unfolded planes.
  • FIG. 7 is a schematic view showing the construction of a sensing unit in a preferred embodiment of the present invention.
  • the flexible measurement and control circuit board 123 has an inertial sensor 122 and other electronic components on the front side.
  • the sensing bracket 121 is attached to the front surface of the flexible measuring and controlling circuit board, and is bent at 90° along the edge of the sensing bracket. After the sensors or electronic components are embedded in the grooves of the planes of the sensing bracket, the whole flexible measuring and controlling is performed.
  • the back of the board faces outward, enveloping the sensing bracket with the sensing and electronics components and covering each plane of the sensing bracket completely and smoothly.
  • the invention considers avoidance or reduction of vibration coupling as a primary consideration when designing the strapdown inertial navigation damping system. If the mechanical structure of the system is unreasonable, the vibrations of the six degrees of freedom of the system are coupled to each other, and the cross vibration of the line vibration and the angular vibration is generated. As a result, the detection data of the inertial measurement system contains strong self-excitation information, which will introduce the system into the pseudo. The motion signal seriously affects the measurement accuracy of the inertial navigation system. In order to reduce the interference of the damper to the angular motion measurement of the system, the angular vibration frequency of the damping system should be as far as possible from the measurement bandwidth of the inertial navigation system. Under broadband random vibration conditions, the lower the damping frequency, the higher the damping efficiency.
  • the unit 14 is all installed therein, and after forming the inner damper assembly, a better damping effect is produced.
  • Figure 9 is a schematic illustration of the positional relationship of the inner damper unit 14 and the sensing assembly after the internal damper unit 14 is constructed in accordance with a preferred embodiment of the present invention.
  • the present embodiment uses six inner damping units 14, that is, six identical damping pads, to be installed. Between the inner wall of the upper casing 16 and the sensing assembly 12, the sensing assembly is suspended at the center of the casing inner cavity, and the deformation force axes of the inner damping units are orthogonal to each other to balance absorption and consume forced vibration from the carrier. .
  • Figure 10 is a schematic illustration of the complete assembly of the miniature inertial measurement system 2.1 in accordance with a preferred embodiment of the present invention. Due to the implementation of the above series of technical measures, the natural frequency, damping coefficient, vibration damping efficiency and mechanical strength of the damper are guaranteed to meet the impact resistance and vibration requirements of the system; the elastic coordinate system and inertia of the miniature inertial measurement system are made.
  • the three coordinate systems of the coordinate system and the solution coordinate system are in an optimal state in which the corresponding coordinate axes are parallel to each other, and the center of mass of the system coincides with the elastic center of the vibration damping device, so as to achieve a high decoupling effect between the respective vibrations, and The natural frequencies are close to each other, and the technical effect of a narrower frequency distribution is obtained.
  • the miniature inertial measurement system of the present invention can be used for an autonomous aircraft such as a drone, a vessel, an underwater automatic detection device, or various vehicles, robots, and the like.
  • the present invention may have other embodiments, for example: (1)
  • the housing structure is not limited to the structure in which the upper case and the lower cover are matched, or the lower case is matched with the upper cover, or the intermediate case is The body is matched with the upper and lower covers; (2) all or part of the six functional modules on the flexible measurement and control circuit board can be integrated, so that The number of flexible measuring and controlling circuit boards can be reduced to six or less.
  • the bracket can also be a rectangular parallelepiped structure, of course, the structure of the circuit board needs to be correspondingly change. It can be seen that the related equivalent replacement technical solutions fall within the protection scope of the present invention.

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  • Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Radar, Positioning & Navigation (AREA)
  • Remote Sensing (AREA)
  • General Engineering & Computer Science (AREA)
  • General Physics & Mathematics (AREA)
  • Mechanical Engineering (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Acoustics & Sound (AREA)
  • Automation & Control Theory (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Gyroscopes (AREA)
  • Navigation (AREA)

Description

一种微型惯性测量系统 技术领域
本发明涉及用于无人机等运载体中的捷联惯性导航技术， 尤其是涉及 一种用于捷联惯性导航中的微型惯性测量系统。 背景技术
捷联式惯性导航是当今迅速发展的一种先进导航技术。 它利用直接固 连在运载体上的陀螺仪、 加速度计等惯性元件测量出运载体相对于惯性参 考系的加速度， 按照牛顿惯性原理进行积分运算， 获得导航坐标系下的速 度、 姿态角和位置信息， 引导运载体从起始点驶向目的地。 捷联惯性导航 技术利用控制计算机将陀螺仪、 加速度计测得的数据进行坐标变换、 求解 微分方程等数学运算， 从姿态矩阵的元素中提取姿态和航向数据， 实现导 航任务。 捷联惯性导航系统利用随时更新的捷联矩阵等数据建立 "数学平 台"， 取代传统的机电式导航平台， 从而大大简化了系统结构， 使系统的体 积和成本大幅度降低， 惯性元件便于安装维护； 此外， 捷联惯性导航系统 不依赖外部系统支持， 自主获得姿态、 速度和位置信息， 也不向外界辐射 任何信息， 具有实时自主， 不受干扰， 不受地域、 时间、 气候条件限制， 以及输出参数全面等优点， 被广泛于航空、 航海、 交通等多种领域。
捷联惯性导航系统通常由一个惯性测量系统、 一个控制计算机、 控制 显示器和相关支持部件构成。 其核心部件惯性测量系统装有陀螺仪和加速 度计。 惯性测量系统的工作原理是： 陀螺仪检测运载体三轴角速率， 加速 度计检测航行器沿着三轴运动的线性加速度， 控制计算机将陀螺仪所测的 角速率信号对时间积分运算， 推算出瞬时航向、 倾角等航行姿态信息， 利 用加速度计测得的加速度信号， 对时间积分运算， 推算出瞬时航行速度信 息； 进行二次积分， 即可推算该时段内航行的距离和位置。
惯性测量系统及其姿态解算技术， 是影响捷联式惯性导航系统性能的 关键技术环节。 这是因为惯性测量及其姿态解算， 是对运载体实施轨迹控 制的前提， 它的精度和效率直接影响导航的时效和精度； 第二， 惯性测量 系统要在严酷的气动环境中直接承受振动、 冲击和角运动， 引发诸多的失 稳和误差效应， 成为捷联惯性导航系统中薄弱环节； 第三， 捷联式惯性导 航系统面临微型化、 产业化方面的挑战， 特别是随着微电子技术的发展， 要求采用中精度甚至低精度的微机电惯性元器件， 到达低成本批量化生产 捷联惯性导航产品的目的。
当运载体趋于小型化、 微型化时， 其基础质量与常规运载体相比有大 幅度减小， 在航行动力环境中受到的激扰和随机振动比常规载体更为剧烈， 系统更不稳定。 惯性测量系统必须在力学结构、 减振设计， 以及微型化工 艺等方面提出针对性技术措施， 克服导航不稳， 精度下降， 甚至电子元器 件使用寿命缩短的缺陷。
图 1 是现有一种小型无人机捷联惯性导航系统中所用的惯性测量系统 结构示意图。 其中采用紧固螺钉将传感支架 1 1紧固在壳体 12 内部， 再用 由四个橡胶垫组成减振单元 13， 从底部将壳体固接在航行器上。 传感支架 由三块相互垂直的陀螺电路板 111、 112、 113组成 （参见图 2 )， 上面分别 安装三个单轴陀螺仪 l l la、 112a, 113a。 其中水平放置的陀螺电路板 1 11 为组合陀螺电路板， 上面除了装有陀螺 111a 以外， 还装有三轴加速度计 l l lb。 三个陀螺仪应安装于三个正交平面上， 它们的敏感轴互相垂直， 构 成测量正交坐标系； 组合陀螺电路板 111上三轴加速度计 111b的测量轴与 该电路板上的陀螺 111a测量轴平行放置。 组合陀螺电路板 111通过接插件 与调理电路板 114及主处理器电路板 115直接连接。
上述惯性测量系统的减振结构等效分析见图 3， 图中质量块 M代表惯性 测量系统， 其质心为 m; 减振单元用 { ，^表示， 其中 表示刚度、 表示 阻尼系数， 下标 i表示减振器中所包含的减振单元的数量， 对于图 1采用 4 个橡胶垫作为减振单元， 则 i = l， 2， 3， 4； B代表航行运载体； P为减振 器的弹性中心。 当运载体 B航行运动时， 对惯性测量系统 m产生基础激励， 减振单元 {^，^吸收并消耗来自运载体 B的强迫振动能量， 以 P点为中心， 作上下弹性运动， 以此减小运载体 B振动对惯性测量系统 m造成的冲击。
上述惯性测量系统存在的问题是：
( 1 )传感支架结构是三块相互分离的电路板， 占用空间大， 三个轴向 刚度明显差异；
( 2 )减振单元安装在惯性测量系统的外部， 不仅额外占用空间， 更重 要的是当惯性测量单元受迫振动时， 由于刚度不均衡， 力学结构不合理， 受振时惯性测量系统容易产生扭转振动；
( 3)减振器的理想作用范围限于单轴方向， 即只能正常衰减来自铅垂 X方向的振动，而对其他方向的减振不能有效抑制，使不同自由度上的线振、 角振之间发生耦合， 减振频带窄。 发明内容 针对现有技术的上述缺陷， 本发明要解决传统惯性测量系统中存在的 占用空间较大、 容易产生扭转振动、 减振频带窄等问题。
为解决上述技术问题， 本发明的技术方案是， 构造一种微型惯性测量 系统， 包括壳体、 传感组件以及减振器； 其特征在于， 所述传感组件包括 刚性传感支架、 装于所述传感支架表面的测控电路板、 以及设于所述测控 电路板上的惯性传感器， 所述惯性传感器包括陀螺仪和加速度计； 所述传 感组件装于所述壳体内； 所述减振器装于所述壳体内并设于所述传感组件 与壳体内壁之间的空隙中。
本发明的优选方案中， 所述传感支架是正方体刚性支架， 在它的至少 一个表面刻有凹槽； 所述测控电路板是柔性测控电路板； 所述柔性测控电 路板上至少有一部分电路元件嵌于所述至少一个表面的凹槽中。
本发明的优选方案中， 在所述传感支架的六个表面均刻有凹槽； 所述 柔性测控电路板共有六块、 并分别覆盖于所述传感支架的六个表面； 每一 块所述柔性测控电路板上的电路元件分别嵌装于其所在传感支架表面的凹 槽中， 使得所述柔性测控电路板平顺地覆盖于所述传感支架的每一个表面。
本发明的优选方案中， 所述传感组件中还包括设于所述柔性测控电路 板上的抗混叠电路和 A/D转换电路； 所述惯性传感器包括三个陀螺仪和一 个加速度计； 所述共六个电路模块分别设于所述六块柔性测控电路板上。
本发明的优选方案中， 所述六块柔性测控电路板为一体结构， 并沿所 述传感支架的棱边作 90° 弯折后完整覆盖于传感支架的各个表面。
本发明的优选方案中， 所述减振器中包括至少两个减振单元， 分别设 于所述传感组件的其中一个表面与所述壳体内壁之间的空隙中。 其中， 所 述减振器中最好包括六个减振单元。
本发明的优选方案中， 所述传感组件被所述六个减振单元悬挂于所述 壳体的内腔中心， 且所述减振器的弹性中心点 P与所述传感组件的质心 m 重合。
本发明的优选方案中， 所述壳体包括下部开口的上壳、 以及装于所述 开口处的下盖。
由于采取了上述技术方案， 本发明具有以下优点： （1 ) 不仅增强了支 架刚性， 还改进了系统力学结构， 实现了三向等刚度减振， 使惯性测量系 统的抗噪能力大幅度提高； （2 ) 改善了惯性测量系统振动特性， 使其固有 频率远离陀螺仪抖动器等敏感器件的工作频率， 惯性传感器安装面的相对 振幅降低到最小； （3 ) 大幅度缩小惯性测量单元体积和重量， 扩大了运载 体的载荷空间。 附图说明
图 1现有小型无人机捷联惯性测量系统示意图。
图 2是图 1中所示惯性测量系统中的传感支架结构示意图。
图 3是图 1所示惯性测量系统中的减振系等效模型示意图。
图 4是本发明一个实施例中减振器所属内减振单元分布示意图，图中 S 为壳体上下、 左右四个内壁。
图 5是本发明一个较佳实施例的传感支架示意图。
图 6是与图 5配合的柔性测控电路板的外形和元器件布置示意图。 图 7是本发明一个较佳实施例中传感组件构成示意图。
图 8是与图 7配合的壳体结构示意图。
图 9是本发明一个较佳实施例中采用的内减振单元与传感组件的位置 关系示意图。
图 10 是本发明一个较佳实施例中微型惯性测量系统的完整装配示意 图。 具体实施方式
剧烈的随机振动是捷联惯性导航系统在运行中面临的主要力学环境， 振动引起系统性能不稳定或电子元器件损坏， 对系统稳定性影响极大。 为 了减小运载体剧烈随机震动引起电子元器件损坏或惯性测量单元性能不稳 定， 除了强化各传感器电路板之间的连接刚度以外， 还要以减振器为阻尼 介质， 将惯性测量单元弹性联结到运载体上， 以获得满意的减震效果。 减 振模式的选取不仅影响着惯导系统的减振性能， 而且也影响着系统的测量 精度， 历来是惯导系统结构设计的重要环节。 本发明从改良传感支架设计 和合理化减振力学结构两个方面着手， 提高微型惯性测量系统的性能。
传感支架是安装陀螺仪和测控电路板及连接线的关键部件， 工作时经 受各种剧烈振动， 其中支架上陀螺仪安装面的相对振幅最大， 其结构的动 态性能将影响到陀螺仪工作的可靠性和精确性， 需要具备一定的静强度、 抗振强度和疲劳寿命。 工艺方面， 要求支架安装方便， 便于加工制造。 合 理设计支架结构， 改善结构的刚度和阻尼特性， 使结构固有频率必须远离 陀螺仪抖动器工作振动频率， 使陀螺仪安装面的相对振幅为最小。 改进支 架设计不能宥于传统思维， 采用大幅度增加壁厚的方法来提高刚度和提高 结构固有频率； 而应通过改善结构的材料、 外形、 结合面等结构设计， 提 高支架的结构刚度和阻尼。 而且， 还要从整体出发， 处理好支架和减振装 置相互制约关系， 还要考虑测控电路板在支架上的安装位置和线路走向。
由前面的陈述可以看出， 为了克服现有惯性测量系统的上述技术缺陷， 本发明采取的技术措施是： 从改进惯性测量系统的力学结构入手， 提供一 个大幅度缩小体积、 三向等刚度减振结构的微型惯性测量系统， 克服三向 刚度不等、 共振激励、 以及产生扭转振动等缺陷， 对捷联惯性导航系统造 成的不良影响。
本发明的一个优选实施例如图 4至图 10所示， 这种微型惯性测量系统 中包括传感组件 12、 减振单元、 上壳 16、 下盖 18等部件， 其中： 传感组件 12 由传感支架 121、 惯性传感器 122、 柔性测控电路板 123 组成： 本实施例中， 传感支架 121 是各平面刻有凹槽， 符合一定比重和刚 度要求的正方体刚性支架。 惯性传感器 122包括陀螺仪和加速度计，其中包括三个陀螺仪和一个加 速度计， 它们悍接在柔性测控电路板 123上。
柔性测控电路板 123应当包含传感器信号预处理功能，其中至少包含抗 混叠电路和 A/D转换电路； 电路板基和连接导线采用柔韧材质， 以耐受 90 ° 的弯折； 柔性测控电路板的外形应当与传感支架的平面展开图形全等， 当它沿着传感支架棱边作 90 ° 弯折后， 整张柔性测控电路板能够完整、 平 顺地覆盖传感支架每一个平面。
具体实施时， 抗混叠电路、 A/D 转换电路、 三个陀螺仪、 一个加速度 计， 共六个电路模块分别设于六块柔性测控电路板上。 每一块柔性测控电 路板上的电路元件分别嵌装于其所在传感支架表面的凹槽中。
上壳 16与下盖 18构成的内腔形状应与传感组件 12的外形相似且空间 略大， 使得壳体各内壁与传感组件对应平面之间留有基本相等的空间， 以 安装内减振单元 14。 内减振器由若干具有适当阻尼特性的内减振单元构成单元 { ，^ 14 组 成， 它们安装在上壳 16内壁 S与传感组件 12的 6个平面之间， 根据运载 体不同振动特性决定其数量， 最多可达 6个。 将传感组件悬挂在壳体内腔 中心， 内减振单元的形变力轴相互正交， 内减振器的弹性中心点 P与传感 组件质心 m重合， 以均衡吸收并消耗来自运载体的强迫振动。 具体实施时， 其中的减振单元由具有一定阻尼作用的弹性材料构成， 可以采用但不限于 弹簧、 橡胶垫、 硅胶、 海绵、 或者其他减振材料。 本发明在一个较佳的实施例中，传感支架选用具有一定比重和刚度的金 属或者非金属材料， 整体加工成一个正方形的传感支架 121， 整体加工而非 组装是为保证支架本身具有足够的刚性， 以降低刚性不足和各向异性的测 量误差参见图 5;
图 6是本发明一个较佳实施例中柔性测控电路板 123平面展开和元器件 布置示意图。 柔性测控电路板 123 的电路板基和连接导线采用柔韧材质， 可以耐受 90° 的弯折；其外形设计成与传感支架外平面展开图全等的形状， 因而具有 6个展开平面。 传感器和其它电子元件悍接在 6个展开平面正面 的适当位置上。
图 7是本发明一个较佳实施例中传感组件构成示意图。柔性测控电路板 123正面悍有惯性传感器 122和其它电子元件。用柔性测控电路板的正面贴 附传感支架 121， 沿着传感支架棱边作 90° 弯折， 将各传感器或电子元件 嵌入传感支架各平面的凹槽之内以后， 整张柔性测控电路板背面朝外， 将 传感支架连同传感、 电子元件包络起来， 并且完整、 平顺地覆盖传感支架 每一个平面。
本发明在设计捷联惯性导航减振系统时，把避免或减小振动耦合作为首 要考虑因素。 如果系统力学结构安排不合理， 系统六自由度上的振动互相 耦合， 产生线振动与角振动交叉激振， 致使惯性测量系统的检测数据内含 有强烈的自身交叉激振信息， 将给系统引入伪运动信号， 严重影响惯导系 统的测量精度。 为了减小减振器对系统角运动测量的干扰， 减振系统的角 振动频率应尽可能远离惯导系统的测量带宽。 在宽带随机振动条件下， 减 振频率越低减振效率越高。
图 8是本发明一个较佳实施例所采用的上壳 16设计，它与下盖 18构成 正方形的内腔为示意清楚图中省略下盖，是容纳传感组件 12与减振单元 14 的场所。 上壳 16与下盖 18构成的壳体内腔的形状， 被设计成与传感组件 12外形相同的正方形， 且比传感组件外形略大。这样设计使得上壳 16与下 盖 18构成的 6个壳体内壁， 与所对应的传感组件 6个外平面之间， 留有形 状和大小基本相同的空间；当把外形基本相同内减振单元 14全部安装其中， 形成内减振器总成之后， 产生了比较好的减振效果。
图 9是本发明一个较佳实施例中采用的全部内减振单元 14构成内减振 器总成后， 与传感组件的位置关系示意图。 为了有效衰减或完全吸收对于 来自前后、 左右、 上下 6个自由度的对传感组件 12的强迫振动， 本实施例 采用 6个内减振单元 14， 即 6个形状相同的减振垫， 安装在上壳 16内壁与 传感组件 12之间， 将传感组件悬挂在壳体内腔中心位置， 且各内减振单元 的形变力轴相互正交， 以均衡吸收并消耗来自运载体的强迫振动。
图 10是本发明一个较佳实施例中关于微型惯性测量系统 2. 1的完整装 配示意图。 由于实行了上述一系列技术措施， 保证了减振器的固有频率、 减振系数、 减振效率、 机械强度等符合系统的抗冲击及振动要求； 使得该 微型惯性测量系统的弹性坐标系、 惯性坐标系、 和求解坐标系三个坐标系， 处于各对应坐标轴相互平行、 系统质心与减振装置的弹性中心重合的最佳 状态， 达到各自由度振动间具有较高的去耦效应， 以及各固有频率相互接 近， 获得较窄频率分布的技术效果。
本发明的微型惯性测量系统可用于无人机等自动驾驶飞行器、船只、水 下自动探测设备， 或是各种车辆、 机器人等。 除了以上实施例以外， 本发 明还可以有其他实施方式， 例如： （1 ) 其中的壳体结构并不限于上壳与下 盖配合的结构， 也可是下壳与上盖配合， 或者是中间壳体与上下盖配合; ( 2 )可将柔性测控电路板上的六个功能模块全部或部分作集成处理， 使得 柔性测控电路板的块数可减为六块以下， 此时传感支架表面设置的凹槽数 量也可相应减少； （3)支架也可以是长方体结构， 当然此时电路板的结构需 作相应改动。 可见， 相关的等同替换技术方案， 均落在本发明要求的保护 范围。

Claims

权 利 要 求
1、 一种微型惯性测量系统， 包括壳体、 传感组件以及减振器； 其特征 在于， 所述传感组件包括刚性传感支架、 装于所述传感支架表面的测控电 路板、 以及设于所述测控电路板上的惯性传感器， 所述惯性传感器包括陀 螺仪和加速度计； 所述传感组件装于所述壳体内； 所述减振器装于所述壳 体内并设于所述传感组件与壳体内壁之间的空隙中。
2、 根据权利要求 1所述的微型惯性测量系统， 其特征在于， 所述传感 支架是正方体刚性支架， 在它的至少一个表面刻有凹槽； 所述测控电路板 是柔性测控电路板； 所述柔性测控电路板上至少有一部分电路元件嵌于所 述至少一个表面的凹槽中。
3、 根据权利要求 2所述的微型惯性测量系统， 其特征在于， 在所述传 感支架的六个表面均刻有凹槽； 所述柔性测控电路板共有六块、 并分别覆 盖于所述传感支架的六个表面； 每一块所述柔性测控电路板上的电路元件 分别嵌装于其所在传感支架表面的凹槽中， 使得所述柔性测控电路板平顺 地覆盖于所述传感支架的每一个表面。
4、 根据权利要求 3所述的微型惯性测量系统， 其特征在于， 所述传感 组件中还包括设于所述柔性测控电路板上的抗混叠电路和 A/D转换电路； 所述惯性传感器包括三个陀螺仪和一个加速度计； 所述共六个电路模块分 别设于所述六块柔性测控电路板上。
5、 根据权利要求 4所述的微型惯性测量系统， 其特征在于， 所述六块 柔性测控电路板为一体结构， 并沿所述传感支架的棱边作 90° 弯折后完整 覆盖于传感支架的各个表面。
6、 根据权利要求 3所述的微型惯性测量系统， 其特征在于， 所述减振 器中包括至少两个减振单元， 分别设于所述传感组件的其中一个表面与所 述壳体内壁之间的空隙中。
7、 根据权利要求 4所述的微型惯性测量系统， 其特征在于， 所述减振 器中包括六个减振单元。
8、 根据权利要求 7所述的微型惯性测量系统， 其特征在于， 所述传感 组件被所述六个减振单元悬挂于所述壳体的内腔中心， 且所述减振器的弹 性中心点 P与所述传感组件的质心 m重合。
9、根据权利要求 1 -8中任一项所述的微型惯性测量系统，其特征在于， 所述壳体包括下部开口的上壳、 以及装于所述开口处的下盖。
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