CN1644456A - Miniature assembled gesture measuring system for mini-satellite - Google Patents

Abstract

A miniature combined attitude measuring system for small satellite is composed of a combined inertial-magnetic measuring unit, a signal processing circuit and a satellite carried computer for storing data and program. Said measuring unit has 3 sensor sets based on microelectromachinal system. Each sensor set has microgyroscope, microaccelerometer and micromagnetometer. Its advantages are high real-time performance, high reliability and high anti-interference power.

Description

The miniature assembled gesture measuring system that is used for microsatellite

Technical field

The invention belongs to minitype spacecraft and space telemetry and control technology field.

Background technology

Microsatellite is developed rapidly in recent years, it have in light weight, volume is little, cost is low, the lead time is short, risk is little and advantage such as advanced technology, having broad application prospects at aspects such as scientific experiment, communication, meteorology, environment measurings, is the important development direction of satellite technology.As the important component part of small satellite technology, the attitude observation and control technology that research and development have high pointing accuracy and degree of stability has crucial meaning.

The sensor that the attitude measurement system of present microsatellite mainly uses comprises sun sensor, star sensor, infrared earth sensor, magnetic sensor etc., wherein, the precision of sun sensor, magnetic sensor and infrared earth sensor is relatively limited, and the precision of star sensor is higher.But response time is longer relatively in their work, and has problem such as visual field.Simultaneously, they receive external influence easily, therefore short term failure may occur in mobile process such as Satellite Orbit Maneuver.In addition, the technology such as GPS that are used for track control at present rely on navigation magnitude external information more, receive influence also just easilier.Therefore, it is significant that research has the microsatellite autonomous navigation technology of better performance.This system not only should be able to not rely on the autonomous operation of ground system, also should have the ability and the real-time of good anti-ectocine.In addition, the developing trend microminiaturization of satellite at present also requires the autonomous navigation system that adopts miniaturization, cheapness and satisfy mission requirements day by day.

Summary of the invention

In order to improve the real-time of microsatellite at the rail attitude measurement, reliability and anti-ectocine ability, the invention provides a kind of miniature assembled gesture measuring system that is used for microsatellite, it is characterized in that: described miniature assembled gesture measuring system comprises inertia/magnetic measurement in a closed series unit, computing machine on the star of signal processing circuit and storage data processor, described inertia/magnetic measurement in a closed series unit is made up of based on the sensor groups of MEMS the cover of three on three normal surfaces that are installed in six matrixes of high precision respectively, each sensor groups includes little gyro, micro-acceleration gauge and little magnetometer, the mouth of described each sensor groups all links to each other with described signal processing circuit, the output signal of each sensor groups is after signal processing circuit is handled, be sent to computing machine on the star, and by being stored in the execution of the data processor in computing machine following steps on the star:

The definition error quaternion is to rotate required quaternion by the attitude estimated valve to actual value, at first utilize microsatellite three axis angular rates and the linear acceleration that records, make up state vector and error vector by described error quaternion, and then obtain system state equation; Simultaneously, the three spool components of magnetic-field intensity under the celestial body system of axes by real-time acquisition make up the measurement equation, then, utilize described system state equation and measure equation, the quantitative data input extended Kalman filter that little gyro, micro-acceleration gauge and little magnetometer are recorded, carry out the attitude of described miniature assembled gesture measuring system and estimate, obtain the real-time attitude information of microsatellite.

In the present invention, described signal processing circuit is integrated in the bottom of six matrixes of high precision, comprise filter circuit, signal amplification circuit, A/D change-over circuit and serial interface, finish signal filtering, amplification and A/D conversion successively, and little gyro, micro-acceleration gauge and the unification of little magnetometer output signal that will be after treatment be transformed into-5V～+ the 5V scope, the serial interface in the signal processing circuit is connected with computing machine on the described star by the space flight connector that is fixed on the described matrix.

The invention has the advantages that:

1) attitude that inertial technology is used for microsatellite is determined, and it is combined with little magnetometer, by the data processing method of design-calculated integrated attitude determination, has realized the miniature assembled gesture measuring system that the two combines.This system can realize that complete autonomous type spatial attitude is determined and navigation, both can not rely on extraneous information, has overcome the drifting problem that pure inertia measurement is brought again greatly.

2) utilize miniature assembled gesture measuring system to carry out determining of microsatellite attitude, effectively improved attitude measurement accuracy, reduced volume, weight and the cost of posture control system.

3) of the present inventionly be designed to realize that the integration measuring system of satellite attitude based on MEMS lays the foundation.

Description of drawings

Fig. 1 is the structured flowchart that is used for the miniature assembled gesture measuring system of microsatellite.

Fig. 2 is the theory structure scheme drawing of inertia/magnetic measurement in a closed series unit.

Fig. 3 is the flow chart of data processing figure of miniature assembled gesture measuring system.

Fig. 4 is the output calibration array mode schematic diagram of miniature assembled gesture measuring system.

Fig. 5 is the feedback compensation array mode schematic diagram of miniature assembled gesture measuring system.

The specific embodiment

Further specify the present invention below in conjunction with accompanying drawing.

As shown in Figure 1, the invention provides a kind of miniature assembled gesture measuring system that is used for microsatellite, comprise computing machine on the star of inertia/magnetic measurement in a closed series unit, signal processing circuit and storage data processor, wherein, the structure of inertia/magnetic measurement in a closed series unit as shown in Figure 2, be made up of based on the sensor groups of MEMS the cover of three on three normal surfaces that are installed in six matrixes of high precision respectively, each sensor groups includes little gyro 1, micro-acceleration gauge 2 and little magnetometer 3.The signal of described each sensor output is analog voltage signal, respectively corresponding celestial body at that time around the cireular frequency of three orthogonal axes, along the linear acceleration and the magnetic-field intensity of three orthogonal directionss.The signal wire (SW) (totally 9 tunnel) of three cover sensor groups is connected on the signal processing circuit 4 by connector separately.Signal processing circuit is integrated in the bottom of six matrixes of high precision, comprise filter circuit, signal amplification circuit, A/D change-over circuit and serial interface, finish signal filtering, amplification and A/D conversion successively, and little gyro, micro-acceleration gauge and the unification of little magnetometer output signal that will be after treatment be transformed into-5V～+ the 5V scope, the serial interface in the signal processing circuit is connected with computing machine on the described star by the space flight connector 5 that is fixed on the described matrix.

From data handing, native system can adopt two kinds of array modes, i.e. output calibration array mode and feedback compensation array mode, groundwork such as Fig. 4, shown in Figure 5.In the output calibration array mode, direct and the MIMU (Micro Inertial Measurement Unit) (MIMU of the result of filtering, comprise above-mentioned 3 little gyros and 3 micro-acceleration gauges) output combine, the compensation output error, and do not influence the mode of operation of system, but owing to do not adopt feedback, and responsive to the filter model error, so require to use more accurate model.And the feedback compensation array mode utilizes filter to estimate error earlier, then feedback compensation MIMU.

At the characteristics of microsatellite, more than the sensor output signal handled through signal processing circuit on star, finish in the computing machine and resolve, detailed process is as follows:

At first, draw the attitude math modeling of group system according to Satellite Attitude Movement model and three axis magnetometer measurement model:

1) kinematics model

$\stackrel{\·}{Q}=\frac{1}{2}{\widetilde{\mathrm{\ω}}}_{\mathrm{bi}}\⊗Q$

$\left[\begin{array}{c}{\stackrel{\·}{q}}_1\\ {\stackrel{\·}{q}}_1\\ {\stackrel{\·}{q}}_1\\ {\stackrel{\·}{q}}_1\end{array}\right]=\left[\begin{array}{cccc}0& {\mathrm{\ω}}_{\mathrm{biz}}& {-\mathrm{\ω}}_{\mathrm{biy}}& {\mathrm{\ω}}_{\mathrm{bix}}\\ -{\mathrm{\ω}}_{\mathrm{biz}}& 0& {\mathrm{\ω}}_{\mathrm{bix}}& {\mathrm{\ω}}_{\mathrm{biy}}\\ {\mathrm{\ω}}_{\mathrm{biy}}& -{\mathrm{\ω}}_{\mathrm{bix}}& 0& {\mathrm{\ω}}_{\mathrm{bix}}\\ -{\mathrm{\ω}}_{\mathrm{bix}}& -{\mathrm{\ω}}_{\mathrm{biy}}& {-\mathrm{\ω}}_{\mathrm{bix}}& 0\end{array}\right]\·\left[\begin{array}{c}{q}_1\\ q_2\\ q_3\\ q_4\end{array}\right]---\left(1\right)$

Q=[q in the formula ₁q ₂q ₃q ₄] ^TFor being tied to the attitude quaternion of carrier coordinate system from inertial coordinate, ${\widetilde{\mathrm{\ω}}}_{\mathrm{bi}}={\left[\begin{array}{ccc}{\mathrm{\ω}}_{\mathrm{bix}}& {\mathrm{\ω}}_{\mathrm{biy}}& {\mathrm{\ω}}_{\mathrm{biz}}\end{array}\right]}^T$ Be the angular rate of carrier with respect to inertial coordinates system.

2) little gyro output model

ω _bi＝u-b-η ₁ (2)

In the formula: ω _BiBe the carrier track angular rate under the perfect condition, u is actual little gyro output, and b is gyro wander, η ₁Gaussian white noise error for gyro wander.

E[η ₁(t)]＝0

$E\left[{\mathrm{\η}}_1\left(t\right){\mathrm{\η}}_1^T\left(t^{\′}\right)\right]=Q_1\left(t\right)\mathrm{\δ}(t-t^{\′})$

Wherein, t is the time.

Again since gyro wander b be static state, η ₂It is the random walk noise of gyro wander.

$\frac{d}{\mathrm{dt}}b={\mathrm{\η}}_2---\left(3\right)$

The feature of this random process satisfies:

E[η ₂(t)]＝0

$E\left[{\mathrm{\η}}_2\left(t\right){\mathrm{\η}}_2^T\left(t^{\′}\right)\right]{=Q}_2\left(t\right)\mathrm{\δ}(t-t^{\′})$

3) magnetometer survey model

$B_b^b=C_i^b\·B_b^i---\left(4\right)$

In the formula: B _b ^bBe the geomagnetic fieldvector under the carrier coordinate system of magnetometer survey, B _b ⁱGround magnetic vector under the inertial system that the international geomagnetic model (IGRF) of serving as reasons calculates, C _i ^bInertial coordinate is tied to the attitude matrix of carrier coordinate system.

Then, carry out attitude algorithm by the integrated attitude determination filter, idiographic flow as shown in Figure 3.

The definition error quaternion is to rotate required quaternion by the attitude estimated valve to actual value:

$\mathrm{\δ}\stackrel{\‾}{q}=\stackrel{\‾}{q}\⊗{\widehat{\stackrel{\‾}{q}}}^{-1}\≈\left[\begin{array}{c}\mathrm{\δq}\\ 1\end{array}\right]---\left(5\right)$

Q=[q q ₄] ^TBe true attitude quaternion,

Estimate quaternion for attitude, δ q is an error quaternion.

According to the said system model, establish state vector and evaluated error and be

State vector: $x=\left[\begin{array}{c}\stackrel{\‾}{q}\\ b\end{array}\right]---\left(6\right)$

Error vector: $\mathrm{\Δx}=\left[\begin{array}{c}\mathrm{\δq}\\ b-\hat{b}\end{array}\right]=\left[\begin{array}{c}\mathrm{\δq}\\ \mathrm{\Δb}\end{array}\right]---\left(7\right)$

By equation of satellite motion

$\frac{d\stackrel{\‾}{q}}{\mathrm{dt}}=\frac{1}{2}\stackrel{\‾}{\mathrm{\ω}}\⊗\stackrel{\‾}{q}---\left(8\right)$

$\frac{d}{\mathrm{dt}}\widehat{\stackrel{\‾}{q}}=\frac{1}{2}\widehat{\stackrel{\‾}{\mathrm{\ω}}}\⊗\widehat{\stackrel{\‾}{q}}---\left(9\right)$

: $\frac{d}{\mathrm{dt}}\mathrm{\δ}\stackrel{\‾}{q}=\frac{1}{2}[\widehat{\stackrel{\‾}{\mathrm{\ω}}}\⊗\mathrm{\δ}\stackrel{\‾}{q}-\mathrm{\δ}\stackrel{\‾}{q}\⊗\widehat{\stackrel{\‾}{\mathrm{\ω}}}]+\frac{1}{2}\mathrm{\δ}\stackrel{\‾}{\mathrm{\ω}}\⊗\mathrm{\δ}\stackrel{\‾}{q}---\left(10\right)$

Wherein: $\mathrm{\δ}\stackrel{\‾}{\mathrm{\ω}}=\left[\begin{array}{c}\mathrm{\ω}-\widehat{\mathrm{\ω}}\\ 0\end{array}\right]$

Ignore second order term, thereby obtain system state equation:

$\frac{d}{\mathrm{dt}}\left[\begin{array}{c}\mathrm{\&delta;q}\\ \mathrm{\&Delta;b}\end{array}\right]=\left[\begin{array}{cc}[\widehat{\mathrm{\&omega;}}\left(t\right)\&times;]& -\frac{1}{2}{\mathrm{<figure-callout\; id="3"\; label="I"\; filenames="A20051001122200082.png"\; state="\{\{state\}\}">I</figure-callout>}}_{3\&times;3}\\ {\mathrm{<figure-callout\; id="3"\; label="O"\; filenames="A20051001122200082.png"\; state="\{\{state\}\}">O</figure-callout>}}_{3\&times;3}& {\mathrm{<figure-callout\; id="3"\; label="O"\; filenames="A20051001122200082.png"\; state="\{\{state\}\}">O</figure-callout>}}_{3\&times;3}\end{array}\right]\&CenterDot;\left[\begin{array}{c}\mathrm{\&delta;q}\\ \mathrm{\&Delta;b}\end{array}\right]+\left[\begin{array}{cc}-\frac{1}{2}{\mathrm{<figure-callout\; id="3"\; label="I"\; filenames="A20051001122200082.png"\; state="\{\{state\}\}">I</figure-callout>}}_{3\&times;3}& {\mathrm{<figure-callout\; id="3"\; label="O"\; filenames="A20051001122200082.png"\; state="\{\{state\}\}">O</figure-callout>}}_{3\&times;3}\\ {\mathrm{<figure-callout\; id="3"\; label="O"\; filenames="A20051001122200082.png"\; state="\{\{state\}\}">O</figure-callout>}}_{3\&times;3}& {\mathrm{<figure-callout\; id="3"\; label="I"\; filenames="A20051001122200082.png"\; state="\{\{state\}\}">I</figure-callout>}}_{3\&times;3}\end{array}\right]\&CenterDot;\left[\begin{array}{c}{\mathrm{\&eta;}}_1\\ {\mathrm{\&eta;}}_2\end{array}\right]---\left(11\right)$

Wherein, Be

The skew symmetry battle array.

Little magnetometer sensitive axes is parallel with three of celestial body system of axes respectively, its measuring amount B _bFor working as the three spool components of geomagnetic field intensity under the celestial body system of axes, measure equation:

${\hat{B}}_m=C\left(\widehat{\stackrel{\&OverBar;}{q}}\right)B_i---\left(12\right)$

In the formula, Be B _bEstimated valve,

Serve as reasons and estimate the attitude matrix that quaternion obtains, B _iBe the magnetic-field intensity that calculates by IGRF.

Vector is measured in definition

$Z=B_b-{\hat{B}}_b=(C\left(\mathrm{\&delta;}\stackrel{\&OverBar;}{q}\right)-I){\hat{B}}_b\&ap;2[\mathrm{\&delta;q}\&times;]{\hat{B}}_b+v---\left(13\right)$

[δ q *] is the skew symmetry battle array of δ q, and ν is for measuring noise.

According to above-mentioned system state equation and measurement equation, the data of utilizing extended Kalman filter that little gyro, micro-acceleration gauge and little magnetometer are recorded on star in the computer system are carried out integrated navigation and are resolved, thereby obtain the real-time attitude information of microsatellite.The further coordinate transformation of above data computing machine on star is handled, and can obtain final in real time at the rail attitude information, for attitude control provides foundation.

The present invention proposes complete assembled gesture measuring system, can effectively improve the reliability and the precision of little measuring system of satellite attitude and reduce its volume and weight based on little gyro, micro-acceleration gauge and the little magnetometer of MEMS.

On the one hand, compare with the single magnetometer attitude determination system of tradition, do not need the recursion kinetics equation, filtering equations is simplified greatly, and calculated amount also significantly descends; And original dynam recursion precision is because influenced by model self precision very big, influenced attitude measurement accuracy in the middle of invisible, and the introducing of gyro information has improved survey precision; On the other hand, estimate gyro wander by filtering algorithm, gyro wander is estimated to feed back to MIMU (Micro Inertial Measurement Unit), just can obtain more accurate angular velocity measurement value, thereby can reduce the deficiency that the MIMU (Micro Inertial Measurement Unit) error accumulates in time, MIMU (Micro Inertial Measurement Unit) has also improved antijamming capability as complete self-aid navigation system simultaneously.In addition, the inertia integrated attitude determination system commonly used with guidance system on the present ground compares, and system of the present invention has effectively overcome the influence of gyro wander equal error, has solved inertia and has decided the problem that appearance can not guarantee long-term accuracy.

The present invention is suitable for space flight test, particularly the application scenarios such as posture control system of microsatellite.Compare with other attitude sensor spares commonly used such as sun sensor, star sensors, it is real-time, is not subject to ectocine, has a good application prospect.

Claims (2)

1. the miniature assembled gesture measuring system that is used for microsatellite, it is characterized in that: described miniature assembled gesture measuring system comprises inertia/magnetic measurement in a closed series unit, computing machine on the star of signal processing circuit and storage data processor, described inertia/magnetic measurement in a closed series unit is made up of based on the sensor groups of MEMS the cover of three on three normal surfaces that are installed in six matrixes of high precision respectively, each sensor groups includes little gyro, micro-acceleration gauge and little magnetometer, the mouth of described each sensor groups all links to each other with described signal processing circuit, the output signal of each sensor groups is after signal processing circuit is handled, be sent to computing machine on the star, and by being stored in the execution of the data processor in computing machine following steps on the star:

The definition error quaternion is to rotate required quaternion by the attitude estimated valve to actual value, at first utilize microsatellite three axis angular rates and the linear acceleration that records, make up state vector and error vector by described error quaternion, and then obtain system state equation; Simultaneously, the three spool components of magnetic-field intensity under the celestial body system of axes by real-time acquisition make up the measurement equation, then, utilize described system state equation and measure equation, the quantitative data input extended Kalman filter that little gyro, micro-acceleration gauge and little magnetometer are recorded, carry out the attitude of described miniature assembled gesture measuring system and estimate, obtain the real-time attitude information of microsatellite.

2. the miniature assembled gesture measuring system that is used for microsatellite according to claim 1, it is characterized in that: described signal processing circuit is integrated in the bottom of six matrixes of high precision, comprise filter circuit, signal amplification circuit, A/D change-over circuit and serial interface, finish signal filtering, amplification and A/D conversion successively, and little gyro, micro-acceleration gauge and the unification of little magnetometer output signal that will be after treatment be transformed into-5V～+ the 5V scope, the serial interface in the signal processing circuit is connected with computing machine on the described star by the space flight connector that is fixed on the described matrix.

Images