阅读说明：本技术 一种轨道转移过程中的姿态确定方法 (Attitude determination method in track transfer process ) 是由 王焕杰 张肖 徐晨 张晓彤 陈浩 刘禹 于 2019-08-26 设计创作，主要内容包括：本发明公开了一种轨道转移过程中的姿态确定方法,该方法包括以下步骤：步骤1：以轨道转移过程中给出的期望推力方向及当前的太阳矢量方向为输入,建立期望姿态参考坐标系,并求得期望姿态参考坐标系到飞行器本体坐标系的姿态转换矩阵；步骤2：当太阳矢量方向与期望推力方向夹角达到平行阈值时,对步骤1建立的期望姿态参考坐标系进行避奇异处理。本发明考虑飞行器姿态在保证推力方向的基础上进行太阳帆板对日指向的约束,建立了目标姿态参考坐标系,实现了姿态控制量的解算。同时充分考虑整个轨道转移过程中太阳方向矢量与飞行器本体的相对姿态关系,对姿态确定策略进行了避奇异处理,保证了姿态确定测量始终能够进行有效数据输出。(The invention discloses a method for determining the posture in the process of rail transfer, which comprises the following steps: step 1: establishing an expected attitude reference coordinate system by taking an expected thrust direction given in the process of orbit transfer and the current sun vector direction as input, and solving an attitude transformation matrix from the expected attitude reference coordinate system to an aircraft body coordinate system; step 2: and (3) when the included angle between the sun vector direction and the expected thrust direction reaches a parallel threshold value, performing singularity avoidance processing on the expected attitude reference coordinate system established in the step (1). According to the method, the solar sailboard is restrained to the sun direction on the basis of ensuring the thrust direction by considering the aircraft attitude, a target attitude reference coordinate system is established, and the attitude control quantity is resolved. Meanwhile, the relative attitude relationship between the sun direction vector and the aircraft body in the whole orbit transfer process is fully considered, singularity avoidance processing is carried out on the attitude determination strategy, and effective data output can be always carried out on attitude determination measurement.)

1. A method for determining the attitude in the process of orbit transfer is characterized by comprising the following steps:

Step 1: establishing an expected attitude reference coordinate system by taking an expected thrust direction given in the process of orbit transfer and the current sun vector direction as input, and solving an attitude transformation matrix from the expected attitude reference coordinate system to an aircraft body coordinate system;

Step 2: and (3) when the included angle between the sun vector direction and the expected thrust direction reaches a parallel threshold value, performing singularity avoidance processing on the expected attitude reference coordinate system established in the step (1).

2. The attitude determination method in the process of orbit transfer according to claim 1, wherein the step 1 specifically comprises the following steps: real-time giving of expected thrust direction by track design in track transfer processMeanwhile, the sun vector direction is obtained by utilizing the satellite-borne sun sensorsuppose a solar panel is + Y along the aircraft body_bmounted on a shaft, able to wind around + Y_bone-dimensional rotation is carried out, and only the posture of the aircraft body is controlled, so that the sun vector direction and the + Y of the aircraft body_bThe included angle of the axes is 90 degrees, namely the alignment direction of the solar sailboard is consistent with the sun vector direction through the one-dimensional rotation of the solar sailboard; plus X of expected thrust direction along aircraft body all the time in orbit transfer process_bDirection; based on the above analysis, a desired pose reference coordinate system O is performed_tX_tY_tZ_tEstablishing;

Setting desired thrust directionAnd the direction of the sun vectorThe description in the coordinate system of the aircraft body is t and s respectively, then:

x_t＝t

y_t＝x_t×s

z_t＝x_t×y_t

So as to obtain a reference coordinate system O of the expected attitude_tX_tY_tZ_tTo the aircraft body coordinate system O_bX_bY_bZ_bAttitude transformation matrix C_bt：

3. The attitude determination method in the process of orbit transfer according to claim 1, wherein the step 2 specifically comprises the following steps: when the included angle between the sun vector direction and the expected thrust direction at the current moment k-1 reaches a set parallel threshold value and the previous moment does not reach the threshold value, y in an expected attitude reference coordinate system at the current moment k_t(k) In the determination, the sun vector direction s (k) at the current moment is not introduced, but s (k-1) at the previous moment is maintained.

Technical Field

The invention relates to a gesture determination technology in orbit transfer, in particular to a gesture determination method in an orbit transfer process.

Background

with the rapid development of aerospace technology, more and more attention is paid to the research in the deep space exploration field. Because the difference between the target track and the initial track entering track is large, in order to guarantee the requirements of subsequent tasks, the fuel consumption needs to be reduced as much as possible in the whole track transfer process, and a track transfer strategy based on an electric thruster is produced. Compared with the traditional chemical thruster, the electric thruster has smaller thrust and can not be ignited only at a far place, and the electric thruster is in an ignition state except a near place section with lower orbital transfer efficiency and track sections such as a ground shadow area. This means that the aircraft attitude needs to be maintained in the fired attitude for a long period of time. In addition, the electric propulsion needs high-power electric energy for supporting, and the aircraft needs to carry out sailboard sun-pointing control.

the current small-thrust orbit transfer research mainly focuses on the aspect of orbit optimization design, so that the attitude determination strategy in the transfer process becomes a problem which needs to be solved urgently at present.

Disclosure of Invention

the invention aims to provide a method for determining the attitude in the track transfer process, which can meet the requirements of thrust direction and solar direction.

In order to achieve the above object, the present invention provides a method for determining an attitude during a track transfer process, which comprises the following steps:

Step 1: establishing an expected attitude reference coordinate system by taking an expected thrust direction given in the process of orbit transfer and the current sun vector direction as input, and solving an attitude transformation matrix from the expected attitude reference coordinate system to an aircraft body coordinate system;

Step 2: and (3) when the included angle between the sun vector direction and the expected thrust direction reaches a parallel threshold value, performing singularity avoidance processing on the expected attitude reference coordinate system established in the step (1).

The attitude determination method in the track transfer process includes the following steps in step 1: real-time giving of expected thrust direction by track design in track transfer processMeanwhile, the sun vector direction is obtained by utilizing the satellite-borne sun sensorSuppose a solar panel is + Y along the aircraft body_bMounted on a shaft, able to wind around + Y_bOne-dimensional rotation is carried out, and only the posture of the aircraft body is controlled, so that the sun vector direction and the + Y of the aircraft body_bThe included angle of the axes is 90 degrees, namely the alignment direction of the solar sailboard is consistent with the sun vector direction through the one-dimensional rotation of the solar sailboard; plus X of expected thrust direction along aircraft body all the time in orbit transfer process_bDirection; based on the above analysis, a desired pose reference coordinate system O is performed_tX_tY_tZ_tEstablishing;

setting desired thrust directionAnd the direction of the sun vectorthe description in the coordinate system of the aircraft body is t and s respectively, then:

x_t＝t

y_t＝x_t×s

z_t＝x_t×y_t

So as to obtain a reference coordinate system O of the expected attitude_tX_tY_tZ_tTo the aircraft body coordinate system O_bX_bY_bZ_bAttitude transformation matrix C_bt：

The attitude determination method in the track transfer process includes the following steps in step 2: when the included angle between the sun vector direction and the expected thrust direction at the current moment k-1 reaches a set parallel threshold value and the previous moment does not reach the threshold value, y in an expected attitude reference coordinate system at the current moment k_t(k) In the determination, the sun vector direction s (k) at the current moment is not introduced, but s (k-1) at the previous moment is maintained.

Compared with the prior art, the invention has the following beneficial effects:

The invention aims to realize the thrust direction protection and the sun-to-sun pointing protection in the process of transferring the continuous thrust track so as to expect the thrust directionAnd direction of sun vectorAnd establishing an expected attitude reference coordinate system for input, and solving an attitude transformation matrix from the expected attitude reference coordinate system to the aircraft body system. Meanwhile, the characteristic that the direction of the sun vector changes slowly in the long-term orbit transfer process is considered, and singularity avoidance processing is carried out on the singularity problem in the attitude determination. Effective input is provided for the attitude control requirements of guaranteeing the thrust direction and the sun direction in the track transfer process.

Drawings

FIG. 1 is a schematic diagram of a desired pose reference frame according to the present invention.

Detailed Description

The invention will be further described by the following specific examples in conjunction with the drawings, which are provided for illustration only and are not intended to limit the scope of the invention.

As shown in fig. 1, the method for determining the attitude in the track transfer process provided by the present invention specifically includes the following steps:

Step 1: and establishing an expected attitude reference coordinate system by taking the expected thrust direction given in the orbit transfer process and the current sun vector direction as input, and giving an attitude conversion matrix from the expected attitude reference coordinate system to the aircraft body coordinate system.

The track design can give the expected thrust direction in real time in the process of transferring the low-thrust trackMeanwhile, the sun vector direction can be obtained by utilizing the satellite-borne sun sensorSuppose the sun sailboard is + Y along the body_bmounted on a shaft, able to wind around + Y_bOne-dimensional rotation is carried out, so long as the posture of the body is controlled, the sun vector direction and the + Y of the body are enabled to be_bThe included angle of the axes is 90 degrees, namely the alignment direction of the solar sailboard is consistent with the sun vector direction through the one-dimensional rotation of the solar sailboard. If the thrust direction always follows the body in the track transfer process plus X_bAnd (4) direction.

Based on the above analysis, a desired pose reference coordinate system O is performed_tX_tY_tZ_tAnd (4) establishing. Setting desired thrust directionAnd the direction of the sun vectorThe description in the coordinate system of the aircraft body is t and s respectively, then

x_t＝t

y_t＝x_t×s

z_t＝x_t×y_t (1)

So as to obtain a reference coordinate system O of the expected attitude_tX_tY_tZ_tTo the aircraft body system O_bX_bY_bZ_bposture ofChange matrix C_bt。

Step 2: considering the continuous change characteristic of the sun vector direction, when the included angle between the sun vector direction and the expected thrust direction reaches a set parallel threshold, the desired reference coordinate system established in the step 1 needs to be subjected to singularity avoidance processing.

During long-term orbit transfer, the direction of the sun vector is continuously and slowly changed, and when the direction of the expected thrust is parallel to the direction of the sun vector, y in the formula (1)_tThe direction of (2) cannot be determined, i.e. a singular problem of pose determination is generated. At the moment, another attitude determination strategy needs to be determined for singularity avoidance processing, the principle is that the thrust direction is preferably ensured to be consistent with the expectation, and then the solar sailboards are made to face the sun as much as possible. In order to avoid the problem that the three-axis attitude angle phase difference obtained when the two attitude determination strategies are switched is large, the following means is adopted to avoid singularity:

Firstly, the parallel condition of the sun vector direction and the expected thrust direction at the current moment is judged. The angle θ (k) between the sun vector direction and the desired thrust direction is arccos (s (k) t (k)). When theta (k) ≦ 1 deg. or theta (k) ≧ 179 deg., and 1 deg. < theta (k-1) <179 deg., establishment of the desired attitude reference coordinate system is as follows

that is, the sun vector direction and the expected thrust direction at the current time k reach the set parallel threshold (1 degree in the text), and when the previous time k-1 does not reach the threshold, the reference coordinate system y of the expected attitude at the current time is carried out_t(k) In the determination, the sun vector direction s (k) at the current moment is not introduced, but s (k-1) at the previous moment is maintained. The present time C thus obtained_bt(k) The three-axis attitude angle obtained by calculation cannot be greatly changed due to singularity. The singularity avoidance cost is that strict sun-to-sun pointing cannot be realized, and the sun-to-sun pointing angle obtained by determining the gesture in the singularity avoidance process does not change along with the direction of the sun vectorAnd the corresponding change occurs, generally speaking, the larger the parallel threshold value design is, the larger the deviation of the solar pointing angle obtained in the singularity avoiding process is. In the actual track continuous transfer process, the singularity avoiding process takes a short time, and the attitude determination caused by the sun vector direction keeping is relatively small.

In conclusion, the attitude determination method in the orbit transfer process provided by the invention can be applied to the long-term orbit transfer task with continuous thrust action. In deep space exploration and other tasks requiring long-time low-thrust orbit transfer, the aircraft is often limited in daily direction besides ensuring a continuous orbital transfer attitude. The method takes the small thrust orbit transfer as a task background, takes the aircraft attitude into consideration to carry out the solar array sun pointing constraint on the basis of ensuring the thrust direction, establishes a target attitude reference coordinate system and realizes the calculation of the attitude control quantity. Meanwhile, the relative attitude relationship between the sun direction vector and the aircraft body in the whole orbit transfer process is fully considered, singularity avoidance processing is carried out on the attitude determination strategy, and effective data output can be always carried out on attitude determination measurement. The attitude determination strategy provided by the invention is simple and feasible, and is easy for engineering application.

While the present invention has been described in detail with reference to the preferred embodiments, it should be understood that the above description should not be taken as limiting the invention. Various modifications and alterations to this invention will become apparent to those skilled in the art upon reading the foregoing description. Accordingly, the scope of the invention should be determined from the following claims.