CN103149030B - Based on the engine plume data acquisition method in-orbit of gyro data - Google Patents

Abstract

The invention discloses a kind of engine plume data acquisition method in-orbit based on gyro data, comprise the following steps: set up the normal three-axis stabilization attitude of spacecraft, the topworks adopting momenttum wheel to control as three axles, magnetic torquer unload, and the attitude sensors such as gyro and infrared or star sensor determine appearance; The engine determining to participate in test opens control point and jet time length; Prepare before engine operation (as drive well heater, drive latching valve etc.); The data that record gyro exports; Test result analysis, analyzes the interference force and moment of the jet generation of engine according to the mass property of gyro data and spacecraft.Adopt the plume data acquisition that present invention achieves engine in-orbit under high vacuum true environment.

Description

On-orbit engine plume data acquisition method based on gyro data

technical field

The invention belongs to the technical field of spacecraft attitude and orbit control, and relates to an on-orbit engine plume data acquisition technology, in particular to a plume acquisition technology based on gyroscope data.

Background technique

The spacecraft is generally equipped with a variety of engines to achieve attitude control and orbit control, such as attitude adjustment, orbit maneuvering, etc. During the operation of the engine, the jet plume may hit the surface of the spacecraft to generate disturbance forces and moments, which will affect the attitude control of the spacecraft Accuracy or track control accuracy, and even more serious effects, such as loss of control. The research shows that to correctly evaluate the effect of the plume on the spacecraft, it is first necessary to accurately describe the vacuum plume field of the engine. There are two main methods for the research of vacuum plume field: numerical simulation method and experimental research method. The numerical simulation method mainly establishes the corresponding mathematical and physical models for different flow states, and adopts the corresponding computer numerical simulation method. The experimental research method is mainly to simulate the space environment plume test on the ground, and the data acquisition technology of the on-orbit engine plume has not been reported.

The accuracy of the numerical simulation method depends on the correctness of the mathematical physical modeling and the selected numerical analysis method, and needs the support of experimental verification. However, the simulation environment of the ground test is inconsistent with the actual high-vacuum environment of the spacecraft in orbit, and it is difficult to realize the simulation environment of the ground test. Related documents are:

【1】Xiao Zejuan et al., Experimental Research on Space Engine Plume Field, Journal of Aerodynamics, December 2008, Volume 26 (2): 480-485. A ground test system was developed to simulate the plume test at an altitude of 100km.

【2】Cheng Xiaoli et al., Research on Plume Pollution of Satellite Orbit-changing Engine, Shanghai Aerospace Science and Technology, No. 5, 2000: 15-18. To solve the problem of high-altitude pollution when satellite orbit-changing engine is working, starting from the theory of molecular motion, a direct simulation Monte The Carlo (DSMC) method is used to simulate and analyze the axisymmetric plume field.

[3] Tang Zhenyu et al., Analysis of the impact of attitude control engine plume liquid phase pollution on spacecraft, Manned Spaceflight, 2011, No. 4: 54~58. Aiming at the liquid phase pollution of liquid attitude control engine plume, a literature review was carried out. Research, the cause of formation, hazards and research methods of droplet formation are given in detail.

The shortcomings of the above-mentioned relevant literature are: the numerical simulation results only focus on the optimization of the simulation algorithm. Although the simulation results provided by foreign literature are compared, there is a lack of engineering test data support; the ground test method only simulates the space environment within 100km , using measuring equipment such as pressure sensors, and spacecraft generally operate in a high vacuum environment of 300km or more, the space environment is different from that within 100km, and the ground test conditions do not meet the high vacuum environment of spacecraft in orbit.

Contents of the invention

The problem solved by the technology of the present invention is: aiming at the deficiencies of the prior art, a technology for acquiring on-orbit engine plume data based on gyroscope data is provided, which realizes the acquisition of on-orbit engine plume data in a real high-vacuum environment.

The technical solution of the present invention is: a kind of on-orbit engine plume data acquisition method based on gyroscope data, comprises the following steps:

(1) Establish the normal three-axis stable attitude of the spacecraft, use the momentum wheel as the actuator for three-axis control, unload the magnetic torque device, and use the gyroscope and infrared or star sensor attitude sensor to determine the attitude;

(2) Determine the start-up control point and jet duration tj of the engine participating in the test on the spacecraft, start recording the data output by the gyroscope before the engine start-up, until the spacecraft enters a three-axis stable attitude after the jet is completed;

(3) According to the gyro output data recorded in step (2), determine the plume disturbance torque generated by the engine jet.

The plume disturbing moment that described engine jet produces is determined according to the following formula:

T _di =J _i (ω _izt1 -ω _izt0 )/(t1-t0)

Among them: T _di is i (i=x, y, z) axis disturbance torque, dimension: Nm;

J _i is the axis inertia of spacecraft i (i=x, y, z), dimension: kg.m ² ;

ω _izt1 is the time t1 of the jet, the inertial angular velocity of the spacecraft i (i=x, y, z) axis, dimension: rad/s;

ω _izt0 is the moment t0 of the jet, the inertial angular velocity of the spacecraft i (i=x, y, z) axis, dimension: rad/s.

Compared with the prior art, the present invention has the following advantages:

1) The test environment of the present invention is a high-vacuum real environment, the spacecraft operates in a steady state on orbit, and the three-axis attitude control adopts momentum wheel control. Except for the engine jet interference used in the test, there is no interference from other jet forces, which is better than the current technology. The plume analysis environment is more realistic.

2) The present invention conducts analysis based on gyro data, and the high precision of the gyro ensures the high precision of the plume analysis results. The present invention has simple algorithm and logic, and strong engineering realizability.

3) The calculation of the algorithm formula of the present invention is simple, which is beneficial to both ground post-processing and on-orbit autonomous direct processing.

Description of drawings

Fig. 1 is the flow chart of the present invention.

Detailed ways

In the present invention, when the spacecraft is in the normal three-axis steady-state flight mode, the momentum wheel is used as the attitude control actuator, the attitude control adopts the PID control law, the engine start control point and the jet time length are selected, and the gyro data of this process is recorded. The variation of the gyro measurement output and the mass characteristics of the spacecraft are used to obtain the effect of the engine plume on the spacecraft.

For realizing above-mentioned process, the present invention comprises the following steps:

(1) Establish a normal three-axis stable attitude of the spacecraft: the momentum wheel is used as the actuator for three-axis control, and the three-axis attitude control uses momentum wheel PID control, magnetic torque device unloading, gyro and infrared or star sensors and other attitude sensors. Attitude, attitude angle and angular velocity are in a steady state.

(2) Determine the start-up control point and injection time length of the engine participating in the test:

The jet direction of the engine participating in the test should be able to interfere with the stars. The selection of the start-up control point is based on the principle that the entire test is within a measurable and controllable arc.

The length of the injection time is determined by the switching accuracy of the engine controller, the steady-state specific impulse distribution of the engine, the relative error of the gyro measurement, and the test accuracy index. For example, AOCC generally uses 82C54 to control the switch of the engine. When the 82C54 counter counts up, an interrupt is generated. AOCC responds to the interrupt and turns off the thruster. The clock pulse unit of 82C54 is t1 second, so the time error caused is less than t1 second. The specific impulse error of the engine changes with the working time. When the working time of the engine is less than 1.0 seconds, the engine has a specific impulse error. The shorter the working time, the greater the error. When the working time is 0.05 seconds, the dispersion is about 50%. When the engine When the working time is greater than or equal to 1.0 seconds, the spread can be approximately considered as 0. Assuming that the steady-state specific impulse distribution of the engine is n1%, the relative error of gyro measurement is n2%, and the test accuracy index is N% (that is, the test result is better than N%), then (n1+n2+N)%<1 and the injection time The length Δt is greater than t1/(1-(n1+n2+N)/100) seconds, so the experimental error is ΔN=1-t1/Δt-(n1+n2)/100, it can be seen that increasing Δt helps to improve the accuracy, It is recommended that the jet duration be greater than 1.0 seconds.

(3) Preparations before engine operation: Turn on the engine heater one lap before the engine is started, and turn off the heater just before the start (for example, 1 minute before the start); turn on the corresponding self-locking valve of the engine before the start. The specific switching time of the heater is determined according to the usage requirements of the engine (for example, 1 minute).

(4) Start to record the data output by the gyroscope before the engine starts to control, until the spacecraft enters the three-axis steady state after the jet is completed. After the injection is completed, close the corresponding self-locking valve of the engine.

Engine injection: observe the temperature change of the catalytic bed of the engine; observe the output of the corresponding pressure sensor of the engine; monitor the attitude angular velocity, attitude angle and gyro output changes of the spacecraft; if the attitude angle is greater than the threshold (such as 3 degrees), stop immediately (such as Directly send the self-locking valve closing command), and the test is over.

(5) Based on the gyro data and the inertia of the spacecraft, analyze the disturbance torque generated by the engine jet.

According to the measurement output of the gyroscope and the inertia of the spacecraft, the disturbance torque T _di (i=x,y,z) can be obtained:

T _di =J _i (ω _izt1 -ω _izt0 )/(t1-t0)

Among them: T _di is i (i=x, y, z) axis disturbance torque, dimension: Nm.

J _i is the axis inertia of spacecraft i (i=x, y, z), dimension: kg.m ² .

t0 and t1 are time, dimension: s.

ω _izt1 is the inertial angular velocity of the spacecraft i (i=x, y, z) axis at the time t1 of the jet, dimension: rad/s.

ω _izt0 is the moment t0 of the jet, the inertial angular velocity of the spacecraft i (i=x, y, z) axis, dimension: rad/s.

Furthermore, the disturbance force generated by the jet of the engine is determined based on the installation position of the engine, the mass characteristics of the spacecraft and the disturbance moment calculated above. The numerical simulation method for plume analysis can be improved.

Example

The present invention utilizes a three-axis stable satellite on-orbit engine at an orbital height of 915km to test the effect of the engine's high-vacuum plume field on the spacecraft, and uses gyroscope data to analyze the test results, which is an engineering test technology that has never been carried out before. . Assuming that t1=2ms, n1=0.1%, n2=0.8%, N=98%, according to the above step (2), it is required that the injection time is greater than t1/(1-(n1+n2+N)%)=2 seconds, take The jet duration is 5.0 seconds, and the other main steps are as follows:

a. Establish the normal three-axis stable attitude of the spacecraft, inject the engine start control point and the length of the injection time.

b. Prepare the engine before working, see step (3) above for details.

c. From the engine start control point to the engine autonomous injection, and record the data of the above step (4).

d. By analyzing the recorded data, the inertial angular velocity output by the gyro in the Z direction before and after the engine jet is ω _izt0 =0.0rad/s, ω _izt1 =8.6651e-4rad/s, and the other directions remain unchanged, while the inertia J _z =2943.7kg. m ² , then the disturbance torque T _dz =0.51015Nm. In this way, the test accuracy can be obtained as 99.06%, which is better than the 98% required by the index.

Parts not described in detail in the present invention belong to the common knowledge of those skilled in the art.

Claims (2)

1., based on an engine plume data acquisition method in-orbit for gyro data, it is characterized in that, comprise the following steps:

(1) set up the normal three-axis stabilization attitude of spacecraft, the topworks adopting momenttum wheel to control as three axles, magnetic torquer unloading, employing gyro and infrared posture sensor or gyro and star sensor determine appearance;

(2) determine that engine spacecraft participating in test opens control point and jet time length tj, engine starts the data recording gyro output before opening control, after completing until jet, spacecraft enters three-axis stabilization attitude; Described jet time length is greater than t/ (1-(n1+n2+N)/100) second, t is the time clock unit controlling tail-off, the steady-state specific impulse of engine scatters as n1%, and gyro to measure relative error is n2%, and test accuracy index is N%;

(3) gyro recorded according to step (2) exports data, determines the plume disturbance torque of the jet generation of engine.

2. a kind of engine plume data acquisition method in-orbit based on gyro data according to claim 1, is characterized in that: the plume disturbance torque of the jet generation of described engine is determined according to following formula:

T _di＝J _i(ω _izt1-ω _izt0)/(t1-t0)

Wherein: T _difor i (i=x, y, z) axle disturbance torque, dimension: N.m;

J _ifor spacecraft i (i=x, y, z) axle inertia, dimension: kg.m ²;

ω _izt1for the jet t1 moment, spacecraft i (i=x, y, z) axle inertia angular velocity, dimension: rad/s;

ω _izt0for the jet t0 moment, spacecraft i (i=x, y, z) axle inertia angular velocity, dimension: rad/s.