CN117734969B - A thrust inversion method for electric propulsion device based on single-frame control moment gyro - Google Patents

Abstract

The present invention relates to a thrust inversion method of an electric propulsion device based on a single-frame control moment gyro, and relates to the technical field of electric propulsion of spacecraft. The present invention adopts a single-frame control moment gyro to control the attitude of a satellite, and an electric propulsion device is offsetly installed on the satellite body. A satellite attitude dynamics model using a single-frame control moment gyro as an actuator of an attitude control system is established, and the satellite attitude dynamics equations during the on-orbit ignition of the electric propulsion device and within T seconds before the ignition are subtracted to obtain an analytical expression for the electric thrust torque of the electric propulsion device. On this basis, the telemetry data within T seconds before and during the ignition are fitted using the least squares method to solve the electric thrust torque, and the thrust of the electric propulsion device is obtained through data processing and error analysis. The present invention can accurately calculate the thrust value of the electric propulsion device when both the satellite attitude and the orbit remain stable, and has the advantages of simple calculation method, short calculation time, high calculation accuracy and easy engineering implementation.

Description

A thrust inversion method for electric propulsion device based on single-frame control moment gyro

Technical Field

The invention relates to the technical field of spacecraft control, in particular to a thrust inversion method of an electric propulsion device based on a single-frame control moment gyro.

Background Art

With the rapid development of spacecraft such as micro-satellites, people have more and more demands for space exploration missions, which also places higher and higher requirements on space propulsion devices. Compared with chemical propulsion, electric propulsion technology is an advanced aerospace propulsion technology that efficiently converts electrical energy into kinetic energy. It has the characteristics of low power consumption, high specific impulse, flexible control, etc., and can provide power for attitude control and orbit maintenance of spacecraft. However, since the thrust of the electric propulsion device is relatively small, generally at the millinewton level, high-precision on-orbit measurement of it has always been a key and difficult problem in research. At present, the thrust of the electric propulsion device is generally calculated by the change of the number of spacecraft orbital elements, but there are disadvantages such as complex calculation methods, long time consumption and low accuracy. At the same time, if the electric propulsion device cannot be turned on for a long time during the on-orbit verification, there will be a situation where the change of the spacecraft orbit is very small due to the small thrust, and the thrust measurement cannot guarantee the accuracy requirement. Based on this, the present invention proposes a thrust inversion method of an electric propulsion device based on a single-frame control moment gyro.

Summary of the invention

In order to solve the problem of low on-orbit thrust calculation accuracy of an electric propulsion device, the present invention provides a thrust inversion method for an electric propulsion device based on a single-frame control moment gyro under the condition of a stable satellite attitude.

To achieve the above object, the present invention provides the following technical solutions:

A thrust inversion method for an electric propulsion device based on a single-frame control moment gyro, the method comprising the following steps:

Step 1: Establish the satellite body coordinate system and the frame coordinate system. The satellite is equipped with orthogonal single-frame control moment gyroscopes along the +X, +Y and +Z axis directions to control the satellite attitude. An electric propulsion device is installed on the satellite body, and a single-frame control moment gyroscope is used as an actuator of the attitude control system.

Step 2: When the satellite attitude is in a three-axis stable state, the uplink remote control command controls the electric propulsion device to ignite for T seconds, and the telemetry data is measured in real time through the gyro, rotor and frame servo system carried by the satellite. The telemetry data includes the satellite body angular velocity, frame angular velocity, frame rotation angle and rotor speed;

Step 3: Establish a satellite attitude dynamics model using a single-frame control moment gyro as the actuator of the attitude control system, and introduce the frame's rotation angular velocity and frame rotation angle into the model;

Step 4: The satellite attitude dynamics equation within _T1 seconds during the ignition of the electric propulsion device is subtracted from the satellite attitude dynamics equation within _T0 seconds before the ignition to obtain the analytical expression of the electric thrust torque;

Step 5: Record the telemetry data within T ₀ seconds before ignition and T ₁ second during ignition of the electric propulsion device, fit the data using the least squares method, and calculate the electric thrust torque generated by the electric propulsion device;

Step 6: Calculate the thrust using the electric thrust torque generated by the electric propulsion device;

Step 7: Calculate the thrust deviation, determine the deviation range, and ultimately determine the thrust generated by the electric propulsion device.

As a further technical solution of the present invention, in step 1, the electric propulsion device is offset installed on the satellite body, and the thrust does not pass through the satellite center of mass; the satellite is orthogonally installed with three single-frame control moment gyroscopes as actuators of the satellite attitude control system.

As a further technical solution of the present invention, in step 3, a satellite attitude dynamics model using a single-frame control moment gyro as an attitude control system actuator is established as follows:

Among them, h _b =[h _bi ]∈R ^3×1 is the angular momentum of the satellite body, is the rate of change of the angular momentum of the satellite body, h _w ＝[h _wi ]∈R ^3×1 is the angular momentum of the single frame moment gyro, ω _bi ＝[ω _bi ]∈R ^3×1 is the angular velocity of the satellite body coordinate system relative to the earth's inertial coordinate system, ω _ri ＝[ω _ri ]∈R ^3×1 is the frame angular velocity, and _ME ＝[ _MEi ]∈R ^3×1 is the external torque;

According to the inertial characteristics of the satellite and momentum wheel, the satellite attitude dynamics model is further written as:

Among them, J = [J _i ] ∈ R ^3×1 is the satellite body moment of inertia, J ₀ = [J _0i ] ∈ R ^3×1 is the rotor moment of inertia, is the angular velocity change rate of the satellite body coordinate system relative to the earth's inertial coordinate system, ω _f =[ω _fi ]∈R ^3×1 is the angular velocity of the rotor, θ _ri is the frame rotation angle, the relationship between the rotor angular momentum and the rotation speed is h _ω =kR, k is the rotor inertia characteristic coefficient, and R = ω _f /2π is the rotor speed.

As a further technical solution of the present invention, in step 4, within T ₀ seconds before the electric propulsion device is ignited, the satellite attitude dynamics equation is:

in, represents the average angular velocity of the satellite body within T ₀ seconds before the electric propulsion device is ignited, represents the average angular velocity of the rotor within T ₀ seconds before ignition, represents the average angular velocity of the frame within T ₀ seconds before ignition, θ _ri0 is the frame rotation angle within T ₀ seconds before ignition, represents the average rotor speed within T ₀ seconds before ignition, represents the satellite body angular velocity change rate within T ₀ seconds before ignition, M _d ＝[M _di ]∈R ^3×1 is the interference torque of the space environment;

During the ignition of the electric propulsion device, within T ₁ second, the satellite attitude dynamics equation is:

in, represents the average angular velocity of the satellite body within T ₁ second during the ignition of the electric propulsion device, represents the average angular velocity of the rotor within T ₁ second during ignition, represents the average angular velocity of the frame within T ₁ second during ignition, θ _ri1 is the frame rotation angle within T ₁ second during ignition, Indicates the average rotor speed within T ₁ second during ignition, represents the rate of change of the satellite body angular velocity within T ₁ second during ignition, _Me = [M _ei ]∈R ^3×1 represents the torque generated by the electric propulsion device;

The satellite attitude dynamics equations within T ₁ seconds during the ignition of the electric propulsion device and before T ₀ are subtracted to obtain the analytical expression of the electric thrust torque:

As a further technical solution of the present invention, in step 5, the telemetry data within T ₀ seconds before ignition of the electric propulsion device and T ₁ second during ignition are recorded, and the telemetry data are fitted using the least squares method to obtain the average angular velocity of the satellite body, the angular velocity change rate, the average angular velocity of the frame, the frame rotation angle and the average speed of the rotor, and are substituted into the electric thrust torque analytical expression of step 4 to calculate the electric thrust torque generated by the electric propulsion device.

As a further technical solution of the present invention, in step 6, the electric propulsion device generates electric thrust torque on the satellite in the X-axis and Z-axis directions, and the thrust is calculated using the electric thrust torque in the X-axis direction: Calculate the thrust using the electric thrust torque in the Z-axis direction:

As a further technical solution of the present invention, in step 7, the deviation is solved according to the result of calculating the thrust in two ways along the X-axis and the Z-axis:

If |D|≤5%, it is considered that the thrust calculation is accurate, and the thrust of the electric propulsion device |F|＝(|F ₁ |+|F ₂ |)/2 is obtained; otherwise, the electric propulsion device start-up ignition command is reissued, and steps 3 to 7 are repeated to perform thrust calculation again.

Compared with the prior art, the beneficial effects of the present invention are as follows: compared with the previous method of using the change of the number of spacecraft orbital elements to calculate the thrust of the electric propulsion device, the present invention has the advantages of simple calculation method, short calculation time, high calculation accuracy and easy engineering operation and implementation. In addition, compared with traditional flywheels and jet thrusters, the single-frame control torque gyro has higher speed stability and pointing accuracy, and smooth output torque, because it can use a smaller rotor to obtain a larger output torque, and can also calculate the thrust of the ultra-high specific impulse electric propulsion device to be developed in the future.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a diagram showing a satellite body coordinate system definition according to an embodiment of the present invention.

FIG. 2 is a schematic diagram of the installation of a single-frame control moment gyroscope provided in an embodiment of the present invention.

FIG3 is a composition diagram of a single-frame control moment gyro system provided by an embodiment of the present invention.

FIG4 is a flow chart of a thrust inversion method for an electric propulsion device based on a single-frame control moment gyro provided in an embodiment of the present invention.

DETAILED DESCRIPTION

The specific implementation of the thrust inversion method of the electric propulsion device based on the single-frame control moment gyro is described in detail below in conjunction with the accompanying drawings. The following description is only used to illustrate the present invention but is not used to limit the scope of the present invention.

In this method, the satellite angular velocity is measured by the gyroscope, the rotor can feedback its own rotation speed in real time, the frame rotation angular velocity and rotation angle are given by the frame servo system, and the satellite on-orbit data are obtained through the ground station. After the electric propulsion device works, the satellite attitude changes, and the single frame control torque gyro is used as an actuator to control the satellite attitude stability. The thrust is calculated using the electric thrust torque generated by the electric propulsion device, the thrust deviation is solved, and finally the thrust generated by the electric propulsion device is determined.

Example 1

Referring to FIG. 1 to FIG. 4 , an embodiment of the present invention provides a thrust inversion method for an electric propulsion device based on a single-frame control moment gyro, the method comprising the following steps:

Step 1: Establish the satellite body coordinate system and the frame coordinate system. The satellite is equipped with orthogonal single-frame control moment gyroscopes along the +X, +Y and +Z axis directions to control the satellite attitude. An electric propulsion device is installed on the satellite body, and a single-frame control moment gyroscope is used as an actuator of the attitude control system.

Specifically, a satellite body coordinate system OXYZ is established, in which the origin O is the satellite center of mass, the +X axis points to the direction of satellite flight, the +Z axis points to the center of the earth, and the +Y axis satisfies the right-hand rule and is perpendicular to the plane formed by XOZ and points to the sun. The electric propulsion device is installed offset on the satellite body, and the thrust does not pass through the satellite center of mass; three single-frame control moment gyroscopes are orthogonally installed on the satellite as actuators of the satellite attitude control system. The nozzle of the electric propulsion device is located in the plane pointed by the -Y axis, and the installation position coordinates are (dxdydz), and the thrust generated is along the +Y axis direction. A frame coordinate system Ox _g y _g z _g is established, in which the origin is the high-speed rotor center of mass, the z _g axis is along the frame axis direction, the x _g axis is along the rotor angular momentum axis direction, and the y _g axis satisfies the right-hand rule and is perpendicular to the plane formed by x _g Oz _g . The satellite is installed with orthogonal single-frame control moment gyroscopes along the +X, +Y and +Z axis directions to control the satellite attitude;

Step 2: When the satellite attitude is in a three-axis stable state, the uplink remote control command controls the electric propulsion device to ignite for T seconds, and the telemetry data is measured in real time through the gyro, rotor and frame servo system carried by the satellite. The telemetry data includes the satellite body angular velocity, frame angular velocity, frame rotation angle and rotor speed;

Specifically, when the satellite attitude is in a three-axis stable state, the electric propulsion device is controlled to ignite T ₁ second through the uplink remote control command from the ground station. The gyroscope carried by the satellite can measure the angular velocity of the current satellite body in real time, and feed it back to the frame supporting the rotor through the frame servo system of the single-frame control moment gyroscope, that is, the ground station can receive telemetry data such as the satellite body angular velocity, frame angular velocity, frame rotation angle, and rotor speed within T ₀ seconds before the ignition of the electric propulsion device and T ₁ second during the ignition.

Step 3: Establish a satellite attitude dynamics model using a single-frame control moment gyro as the actuator of the attitude control system, and introduce the frame's rotation angular velocity and frame rotation angle into the model;

Specifically, the satellite attitude dynamics model using a single-frame control moment gyro as the actuator of the attitude control system is established as follows:

Among them, h _b =[h _bi ]∈R ^3×1 is the angular momentum of the satellite body, is the rate of change of the angular momentum of the satellite body, h _w ＝[h _wi ]∈R ^3×1 is the angular momentum of the single-frame moment gyro, ω _bi ＝[ω _bi ]∈R ^3×1 is the angular velocity of the satellite body coordinate system relative to the Earth’s inertial coordinate system, ω _ri ＝[ω _ri ]∈R ^3×1 is the frame angular velocity, and _ME ＝[ _MEi ]∈R ^3×1 is the external torque.

According to the inertial characteristics of the satellite and momentum wheel, the satellite attitude dynamics model is further written as:

Among them, J = [J _i ] ∈ R ^3×1 is the satellite body moment of inertia, J ₀ = [J _0i ] ∈ R ^3×1 is the rotor moment of inertia, is the angular velocity change rate of the satellite body coordinate system relative to the earth's inertial coordinate system, ω _f =[ω _fi ]∈R ^3×1 is the angular velocity of the rotor, θ _ri is the frame rotation angle (assuming that the frame rotation angle is small and does not meet the singularity condition), the relationship between the rotor angular momentum and the speed is h _ω =kR, k is the rotor inertia characteristic coefficient, and its value is determined by the specific rotor model, and R = ω _f /2π is the rotor speed.

Step 4: The satellite attitude dynamics equation within _T1 seconds during the ignition of the electric propulsion device is subtracted from the satellite attitude dynamics equation within _T0 seconds before the ignition to obtain the analytical expression of the electric thrust torque;

Specifically, the ignition time of the electric propulsion device for on-orbit verification is short. Before and after the ignition, the satellite attitude maintains a three-axis stable state, and the single-frame control torque gyro rotor is in a non-saturated normal working state. Then, within T ₀ seconds before the ignition of the electric propulsion device, the satellite attitude dynamics equation is:

in, represents the average angular velocity of the satellite body within T ₀ seconds before the electric propulsion device is ignited, represents the average angular velocity of the rotor within T ₀ seconds before ignition, represents the average angular velocity of the frame within T ₀ seconds before ignition, θ _ri0 is the frame rotation angle within T ₀ seconds before ignition, represents the average rotor speed within T ₀ seconds before ignition, represents the satellite body angular velocity change rate within T ₀ seconds before ignition, and M _d = [M _di ]∈R ^3×1 is the interference torque of the space environment.

During the ignition of the electric propulsion device, within T ₁ second, the satellite attitude dynamics equation is:

in, represents the average angular velocity of the satellite body within T ₁ second during the ignition of the electric propulsion device, represents the average angular velocity of the rotor within T ₁ second during ignition, represents the average angular velocity of the frame within T ₁ second during ignition, θ _ri1 is the frame rotation angle within T ₁ second during ignition, Indicates the average rotor speed within T ₁ second during ignition, represents the rate of change of the satellite body angular velocity within T ₁ second during ignition, and _Me = [ _Mei ]∈R ^3×1 represents the torque generated by the electric propulsion device.

The satellite attitude dynamics equations within T ₁ seconds during the ignition of the electric propulsion device and before T ₀ are subtracted to obtain the analytical expression of the electric thrust torque:

Step 5: Record the telemetry data within T ₀ seconds before ignition and T ₁ second during ignition of the electric propulsion device, fit the data using the least squares method, and calculate the electric thrust torque generated by the electric propulsion device;

Specifically, the telemetry data within T ₀ seconds before the ignition of the electric propulsion device and T ₁ second during the ignition are recorded, and the telemetry data are fitted by the least square method to obtain the average angular velocity of the satellite body, the angular velocity change rate, the frame average angular velocity, the frame rotation angle and the average rotation speed of the rotor, and the electric thrust torque generated by the electric propulsion device is calculated by substituting them into the electric thrust torque analytical expression of step 4;

Step 6: Calculate the thrust using the electric thrust torque generated by the electric propulsion device;

Specifically, the electric propulsion device generates electric thrust torque on the satellite in the X-axis and Z-axis directions, and the thrust is calculated using the electric thrust torque in the X-axis direction:

Calculate the thrust using the electric thrust torque in the Z-axis direction:

Step 7: Calculate the thrust deviation, determine the deviation range, and ultimately determine the thrust generated by the electric propulsion device.

Specifically, the deviation is calculated based on the results of the thrust calculations in two ways: X-axis and Z-axis:

If |D|≤5%, it is considered that the thrust calculation is accurate, and the thrust of the electric propulsion device |F|＝(|F ₁ |+|F ₂ |)/2 is obtained; otherwise, the electric propulsion device start-up ignition command is reissued, and steps 3 to 7 are repeated to perform thrust calculation again.

The above are only preferred embodiments of the present invention, and are not intended to limit the patent scope of the present invention. Any equivalent structure or equivalent process transformation made using the contents of the present invention specification and drawings, or directly or indirectly applied in other related technical fields, are also included in the patent protection scope of the present invention.

Claims (6)

1. The thrust inversion method of the electric propulsion device based on the single-frame control moment gyro is characterized by comprising the following steps of:

Step 1: establishing a satellite body coordinate system and a frame coordinate system, and installing a single-frame control moment gyro of an orthogonal configuration along the +X, +Y and +Z axis directions to control the satellite attitude, wherein an electric propulsion device is installed on the satellite body, and the single-frame control moment gyro is used as an actuating mechanism of an attitude control system;

step 2: when the satellite attitude is in a triaxial steady state, an uplink remote control instruction controls an electric propulsion device to ignite for T seconds, and telemetry data are measured in real time through a gyroscope, a rotor and a frame servo system carried by the satellite, wherein the telemetry data comprise the angular speed of a satellite body, the angular speed of a frame, the rotation angle of the frame and the rotation speed of the rotor;

Step 3: establishing a satellite attitude dynamics model by using a single-frame control moment gyro as an actuating mechanism of an attitude control system, and introducing the rotation angular speed and the rotation angle of the frame into the model;

step 4: the method comprises the steps of utilizing a satellite attitude kinetic equation within T ₁ seconds in ignition of an electric propulsion device to make a difference with a satellite attitude kinetic equation within T ₀ seconds before ignition to obtain an analysis expression of electric propulsion torque;

step 5: recording telemetry data in T ₀ seconds before ignition and T ₁ seconds during ignition of the electric propulsion device, and calculating to obtain electric propulsion torque generated by the electric propulsion device by using least square fitting data;

step 6: calculating thrust by utilizing an electric thrust moment generated by the electric propulsion device;

step 7: calculating thrust deviation, judging a deviation range, and finally determining the thrust generated by the electric propulsion device;

In the step 3, a satellite attitude dynamics model adopting a single-frame control moment gyro as an actuating mechanism of an attitude control system is established as follows:

Wherein h _b＝[h_bi]∈R^3×1 is the angular momentum of the satellite body, H _w＝[h_wi]∈R^3×1 is the angular momentum of the single-frame moment gyro, omega _bi＝[ω_bi]∈R^3×1 is the angular velocity of the satellite body coordinate system relative to the earth inertial coordinate system, omega _ri＝[ω_ri]∈R^3×1 is the frame angular velocity, and M _E＝[M_Ei]∈R^3×1 is the external moment;

According to the inertial characteristics of the satellite and the momentum wheel, the satellite attitude dynamics model is further written as:

Wherein J= [ J _i]∈R^3×1 ] is the moment of inertia of the satellite body, J ₀＝[J_0i]∈R^3×1 is the moment of inertia of the rotor, The angular velocity change rate of the satellite body coordinate system relative to the earth inertia coordinate system is ω _f＝[ω_fi]∈R^3×1, the angular velocity of the rotor is θ _ri, the relation between the angular momentum of the rotor and the rotational speed is h _ω =kr, k is a rotor inertia characteristic coefficient, and r=ω _f/2pi is the rotational speed of the rotor.

2. The method for inverting the thrust of the electric propulsion device based on the single-frame control moment gyro according to claim 1, wherein in the step 1, the electric propulsion device is installed on a satellite body in a biased manner, and the thrust does not pass through the mass center of the satellite; three single-frame control moment gyroscopes are orthogonally arranged on the satellite and serve as actuating mechanisms of a satellite attitude control system.

3. The thrust reversal method of an electric propulsion device based on a single-frame control moment gyro according to claim 1, wherein in step 4, the satellite attitude dynamics equation is that within T ₀ seconds before the electric propulsion device fires:

Wherein, The average angular velocity of the satellite body within T ₀ seconds before ignition of the electric propulsion unit is indicated, The average angular velocity of the inner rotor is shown for T ₀ seconds prior to ignition,Represents the average angular velocity of the frame within T ₀ seconds before ignition, theta _ri0 is the rotation angle of the frame within T ₀ seconds before ignition,Represents the average rotor speed in T ₀ seconds before ignition,The change rate of the angular velocity of the satellite body within T ₀ seconds before ignition is shown, and M _d＝[M_di]∈R^3×1 is the disturbance moment of the space environment;

Within T ₁ seconds in ignition of the electric propulsion device, the satellite attitude dynamics equation is as follows:

Wherein, The average angular velocity of the satellite body within T ₁ seconds of ignition of the electric propulsion unit is shown, The average angular velocity of the rotor during T ₁ seconds in ignition is indicated,Represents the average angular velocity of the frame within T ₁ seconds in ignition, θ _ri1 is the rotation angle of the frame within T ₁ seconds in ignition,Represents the average rotor speed in T ₁ seconds during ignition,The change rate of the angular velocity of the satellite body within T ₁ seconds in ignition is shown, and M _e＝[M_ei]∈R^3×1 is the moment generated by the electric propulsion device;

The satellite attitude kinetic equation in T ₁ seconds in the ignition of the electric propulsion device and in T ₀ before the ignition is subjected to difference value to obtain an electric propulsion moment analysis expression:

4. The thrust reversal method of an electric propulsion device based on a single-frame control moment gyro according to claim 1, wherein in step 5, telemetry data in T ₀ seconds before ignition and T ₁ seconds during ignition of the electric propulsion device are recorded, the telemetry data are fitted by a least square method, average angular velocity of a satellite body, change rate of angular velocity, average angular velocity of a frame, rotation angle of the frame and average rotation speed of a rotor are obtained, and electric propulsion moment generated by the electric propulsion device is calculated by substituting the electric propulsion moment analytic expression in step 4.

5. The thrust inversion method of an electric propulsion device based on a single-frame control moment gyro according to claim 1, wherein in step 6, the electric propulsion device generates electric propulsion moment to satellites in the X-axis and Z-axis directions, and thrust is calculated by using the electric propulsion moment in the X-axis direction: calculating thrust by utilizing the electric thrust moment in the Z-axis direction:

6. The thrust inversion method of an electric propulsion device based on a single-frame control moment gyro according to claim 1, wherein in step 7, the calculated deviation of the result of the thrust is calculated according to two modes of an X axis and a Z axis:

if the D is less than or equal to 5%, the thrust measurement is considered to be accurate, and the thrust |F|= (|F ₁|+|F₂ |)/2 of the electric propulsion device is obtained; otherwise, the starting ignition instruction of the electric propulsion device is sent again, and the steps 3 to 7 are repeated to calculate the thrust again.