Real-time characterization method for the simulation process of the combined action of space atomic oxygen and ultraviolet irradiation

Abstract

The present invention provides a real-time characterization method for the simulation process of the combined action of space atomic oxygen and ultraviolet radiation, belonging to the field of aerospace technology. When the present invention conducts simulation characterization on an in-orbit space vehicle in the combined environment of space atomic oxygen and ultraviolet radiation to study the ultraviolet and atomic oxygen effects, it characterizes the operating attitude of the vehicle in the form of quaternions, and jointly determines the surface of the vehicle that is subjected to ultraviolet electromagnetic radiation according to the operating attitude of the space vehicle and the incident direction of solar ultraviolet radiation. It characterizes the space position and operating speed of the vehicle, and jointly determines the interaction between the vehicle and atomic oxygen in combination with the operating attitude of the vehicle and the space atomic oxygen environment obtained according to the atomic oxygen calculation model. The present invention comprehensively considers the windward side and sun-facing side of the space vehicle, and can relatively realistically consider the real-time quantitative characterization of the space environment in which the space vehicle is exposed to sunlight and faces atomic oxygen during its in-orbit operation.

Description

Technical Field

[0001] The present invention relates to the field of aerospace technology, and in particular, to a method for real-time characterization of the simulation process of the combined action of space atomic oxygen and ultraviolet irradiation. Background Technique

[0002] Atomic oxygen is an important component of the Earth's upper atmosphere and poses a serious threat to spacecraft operating in low Earth orbit. For example, when atomic oxygen impacts the surface of a spacecraft, it may undergo specular elastic reflection, or its energy and kinetic energy may change to produce diffuse scattering; the incident atomic oxygen may react with other substances attached to the surface of the spacecraft, such as colliding with NO at high speed to form an excited nitrogen oxide compound (NO2 * ), when NO2 * is dissociated from the surface of the spacecraft and returns to the ground state, the excited energy is released in the form of photons, generating orange-red glow, which will affect the stealth of the spacecraft. Generally speaking, the harmful effects of atomic oxygen are mainly manifested as directly reacting with the surface material of the spacecraft to form volatile reaction products, producing an erosion effect. The result not only leads to mass loss of the material, but also deteriorates the surface performance, thus becoming an important factor affecting the on-orbit life and reliability of the spacecraft. Many polymers, carbon fiber composites, and Ag, etc., are sensitive materials that are easily damaged by atomic oxygen. The erosion thickness Δx of the material by atomic oxygen can be calculated as the product of the reaction coefficient R e of atomic oxygen and the fluence Φ, that is, Δx = R e ·Φ. For example, the reaction coefficient R e of atomic oxygen with polyimide = 3.0×10 -24 cm 3 / atom, and further, the erosion thickness after exposure at an altitude of 400 km for 1 year can reach dozens of μm, indicating that a polyimide film with a relatively thin thickness may disappear within 1 year.

[0003] To effectively prevent the harmful effects of atomic oxygen, it is necessary to conduct in-depth research on the interaction between atomic oxygen and spacecraft surface materials, which is of great significance for improving the in-orbit service reliability and lifespan of low-Earth orbit spacecraft. An effective research approach is to analyze through ground simulation tests and establish the relationship between space atomic oxygen environmental parameters and the performance degradation of spacecraft outer surface materials, serving as a necessary basis for predicting the in-orbit service lifespan of spacecraft. The prerequisite for solving this problem is to quantitatively characterize the space atomic oxygen environment. The distribution of space atomic oxygen is related to many factors such as orbital altitude, season, local solar time at the location of the spacecraft, and solar activity. Generally, the number density of atomic oxygen decreases exponentially with increasing altitude. During the high solar activity years, the number density of atomic oxygen increases. The higher the orbital altitude, the more obvious the difference in the atomic oxygen number density between high and low solar activity years. For example, at an altitude of 1000 km, the difference between the two is about 4 orders of magnitude. The rotation of the Earth can cause diurnal variations in the atomic oxygen number density. At a certain altitude, the atomic oxygen number density often reaches its maximum value at local solar time 14:00 and its minimum value at local solar time 4:00. Since the Earth revolves around the sun in the ecliptic plane and passes through the solar equatorial plane twice a year, there are obvious semi-annual variations in atmospheric temperature and atomic oxygen number density. The atmospheric temperature and atomic oxygen number density reach their maximum values in October each year and their minimum values in July.

[0004] Solar ultraviolet radiation has an important impact on the thermospheric atmosphere and is the root cause of the generation of atomic oxygen. The solar ultraviolet radiation energy fluence refers to the solar ultraviolet radiation energy received per unit area of a vertical surface within a given time, and its calculation formula is as follows: H s = E λ ·t; where H s is the solar ultraviolet radiation energy fluence, with the unit of J / m 2 ; E λ is the solar ultraviolet irradiance, with the unit of W / m 2 ; t is the irradiation time, with the unit of seconds. During ground simulation tests, the solar electromagnetic radiation energy fluence is often characterized in units of equivalent sun hours (esh). This is equivalent to the solar electromagnetic radiation energy received per unit area when irradiated for 1 h under the condition of 1 solar constant. To define the acceleration test factor, the "equivalent solar irradiance" is often used to characterize the actual spectral irradiance. The equivalent solar irradiance is the ratio of the actual spectral irradiance during the test to the solar spectral irradiance in the corresponding band, and the test acceleration factor can be obtained from the multiple of the equivalent solar irradiance relative to the solar spectral irradiance. As mentioned above, the spectral irradiances of solar near-ultraviolet and far-ultraviolet radiation are 118 W / m 2 and 0.1 W / m 2 respectively, so the corresponding irradiation energy fluences can be calculated by substituting them into the above formula.

[0005] The energy of atomic oxygen in space is about 5 eV, and its chemical properties are active. It can cause severe oxidative erosion on the surface materials of spacecraft, resulting in the degradation or even failure of material properties. The wavelength of space ultraviolet light is generally in the range of 10 nm - 400 nm. Long-term ultraviolet radiation can cause the material to fade and crack, and the thermal, optical, mechanical and other properties will decline. Low Earth Orbit spacecraft will be irradiated by atomic oxygen and ultraviolet during flight. This comprehensive space environment will cause the degradation of the surface material properties of the spacecraft, which may endanger the safety of spacecraft operation or reduce the service life of the spacecraft. Sensitive surface materials such as thermal control coatings, multi-layer insulators and optical surfaces are particularly vulnerable to influence. Minor surface property changes will have a greater impact on their functions. In addition, the breakage of chemical bonds will also generate some new reactive groups on the material surface, thus promoting the erosion of atomic oxygen on the material. Summary of the Invention

[0006] During the on-orbit operation of a spacecraft, it will be affected by the combined action of atomic oxygen and ultraviolet radiation. How to real-time characterize the comprehensive environment of atomic oxygen and ultraviolet electromagnetic radiation is the technical problem to be solved by the present invention.

[0007] To solve the above problems, the present invention provides a real-time characterization method for the simulation process of the combined action of space atomic oxygen and ultraviolet radiation, including:

[0008] Determine the calculation range of real-time simulation according to the simulation start time and the simulation end time;

[0009] According to the simulation step size, continuously accumulate the simulation step size from the simulation start time to obtain the stepped simulation time;

[0010] Solve the real-time quantitative characterization of the comprehensive environment of space atomic oxygen and ultraviolet radiation at each simulation time, specifically including:

[0011] Obtain the incident direction of solar ultraviolet electromagnetic radiation in space,

[0012] For each satellite in operation at the same simulation time, obtain the three-dimensional coordinates and running speed parameters of the satellite in the inertial system. The three-dimensional coordinates and the running speed parameters are used to characterize the position and running speed of the spacecraft,

[0013] Obtain the quaternion attitude parameters of the satellite. The quaternion attitude parameters are used to characterize the running attitude of the spacecraft,

[0014] Obtain the atomic oxygen number density according to the atomic oxygen calculation model. The atomic oxygen number density is used to characterize the space atomic oxygen environment,

[0015] Analyze the surface of the spacecraft affected by the ultraviolet electromagnetic radiation according to the operating attitude of the spacecraft and the incident direction of the solar ultraviolet electromagnetic radiation, and analyze the atomic oxygen effect on the spacecraft according to the operating speed and attitude of the spacecraft and the space atomic oxygen environment.

[0016] Preferably, the method for obtaining the incident direction of the solar ultraviolet electromagnetic radiation includes: assuming that the sunlight received by the spacecraft operating near the Earth is parallel light, obtaining the spatial position coordinates of the sun in the J2000 geocentric coordinate system at the simulation time according to the DE430 ephemeris published by JPL, and obtaining the direction vector pointing from the geocenter to the solar centroid, taking the negative of the direction vector and normalizing it to obtain the direction vector of the solar light in the J2000 geocentric coordinate system, and the direction vector is the incident direction of the solar ultraviolet electromagnetic radiation.

[0017] Preferably, the method for obtaining the position and operating speed of the spacecraft includes: obtaining the orbital elements of the spacecraft operating in orbit according to the two-body motion equation and Kepler's three laws, and obtaining the position and operating speed of the spacecraft in the inertial system according to the orbital elements.

[0018] Preferably, obtaining the quaternion attitude parameters of the satellite includes:

[0019] Define the orbital coordinate system as O0X0Y0Z0, and the direction vector of the rotation axis e in the orbital coordinate system is e = (C x , C y , C z ), where (C x , C y , C z ) represents the direction cosine, and the orbital coordinate system rotates by an angle α around the rotation axis e to obtain the satellite body coordinate system O b X b Y b Z b , then the quaternion attitude parameter q of the satellite is:

[0020]

[0021] Among them, q1, q2, and q3 are vector components, q0 is the scalar component, and q1 2 + q2 2 + q3 2 + q0 2 = 1.

[0022] Preferably, the atomic oxygen calculation model includes the NRLMSISE-00 atmospheric model. The inputs of the NRLMSISE-00 atmospheric model include the number of days from January 1 of the current year to the current day, the number of seconds from 00:00:00 of the current day to the solution time, geographical longitude, geographical latitude, altitude, the solar 10.7 cm radiation flux of the previous day, the average solar 10.7 cm radiation flux over 81 days, and an 8-digit array obtained based on the average geomagnetic index of the current day and the 20 3-hour average geomagnetic indices before the solution time.

[0023] Preferably, the geographical longitude and the geographical latitude are obtained by coordinate system transformation, which transforms the three-dimensional coordinates of the satellite in the inertial system into the geodetic coordinates in the WGS-84 coordinate system.

[0024] Preferably, the coordinate system transformation includes the transformation between the J2000 coordinate system and the Earth-centered Earth-fixed coordinate system. The overall transformation matrix HG from the J2000 coordinate system to the Earth-centered Earth-fixed coordinate system is determined according to the second formula, and the second formula is:

[0025] HG = (PR)(NR)(ER)(EP);

[0026] In the formula, PR is the precession matrix, NR is the nutation matrix, ER is the Earth rotation matrix, and EP is the polar motion matrix;

[0027] The precession matrix is obtained through the third formula, and the third formula is:

[0028] PR = R z (-z A )R y (-θ A )R z (-ζ A );

[0029] In the formula: R z represents the Euler angle coordinate transformation of the Z-axis of the J2000 coordinate system, R y is the Euler angle coordinate transformation of the Y-axis of the J2000 coordinate system, z A , θ A , ζ A are precession constants, which are calculated through the fourth formula, and the fourth formula is:

[0030]

[0031] In the fourth formula: T is the Julian century number starting from J2000.0;

[0032] The nutation matrix is obtained through the fifth formula, and the fifth formula is:

[0033] NR = RX (-ε A )R z (-Δψ)R X (-ε A +Δε);

[0034] In the fifth formula: ε A is the mean obliquity of the ecliptic, Δψ is the nutation in longitude, Δε is the nutation in obliquity, and R X is the Euler angle coordinate transformation for the X-axis of the J2000 coordinate system;

[0035] The Earth rotation matrix is obtained through the tenth formula, and the tenth formula is:

[0036] ER = R z (θ G );

[0037] In the tenth formula, θ G is the Greenwich sidereal time;

[0038] The polar motion matrix is obtained through the thirteenth formula, and the thirteenth formula is:

[0039] EP = R Y (-x p )R X (-y p );

[0040] In the formula, x p , y p are the geocentric polar coordinates.

[0041] Preferably, the coordinate system transformation further includes the transformation between the geocentric geodetic coordinate system and the Earth-centered, Earth-fixed coordinate system. The geocentric geodetic coordinate system includes the WGS-84 coordinate system. Let the coordinates of the target in the WGS-84 coordinate system be (lon, lat, alt). The coordinates (x d , y d , z d ) of the target in the Earth-centered, Earth-fixed coordinate system are determined according to the fourteenth formula, and the fourteenth formula is:

[0042]

[0043] In the fourteenth formula: e 2 = (a 2 - b 2 ) / a 2 , a is the semi-major axis of the WGS-84 ellipsoid corresponding to the WGS-84 coordinate system, b is the semi-minor axis of the WGS-84 ellipsoid, N is the normal length of the WGS-84 ellipsoid, and e is the eccentricity of the WGS-84 ellipsoid. Among them,

[0044] Preferably, the output of the NRLMSISE-00 atmospheric model includes the number densities of N2, O2, He, Ar, N, H, O, and ionospheric positive oxygen ions O + , the neutral atmospheric temperature, and the total atmospheric density.

[0045] Preferably, the time is characterized by the Julian day, which is calculated according to the first formula, and the first formula includes:

[0046]

[0047] where: JD represents the Julian day, INT represents taking the integer, Year represents the year in the Gregorian calendar, Month represents the month, Day represents the day, Hour represents the hour, Minute represents the minute, and Second represents the second.

[0048] The advantages of the present invention compared with the prior art are as follows:

[0049] When the present invention conducts simulation characterization on an in-orbit space vehicle in the comprehensive environment of space atomic oxygen and ultraviolet radiation to study the ultraviolet and atomic oxygen effects, the running attitude of the space vehicle is characterized in a quaternion manner, and the surface of the space vehicle affected by ultraviolet electromagnetic radiation is jointly determined according to the running attitude of the space vehicle and the incident direction of solar ultraviolet electromagnetic radiation. The spatial position and running speed of the space vehicle are characterized, and the interaction between the space vehicle and atomic oxygen is jointly determined in combination with the running attitude of the space vehicle and the space atomic oxygen environment obtained according to the atomic oxygen calculation model. The present invention comprehensively considers the windward side and sun-facing side of the space vehicle, and can more realistically consider the real-time quantitative characterization of the space environment of the space vehicle being exposed to sunlight and facing atomic oxygen during in-orbit operation. Description of the Drawings

[0050] Figure 1 is a flowchart of the real-time characterization method for the simulation process of the comprehensive action of space atomic oxygen and ultraviolet irradiation in an embodiment of the present invention;

[0051] Figure 2 is a schematic diagram of the Earth-centered Earth-fixed coordinate system in an embodiment of the present invention;

[0052] Figure 3 is a schematic diagram of the WGS-84 coordinate system in an embodiment of the present invention;

[0053] Figure 4 is a schematic diagram of the J2000 inertial coordinate system and the satellite body coordinate system in an embodiment of the present invention;

[0054] Figure 5Schematic diagram of the conversion relationship between the J2000 coordinate system and the Earth-centered Earth-fixed coordinate system in the embodiments of the present invention;

[0055] Figure 6 Schematic diagram of the Euler angle attitude change in the embodiments of the present invention;

[0056] Figure 7 Schematic diagram of the quaternion attitude change in the embodiments of the present invention;

[0057] Figure 8 Schematic diagram of the quaternion coordinate transformation in the embodiments of the present invention;

[0058] Figure 9 Schematic diagram of the coordinate system transformation represented by Euler angles in the embodiments of the present invention;

[0059] Figure 10 In the embodiments of the present invention Figure 4 Display of the attitude change during the space operation of the satellite;

[0060] Figure 11 Real-time change curve graph of the position and velocity of the satellite in the embodiments of the present invention;

[0061] Figure 12 Real-time change curve graph of the atomic oxygen of the satellite in the embodiments of the present invention;

[0062] Figure 13 Solar ray vector in the embodiments of the present invention.

[0063] Explanation of reference numerals:

[0064] 1. Greenwich meridian plane; 2. Target point; 3. Equator; 4. Mean equinox point at J2000. Detailed implementation manners

[0065] To make the above objects, features, and advantages of the present invention more obvious and understandable, the following detailed description of the specific embodiments of the present invention will be given with reference to the accompanying drawings.

[0066] During the on-orbit operation of a space vehicle, both the motion state, including the space position and the motion speed, are changing, the motion attitude is also changing in real time, and the space atomic oxygen environment it is in also changes with the different space positions of the space vehicle. Therefore, it is necessary to perform real-time characterization calculations on various parameters of the space vehicle. A real-time characterization method for the comprehensive action simulation process of space atomic oxygen and ultraviolet irradiation in the embodiments of the present invention includes:

[0067] Determine the calculation range of the real-time simulation according to the simulation start time and the simulation end time;

[0068] According to the simulation step size, continuously accumulate the simulation step size from the simulation start time to obtain the stepped simulation time;

[0069] For real-time quantitative characterization of the comprehensive environment of atomic oxygen and ultraviolet radiation in space at each simulation time, see Figure 1 As shown below, which specifically includes:

[0070] Obtain the incident direction of solar ultraviolet electromagnetic radiation in space.

[0071] For each satellite in operation at the same simulation time, obtain the three-dimensional coordinates and operating speed parameters of the satellite in the inertial system. The three-dimensional coordinates and the operating speed parameters are used to characterize the position and operating speed of the spacecraft.

[0072] Obtain the quaternion attitude parameters of the satellite. The quaternion attitude parameters are used to characterize the operating attitude of the spacecraft.

[0073] Obtain the atomic oxygen number density according to the atomic oxygen calculation model. The atomic oxygen number density is used to characterize the space atomic oxygen environment.

[0074] Based on the operating attitude of the spacecraft and the incident direction of the solar ultraviolet electromagnetic radiation, analyze the surface of the spacecraft affected by the ultraviolet electromagnetic radiation. Based on the operating speed and operating attitude of the spacecraft and the space atomic oxygen environment, analyze the atomic oxygen effect on the spacecraft.

[0075] During the on-orbit operation of a spacecraft, especially a low-orbit spacecraft, due to its periodic circular motion around the Earth, the spacecraft itself will be subjected to periodic solar ultraviolet radiation during operation. The combination of ultraviolet high-energy irradiation and the atomic oxygen environment will cause more serious harm to the spacecraft. Ultraviolet and atomic oxygen mainly act on the surface of the spacecraft. Therefore, when considering the ultraviolet and atomic oxygen effects, it is necessary to consider the windward side and sun-exposed side of the spacecraft, so as to achieve a more realistic real-time quantitative characterization of the space environment of the spacecraft being sun-exposed and facing atomic oxygen during on-orbit operation.

[0076] Embodiments of the present invention solve the technical problem of being able to characterize the comprehensive environment of atomic oxygen and ultraviolet electromagnetic radiation on the sun-exposed surface and the windward surface of a space vehicle in real time, considering the incident direction of ultraviolet electromagnetic radiation and the simulation characterization of the atomic oxygen on the windward surface. Specifically, for the problem of the incident direction of ultraviolet electromagnetic radiation, during the operation of the space vehicle in space, it is not the case that each surface of the space vehicle uniformly receives solar ultraviolet electromagnetic radiation. Instead, the specific surface of the space vehicle that is affected by ultraviolet electromagnetic radiation is jointly determined by the operating attitude of the space vehicle and the direction of the solar parallel light. This embodiment combines the satellite attitude and the ultraviolet incident direction in space to determine the surface of the satellite that is affected by ultraviolet rays. For the problem of the simulation characterization of the atomic oxygen on the windward surface, it is necessary to analyze and characterize by combining the operating speed and attitude of the space vehicle, as well as the space atomic oxygen density environment in which the space vehicle is located, to determine the atomic oxygen effect on the space vehicle in orbit. To solve the atomic oxygen environment problem on the windward surface, this embodiment calculates the number density of atomic oxygen through the position of the space vehicle, the relative speed of the space vehicle's movement to the atomic oxygen on the windward side, and in combination with the atomic oxygen model, and analyzes the atomic oxygen effect on the space vehicle.

[0077] First, the time involved in the embodiments of the present invention is described. In this embodiment, the time calibration mainly faces two directions. One is for the user to use the Gregorian calendar year, month, day, hour, minute, and second, and the Coordinated Universal Time (UTC) is used to define the operating time of the space vehicle, the start time and end time of the simulation calculation. The other is for the internal simulation calculation process of this embodiment. For the convenience and accuracy of the simulation calculation, continuous real numbers are required for the time calculation. In this embodiment, the Julian day is used to convert the time into continuous linear real numbers.

[0078] Coordinated Universal Time is also known as the world's unified time. Based on the atomic second, at the 13th General Conference on Weights and Measures in 1967, a resolution was passed to adopt the following definition to replace the astronomical definition of the second: One second is the time duration of 9,192,631,770 cycles of radiation corresponding to the transition between the two hyperfine levels of the ground state of the cesium-133 atom. International Atomic Time is an international reference time scale based on the above definition of the second and belongs to the International System of Units (SI). To ensure that the difference between Coordinated Universal Time and Universal Time does not exceed 0.9 seconds, the International Earth Rotation Service in Paris will decide to add leap seconds to Coordinated Universal Time when necessary.

[0079] The Julian day is the number of days starting from 12h on January 1, 4573 BC. By converting the Gregorian calendar year, month, and day into Julian days, it is possible to obtain the description of the difference between two Gregorian calendar moments under a unified specification, which is convenient for formulating the simulation step size. Similarly, the Julian day is also applicable to the calculation of the Earth's rotation angle.

[0080] The method for calculating the Julian day from the Gregorian calendar is to calculate the Julian day using the first formula. The first formula is as follows:

[0081]

[0082] In the first formula: JD represents the Julian day, INT() represents taking the integer, Year represents the year in the Gregorian calendar, Month represents the month, Day represents the day, Hour represents the hour, Minute represents the minute, and Second represents the second.

[0083] In the embodiments of the present invention, different space calibration coordinate systems need to be established when calculating the position and velocity of a spacecraft. The following is an explanation for different coordinate system systems.

[0084] 1) Earth-centered, Earth-fixed coordinate system. In order to transform the geodetic coordinates with a certain point as the fulcrum to the epoch coordinate system where the satellite is located, first, the coordinates of the target point in the geocentric geodetic coordinate system need to be transformed to the Earth-centered, Earth-fixed coordinate system. The definition of the Earth-centered, Earth-fixed coordinate system is as follows: See Figure 2 As shown, the origin of the Earth-centered, Earth-fixed coordinate system coincides with the Earth's mass center. The Z-axis points to the Earth's North Pole, the X-axis points to the intersection of the equator 3 and the Greenwich meridian plane 1, and the Y-axis, together with the X-axis and Z-axis, forms a right-handed rectangular coordinate system.

[0085] 2) Geocentric geodetic coordinate system. When using the space atomic oxygen environment model, geodetic longitude and latitude are often used. In this embodiment, subsequent calculations are carried out according to the WGS-84 coordinate system and its corresponding WGS-84 Earth ellipsoid model. The definition of the coordinate system is as follows: WGS-84 is a centroid coordinate system, and its origin coincides with the Earth's mass center. See Figure 3 As shown, its Z-axis points to the direction of the Conventional Terrestrial Pole (CTP) defined by the Bureau International de l'Heure in 1984.0. The X-axis points to the intersection of the conventional meridian plane and the intersection of the conventional terrestrial pole and the equator 3. The Y-axis, together with the X-axis and Z-axis, forms a right-handed rectangular coordinate. The geodetic coordinates (lon, lat, alt) in the WGS-84 coordinate system are the longitude, latitude, and altitude of the target in the WGS-84 coordinate system. The geodetic longitude lon is the angle between the geodetic meridian plane of the target point 2 and the prime meridian plane. The geodetic latitude lat is the angle between the normal of the WGS-84 ellipsoid passing through the target point 2 and the equator 3 plane. Figure 3 One of the dotted lines forming the included angle lat represents a line on the equator 3 plane, and alt is the normal distance between the target point 2 and the WGS-84 ellipsoid surface, where Figure 3 the light-colored lines represent the ellipsoid surface.

[0086] Corresponding to the WGS-84 coordinate system is the WGS-84 ellipsoid. The parameters of the WGS-84 ellipsoid are shown in Table 1, where a is the semi-major axis of the ellipsoid, b is the semi-minor axis of the ellipsoid, e is the eccentricity, and f is the true anomaly.

[0087] Table 1 WGS-84 Ellipsoid Parameters

[0088]

[0089] 3) Epoch Geocentric Inertial Coordinate System. The epoch inertial coordinate system adopts the J2000 coordinate system. The celestial equator 3 and the equinox at the J2000 (12:00 on January 1, 2000) are used to define the celestial reference coordinate system. See Figure 4 As shown, the origin of the J2000 coordinate system coincides with the Earth's mass center, the X-axis points to the mean equinox at the J2000 moment 4, the Z-axis points to the North Pole, and the Y-axis forms a right-handed rectangular coordinate system with the X-axis and Z-axis. It is an inertial coordinate system.

[0090] 4) Satellite Body Coordinate System. The coordinate system is defined as follows. See Figure 4 As shown, the origin of the coordinate system is located at the satellite's mass center. The Xs-axis, Ys-axis, and Zs-axis are the three principal inertia axes of the satellite body respectively. In the absence of attitude deviation, the three coordinate axes (Xs-axis, Ys-axis, Zs-axis) of the satellite body coordinate system coincide with the satellite orbit coordinate system. At this time, the Zs-axis of the satellite points to the Earth's center, the direction of the Xs-axis coincides with the satellite's motion direction, the Ys-axis points to the negative normal direction of the satellite orbit plane, and forms a right-handed rectangular coordinate system with the Xs-axis and Zs-axis. Among them, α represents the orbital inclination, β represents the longitude of the ascending node, and δ represents the argument of perigee.

[0091] Since the geodetic coordinates of latitude, longitude, and altitude are required for subsequent calculation of the atomic oxygen number density according to the atomic oxygen calculation model, this requires coordinate system transformation. Therefore, the coordinate system transformation is introduced below, including the transformation between the J2000 coordinate system and the geocentric earth-fixed coordinate system. When transforming from the J2000 coordinate system to the geocentric earth-fixed coordinate system, the effects of precession, nutation, earth rotation, and polar motion need to be considered. Among them, the earth rotation can be represented by the Greenwich sidereal hour angle. The conversion relationship is shown in Figure 5 As shown. Among them, the overall transformation matrix HG from the J2000 coordinate system to the geocentric earth-fixed coordinate system can be determined according to the second formula. The second formula is:

[0092] HG = (PR)(NR)(ER)(EP);

[0093] In the second formula, PR is the precession matrix, NR is the nutation matrix, ER is the earth rotation matrix, and EP is the polar motion matrix. They are introduced separately below.

[0094] 1) Precession Matrix: Precession is the long-term motion of the Earth's axis of rotation under the gravitational forces of the sun, moon, and other planets. The precession matrix can be obtained according to the third formula. The third formula is:

[0095] PR = R z (-z A )R y(-θ A )R z (-ζ A );

[0096] In the third formula: R z is the Euler angle coordinate transformation for the Z-axis of the J2000 coordinate system, and R y is the Euler angle coordinate transformation for the Y-axis of the J2000 coordinate system, z A , θ A , ζ A are precession constants, which are the successive transformation angles of the three precession constants corresponding to the coordinate axes. The z A , θ A , ζ A can be obtained through the fourth formula. The fourth formula is:

[0097]

[0098] In the fourth formula: T is the Julian century number starting from J2000.0 (TD = 2451545.0). That is, if the Julian day JD of a certain moment, then the corresponding moment's

[0099] 2) Nutation matrix: Nutation is the short-term wobbling of the Earth's axis of rotation caused by external forces. The nutation matrix is obtained through the fifth formula. The fifth formula is:

[0100] NR = R X (-ε A )R z (-Δψ)R X (-ε A +Δε);

[0101] In the fifth formula: R X is the Euler angle coordinate transformation for the X-axis of the J2000 coordinate system, ε A is the mean obliquity of the ecliptic at the epoch, Δψ is the nutation in longitude, and Δε is the nutation in obliquity. They can be obtained by interpolation through the ephemeris file of DE405 using the sixth formula and the seventh formula. The sixth formula is:

[0102] ε A = 84381”.448 - 46.8150T - 0.0059T 2 + 0.001813T 3 ;

[0103] The seventh formula is:

[0104]

[0105] σ in the seventh formula iIt can be obtained by the eighth formula, and the eighth formula is:

[0106] σ i = a 1j l + a 2j l'+ a 3j F + a 4j D + a 5j Ω;

[0107] In the eighth formula: l is the mean anomaly of the moon, l’ is the mean anomaly of the sun, F is the mean elongation of the moon from the ascending node, D is the mean angular distance between the sun and the moon, and Ω is the ecliptic longitude of the moon. These parameters can be obtained by the ninth formula, and the ninth formula is:

[0108]

[0109] 3) Earth rotation matrix: The main difference between the J2000 coordinate system and the Earth-fixed coordinate system is that the Earth-fixed coordinate system takes into account the Earth's rotation. Therefore, an accurate Earth rotation matrix is particularly important. The Earth rotation matrix is obtained through the tenth formula, and the tenth formula is:

[0110] ER = R z (θ G );

[0111] In the tenth formula, θ G is the Greenwich sidereal time and can be calculated by the eleventh formula. The eleventh formula is:

[0112] θ G = 15 × [6 h 41'50”.54841 + 860184”.812866T + 0”0931047T 2 - 6”.2 × 10 -6 T 3 + Δψcos(ε M + Δε)];

[0113] In the eleventh formula, ε M is the obliquity of the ecliptic and can be calculated by the twelfth formula. The twelfth formula is:

[0114] ε M = 23°26′21.″448 - 46.″8150T - 0.″00059T 2 + 0.″001813T 3 .

[0115] 4) Polar motion matrix: The position of the Earth's axis of rotation relative to the Earth changes over time, which will change the position of the observer's celestial pole on the celestial sphere, known as polar motion. The polar motion matrix is calculated according to the thirteenth formula, and the thirteenth formula is:

[0116] EP = R Y (-x p )R X (-y p );

[0117] In the thirteenth formula: x p , y p are geodetic coordinates.

[0118] In an embodiment of the present invention, the coordinate system transformation further includes the transformation between the geocentric geodetic coordinate system and the geocentric earth-fixed rectangular coordinate system. Specifically, assuming that the coordinates of the target in the geocentric geodetic coordinate system are (lon, lat, alt), then the coordinates (x d , y d , z d ) of the target in the geocentric earth-fixed coordinate system can be expressed by the fourteenth formula, and the fourteenth formula is:

[0119]

[0120] In the fourteenth formula: e 2 =(a 2 -b 2 ) / a 2 , see the definition of the WGS-84 ellipsoid. N is the normal length, where

[0121] Similarly, the geodetic coordinates can be solved by inversely solving the longitude, latitude and altitude from the coordinates in the geocentric earth-fixed coordinate system. The geodetic coordinates (lon, lat, alt) in the geocentric geodetic coordinate system obtained are as shown in the fifteenth formula:

[0122]

[0123] In some embodiments of the present invention, the method for obtaining the incident direction of solar ultraviolet electromagnetic radiation includes: considering that the distance between the Earth's orbit and the Sun is much greater than the radii of the Sun and the Earth, so in this embodiment, it is assumed that the sunlight received by the spacecraft operating near the Earth is parallel light. Therefore, the present invention calculates the position of the Sun in the J2000 coordinate system relative to the Earth, calculates the direction vector of the sunlight, and determines the incident direction of the space ultraviolet electromagnetic radiation based on this. The specific calculation method is to obtain the spatial position coordinates of the Sun in the J2000 geocentric coordinate system at the simulation time according to the DE430 ephemeris published by JPL, and obtain the direction vector pointing from the geocenter to the center of mass of the Sun. Take the negative of the direction vector and normalize it to obtain the direction vector of the solar rays in the J2000 geocentric coordinate system.

[0124] In some embodiments of the present invention, the method for obtaining the spatial position and motion speed of a spacecraft is as follows: The spacecraft in orbit is initialized using Keplerian orbital elements for the convenience of user setting. According to the two-body motion equation and Kepler's three laws, the orbital elements of the spacecraft in orbit are obtained. The solution expression of the orbital elements is shown in the sixteenth formula, and based on the orbital elements, the position r0 and velocity v0 of the spacecraft in the inertial system are obtained, as shown in the seventeenth formula to the twentieth formula.

[0125] Among them, the sixteenth formula is:

[0126]

[0127] In the sixteenth formula, a is the semi-major axis, which is used to describe the size of the orbit; e is the eccentricity, which is used to describe the shape of the orbit; i is the inclination angle, which is used to describe the inclination angle of the orbital plane relative to the equatorial plane; Ω is the right ascension of the ascending node, which is used to describe the position of the ascending node relative to the vernal equinox; ω is the argument of perigee, which is used to describe the position of the perigee relative to the ascending node; f is the true anomaly, which is used to describe the position of the satellite relative to the perigee.

[0128] The velocity vector of the spacecraft is shown in the seventeenth formula, and the seventeenth formula is:

[0129]

[0130] Among them, v x 、v y 、v z are the x-axis component, y-axis component, and z-axis component of the velocity respectively.

[0131] The position vector of the spacecraft is shown in the eighteenth formula, and the eighteenth formula is:

[0132]

[0133] Among them, r x 、r y 、r z are the x-axis component, y-axis component, and z-axis component of the position respectively.

[0134] Then the magnitude of the velocity of the spacecraft is shown in the nineteenth formula, and the nineteenth formula is:

[0135]

[0136] The magnitude of the position of the spacecraft is shown in the twentieth formula, and the twentieth formula is:

[0137]

[0138] The angular momentum H of the orbit and its magnitude are shown in the twenty - first formula and the twenty - second formula respectively. The twenty - first formula is:

[0139]

[0140] The twenty - second formula shown is:

[0141]

[0142] Where H x 、H y 、H z are the x - axis component, y - axis component, and z - axis component of the orbital angular momentum respectively.

[0143] In the embodiments of the present invention, the motion attitude of the space vehicle is actually the relative transformation relationship between the body coordinate system of the space vehicle and the reference coordinate system. In the prior art, the Euler angle description methods such as pitch angle, yaw angle, and roll angle are used. The advantage is intuitiveness and convenience for users, but the disadvantage is the singularity problem, which affects the calculation.

[0144] The Euler angles refer to that any two coordinate systems in space can make the three coordinate axes of the two coordinate systems coincide respectively through several rotations. During the rotation process, the angle of each rotation is called the Euler angle. In the attitude description of an on - orbit spacecraft, any attitude transformation can be achieved by rotating the corresponding Euler angles around the three coordinate axes in sequence. Referring to the description of the aircraft attitude in the aviation field, the Euler angle transformation of the satellite has specific physical meanings, that is, the yaw angle Yaw is represented by the symbol φ, the pitch angle Pitch is represented by the symbol θ, and the roll angle Roll is represented by the symbol as shown in Figure 6 .

[0145] In the description of the satellite attitude, let the reference coordinate system be the orbital coordinate system O O X O Y O Z O , and the satellite body coordinate system O b X b Y b Z b . Then in the attitude transformation, the satellite body coordinate system is obtained by rotating the orbital coordinate system. Usually, the Euler angles of the satellite are specified by rotating in the rotation order of Z - X - Y respectively. φ is the yaw angle, which is the angle between the projection of the satellite's rolling axis X b (pointing to the satellite's velocity direction) on the local horizontal plane and the orbital X O axis, is the roll angle, which is the angle between the satellite's pitch axis Y b and its projection on the local horizontal plane, and θ is the pitch angle, which is the angle between the satellite's rolling axis Xb The angle between it and the projection on the local horizontal plane. Thus, an attitude transformation matrix as shown in the twenty-third formula can be obtained. The twenty-third formula is:

[0146]

[0147] In the twenty-third formula, for simplicity of representation, c represents cos and s represents sin.

[0148] Although the Euler angle representation of the attitude is simple, direct, and easy to observe, there is a singularity problem, that is, there is a certain attitude that can be represented by two sets of Euler angle values, losing uniqueness. To solve this problem, quaternions are introduced to describe the attitude problem. In some embodiments of the present invention, a quaternion attitude description method is used and the output of the attitude change of the space vehicle in space is calculated.

[0149] The quaternion q consists of a scalar part and a vector part, and its form is as shown in the twenty-fourth formula. The twenty-fourth formula is:

[0150] q = q0 + q1i + q2j + q3k = [q s ,q v ;

[0151] In the twenty-fourth formula, q s represents the scalar part with a unit of 1; q v represents the vector part, and i, j, k are the imaginary part units. i, j, k are pairwise orthogonal unit vectors;

[0152] The geometric description of the quaternion is as follows. Define the orbital coordinate system as O0X0Y0Z0. The direction vector of the rotation axis e in the orbital coordinate system is e = (C x ,C y ,C z ), where (C x ,C y ,C z ) represents the direction cosine. The orbital coordinate system rotates by an angle α around the rotation axis e to obtain the satellite body coordinate system O b X b Y b Z b . Referring to Figure 7, the satellite attitude quaternion is as shown in the twenty-fifth formula. The twenty-fifth formula is:

[0153]

[0154] In the twenty-fifth formula, q1, q2, q3 are vector components, q0 is the scalar component, and q1 2 +q2 2 + q3 2 +q0 2= 1;

[0155] The complex number expression of a quaternion is as shown in the twenty-sixth formula, and the twenty-sixth formula is:

[0156] q = cos(θ / 2) + nsin(θ / 2);

[0157] According to Euler's formula, the exponential form of a quaternion is as shown in the twenty-seventh formula, and the twenty-seventh formula is:

[0158] The norm representation of a quaternion is as shown in the twenty-eighth formula, and the twenty-eighth formula is:

[0159]

[0160] The matrix form of a quaternion is as shown in the twenty-ninth formula, and the twenty-ninth formula is:

[0161]

[0162] The calculation of a quaternion is as shown in the thirtieth formula, and the thirtieth formula is:

[0163] q1 + q2 = [q s1 + q s2 q v1 + q v2

[0164] λq = [λq s , λq v

[0165]

[0166] q * = [q s , -q v

[0167]

[0168] q -1 = q * / ||q|| 2 ;

[0169] In the thirtieth formula, λ is a scalar, q is the conjugate quaternion of q, and ||q|| is the modulus of the quaternion. In this embodiment, is used to represent quaternion multiplication. When ||q|| = 1, it is a unit quaternion, which is an important constraint condition for quaternion calculation and solution. If a quaternion is a unit quaternion, the inverse of the quaternion is equal to its conjugate.

[0170] The quaternion multiplication is represented by the thirty-first formula, and the thirty-first formula is:​​​

[0171]

[0172] For any three unit quaternions p, q, and z, their multiplication can be written in the form of the thirty-second formula, and the thirty-second formula is:

[0173]

[0174] According to Euler's theorem, see Figure 8 as shown, the motion of any rigid body can be achieved by superimposing the rotation about an axis ( Figure 8 the r-axis in it) and the translation parallel to this axis. The coordinate transformation from the coordinate system O-xyz rotating by an angle n to the coordinate system O-x'y'z' can be represented by a quaternion as shown in the thirty-third formula, and the thirty-third formula is:

[0175]

[0176] If the coordinate transformation matrix is R, then R is represented as shown in the thirty-fourth formula, and the thirty-fourth formula is:

[0177]

[0178] According to Euler's theorem, the transformation of the coordinate system can also be represented by Euler angles ψ. Assuming that the two coordinate systems rotate in the order of ZYX, see Figure 9 as shown, the transformation matrix represented by Euler angles is as shown in the thirty-fifth formula, and the thirty-fifth formula is:

[0179]

[0180] When the attitude quaternion is q, the (Euler angle value range ±90°) thirty-sixth formula can be obtained, and the thirty-sixth formula is:

[0181]

[0182]

[0183]

[0184] The conversion formula from Euler angles to quaternions is as shown in the thirty-seventh formula, and the thirty-seventh formula is:

[0185]

[0186] The spacecraft attitude model mainly includes a kinematic model and a dynamic model, mainly based on the orbital coordinate system and the body coordinate system as a reference.

[0187] In order to evaluate the interaction between atomic oxygen and spacecraft, it is necessary to calculate the flux and fluence of atomic oxygen. The number density of atomic oxygen is the basis for calculating the spatial atomic oxygen flux and cumulative flux, and can be determined through a thermospheric atmosphere model. In some embodiments of the present invention, the atomic oxygen environment characterization method is to obtain the number density of atomic oxygen using the NRLMSISE-00 atmospheric model.

[0188] The NRLMSISE-00 atmospheric model was developed by the Naval Research Laboratory (NRL) of the United States in 2000 based on the MSISE-90 model. The NRLMSISE-00 simulation model is used for orbit prediction of near-Earth spacecraft. MSIS refers to mass spectrometer and incoherent scatter radar, and the E indicates that the model covers from the ground to the exospheric bottom layer, while the early models only covered up to the thermosphere.

[0189] The NRLMSISE-00 atmospheric model has a total of 8 input items, including: the number of days from January 1 of the current year to the current day, the number of seconds from 00:00:00 of the current day to the solution time, geographical longitude, latitude, altitude, the solar radiation flux at 10.7 cm on the previous day (F 10.7 ), the average F over 81 days (3 solar rotation periods, with the current day as the midpoint) 10.7 , and an 8-digit array calculated from the average geomagnetic index (A p ) on the current day and 20 three-hour averages of A before the solution time p . The outputs of the NRLMSISE-00 atmospheric model include the number densities of N2, O2, He, Ar, N, H, O, and the ionospheric positive oxygen ion O + , the neutral atmospheric temperature, and the total atmospheric density.

[0190] In the present invention, as the simulation time progresses, the motion state of the spacecraft during its on-orbit operation and the spatial atomic oxygen and ultraviolet radiation environment in which the spacecraft is located are characterized in real time. The motion state of the spacecraft includes changes in the spatial position, motion speed, and motion attitude of the spacecraft. In the specific implementation manner, the six orbital elements of the initial orbit of the spacecraft specifically include: perigee 400 km, apogee 400 km, inclination 60 degrees, right ascension of the ascending node 0 degrees, argument of perigee 0 degrees, and true anomaly at perigee 0 degrees. The mission start time is 12:00:00 on January 1, 2017. The simulation step size is 20 seconds. The initial attitude orbital quaternion of the spacecraft is described as 1, 0, 0, 0.

[0191] According to the real-time quantification and characterization method in the present invention, the comprehensive real-time quantification and characterization data of atomic oxygen and ultraviolet considering the ultraviolet radiation direction and the positive windward surface are obtained and are successively presented as Figures 10 - 13As shown in the figure. Among them, SAT_1 in the figure represents the name of the satellite. The method for real-time characterization specifically includes:

[0192] Determine the calculation range of real-time simulation according to the start time and end time of the simulation. According to the simulation step size, continuously accumulate the simulation step size from the start time of the simulation to obtain the stepped simulation time. Corresponding to the simulation time, solve the real-time quantitative characterization of the space atomic oxygen and ultraviolet radiation environment at each moment;

[0193] For the calculation of each moment, first obtain the direction of ultraviolet radiation in space through the method for obtaining the incident direction of ultraviolet electromagnetic radiation described above, as shown in Figure 13 shown. Then, for each operating satellite at the same moment, use the method for calculating the spatial position and velocity of the spacecraft described above to calculate the three-dimensional position coordinates and operating velocity parameters in the inertial system, as shown in Figure 11 shown. Use the method for describing the attitude motion of the satellite with quaternions described above to obtain the quaternion attitude parameters of the satellite. Use the coordinate system transformation method described above to transform the position of the satellite in the inertial system to the geodetic coordinates under WGS-84 to obtain the two-dimensional coordinates of the spacecraft. Connect them in chronological order to obtain the sub-satellite point track of the spacecraft. Use the geodetic coordinates in the calculation of the atomic oxygen number density in the atomic oxygen calculation model NRLMSISE00, as shown in Figure 12 shown. It is the real-time change curve of atomic oxygen. In the figure, CM-3 represents per cubic centimeter (cm -3 ). Calculate the real-time change quantitative characterization at each moment in turn to obtain the data that needs to be calculated in this embodiment, so as to solve the problem of real-time quantitative characterization of atomic oxygen and ultraviolet for spacecraft operating in space.

[0194] Although the present disclosure is disclosed as above, the protection scope of the present disclosure is not limited thereto. Those skilled in the art can make various changes and modifications without departing from the spirit and scope of the present disclosure, and these changes and modifications will all fall within the protection scope of the present invention.

Claims

1. A real-time characterization method for the simulation process of the combined action of space atomic oxygen and ultraviolet irradiation, characterized in that Including: Determine the calculation range of real-time simulation according to the simulation start time and the simulation end time; According to the simulation step size, continuously accumulate the simulation step size from the simulation start time to obtain the stepped simulation time; Solve the real-time quantitative characterization of the comprehensive environment of space atomic oxygen and ultraviolet radiation at each simulation time, specifically including: Obtain the incident direction of solar ultraviolet electromagnetic radiation in space, For each operating satellite at the same simulation time, obtain the three-dimensional coordinates and operating speed parameters of the satellite in the inertial system, and the three-dimensional coordinates and the operating speed parameters are used to characterize the position and operating speed of the space vehicle; Obtain the quaternion attitude parameters of the satellite, and the quaternion attitude parameters are used to characterize the operating attitude of the space vehicle; Obtain the atomic oxygen number density according to the atomic oxygen calculation model, and the atomic oxygen number density is used to characterize the space atomic oxygen environment; According to the operating attitude of the space vehicle and the incident direction of the solar ultraviolet electromagnetic radiation, analyze the surface of the space vehicle affected by the ultraviolet electromagnetic radiation, and according to the operating speed and operating attitude of the space vehicle and the space atomic oxygen environment, analyze the atomic oxygen action received by the space vehicle.

2. The real-time characterization method for the simulation process of the combined action of space atomic oxygen and ultraviolet irradiation according to claim 1, wherein, The method for obtaining the incident direction of the solar ultraviolet electromagnetic radiation includes: assuming that the sunlight received by the space vehicle operating near the earth is parallel light, obtaining the space position coordinates of the sun at the simulation time in the J2000 geocentric coordinate system according to the DE430 ephemeris published by JPL, and obtaining the direction vector from the earth center to the solar centroid, taking the inverse of the direction vector and normalizing it to obtain the direction vector of the solar ray in the J2000 geocentric coordinate system, and the direction vector is the incident direction of the solar ultraviolet electromagnetic radiation.

3. The real-time characterization method for the simulation process of the combined action of space atomic oxygen and ultraviolet irradiation according to claim 1, characterized in that The method for obtaining the position and operating speed of the space vehicle includes: obtaining the orbital elements of the space vehicle operating in orbit according to the two-body motion equation and Kepler's three laws, and obtaining the position and operating speed of the space vehicle in the inertial system according to the orbital elements.

4. The real-time characterization method for the simulation process of the combined action of space atomic oxygen and ultraviolet irradiation according to claim 1, characterized in that The obtaining of the quaternion attitude parameters of the satellite includes: Define the orbital coordinate system as O0X0Y0Z0, and the direction vector of the rotation axis e in the orbital coordinate system is e = (C x , C y , C z ), where (C x , C y , C z ) represents the direction cosine. The orbital coordinate system rotates by an angle α around the rotation axis e to obtain the satellite body coordinate system O b X b Y b Z b . Then the quaternion attitude parameter q of the satellite is: where q1, q2, and q3 are vector components, q0 is a scalar component, and q1 2 + q2 2 + q3 2 + q0 2 = 1.

5. The real-time characterization method for the simulation process of the combined action of space atomic oxygen and ultraviolet irradiation according to claim 1, characterized in that, The atomic oxygen calculation model includes the NRLMSISE-00 atmospheric model, and the inputs of the NRLMSISE-00 atmospheric model include the number of days from January 1 of the current year to the current day, the number of seconds from 00:00:00 of the current day to the solution time, geographical longitude, geographical latitude, altitude, the solar 10.7 cm radiation flux of the previous day, the average solar 10.7 cm radiation flux of 81 days, and an 8-digit array obtained according to the average geomagnetic index of the current day and 20 3-hour average geomagnetic indices before the solution time.

6. The real-time characterization method for the simulation process of the combined action of space atomic oxygen and ultraviolet irradiation according to claim 5, characterized in that The geographical longitude and the geographical latitude are obtained by coordinate system transformation, and the three-dimensional coordinates of the satellite in the inertial system are transformed into the geodetic height coordinates in the WGS-84 coordinate system.

7. The real-time characterization method for the simulation process of the combined action of space atomic oxygen and ultraviolet irradiation according to claim 6, characterized in that The coordinate system transformation includes the transformation between the J2000 coordinate system and the Earth-centered Earth-fixed coordinate system, and the overall transformation matrix HG from the J2000 coordinate system to the Earth-centered Earth-fixed coordinate system is determined according to the second formula, and the second formula is: HG = (PR)(NR)(ER)(EP); wherein, PR is the precession matrix, NR is the nutation matrix, ER is the Earth rotation matrix, and EP is the polar motion matrix; The precession matrix is obtained through the third formula, and the third formula is: PR = R z (-z A )R y (-θ A )R z (-ζ A ); Where: R z represents the Euler angle coordinate transformation of the Z axis of the J2000 coordinate system, R y is the Euler angle coordinate transformation of the Y axis of the J2000 coordinate system, z A , θ A , ζ A is the precession constant, which is calculated by the fourth formula. The fourth formula is: In the fourth formula: T is the Julian century number starting from J2000.0; The nutation matrix is obtained through the fifth formula, and the fifth formula is: NR = R X (-ε A )R z (-Δψ)R X (-ε A +Δε); In the fifth formula: ε A is the mean obliquity of the ecliptic at epoch, Δψ is the nutation in longitude, Δε is the nutation in obliquity, and R X is the Euler angle coordinate transformation for the X-axis of the J2000 coordinate system; The Earth rotation matrix is obtained through the tenth formula, and the tenth formula is: ER = R z (θ G ); In the tenth formula, θ G is Greenwich sidereal time; The polar motion matrix is obtained through the thirteenth formula, and the thirteenth formula is: EP = R Y (-x p )R X (-y p ); where x p and y p are geodetic polar coordinates.

8. The real-time characterization method for the simulation process of the combined action of space atomic oxygen and ultraviolet irradiation according to claim 7, characterized in that The coordinate system transformation further includes the transformation between the geocentric geodetic coordinate system and the Earth-centered, Earth-fixed coordinate system. The geocentric geodetic coordinate system includes the WGS-84 coordinate system. Suppose the coordinates of the target in the WGS-84 coordinate system are (lon, lat, alt). Determine the coordinates (x d , y d , z d ) of the target in the Earth-centered, Earth-fixed coordinate system according to the fourteenth formula. The fourteenth formula is: In the fourteenth formula: e 2 =(a 2 -b 2 ) / a 2 , where a is the semi-major axis of the WGS-84 ellipsoid corresponding to the WGS-84 coordinate system, b is the semi-minor axis of the WGS-84 ellipsoid, N is the normal length of the WGS-84 ellipsoid, and e is the eccentricity of the WGS-84 ellipsoid. Among them, 9. The real-time characterization method for the simulation process of the combined action of space atomic oxygen and ultraviolet irradiation according to claim 5, characterized in that The output of the NRLMSISE-00 atmospheric model includes the number densities of N2, O2, He, Ar, N, H, O, and the ionospheric positive oxygen ion O + , the neutral atmospheric temperature, and the total atmospheric density.

10. The real-time characterization method for the simulation process of the combined action of space atomic oxygen and ultraviolet irradiation according to claim 1, characterized in that The time is represented by the Julian day, and the Julian day is calculated according to the first formula, and the first formula includes: where: JD represents the Julian day, INT represents taking the integer, Year represents the year in the Gregorian calendar, Month represents the month, Day represents the day, Hour represents the hour, Minute represents the minute, and Second represents the second.

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