A satellite fuel evaluation method and device

Abstract

The present invention discloses a satellite fuel assessment method and device, the method comprising: determining thruster thrust data during the test of the satellite electronic integration setting to be estimated; performing the first precise orbit determination on the satellite to be measured; performing the first east-west position keeping on the satellite to be measured, recording the first thruster group used, the first number of ignition pulses, the first ignition control cycle and the first total ignition time; performing the second precise orbit determination on the satellite to be measured; solving the first speed increment of the first east-west position keeping based on the data of the first precise orbit determination and the second precise orbit determination; calculating the first remaining fuel mass of the satellite to be tested based on the first speed increment, the dry weight of the satellite to be tested, and the thrust of the first thruster group. The satellite fuel assessment method provided in the embodiment of the present application has a high accuracy in assessing the remaining satellite fuel.

Description

Technical Field

[0001] The invention belongs to the technical field of spacecraft orbit design, and in particular relates to a satellite fuel evaluation method and device. Background Art

[0002] The construction, launch, and operation and management of geostationary communication satellites (hereinafter referred to as communication satellites or satellites) cost tens of billions of yuan. It is necessary to control launch risks and coordinate orbital location resources. Therefore, how to make the best use of an in-orbit communication satellite and maximize its commercial value is the goal pursued by communication satellite operators around the world.

[0003] At present, the main factor restricting the actual on-orbit life of communication satellites is fuel exhaustion. At the same time, the availability and reliability of satellite platform equipment and payloads must also be considered. It is necessary to maximize the effectiveness of the limited satellite life, and at the same time ensure that the satellite is properly disposed of at the end of its life, in accordance with international and domestic requirements for the deorbiting of communication satellites, and avoid becoming space junk in geostationary orbit. Therefore, the remaining mass of onboard fuel is one of the core data that satellite operators pay attention to. The current mainstream methods for evaluating satellite onboard fuel include the bookkeeping method, the PVT (Pressure, Volume, Temperature) method, and the specific heat capacity method. The above methods all measure the relevant data through on-board sensors and perform data modeling to calculate the remaining mass of satellite fuel.

[0004] The existing method of calculating the mass of remaining fuel of a satellite will lead to errors in calculating the remaining fuel of the satellite when the on-board sensors are damaged or the data modeling is incorrect. Summary of the invention

[0005] The embodiment of the present invention provides a method for calculating the remaining fuel of a satellite using orbital dynamics, so as to solve the problem in the prior art that the satellite's remaining fuel is calculated incorrectly due to damage to the onboard sensor or data modeling errors.

[0006] In order to solve the above technical problems, the present invention discloses a satellite fuel evaluation method, wherein the method comprises: determining thruster thrust data during the electronic integration setting test of the satellite to be estimated; performing the first precise orbit determination on the satellite to be tested; performing the first east-west position keeping on the satellite to be tested, recording the first thruster group used, the first ignition pulse number, the first ignition control cycle and the first total ignition time; performing the second precise orbit determination on the satellite to be tested; solving the first velocity increment of the first east-west position keeping based on the data of the first precise orbit determination and the second precise orbit determination; calculating the first remaining fuel mass of the satellite to be tested based on the first velocity increment, the dry weight of the satellite to be tested and the thrust of the first thruster group.

[0007] Optionally, the method also includes: performing a second east-west position keeping on the satellite to be tested, recording the second thruster group used, the second number of ignition pulses, the second ignition control cycle and the second total ignition duration; performing a third precise orbit determination on the satellite to be tested; solving a second velocity increment of the second east-west position keeping based on the data of the second precise orbit determination and the third precise orbit determination; calculating a second remaining fuel mass of the satellite to be tested based on the third precise orbit determination data and the velocity increment; and estimating the remaining fuel mass calculation error.

[0008] Optionally, the step of estimating the calculation error of the remaining fuel mass includes: respectively estimating the thrust error of the thruster, the thruster ignition duration error, the semi-major axis of the orbit and the tangential position-keeping maneuvering error; and determining the estimated calculation error of the remaining fuel mass based on the thrust error of the thruster, the thruster ignition duration error, the semi-major axis of the orbit and the tangential position-keeping maneuvering error.

[0009] Optionally, the step of estimating the thrust error of the thruster includes: forming two simultaneous equations by collecting two consecutive east-west position control data; solving the optimal solution of the two thruster efficiencies and the target satellite mass that satisfies the minimum residual of the two simultaneous equations by a nonlinear least squares solution method; and determining the modulus of the minimum residual as the thrust error of the thruster.

[0010] Optionally, the orbit semi-major axis error is the sum of the orbit determination accuracy error and the orbit extrapolation error.

[0011] Optionally, the method for estimating the thruster ignition duration error includes at least one of the following: a successive accumulation method of ignition pulse errors and a telemetry equivalent recording method of ignition pulse duration.

[0012] In order to solve the above technical problems, the present invention discloses a satellite fuel evaluation device, wherein the device includes: a determination module, used to determine the thruster thrust data during the electronic integrated setting test of the satellite to be estimated; a first orbit determination module, used to perform the first precise orbit determination on the satellite to be tested; a first holding module, used to perform the first east-west holding on the satellite to be tested, and record the first thruster group used, the first ignition pulse number, the first ignition control cycle and the first total ignition time; a second orbit determination module, used to perform the second precise orbit determination on the satellite to be tested; a first solution module, used to solve the first speed increment of the first east-west holding based on the data of the first precise orbit determination and the second precise orbit determination; a first calculation module, used to calculate the first remaining fuel mass of the satellite to be tested based on the first speed increment, the dry weight of the satellite to be tested, and the thrust of the first thruster group.

[0013] Optionally, the device also includes: a second position keeping module, used to perform a second east-west position keeping on the satellite to be tested, and record the second thruster group used, the second number of ignition pulses, the second ignition control cycle and the second total ignition duration; a second orbit determination module, used to perform a third precise orbit determination on the satellite to be tested; a second solution module, used to solve the second speed increment of the second east-west position keeping based on the data of the second precise orbit determination and the third precise orbit determination; a second calculation module, used to calculate the second remaining fuel mass of the satellite to be tested based on the third precise orbit determination data and the speed increment; and an estimation module, used to estimate the calculation error of the remaining fuel mass.

[0014] Optionally, the estimation module includes: a first submodule, used to estimate the thrust error of the thruster, the thruster ignition duration error, the semi-major axis and the tangential position-keeping maneuvering error respectively; and a second submodule, used to determine the estimated remaining fuel mass calculation error based on the thrust error of the thruster, the thruster ignition duration error, the semi-major axis of the orbit and the tangential position-keeping maneuvering error.

[0015] Optionally, when the first submodule estimates the thrust error of the thruster, it is specifically used to: form two simultaneous equations by collecting two consecutive east-west position control data; solve the optimal solution of the two thruster efficiencies and the target satellite mass that satisfies the minimum residual of the two simultaneous equations through a nonlinear least squares solution method; and determine the modulus of the minimum residual as the thruster thrust error.

[0016] The satellite fuel evaluation scheme disclosed in the embodiment of the present invention determines the thruster thrust data during the test of the satellite electronic integration setting to be estimated; performs the first precise orbit determination on the satellite to be measured; performs the first east-west position protection on the satellite to be measured, records the first thruster group used, the first ignition pulse number, the first ignition control cycle and the first ignition total duration; performs the second precise orbit determination on the satellite to be measured; based on the data of the first precise orbit determination and the second precise orbit determination, solves the first speed increment of the first east-west position protection; based on the first speed increment, the dry weight of the satellite to be measured, and the thrust of the first thruster group, calculates the first remaining fuel mass of the satellite to be tested, and the remaining fuel can be estimated without relying on the on-board sensor. This satellite fuel evaluation method can avoid the problem of being unable to calculate the remaining fuel due to the failure of the on-board sensor, and the estimated satellite fuel remaining amount has high accuracy. BRIEF DESCRIPTION OF THE DRAWINGS

[0017] Figure 1 A flowchart of the steps of the satellite fuel assessment method provided by an embodiment of the present invention;

[0018] Figure 2 This is a schematic diagram of the structure of a satellite fuel evaluation device provided in an embodiment of the present invention. DETAILED DESCRIPTION

[0019] The present invention will be described in detail below based on specific embodiments and in conjunction with the accompanying drawings. It should be understood that the specific embodiments described herein are only used to explain the present invention and are not used to limit the present invention.

[0020] like Figure 1 As shown, the satellite fuel evaluation method of the embodiment of the present invention includes the following steps:

[0021] Step 101: Determine thruster thrust data during a satellite electronic integration setting test to be estimated.

[0022] An exemplary thruster thrust data is shown in Table 1:

[0023] Table 1

[0024] Jet Fx(cow) Fy(cow) Fz(cow) 2A 9.227 1.423 -3.583 3A 9.528 -1.589 2.587 4A -9.53 1.577 2.587 5A -9.233 -1.427 -3.567 6A 0 -9.136 -4.066 7A -0.004 -9.51 3.091 2B -9.527 -1.579 2.596 3B -9.229 1.425 -3.577 4B 9.23 -1.429 -3.572 5B 0.525 1.594 2.595 6B 0 9.511 3.088 7B -0.006 9.141 -4.055

[0025] The thrust in this table is defined in the satellite body coordinate system: X is the direction of motion (approximately east); Y is approximately south; and Z points to the center of the earth.

[0026] By analyzing the thrust data of the thruster, it can be seen that the thrust output of the thruster has a gradual stabilization process from small to large, and the thrust has a conversion efficiency curve according to the ignition pulse duration. The curve changes in the trend that when the ignition pulse is in the range of 0 to 200ms, the thrust efficiency is linear, and the ignition pulse can reach 100% when it is 200ms. When the ignition pulse is greater than 200ms, the thrust efficiency is constant at 100%.

[0027] In the constant thrust output range, the satellite will adopt a real-time ignition shutdown strategy due to attitude control needs. The number of ignition shutdowns and the total duration are calculated in real time based on the attitude control needs at the time. The telemetry does not save the thruster's entire working parameters, which makes ground thrust modeling difficult during steady-state ignition.

[0028] However, in the normal east-west position-keeping scenario, the thruster's single ignition pulse width remains constant at 256ms, and the thruster's operating control is relatively stable in the position-keeping mode. Unignited pulses can be detected through telemetry, which is beneficial to ground thrust modeling.

[0029] Step 102: Perform the first precise orbit determination on the satellite to be measured.

[0030] Step 103: Perform the first east-west position keeping of the satellite to be tested, and record the first thruster group used, the first ignition pulse number, the first ignition control period and the first total ignition time.

[0031] Table 2 is an exemplary table of working parameters of a certain satellite thruster:

[0032]

[0033] It should be noted that this table is for calculation verification. For other satellite operating parameters, please refer to the satellite binding parameter document.

[0034] Effective ignition time of chemical thruster of a certain satellite platform T ON :

[0035] Normal Mode:

[0036]

[0037] Step 104: Perform a second precise orbit determination on the satellite to be measured.

[0038] Step 105: Based on the data of the first precise orbit determination and the second precise orbit determination, the first velocity increment of the first east-west position protection is solved.

[0039] The first speed increment can be expressed as ΔV T , ΔV R , ΔV N .

[0040] The known conditions are as follows:

[0041] 1. The number of orbital elements at the orbit determination epoch (T0) after position preservation.

[0042] 2. Extrapolate the orbit elements before the position-keeping orbit determination to the orbit elements after the position-keeping orbit determination epoch (T0).

[0043] The onboard control engine usually uses a single and dual hydrazine engine, which has a thrust of the Newton order. The time of each ignition control is very small relative to the 24h orbital period, and can usually be considered as a pulse. Another possible control engine is an electric rocket, which has a very small thrust. The error of the pulse approximation may be slightly larger, which can be corrected. For pulse orbit change, the integration of the perturbation equation can be obtained:

[0044]

[0045]

[0046]

[0047]

[0048]

[0049]

[0050] Where, ΔV T , ΔV R , ΔV N It is the component of velocity increment in the tangential T, radial R and normal N directions. Sis the stationary orbital velocity. The tangential velocity component ΔV T What changes are the semi-major axis and eccentricity of the satellite orbital elements, and the radial velocity component ΔV R Change the eccentricity and longitude in the satellite orbital elements, the normal velocity component ΔV N What changes is the inclination vector in the satellite orbit element. It can be found that the changes in the semi-major axis and inclination orbit parameters are only related to ΔV T , ΔV N One-to-one correspondence, no coupling effect with other velocity increments.

[0051] According to formulas (1), (4) and (5), the satellite tangential maneuver Dvt is obtained by solving the actual change of the semi-major axis before and after the east-west position preservation; the satellite normal maneuver Dvn is obtained by solving the actual change of the inclination angle vector before and after the north-south position preservation.

[0052] Step 106: Calculate a first remaining fuel mass of the satellite to be tested based on the first velocity increment, the dry weight of the satellite to be tested, and the thrust of the first thruster group.

[0053] The velocity increment (Dv) (m / s) determined by the orbit determination data; thruster thrust (F) (N), where the thruster thrust comes from the satellite binding parameters and is linked to the thruster configuration; Mi: the satellite dry weight (Kg) can be entered into the following formula to calculate the remaining fuel mass of the satellite:

[0054] M f =(F*T on ) / Dv-M i ——(7)

[0055] Among them, M f To obtain the remaining fuel mass (after ignition), the unit is (kg).

[0056] The satellite fuel assessment method provided in the embodiment of the present application is that the satellite performs two east-west position protections in the normal mode, and performs precise orbit determination before and after the satellite position protection, then calculates the satellite orbit semi-major axis increment based on the precise orbit determination result, and then calculates the satellite total mass range and thruster efficiency in combination with parameters such as thruster ignition duration and thruster nominal thrust, and finally subtracts the satellite dry weight from the satellite total mass to obtain the satellite remaining fuel mass, which can avoid the problem of being unable to calculate the remaining fuel due to a sensor failure on the satellite, and the assessed satellite remaining fuel amount is highly accurate.

[0057] In an optional embodiment, the satellite fuel assessment method provided in the embodiment of the present application further includes the following steps:

[0058] S1: Perform the second east-west position keeping of the satellite to be tested, and record the second thruster group used, the second ignition pulse number, the second ignition control cycle and the second total ignition time;

[0059] S2: Conduct the third precise orbit determination of the satellite to be measured;

[0060] S3: based on the data of the second precise orbit determination and the third precise orbit determination, solving the second velocity increment of the second east-west position protection;

[0061] S4: Calculate the second remaining fuel mass of the satellite to be tested based on the third precise orbit determination data and velocity increment;

[0062] S5: Estimated remaining fuel mass calculation error.

[0063] This method of estimating the satellite's remaining fuel mass by performing multiple precise orbit determinations on the satellite to be tested is more accurate and reliable than estimating the satellite's remaining fuel mass by performing only two precise orbit determinations.

[0064] In an optional embodiment, the step of estimating the remaining fuel mass calculation error may include:

[0065] The thruster thrust error, thruster ignition duration error, semi-major axis and tangential position-keeping maneuvering errors are estimated respectively; based on the thruster thrust error, thruster ignition duration error, orbital semi-major axis and tangential position-keeping maneuvering errors, the calculation error of the estimated remaining fuel mass is determined.

[0066] The error in the remaining fuel mass can be estimated using the following error function:

[0067]

[0068]

[0069] Among them, dF is the thruster error, dTon is the thruster ignition duration error, and d(Dvt) is the major axis da and tangential position-keeping maneuver error.

[0070] In an optional embodiment, the method of estimating the thrust error of the thruster may be as follows:

[0071] Two simultaneous equations are formed by collecting two consecutive east-west position control data;

[0072] The optimal solution of the two thruster efficiencies and the target satellite mass that satisfies the minimum residual of the two simultaneous equations is obtained by using the nonlinear least squares solution method.

[0073] The magnitude of the minimum residual is determined as the thruster thrust error.

[0074] The thruster error dF can be estimated as follows:

[0075] Assuming that the east-west position-protecting main thrusters are 2B and 3B, the attitude control impulses generated by the stabilizing thrust F2B and F3B attitude control thrusters 2B, 3B, 4B, and 5B cancel each other out in the east-west direction and are approximately 0 and can be ignored. The thruster efficiency is η.

[0076] There are several data processing algorithms as follows:

[0077] 1) Double position protection method:

[0078] By collecting two consecutive east-west position control data, two simultaneous equations are formed to solve two unknown quantities, including the two thruster efficiencies η and the final satellite mass M0, which are substituted into the following set of equations for solution.

[0079]

[0080]

[0081] Where T 2B1 、T 2B0 , T 3B1 , T 3B2 represents the ignition duration of the first and second positioning maneuvers thrusters 2B and 3B, which are obtained from the telemetry data as known quantities; ΔV0 and ΔV1 represent the maneuvering momentum of the two positioning maneuvers orbit determination solutions, which are obtained from the orbit determination data as known quantities; F 2B , F 3B As the known thruster steady thrust, obtained from the parameter binding file.

[0082] ΔM1 represents the theoretical fuel consumption of the second position-keeping maneuver,

[0083]

[0084] Isp=275 g0=9.80665

[0085] The optimal solution of η and M0 that satisfies the minimum residual of the simultaneous equations (9), (10) and (11) can be solved by the nonlinear least squares solution method. The modulus of the minimum residual corresponds to the deviation of the physical quantity Dv, which can be used to estimate the error of this solution.

[0086] 2) Three-way position protection method:

[0087] This method is to model the satellite thruster thrust separately and solve the thrust as an unknown variable.

[0088] By collecting three consecutive east-west position control data, three simultaneous equations are formed to solve three unknown quantities, including the two thrusters F 2B , F 3B The thrust and the final satellite mass M0 are substituted into the following set of equations and solved.

[0089]

[0090]

[0091]

[0092] Where T 2B1 , T 2B2 , T 2B3 , T 3B1 , T 3B2 , T 3B3 =ΔV1, ΔV2, ΔV3 represent the actual maneuvering amount of the three-position-keeping maneuvers, which are obtained from the orbit determination data as known quantities. ΔM1 represents the theoretical fuel consumption of the second-position-keeping maneuver.

[0093]

[0094]

[0095] Isp=275×0.92

[0096]

[0097] In an optional embodiment, the method for estimating the thruster ignition duration error includes at least one of the following: a method of successive accumulation of ignition pulse errors and a method of telemetry equivalent recording of ignition pulse duration.

[0098] The estimation method of thruster ignition time error dTon is as follows:

[0099] Ignition pulse error successive accumulation method:

[0100] The thruster ignition duration is in seconds, and according to the number of significant digits, the error limit is less than or equal to 0.0005 seconds. Assuming that the thruster Dv = 0.1 m / s, the thruster ignition pulse is about 30 times, and the total dTon = 2*0.0005*30 = 0.03 seconds.

[0101] Ignition pulse duration telemetry equivalent recording method:

[0102] The ignition duration count of each thruster is transmitted by telemetry, with a resolution equivalent to 2ms. The thruster ignition duration Ton is obtained by recording the difference between the corresponding thruster ignition duration counts before and after the position protection, dTon = 0.002 seconds.

[0103] The estimation method of the semi-major axis da and the tangential position-keeping maneuver error d(Dvt) is described as follows:

[0104] Orbit semi-major axis error = orbit determination accuracy error + orbit extrapolation error;

[0105] Taking the synchronous orbit satellite fixed at 125 degrees east longitude as an example, according to the current geographical distribution of three stations of a certain orbit determination system, the ranging noise of each station is 18 meters, the system offset is 60 meters, and one ranging data is saved for 1 minute. The ranging time is 6, 12, and 24 hours respectively, and the six orbital elements, the system differences of station one and station three are solved at the same time. By applying the covariance matrix analysis method, the orbit determination error analysis data of synchronous orbit satellites can be obtained. According to the calculation under the above conditions, as long as 24 hours of ranging and orbit determination are carried out, the orbit determination error of the semi-major axis of the satellite orbit can be controlled within 3 meters, and the orbit determination accuracy of the semi-major axis is higher than the orbit determination accuracy of the inclination angle. Therefore, the east-west position is preferred to ensure the implementation of the satellite remaining fuel assessment.

[0106] Precision ephemeris deduction and orbit determination software need to use a consistent orbital dynamics model with the same model error, truncation error and rounding error. In this way, when calculating the orbital parameter difference, the fixed deviation introduced by the software tool is offset. In the dynamics model, it is recommended to use the Earth gravity field model of order 21×21 or above, the JPL-DE405 solar system sun, moon and planet digital ephemeris, enable polar motion, leap second, UTC-UT1 time difference correction and solar light pressure and other perturbation acceleration corrections, and the extrapolated ephemeris position error of one day is about 1 meter, and the maximum error converted to the semi-major axis of the orbit is about 2 meters.

[0107] Figure 2 The structure block diagram of a satellite fuel evaluation device provided by an embodiment of the present invention is shown in FIG. Figure 2 As shown, the satellite fuel evaluation device of the embodiment of the present invention includes the following modules: a determination module 201, used to determine the thruster thrust data during the test of the electronic integrated setting of the satellite to be estimated; a first orbit determination module 202, used to perform the first precise orbit determination on the satellite to be tested; a first holding module 203, used to perform the first east-west holding on the satellite to be tested, and record the first thruster group used, the first ignition pulse number, the first ignition control cycle and the first ignition total duration; a second orbit determination module 204, used to perform the second precise orbit determination on the satellite to be tested; a first solving module 205, used to solve the first speed increment of the first east-west holding based on the data of the first precise orbit determination and the second precise orbit determination; a first calculation module 206, used to calculate the first remaining fuel mass of the satellite to be tested based on the first speed increment, the dry weight of the satellite to be tested, and the thrust of the first thruster group.

[0108] Optionally, the device also includes: a second position keeping module, used to perform a second east-west position keeping on the satellite to be tested, and record the second thruster group used, the second number of ignition pulses, the second ignition control cycle and the second total ignition time; a second orbit determination module, used to perform a third precise orbit determination on the satellite to be tested; a second solution module, used to solve the second velocity increment of the second east-west position keeping based on the data of the second precise orbit determination and the third precise orbit determination; a second calculation module, used to calculate the second remaining fuel mass of the satellite to be tested based on the third precise orbit determination data and the velocity increment; an estimation module, used to estimate the remaining fuel mass calculation error.

[0109] Optionally, the estimation module includes: a first submodule, used to estimate the thrust error of the thruster, the thruster ignition duration error, the semi-major axis and the tangential position-keeping maneuvering error respectively; and a second submodule, used to determine the estimated remaining fuel mass calculation error based on the thrust error of the thruster, the thruster ignition duration error, the semi-major axis of the orbit and the tangential position-keeping maneuvering error.

[0110] Optionally, when the first submodule estimates the thrust error of the thruster, it is specifically used to: form two simultaneous equations by collecting two consecutive east-west position control data; solve the optimal solution of the two thruster efficiencies and the target satellite mass that satisfy the minimum residual of the two simultaneous equations through a nonlinear least squares solution method; and determine the modulus of the minimum residual as the thruster thrust error.

[0111] The satellite remaining fuel assessment device provided in the embodiment of the present application can be used to execute the relevant processes of the above-mentioned satellite remaining fuel assessment method. The relevant parts between the two can be referred to the relevant description in the above-mentioned satellite remaining fuel assessment method embodiment, which will not be repeated here.

[0112] It should be noted that the above description is only a preferred embodiment of the present invention. It should be understood that for those skilled in the art, several changes and improvements can be made without departing from the technical concept of the present invention, which are all included in the protection scope of the present invention.

[0113] The various embodiments in this specification are described in a progressive manner, and each embodiment focuses on the differences from other embodiments. The same or similar parts between the various embodiments can be referenced to each other.

[0114] The above description is only the best specific implementation mode of the present invention, but the protection scope of the present invention is not limited thereto. Any changes or substitutions that can be easily thought of by any technician familiar with the technical field within the technical scope disclosed by the present invention should be covered within the protection scope of the present invention.

[0115] The contents not described in detail in the specification of the present invention belong to the common knowledge of those skilled in the art.

Claims

1. A satellite fuel assessment method, characterized in that: The method comprises: Determine thruster thrust data during satellite electronics integration setup testing to be estimated; Carry out the first precise orbit determination for the satellite to be measured; Perform the first east-west position keeping of the satellite to be tested, and record the first thruster group used, the first ignition pulse number, the first ignition control cycle and the first total ignition time; Conduct a second precise orbit determination for the satellite to be measured; Based on the data of the first precise orbit determination and the second precise orbit determination, solving the first velocity increment of the first east-west position protection; A first remaining fuel mass of the satellite to be tested is calculated based on the first speed increment, the dry weight of the satellite to be tested, and the thrust of the first thruster group.

2. The method according to claim 1, characterized in that The method further comprises: Perform a second east-west position maintenance on the satellite to be tested, and record the second thruster group used, the second ignition pulse number, the second ignition control cycle, and the second total ignition time; Carry out the third precise orbit determination for the satellite to be measured; Based on the data of the second precise orbit determination and the third precise orbit determination, solving the second velocity increment of the second east-west position protection; Calculating a second remaining fuel mass of the satellite to be tested based on the third precise orbit determination data and the velocity increment; Estimated remaining fuel mass calculation error.

3. The method according to claim 2, characterized in that The step of estimating the remaining fuel mass calculation error comprises: Estimate thruster thrust error, thruster ignition time error, semi-major axis and tangential position-keeping maneuver errors respectively; Based on the thruster thrust error, the thruster ignition duration error, the orbit semi-major axis and the tangential position-keeping maneuvering error, the estimated remaining fuel mass calculation error is determined.

4. The method according to claim 3, characterized in that The step of estimating the thrust error of the thruster comprises: Two simultaneous equations are formed by collecting two consecutive east-west position control data; By using a nonlinear least squares solution method, an optimal solution of the two thruster efficiencies and the target satellite mass that satisfies the minimum residual of the two simultaneous equations is obtained; The modulus of the minimum residual is determined as the thruster thrust error.

5. The method according to claim 3, characterized in that: The orbit semi-major axis error is the sum of the orbit determination accuracy error and the orbit extrapolation error.

6. The method according to claim 3, characterized in that The method of estimating the thruster ignition duration error includes at least one of the following: Successive accumulation method of ignition pulse error and telemetry equivalent recording method of ignition pulse duration.

7. A satellite fuel evaluation device, characterized in that: The device comprises: a determination module, for determining thruster thrust data during a satellite electronic integration setting test to be estimated; The first orbit determination module is used to perform the first precise orbit determination on the satellite to be measured; The first position keeping module is used to perform the first east-west position keeping of the satellite to be tested, and record the first thruster group used, the first ignition pulse number, the first ignition control cycle and the first total ignition time; The second orbit determination module is used to perform a second precise orbit determination on the satellite to be measured; A first solving module is used to solve a first velocity increment of a first east-west position protection based on data of the first precise orbit determination and the second precise orbit determination; The first calculation module is used to calculate a first remaining fuel mass of the satellite to be tested based on the first speed increment, the dry weight of the satellite to be tested, and the thrust of the first thruster group.

8. The device according to claim 7, characterized in that The device also includes: The second position-keeping module is used to perform the second east-west position-keeping of the satellite to be tested, and record the second thruster group used, the second ignition pulse number, the second ignition control cycle, and the second total ignition time; The second orbit determination module is used to perform the third precise orbit determination on the satellite to be measured; A second solving module is used to solve a second velocity increment of a second east-west position protection based on data of the second precise orbit determination and the third precise orbit determination; A second calculation module is used to calculate a second remaining fuel mass of the satellite to be tested based on the third precise orbit determination data and the velocity increment; The estimation module is used to estimate the remaining fuel mass calculation error.

9. The device according to claim 8, characterized in that The estimation module comprises: The first submodule is used to estimate thruster thrust error, thruster ignition time error, semi-major axis and tangential position-keeping maneuvering error respectively; The second submodule is used to determine the calculation error of the estimated remaining fuel mass based on the thrust error of the thruster, the ignition time error of the thruster, the semi-major axis of the orbit and the tangential position-keeping maneuvering error.

10. The device according to claim 9, characterized in that When the first submodule estimates the thrust error of the thruster, it is specifically used to: Two simultaneous equations are formed by collecting two consecutive east-west position control data; By using a nonlinear least squares solution method, an optimal solution of the two thruster efficiencies and the target satellite mass that satisfies the minimum residual of the two simultaneous equations is obtained; The modulus of the minimum residual is determined as the thruster thrust error.

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