Method for determining the amount of non-isobaric air cushion propellant in parallel tanks and the center of mass of satellite

Abstract

The present invention relates to a method for determining the amount of non-isobaric air cushion propellant filling in parallel tanks and the satellite mass center, and the steps include: obtaining input parameters through experiments, tests and simulations; carrying out balanced filling to obtain the propellant filling mass, propellant mass center and satellite mass center coordinates of each tank in the launch state during balanced filling; calculating the mass center adjustment capacity during unbalanced filling, and sequentially obtaining the propellant filling mass, propellant mass center and satellite mass center coordinates of each tank when each tank is filled according to the maximum filling mass; calculating the intersection of the engine thrust vector and the XY plane at the satellite mass center; based on the determined propellant filling amount, air cushion pressure and satellite mass center, performing propellant and helium filling and subsequent on-orbit operations. The present invention introduces a satellite mass center relationship diagram, which can vividly display the positional relationship between the satellite mass center and the constraint conditions at different times, which is easy to understand and quickly guide engineering applications.

Description

Technical Field

[0001] The invention relates to a method for determining the filling amount of non-isobaric air cushion propellant in parallel tanks and the center of mass of a satellite, and belongs to the overall technical field of missile systems. Background Art

[0002] The design of the propellant filling amount and the satellite's center of mass is an important task after all satellite assembly work is completed and before the satellite is launched. The weight of liquid propellant accounts for a high proportion when the satellite is in the launch state. For example, the weight of geosynchronous orbit satellite propellant accounts for about 40%. Therefore, propellant filling will also significantly affect the mass characteristics of the satellite in the launch state (such as weight, center of mass position, etc.), so the calculation of the satellite propellant filling amount and the determination of the satellite's center of mass position are carried out jointly. Different mass characteristics at launch will further affect the flight safety of the launch vehicle in the active stage of the launch and the satellite attitude stability during the separation of the rocket and the satellite.

[0003] In the past, satellites with bipropellant chemical propulsion systems all adopted a scheme of loading a single propellant in a single tank. The design of the propellant filling amount and the satellite's center of mass is relatively simple. After allocating the filling amounts of the oxidizer and the combustion agent according to the mixing ratio of the chemical propulsion subsystem, the calculation of the filling amount of each tank is completed. When the tank position has been designed in advance, the center of mass of the satellite at launch is also determined with the filling amount. In addition, the mass of the propellant in each tank cannot be adjusted from the launch of the satellite to the start of the satellite orbit change.

[0004] In recent years, with the increase in satellite propellant carrying capacity and the optimization of satellite configuration, more and more bipropellant chemical propulsion system satellites have adopted a parallel tank solution, that is, the oxidizer and the fuel are each loaded in two tanks, and the two tanks discharge propellants simultaneously when in orbit. Compared with a single propellant loaded in a single tank, the propellant filling amount and satellite center of mass design of parallel tanks are more complicated. In addition to determining the filling amount of oxidizer and fuel, it is also necessary to distribute the weight of the same propellant between the two parallel tanks. And the different filling amounts of multiple tanks also have a more complex and variable impact on the center of mass of the satellite during launch.

[0005] In the prior art, some technologies do not quantitatively establish the relationship between the filling amount, the initial center of mass, and the engine interference torque, that is, they do not provide a specific quantitative design method for reducing the interference torque of the orbit change segment by reducing the offset of the satellite's initial center of mass. Although some technologies quantitatively calculate the interference torque, they identify the initial center of mass at the time of satellite launch, that is, the initial center of mass during the orbit change process, that is, they do not consider the difference in satellite mass characteristics at the two states of satellite launch and the initial moment of orbit change. Some technologies only consider the independent control of the propellant filling amount in each tank during launch, and do not consider the ability of the tank air cushion to regulate the propellant mass and the satellite center of mass.

[0006] The satellite tank will not be 100% filled with liquid propellant during launch, but a small portion of the volume will be reserved to be filled with low-density inert gas (such as helium) as an air cushion. The air cushion has two functions: one is to act as a compressible substance to absorb the volume change of the liquid in the tank due to thermal expansion and contraction to ensure the tank pressure is safe, and the other is to form an extrusion force to enable the tank to discharge propellant outward uninterruptedly. The satellite tank generally has two ports, the lower port for filling and discharging propellant, and the upper port for filling gas into the tank. In the parallel tank propulsion system, each port of each tank will be equipped with a valve. Therefore, in addition to being able to independently control the propellant filling amount of each tank, the gas inlet and outlet can also be independently controlled through the upper port valve of each tank, so as to achieve the purpose of filling pressure of different air cushions for each tank.

[0007] The mass characteristic design during satellite launch and the mass characteristic design during satellite orbit change need to consider different constraints. The former needs to meet the flight safety constraints of the launch vehicle and the satellite attitude stability constraints during satellite-rocket separation (the greater the deviation of the satellite's center of mass, the greater the deviation of the satellite's angular velocity after the satellite-rocket separation device launches the satellite, and the more difficult it is for the satellite to establish a stable attitude). The latter needs to meet the attitude stability constraints during satellite orbit change (mainly related to the interference torque of the orbit control engine, that is, the torque of the engine thrust vector on the satellite's center of mass).

[0008] The center of mass adjustment of a traditional satellite loaded with a single propellant and a single tank can only be achieved by adding and changing the position of the counterweight. For parallel tank satellites, in addition to using traditional counterweights, the propellant is similar to a floating, virtual "counterweight", which can adjust the satellite's center of mass within a certain range by distributing the same propellant differently between two parallel tanks. Moreover, by reasonably designing the tank air cushion pressure and opening the relevant valves at the right time, this floating, virtual "counterweight" can adjust its position and weight in real time and in a controlled manner on orbit to further adjust the satellite's center of mass. Summary of the invention

[0009] The technical problem solved by the present invention is: to overcome the shortcomings of the prior art, to propose a method for determining the non-isobaric air cushion propellant filling amount of parallel tanks and the satellite mass center, to adopt a strategy of independently controlling the propellant filling amount and air cushion pressure of each tank of the parallel tanks, to use the air cushion as a driving force for adjusting the distribution of the propellant in different tanks, to respectively design different satellite mass centers in the satellite launch state and at the beginning of the orbit change, and to simultaneously meet the multiple constraints on the satellite mass center of the carrier rocket flight safety in the satellite launch state and active phase, and the satellite attitude stability during the satellite-rocket separation and orbit change.

[0010] The solution of the present invention is:

[0011] A method for determining the amount of parallel tank non-isobaric air cushion propellant and the satellite mass center, the steps comprising:

[0012] S1. Obtain input parameters through experiments, tests and simulations, mainly including the mass and center of mass of the satellite before filling, the performance parameters of the tank and engine and the installation parameters on the satellite, the density of the propellant, the optional number and installation position of the counterweights, and the constraints of the launch vehicle on the satellite center of mass;

[0013] S2. Conduct balanced filling, i.e., preliminary analysis when the weight of the propellants in the parallel tanks is equal, and obtain the propellant filling mass of each tank during balanced filling, the propellant center of mass, and the coordinates of the satellite center of mass in the launch state;

[0014] S3. Calculate the center of mass adjustment capability during non-balanced filling, and sequentially obtain the propellant filling mass, propellant center of mass, and satellite center of mass coordinates of each tank in the launch state when each tank is filled according to the maximum filling mass;

[0015] S4. Calculate the intersection of the engine thrust vector and the XY plane at the satellite mass center according to the conversion relationship between the engine thrust vector and the engine body coordinate system, the engine precision measuring mirror coordinate system and the satellite mechanical coordinate system;

[0016] S5, using the satellite mass center coordinates calculated by S2, S3, and S4, the intersection of the engine thrust vector and the XY plane at the satellite mass center, and the constraints of the launch vehicle on the satellite mass center as input, draw a satellite mass center relationship diagram;

[0017] S6. Determine whether it is necessary to add counterweights and whether it is necessary to unbalance the parallel tanks to ensure that the satellite mass center meets the constraints of the launch vehicle in the launch state through the positional relationship between the points and the domains in the satellite mass center relationship diagram drawn in S5; if it is necessary to add counterweights, determine the number, weight and position of the counterweights;

[0018] S7, select a point in the feasible domain of the satellite mass center relationship diagram drawn in S5 as the satellite X and Y center of mass in the launch state, and calculate the propellant filling amount of each tank and the satellite Z center of mass in the launch state;

[0019] S8. Select a point closest to the engine thrust vector in the feasible domain of the satellite mass center relationship diagram drawn in S5 as the satellite X and Y mass centers at the beginning of the orbit change, and calculate the propellant loading of each tank, the satellite Z mass center, and the engine interference torque;

[0020] S9. Calculate the mass of helium based on the pressure of the parallel tank at the beginning of the orbit change, and infer the air cushion pressure at the launch state based on the helium mass conservation principle;

[0021] S10. Based on the determined propellant filling amount, air cushion pressure and satellite center of mass, perform propellant and helium filling and subsequent on-orbit operations.

[0022] Furthermore, in S2, the specific method is:

[0023] S2.1 Calculate the total mass of oxidizer and total mass of fuel based on the satellite launch weight, dry satellite weight, and system mixing ratio;

[0024] S2.2 Calculate the filling mass and filling volume of the tank during balanced filling;

[0025] S2.3 Assume that the X and Y centroids of the propellant in the tank are on the tank central axis, that is, the X and Y centroids of the propellant are the same as the X and Y coordinates of the tank central axis;

[0026] S2.4 interpolate and calculate the Z-axis center of mass of the propellant in each tank, and integrate the data into the X, Y, and Z-axis center of mass coordinates of the propellant;

[0027] S2.5 Calculate the satellite center of mass coordinates in the launch state after balanced filling.

[0028] Furthermore, in S3, the specific method is:

[0029] S3.1 Calculate the maximum filling mass of each tank;

[0030] S3.2 Calculate the filling mass of the propellant in each tank of the oxygen tank a and the fuel tank a at the maximum filling mass;

[0031] S3.3 Calculate the filling volume of each tank of oxygen tank a and fuel tank a at the maximum filling mass;

[0032] S3.4 Calculate the satellite center of mass coordinates during unbalanced refueling;

[0033] S3.5 Calculate the filling mass of the propellant in each tank of oxygen tank b, fuel tank a, oxygen tank a, fuel tank b, oxygen tank b, fuel tank b at the maximum filling mass in sequence;

[0034] S3.6 Repeat steps S3.3 and S3.4 to calculate the satellite center of mass coordinates of oxygen box b, fuel box a, oxygen box a, fuel box b, oxygen box b, fuel box b at the maximum filling mass.

[0035] Furthermore, in S4, the specific method is:

[0036] S4.1 define a vector in the Z direction and a unit vector in the Z direction using the Z coordinate of the satellite's center of mass in the balanced and fueled launch state as the modulus;

[0037] S4.2 Calculate the engine thrust vector in the engine body coordinate system;

[0038] S4.3 Coordinate rotation transformation matrix between the engine precision measuring mirror coordinate system and the satellite mechanical coordinate system;

[0039] S4.4 Calculate the thrust vector in the satellite mechanical coordinate system when the engine is operating;

[0040] S4.5 Calculate the position coordinates of the engine thrust action point in the engine precision measuring mirror coordinate system;

[0041] S4.6 Calculate the vector of the engine thrust application point in the satellite mechanical coordinate system;

[0042] S4.7 Calculate intermediate variables;

[0043] S4.8 Calculate the coordinates of the intersection of the engine thrust vector and the XY plane at the satellite's center of mass.

[0044] Furthermore, in S5, the specific method is:

[0045] S5.1 In the XY plane coordinate system, draw point P1 according to the X and Y coordinates in the dry satellite mass center coordinate XYZ_s_1, which is called the "satellite dry satellite mass center";

[0046] S5.2 In the XY plane coordinate system, plot point P2 based on the X and Y coordinates of the centroid coordinates in S2.5, which is called the "balanced refueling satellite launch centroid";

[0047] S5.3 In the XY plane coordinate system, plot points P3, P4, P5, and P6 based on the X and Y coordinates of the centroid coordinates in S3.4 and S3.6, which are called the "unbalanced refueling satellite launch limit centroid";

[0048] S5.4 Connect points P3, P4, P5, and P6 to form a quadrilateral domain Q, which is called the "unbalanced refueling satellite launch center of mass adjustment domain";

[0049] S5.5 In the XY plane coordinate system, plot point P7 based on the X and Y coordinates of the coordinates in S4.8, which is called the "moment point of the engine thrust about the center of mass at the initial moment of the orbit change";

[0050] S5.6 In the XY plane coordinate system, a circular region R is drawn with the origin O as the center and the maximum permissible value of the satellite's lateral center of mass under launch conditions by the launch vehicle as the radius. This is called the "rocket-constrained satellite launch center of mass feasible region".

[0051] Furthermore, if the "unbalanced filling satellite launch center of mass adjustment domain" has no intersection with the "rocket constrained satellite launch center of mass feasible domain", the traditional method of adding counterweights must be adopted. Adding counterweights can translate the position of the "satellite dry star center of mass", and at the same time, the translation of the positions of the "balanced filling satellite launch center of mass" and the "unbalanced filling satellite launch center of mass adjustment domain" can be achieved, and the "unbalanced filling satellite launch center of mass adjustment domain" can be made to intersect with the "rocket constrained satellite launch center of mass feasible domain".

[0052] Furthermore, in S7, the specific steps are:

[0053] S7.1 In the satellite centroid relationship diagram, select a point in the intersection of the quadrilateral domain Q and the circular domain R as the satellite centroid in the launch state;

[0054] S7.2 Calculate the amount of propellant to be filled in each tank during launch.

[0055] S7.3 Calculate the filling volume of each tank in the launch state;

[0056] S7.4 Calculate the Z-center of mass of the propellant in each tank;

[0057] S7.5 Calculate the Z-axis centroid coordinates of the satellite in the launch state;

[0058] S7.6 The X, Y, and Z coordinate vectors of the satellite's center of mass when synthesized in the launch state.

[0059] Further, in S8, the specific method is:

[0060] S8.1 In the satellite centroid relationship diagram, select the point closest to point P7 in the intersection of the quadrilateral domain Q and the circular domain R as the satellite centroid at the start of the orbit change;

[0061] S8.2 Calculate the amount of propellant filled in each tank at the start of the orbit change;

[0062] S8.3 Calculate the filling volume of each tank at the beginning of the track change;

[0063] S8.4 Calculate the Z-center of mass of the propellant in each tank by interpolation;

[0064] S8.5 Calculate the Z-axis centroid coordinate of the satellite at the beginning of the orbit change;

[0065] S8.6 synthesizes the X, Y coordinates of the point PT in step S8.1 and the Z-axis center of mass coordinates to obtain the X, Y, Z coordinate vector of the satellite center of mass at the start of the orbit change;

[0066] S8.7 Calculate the engine disturbance torque at the start of the track change.

[0067] Furthermore, in S9, the specific method is:

[0068] S9.1 At the beginning of the orbit change, the downstream valves of the two tanks of the same propellant are both open, and the outlets of the two tanks are connected, so the pressures are equal. The pressures of oxygen tank a and oxygen tank b at the beginning of the orbit change are designed to be a certain value;

[0069] S9.2 Calculate the mass of helium in oxygen tank a and oxygen tank b at the start of the maneuver;

[0070] S9.3 The mass of helium in each tank at the beginning of the orbit change and in the launch state is conserved. Calculate the helium pressure of oxygen tank a and oxygen tank b in the launch state, and obtain the helium pressure of fuel tank a and fuel tank b in the launch state.

[0071] The beneficial effects of the present invention compared with the prior art are:

[0072] (1) In the prior art, the design method of the propellant loading amount and the center of mass of the parallel tank satellite does not consider the difference in satellite mass characteristics at the time of satellite launch and the initial time of orbit change, and does not consider the insufficient adjustment ability of the filling pressure of the tank air cushion on the propellant loading amount and the satellite center of mass. The present invention adopts a strategy of independently controlling the propellant loading mass and the air cushion pressure of each tank at different times, and uses the air cushion as the driving force for adjusting the distribution of propellants in different tanks. The satellite center of mass is designed at different times of satellite launch and the initial time of orbit change, respectively, so as to meet multiple constraints at the same time.

[0073] (2) The present invention is based on the current hardware configuration of parallel tank satellites. By controlling the self-locking valves at the upstream and downstream of the tanks respectively, the filling and pressurization of propellant and helium can be completed, and the satellite center of mass design that satisfies the dual constraints of the launch state and the start of the orbit change can be realized, which is highly economical in terms of cost.

[0074] (3) The present invention introduces a satellite centroid relationship diagram, which can vividly display the positional relationship between the satellite centroid and the constraint conditions at different times, making it easy to understand and quickly guide engineering applications;

[0075] (4) The calculation links of each part of the present invention are fully parameterized, and calculation methods such as iterative loops and judgment branches are introduced to facilitate the compilation of calculation programs, thereby reducing the calculation time;

[0076] (5) The present invention comprehensively considers factors such as engine performance parameters and satellite installation accuracy errors, closely follows engineering practice, is convenient for engineering application, and helps improve calculation accuracy. BRIEF DESCRIPTION OF THE DRAWINGS

[0077] Figure 1 This is a schematic diagram of the parallel tank satellite configuration of the present invention;

[0078] Figure 2 This is a schematic diagram of the parallel tank chemical propulsion system of the present invention;

[0079] Figure 3 This is a satellite center of mass relationship diagram of the present invention. DETAILED DESCRIPTION

[0080] The present invention will be further described below in conjunction with the embodiments.

[0081] The method of determining the amount of non-isobaric air cushion propellant in the parallel tank and the satellite center of mass, such as Figure 1-3As shown,

[0082] Attached Figure 1 It is a schematic diagram of the parallel tank satellite configuration.

[0083] A parallel tank satellite generally contains 4 tanks, including 2 oxygen tanks and 2 fuel tanks, namely oxygen tank a MON-a, oxygen tank b MON-b, fuel tank a MMH-a, and fuel tank b MMH-b. The engine is arranged on the satellite bottom plate.

[0084] Attached Figure 2 This is the schematic diagram of a parallel tank chemical propulsion system. The names of each unit are shown in the legend.

[0085] Attached Figure 3 It is the satellite centroid relationship diagram.

[0086] No counterweight is added: the center of mass of the dry satellite is point P1, the center of mass of the satellite in the balanced filling launch state is point P2, and the satellite center of mass adjustment capacity in the unbalanced filling launch state is the quadrilateral domain Q composed of points P3~P6.

[0087] Add counterweights: the center of mass of the dry satellite is point P1', the center of mass of the satellite in the balanced filling launch state is point P2', and the satellite center of mass adjustment capacity in the unbalanced filling launch state is the quadrilateral domain Q' composed of points P3'~P6'.

[0088] It includes two parts: the first is the calculation of the amount of propellant filling and the center of mass of the satellite. The second is the assembly of the counterweight (if any), the filling of propellant and helium, and the on-orbit operation.

[0089] Part_1 Propellant filling amount and satellite center of mass calculation

[0090] The calculation of propellant filling amount and satellite mass center includes 10 steps, as follows:

[0091] Calcu_1: Get input parameters

[0092] a) System mixing ratio η. This parameter can be estimated by the orbit control engine mixing ratio, attitude control thruster mixing ratio, etc.

[0093] b) The measured volume V_t of the tank (the four values ​​in the parameters are the values ​​of oxygen tank a, oxygen tank b, fuel tank a, and fuel tank b respectively), the X and Y coordinates XY_t of the tank center axis (parallel to the Z axis of the satellite mechanical coordinate system) in the satellite mechanical coordinate system, the density of the propellant ρ_p, and the maximum allowable filling ratio c_t of the tank.

[0094] c) The height of the liquid center of mass Z_p at different filling volumes of the tank, including multiple sets of corresponding relationships between different specific filling volumes and the height of the liquid center of mass (V_p_ i , Z_p_ i), which can be used to interpolate the height of the liquid center of mass for any filling volume.

[0095] d) Satellite launch mass m_s_3.

[0096] e) Before propellant filling, that is, the weight of the satellite stem is m_s_1, and the coordinates of the satellite stem center of mass in the satellite mechanical coordinate system are XYZ_s_1.

[0097] f) orbit control engine thrust F, orbit control engine thrust deflection angle α, thrust vector lateral displacement angle β, thrust vector deflection position angle γ, thrust vector lateral displacement δ, and length from mounting surface to nozzle l_1.

[0098] g) The angle matrix Г between the axes of the precision measuring mirror coordinate system and the axes of the satellite mechanical coordinate system, and the coordinates of the engine nozzle center in the satellite mechanical coordinate system Δ=[Δ1,Δ2,Δ3].

[0099] h) The maximum permissible value of the satellite's lateral center of mass in the launch vehicle is r_L, that is, the root mean square of the satellite's X and Y center of mass in the launch state does not exceed r_L.

[0100] i) The temperature of the helium in the tank is T_He.

[0101] j) Input parameters of the counterweight: the maximum number of counterweights that can be installed on the satellite, m, the mass of m sets of counterweights, m k =[m k1 m k2 ... m km ], the center of mass coordinates of the m-group counterweight in the satellite mechanical coordinate system

[0102] Calcu_2: Preliminary analysis when balancing a raise

[0103] a) Calculate the total mass of the oxidant m_o and the total mass of the fuel m_f.

[0104] m_o=(m_s_3-m_s_1)*η*(1+η) (1)

[0105] m_f=m_s_3-m_s_1-m_o (2)

[0106] b) Calculate the filling mass m_p and filling volume V_p of the tank when balanced filling (i.e., filling of equal mass of propellant in the same type of propellant tank).

[0107] m_p=0.5*[m_o,m_o,m_f,m_f] (3)

[0108] V_p=m_p. / ρ_p (4)

[0109] Here, the symbol ". / " represents array division.

[0110] c) Assume that the X and Y center of mass of the propellant in the tank is on the central axis of the tank, that is, the X and Y center of mass coordinates of the propellant are XY_t.

[0111] d) Using V_p as input, interpolate and calculate the Z-axis center of mass of the propellant in each tank, and integrate the data into the X, Y, and Z-axis center of mass coordinates of the propellant, XYZ_t.

[0112] e) Calculate the satellite center of mass coordinates at launch state after balanced filling

[0113] XYZ_s_3=(m_s_1*XYZ_s_1+m_p*XYZ_t) / m_s_3 (5)

[0114] Calcu_3: Calculates the center of mass adjustment capability during unbalanced filling

[0115] a) Calculate the maximum filling mass of each tank.

[0116] m_p_ma=c_t.*(ρ_p.*V_t) (6)

[0117] b) Calculate the filling mass m_p of the propellant in each tank of oxygen tank a and fuel tank a at the maximum filling mass.

[0118] m_p=[m_p_ma(1),m_o-m_p_ma(1),m_p_ma(3),m_f-m_p_ma(3)] (7)

[0119] c) Similar to formula (5), calculate the filling volume V_p of each tank of oxygen tank a and fuel tank a at the maximum filling mass.

[0120] d) Similar to steps c), d), and e) of Calcu_2, calculate the satellite center of mass coordinates during unbalanced filling.

[0121] e) Calculate the filling mass m_p of the propellant in each tank of oxygen tank b, fuel tank a, oxygen tank a, fuel tank b, oxygen tank b, fuel tank b at the maximum filling mass.

[0122] m_p=[m_o-m_p_ma(2),m_p_ma(2),m_p_ma(3),m_f-m_p_ma(3)] (8)

[0123] m_p=[m_p_ma(1),m_o-m_p_ma(1),m_f-m_p_ma(4),m_p_ma(4)] (9)

[0124] m_p=[m_o-m_p_ma(2),m_p_ma(2),m_f-m_p_ma(4),m_p_ma(4)] (10)

[0125] f) Repeat steps c) and d) to calculate the satellite center of mass coordinates of oxygen box b, fuel box a, oxygen box a, fuel box b, oxygen box b, fuel box b at the maximum filling mass.

[0126] Calcu_4: Calculates the intersection of the engine thrust vector and the XY plane at the satellite's center of mass

[0127] a) Define vector N = [0,0,XYZ_s_3(3)], where XYZ_s_3(3) is the Z coordinate of the satellite's center of mass at launch after balanced filling in step e) of Calcu_2; define vector VP = [0,0,1].

[0128] b) Calculate the vector of engine thrust in the engine body coordinate system

[0129] F_e=F·[cos(α),sin(α)·cos(γ),sin(α)·sin(γ)] T (11)

[0130] c) Coordinate rotation transformation matrix T between the engine precision measuring mirror coordinate system and the satellite mechanical coordinate system

[0131]

[0132] Where ξ1, ψ1, υ1, etc. are the elements of the angle matrix Г between the axes of the precision mirror coordinate system and the axes of the satellite mechanical coordinate system.

[0133] d) Thrust vector in the satellite mechanical coordinate system when the engine is working

[0134] F_s=T·F_e (13)

[0135] e) The position coordinates of the engine thrust action point in the engine precision measuring mirror coordinate system are:

[0136] P_c=[0,δ·cosβ,δ·sinβ] T +[l_1,0,0] T (14)

[0137] f) The vector of the engine thrust action point in the satellite mechanical coordinate system is:

[0138] P_s=T·P_c+[Δ1,Δ2,Δ3] T (15)

[0139] g) Calculate the intermediate variable t

[0140] t=(N-P_s')*VP' / (VP*F_s) (16)

[0141] h) Calculate the intersection coordinates P of the engine thrust vector and the XY plane at the satellite's mass center

[0142] P=(P_s+t*F_s)' (17)

[0143] Calcu_5: Plotting satellite centroid relationships

[0144] a) In the XY plane coordinate system, draw point P1 according to the X and Y coordinates in the center of mass coordinate XYZ_s_1, which is called the "satellite dry star center of mass".

[0145] b) In the XY plane coordinate system, draw point P2 according to the X and Y coordinates of the center of mass coordinates XYZ_s_3 in step e) of Calcu_2, which is called the "balanced filling satellite launch center of mass".

[0146] c) In the XY plane coordinate system, draw points P3, P4, P5, and P6 based on the X and Y coordinates of the centroid coordinates XYZ_s_3 in steps d) and f) of Calcu_3, which are called the "unbalanced refueling satellite launch limit centroid."

[0147] d) Connect points P3, P4, P5, and P6 to form a quadrilateral domain Q, which is called the "unbalanced refueling satellite launch center of mass adjustment domain."

[0148] e) In the XY plane coordinate system, draw point P7 according to the X and Y coordinates of coordinate P in step h) of Calcu_4, which is called the "moment point of engine thrust about the center of mass at the initial moment of orbit change".

[0149] f) In the XY plane coordinate system, a circular area R is drawn with the origin O as the center and the maximum allowable value r_L of the satellite's lateral center of mass under the launch state of the carrier rocket as the radius. This is called the "feasible region of the satellite launch center of mass constrained by the rocket."

[0150] Calcu_6: Determine whether additional counterweights are needed. If additional counterweights are needed, determine the number, weight, and location of the counterweights.

[0151] Analyze the satellite centroid relationship diagram drawn in Calcu_5. Generally, point P2 is within the quadrilateral domain Q. Analyze the positional relationship between point P2, quadrilateral domain Q, and circular domain R:

[0152] 1) If point P2 is within the circular region R, then the satellite mass center can meet the constraints of the launch vehicle under the launch state without adding counterweight blocks and when the parallel tanks are balanced and filled;

[0153] 2) If point P2 is not within the circular domain R, but the quadrilateral domain Q intersects the circular domain R, then the satellite mass center can meet the constraints of the launch vehicle under the launch state when the counterweight is not added and the parallel tanks are unbalancedly filled;

[0154] 3) If the quadrilateral domain Q does not intersect the circular domain R, a counterweight must be added to ensure that the satellite's center of mass meets the constraints of the launch vehicle in the launch state.

[0155] The steps in this section are as follows:

[0156] a) Selection of counterweight position: In the XY plane coordinate system, draw a line from point P1 to the coordinate origin O, and continue to extend it until it exceeds the X and Y outer envelope of the satellite, and select several groups of counterweight installation holes that are closer to the above line segments in the X and Y directions as candidates.

[0157] b) Set n=1.

[0158] c) Within the range determined in step a), calculate the new satellite dry mass m_s_1_n and center of mass coordinates XYZ_s_1_n after adding n sets of counterweights.

[0159] m_s_1_n=m_s_1+m k [1] (n,1) (18)

[0160] XYZ_s_1_n=[m_s_1*XYZ_s_1+m k *XYZ_k] / m_s_1_n (19)

[0161] In the formula, [1] (n,1) It is a matrix with n rows and 1 column containing all 1s.

[0162] d) Update the weight m_s_1 of the satellite stem and the coordinates XYZ_s_1 of the satellite stem center of mass in step e) of Calcu_1 to the values ​​of m_s_1_n and XYZ_s_1_n.

[0163] e) Re-carry out the work of Calcu_2, Calcu_3, Calcu_4 and Calcu_5.

[0164] f) Determine whether the quadrilateral domain Q intersects with the circular domain R. If not, proceed to step (ga); if so, proceed to step (gb).

[0165] ga) Add 1 to n and go to step c) of Calcu_6.

[0166] gb)End Calcu_6.

[0167] Calcu_7: Calculates the satellite center of mass and propellant charge in the launch state

[0168] Calcu_6 obtained the design condition that the “unbalanced refueling satellite launch center of mass adjustment domain” and the “rocket-constrained satellite launch center of mass feasible domain” have an intersection.

[0169] The steps in this section are as follows:

[0170] a) In the satellite centroid relationship diagram, select a point in the feasible domain (the intersection of the quadrilateral domain Q and the circular domain R) as the satellite's X and Y centroid in the launch state, record it as point PL, and record the X and Y coordinate vector PL of the point.

[0171] b) Calculate the propellant filling quantity of each tank using the following formula:

[0172]

[0173] c) Calculate the filling volume V_p_L of each tank during launch

[0174] V_p_L=m_p_L. / ρ_p (21)

[0175] d) Using V_p_L as input, interpolate and calculate the Z-center of mass Z_p_L of the propellant in each tank.

[0176] e) Calculate the Z coordinates of the satellite's center of mass when in the launch state.

[0177] Z_s_3=[m_s_1*XYZ_s_1(3)+m_p_L*Z_p_L] / m_s_3 (22)

[0178] f) Combine the X, Y coordinates of point PL in step a) with Z_s_3 to obtain the X, Y, Z coordinate vector XYZ_s_L of the satellite's center of mass in the launch state.

[0179] Calcu_8: Calculates the satellite center of mass and propellant filling amount at the beginning of orbit change

[0180] The steps in this section are as follows:

[0181] a) In the satellite centroid relationship diagram, select the point closest to point P7 in the quadrilateral domain Q as the satellite's X and Y centroid at the beginning of the orbit change, record it as point PT, and record the X and Y coordinate vector PT of the point.

[0182] b) Calculate the amount of propellant in each tank at the start of the orbit change using the following formula:

[0183]

[0184] c) Calculate the filling volume V_p_T of each tank at the beginning of the track change

[0185] V_p_T=m_p_T. / ρ_p (24)

[0186] d) Using V_p_T as input, interpolate and calculate the Z-center of mass Z_p_T of the propellant in each tank.

[0187] e) Calculate the Z coordinate of the satellite’s center of mass when the orbit change begins.

[0188] Z_s_T=[m_s_1*XYZ_s_1(3)+m_p_T*Z_p_T] / m_s_3 (25)

[0189] f) Combine the X, Y coordinates of the midpoint PT in step a) with Z_s_T to obtain the X, Y, Z coordinate vector XYZ_s_T of the satellite's center of mass at the beginning of the orbit change.

[0190] g) Calculate the engine interference torque at the beginning of the track change

[0191] M=F_s×(P_s-XYZ_s_T) (26)

[0192] Calcu_9: Calculates the air cushion pressure during launch

[0193] a) At the beginning of the orbit change, the downstream valves of the two tanks of the same propellant are both open, and the outlets of the two tanks are connected, so the pressure is equal. At this time, the tank pressure can be taken in the same way as the traditional satellite with a single propellant loaded by a single tank. The pressure of oxygen tank a and oxygen tank b at the beginning of the orbit change is designed to be P_o_T.

[0194] b) Calculate the helium masses m_He_o_a and m_He_o_b in oxygen tank a and oxygen tank b respectively at the beginning of the orbit change.

[0195] Z=1+1.9572205×P_o_T×T_He -1.062213 (27)

[0196] ρ_He=P_o_T×M / (Z×T_He×R) (28)

[0197] m_He_o_a=ρ_He×[V_t(1)-V_p_T(1)] (29)

[0198] m_He_o_b=ρ_He×[V_t(2)-V_p_T(2)] (30)

[0199] Where Z is the compressibility factor of helium, R is the gas constant, M is the molar mass of helium, and ρ_He is the density of helium.

[0200] c) The mass of helium in each tank is conserved at the beginning of the orbit change and in the launch state. According to the following formula, calculate the helium pressures P_a_o_L and P_b_o_L of oxygen tank a and oxygen tank b respectively in the launch state.

[0201] m_He_o_a / [V_t(1)-V_p_L(1)]=

[0202] P_a_o_L×M / [(1+1.9572205×P_a_o_L×T_He -1.062213 )×T_He×R] (32)

[0203] In the above formula, only P_a_o_L is an unknown variable, and the other variables have been solved in the previous steps. Similarly, P_b_o_L can be solved.

[0204] d) Similar to oxygen box a and oxygen box b, the helium pressures P_a_f_L and P_b_f_L of fuel box a and fuel box b respectively during the launch state can be solved.

[0205] Calcu_10: Summarizes calculation results

[0206] a) If additional counterweights are needed, output the number of counterweights n and the mass of the counterweights m k =[m k1 m k2 ...m kn ], the center of mass coordinates of the counterweight in the satellite mechanical coordinate system

[0207] b) The propellant filling mass m_p_L of each tank during launch.

[0208] c) The helium filling pressure of each tank during launch state is P_a_o_L, P_b_o_L, P_a_f_L, and P_b_f_L.

[0209] Part_2 Assembly of counterweight (if any), filling of propellant and helium, and on-orbit operation

[0210] It includes the following steps:

[0211] a) If additional counterweights are required, install the counterweights on the satellite in sequence according to the calculation results of step a) of Calcu_10 in Part_1.

[0212] b) Open the self-locking valves and the filling and exhaust valves upstream and downstream of the four storage tanks to replace the air in the tanks with high-purity helium. Close the self-locking valves upstream and downstream of the four storage tanks.

[0213] c) Open the self-locking valves upstream and downstream of oxygen tank a, add oxidant to oxygen tank a through the oxidant filling and discharging valve downstream of the oxygen tank until the oxidant filling mass reaches m_p_L(1), and close the self-locking valve downstream of oxygen tank a.

[0214] d) Fill oxygen tank a with high-pressure helium through the helium filling and discharging valve upstream of the oxygen tank until the pressure reaches P_a_o_L, and close the self-locking valve upstream of the oxygen tank a.

[0215] e) Similarly, complete the propellant and helium filling of oxygen tank b, fuel tank a, and fuel tank b according to the calculation results of steps b) and c) of Calcu_10 in Part_1.

[0216] f) For satellite launch, satellite-rocket separation, etc., a short time before the first orbit change (for example 30 minutes), open the self-locking valves downstream of the four tanks. At this time, the pressure between the two tanks filled with the same propellant will gradually approach. When the pressure between the two tanks no longer changes, the subsequent orbit change can be carried out.

[0217] Although the present invention has been disclosed as above in the form of a preferred embodiment, it is not intended to limit the present invention. Any person skilled in the art may make possible changes and modifications to the technical solution of the present invention by using the methods and technical contents disclosed above without departing from the spirit and scope of the present invention. Therefore, any simple modifications, equivalent changes and modifications made to the above embodiments based on the technical essence of the present invention without departing from the content of the technical solution of the present invention shall fall within the protection scope of the technical solution of the present invention.

Claims

1. A method for determining the amount of non-isobaric air cushion propellant in parallel tanks and the center of mass of a satellite, characterized in that the steps include: S1. Obtain input parameters through experiments, tests and simulations, mainly including the mass and center of mass of the satellite before filling, the performance parameters of the tank and engine and the installation parameters on the satellite, the density of the propellant, the optional number and installation position of the counterweights, and the constraints of the launch vehicle on the satellite center of mass; S2. Conduct balanced filling, i.e., preliminary analysis when the weight of the propellants in the parallel tanks is equal, and obtain the propellant filling mass of each tank during balanced filling, the propellant center of mass, and the coordinates of the satellite center of mass in the launch state; S3. Calculate the center of mass adjustment capability during non-balanced filling, and sequentially obtain the propellant filling mass, propellant center of mass, and satellite center of mass coordinates of each tank in the launch state when each tank is filled according to the maximum filling mass; In S3, the specific method is: S3.1 Calculate the maximum filling mass of each tank; S3.2 Calculate the filling mass of the propellant in each tank of the oxygen tank a and the fuel tank a at the maximum filling mass; S3.3 Calculate the filling volume of each tank of oxygen tank a and fuel tank a at the maximum filling mass; S3.4 Calculate the satellite center of mass coordinates during unbalanced refueling; S3.5 Calculate the propellant filling mass of each tank at the maximum filling mass of oxygen tank b, fuel tank a, oxygen tank a, fuel tank b, oxygen tank b, fuel tank b in turn; S3.6 Repeat steps S3.3 and S3.4 to calculate the satellite center of mass coordinates of oxygen tank b, fuel tank a, oxygen tank a, fuel tank b, oxygen tank b, fuel tank b at the maximum filling mass; S4. Calculate the intersection of the engine thrust vector and the XY plane at the satellite mass center according to the conversion relationship between the engine thrust vector and the engine body coordinate system, the engine precision measuring mirror coordinate system and the satellite mechanical coordinate system; In S4, the specific method is: S4.1 define a vector in the Z direction and a unit vector in the Z direction using the Z coordinate of the satellite's center of mass in the balanced and fueled launch state as the modulus; S4.2 Calculate the vector of the engine thrust in the engine body coordinate system; S4.3 Coordinate rotation transformation matrix between the engine precision measuring mirror coordinate system and the satellite mechanical coordinate system; S4.4 Calculate the thrust vector in the satellite mechanical coordinate system when the engine is operating; S4.5 Calculate the position coordinates of the engine thrust action point in the engine precision measuring mirror coordinate system; S4.6 Calculate the vector of the engine thrust application point in the satellite mechanical coordinate system; S4.7 Calculate intermediate variables; S4.8 Calculate the coordinates of the intersection of the engine thrust vector and the XY plane at the satellite's center of mass; S5, using the satellite mass center coordinates calculated by S2, S3, and S4, the intersection of the engine thrust vector and the XY plane at the satellite mass center, and the constraints of the launch vehicle on the satellite mass center as input, draw a satellite mass center relationship diagram; In S5, the specific method is: S5.1 In the XY plane coordinate system, draw point P1 according to the X and Y coordinates in the dry satellite mass center coordinate XYZ_s_1, which is called the "satellite dry satellite mass center"; S5.2 In the XY plane coordinate system, draw point P2 based on the X and Y coordinates of the centroid coordinates in S2.5, which is called the "balanced refueling satellite launch centroid"; S5.3 In the XY plane coordinate system, points P3, P4, P5, and P6 are plotted based on the X and Y coordinates of the centroid coordinates in S3.4 and S3.6, which are called the "unbalanced refueling satellite launch limit centroid"; S5.4 Connect points P3, P4, P5, and P6 to form a quadrilateral domain Q, which is called the "unbalanced refueling satellite launch center of mass adjustment domain"; S5.5 In the XY plane coordinate system, plot point P7 based on the X and Y coordinates of the coordinates in S4.8, which is called the "moment point of the engine thrust about the center of mass at the initial moment of the orbit change"; S5.6 In the XY plane coordinate system, a circle R is drawn with the origin O as the center and the maximum permissible value of the satellite's lateral center of mass under launch by the launch vehicle as the radius, which is called the "feasible region of satellite launch center of mass under rocket constraints"; S6. Determine whether it is necessary to add counterweights and whether it is necessary to unbalance the parallel tanks to ensure that the satellite mass center meets the constraints of the launch vehicle in the launch state through the positional relationship between the points and the domains in the satellite mass center relationship diagram drawn in S5; if it is necessary to add counterweights, determine the number, weight and position of the counterweights; S7, select a point in the feasible domain of the satellite mass center relationship diagram drawn in S5 as the satellite X and Y center of mass in the launch state, and calculate the propellant filling amount of each tank and the satellite Z center of mass in the launch state; S8. Select a point closest to the engine thrust vector in the feasible domain of the satellite mass center relationship diagram drawn in S5 as the satellite X and Y mass centers at the beginning of the orbit change, and calculate the propellant loading of each tank, the satellite Z mass center, and the engine interference torque; S9. Calculate the mass of helium based on the pressure of the parallel tank at the beginning of the orbit change, and infer the air cushion pressure at the launch state based on the helium mass conservation principle; S10. Based on the determined propellant filling amount, air cushion pressure and satellite center of mass, perform propellant and helium filling and subsequent on-orbit operations.

2. The method for determining the amount of parallel tank non-isobaric air cushion propellant and the satellite mass center according to claim 1, characterized in that: In S2, the specific method is: S2.1 Calculate the total mass of oxidizer and total mass of fuel based on the satellite launch weight, dry satellite weight, and system mixing ratio; S2.2 Calculate the filling mass and filling volume of the tank during balanced filling; S2.3 Assume that the X and Y centroids of the propellant in the tank are on the tank central axis, that is, the X and Y centroids of the propellant are the same as the X and Y coordinates of the tank central axis; S2.4 interpolate and calculate the Z-axis center of mass of the propellant in each tank, and integrate the data into the X, Y, and Z-axis center of mass coordinates of the propellant; S2.5 Calculate the satellite center of mass coordinates in the launch state after balanced filling.

3. The method for determining the amount of parallel tank non-isobaric air cushion propellant and the satellite mass center according to claim 1, characterized in that: If the "unbalanced refueling satellite launch center of mass adjustment domain" has no intersection with the "rocket-constrained satellite launch center of mass feasible domain", the traditional method of adding counterweights must be adopted. Adding counterweights can translate the position of the "satellite dry star center of mass", and at the same time, the translation of the positions of the "balanced refueling satellite launch center of mass" and the "unbalanced refueling satellite launch center of mass adjustment domain" can be achieved, and the "unbalanced refueling satellite launch center of mass adjustment domain" and the "rocket-constrained satellite launch center of mass feasible domain" can be made to intersect.

4. The method for determining the amount of parallel tank non-isobaric air cushion propellant and the satellite mass center according to claim 1, characterized in that: In S7, the specific steps are: S7.1 In the satellite centroid relationship diagram, select a point in the intersection of the quadrilateral domain Q and the circular domain R as the satellite centroid in the launch state; S7.2 Calculate the amount of propellant to be filled in each tank during launch. S7.3 Calculate the filling volume of each tank in the launch state; S7.4 Calculate the Z-center of mass of the propellant in each tank; S7.5 Calculate the Z-axis centroid coordinates of the satellite in the launch state; S7.6 The X, Y, and Z coordinate vectors of the satellite's center of mass when synthesized in the launch state.

5. The method for determining the amount of parallel tank non-isobaric air cushion propellant and the satellite mass center according to claim 1, characterized in that: In S8, the specific method is: S8.1 In the satellite centroid relationship diagram, select the point closest to point P7 in the intersection of the quadrilateral domain Q and the circular domain R as the satellite centroid at the start of the orbit change; S8.2 Calculate the amount of propellant filled in each tank at the start of the orbit change; S8.3 Calculate the filling volume of each tank at the beginning of the track change; S8.4 Calculate the Z-center of mass of the propellant in each tank by interpolation; S8.5 Calculate the Z-axis centroid coordinate of the satellite at the beginning of the orbit change; S8.6 synthesizes the X, Y coordinates of the point PT in step S8.1 and the Z-axis center of mass coordinates to obtain the X, Y, Z coordinate vector of the satellite center of mass at the start of the orbit change; S8.7 Calculate the engine disturbance torque at the start of the track change.

6. The method for determining the amount of parallel tank non-isobaric air cushion propellant and the satellite mass center according to claim 1, characterized in that: In S9, the specific method is: S9.1 At the beginning of the orbit change, the downstream valves of the two tanks of the same propellant are both open, and the outlets of the two tanks are connected, so the pressures are equal. The pressures of oxygen tank a and oxygen tank b at the beginning of the orbit change are designed to be a certain value; S9.2 Calculate the mass of helium in oxygen tank a and oxygen tank b at the start of the maneuver; S9.3 The mass of helium in each tank at the beginning of the orbit change and in the launch state is conserved. Calculate the helium pressure of oxygen tank a and oxygen tank b in the launch state, and obtain the helium pressure of fuel tank a and fuel tank b in the launch state.

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