Method and system for inhibiting passive evaporation capacity of cryogenic propellant in microgravity environment

Abstract

The invention provides a method and a system for inhibiting passive evaporation capacity of a low-temperature propellant in a microgravity environment, and the method comprises the steps: S1, designing a transfer orbit form according to an operation mission profile of a transfer vehicle, determining a perigee and an apogee of the transfer orbit, and calculating a transfer orbit period; s2, performing dynamic simulation calculation through a transfer orbit equation, a sun position and an earth shadow model to obtain the in-orbit sunshine duration of the transfer vehicle; s3, calculating data of space external heat flow borne by the surface of the transfer carrier and the temperature of the outer surface of the multi-layer heat insulation material; s4, calculating the heat flux density of the multi-layer heat insulation material in different arrangement forms, and selecting the arrangement form of the multi-layer heat insulation material with the optimal heat insulation effect; and S5, evaluating whether the low-temperature propellant carried by the vehicle meets the allowable use amount of the trajectory, and continuously carrying out iterative calculation by optimizing the transfer trajectory until the low-temperature propellant carried by the vehicle meets the allowable use amount of the trajectory.

Description

TECHNICAL FIELD

[0001] The application relates to a low-temperature propellant passive evaporation amount inhibition method in a microgravity environment and belongs to the field of overall design of space transportation vehicles. BACKGROUND

[0002] In a traditional one-time arrival mode, a rocket final stage and a payload need to be sent into space by the same rocket base stage. The rocket needs to simultaneously consider the task requirements of the two different phases of entering space and space transfer, which greatly limits the theoretical upper limit of the transportation capacity. A new type of space transfer transportation system completely decouples the entering space and space transfer processes in a relay transportation mode, and adds two major elements, a space refueling device and a transfer vehicle. In the new space transfer transportation relay mode, the dry weight of the transfer vehicle in the entering space phase and the payload can be designed according to the maximum low-orbit carrying capacity of the carrier rocket; the permanent orbit reuse of the transfer vehicle in the space transfer phase greatly improves the structural efficiency. The use of on-orbit refueling to increase the available amount of propellant on demand can achieve a fold increase in the carrying capacity of medium-high orbit and deep space missions under the same launch scale.

[0003] Low-temperature propellant is an important option for high-performance propellant of transfer vehicles due to its high specific impulse, non-toxicity and non-pollution. However, it is difficult to store for a long time due to its low boiling point and easy vaporization under heating, which limits its application in space to some extent. The space transfer vehicle is in a long-term on-orbit state when performing lunar transfer, Mars transfer and large-scale deep space exploration missions. It is affected by the space thermal environment such as solar radiation, earth infrared radiation, planetary albedo and black background, which brings difficulties to the long-term storage of low-temperature propellant.

[0004] On-orbit evaporation amount control of low-temperature propellant refers to the comprehensive use of various space thermal management technologies to reduce heat leakage from the orbit environment to the low-temperature tank, effectively absorb, transfer and utilize propellant vapor cold, so as to reduce the evaporation amount of the propellant and control the pressure of the tank. The passive evaporation amount control scheme refers to the evaporation amount control scheme implemented only by the adiabatic method. The adiabatic method based on foam material is widely used in low-temperature carrier rockets and is mature in technology. Multilayer insulation (MLI) combines high reflectivity film material and low thermal conductivity spacer material, and has good evaporation amount control effect in the vacuum solar radiation environment, but only the side wall of the Centaur upper stage has applied this technology. At present, the research in this field in China is relatively lagging behind, which limits the long-term on-orbit application of low-temperature propellant in the new type of space transfer transportation system. SUMMARY

[0005] The technical problems to be solved by the present application are: overcoming the above-mentioned shortcomings of the prior art, providing a rapid engineering calculation method and system suitable for long-term on-orbit passive thermal control design of a new type of space transfer vehicle and low-temperature propellant evaporation amount estimation, supporting the transfer vehicle carrying capacity design and trajectory optimization calculation of the new type of space transfer system.

[0006] The technical solution adopted by the present application is: a low-temperature propellant passive evaporation amount suppression method in a microgravity environment, comprising:

[0007] S1, according to the transfer vehicle operation task profile, design the transfer orbit form and determine the perigee and apogee of the transfer orbit, and calculate the transfer orbit period;

[0008] S2, through dynamic simulation calculation of the transfer orbit equation, the sun position and the earth shadow model, the on-orbit sunlight duration of the transfer vehicle is obtained;

[0009] S3, calculate the space external heat flux data and the outer surface temperature of the multi-layer thermal insulation material received by the surface of the transfer vehicle;

[0010] S4, calculate the heat flux density of the multi-layer thermal insulation material under different arrangement forms, and select the arrangement form of the multi-layer thermal insulation material with the best heat insulation effect;

[0011] S5, evaluate whether the low-temperature propellant carried by the vehicle meets the allowable amount of trajectory, and continuously iterate the calculation by optimizing the transfer trajectory until the low-temperature propellant carried by the vehicle meets the allowable amount of trajectory.

[0012] Further, the calculation of the space external heat flux data and the outer surface temperature of the multi-layer thermal insulation material received by the surface of the transfer vehicle comprises:

[0013] S2.1, according to the position, speed and attitude angle information of the transfer vehicle in the launch inertial coordinate system during the on-orbit period, the vector of the transfer vehicle and the sunlight direction, the vector of the transfer vehicle and the earth radiation heat, and the ratio of the flight height to the earth radius are calculated, and the space external heat flux data received by the surface of the transfer vehicle is calculated;

[0014] S2.2, assuming that the surface of the multi-layer thermal insulation material of the transfer vehicle reaches thermal equilibrium when it is illuminated, according to the space external heat flux data received by the surface of the transfer vehicle, the MLI outer surface film absorption ratio and the emissivity, the outer surface temperature of the multi-layer thermal insulation material is obtained.

[0015] Further, the calculation of the heat flux density of the multi-layer thermal insulation material under different arrangement forms, and the selection of the arrangement form of the multi-layer thermal insulation material with the best heat insulation effect, comprises:

[0016] S3.1, given the arrangement form of the multi-layer insulation material, including the number of radiation screens and the thickness of the spacing layer, taking the outer surface temperature of the multi-layer insulation material obtained in S2.2 as the temperature boundary condition, setting the initial temperature distribution in the multi-layer insulation material; according to the initial temperature distribution in the multi-layer insulation material, the heat flux density of radiation heat transfer, gas heat transfer and solid heat conduction in each layer of the multi-layer insulation material is calculated;

[0017] S3.2, the temperature distribution in the multi-layer insulation material is calculated according to the heat flux density, and compared with the set initial temperature distribution in the multi-layer insulation material. If the deviation between the calculated temperature distribution in the multi-layer insulation material and the set initial temperature distribution in the multi-layer insulation material exceeds the set threshold, the temperature distribution in the multi-layer insulation material is re-set, and the iterative calculation is performed until the calculated temperature distribution in the multi-layer insulation material is equal to the set temperature distribution in the multi-layer insulation material, and the total heat flux density under the arrangement form of the multi-layer insulation material given in S3.1 is obtained;

[0018] S3.3, taking the process constraints, mass or size constraints of the multi-layer insulation material as constraint conditions, repeating S3.1 and 3.2, and calculating the heat flux density of the multi-layer insulation under different arrangement forms;

[0019] S3.4, comparing the heat flux density of the multi-layer insulation material under different arrangement forms, and selecting the arrangement form of the MLI material with the best heat insulation effect.

[0020] Further, whether the low-temperature propellant carried by the carrier satisfies the allowed amount of trajectory is evaluated by optimizing the transfer trajectory, and the iterative calculation is continuously performed until the low-temperature propellant carried by the carrier satisfies the allowed amount of trajectory, comprising:

[0021] According to the arrangement form of the multi-layer insulation material with the best heat insulation effect selected in S3.4 and the on-orbit sunshine duration obtained in S2, the on-orbit evaporation amount of the low-temperature propellant is calculated;

[0022] According to the on-orbit evaporation amount of the low-temperature propellant, the propellant consumption amount of the transfer carrier on the transfer orbit is calculated, and the accurate transfer trajectory form is determined;

[0023] According to the trajectory data and the safety margin coefficient of the accurate transfer trajectory form, whether the low-temperature propellant carried by the transfer carrier satisfies the allowed amount of trajectory is evaluated. If the low-temperature propellant carried by the transfer carrier does not satisfy the allowed amount of trajectory, the method ends. If the low-temperature propellant carried by the transfer carrier does not satisfy the allowed amount of trajectory, the transfer trajectory form is redesigned, and S1-S5 is repeated until the low-temperature propellant carried by the carrier satisfies the allowed amount of trajectory.

[0024] A low-temperature propellant passive evaporation amount suppression system in a microgravity environment, comprising:

[0025] A trajectory module is configured to design a transfer orbit form and determine a transfer orbit perigee and apogee according to a transfer vehicle operation mission profile, and calculate a transfer orbit period; the transfer vehicle in-orbit sunlight duration is obtained through dynamic simulation calculation of a transfer orbit equation, a sun position and an earth shadow model;

[0026] A space thermal environment module is configured to calculate space external heat flux data and a multi-layer insulation material outer surface temperature received by the transfer vehicle surface;

[0027] An MLI material module is configured to calculate heat flux density of the multi-layer insulation material in different arrangement forms, and select an optimal arrangement form of the multi-layer insulation material in terms of heat insulation effect;

[0028] An optimization iteration module is configured to evaluate whether the low-temperature propellant carried by the vehicle meets the allowable amount of trajectory, and continuously iteratively calculate by optimizing the transfer trajectory until the low-temperature propellant carried by the vehicle meets the allowable amount of trajectory.

[0029] Further, the space thermal environment module comprises:

[0030] According to position, velocity and attitude angle information of the transfer vehicle in the launch inertial coordinate system during the in-orbit period, a vector of the transfer vehicle and sunlight, a vector of the transfer vehicle and earth radiation heat, and a ratio of flight height to earth radius are calculated to obtain space external heat flux data received by the transfer vehicle surface;

[0031] According to the space external heat flux data received by the transfer vehicle surface, an MLI outer surface film absorption ratio and emissivity, an MLI outer surface temperature is obtained.

[0032] Further, the MLI material module comprises:

[0033] The heat flux density of the multi-layer insulation material in different arrangement forms is calculated with process constraints, mass or size constraints of the multi-layer insulation material as constraint conditions;

[0034] The heat flux density of the multi-layer insulation material in different arrangement forms is compared, and an optimal arrangement form of the MLI material in terms of heat insulation effect is selected.

[0035] Further, the method of calculating the heat flux density of the multi-layer insulation material in different arrangement forms with process constraints, mass or size constraints of the multi-layer insulation material as constraint conditions comprises:

[0036] With constraints on the process, quality, or size of the multi-layer thermal insulation material as conditions, and given the arrangement of the multi-layer thermal insulation material, including the number of radiation screen layers and the thickness of the spacer layer, the outer surface temperature of the multi-layer thermal insulation material calculated in the space thermal environment module is used as the temperature boundary condition to set the initial temperature distribution within the multi-layer thermal insulation material; based on the initial temperature distribution within the multi-layer thermal insulation material, the heat flux density of radiative heat transfer, gas heat transfer, and solid heat conduction in each layer of the multi-layer thermal insulation material is calculated.

[0037] The temperature distribution inside the multilayer insulation material is calculated based on the heat flux density and compared with the initial temperature distribution inside the multilayer insulation material. If the deviation between the calculated temperature distribution inside the multilayer insulation material and the initial temperature distribution inside the multilayer insulation material exceeds a set threshold, the temperature distribution inside the multilayer insulation material is reset and iteratively calculated until the calculated temperature distribution inside the multilayer insulation material is equal to the initial temperature distribution inside the multilayer insulation material, thus obtaining the total heat flux density under the given arrangement of multilayer insulation materials.

[0038] Furthermore, the optimization iteration module includes:

[0039] Based on the optimal MLI material arrangement with the best thermal insulation effect selected in the MLI material module and the on-orbit sunshine duration of the transfer vehicle calculated in the ballistic module, the on-orbit evaporation amount of cryogenic propellant is calculated.

[0040] Based on the on-orbit evaporation of cryogenic propellant, calculate the propellant consumption of the transfer vehicle on the transfer trajectory and determine the precise transfer trajectory.

[0041] Based on the ballistic data and safety margin factor of the precise transfer trajectory, assess whether the cryogenic propellant carried by the transfer vehicle meets the ballistic allowable amount.

[0042] Furthermore, the optimization iteration module also includes:

[0043] If the cryogenic propellant carried by the transfer vehicle does not meet the ballistic allowable amount, it returns to the ballistic module and the transfer trajectory is redesigned.

[0044] The advantages of this invention compared to the prior art are:

[0045] (1) This invention adopts an engineering design and calculation method for suppressing the long-term passive evaporation of cryogenic propellants in orbit. Through joint iterative calculation of flight trajectory, space thermal environment, and MLI insulation performance, it achieves rapid iterative optimization of cryogenic propellant evaporation control.

[0046] (2) This invention proposes a coupled iterative optimization design method and process for ballistic-thermal environment-MLI insulation performance of new space transfer vehicles for long-term on-orbit operation, and develops a rapid engineering calculation method for passive thermal protection design and prediction of cryogenic propellant evaporation, providing a technical basis for the overall design of new space transfer vehicles. Attached Figure Description

[0047] Figure 1 This is a flowchart of the method of the present invention;

[0048] Figure 2 This is a schematic diagram of the passive heat-resistant material covering of the storage tank;

[0049] Figure 3 A curve showing the estimated lateral temperature distribution of the storage tank. Detailed Implementation

[0050] The present invention will be described in conjunction with the accompanying drawings.

[0051] like Figure 1 As shown, a method for suppressing passive evaporation of cryogenic propellants under microgravity conditions includes the following steps:

[0052] Based on the mission profile of the new space transfer vehicle, the flight trajectory was initially designed. The space thermal environment parameters were determined based on the trajectory parameters. The MLI material optimization design was carried out with the thermal environment parameters as input conditions, and the heat leakage of the tank was calculated. The predicted propellant evaporation rate was incorporated into the trajectory design to form a rapid iterative optimization loop.

[0053] Step 1: Unlike traditional ballistic trajectory designs, the on-orbit transfer trajectory scheme of the new space transfer vehicle does not require calculations starting from launch site takeoff. Instead, it uses the stationary orbit as the starting point and the target orbit as the ending point. The on-orbit transfer trajectory is characterized by a Hohmann-like transfer, meaning that the engine ignites at the start of the transfer to provide the necessary velocity increment, then enters a long taxiing phase. Upon approaching the target orbit, it ignites again to decelerate, thus being captured by the target orbit. The main steps are as follows:

[0054] (1.1) Based on the mission profile of the transfer vehicle, design the transfer trajectory and determine the perigee and apogee of the transfer trajectory, and calculate the transfer trajectory period;

[0055] (1.2) The on-orbit sunshine duration of the transfer vehicle was obtained by dynamic simulation calculation using the transfer orbit equation, solar position and Earth shadow model.

[0056] Step Two: When a space transfer vehicle performs an orbital transfer mission, it is affected by the space thermal environment, including solar radiation, Earth's infrared radiation, planetary albedo, and black background. Therefore, it is necessary to calculate the magnitude of the space heat flux acting on the surface of the launch vehicle. The specific steps are as follows:

[0057] (2.1) The position, velocity and attitude angle information of the transfer vehicle in the launch inertial coordinate system during the transfer vehicle in orbit is provided according to step (1.1), the vector of the transfer vehicle and the direction of sunlight, the vector of the transfer vehicle and the earth's radiant heat, and the ratio of the flight height to the earth's radius are calculated, and the space external heat flow data on the surface of the transfer vehicle is calculated;

[0058] (2.2) Assuming that the multi-layer insulation material (MLI) surface of the transfer vehicle reaches thermal equilibrium when it is illuminated, the multi-layer insulation material outer surface temperature is obtained according to the space external heat flow data on the surface of the transfer vehicle, the MLI outer surface film absorption ratio and the emissivity.

[0059] Step three: the MLI material is composed of multiple radiation screens and spacer materials, and the heat transfer mechanisms involved are radiation, convective heat transfer and thermal conduction. Based on the traditional heat analysis method, each layer is taken as a node, and then the heat transfer effects of radiation, gas heat transfer and solid heat conduction in each layer are analyzed, and finally the total heat transfer amount is obtained, and the equivalent thermal conductivity of the MLI material is obtained. The specific steps are as follows:

[0060] (3.1) First, give the MLI arrangement form (number of radiation screens and thickness of spacer layer), take the multi-layer insulation material outer surface temperature obtained in step (2.2) as the temperature boundary condition, set the initial temperature distribution of the MLI material, and calculate the heat flux density of radiation heat transfer, gas heat transfer and solid heat conduction in each layer of the MLI material according to the initial temperature distribution;

[0061] (3.2) According to the heat flux density, the temperature distribution in the multi-layer insulation material is calculated, and compared with the set initial temperature distribution in the multi-layer insulation material. If the deviation between the calculated temperature distribution in the multi-layer insulation material and the set initial temperature distribution in the multi-layer insulation material exceeds the set threshold, the initial temperature distribution in the multi-layer insulation material is reset, and the iterative calculation is performed until the calculated temperature distribution in the multi-layer insulation material is equal to the set temperature distribution in the multi-layer insulation material, and the total heat flux density under the given arrangement form of the multi-layer insulation material in step 3.1 is obtained;

[0062] (3.3) Based on the MLI process constraints and mass (or size) constraints, steps 3.1 and 3.2 are repeated to calculate the heat flux density under different arrangement forms of the MLI material;

[0063] (3.4) Compare the heat flux densities under different arrangement forms of the MLI material, and select the arrangement form of the MLI material with the best heat insulation effect.

[0064] Step four:

[0065] (4.1) According to the MLI arrangement form calculated in step (3.4) and the sunlight condition of the transfer vehicle during the transfer orbit operation obtained in step (1.2), the on-orbit evaporation amount of the cryogenic propellant is calculated;

[0066] (4.2) The propellant consumption of the transfer vehicle on the transfer orbit is calculated according to the on-orbit evaporation amount of the cryogenic propellant obtained in step (4.1), and the accurate transfer trajectory form is determined;

[0067] (4.3) Based on the trajectory data of the accurate transfer trajectory form and the safety margin coefficient, it is evaluated whether the cryogenic propellant carried by the vehicle meets the trajectory allowable amount, if not, the transfer trajectory is redesigned, and steps one to four are repeated until the cryogenic propellant carried by the vehicle meets the trajectory allowable amount.

[0068] A low-temperature propellant passive evaporation amount suppression system in a microgravity environment, comprising:

[0069] A trajectory module for designing the transfer orbit form and determining the perigee and apogee of the transfer orbit according to the transfer vehicle operation mission profile, and calculating the transfer orbit period; the on-orbit sunlight duration of the transfer vehicle is obtained through dynamic simulation calculation of the transfer orbit equation, the sun position and the earth shadow model;

[0070] A space thermal environment module for calculating the space external heat flux data received by the surface of the transfer vehicle and the outer surface temperature of the multi-layer thermal insulation material;

[0071] An MLI material module for calculating the heat flux density of the multi-layer thermal insulation material under different arrangement forms, and selecting the arrangement form of the multi-layer thermal insulation material with the best heat insulation effect;

[0072] An optimization iteration module for evaluating whether the cryogenic propellant carried by the vehicle meets the trajectory allowable amount, and continuously iterating the calculation by optimizing the transfer trajectory until the cryogenic propellant carried by the vehicle meets the trajectory allowable amount.

[0073] Further, the space thermal environment module comprises:

[0074] According to the position, velocity and attitude angle information of the transfer vehicle during the on-orbit period in the launch inertial coordinate system, the vector of the transfer vehicle and the sunlight direction, the vector of the transfer vehicle and the earth radiation heat, and the ratio of the flight height to the earth radius are calculated, and the space external heat flux data received by the surface of the transfer vehicle is calculated;

[0075] According to the space external heat flux data received by the surface of the transfer vehicle, the MLI outer surface film absorption ratio and the emissivity, the outer surface temperature of the multi-layer thermal insulation material is obtained.

[0076] Further, the MLI material module comprises:

[0077] calculating heat flux density of the multi-layer insulation under different arrangement forms as constraint conditions of process constraints, mass or size constraints of the multi-layer insulation material;

[0078] comparing heat flux densities of the multi-layer insulation material under different arrangement forms, and selecting an arrangement form of the multi-layer insulation material with optimal heat insulation effect.

[0079] Further, the method for calculating heat flux density of the multi-layer insulation under different arrangement forms as constraint conditions of process constraints, mass or size constraints of the multi-layer insulation material comprises:

[0080] given the arrangement form of the multi-layer insulation material, including the number of radiation screens and the thickness of the spacer layer, setting the initial temperature distribution in the multi-layer insulation material as a temperature boundary condition according to the outer surface temperature of the multi-layer insulation material calculated by the space thermal environment module, calculating the heat flux density of radiation heat transfer, gas heat transfer and solid heat conduction in each layer of the multi-layer insulation material according to the initial temperature distribution in the multi-layer insulation material;

[0081] calculating the temperature distribution in the multi-layer insulation material according to the heat flux density, and comparing it with the set initial temperature distribution in the multi-layer insulation material, if the deviation between the calculated temperature distribution in the multi-layer insulation material and the set initial temperature distribution in the multi-layer insulation material exceeds the set threshold, re-setting the temperature distribution in the multi-layer insulation material and iteratively calculating until the calculated temperature distribution in the multi-layer insulation material is equal to the set temperature distribution in the multi-layer insulation material, and obtaining the total heat flux density under the given arrangement form of the multi-layer insulation material.

[0082] Further, the optimization iteration module comprises:

[0083] calculating the on-orbit evaporation amount of the cryogenic propellant according to the arrangement form of the MLI material selected in the MLI material module and the on-orbit sunlight duration of the transfer vehicle calculated in the trajectory module;

[0084] calculating the propellant consumption of the transfer vehicle on the transfer orbit according to the on-orbit evaporation amount of the cryogenic propellant, and determining the accurate transfer trajectory form;

[0085] evaluating whether the cryogenic propellant carried by the transfer vehicle meets the allowable amount of the trajectory according to the trajectory data of the accurate transfer trajectory form and the safety margin coefficient.

[0086] Further, the optimization iteration module further comprises:

[0087] If the cryogenic propellant carried by the transfer vehicle does not meet the allowable amount of the trajectory, return to the trajectory module to redesign the transfer trajectory form.

[0088] Embodiment

[0089] Step one: Based on the space transfer vehicle, the 400km to 20000km elliptical orbit round trip mission is analyzed. The elliptical orbit period is calculated by using the third law of Kepler:

[0090]

[0091] In the formula, a is the semi-major axis of the orbit, G is the gravitational constant, and M is the mass of the earth. The orbit period is about 21230s.

[0092] In the elliptical orbit, the distance between the spacecraft and the earth is constantly changing, and the sunlight time depends on the cumulative time of the spacecraft outside the earth's shadow in the orbit period. The sunlight time is obtained by dynamic calculation through the orbit equation, the sun position and the earth shadow model, and the simulation calculation obtains the sunlight time of 19000s.

[0093] Step two: The space transfer vehicle is mainly affected by solar radiation, earth albedo and earth infrared radiation heat flux in long-term on-orbit flight profile, and the heat flux density of the three is 1353W / m 2 , 459W / m 2 , 320W / m 2 , according to the ballistic characteristics, the peak surface heat flux of the space transfer vehicle is not more than 1.5kW / m 2 .

[0094] The outer surface of the multilayer thermal insulation material used at present is double-sided aluminum-coated polyimide film, the thickness of which has a solar absorption ratio of 0.12 and a hemispherical emissivity of 0.06. It is assumed that the surface of the multilayer thermal insulation material quickly reaches thermal equilibrium when it is exposed to light, i.e. the heat absorbed by the surface of the multilayer thermal insulation material Q 吸收 is equal to the surface radiation heat Q 辐射 , according to the Stefan-Boltzmann law, the peak temperature of the outer surface of the space transfer vehicle is calculated:

[0095] Q 吸收 = Q 辐射

[0096] α s Q=εσT 4

[0097] In the formula, α s is the solar absorption rate, Q is the peak heat flux received by the outer surface of the vehicle, ε is the surface emissivity of the material, σ is the Stefan-Boltzmann constant, and T is the outer surface temperature. The calculated peak temperature of the outer surface of the vehicle is 200℃.

[0098] Step three: The heat leakage analysis is carried out for the liquid hydrogen tank coated with foamed material and multilayer insulation material, and the schematic diagram of the insulation material coating is asFigure 2 As shown. The tank wall is 5mm thick and made of stainless steel. A 20mm thick foam layer is tightly wrapped around the wall, made of polyurethane foam. MLI material is then applied over the foam layer. The arrangement can be a variable density arrangement as shown in the figure, or multiple layers of insulation material of the same density can be used for wrapping.

[0099] The temperature boundary condition within the computational region is 25K (liquid hydrogen temperature range), and the temperature boundary condition outside the region is 473.15K (the peak temperature calculated in step two), corresponding to the input conditions of the space thermal environment module. A one-dimensional steady-state heat transfer model is established for the above region, where the tank wall and foamed material are purely thermally conductive models. The heat flux density of each unit in the multilayer insulation material is calculated using the following formula:

[0100] Heat transfer via radiation from the reflective layer:

[0101]

[0102] Among them, T H and T C ε represents the temperature of the surface of the heat-reflecting layer and the surface of the cold-reflecting layer, respectively. H and ε C The emissivity of the heat-reflecting layer and the cold-reflecting layer are respectively; the emissivity of the double-sided aluminized polyimide film is 0.06, and σ is...

[0103] The Stefan-Boltzmann constant has a value of 5.67 × 10⁻⁶. -8 W / (m 2 K 4 ).

[0104] Thermal conductivity of the spacer layer solid:

[0105]

[0106] Where C2 is an empirical constant related to the spacer material, taken as 0.025, and f is the sparseness of the spacer material, taken as 0.02. D x The thickness of the spacer material between the two radiating layers is given. Due to manufacturing limitations, k is the thermal conductivity of the spacer material, taken as 0.05 W / (m). 2 K).

[0107] In a long-term on-orbit environment, gas heat transfer can be neglected. Ultimately, the heat flux density q of each unit... Lt for

[0108] q Lt =q r +q sc ,

[0109] Since the heat flux of the emitting layer is nonlinearly related to the temperature, the heat transfer in the MLI material needs to be solved by iteration. Assuming that the size of the low-density MLI material is 8 units / cm, the size of the medium-density layer MLI material is 12 units / cm, and the size of the high-density layer MLI material is 16 units / cm, the heat transfer calculation is carried out for the arrangement in the arrangement, the full low-density MLI material arrangement and the full high-density MLI material arrangement in Figure 2 , and the total number of multilayer insulation layers is 60 units, and the results are shown in Figure 3 .

[0110] When the full high-density MLI material arrangement is used, the thickness of the insulation layer is 37.5 mm, and the calculated heat flux is 1.7597 W / m 2 ; when the full low-density MLI material arrangement is used, the thickness of the insulation layer is 75 mm, and the calculated heat flux is 1.6112 W / m 2 ; when the variable-density MLI material arrangement is used, the thickness of the insulation layer is about 54.2 mm, and the calculated heat flux is 1.6395 W / m 2 . It can be seen that the heat flux of the full high-density MLI arrangement is larger, the heat flux of the full low-density MLI arrangement is lower, but the overall material thickness is larger, the heat flux of the variable-density MLI arrangement is slightly higher than that of the full low-density MLI arrangement, but the material thickness is moderate, so the variable-density MLI arrangement is selected, and the calculated heat flux of the transfer vehicle is 1.6395 W / m 2 .

[0111] Step four: based on the heat flux calculated by the MLI module and the illumination time calculated by the trajectory module, the evaporation amount of the low-temperature propellant under the flight profile of the transfer vehicle is estimated, the trajectory module is fed back for accurate trajectory analysis, the long-term on-orbit evaporation amount of the low-temperature propellant is recalculated, if the evaporation amount does not meet the allowable requirement of the trajectory, the control measures such as optimizing the trajectory design and optimizing the MLI configuration are used to reduce the heat flux of the tank wall, and the iterative calculation is continuously carried out until the allowable requirement of the trajectory is met.

[0112] The part of the application not described in detail belongs to the known technology of those skilled in the art.

Claims

1. A method for suppressing passive evaporation of cryogenic propellants in a microgravity environment, characterized in that, The method comprises the following steps: S1. According to the transfer vehicle operation task profile, the transfer orbit form is designed, the transfer orbit perigee and apogee are determined, and the transfer orbit period is calculated; S2. The transfer vehicle on-orbit sunlight duration is obtained through dynamic simulation calculation of the transfer orbit equation, the sun position and the earth shadow model; S3. The space external heat flow data and the outer surface temperature of the multilayer insulation material on the surface of the transfer vehicle are calculated; S4. The heat flux density of the multilayer insulation material under different arrangement forms is calculated, and the arrangement form of the multilayer insulation material with the optimal heat insulation effect is selected; S5. Whether the low-temperature propellant carried by the vehicle meets the allowable amount of trajectory is evaluated, the transfer trajectory is optimized, and the iterative calculation is continuously performed until the low-temperature propellant carried by the vehicle meets the allowable amount of trajectory.

2. The method of claim 1, wherein: The calculation of the space external heat flow data and the outer surface temperature of the multilayer insulation material on the surface of the transfer vehicle comprises the following steps: S2.

1. According to the position, speed and attitude angle information of the transfer vehicle in the launch inertial coordinate system during the on-orbit period, the vector of the transfer vehicle and the sunlight direction, the vector of the transfer vehicle and the earth radiation heat, and the ratio of the flight height to the earth radius are calculated, and the space external heat flow data on the surface of the transfer vehicle are calculated; S2.

2. Assuming that the surface of the multilayer insulation material of the transfer vehicle reaches thermal equilibrium when it is illuminated, the outer surface temperature of the multilayer insulation material is obtained according to the space external heat flow data on the surface of the transfer vehicle, the MLI outer surface film absorption ratio and the emissivity.

3. The method of claim 2, wherein the method is characterized by: The calculation of the heat flux density of the multilayer insulation material under different arrangement forms and the selection of the arrangement form of the multilayer insulation material with the optimal heat insulation effect comprise the following steps: S3.

1. The arrangement form of the multilayer insulation material is given, including the given number of radiation screens and the thickness of the spacer layer, the outer surface temperature of the multilayer insulation material obtained in S2.2 is taken as the temperature boundary condition, and the initial temperature distribution in the multilayer insulation material is set; according to the initial temperature distribution in the multilayer insulation material, the heat flux density of radiation heat exchange, gas heat transfer and solid heat conduction in each layer of the multilayer insulation material is calculated; S3.

2. The temperature distribution in the multilayer insulation material is calculated according to the heat flux density, and is compared with the set initial temperature distribution in the multilayer insulation material; if the deviation between the calculated temperature distribution in the multilayer insulation material and the set initial temperature distribution in the multilayer insulation material exceeds the set threshold, the initial temperature distribution in the multilayer insulation material is reset, and the iterative calculation is performed until the calculated temperature distribution in the multilayer insulation material is equal to the set temperature distribution in the multilayer insulation material, and the total heat flux density under the arrangement form of the multilayer insulation material given in S3.1 is obtained; S3.

3. With the process constraints, mass or size constraints of the multilayer insulation material as constraint conditions, S3.1 and S3.2 are repeated to calculate the heat flux density of the multilayer insulation material under different arrangement forms; S3.

4. The heat flux densities of the multilayer insulation material under different arrangement forms are compared, and the arrangement form of the MLI material with the optimal heat insulation effect is selected.

4. The method of claim 3, wherein the method is characterized by: The evaluation of whether the low-temperature propellant carried by the vehicle meets the allowable amount of trajectory, the continuous iterative calculation through the optimization of the transfer trajectory, and the satisfaction of the low-temperature propellant carried by the vehicle to the allowable amount of trajectory comprise the following steps: According to the multi-layer insulation material arrangement form selected in S3.4 with the optimal adiabatic effect and the on-orbit sunlight duration of the transfer vehicle obtained in S2, the on-orbit evaporation amount of the cryogenic propellant is calculated; According to the on-orbit evaporation amount of the cryogenic propellant, the propellant consumption amount of the transfer vehicle on the transfer orbit is calculated, and the accurate transfer trajectory form is determined; According to the trajectory data of the accurate transfer trajectory form and the safety margin coefficient, whether the cryogenic propellant carried by the transfer vehicle meets the allowable amount of the trajectory is evaluated; if the cryogenic propellant carried by the transfer vehicle does not meet the allowable amount of the trajectory, the method ends; if the cryogenic propellant carried by the transfer vehicle does not meet the allowable amount of the trajectory, the transfer trajectory form is redesigned, and S1-S5 is repeated until the cryogenic propellant carried by the transfer vehicle meets the allowable amount of the trajectory.

5. A passive evaporation mitigation system for cryogenic propellants in microgravity, characterized in that, It comprises: a trajectory module, configured to design a transfer orbit form according to a transfer vehicle operation mission profile, determine a transfer orbit perigee and apogee, and calculate a transfer orbit period; the on-orbit sunlight duration of the transfer vehicle is obtained through dynamic simulation calculation of a transfer orbit equation, a sun position and an earth shadow model; a space thermal environment module, configured to calculate space external heat flux data and multi-layer insulation material outer surface temperature received by the transfer vehicle surface; an MLI material module, configured to calculate heat flux density of the multi-layer insulation material under different arrangement forms, and select a multi-layer insulation material arrangement form with the optimal adiabatic effect; an optimization iteration module, configured to evaluate whether the cryogenic propellant carried by the transfer vehicle meets the allowable amount of the trajectory, continuously iteratively calculate by optimizing the transfer trajectory, and until the cryogenic propellant carried by the transfer vehicle meets the allowable amount of the trajectory.

6. The system of claim 5, wherein: The space thermal environment module comprises: according to the position, speed and attitude angle information of the transfer vehicle during on-orbit period in the launch inertial coordinate system, the vector of the transfer vehicle and sunlight, the vector of the transfer vehicle and earth radiation heat, and the ratio of flight height to earth radius are calculated, and the space external heat flux data received by the transfer vehicle surface is calculated; according to the space external heat flux data received by the transfer vehicle surface, the MLI outer surface film absorption ratio and the emissivity, the outer surface temperature of the multi-layer insulation material is obtained.

7. The system of claim 6, wherein: The MLI material module comprises: with the process constraint, mass or size constraint of the multi-layer insulation material as the constraint condition, the heat flux density of the multi-layer insulation under different arrangement forms is calculated; the heat flux density of the multi-layer insulation material under different arrangement forms is compared, and the MLI material arrangement form with the optimal adiabatic effect is selected.

8. The method of claim 7, wherein the method is characterized by: The method for calculating the heat flux density of the multi-layer insulation under different arrangement forms with the process constraint, mass or size constraint of the multi-layer insulation material as the constraint condition comprises: with the process constraint, mass or size constraint of the multi-layer insulation material as the constraint condition, the multi-layer insulation material arrangement form is given, including the given number of radiation screens and the thickness of the spacer layer, the outer surface temperature of the multi-layer insulation material calculated in the space thermal environment module is taken as the temperature boundary condition, and the initial temperature distribution in the multi-layer insulation material is set; according to the initial temperature distribution in the multi-layer insulation material, the heat flux density of radiation heat transfer, gas heat transfer and solid heat conduction in each layer of the multi-layer insulation material is calculated; The temperature distribution in the multi-layer thermal insulation material is calculated according to the heat flux density, and is compared with the set initial temperature distribution in the multi-layer thermal insulation material; if the deviation between the calculated temperature distribution in the multi-layer thermal insulation material and the set initial temperature distribution in the multi-layer thermal insulation material exceeds a set threshold value, the temperature distribution in the multi-layer thermal insulation material is re-set, and iterative calculation is performed until the calculated temperature distribution in the multi-layer thermal insulation material is equal to the set temperature distribution in the multi-layer thermal insulation material, so as to obtain the total heat flux density under the given arrangement form of the multi-layer thermal insulation material.

9. The system of claim 7, wherein: The optimization iteration module comprises: According to the arrangement form of the MLI material selected from the MLI material module and having the optimal heat insulation effect and the calculated on-orbit sunlight duration of the transfer carrier in the ballistic module, the on-orbit evaporation amount of the cryogenic propellant is calculated; According to the on-orbit evaporation amount of the cryogenic propellant, the propellant consumption amount of the transfer carrier on the transfer orbit is calculated, and the accurate transfer ballistic form is determined; According to the ballistic data of the accurate transfer ballistic form and the safety margin coefficient, whether the cryogenic propellant carried by the transfer carrier meets the allowable amount of the ballistic is evaluated.

10. The system of claim 9, wherein: The optimization iteration module further comprises: If the cryogenic propellant carried by the transfer carrier does not meet the allowable amount of the ballistic, the ballistic module is returned, and the transfer ballistic form is re-designed.

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