Satellite fairing safety temperature drop calculation method

Abstract

This invention discloses a method for calculating the safe temperature drop of a satellite fairing, including: S1 calculating the area and volume of the satellite fairing; S2 calculating the heat transfer coefficient of the satellite fairing; S3 calculating the heat dissipation of each structural parameter of the fairing assembly; S4 calculating the heat loss of the fairing assembly during transportation; S5 calculating the heat capacity of the fairing assembly; S6 calculating the temperature drop of the fairing assembly based on the principle of heat conservation between the heat lost and the heat reduced by the fairing assembly; and S7 plotting the temperature drop curve over time. This invention can accurately calculate the temperature drop inside the fairing assembly during relocation and tower opening, providing a precise and quantitative basis for air conditioning temperature and humidity control. Furthermore, by introducing parameters of temperature drop, time, and ambient temperature, a functional relationship is established, providing a reasonable and feasible technical approach for calculating the fairing temperature drop, meeting usage requirements, and establishing a temperature drop table curve calculation method for easy engineering application.

Description

Technical Field

[0001] This invention belongs to the technical field of fairing temperature drop, specifically relating to a method for calculating the safe temperature drop of a satellite fairing. Background Technology

[0002] The fairing is a crucial protective device during satellite transfer and launch, serving functions such as protection, moisture prevention, and thermal insulation. It acts as a vital barrier for the protection of the satellite's precision electronic equipment. After the satellite is assembled, the resulting fairing assembly relies on its initial temperature and thermal inertia to maintain the internal temperature of the fairing during the transfer from the technical facility to the launch pad (or during tower opening), without the need for air conditioning. The temperature drop inside the fairing during transport is closely related to factors such as its initial temperature, structure, outdoor temperature, and transport time. If the temperature drop is excessive, exceeding safety requirements, it can adversely affect the satellite.

[0003] Currently, the satellite fairing assembly is difficult to meet the 15°C requirement when the outdoor ambient temperature is low, which may have a safety impact on the operation of products on the satellite and the prevention of condensation. At the same time, the selection of the transfer time window is also affected by the outdoor ambient temperature. Therefore, considering the temperature requirements of the satellite fairing assembly, the outdoor ambient temperature, and the transfer time requirements, it is necessary to pre-adjust the temperature of the satellite fairing assembly in advance. In addition, considering the pre-adjustment temperature capability of the plant's air conditioning, it is necessary to accurately predict the temperature drop of the fairing to provide a quantitative reference for the air conditioning temperature adjustment before transfer and tower opening.

[0004] However, existing technologies lack a specific solution for calculating the temperature drop of satellite fairings; similar heat transfer calculation schemes require establishing complex finite element models and calculating heat flux density by setting heat transfer temperatures. The calculation software used is commercial software, which suffers from problems such as long calculation time, complex calculation process, lack of specificity, high professional requirements, and limited applicability. Summary of the Invention

[0005] The purpose of this invention is to address the aforementioned shortcomings in the prior art by providing a method for calculating the safe temperature drop of a satellite fairing, thereby solving the problem of the lack of a solution for calculating the temperature drop of a satellite fairing in the prior art.

[0006] To achieve the above objectives, the technical solution adopted by the present invention is as follows:

[0007] A method for calculating the safe temperature drop of a satellite fairing, comprising the following steps:

[0008] S1. Calculate the area and volume of the satellite fairing based on its structural parameters;

[0009] S2. Calculate the heat transfer coefficient of the satellite fairing based on the thermal resistance of the enclosure structure.

[0010] S3. Based on the heat transfer coefficient of the satellite fairing, the heat dissipation of each structural parameter of the fairing assembly is calculated using the equivalent calculation method.

[0011] S4. Calculate the heat loss of the satellite fairing assembly during transportation based on the heat dissipation of each structural parameter of the satellite fairing.

[0012] S5. Calculate the heat capacity of the star-shaped array based on the heat storage medium inside the star-shaped array.

[0013] S6. Calculate the heat loss and heat reduction of the star-shaped array assembly, respectively, and calculate the temperature drop of the star-shaped array assembly based on the principle of heat conservation between the heat loss and heat reduction of the star-shaped array assembly.

[0014] S7. Calculate the temperature drop of the star shield assembly over time under different outdoor ambient temperatures, and plot the temperature drop curve over time.

[0015] Furthermore, the structural parameters of the satellite fairing in step S1 include the cylindrical section, the conical section, the end cap, and the transport vehicle base, and the material parameters of the satellite fairing include foam and fiberglass.

[0016] Furthermore, step S2 specifically includes the following:

[0017] Based on the internal surface thermal resistance R 内 Fiberglass wall panel thermal resistance R 壁板 Thermal resistance R of insulation foam 泡沫 and external surface thermal resistance R 外 Calculate the thermal resistance R of the fairing enclosure structure. 罩 :

[0018] R 罩 =R 内 +R 壁板 +R 泡沫 +R 外

[0019] Thermal resistance R of fairing enclosure structure 罩 Calculate the heat transfer coefficient K of the satellite fairing 罩 :

[0020]

[0021]

[0022] Among them, K 车 R is the heat transfer coefficient of the transport vehicle. 车 For the thermal resistance of the transport vehicle.

[0023] Furthermore, in step S3, the heat dissipation of the cylindrical section, conical section, end, and transport vehicle base is calculated using the following formulas:

[0024] Q = K·S

[0025] Where K is the heat transfer coefficient, S is the area, and Q is the heat dissipation.

[0026] Further, in step S4, the heat loss W during the transport of the star-shaped array assembly is calculated. 罩 ,include:

[0027] W 罩 = (Q1 + Q2 + Q3 + Q4) × 12.79

[0028] Where Q1 is the heat dissipation of the cylindrical section, Q2 is the heat dissipation of the conical section, Q3 is the heat dissipation of the end, and Q4 is the heat dissipation of the transport vehicle base.

[0029] Further, in step S5, the heat capacity C of the star-shaped array assembly is calculated. 星罩 ,include:

[0030] C 星罩 =A1×α1+A2×α2+A3×α3

[0031] Where A1 is the weight percentage of the thermal insulation foam, A2 is the weight percentage of the fiberglass wall, and A3 is the weight percentage of the air; α1, α2, and α3 are the specific heats of the thermal insulation foam, fiberglass wall, and air, respectively.

[0032] Furthermore, step S6 specifically includes the following:

[0033] Calculate the heat loss Q1 of the star shield assembly based on the heat loss during transportation:

[0034] Q1 = W 罩 ×t

[0035] Where t is time;

[0036] Based on the heat capacity calculation of the star shield assembly, the heat reduction Q2 of the star shield assembly is:

[0037] Q2 = C 星罩 ×G×Δt

[0038] Where G is the mass of the star-shaped shield assembly, and Δt is the temperature drop;

[0039] Based on the principle of heat conservation between the heat lost by the star-shaped array and the heat reduced by the star-shaped array, the temperature drop Δt of the star-shaped array is calculated.

[0040] Q1 = Q2

[0041]

[0042] The satellite fairing safety temperature drop calculation method provided by this invention has the following beneficial effects:

[0043] This invention can accurately calculate the temperature drop inside the hood assembly during the transfer and tower opening process, providing a precise and quantitative basis for air conditioning temperature and humidity control.

[0044] The fairing safety temperature drop calculation method of the present invention introduces parameters such as temperature drop, time, and ambient temperature by establishing an equivalent heat transfer method, optimizing and simplifying the method, and using the high-altitude wind force additional coefficient method, and establishes a functional relationship. This provides a reasonable and feasible technical approach for calculating the fairing temperature drop, making the calculation results accurate to within 2%, meeting the usage requirements, and establishing a temperature drop table curve calculation method, which is convenient for engineering applications.

[0045] This invention calculates the safe temperature drop inside the satellite shield during the relocation and tower opening process of the satellite shield assembly, thereby ensuring that the satellite does not condense during the relocation and tower opening process and remains within the required range.

[0046] This invention applies equivalent heat transfer coefficients to the bottom of the satellite fairing, the transport vehicle, and the section docked with the rocket. It ignores the satellite's own structure, which includes insulation and air gaps, simplifying and optimizing the calculation method. This invention pioneers an effective method for calculating the overall safe temperature drop of the satellite fairing assembly. Attached Figure Description

[0047] Figure 1 This is a schematic diagram of the fairing structure of the present invention.

[0048] Figure 2 This invention illustrates the temperature drop of the star cover assembly over time under different outdoor temperatures. Detailed Implementation

[0049] The specific embodiments of the present invention are described below to enable those skilled in the art to understand the present invention. However, it should be understood that the present invention is not limited to the scope of the specific embodiments. For those skilled in the art, various changes are obvious as long as they are within the spirit and scope of the present invention as defined and determined by the appended claims. All inventions utilizing the concept of the present invention are protected.

[0050] Example 1: The method for calculating the safe temperature drop of the satellite fairing in this scheme specifically includes the following steps:

[0051] Step S1: Calculate the area and volume of the satellite fairing based on its structural parameters;

[0052] The structural parameters of the satellite fairing in this step include the cylindrical section, the conical section, the end cap, and the base of the transport vehicle. The material parameters of the satellite fairing include foam and fiberglass.

[0053] For details, please refer to Figure 1 The satellite fairing is mainly made of fiberglass with a thickness of 3mm. It is covered with thermal insulation foam material with a thickness of 55mm. The structure includes cylindrical sections, conical sections, and end caps. The main material properties are shown in Table 1, and the structural dimensions are shown in Table 2.

[0054] Table 1. Composition of rectifier and transport vehicle

[0055]

[0056] Table 2 Fairing Structural Dimensions

[0057]

[0058]

[0059] Step S2: Calculate the heat transfer coefficient of the satellite fairing based on the thermal resistance of its enclosure structure.

[0060] Based on the internal surface thermal resistance R 内 Fiberglass wall panel thermal resistance R 壁板 Thermal resistance R of insulation foam 泡沫 and external surface thermal resistance R 外 Calculate the thermal resistance R of the fairing enclosure structure. 罩 :

[0061] R 罩 =R 内 +R 壁板 +R 泡沫 +R 外

[0062] Among them, R 内 =0.11m 2 ·K / W; R 外 =0.04m 2 K / W;

[0063]

[0064]

[0065] Where h is the material thickness in mm; λ is the thermal conductivity of the material in W / (m·K); α λ This is a correction factor for thermal conductivity;

[0066] Thermal resistance R of fairing enclosure structure 罩 Calculate the heat transfer coefficient K of the satellite fairing 罩 :

[0067]

[0068]

[0069] Among them, K 车 R is the heat transfer coefficient of the transport vehicle. 车 For the thermal resistance of the transport vehicle.

[0070] Step S3: Based on the heat transfer coefficient of the satellite fairing, the heat dissipation of each structural parameter of the fairing assembly is calculated using the equivalent calculation method. This specifically includes the following:

[0071] The fairing's heat dissipation during relocation and transportation mainly involves two parts: heat dissipation through the fairing body to the external environment, and heat transfer to the outside environment through the transport vehicle's base. The specific heat dissipation amounts for these two parts are calculated using the following formula:

[0072] Q = K·S

[0073] Where K is the heat transfer coefficient, in W / (m²). 2 ·K); S is the area, in meters. 2 Q represents heat dissipation;

[0074] The heat dissipation of the cylindrical section, conical section, end, and transport vehicle base were calculated using the above formulas and are 33.31 W / K, 10.06 W / K, 6.03 W / K, and 58.67 W / K, respectively, as shown in Table 3.

[0075] Table 3. Fairing Structural Dimensions

[0076]

[0077] The lower space of the satellite array is sealed by the base of the transport vehicle during the transfer. After the satellite array is hoisted onto the tower, it is sealed by the instrument compartment on the upper part of the rocket. Both are made of metal and have the same area. The heat transfer can be calculated using equivalent values.

[0078] By applying reasonable equivalence to the structure of the shield and the bottom seal of the star-shield assembly, this equivalent calculation method is reasonable and feasible, establishing a reasonable and feasible technical approach for calculating the temperature drop of the star-shield assembly.

[0079] Step S4: Calculate the heat loss W of the satellite fairing assembly during transportation based on the heat dissipation of each structural parameter. 罩 ,include:

[0080] W 罩 = (Q1 + Q2 + Q3 + Q4) × 12.79

[0081] Where Q1 is the heat dissipation of the cylindrical section, Q2 is the heat dissipation of the conical section, Q3 is the heat dissipation of the end, and Q4 is the heat dissipation of the transport vehicle base.

[0082] In this embodiment, the outdoor temperature during the transitional season is 10.2℃, the temperature inside the enclosure at the start of transport is 21.88℃, the temperature inside the factory building outside the enclosure is 24.1℃, the average temperature is 22.99℃, the heat transfer temperature difference is 12.79℃, and the transfer time t from opening the factory gate to transporting to the tower is 65 minutes. Substituting the data in Table 3 into the formula of this step, the heat consumption of the star-enclosure assembly is calculated, and the result is:

[0083] W 置 =(33.31+10.06+6.03+58.67)×12.79=1382W=1.382kJ / s

[0084] Step S5: Calculate the heat capacity C of the star-shaped array assembly based on the heat storage medium inside the star-shaped array assembly. 星罩 ,include:

[0085] C 星罩 =A1×α1+A2×α2+A3×α3

[0086] Where A1 is the weight percentage of thermal insulation foam, A2 is the weight percentage of fiberglass wall, and A3 is the weight percentage of air; α1, α2, and α3 are the specific heats of thermal insulation foam, fiberglass wall, and air, respectively.

[0087] Specifically, the heat storage medium within the satellite enclosure mainly includes insulating foam, fiberglass panels, air, and the satellite itself. Because the satellite has thermal control materials and is isolated from the external environment by an air gap, fiberglass panels, and insulating foam, heat transfer is relatively small, so the satellite's heat storage is not considered here. The specific heat of air is 1.01 kJ / (kg·K), and its volume is approximately 40.14 m³. 3 Its density is 1.2 kg / m³. 3 The thermal insulation foam has a specific heat of 1.47 kJ / (kg·K) and a volume of 4.58 m³. 3 Its density is 40 kg / m³ 3 The fiberglass weighs approximately 1300 kg and has a specific heat capacity of 1.26 kJ / (kg·K). Substituting the data from Table 4 into the formula for this step, we can obtain:

[0088] C 星罩 =1.47×11.97%+1.26×84.88%+1.01×3.15%=1.277kJ / (kg·K)

[0089] The method for calculating the average specific heat capacity of the shroud is to determine the percentage by weight, and then use a weighted average algorithm to calculate the weighted average specific heat capacity of the shroud.

[0090] An algorithm combining optimization and simplification was adopted. The satellite has thermal control measures and is isolated from the surrounding air layer, with good thermal insulation. A simplified algorithm that ignores the satellite body was used. The calculation results show that the algorithm is reasonable and feasible for calculating this type of problem.

[0091] Table 4. Calculation of Equivalent Specific Heat of Rectifier

[0092] Serial Number Material proportion(%) Specific heat kJ / (kg·K) 1 Thermal insulation foam 11.97 1.47 2 Fiberglass wall 84.88 1.26 3 Air 3.15 1.01 4 Star Cover 1.277

[0093] Step S6: Calculate the heat loss and heat reduction of the star-shaped array assembly, and based on the principle of heat conservation between the heat loss and heat reduction of the star-shaped array assembly, calculate the temperature drop of the star-shaped array assembly, which specifically includes the following:

[0094] Calculate the heat loss Q1 of the star shield assembly based on the heat loss during transportation:

[0095] Q1 = W 罩 ×t

[0096] Where t is time;

[0097] Based on the heat capacity calculation of the star shield assembly, the heat reduction Q2 of the star shield assembly is:

[0098] Q2 = C 星罩 ×G×Δt

[0099] Where G is the mass of the star-shaped shield assembly, and Δt is the temperature drop;

[0100] Based on the principle of heat conservation between the heat lost and the heat reduced by the fairing assembly, i.e., assuming no mass exchange between the fairing and the outside environment, the heat lost by the fairing manifests as a decrease in temperature inside the fairing. Therefore, the temperature drop Δt of the fairing assembly is calculated:

[0101] Q1 = Q2

[0102]

[0103] After substituting the data, the temperature drop Δt of the star-shaped shield assembly was calculated:

[0104]

[0105] As can be seen from the calculation formula in this step, the introduction of the time parameter in this step can establish a functional relationship between the time parameter and the temperature drop.

[0106] Furthermore, the calculated temperature drop Δt was compared and analyzed with the actual measurement results. The temperature of the star cover assembly when it left the factory was 21.88℃, and when it was transported to the tower, it was 19.09℃, with a time of 65 minutes. The actual measured temperature difference was 2.79℃, and the calculated temperature difference was 2.76℃. The absolute value of the calculated temperature difference and the actual measured temperature difference was 0.03℃, and the relative deviation was 1.1%, which is almost completely consistent with the actual result. This verifies the scientific nature of the calculation method of this invention and the correctness of the calculation results.

[0107] Step S7: Calculate the temperature drop of the satellite array assembly over time under different outdoor ambient temperatures, and plot the temperature drop curve over time. This includes the following:

[0108] This step is a specific application of steps S1 to S6. When the average temperature of the hood is 22.99℃, the calculation method of step S6 is used to calculate the changes of the star hood assembly with time under different outdoor ambient temperatures. The calculated data is shown in Table 5.

[0109] Table 5. Temperature drop calculation for the satellite shield assembly under different outdoor environments.

[0110]

[0111] The application of the temperature drop calculation results in this embodiment in engineering is as follows:

[0112] Based on the calculation results in Table 5, plot the temperature drop curve over time, as follows: Figure 2 As shown, the horizontal axis represents time, i.e., the time without air conditioning, in minutes; the left side represents the temperature drop inside the fairing, in degrees Celsius; and the right side represents different outdoor ambient temperatures.

[0113] In practical engineering applications, the temperature drop inside the fairing can be quickly determined by checking the time on the horizontal axis of the temperature drop versus time curve and the corresponding outdoor ambient temperature curves. By consulting the fairing temperature drop table and fairing temperature drop curve, the temperature drop changes inside the fairing can be easily and quickly monitored, allowing for the selection of appropriate time windows for tasks such as opening the launch tower and rocket relocation.

[0114] As a second application of this embodiment, this embodiment can further realize the application of the temperature drop calculation model of the star-shaped array, as follows:

[0115] The temperature drop calculation model of the radiator assembly was applied to a general inspection process. Before the test, the temperature inside the radiator was adjusted to 24℃, the average temperature of the radiator was 23.0℃, and the outdoor ambient temperature was 9℃. After the tower was opened and the air ducts were retracted, the duration without air conditioning support was approximately 75 minutes. Based on the temperature drop calculation results of steps S1 to S7 of this embodiment and the temperature drop change curve over time, the calculated temperature drop inside the radiator was approximately 3.5℃. When the wind was strong at a high altitude in the open field, with a wind force additional rate of 10%, the corrected temperature drop result of the radiator was 3.85℃. The actual temperature drop result inside the radiator during the monitoring process was 3.90℃. The predicted temperature drop value was consistent with the actual monitored temperature drop result, with an absolute error of only 0.05℃ and a relative error of 1.2%, which was almost completely consistent with the actual result. Therefore, this embodiment can provide quantitative reference for further temperature control scheme formulation and air conditioning environment protection.

[0116] Although specific embodiments of the invention have been described in detail with reference to the accompanying drawings, this should not be construed as limiting the scope of protection of this patent. Various modifications and variations that can be made by a person skilled in the art without inventive effort within the scope described in the claims still fall within the scope of protection of this patent.

Claims

1. A method for calculating the safe temperature drop of a satellite fairing, characterized in that, Includes the following steps: S1. Calculate the area and volume of the satellite fairing based on its structural parameters; S2. Calculate the heat transfer coefficient of the satellite fairing based on the thermal resistance of the enclosure structure. S3. Based on the heat transfer coefficient of the satellite fairing, the heat dissipation of each structural parameter of the fairing assembly is calculated using the equivalent calculation method. In step S3, the heat dissipation of the cylindrical section, conical section, end, and transport vehicle base is calculated using the following formulas respectively: Where K is the heat transfer coefficient, S is the area, and Q is the heat dissipation; S4. Calculate the heat loss during transportation of the satellite fairing assembly based on the heat dissipation of each structural parameter; step S4 calculates the heat loss W during transportation of the satellite fairing assembly. 罩 ,include: W 罩 =(Q1+ Q2+Q3+Q4)×12.79 Where Q1 is the heat dissipation of the cylindrical section, Q2 is the heat dissipation of the conical section, Q3 is the heat dissipation of the end, and Q4 is the heat dissipation of the transport vehicle base. S5. Calculate the heat capacity of the star-shaped enclosure assembly based on the heat storage medium within the assembly; in step S5, the heat capacity C of the star-shaped enclosure assembly is calculated. 星罩 ,include: C 星罩 =A1×α1+ A2×α2+ A3×α3 Where A1 is the weight percentage of thermal insulation foam, A2 is the weight percentage of fiberglass wall, and A3 is the weight percentage of air; α1, α2, and α3 are the specific heats of thermal insulation foam, fiberglass wall, and air, respectively. S6. Calculate the heat loss and heat reduction of the star-shaped array assembly, respectively, and calculate the temperature drop of the star-shaped array assembly based on the principle of heat conservation between the heat loss and heat reduction of the star-shaped array assembly. S7. Calculate the temperature drop of the star shield assembly over time under different outdoor ambient temperatures, and plot the temperature drop curve over time.

2. The method for calculating the safe temperature drop of a satellite fairing according to claim 1, characterized in that, The structural parameters of the satellite fairing in step S1 include the cylindrical section, the conical section, the end head, and the transport vehicle base. The material parameters of the satellite fairing include foam and fiberglass.

3. The method for calculating the safe temperature drop of a satellite fairing according to claim 1, characterized in that, Step S2 specifically includes the following: Based on the internal surface thermal resistance R 内 Fiberglass wall panel thermal resistance R 壁板 Thermal resistance R of insulation foam 泡沫 and external surface thermal resistance R 外 Calculate the thermal resistance of the fairing enclosure structure. : Thermal resistance of fairing enclosure structure Calculate the heat transfer coefficient of the satellite fairing : in, The heat transfer coefficient of the transport vehicle, For the thermal resistance of the transport vehicle.

4. The method for calculating the safe temperature drop of a satellite fairing according to claim 1, characterized in that, Step S6 specifically includes the following: Calculate the heat loss of the star-shaped array assembly based on the heat loss during its transportation. : in, t For time; Based on the heat capacity calculation of the star-shaped shield assembly, the heat reduction of the star-shaped shield assembly... Q 2: in, G For the mass of the star shield assembly, To lower the temperature; Based on the principle of heat conservation between the heat lost and the heat reduced by the star-shaped enclosure assembly, the temperature drop of the star-shaped enclosure assembly is calculated. ; = 。

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