A method and system for estimating the mass of a two-stage-to-orbit spacecraft

Abstract

This invention belongs to the field of aircraft design and relates to a method and system for estimating the mass of a two-stage-to-orbit spacecraft. The method involves first determining the mission profile and payload, then calculating the mass of the second-stage spacecraft, followed by the fuel mass and takeoff mass of the first-stage spacecraft, and finally the structural mass of the first-stage spacecraft. Based on the structural mass of the first-stage spacecraft, the takeoff mass is recalculated. An error calculation is performed on the two calculated takeoff masses. If the error requirement is met, the iteration is considered converged, and the final estimated mass is the first-stage spacecraft structural mass and fuel mass corresponding to the final takeoff mass. If the error does not meet the requirement, the first-stage spacecraft structural mass from the previous iteration is used to recalculate the takeoff mass until the error requirement is met. By using iterative calculation, the error range is continuously narrowed until the accuracy requirement is met, maximizing the approximation of the actual spacecraft mass and providing a reliable data foundation for subsequent processes such as spacecraft design.

Description

Technical Field

[0001] This invention belongs to the field of aircraft design, specifically relating to a method and system for estimating the mass of a two-stage-to-orbit aerospace vehicle. Background Technology

[0002] Horizontal takeoff and landing (LTL) hypersonic weapons and low-cost reusable space-to-ground vehicles are strategic assets for achieving future air superiority, and countries worldwide are focusing on developing LTL reusable combined-powered space vehicles. Vehicle mass estimation is the first step in vehicle design; it allows for the calculation of the required fuel mass, structural mass, and approximate dimensions. While there is considerable research on aircraft mass estimation methods, research on scale assessment methods for complete two-stage-to-orbit (TTO) space-to-ground vehicles is relatively limited.

[0003] A search revealed that among existing patents, the invention patent with publication number CN 115758692 A discloses a method for estimating the mass of a two-stage-to-orbit aerospace vehicle based on statistical analysis. This method uses existing data to fit the relationship between the structural mass and takeoff mass of the first and second stages of the vehicle. It establishes a functional relationship between the fuel mass and the total takeoff mass of the first and second stages based on the vehicle's motion equations, and solves for the vehicle's mass by solving a series of equations. However, this invention's mass assessment does not consider many factors, including the influence of the first-stage vehicle's structural dimensions on its structural mass, the mass of subsystems, and the changes in engine performance parameters caused by the combined propulsion mode transition in the first stage. Furthermore, calculating the relationship between structural mass and takeoff mass using a fitting method can lead to large errors. Summary of the Invention

[0004] The purpose of this invention is to provide a method and system for estimating the mass of a two-stage-to-orbit spacecraft, which solves the problems of insufficient accuracy in mass assessment, incomplete consideration of many key factors, and changes in engine parameters caused by mode transitions of the first-stage spacecraft in the current technology.

[0005] This invention is achieved through the following technical solution:

[0006] This invention discloses a method for estimating the mass of a two-stage-to-orbit spacecraft, comprising the following steps:

[0007] S1. Determine the mission profile and payload mass of the two-stage-to-orbit aerospace vehicle;

[0008] S2. Based on the developed structural mass and engine mass parameters of the second-stage rocket, estimate the structural mass of the second-stage spacecraft of the two-stage-to-orbit aerospace vehicle;

[0009] S3. Based on the mission profile determined in S1, calculate the mass ratio of the second-stage vehicle according to the relationship between the vehicle acceleration process and propellant consumption, and then calculate the propellant mass of the second-stage vehicle.

[0010] Calculate the mass of the second-stage aircraft based on the propellant mass of the second-stage aircraft and the structural mass of the second-stage aircraft obtained from S2.

[0011] S4. Give an initial guess to the structural mass of the first-stage aircraft. Based on the load mass of S1, the mass of the second-stage aircraft in S3, and the operating parameters of the engines in different modes, calculate the fuel mass of the first-stage aircraft according to the constructed fuel mass estimation model of the first-stage aircraft.

[0012] S5. Calculate the takeoff mass of the first-stage aircraft;

[0013] S6. Calculate the structural mass of the first-stage aircraft according to the method for estimating the structural mass of the first-stage aircraft.

[0014] S7. Recalculate the takeoff mass based on the structural mass of the first-stage aircraft obtained in S6;

[0015] S8. Calculate the error of the takeoff mass obtained from S7 and S5. If the error requirement is met, the iteration is considered to have converged. Then, the first-stage aircraft structural mass and first-stage aircraft fuel mass corresponding to the takeoff mass obtained from S7 are used as the final mass estimate.

[0016] If the error does not meet the requirements, return the structural mass of the first-stage aircraft from the previous iteration to S4, and repeat S4-S7 until the error requirements are met.

[0017] Furthermore, in S3, the mass ratio of the second-stage aircraft is calculated based on the relationship between the constructed aircraft acceleration process and propellant consumption, expressed as:

[0018]

[0019] The relationship between the spacecraft's acceleration process and propellant consumption is defined during the construction phase, with the velocity and altitude of each flight node in the mission profile recorded as follows: ,in Indicates takeoff status. Define the required separation point height and velocity for the mission. arrive The flight process is the first Modality, defined as the first The flight mass at the node is ;

[0020] but Indicates that the second-stage aircraft is in the... The velocity at the node, Indicates that the second-stage aircraft is in the... The velocity at the node, Indicates that the second-stage aircraft is in the... Height at the node Indicates that the second-stage aircraft is in the... Height at the node; It is the acceleration due to gravity; Specific impulse of the rocket engine;

[0021] The propellant mass of the second-stage aircraft is calculated based on the mass ratio of the second-stage aircraft, using the following expression:

[0022]

[0023] in, For the propellant mass of the second-stage aircraft, To estimate the structural mass of the second-stage aircraft, For load mass.

[0024] Furthermore, in S3, the mass of the second-stage vehicle is calculated based on the propellant mass of the second-stage vehicle and the structural mass of the second-stage vehicle obtained in S2; specifically:

[0025]

[0026] It is a second-level aircraft mass.

[0027] Furthermore, in S4, the expression for the first-stage aircraft fuel mass estimation model is as follows:

[0028]

[0029] in, For the fuel mass of a first-class aircraft, For the structural mass of a first-class aircraft, For load mass, For a second-level aircraft mass, The total mass ratio of all modes for propulsion of a first-stage aircraft.

[0030] Furthermore, the formula for calculating the total mass ratio of all modes of propulsion for a first-stage aircraft is as follows:

[0031]

[0032] in, The mass ratio of different modes of a first-stage aircraft, where i represents the mass ratio of the first-stage aircraft. Modality, where s represents the final node.

[0033] Furthermore, in S5, the formula for calculating takeoff mass is:

[0034]

[0035] in, For the structural mass of a first-class aircraft, For load mass, For a second-level aircraft mass, For the fuel mass of a first-class aircraft, For takeoff mass.

[0036] Furthermore, in S6, the structural mass of the first-stage aircraft is calculated according to the first-stage aircraft structural mass estimation method, specifically as follows:

[0037] The formula for calculating the structural mass of a first-stage aircraft is:

[0038]

[0039] in, For the sake of fuselage quality, For wing mass, For the tail mass, For landing quality, For the quality of other systems.

[0040] Furthermore, fuselage weight:

[0041]

[0042] Wing mass:

[0043]

[0044] Tail mass:

[0045]

[0046] Landing quality:

[0047]

[0048] in, For aspect ratio, For fuselage height, For fuselage length, Mach number, For overload limit, For the rudder area, The total area of ​​the control surfaces. The vertical tail area, The area of ​​the trapezoidal wing. For the relative thickness of the airfoil, For the width of the fuselage, For takeoff mass, For wingtip to tip ratio, The MAC wing sweep angle is 25%. The MAC tail sweep angle is 25%. For wingtip to tip ratio, For takeoff mass.

[0049] Furthermore, in S8, the formula for calculating the error is:

[0050]

[0051] in, The takeoff mass calculated in S5. This is the takeoff mass recalculated in S7.

[0052] This invention also discloses a two-stage-to-orbit spacecraft mass estimation system for implementing the aforementioned two-stage-to-orbit spacecraft mass estimation method, comprising:

[0053] The mission module is used to set the mission profile and payload mass of a two-stage-to-orbit aerospace vehicle.

[0054] The second-stage spacecraft structural mass prediction module predicts the structural mass of the second-stage spacecraft based on the developed second-stage rocket structural mass and engine mass parameters.

[0055] The second-stage vehicle mass calculation module is used to calculate the second-stage vehicle mass ratio and the propellant mass of the second-stage vehicle based on the relationship between the vehicle acceleration process and propellant consumption.

[0056] The mass of the second-stage aircraft is calculated based on the propellant mass and structural mass of the second-stage aircraft.

[0057] The first-stage vehicle fuel mass estimation module is used to calculate the first-stage vehicle fuel mass based on the payload mass, second-stage vehicle mass, and engine operating parameters of different modes, according to the constructed first-stage vehicle fuel mass estimation model.

[0058] The first-stage aircraft structural mass calculation module is used to calculate the structural mass of the first-stage aircraft according to the first-stage aircraft structural mass estimation method.

[0059] Takeoff mass calculation module, used to calculate takeoff mass;

[0060] And recalculate the takeoff mass based on the first-stage aircraft structural mass obtained from the first-stage aircraft structural mass calculation module;

[0061] The error calculation module is used to calculate the error between the two takeoff masses obtained by the takeoff mass calculation module.

[0062] The judgment module is used to determine whether the error requirement is met. If the error requirement is met, the iteration is considered to have converged. The final mass estimate is then taken as the first-stage aircraft structural mass and first-stage aircraft fuel mass corresponding to the last takeoff mass.

[0063] If the error does not meet the requirements, the structural mass of the first-stage aircraft from the previous iteration is input into the first-stage aircraft fuel mass estimation module, the takeoff mass calculation module, and the error calculation module for repeated calculation until the error requirements are met.

[0064] Compared with the prior art, the present invention has the following beneficial technical effects:

[0065] This invention discloses a method for estimating the mass of a two-stage-to-orbit spacecraft. Through multi-step, meticulous calculations, it comprehensively considers numerous key factors such as mission profiles, the structure of each stage of the spacecraft, and engine operating parameters. It is not a simple empirical estimation; each step is cross-checked, resulting in a more accurate final mass estimate. Iterative calculations continuously narrow the error range until the required accuracy is met, maximizing the fit with the actual spacecraft mass and providing a reliable data foundation for subsequent processes such as spacecraft design and ballistic calculations.

[0066] Based on existing parameters of second-stage rockets, and drawing on data from mature technological systems, this approach enables the estimation of spacecraft capabilities. It not only fits the existing technological framework but also flexibly addresses the specific design requirements of new two-stage-to-orbit spacecraft. Furthermore, the constructed model can be flexibly adjusted according to different mission profiles, engine operating modes, and other variables, making it applicable to various types of two-stage-to-orbit missions.

[0067] By incorporating a series of factors, such as the relationship between the acceleration process of an aircraft and propellant consumption, the structural mass of each stage of the aircraft, engine parameters, and payload mass, into a unified calculation system, the limitations of evaluating each subsystem individually are broken, forming a comprehensive and systematic quality assessment process. This ensures that the coordination between the various parts is reflected in the quality prediction stage, which helps to create a well-balanced overall aircraft architecture.

[0068] Furthermore, when estimating the fuel mass of the first-stage spacecraft using the total mass ratio of all modes of propulsion of the first-stage spacecraft, the performance changes of different engine modes of the first-stage spacecraft are comprehensively considered, which is consistent with the actual flight conditions of two-stage-to-orbit aerospace vehicles and can improve the accuracy of mass estimation. At the same time, this calculation method can be applied to other combined propulsion configurations.

[0069] Furthermore, when calculating the mass of the first-stage spacecraft, a comprehensive scale assessment of the complete concepts, including the structural mass, subsystem mass, and fuel mass of the two-stage-to-orbit spacecraft, was considered, making the calculated structural mass of the first-stage spacecraft more consistent with reality. Attached Figure Description

[0070] Figure 1 This is a flowchart of the overall process for estimating the mass of a two-stage-to-orbit aerospace vehicle provided in this embodiment of the disclosure;

[0071] Figure 2 This is a mission profile for a second-stage aircraft.

[0072] Figure 3 This is a mission profile for a first-stage aircraft. Detailed Implementation

[0073] To make the objectives, technical solutions, and advantages of the present invention clearer, the following detailed description is provided in conjunction with the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are merely illustrative of the invention and are not intended to limit the invention; that is, the described embodiments are only a part of the embodiments of the present invention, and not all of them.

[0074] The detailed description of the embodiments of the present invention provided in the following figures is not intended to limit the scope of the claimed invention, but merely to illustrate one selected embodiment of the invention. All other embodiments obtained by those skilled in the art based on the figures and embodiments of the present invention without inventive effort are within the scope of protection of the present invention.

[0075] The features and performance of the present invention will be further described in detail below with reference to embodiments.

[0076] like Figure 1 As shown, this invention discloses a method for estimating the mass of a two-stage-to-orbit spacecraft, which mainly includes the following:

[0077] First, determine the payload mass and flight profile of the flight mission. Starting from the mass estimation of the second-stage vehicle, estimate the structural mass of the second-stage vehicle based on experience, determine the mass ratio based on the rocket engine performance and mission profile, and then calculate the mass of the second-stage vehicle.

[0078] Then, the second-stage vehicle and the payload are used together as the load of the first-stage vehicle. The mass ratio of the first-stage vehicle is calculated based on the engine performance mode transition point. The size of the first-stage vehicle is estimated and the structural mass of the first-stage vehicle is calculated. Thus, the mass of the first-stage vehicle and the takeoff mass are calculated. The process is then iterative and converges.

[0079] like Figure 2 and Figure 3 As shown, based on the overall parameters of various developed reusable space transportation systems, and based on the performance boundaries of rocket engines and combined propulsion, the mission profile of a two-stage-to-orbit spacecraft is determined:

[0080] The spacecraft takes off horizontally from the ground, powered by six pre-cooled air turbo rocket engines (PATR). At 18 km (Ma 3.5), the PATR engines shut down, the PATR flow channel splitters close, and the subsonic ramjet mode is activated. At 24 km (Ma 4.0), the subsonic ramjet engines shut down, and the scramjet mode is activated. The first-stage spacecraft carries the second-stage spacecraft to 31 km (Ma 6.8), where the first and second stages separate, and the first-stage spacecraft returns without power. The second-stage spacecraft is propelled by two hydrogen-oxygen rocket engines and ultimately accelerates to low Earth orbit.

[0081] During the design process, it is necessary to first construct the general equations for the acceleration motion of the aircraft, specifically:

[0082] Let the velocity and altitude of each flight node in the mission profile be denoted as . ,in Indicates takeoff status. Define the required separation point height and velocity for the mission. arrive The flight process is the first Modality, defined as the first The flight mass at the node is quality ratio Speed ​​increment The height increment is Specific quality estimation methods will be established subsequently.

[0083] The general equation for the relationship between aircraft acceleration and propellant consumption rate is:

[0084] (1)

[0085] In the formula: For the mass of the aircraft, For the speed of the aircraft, For engine thrust, For aircraft drag, It is the acceleration due to gravity. The ascent angle.

[0086] Define effective specific impulse as , This refers to the engine's specific impulse.

[0087] The general equation for aircraft acceleration and propellant consumption rate is transformed into:

[0088] (2)

[0089] Compared to engine specific impulse, effective specific impulse, which takes into account flight drag, better reflects the acceleration performance of an aircraft. The larger the value, the closer the effective specific impulse is to the engine's specific impulse.

[0090] The process for determining the structural mass of a second-stage aircraft is as follows:

[0091] Determine the load mass Based on the second-stage rocket structure mass and rocket engine mass of various developed rocket types, the second-stage spacecraft structure mass of a two-stage-to-orbit aerospace vehicle is estimated. .

[0092] The calculation process for the mass ratio and total mass of a second-stage aircraft is as follows:

[0093] With load mass and the estimated structural mass of the second-stage aircraft Based on this, the mass ratio of the second-stage aircraft is calculated by analyzing the relationship between the acceleration process and propellant consumption, thereby calculating the propellant mass of the second-stage aircraft. The details are as follows:

[0094] Because the thrust-to-drag ratio of the spacecraft is large in rocket mode, it can be considered that... In addition, rocket engines Since it is approximately constant during flight, integrating both sides of equation (2) yields:

[0095] (3)

[0096] The second term in the above formula is the gravitational loss. Its value is relative to the velocity increment Generally small in size and not accurately obtainable during the conceptual design phase, it is therefore ignored in the initial estimate. The average velocity of the rocket modes is used to estimate Then we have:

[0097] (4)

[0098] Combining formulas (3) and (4), we obtain the mass ratio of the second-stage aircraft:

[0099] (5)

[0100] The mass ratio of a second-stage aircraft can also be expressed as the ratio of the total mass of the aircraft to the mass of the non-propellant portion, from which we can derive:

[0101] (6)

[0102] in, To estimate the propellant mass of the second-stage vehicle, To estimate the structural mass of the second-stage aircraft, For load mass.

[0103] Therefore, by combining formulas (5) and (6), the mass of the propellant in the second-stage aircraft can be calculated, and thus the mass of the second-stage aircraft can be calculated. Specifically:

[0104] .

[0105] Next, we will introduce the relevant content of the first-level aircraft.

[0106] First, construct the total mass ratio of all modes of the first-stage propulsion system, specifically:

[0107] In the air-breathing mode, the thrust of the spacecraft is relatively small, and the drag loss cannot be ignored. Therefore, a different estimation method than that used for the rocket mode is adopted. By multiplying both sides of equation (2) by V and dividing by M, we get:

[0108] (7)

[0109] In the formula, TV represents the engine power, and the engine thermal efficiency is assumed to be... If the heat value is q, then we have

[0110] (8)

[0111] Substituting into equation (7) and taking the drag ratio as a constant, we get:

[0112] (9)

[0113] It can ultimately be written in a unified form:

[0114] (10)

[0115] In equation (10), gravity loss The acceleration performance of a vehicle is determined by its mean effective specific impulse. For air-breathing propulsion mode, the mean effective specific impulse is:

[0116] (11)

[0117] For the PATR engine, a precise component-level model of the engine has been established, and simulation data for various operating conditions has been obtained, with the average specific impulse calculated. The average specific impulse of subsonic ramjet and supersonic ramjet engines is calculated based on combustion efficiency and fuel calorific value.

[0118] Calculate the total mass ratio of all modes of propulsion for a first-stage aircraft:

[0119] (12)

[0120] In equation (12), For load mass, For a second-level aircraft mass, For the fuel mass of a first-class aircraft, For the structural mass of a first-class aircraft, The mass ratio of different engine modes of a first-stage aircraft; i represents the mass ratio of the first-stage aircraft. Modality, where s represents the final node.

[0121] The fuel mass estimation model for a first-stage aircraft can be derived from formula (12). The expression is:

[0122] .

[0123] Design an algorithm for estimating the structural mass of a first-stage aircraft:

[0124] By combining the parameters of various reusable space launch vehicles that have been developed, a structural mass formula based on dimensions is formed for each component.

[0125] Body weight:

[0126] (13)

[0127] Wing mass:

[0128] (14)

[0129] Tail mass:

[0130] (15)

[0131] Landing quality:

[0132] (16)

[0133] in, For aspect ratio, For fuselage height, For fuselage length, Mach number, For overload limit, For the rudder area, The total area of ​​the control surfaces. The vertical tail area, The area of ​​the trapezoidal wing. For the relative thickness of the airfoil, For the width of the fuselage, For takeoff mass, For wingtip to tip ratio, The MAC wing sweep angle is 25%. The MAC tail sweep angle is 25%. The ratio of wingtip to tip.

[0134] The structural mass of the first-stage aircraft is:

[0135] (17)

[0136] in, For the quality of other systems, specifically including the quality of thermal management, landing gear, load system, control system, engine, etc.

[0137] Takeoff mass:

[0138] (18)

[0139] in, For the structural mass of a first-class aircraft, For load mass, For a second-level aircraft mass, This refers to the fuel mass of a first-class aircraft.

[0140] The preceding calculations yield the mass ratio and structural mass of the aircraft. However, the calculation of the structural mass of the first-stage aircraft requires the takeoff mass, and the fuel mass of the first-stage aircraft needs to be calculated using the structural mass and mass ratio. The takeoff mass is the sum of the masses of the first and second-stage aircraft. Therefore, the entire process requires continuous iterative loops until convergence.

[0141] (1) Given the structural mass of the first-stage aircraft An initial guess;

[0142] (2) Based on the ratio of the structural mass of the first-stage aircraft to the total mass of all modes of propulsion of the first-stage aircraft. Calculate the fuel mass of the first-stage aircraft:

[0143] (19)

[0144] (3) Calculate the takeoff mass of the aircraft according to formula (18). ;

[0145] (4) Calculate the structural mass of the first-stage aircraft according to the first-stage aircraft structural mass estimation method, i.e., formula (17);

[0146] (5) Based on formula (18) and the structural mass of the first-stage aircraft in the previous iteration in step (4). Recalculate takeoff mass ;

[0147] (6) Calculate the relative error When the error is less than 0.5%, the iteration is considered to have converged, and the first-stage aircraft structural mass and first-stage aircraft fuel mass corresponding to the last takeoff mass are used as the final mass estimate.

[0148] If the error is greater than or equal to 0.5%, the structural mass of the first-stage spacecraft from the previous iteration will be adjusted. Return to step (2) and continue iterative calculation until convergence.

[0149] This invention also discloses a two-stage-to-orbit spacecraft mass estimation system for implementing the aforementioned two-stage-to-orbit spacecraft mass estimation method, comprising:

[0150] The mission module is used to set the mission profile and payload mass of a two-stage-to-orbit aerospace vehicle.

[0151] The second-stage spacecraft structural mass prediction module predicts the structural mass of the second-stage spacecraft based on the developed second-stage rocket structural mass and engine mass parameters.

[0152] The second-stage vehicle mass calculation module is used to calculate the second-stage vehicle mass ratio and the propellant mass of the second-stage vehicle based on the relationship between the vehicle acceleration process and propellant consumption.

[0153] The mass of the second-stage aircraft is calculated based on the propellant mass and structural mass of the second-stage aircraft.

[0154] The first-stage vehicle fuel mass estimation module is used to calculate the first-stage vehicle fuel mass based on the payload mass, second-stage vehicle mass, and engine operating parameters of different modes, according to the constructed first-stage vehicle fuel mass estimation model.

[0155] The first-stage aircraft structural mass calculation module is used to calculate the structural mass of the first-stage aircraft according to the first-stage aircraft structural mass estimation method.

[0156] Takeoff mass calculation module, used to calculate takeoff mass;

[0157] And recalculate the takeoff mass based on the first-stage aircraft structural mass obtained from the first-stage aircraft structural mass calculation module;

[0158] The error calculation module is used to calculate the error between the two takeoff masses obtained by the takeoff mass calculation module.

[0159] The judgment module is used to determine whether the error requirement is met. If the error requirement is met, the iteration is considered to have converged. The final mass estimate is then taken as the first-stage aircraft structural mass and first-stage aircraft fuel mass corresponding to the last takeoff mass.

[0160] If the error does not meet the requirements, the structural mass of the first-stage aircraft from the previous iteration is input into the first-stage aircraft fuel mass estimation module, the takeoff mass calculation module, and the error calculation module for repeated calculation until the error requirements are met.

[0161] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention and not to limit it. Although the present invention has been described in detail with reference to the above embodiments, those skilled in the art should understand that modifications or equivalent substitutions can still be made to the specific implementation of the present invention. Any modifications or equivalent substitutions that do not depart from the spirit and scope of the present invention should be covered within the scope of protection of the claims of the present invention.

Claims

1. A method for estimating the mass of a two-stage-to-orbit spacecraft, characterized by comprising the following steps: S1. Determine the mission profile and payload mass of the two-stage-to-orbit aerospace vehicle; S2. Based on the developed structural mass and engine mass parameters of the second-stage rocket, estimate the structural mass of the second-stage spacecraft of the two-stage-to-orbit aerospace vehicle; S3. Based on the mission profile determined in S1, calculate the mass ratio of the second-stage vehicle according to the relationship between the vehicle acceleration process and propellant consumption, and then calculate the propellant mass of the second-stage vehicle. Calculate the mass of the second-stage aircraft based on the propellant mass of the second-stage aircraft and the structural mass of the second-stage aircraft obtained from S2. S4. Give an initial guess to the structural mass of the first-stage aircraft. Based on the load mass of S1, the mass of the second-stage aircraft in S3, and the operating parameters of the engines in different modes, calculate the fuel mass of the first-stage aircraft according to the constructed fuel mass estimation model of the first-stage aircraft. S5. Calculate the takeoff mass of the first-stage aircraft; S6. Calculate the structural mass of the first-stage aircraft according to the method for estimating the structural mass of the first-stage aircraft. S7. Recalculate the takeoff mass based on the structural mass of the first-stage aircraft obtained in S6; S8. Calculate the error of the takeoff mass obtained from S7 and S5. If the error requirement is met, the iteration is considered to have converged. Then, the first-stage aircraft structural mass and first-stage aircraft fuel mass corresponding to the takeoff mass obtained from S7 are used as the final mass estimate. If the error does not meet the requirements, return the structural mass of the first-stage aircraft from the previous iteration to S4, and repeat S4-S7 until the error requirements are met. In S3, the mass ratio of the second-stage aircraft is calculated based on the relationship between the constructed aircraft acceleration process and propellant consumption. The expression is: The relationship between the spacecraft's acceleration process and propellant consumption is defined during the construction phase, with the velocity and altitude of each flight node in the mission profile recorded as follows: ,in Indicates takeoff status. Define the required separation point height and velocity for the mission. arrive The flight process is the first Modality, defined as the first The flight mass at the node is ; but Indicates that the second-stage aircraft is in the... The velocity at the node, Indicates that the second-stage aircraft is in the... The velocity at the node, Indicates that the second-stage aircraft is in the... Height at the node Indicates that the second-stage aircraft is in the... Height at the node; It is the acceleration due to gravity; Specific impulse of the rocket engine; The propellant mass of the second-stage aircraft is calculated based on the mass ratio of the second-stage aircraft, using the following expression: in, For the propellant mass of the second-stage aircraft, To estimate the structural mass of the second-stage aircraft, For load mass; In S4, the expression for the first-stage aircraft fuel mass estimation model is as follows: in, For the fuel mass of a first-class aircraft, For the structural mass of a first-class aircraft, For load mass, For a second-level aircraft mass, The total mass ratio of all modes for propulsion of a first-stage aircraft.

2. The method for estimating the mass of a two-stage-to-orbit spacecraft according to claim 1, characterized in that, in S3, the mass of the second-stage spacecraft is calculated based on the propellant mass of the second-stage spacecraft and the structural mass of the second-stage spacecraft obtained in S2; specifically: 。 3. The method for estimating the mass of a two-stage-to-orbit spacecraft according to claim 1, characterized in that the expression for calculating the total mass ratio of all modes of the first-stage spacecraft propulsion is: in, The mass ratio of different modes of a first-stage aircraft, where i represents the mass ratio of the first-stage aircraft. Modality, where s represents the final node.

4. The method for estimating the mass of a two-stage-to-orbit aerospace vehicle according to claim 1, characterized in that, in S5, the expression for calculating the takeoff mass is: in, For takeoff mass.

5. The method for estimating the mass of a two-stage-to-orbit spacecraft according to claim 1, characterized in that, in S6, the structural mass of the first-stage spacecraft is calculated according to the method for estimating the structural mass of the first-stage spacecraft, specifically as follows: The formula for calculating the structural mass of a first-stage aircraft is: in, For the sake of fuselage quality, For wing mass, For the tail mass, For landing quality, For the quality of other systems.

6. The method for estimating the mass of a two-stage-to-orbit spacecraft according to claim 5, characterized in that the fuselage mass is: Wing mass: Tail mass: Landing quality: in, For aspect ratio, For fuselage height, For fuselage length, Mach number, For overload limit, For the rudder area, The total area of ​​the control surfaces. The vertical tail area, The area of ​​the trapezoidal wing. For the relative thickness of the airfoil, For the width of the fuselage, For takeoff mass, For wingtip to tip ratio, The MAC wing sweep angle is 25%. The MAC tail sweep angle is 25%. For wingtip to tip ratio, For takeoff mass.

7. The method for estimating the mass of a two-stage-to-orbit spacecraft according to claim 1, characterized in that, in S8, the formula for calculating the error is: in, The takeoff mass calculated in S5. This is the takeoff mass recalculated in S7.

8. A two-stage-to-orbit spacecraft mass estimation system that implements the two-stage-to-orbit spacecraft mass estimation method according to any one of claims 1-7, characterized in that it comprises: The mission module is used to set the mission profile and payload mass of a two-stage-to-orbit aerospace vehicle. The second-stage spacecraft structural mass prediction module predicts the structural mass of the second-stage spacecraft based on the developed second-stage rocket structural mass and engine mass parameters. The second-stage vehicle mass calculation module is used to calculate the second-stage vehicle mass ratio and the propellant mass of the second-stage vehicle based on the relationship between the vehicle acceleration process and propellant consumption. The mass of the second-stage aircraft is calculated based on the propellant mass and structural mass of the second-stage aircraft. The first-stage vehicle fuel mass estimation module is used to calculate the first-stage vehicle fuel mass based on the payload mass, second-stage vehicle mass, and engine operating parameters of different modes, according to the constructed first-stage vehicle fuel mass estimation model. The first-stage aircraft structural mass calculation module is used to calculate the structural mass of the first-stage aircraft according to the first-stage aircraft structural mass estimation method. Takeoff mass calculation module, used to calculate takeoff mass; The takeoff mass is recalculated based on the structural mass of the first-stage aircraft obtained from the first-stage aircraft structural mass calculation module. The error calculation module is used to calculate the error between the two takeoff masses obtained by the takeoff mass calculation module. The judgment module is used to determine whether the error requirement is met. If the error requirement is met, the iteration is considered to have converged. The final mass estimate is then taken as the first-stage aircraft structural mass and first-stage aircraft fuel mass corresponding to the last takeoff mass. If the error does not meet the requirements, the structural mass of the first-stage aircraft from the previous iteration is input into the first-stage aircraft fuel mass estimation module, the takeoff mass calculation module, and the error calculation module for repeated calculation until the error requirements are met.

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