Two-stage injection aerospace vehicle mass estimation method and system

Abstract

The invention belongs to the field of aircraft design, and relates to a two-stage injection aerospace vehicle mass estimation method and system. The method comprises the following steps: firstly, determining a mission profile and a load, then calculating the mass of a secondary aircraft, then calculating the fuel mass of a primary aircraft and the take-off mass of the primary aircraft, then calculating the structural mass of the primary aircraft, and recalculating the take-off mass according to the structural mass of the primary aircraft; error calculation is carried out on the takeoff masses calculated twice, if the error requirement is met, iteration convergence is considered, and the first-level aircraft structure mass and the first-level aircraft fuel mass corresponding to the takeoff masses obtained finally serve as final mass estimation values; and if the error does not meet the requirement, returning the primary aircraft structure mass of the last iteration to recalculate the take-off mass until the error requirement is met. And the error range is continuously reduced by utilizing iterative calculation until the precision requirement is met, the actual aircraft quality is fitted to the maximum extent, and a reliable data basis is provided for subsequent processes such as aircraft design.

Description

Technical Field

[0001] The present invention belongs to the field of aircraft design, and in particular relates to a method and system for estimating the mass of a two-stage orbital aerospace vehicle. Background Art

[0002] Horizontal take-off and landing near-space hypersonic weapons and low-cost reusable space shuttles are strategic weapons for achieving space superiority in the future. Countries around the world are focusing on developing horizontal take-off and landing reusable combined-power space shuttles. Aircraft mass estimation is the first step in aircraft design. Through mass estimation, the fuel mass required for the aircraft, the aircraft structure mass and approximate size can be calculated. Currently, there are many studies on aircraft mass estimation methods, but there are few studies on the scale assessment methods of the complete concept of a two-stage orbital space shuttle.

[0003] After searching, among the existing patents, the invention patent with publication number CN 115758692 A discloses a two-stage orbital aerospace vehicle mass estimation method based on statistical analysis. The relationship between the structural mass and take-off mass of the first and second stage aircraft is fitted by existing data, and the functional relationship between the fuel mass and the total take-off mass of the first and second stage aircraft is established according to the aircraft motion equation. The aircraft mass is solved by simultaneous equations. The quality assessment of this invention does not take into account many factors, including the influence of the structural size of the first stage aircraft on the structural mass, the mass of the subsystem, and the change of engine performance parameters when the first stage aircraft adopts combined power mode conversion. Secondly, the relationship between the structural mass and the take-off mass calculated by the fitting method will lead to large errors. Summary of the invention

[0004] The purpose of the present invention is to provide a two-stage orbital aerospace vehicle mass estimation method and system, which solves the problems in the current technology such as insufficient quality assessment accuracy, insufficient consideration of many key factors, and changes in engine parameters caused by the mode conversion of the first-stage aircraft.

[0005] The present invention is achieved through the following technical solutions: The present invention discloses a method for estimating the mass of a two-stage orbital aerospace vehicle, comprising the following steps: S1. Determine the mission profile and payload mass of the two-stage orbital aerospace vehicle; S2. Estimate the mass of the second stage structure of the two-stage orbital aerospace vehicle based on the developed second stage rocket structure mass and engine mass parameters; S3, based on the mission profile determined in S1, the secondary vehicle mass ratio is calculated according to the relationship between the vehicle acceleration process and the propellant consumption, thereby calculating the secondary vehicle propellant mass; Calculate the mass of the secondary vehicle based on the mass of the secondary vehicle propellant and the mass of the secondary vehicle structure obtained by S2; S4, give an initial guess value for the mass of the first-stage aircraft structure, and calculate the first-stage aircraft fuel mass according to the constructed first-stage aircraft fuel mass estimation model based on the load mass of S1, the second-stage aircraft mass of S3 and the operating parameters of the engine in different modes; S5. Calculate the take-off mass of the first-stage aircraft; S6. Calculate the mass of the first-level aircraft structure according to the first-level aircraft structure mass estimation method; S7, recalculating the take-off mass based on the primary aircraft structure mass obtained in S6; S8, performing error calculation on the takeoff masses obtained in S7 and S5. If the error requirement is met, the iteration is considered to have converged. The first-stage aircraft structure mass and the first-stage aircraft fuel mass corresponding to the takeoff mass obtained in S7 are used as the final mass estimation value; If the error does not meet the requirement, the first-level aircraft structure mass of the previous iteration is returned to S4, and S4-S7 are repeated until the error requirement is met.

[0006] Further, in S3, the secondary vehicle mass ratio is calculated based on the relationship between the constructed vehicle acceleration process and the propellant consumption, and the expression is:

[0007] When constructing the relationship between the vehicle acceleration process and the propellant consumption, the speed and altitude of each flight node of the mission profile are recorded as ,in Indicates take-off status. For the separation point height and speed required by the mission, define arrive The flight process is Modal, define The flight mass at the node is ; but Indicates that the secondary aircraft is in the The speed at the node, Indicates that the secondary aircraft is in the The speed at the node, Indicates that the secondary aircraft is in the The height of the node, Indicates that the secondary aircraft is in the Height at the node; is the acceleration due to gravity; is the specific impulse of the rocket engine; The mass of the secondary vehicle propellant is calculated based on the secondary vehicle mass ratio, and the expression is:

[0008] in, is the mass of the second stage vehicle propellant, is the estimated secondary vehicle structure mass, is the load mass.

[0009] Further, in S3, the mass of the secondary vehicle is calculated according to the mass of the secondary vehicle propellant and the mass of the secondary vehicle structure obtained in S2; specifically:

[0010] is the secondary vehicle mass.

[0011] Further, in S4, the first-stage aircraft fuel mass estimation model expression is constructed as follows:

[0012] in, is the mass of primary aircraft fuel, is the mass of the first-stage aircraft structure, is the load mass, is the secondary vehicle mass, is the total mass ratio of all modes of the first-stage vehicle propulsion.

[0013] Furthermore, the calculation expression of the total mass ratio of all modes of the first-stage aircraft propulsion is:

[0014] in, is the mass ratio of different modes of the first-stage aircraft, i represents the modality, s represents the final node.

[0015] Furthermore, in S5, the calculation expression of take-off mass is:

[0016] in, is the mass of the first-stage aircraft structure, is the load mass, is the secondary vehicle mass, is the mass of primary aircraft fuel, For take-off mass.

[0017] Further, in S6, the first-level aircraft structure mass is calculated according to the first-level aircraft structure mass estimation method, specifically: The calculation formula for the mass of the first-stage aircraft structure is:

[0018] in, is the body mass, is the wing mass, is the vertical tail mass, is the landing mass, For other system qualities.

[0019] Further, the quality of the fuselage:

[0020] Wing mass:

[0021] Vertical tail mass:

[0022] Landing Quality:

[0023] in, is the aspect ratio, is the fuselage height, is the fuselage length, is the Mach number, is the overload limit, is the rudder area, is the total control surface area, is the vertical tail area, is the area of ​​the trapezoidal wing, is the relative thickness of the airfoil, is the fuselage width, is the take-off mass, The ratio of wing tip to wing tip. The wing sweep angle is 25% MAC. The vertical tail sweep angle is 25% MAC. The ratio of wing tip to wing tip. For take-off mass.

[0024] Further, in S8, the error calculation formula is:

[0025] in, is the take-off mass calculated in S5, This is the takeoff mass recalculated in S7.

[0026] The present invention also discloses a two-stage orbital aerospace vehicle mass estimation system for realizing the two-stage orbital aerospace vehicle mass estimation method, comprising: Mission module, used to set the mission profile and payload mass of the two-stage orbital aerospace vehicle; The second-stage vehicle structure mass estimation module estimates the second-stage vehicle structure mass of a two-stage orbital aerospace vehicle based on the developed second-stage rocket structure mass and engine mass parameters; A secondary vehicle mass calculation module is used to calculate the secondary vehicle mass ratio and the secondary vehicle propellant mass according to the relationship between the constructed vehicle acceleration process and the propellant consumption; Calculate the mass of the secondary vehicle based on the mass of the secondary vehicle propellant and the mass of the secondary vehicle structure; A first-stage aircraft fuel mass estimation module, used for calculating the first-stage aircraft fuel mass according to a constructed first-stage aircraft fuel mass estimation model based on the load mass, the second-stage aircraft mass and different mode engine operating parameters; A first-level aircraft structure mass calculation module, used to calculate the first-level aircraft structure mass according to a first-level aircraft structure mass estimation method; Take-off mass calculation module, used to calculate take-off mass; and recalculating the take-off mass according to the first-level aircraft structure mass obtained by the first-level aircraft structure mass calculation module; An error calculation module, used for performing error calculation on the two take-off masses obtained by the take-off mass calculation module; A judgment module is used to judge whether the error requirement is met. If the error meets the requirement, it is considered that the iteration has converged, and the first-stage aircraft structure mass and the first-stage aircraft fuel mass corresponding to the last obtained takeoff mass are used as the final mass estimation value; If the error does not meet the requirement, the first-stage aircraft structure mass of the previous iteration is input into the first-stage aircraft fuel mass estimation module, and the takeoff mass calculation module and the error calculation module repeat the calculation until the error requirement is met.

[0027] Compared with the prior art, the present invention has the following beneficial technical effects: The present invention discloses a method for estimating the mass of a two-stage spacecraft entering orbit. Through multi-step precise calculations, it comprehensively considers many key factors such as mission profile, various levels of aircraft structure, engine operating parameters, etc. It is not a simple empirical estimation. Each link is verified mutually, so that the final mass estimation value is more accurate. It uses iterative calculations to continuously narrow the error range until the accuracy requirements are met, and the actual aircraft mass is matched to the greatest extent, providing a reliable data basis for subsequent processes such as aircraft design and trajectory calculation.

[0028] Based on the existing two-stage rocket parameters, the estimation work can be carried out based on the data of mature technical systems, which can not only adapt to the existing technical framework, but also flexibly respond to the special design requirements of new two-stage orbital aerospace vehicles. Moreover, the constructed model can be flexibly adjusted according to different mission profiles, engine working modes and other variables, and can be used for various types of two-stage orbital missions.

[0029] The relationship between the vehicle acceleration process and propellant consumption, the structural quality of each level of the vehicle, engine parameters, payload mass and a series of other factors are incorporated into a unified calculation system, breaking the limitations of the individual evaluation of each subsystem and forming a comprehensive and systematic quality assessment process. It ensures that the coordination and cooperation between the various parts is reflected in the quality estimation link, which helps to create an overall aircraft architecture with balanced performance.

[0030] Furthermore, when using the total mass ratio of all modes of propulsion of the first-stage vehicle to estimate the fuel mass of the first-stage vehicle, the performance changes of different engine modes of the first-stage vehicle are comprehensively considered, which is in line with the actual flight conditions of the two-stage orbital aerospace vehicle and can improve the accuracy of mass estimation. At the same time, this calculation method can be applied to a variety of other combined power forms.

[0031] Furthermore, when calculating the mass of the first-stage spacecraft, the scale assessment of the complete concepts such as the structural mass of the two-stage orbital aerospace vehicle, the mass of the subsystems, and the mass of the fuel is taken into consideration, so that the calculated structural mass of the first-stage spacecraft is more in line with the actual situation. BRIEF DESCRIPTION OF THE DRAWINGS

[0032] Figure 1 It is an overall flow chart of the mass estimation method of a two-stage orbital aerospace vehicle provided by an embodiment of the present disclosure; Figure 2 It is the mission profile of the secondary vehicle; Figure 3 This is the mission profile of the first-stage aircraft. DETAILED DESCRIPTION

[0033] In order to make the purpose, technical solution and advantages of the present invention more clear, the following is further described in detail with reference to the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are only used to explain the present invention and are not used to limit the present invention, that is, the embodiments described are only part of the embodiments of the present invention, not all embodiments.

[0034] The detailed description of the embodiments of the present invention provided in the following drawings is not intended to limit the scope of the invention claimed for protection, but merely represents a selected embodiment of the present invention. Based on the drawings and embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative work are within the scope of protection of the present invention.

[0035] The features and performances of the present invention are further described in detail below in conjunction with the embodiments.

[0036] like Figure 1 As shown, the present invention discloses a method for estimating the mass of a two-stage orbital aerospace vehicle, which mainly includes the following contents: First, determine the payload mass and flight profile of the mission, start from the second-stage vehicle mass estimate, estimate the second-stage vehicle structure mass based on experience, determine the mass ratio based on the rocket engine performance and the mission profile, and thus calculate the second-stage vehicle mass.

[0037] Then, the secondary aircraft and the payload are taken together as the load of the first-stage aircraft. The mass ratio of the first-stage aircraft is calculated according to the modal transition point of the engine performance. The size of the first-stage aircraft is estimated and the structural mass of the first-stage aircraft is calculated. Thus, the mass and take-off mass of the first-stage aircraft are calculated and converged after iterative cycles.

[0038] like Figure 2 and Figure 3 As shown, based on the overall parameters of the developed multi-type reusable space-to-space round-trip transportation system and the performance boundary of the rocket engine and combined power, the mission profile of the two-stage orbital space vehicle is determined: The vehicle took off horizontally from the ground, using 6 pre-cooling air turbo rocket engines (PATR); at 18km and 3.5Ma, the PATR engine was shut down, the PATR flow channel diverter was closed, and the sub-ramjet mode was started; at 24km and 4.0Ma, the sub-ramjet engine was shut down and the scramjet mode was started. The first-stage vehicle carried the second-stage vehicle to 31km and 6.8Ma, and the first and second stages separated, and the first-stage vehicle returned without power. The second-stage vehicle was propelled by 2 hydrogen-oxygen rocket engines and eventually accelerated to reach low-Earth orbit.

[0039] When designing, you need to first construct the general equation of aircraft acceleration motion, specifically: The speed and altitude of each flight node of the mission profile are recorded as ,in Indicates take-off status. For the separation point height and speed required by the mission, define arrive The flight process is Modal, define The flight mass at the node is , quality ratio , speed increment , the height increment is , and subsequently establish a specific quality estimation method.

[0040] The general equation for vehicle acceleration and propellant consumption rate is: (1) Where: is the mass of the aircraft, is the aircraft speed, is the engine thrust, is the aircraft drag, is the acceleration due to gravity, is the climbing angle.

[0041] Define the effective specific impulse as , is the engine specific impulse.

[0042] The general equation of vehicle acceleration and propellant consumption rate becomes: (2) Compared with the engine specific impulse, the effective specific impulse considering the flight resistance can better reflect the acceleration performance of the aircraft. The larger it is, the closer the effective specific impulse is to the engine specific impulse.

[0043] The specific process of determining the quality of the secondary aircraft structure is as follows: Determine the load mass , based on the second-stage rocket structure mass and rocket engine mass of multiple rockets that have been developed, the second-stage rocket structure mass of the two-stage spacecraft is estimated .

[0044] The calculation process of the secondary vehicle mass ratio and total mass is as follows: Load mass and the estimated secondary vehicle structure mass Based on the analysis of the relationship between the vehicle acceleration process and the propellant consumption, the secondary vehicle mass ratio is calculated, thereby calculating the secondary vehicle propellant mass. The details are as follows: Since the thrust-to-drag ratio of the vehicle is large in rocket mode, it can be considered that In addition, the rocket engine It is approximately constant during the flight, so integrating both ends of equation (2) yields: (3) The second term in the above formula is the gravity loss. , whose value is relative to the velocity increment It is usually small and cannot be accurately obtained in the conceptual design stage, so it is ignored in the initial estimate and the As the mean velocity of the rocket mode, it is used to estimate , then: (4) Combining formula (3) and formula (4), we get the mass ratio of the secondary aircraft: (5) The secondary vehicle mass ratio can also be expressed as the ratio of the total vehicle mass to the non-propellant mass, which can be obtained as: (6) in, is the estimated second stage propellant mass, is the estimated secondary vehicle structure mass, is the load mass.

[0045] Therefore, by combining formula (5) and formula (6), the mass of the second-stage vehicle propellant can be calculated, and then the mass of the second-stage vehicle can be calculated. Specifically: .

[0046] Next, I will introduce the relevant content of the first-level aircraft.

[0047] First, construct the total mass ratio of all modes of the first-stage aircraft propulsion, specifically: In the air-breathing mode, the thrust-drag ratio of the aircraft is relatively small, and the drag loss cannot be ignored. Therefore, a different estimation method from the rocket mode in the previous article is adopted. Multiply both ends of equation (2) by V and divide by M to obtain: (7) In the formula, TV represents the engine power, and the engine thermal efficiency is , the calorific value is q, then (8) Substituting into equation (7), taking the thrust-to-drag ratio as a constant, the integral is: (9) It can finally be written in unified form: (10) In formula (10), gravity loss , the reaction vehicle acceleration performance is the average effective specific impulse. For the air-breathing propulsion mode, the average effective specific impulse is: (11) For the PATR engine, an accurate component-level model of the engine has been established, and simulation data of each operating point has been obtained to take the average specific impulse The average specific impulse of ramjet and scramjet engines is calculated based on combustion efficiency and fuel calorific value.

[0048] Calculate the total mass ratio of all modes of the first stage vehicle propulsion: (12) In formula (12), is the load mass, is the secondary vehicle mass, is the mass of primary aircraft fuel, is the mass of the first-stage aircraft structure, is the mass ratio of different engine modes of the first-stage aircraft; i represents the modality, s represents the final node.

[0049] According to formula (12), the first-stage aircraft fuel mass estimation model can be derived. The expression is: .

[0050] Design an algorithm for estimating the mass of a first-stage aircraft structure: Combining the developed parameters of various types of reusable space launch vehicles, a size-based structural mass formula for each component is formed.

[0051] Body quality: (13) Wing mass: (14) Vertical tail mass: (15) Landing Quality: (16) in, is the aspect ratio, is the fuselage height, is the fuselage length, is the Mach number, is the overload limit, is the rudder area, is the total control surface area, is the vertical tail area, is the area of ​​the trapezoidal wing, is the relative thickness of the airfoil, is the fuselage width, is the take-off mass, The ratio of wing tip to wing tip. The wing sweep angle is 25% MAC. The vertical tail sweep angle is 25% MAC. The ratio of wing tip to wing tip.

[0052] The mass of the first-stage aircraft structure is: (17) in, The quality of other systems, including thermal management, landing gear, load system, control system, engine, etc.

[0053] Take-off mass: (18) in, is the mass of the first-stage aircraft structure, is the load mass, is the secondary vehicle mass, It is the fuel mass of the first-level aircraft.

[0054] Through the previous calculations, the mass ratio and structural mass of the aircraft can be obtained. However, the structural mass calculation of the first-stage aircraft requires the takeoff mass, and the fuel mass of the first-stage aircraft needs to be calculated through the structural mass and mass ratio. The takeoff mass is the sum of the masses of the first and second-stage aircraft. Therefore, the whole process needs to be iterated and looped continuously until convergence.

[0055] (1) Given the mass of the first-stage aircraft structure An initial guess; (2) Based on the ratio of the first-stage vehicle structure mass to the total modal mass of the first-stage vehicle propulsion , calculate the first-stage aircraft fuel mass: (19) (3) Calculate the takeoff mass of the aircraft according to formula (18): ; (4) Calculate the mass of the first-stage aircraft structure according to the first-stage aircraft structure mass estimation method, i.e., formula (17); (5) According to formula (18) and the mass of the first-stage aircraft structure in the previous iteration in step (4): Recalculate takeoff mass ; (6) Calculate relative error When the error is less than 0.5%, the iteration is considered to have converged, and the first-stage aircraft structure mass and the first-stage aircraft fuel mass corresponding to the last takeoff mass obtained are used as the final mass estimation value; If the error is greater than or equal to 0.5%, the first-level aircraft structure quality of the previous iteration will be Return to step (2) and continue iterative calculation until convergence.

[0056] The present invention also discloses a two-stage orbital aerospace vehicle mass estimation system for realizing the two-stage orbital aerospace vehicle mass estimation method, comprising: Mission module, used to set the mission profile and payload mass of the two-stage orbital aerospace vehicle; The second-stage vehicle structure mass estimation module estimates the second-stage vehicle structure mass of a two-stage orbital aerospace vehicle based on the developed second-stage rocket structure mass and engine mass parameters; A secondary vehicle mass calculation module is used to calculate the secondary vehicle mass ratio and the secondary vehicle propellant mass according to the relationship between the constructed vehicle acceleration process and the propellant consumption; Calculate the mass of the secondary vehicle based on the mass of the secondary vehicle propellant and the mass of the secondary vehicle structure; A first-stage aircraft fuel mass estimation module, used for calculating the first-stage aircraft fuel mass according to a constructed first-stage aircraft fuel mass estimation model based on the load mass, the second-stage aircraft mass and different mode engine operating parameters; A first-level aircraft structure mass calculation module, used to calculate the first-level aircraft structure mass according to a first-level aircraft structure mass estimation method; Take-off mass calculation module, used to calculate take-off mass; and recalculating the take-off mass according to the first-level aircraft structure mass obtained by the first-level aircraft structure mass calculation module; An error calculation module, used for performing error calculation on the two take-off masses obtained by the take-off mass calculation module; A judgment module is used to judge whether the error requirement is met. If the error meets the requirement, it is considered that the iteration has converged, and the first-stage aircraft structure mass and the first-stage aircraft fuel mass corresponding to the last obtained takeoff mass are used as the final mass estimation value; If the error does not meet the requirement, the first-stage aircraft structure mass of the previous iteration is input into the first-stage aircraft fuel mass estimation module, and the takeoff mass calculation module and the error calculation module repeat the calculation until the error requirement is met.

[0057] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention rather than to limit it. Although the present invention has been described in detail with reference to the above embodiments, ordinary technicians in the relevant field should understand that the specific implementation methods of the present invention can still be modified or replaced by equivalents. Any modification or equivalent replacement that does 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 orbital aerospace vehicle, characterized in that it comprises the following steps: S1. Determine the mission profile and payload mass of the two-stage orbital aerospace vehicle; S2. Estimate the mass of the second stage structure of the two-stage orbital aerospace vehicle based on the developed second stage rocket structure mass and engine mass parameters; S3, based on the mission profile determined in S1, the secondary vehicle mass ratio is calculated according to the relationship between the vehicle acceleration process and the propellant consumption, thereby calculating the secondary vehicle propellant mass; Calculate the mass of the secondary vehicle based on the mass of the secondary vehicle propellant and the mass of the secondary vehicle structure obtained by S2; S4, give an initial guess value for the mass of the first-stage aircraft structure, and calculate the first-stage aircraft fuel mass according to the constructed first-stage aircraft fuel mass estimation model based on the load mass of S1, the second-stage aircraft mass of S3 and the operating parameters of the engine in different modes; S5. Calculate the take-off mass of the first-stage aircraft; S6. Calculate the mass of the first-level aircraft structure according to the first-level aircraft structure mass estimation method; S7, recalculating the take-off mass based on the primary aircraft structure mass obtained in S6; S8, performing error calculation on the takeoff masses obtained in S7 and S5. If the error requirement is met, the iteration is considered to have converged. The first-stage aircraft structure mass and the first-stage aircraft fuel mass corresponding to the takeoff mass obtained in S7 are used as the final mass estimation value; If the error does not meet the requirement, the first-level aircraft structure mass of the previous iteration is returned to S4, and S4-S7 are repeated until the error requirement is met.

2. A two-stage orbital aerospace vehicle mass estimation method according to claim 1, characterized in that, in S3, the mass ratio of the second-stage vehicle is calculated according to the relationship between the vehicle acceleration process and the propellant consumption, and the expression is: When constructing the relationship between the vehicle acceleration process and the propellant consumption, the speed and altitude of each flight node of the mission profile are recorded as ,in Indicates take-off status. For the separation point height and speed required by the mission, define arrive The flight process is Modal, define The flight mass at the node is ; but Indicates that the secondary aircraft is in the The speed at the node, Indicates that the secondary aircraft is in the The speed at the node, Indicates that the secondary aircraft is in the The height of the node, Indicates that the secondary aircraft is in the Height at the node; is the acceleration due to gravity; is the specific impulse of the rocket engine; The mass of the secondary vehicle propellant is calculated based on the secondary vehicle mass ratio, and the expression is: in, is the mass of the second stage vehicle propellant, is the estimated secondary vehicle structure mass, is the load mass.

3. The method for estimating the mass of a two-stage orbital aerospace vehicle according to claim 2 is characterized in that, in S3, the mass of the secondary vehicle is calculated according to the mass of the propellant of the secondary vehicle and the mass of the secondary vehicle structure obtained in S2; specifically: is the secondary vehicle mass.

4. The method for estimating the mass of a two-stage orbital aerospace vehicle according to claim 1, characterized in that, in S4, the first-stage vehicle fuel mass estimation model expression constructed is: in, is the mass of primary aircraft fuel, is the mass of the first-stage aircraft structure, is the load mass, is the secondary vehicle mass, is the total mass ratio of all modes of the first-stage vehicle propulsion.

5. A two-stage orbital aerospace vehicle mass estimation method according to claim 4, characterized in that the calculation expression of the total modal mass ratio of the first-stage vehicle propulsion is: in, is the mass ratio of different modes of the first-stage aircraft, i represents the modality, s represents the final node.

6. The method for estimating mass of a two-stage orbital aerospace vehicle according to claim 1, characterized in that, in S5, the calculation expression of the take-off mass is: in, is the mass of the first-stage aircraft structure, is the load mass, is the secondary vehicle mass, is the mass of primary aircraft fuel, For take-off mass.

7. The method for estimating the mass of a two-stage orbital aerospace vehicle according to claim 1, characterized in that, in S6, the mass of the first-stage vehicle structure is calculated according to the method for estimating the mass of the first-stage vehicle structure, specifically: The calculation formula for the mass of the first-stage aircraft structure is: in, is the body mass, is the wing mass, is the vertical tail mass, is the landing mass, For other system qualities.

8. The method for estimating the mass of a two-stage orbital aerospace vehicle according to claim 7, characterized in that the mass of the fuselage is: Wing mass: Vertical tail mass: Landing Quality: in, is the aspect ratio, is the fuselage height, is the fuselage length, is the Mach number, is the overload limit, is the rudder area, is the total control surface area, is the vertical tail area, is the area of ​​the trapezoidal wing, is the relative thickness of the airfoil, is the fuselage width, is the take-off mass, The ratio of wing tip to wing tip. The wing sweep angle is 25% MAC. The vertical tail sweep angle is 25% MAC. The ratio of wing tip to wing tip. For take-off mass.

9. The method for estimating mass of a two-stage orbital aerospace vehicle according to claim 1, characterized in that, in S8, the error calculation formula is: in, is the take-off mass calculated in S5, This is the takeoff mass recalculated in S7.

10. A two-stage orbital aerospace vehicle mass estimation system for implementing the two-stage orbital aerospace vehicle mass estimation method according to any one of claims 1 to 9, characterized in that it comprises: Mission module, used to set the mission profile and payload mass of the two-stage orbital aerospace vehicle; The second-stage vehicle structure mass estimation module estimates the second-stage vehicle structure mass of a two-stage orbital aerospace vehicle based on the developed second-stage rocket structure mass and engine mass parameters; A secondary vehicle mass calculation module is used to calculate the secondary vehicle mass ratio and the secondary vehicle propellant mass according to the relationship between the constructed vehicle acceleration process and the propellant consumption; Calculate the mass of the secondary vehicle based on the mass of the secondary vehicle propellant and the mass of the secondary vehicle structure; A first-stage aircraft fuel mass estimation module, used for calculating the first-stage aircraft fuel mass according to a constructed first-stage aircraft fuel mass estimation model based on the load mass, the second-stage aircraft mass and different mode engine operating parameters; A first-level aircraft structure mass calculation module, used to calculate the first-level aircraft structure mass according to a first-level aircraft structure mass estimation method; Take-off mass calculation module, used to calculate take-off mass; Recalculate the take-off mass according to the first-stage aircraft structure mass obtained by the first-stage aircraft structure mass calculation module; An error calculation module, used for performing error calculation on the two take-off masses obtained by the take-off mass calculation module; A judgment module is used to judge whether the error requirement is met. If the error meets the requirement, it is considered that the iteration has converged, and the first-stage aircraft structure mass and the first-stage aircraft fuel mass corresponding to the last obtained takeoff mass are used as the final mass estimation value; If the error does not meet the requirement, the first-stage aircraft structure mass of the previous iteration is input into the first-stage aircraft fuel mass estimation module, and the takeoff mass calculation module and the error calculation module repeat the calculation until the error requirement is met.

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