Method for estimating initial mass ratio of two-stage injection aerospace vehicle

Abstract

The invention belongs to the field of aerospace, and relates to a method for estimating an initial mass ratio of a two-stage orbit-injection aerospace vehicle, which comprises the following steps of: preliminarily selecting an energy distribution ratio alpha according to a first-stage initial mass ratio and a second-stage initial mass ratio of the two-stage orbit-injection aerospace vehicle; calculating the initial mass ratio gamma of the two-stage injection aerospace vehicle through the energy distribution ratio alpha; adjusting different energy distribution ratios alpha, and carrying out trade-off optimization design to obtain an optimal interval of the energy distribution ratio alpha, the air weight coefficient pie1 of the first-stage aerospace vehicle and the air weight coefficient pie2 of the second-stage aerospace vehicle for subsequent detailed design analysis; and the initial mass ratio of the takeoff weight to the load weight of the two-stage injection aerospace vehicle can be quickly calculated. And design, research and development personnel can better know and master the overall performance of the aircraft, so that optimization design is carried out. Meanwhile, the method is also suitable for estimating the initial mass ratio of the takeoff weight to the load weight of other types of aerospace vehicles.

Description

Technical Field

[0001] The present application belongs to the field of aerospace, and in particular to a method for estimating the initial mass ratio of a two-stage orbital aerospace vehicle. Background Art

[0002] With the continuous development of aerospace technology, aerospace vehicles, as a means of transportation that can achieve round-trip between the earth and the sky and enter the space orbit, have attracted more and more attention. In the design of a two-stage orbital aerospace vehicle, the ratio of the takeoff weight of the aerospace vehicle to the load weight (hereinafter referred to as the "initial mass ratio") is an important parameter. The smaller the parameter, the stronger the carrying capacity of the aerospace vehicle and the lower the carrying cost. It directly affects the performance and carrying capacity of the aircraft. The initial mass ratio of the aerospace vehicle is an important indicator to measure its performance, which directly affects the carrying capacity, launch cost and safety of the aircraft. Therefore, how to accurately estimate the initial mass ratio of the two-stage orbital aerospace vehicle is of great significance. At present, the estimation method of the initial mass ratio of the aerospace vehicle mainly relies on empirical formulas and simulation experiments. In the design of aerospace vehicles, the initial mass ratio is an important parameter for evaluating the performance of the aircraft. However, the existing estimation methods are often complicated and require a lot of calculations and data support, which brings inconvenience to the design work.

[0003] Therefore, it is imperative to develop a simple and accurate method for estimating the initial mass ratio. Summary of the invention

[0004] The purpose of this application is to provide a method for estimating the initial mass ratio of a two-stage orbital aerospace launch vehicle to solve the problem that the existing initial mass ratio estimation methods are relatively complex and inconvenient.

[0005] The technical solution of the present application is: a method for estimating the initial mass ratio of a two-stage orbital space launch vehicle, comprising:

[0006] Conduct force analysis and force equation analysis of the two-stage orbital space launch vehicle, and simplify the results of force analysis and force equation analysis in combination with the flight status;

[0007] Obtain the fuel and flight-launch integration design data of the first-stage carrier and the flight data of the two-stage space launch vehicle, calculate the first-stage weight consumption of the two-stage space launch vehicle; then combine the empty weight coefficient Π of the first-stage carrier e1 Calculate the initial mass ratio Γ1 of the first stage of the two-stage orbital space launch vehicle;

[0008] Obtain the fuel of the second-stage carrier, the performance data of the selected rocket power and the flight-launch integration design data, calculate the second-stage weight consumption of the second-stage carrier, and then combine the empty weight coefficient Π of the second-stage carrier e2Calculate the initial mass ratio Γ2 of the second stage of the two-stage orbital space launch vehicle;

[0009] The energy distribution ratio α is preliminarily selected according to the initial mass ratio of the first stage and the initial mass ratio of the second stage of the two-stage space vehicle for orbital entry; α represents a portion of the total orbital energy provided by the first stage vehicle, and then the initial mass ratio Γ of the two-stage space vehicle for orbital entry is calculated by the energy distribution ratio α;

[0010] Adjust different energy distribution ratios α, make trade-offs and find the optimal design, and obtain the energy distribution ratio α, the empty weight coefficient Π of the first-stage aerospace vehicle e1 and the empty weight coefficient of the second-stage aerospace vehicle Π e2 The optimal interval is provided for subsequent detailed design analysis.

[0011] Preferably, the specific method for performing trade-off optimization design is: obtain the current design standard range of the mass ratio of the two-stage orbital aerospace vehicle, and set the allocation interval 0-1 of the step amount and the energy distribution ratio α, adjust the energy distribution ratio α in turn according to the step amount, repeat the design of the initial mass ratio of the two-stage orbital aerospace vehicle, and compare the newly obtained minimum initial mass ratio of the two-stage orbital aerospace vehicle with the design standard range of the mass ratio of the two-stage orbital aerospace vehicle to determine whether it meets the design standard range requirements of the mass ratio of the two-stage orbital aerospace vehicle. If so, the corresponding parameters are included in the preferred interval; if not, the corresponding parameters are not included in the preferred interval; after completing all allocations in the allocation interval 0-1, the energy distribution ratio α and the empty weight coefficient Π of the first-stage aerospace vehicle are obtained. e1 、The empty weight coefficient of the second-stage aerospace vehicle Π e2 The optimal intervals of the main influencing parameters are as follows.

[0012] Preferably, the specific calculation method of the first-stage initial mass ratio Γ1 is:

[0013] Determine the fuel type used by the air-breathing propulsion of the first stage of the two-stage orbital space launch vehicle and determine the fuel calorific value h PR ;

[0014] According to the flight and launch integration design data of the two-stage space launch vehicle, the first-stage launch vehicle The value of the first stage carrier is the efficiency of the first-stage space launch vehicle / air-breathing power integrated design, where D is the drag and D e is the installed or external drag, F is the non-installed engine thrust of the first-stage aerospace vehicle;

[0015] The effective total efficiency of the first-stage carrier's air-breathing power is determined based on the performance data of the engine unit providing air-breathing power and the flight-engine integration design of the first-stage carrier. η0 is the total efficiency of the first stage vehicle air-breathing power;

[0016] The orbit ratio was preliminarily determined based on the flight data of the two-stage space launch vehicle.

[0017] Calculate the first stage weight consumption of the two-stage orbital aerospace vehicle based on the effective total efficiency of the first stage air-breathing power and the orbit ratio;

[0018] The empty weight coefficient Π of the first stage of the two-stage orbital aerospace vehicle is preliminarily selected according to the design requirements. e1 ; Then, the initial mass ratio of the first stage of the two-stage orbital aerospace launch vehicle is calculated through the first-stage weight consumption and the empty weight coefficient of the first-stage launch vehicle.

[0019] Preferably, the calculation formula for the first-level weight consumption is:

[0020]

[0021] The calculation formula for the initial mass ratio of the first stage of the two-stage orbital space launch vehicle is:

[0022]

[0023] In the formula, m i1 =m p1 +m e1 +m f1 , Subscript 1 represents the first stage carrier, m represents the mass of the aircraft, in kg; subscripts i, e, f, and p represent the initial takeoff state, empty state, fuel, and load of the aircraft, respectively; Γ represents the initial mass ratio, g0 is the gravitational acceleration on the earth's surface; r0 is the average radius of the earth; r is the flight radius of the two-stage orbital aerospace launch vehicle.

[0024] Preferably, the specific calculation method of the second stage initial mass ratio Γ2 is:

[0025] According to the type of fuel used in the second stage rocket power of the two-stage space launch vehicle, the rocket power specific impulse I is determined. sp ;

[0026] According to the flight and launch integrated design data of the two-stage space launch vehicle, the second-stage launch vehicle The value of is the efficiency of the integrated design of the second-stage vehicle; where D is the drag, D e is the installed or external drag, F is the non-installed engine thrust of the second-stage aerospace vehicle;

[0027] The effective specific impulse of the second-stage carrier rocket power is determined based on the performance data of the selected rocket power and the flight-engine integration design of the second-stage carrier.

[0028] Calculate the second stage weight consumption of the two-stage orbital aerospace vehicle based on the effective specific impulse of the second stage vehicle rocket power;

[0029] The empty weight coefficient Π of the second stage of the two-stage orbital space launch vehicle is preliminarily selected according to the design requirements. e2 ;

[0030] The initial mass ratio of the second stage of the two-stage orbital aerospace launch vehicle is calculated based on the second stage weight consumption and the empty weight coefficient of the second stage launch vehicle.

[0031] Preferably, the formula for the second-level weight consumption is:

[0032]

[0033] The formula for the second stage initial mass ratio is:

[0034]

[0035] Subscript 2 indicates the second stage carrier, subscripts i, e, f, and p indicate the aircraft's initial takeoff state, empty state, fuel, and load, respectively; m i2 =m p2 +m e2 +m f2 , m p1 =m i2 ;

[0036] After conversion, we get:

[0037]

[0038] The initial mass ratio estimation method of the two-stage orbital aerospace vehicle of the present application can quickly calculate the takeoff weight and the initial mass ratio of the payload weight of the two-stage orbital aerospace vehicle. It helps designers and developers to better understand and master the overall performance of the aircraft, so as to optimize the design. At the same time, this method is also applicable to the estimation of the initial mass ratio of the takeoff weight and payload weight of other types of aerospace vehicles. It can effectively improve the efficiency and accuracy of aerospace vehicle design and development. BRIEF DESCRIPTION OF THE DRAWINGS

[0039] In order to more clearly illustrate the technical solution provided by the present application, the following is a brief introduction to the accompanying drawings. Obviously, the accompanying drawings described below are only some embodiments of the present application.

[0040] Figure 1 This is a schematic diagram of the overall process of this application;

[0041] Figure 2 This is a schematic diagram of the forces acting on the two-stage orbital space launch vehicle during flight;

[0042] Figure 3 This is a curve diagram of the corresponding relationship between the initial mass ratio and the energy distribution ratio of the two-stage orbital space launch vehicle in this application. DETAILED DESCRIPTION

[0043] The following will be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0044] A method for estimating the initial mass ratio of a two-stage orbital aerospace vehicle is based on the basic principles and parameters of the two-stage orbital aerospace vehicle design. Necessary simplifications and calculations are performed through the mathematical formula of the two-stage orbital aerospace vehicle's climbing flight phase, thereby improving the accuracy and efficiency of the estimation.

[0045] like Figure 1 As shown, the specific steps include:

[0046] Step S100, perform force analysis and force equation analysis on the two-stage orbital aerospace vehicle, and simplify the results of the force analysis and force equation analysis in combination with the flight status, such as ignoring the lift and flight track angle during the climbing process.

[0047] The aerospace vehicle consists of a first-stage vehicle, a second-stage vehicle and a payload. The first-stage vehicle uses air-breathing power, and the second-stage vehicle uses rocket power. At the separation point of the aerospace vehicle, the first-stage and second-stage vehicles separate, and the second-stage vehicle continues to move the payload into low-Earth orbit. The aerospace vehicle climbs at a dynamic speed and altitude. In this flight state, the speed and geometric altitude of the aircraft will change over time. According to the general force conditions of the aircraft, such as Figure 2 As shown (force equation).

[0048] Step S200, obtain the fuel of the first-stage carrier, the flight-engine integrated design data and the flight data of the two-stage space vehicle, and calculate the first-stage weight consumption of the two-stage space vehicle. Then, combine the empty weight coefficient Π of the first-stage carrier e1 Calculate the initial mass ratio Γ1 of the first stage of the two-stage orbital space launch vehicle.

[0049] The specific calculation method of the first-stage initial mass ratio Γ1 is:

[0050] Step S210, determine the fuel type used by the air-breathing power of the first stage of the two-stage space launch vehicle. Generally, hydrocarbon fuel or liquid hydrogen fuel is selected according to the mission characteristics and actual needs, and determine the fuel calorific value h PR , the unit is J / kg;

[0051] Step S220, according to the flight and launch integration design data of the two-stage space launch vehicle, that is, according to the preliminary design results or actual conditions, determine the first-stage launch vehicle The value of the first stage carrier. is the efficiency of the first-stage space launch vehicle / air-breathing power integrated design, where D is the drag and D e is the installed or external drag, and F is the non-installed engine thrust of the first-stage aerospace vehicle.

[0052] Step S230, determining the effective total efficiency of the first-stage carrier's air-breathing power according to the performance data of the air-breathing power provided by the engine unit and the flight-engine integration design of the first-stage carrier η0 is the total efficiency of the first stage vehicle air-breathing power;

[0053] Step S240: preliminarily determine the orbit ratio based on the flight data of the two-stage space launch vehicle.

[0054] Step S250, calculating the first-stage weight consumption of the two-stage space launch vehicle according to the effective total efficiency of the air-breathing power of the first-stage launch vehicle and the orbit ratio, the calculation formula is:

[0055]

[0056] Step S260: preliminarily select the empty weight coefficient Π of the first stage of the two-stage orbital space launch vehicle according to the design requirements. e1 Then, the initial mass ratio of the first stage of the two-stage orbital aerospace vehicle is calculated by the first stage weight consumption and the empty weight coefficient of the first stage vehicle. The specific formula is:

[0057]

[0058] In the formula, m i1 =m p1 +m e1 +m f1 , Subscript 1 represents the first stage carrier, m represents the mass of the aircraft, in kg; subscripts i, e, f, and p represent the initial takeoff state, empty state, fuel, and load of the aircraft, respectively; Γ represents the initial mass ratio, g0 is the gravitational acceleration on the earth's surface, in m / s2 ; r0 is the average radius of the earth, which is 6371km; r is the flight radius of the two-stage orbital aerospace launch vehicle.

[0059] According to the weight consumption estimation formula of climbing air-breathing powered aerospace vehicles, the following situations are used when the first stage uses air-breathing power:

[0060]

[0061] Step S300, obtain the fuel of the second-stage carrier, the performance data of the selected rocket power and the flight-launch integration design data, calculate the second-stage weight consumption of the second-stage carrier, and then combine the empty weight coefficient Π of the second-stage carrier e2 Calculate the initial mass ratio Γ2 of the second stage of the two-stage orbital space launch vehicle.

[0062] The specific calculation method is:

[0063] Step S310, determining the rocket power specific impulse I according to the fuel type used for the second stage rocket power of the two-stage space launch vehicle sp ;

[0064] Step S320, according to the flight-launch integration design data of the two-stage space launch vehicle, that is, according to the preliminary design results or actual conditions, determine the second-stage launch vehicle The value of is the efficiency of the integrated design of the second stage vehicle. Where D is the drag, D e is the installed or external drag, and F is the non-installed engine thrust of the second-stage aerospace vehicle.

[0065] Step S330: determine the effective specific impulse of the rocket power of the second-stage carrier according to the performance data of the selected rocket power and the flight-engine integration design of the second-stage carrier.

[0066] Step S340, calculating the second-stage weight consumption of the two-stage space launch vehicle according to the effective specific impulse of the second-stage launch vehicle rocket power, the specific formula is:

[0067]

[0068] Step S350: preliminarily select the empty weight coefficient Π of the second stage of the two-stage orbital space vehicle according to the design requirements. e2 ;

[0069] Step S360, calculating the second stage initial mass ratio of the two-stage space launch vehicle according to the second stage weight consumption and the empty weight coefficient of the second stage launch vehicle, the calculation formula is:

[0070]

[0071] Where, subscript 2 represents the second stage carrier, m i2 =m p2 +m e2 +m f2 , m p1 =m i2 ;

[0072] After conversion, we get:

[0073]

[0074] According to the weight consumption estimation formula of climbing rocket-powered aerospace vehicles, the following situations are used when the second stage uses rocket power:

[0075]

[0076] At the same time, due to:

[0077]

[0078] So we have:

[0079] Step S400, preliminarily selecting an energy distribution ratio α according to the initial mass ratio of the first stage and the initial mass ratio of the second stage of the two-stage space launch vehicle; the energy distribution ratio α is set empirically, and α represents a portion of the total orbital energy provided by the first stage vehicle. It is a part of the orbital energy provided by the second-stage carrier, and then the initial mass ratio Γ of the two-stage orbital aerospace carrier is calculated by the energy distribution ratio α.

[0080] The specific design is to convert the initial mass ratio of the first stage and the initial mass ratio of the second stage:

[0081]

[0082] make The initial mass ratio Γ of the two-stage orbital aerospace launch vehicle based on air-breathing power and rocket power is estimated as follows:

[0083]

[0084] Step S500, repeating steps S200-S400, adjusting different energy allocation ratios α, performing trade-off optimization design, and obtaining the energy allocation ratio α, the first-stage space vehicle empty weight coefficient Π e1 、The empty weight coefficient of the second-stage aerospace vehicle Π e2 The optimal intervals of the main influencing parameters are provided for subsequent detailed design analysis.

[0085] The specific method for performing trade-off optimization design is as follows: obtaining the current design standard range of the mass ratio of the two-stage orbital aerospace vehicle, and setting the allocation interval 0-1 of the step amount and the energy distribution ratio α, adjusting the energy distribution ratio α in turn according to the step amount, repeating the design of the initial mass ratio of the two-stage orbital aerospace vehicle, and comparing the newly obtained minimum initial mass ratio of the two-stage orbital aerospace vehicle with the design standard range of the mass ratio of the two-stage orbital aerospace vehicle to determine whether it meets the design standard range requirements of the mass ratio of the two-stage orbital aerospace vehicle. If so, the corresponding parameters are included in the preferred interval; if not, the corresponding parameters are not included in the preferred interval; after completing all allocations in the allocation interval 0-1, the energy distribution ratio α and the empty weight coefficient Π of the first-stage aerospace vehicle are obtained. e1 、The empty weight coefficient of the second-stage aerospace vehicle Π e2 The optimal intervals of the main influencing parameters are as follows.

[0086] The following is a specific example:

[0087] like Figure 3 As shown: The first-stage space vehicle uses liquid hydrogen fuel for air-breathing propulsion, and its fuel calorific value is h PR =120000 kJ / kg, effective total efficiency of the first-stage aerospace vehicle air-breathing power The effective specific impulse of the second-stage aerospace launch vehicle rocket This example Figure 2 It can be seen that when the energy distribution ratio α is 0.55, the initial mass ratio Γ of the two-stage orbital space launch vehicle can take the minimum value, that is, Λ1=0.7664, Λ2=2.0457, Π e1 =0.45, Π e2 =0.25 when Γ min =14.194; Λ1=0.7664, Λ2=2.0457, Π e1 =0.40,Π e2 =0.20 when Γ min =9.912. It can be seen that the minimum value of the initial mass ratio Γ of the two-stage space launch vehicle has a significant effect on the energy distribution ratio α and the empty weight coefficient Π of the first-stage space launch vehicle. e1 、The empty weight coefficient of the second-stage aerospace vehicle Π e2 All of them are very sensitive and require careful weighing and research to obtain and determine the performance and carrying capacity of a relatively excellent two-stage orbital aerospace launch vehicle.

[0088] Through the above design, the takeoff weight and initial mass ratio of the payload weight of the two-stage orbital aerospace vehicle can be quickly calculated. This helps designers and developers to better understand and master the overall performance of the aircraft, so as to optimize the design. At the same time, this method is also applicable to the estimation of the takeoff weight and initial mass ratio of the payload weight of other types of aerospace vehicles. It can effectively improve the efficiency and accuracy of aerospace vehicle design and development.

[0089] Finally, it should be noted that: the drawings of the embodiments disclosed in the present invention only involve structures related to the embodiments disclosed in the present invention, and other structures can refer to the general design. In the absence of conflict, the same embodiment and different embodiments of the present invention can be combined with each other;

[0090] Finally: The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc. made within the spirit and principles of the present invention should be included in the protection scope of the present invention.

Claims

1. A method for estimating the initial mass ratio of a two-stage space launch vehicle, characterized in that: include: Conduct force analysis and force equation analysis of the two-stage orbital space launch vehicle, and simplify the results of force analysis and force equation analysis in combination with the flight status; Obtain the fuel and flight-launch integration design data of the first-stage carrier and the flight data of the two-stage space launch vehicle, calculate the first-stage weight consumption of the two-stage space launch vehicle; then combine the empty weight coefficient Π of the first-stage carrier e1 Calculate the initial mass ratio Γ1 of the first stage of the two-stage orbital space launch vehicle; Obtain the fuel of the second-stage carrier, the performance data of the selected rocket power and the flight-launch integration design data, calculate the second-stage weight consumption of the second-stage carrier, and then combine the empty weight coefficient Π of the second-stage carrier e2 Calculate the initial mass ratio Γ2 of the second stage of the two-stage orbital space launch vehicle; The energy distribution ratio α is preliminarily selected according to the initial mass ratio of the first stage and the initial mass ratio of the second stage of the two-stage space vehicle for orbital entry; α represents a portion of the total orbital energy provided by the first stage vehicle, and then the initial mass ratio Γ of the two-stage space vehicle for orbital entry is calculated by the energy distribution ratio α; Adjust different energy distribution ratios α, make trade-offs and find the optimal design, and obtain the energy distribution ratio α, the empty weight coefficient Π of the first-stage aerospace vehicle e1 and the empty weight coefficient of the second-stage aerospace vehicle Π e2 The optimal interval is provided for subsequent detailed design analysis.

2. The method for estimating the initial mass ratio of a two-stage orbital space launch vehicle according to claim 1, characterized in that: The specific method for performing trade-off optimization design is as follows: obtaining the current design standard range of the mass ratio of the two-stage orbital aerospace vehicle, and setting the allocation interval 0-1 of the step amount and the energy distribution ratio α, adjusting the energy distribution ratio α in turn according to the step amount, repeating the design of the initial mass ratio of the two-stage orbital aerospace vehicle, and comparing the newly obtained minimum initial mass ratio of the two-stage orbital aerospace vehicle with the design standard range of the mass ratio of the two-stage orbital aerospace vehicle to determine whether it meets the design standard range requirements of the mass ratio of the two-stage orbital aerospace vehicle. If so, the corresponding parameters are included in the preferred interval; if not, the corresponding parameters are not included in the preferred interval; after completing all allocations in the allocation interval 0-1, the energy distribution ratio α and the empty weight coefficient Π of the first-stage aerospace vehicle are obtained. e1 、The empty weight coefficient of the second-stage aerospace vehicle Π e2 The optimal intervals of the main influencing parameters are as follows.

3. The method for estimating the initial mass ratio of a two-stage orbital space launch vehicle according to claim 1, characterized in that: The specific calculation method of the first-stage initial mass ratio Γ1 is: Determine the fuel type used by the air-breathing propulsion of the first stage of the two-stage orbital space launch vehicle and determine the fuel calorific value h PR ; According to the flight and launch integration design data of the two-stage space launch vehicle, the first-stage launch vehicle The value of the first stage carrier is the efficiency of the first-stage space launch vehicle / air-breathing power integrated design, where D is the drag and D e is the installed or external drag, F is the non-installed engine thrust of the first-stage aerospace vehicle; The effective total efficiency of the first-stage carrier's air-breathing power is determined based on the performance data of the engine unit providing air-breathing power and the flight-engine integration design of the first-stage carrier. η0 is the total efficiency of the first stage vehicle air-breathing power; The orbit ratio was preliminarily determined based on the flight data of the two-stage space launch vehicle. Calculate the first stage weight consumption of the two-stage orbital aerospace vehicle based on the effective total efficiency of the first stage air-breathing power and the orbit ratio; The empty weight coefficient Π of the first stage of the two-stage orbital aerospace vehicle is preliminarily selected according to the design requirements. e1 ; Then, the initial mass ratio of the first stage of the two-stage orbital aerospace launch vehicle is calculated through the first-stage weight consumption and the empty weight coefficient of the first-stage launch vehicle.

4. The method for estimating the initial mass ratio of a two-stage orbital space launch vehicle according to claim 3, characterized in that: The calculation formula for the first level weight consumption is: The calculation formula for the initial mass ratio of the first stage of the two-stage orbital space launch vehicle is: In the formula, m i1 =m p1 +m e1 +m f1 , Subscript 1 indicates the first stage carrier, m indicates the mass of the aircraft in kg; subscripts i, e, f, and p indicate the initial takeoff state, empty state, fuel, and load of the aircraft, respectively; Γ represents the initial mass ratio, g0 is the gravitational acceleration on the earth's surface; r0 is the average radius of the earth; r is the flight radius of the two-stage orbital aerospace launch vehicle.

5. The method for estimating the initial mass ratio of a two-stage orbital space launch vehicle according to claim 1, characterized in that: The specific calculation method of the second stage initial mass ratio Γ2 is: According to the type of fuel used in the second stage rocket power of the two-stage space launch vehicle, the rocket power specific impulse I is determined. sp ; According to the flight and launch integrated design data of the two-stage space launch vehicle, the second-stage launch vehicle The value of is the efficiency of the integrated design of the second-stage vehicle; where D is the drag, D e is the installed or external drag, F is the non-installed engine thrust of the second-stage aerospace vehicle; The effective specific impulse of the second-stage carrier rocket power is determined based on the performance data of the selected rocket power and the flight-engine integration design of the second-stage carrier. Calculate the second stage weight consumption of the two-stage orbital aerospace vehicle based on the effective specific impulse of the second stage vehicle rocket power; The empty weight coefficient Π of the second stage of the two-stage orbital space launch vehicle is preliminarily selected according to the design requirements. e2 ; The initial mass ratio of the second stage of the two-stage orbital aerospace launch vehicle is calculated based on the second stage weight consumption and the empty weight coefficient of the second stage launch vehicle.

6. The method for estimating the initial mass ratio of a two-stage orbital space launch vehicle according to claim 5, characterized in that: The formula for the second level weight consumption is: The formula for the second stage initial mass ratio is: Subscript 2 indicates the second stage carrier, subscripts i, e, f, and p indicate the aircraft's initial takeoff state, empty state, fuel, and load, respectively; m i2 =m p2 +m e2 +m f2 , m p1 =m i2 ; After conversion, we get:

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