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

Abstract

The invention belongs to the field of aerospace, and particularly relates to a method for estimating the initial mass ratio of a single-stage injection aerospace vehicle, and the method comprises the steps: preliminarily selecting an energy distribution ratio alpha according to the first-stage weight consumption and the second-stage weight consumption of the single-stage injection aerospace vehicle; the energy distribution ratio alpha is set through experience, alpha represents a part of total orbit energy provided by the first-stage vehicle, # imgabs0 # is a part of orbit energy provided by the second-stage vehicle, and then the initial mass ratio gamma of the single-stage injection aerospace vehicle is calculated through the energy distribution ratio alpha. And the initial mass ratio of the takeoff weight to the load weight of the single-stage injection aerospace vehicle can be quickly calculated. Therefore, design and 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 relates to the field of aerospace, and in particular to a method for estimating the initial mass ratio of a single-stage orbital aerospace vehicle. Background Art

[0002] With the continuous development of aerospace technology, space launch vehicles, as a means of transport capable of traveling between Earth and space and entering orbit, have attracted increasing attention. In the design of single-stage orbital launch vehicles, the ratio of the launch vehicle's takeoff weight to its payload weight (hereinafter referred to as the "initial mass ratio") is a critical parameter. A smaller initial mass ratio indicates a higher carrying capacity and lower launch costs, directly impacting the vehicle's performance and payload capacity. The initial mass ratio of a space launch vehicle is a key performance metric, directly impacting its payload capacity, launch cost, and safety. Therefore, accurately estimating the initial mass ratio of single-stage orbital launch vehicles using a relay of air-breathing and rocket propulsion is of great significance. Currently, methods for estimating the initial mass ratio of space launch vehicles primarily rely on empirical formulas and simulation experiments. In space launch vehicle design, the initial mass ratio is a crucial parameter for evaluating vehicle performance. However, existing estimation methods are often complex and require extensive computational and data support, which can be inconvenient for 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 single-stage orbital aerospace launch vehicle to solve the problem that the existing initial mass ratio estimation method is relatively complex and inconvenient.

[0005] The technical solution of this application is: a method for estimating the initial mass ratio of a single-stage orbital aerospace vehicle, comprising:

[0006] Conduct force analysis and force equation analysis of a single-stage orbital launch vehicle, and simplify the results of force analysis and force equation analysis based on flight conditions;

[0007] Obtain the fuel type of the first-stage carrier and the flight-launch integrated design data of the single-stage orbital aerospace carrier, and then calculate the first-stage weight consumption of the single-stage orbital aerospace carrier in combination with the orbit ratio, and then select the empty weight coefficient Π of the first-stage carrier according to the design requirements e1 ;

[0008] Obtain the fuel type of the second-stage vehicle and the flight-launch integration design data of the second-stage vehicle to determine the effective specific impulse of the rocket power Calculate the second stage weight consumption of the second stage vehicle;

[0009] The energy distribution ratio α is preliminarily selected based on the weight consumption of the first stage and the second stage of the single-stage orbital aerospace vehicle. The energy distribution ratio α is set empirically and represents the 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 vehicle, and then the initial mass ratio Γ of the single-stage space vehicle 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 first-stage aerospace vehicle empty weight coefficient π e1 and the empty weight coefficient π of the second-stage aerospace vehicle e2 The optimal interval is provided for subsequent detailed design analysis.

[0011] Preferably, the force equation of the single-stage orbital aerospace vehicle is:

[0012] m i1 =m p +m e +m f1 +m f2

[0013] m i2 =m p +m e +m f2 =m i -m f1

[0014] m i =m i1

[0015] m i2 =m i1 -m f1

[0016] ∏ f =∏ f1 +∏ f2 ,

[0017] Where m is the mass of the aircraft, in kg;

[0018] ∏ is the coefficient;

[0019] The subscripts 1 and 2 represent the first-stage carrier and the second-stage carrier, respectively;

[0020] The subscripts i, e, f, and p represent the aircraft's initial takeoff state, empty state, fuel, and load, respectively; Γ represents the initial mass ratio;

[0021] According to the force equation of the single-stage orbital aerospace vehicle, we can obtain:

[0022]

[0023] Preferably, the calculation method of the first-level weight consumption is:

[0024] Determine the fuel type used by the air-breathing propulsion system of the first stage of a single-stage orbital space launch vehicle, and determine the fuel calorific value h according to the fuel type. PR ;

[0025] Determine the first stage carrier based on the preliminary design results or actual conditions The value of the first stage carrier is the efficiency of the first stage 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;

[0026] The total effective 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's air-breathing power;

[0027] The orbit ratio was preliminarily determined based on the flight data of the single-stage space launch vehicle. The first-stage weight consumption of a single-stage space launch vehicle is calculated based on the effective total efficiency of the first-stage air-breathing propulsion and the orbit ratio. The first-stage weight consumption calculation formula is:

[0028]

[0029] Where g0 is the acceleration of gravity on the Earth's surface, in m / s 2 ; r0 is the average radius of the earth; η0 is the total efficiency of the air-breathing power at the first stage of power.

[0030] Preferably, the rocket power effective specific impulse The method to obtain is:

[0031] According to the type of fuel used by the rocket power during the second-stage power operation, the rocket power specific impulse I of the corresponding second-stage carrier is determined. sp ;

[0032] Find the flight-launch integration design data corresponding to the second-stage carrier of the single-stage orbital aerospace carrier from the completed design data, and determine the integrated design efficiency of the second carrier based on the flight-launch integration design data corresponding to the second-stage carrier

[0033] The rocket power specific impulse of the second stage carrier is I sp and the integrated design efficiency of the second carrier Combined to obtain the effective specific impulse of the second-stage carrier rocket

[0034] Preferably, the calculation formula for the second-stage weight consumption of the second-stage carrier is:

[0035]

[0036] Preferably, the calculation formula for the initial mass ratio Γ of the single-stage orbital space launch vehicle is:

[0037]

[0038] Where,

[0039] The single-stage orbital aerospace vehicle initial mass ratio estimation method disclosed herein can quickly calculate the takeoff weight to payload initial mass ratio of a single-stage orbital aerospace vehicle. This helps designers and developers better understand and grasp the overall performance of the vehicle, thereby optimizing its design. The method is also applicable to estimating the takeoff weight to payload initial mass ratio of other types of aerospace vehicles. The implementation of this invention can effectively improve the efficiency and accuracy of aerospace vehicle design and development. BRIEF DESCRIPTION OF THE DRAWINGS

[0040] In order to more clearly illustrate the technical solutions provided by this application, the following is a brief introduction to the accompanying drawings. Obviously, the accompanying drawings described below are only some embodiments of this application.

[0041] Figure 1 This is a schematic diagram of the overall process of this application;

[0042] Figure 2 This is a schematic diagram of the forces acting on the single-stage orbital launch vehicle during flight;

[0043] Figure 3 This is a curve diagram of the corresponding relationship between the initial mass ratio and energy distribution ratio of the single-stage orbital aerospace launch vehicle in this application. DETAILED DESCRIPTION

[0044] The following will clearly and completely describe the technical solutions in the embodiments of the present invention in conjunction with the accompanying drawings. 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 making creative efforts are within the scope of protection of the present invention.

[0045] A method for estimating the initial mass ratio of a single-stage orbital space launch vehicle is proposed. The method is based on the basic principles and parameters of the single-stage orbital space launch vehicle design and performs necessary simplifications and calculations through the mathematical formula of the single-stage orbital space launch vehicle's climbing flight phase, thereby improving the accuracy and efficiency of the estimation.

[0046] like Figure 1 As shown, the specific steps include:

[0047] Step S100: Perform force analysis and force equation analysis on the single-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.

[0048] The aerospace vehicle propulsion system consists of the first stage propulsion and the second stage propulsion. The first stage propulsion adopts air-breathing propulsion and the second stage propulsion adopts rocket propulsion. Before and after the aerospace vehicle propulsion mode conversion, the aerospace vehicle does not separate, but the propulsion system is used for relay propulsion. The aerospace vehicle climbs at a dynamic speed and altitude. In this flight state, the speed and geometric altitude of the aircraft will change with time. According to the general force conditions of the aircraft, Figure 2 As shown (force equation).

[0049] The force equation of a single-stage orbital aerospace vehicle is:

[0050] m i1 =m p +m e +m f1 +m f2

[0051] m i2 =m p +m e +m f2 =m i -m f1

[0052] m i =m i1

[0053] m i2 =m i1 -m f1

[0054] ∏ f =∏ f1 +∏ f2 ,

[0055] Where m is the mass of the aircraft, in kg;

[0056] π is the coefficient;

[0057] The subscripts 1 and 2 represent the first-stage carrier and the second-stage carrier, respectively;

[0058] The subscripts i, e, f, and p represent the aircraft’s initial takeoff state, empty state, fuel, and load, respectively;

[0059] Γ represents the initial mass ratio.

[0060] According to the force equation of the single-stage orbital aerospace vehicle, we can obtain:

[0061]

[0062] Step S200: Obtain the fuel type of the first-stage carrier and the flight-launch integration design data of the single-stage orbital aerospace carrier, and then calculate the first-stage weight consumption of the single-stage orbital aerospace carrier in combination with the orbit ratio, and then select the empty weight coefficient Π of the first-stage carrier according to the design requirements. e1 .

[0063] Preferably, the calculation method of the first-level weight consumption is:

[0064] Step S210: Determine the fuel type used by the first stage of the single-stage orbital space vehicle. Generally, hydrocarbon fuel or liquid hydrogen fuel is selected according to the mission characteristics and actual needs. Determine the fuel calorific value h PR , unit is J / kg;

[0065] Step S220: Determine the flight and launch integration design data of the single-stage space launch vehicle based on the preliminary design results or actual conditions. The value of the first stage carrier. is the efficiency of the first stage 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.

[0066] Step S230: Determine the total effective efficiency of the first stage air-breathing power according to the performance data of the air-breathing power provided by the engine unit and the design of the first stage vehicle's flight and engine integration. η0 is the total efficiency of the first-stage vehicle's air-breathing power;

[0067] Step S240: Preliminary determination of the orbit ratio based on the flight data of the single-stage space launch vehicle.

[0068] Step S250: Calculate the first-stage weight consumption of the single-stage space launch vehicle based on the effective total efficiency of the first-stage air-breathing power and the orbit ratio. The calculation formula for the first-stage weight consumption is:

[0069]

[0070] Where g0 is the acceleration of gravity on the Earth's surface, in m / s 2 ; r0 is the average radius of the earth, which is 6371km; η0 is the total efficiency of the air-breathing power at the first stage.

[0071] Step S300: Obtain the fuel type of the second-stage carrier and the flight-launch integration design data of the second-stage carrier to determine the effective specific impulse of the rocket power. Calculate the second stage weight consumption for the second stage vehicle.

[0072] Preferably, the rocket power effective specific impulse The method to obtain is:

[0073] According to the type of fuel used by the rocket power during the second-stage power operation, the rocket power specific impulse I of the corresponding second-stage carrier is determined. sp ;

[0074] Find the flight-launch integration design data corresponding to the second-stage carrier of the single-stage orbital aerospace carrier from the completed design data, and determine the integrated design efficiency of the second carrier based on the flight-launch integration design data corresponding to the second-stage carrier

[0075] The rocket power specific impulse of the second stage carrier is I sp and the integrated design efficiency of the second carrier Combined to obtain the effective specific impulse of the second-stage carrier rocket

[0076] Preferably, the calculation formula for the second-stage weight consumption of the second-stage carrier is:

[0077]

[0078] Step S400: Preliminarily select an energy distribution ratio α based on the weight consumption of the first stage and the second stage of the single-stage space launch vehicle. The energy distribution ratio α is set empirically and represents a portion of the total orbital energy provided by the first stage. It is a part of the orbital energy provided by the second-stage carrier, and then the initial mass ratio Γ of the single-stage orbital aerospace carrier is calculated by the energy distribution ratio α.

[0079] Preferably, the calculation formula for the initial mass ratio Γ of the single-stage orbital space launch vehicle is:

[0080] The energy conversion formula for the first-level weight consumption and the second-level weight consumption is:

[0081]

[0082] Introducing the energy distribution ratio α into the energy conversion formula, we get:

[0083]

[0084] At the same time, due to: Where, π f1 is the fuel coefficient of the first stage carrier, Π f2 is the fuel coefficient of the second stage vehicle.

[0085] Transforming the force equation of the single-stage orbital aerospace vehicle, we obtain:

[0086]

[0087] That is, the initial mass ratio Γ is:

[0088]

[0089] Substituting into the energy conversion formula we get:

[0090]

[0091] After conversion we get:

[0092]

[0093] And then another: The calculation formula for the initial mass ratio Γ of a single-stage space launch vehicle with relay propulsion of air-breathing power and rocket power is:

[0094]

[0095] After conversion we get:

[0096]

[0097] Step S500, repeat steps S200-S400, adjust different energy distribution ratios α, perform trade-off optimization design, and obtain the optimal energy distribution ratio α and the first-stage aerospace vehicle empty weight coefficient ∏ 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.

[0098] The following is an example to illustrate:

[0099] like Figure 3 The first stage carrier uses hydrocarbon fuel, the fuel calorific value h PR =44200kJ / kg, first stage vehicle effective total efficiency Second-stage carrier rocket power effective specific impulse This example Figure 2 It can be seen that when the energy distribution ratio α is 0.30, the initial mass ratio Γ of the single-stage space launch vehicle can take the minimum value Γ min , that is, Λ1=1.4537, Λ2=1.6125, Π e =0.155 Γ min =6.38; Λ1=1.4537, Λ2=1.6125, ∏ e =0.175 Γ min =7.311; Λ1=1.4537, Λ2=1.6125, ∏ e =0.195 Γ min =8.563. It can be seen that the minimum value of the initial mass ratio Γ of the single-stage orbital aerospace vehicle has a significant impact on the energy distribution ratio α and the empty weight coefficient ∏ of the single-stage orbital aerospace vehicle. e All of these are very sensitive and require careful weighing and research to obtain and determine the performance and carrying capacity of a relatively excellent single-stage orbital aerospace launch vehicle.

[0100] The above design enables rapid calculation of the takeoff weight to payload initial mass ratio for a single-stage orbital aerospace vehicle. This helps designers and developers better understand and grasp the overall performance of the vehicle, enabling them to optimize their designs. This method is also applicable to estimating the takeoff weight to payload initial mass ratio for other types of aerospace vehicles. Implementation of this invention can effectively improve the efficiency and accuracy of aerospace vehicle design and development.

[0101] 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. Other structures can refer to common designs. In the absence of conflicts, the same embodiment and different embodiments of the present invention can be combined with each other.

[0102] 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 scope of protection of the present invention.

Claims

1. A method for estimating the initial mass ratio of a single-stage orbital aerospace vehicle, characterized in that: include: Conduct force analysis and force equation analysis of a single-stage orbital launch vehicle, and simplify the results of force analysis and force equation analysis based on flight conditions; Obtain the fuel type of the first-stage vehicle and the flight-launch integration design data of the single-stage orbital aerospace vehicle, and then calculate the first-stage weight consumption of the single-stage orbital aerospace vehicle in combination with the orbital ratio. Then, select the empty weight coefficient ∏e1 of the first-stage vehicle according to the design requirements; Obtain the fuel type of the second-stage vehicle and the flight-launch integration design data of the second-stage vehicle to determine the effective specific impulse of the rocket power Calculate the second stage weight consumption of the second stage vehicle; The energy distribution ratio α is preliminarily selected based on the weight consumption of the first stage and the second stage of the single-stage orbital aerospace vehicle. The energy distribution ratio α is set empirically and represents the 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 vehicle, and then the initial mass ratio Γ of the single-stage space vehicle is calculated by the energy distribution ratio α; By adjusting different energy distribution ratios α and performing trade-off optimization design, the optimal ranges of the energy distribution ratio α, the first-stage aerospace vehicle empty weight coefficient ∏e1, and the second-stage aerospace vehicle empty weight coefficient ∏e2 are obtained for subsequent detailed design analysis.

2. The method for estimating the initial mass ratio of a single-stage orbital launch vehicle according to claim 1, wherein: The force equation of the single-stage orbital aerospace vehicle is: m i1 =m p +m e +m f1 +m f2 m i2 =m p +m e +m f2 =m i -m f1 m i =m i1 m i2 =m i1 -m f1 ∏ f ＝∏ f1 +∏ f2 ， Where m is the mass of the aircraft, in kg; π is the coefficient; The subscripts 1 and 2 represent the first-stage carrier and the second-stage carrier, respectively; The subscripts i, e, f, and p represent the aircraft's initial takeoff state, empty state, fuel, and load, respectively; Γ represents the initial mass ratio; According to the force equation of the single-stage orbital aerospace vehicle, we can obtain:

3. The method for estimating the initial mass ratio of a single-stage orbital launch vehicle according to claim 2, wherein: The calculation method of the first-level weight consumption is: Determine the fuel type used by the air-breathing propulsion system of the first stage of a single-stage orbital space launch vehicle, and determine the fuel calorific value h according to the fuel type. PR ; Determine the first stage carrier based on the preliminary design results or actual conditions The value of the first stage carrier is the efficiency of the first stage 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 total effective 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's air-breathing power; The orbit ratio was preliminarily determined based on the flight data of the single-stage space launch vehicle. The first-stage weight consumption of a single-stage space launch vehicle is calculated based on the effective total efficiency of the first-stage air-breathing propulsion and the orbit ratio. The first-stage weight consumption calculation formula is: Where g0 is the acceleration of gravity on the Earth's surface, in m / s 2 ; r0 is the average radius of the earth; η0 is the total efficiency of the air-breathing power at the first stage of power.

4. The method for estimating the initial mass ratio of a single-stage orbital launch vehicle according to claim 3, wherein: The rocket power effective specific impulse The method to obtain is: According to the type of fuel used by the rocket power during the second-stage power operation, the rocket power specific impulse I of the corresponding second-stage carrier is determined. sp ; Find the flight-launch integration design data corresponding to the second-stage carrier of the single-stage orbital aerospace carrier from the completed design data, and determine the integrated design efficiency of the second carrier based on the flight-launch integration design data corresponding to the second-stage carrier The rocket power specific impulse of the second stage carrier is I sp and the integrated design efficiency of the second carrier Combined to obtain the effective specific impulse of the second-stage carrier rocket 5. The method for estimating the initial mass ratio of a single-stage orbital launch vehicle according to claim 4, wherein: The calculation formula for the second-stage weight consumption of the second-stage carrier is:

6. The method for estimating the initial mass ratio of a single-stage orbital launch vehicle according to claim 4, wherein: The calculation formula of the initial mass ratio Γ of the single-stage orbital space launch vehicle is: Where,

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