A method for estimating mass of aerospace vehicles

Abstract

The present application belongs to the field of overall design of single-stage orbital aerospace vehicles, and particularly relates to a method for estimating the mass of an aerospace vehicle. It includes: step one, obtaining the general equation of the vehicle acceleration process; step two, constructing a mass ratio estimation model of the rocket propulsion mode according to the general equation of the vehicle acceleration process; step three, constructing a mass ratio estimation model of the air-breathing propulsion mode according to the general equation of the vehicle acceleration process and the mass ratio estimation model of the rocket propulsion mode; step four, constructing a mass estimation model of the single-stage orbital aerospace vehicle according to the mass ratio estimation model of the rocket propulsion mode and the mass ratio estimation model of the air-breathing propulsion mode. The present application completes the estimation of the total mass of the single-stage orbital aerospace vehicle, and can quickly obtain the law between the power mode transition point, the vehicle structure mass fraction, the orbital load and the total mass of the vehicle, thereby improving the convergence speed of the solution.

Description

Technical Field

[0001] The present application belongs to the field of overall design of single-stage orbital aerospace vehicles, and in particular, relates to a method for estimating the mass of an aerospace vehicle. Background Art

[0002] A single-stage orbital aerospace vehicle is a reusable spacecraft used to deliver an orbital payload to orbital space. At present, there is a lack of research on conceptual design methods suitable for single-stage orbital vehicles at home and abroad, and there is no realistic single-stage orbital vehicle for reference. The research work on the single-stage orbital aerospace vehicle scheme mainly focuses on the scale assessment of the entire system and the selection of aerodynamic layout. In terms of vehicle mass assessment research, the research work that has been carried out can be divided into two categories: one is to obtain the relationship between the size and mass of the vehicle through the statistical data of the overall parameters of the vehicle, using the regression analysis method. This type of research focuses on quality assessment and does not consider whether the vehicle has actual orbital capability; the other is to analyze the propellant consumption rate based on the assumed structural mass coefficient under certain load requirements through the energy conservation relationship, thereby analyzing the mass scale of the vehicle. The existing technology is difficult to analyze the factors affecting the takeoff total weight of a single-stage orbital aerospace vehicle, and the cycle is relatively long, which is difficult to meet the needs of rapid demonstration.

[0003] Therefore, it is desired to have a technical solution to overcome or at least alleviate at least one of the above-mentioned defects of the prior art. Summary of the invention

[0004] The purpose of this application is to provide a method for estimating the mass of an aerospace vehicle to solve at least one problem existing in the prior art.

[0005] The technical solution of this application is:

[0006] A method for estimating the mass of an aerospace vehicle, comprising:

[0007] Step 1: Obtain the general equation of the aircraft acceleration process;

[0008] Step 2: construct a mass ratio estimation model of the rocket propulsion mode according to the general equation of the aircraft acceleration process;

[0009] Step 3: construct a mass ratio estimation model for an air-breathing propulsion mode according to the general equation of the aircraft acceleration process and the mass ratio estimation model for the rocket propulsion mode;

[0010] Step 4: Construct a mass estimation model for a single-stage orbital aerospace vehicle based on the mass ratio estimation model of the rocket propulsion mode and the mass ratio estimation model of the air-breathing propulsion mode.

[0011] In at least one embodiment of the present application, in step 1, the general equation for obtaining the aircraft acceleration process includes:

[0012] A general relationship between the acceleration and propellant consumption rate of a general aircraft is established, where Newton's second law is applied in the direction of flight speed and the thrust is expressed in terms of engine specific impulse:

[0013]

[0014] Define the effective specific impulse as:

[0015] I eff =I sp (1-D / T)

[0016] The general equation for the aircraft acceleration process is obtained as follows:

[0017]

[0018] In the formula, m is the mass of the aircraft, V is the speed of the aircraft, and I sp is the engine specific impulse, T is the engine thrust, D is the vehicle drag, g is the acceleration due to gravity, and γ is the vehicle climb angle.

[0019] In at least one embodiment of the present application, in step 2, constructing a mass ratio estimation model of a rocket propulsion mode according to a general equation of the aircraft acceleration process includes:

[0020] Since the thrust-to-drag ratio of the aircraft is large in the rocket propulsion mode, it is considered that the effective specific impulse I eff Approximately equal to I sp , rocket engine I sp It is approximately constant during the flight process. The first transformation equation is obtained by integrating the general equation of the aircraft acceleration process:

[0021]

[0022] Where ΔV i is the velocity increment of the ith mode, μ i is the mass ratio of the ith mode;

[0023] in,

[0024] ΔV i =V i+1 -V i

[0025] μ i =m i / m i+1

[0026]

[0027] Where ΔV g It is the gravity loss, which is generally small compared to the velocity increment ΔV and is ignored in the initial estimation;

[0028] The average velocity V of the i-th mode m Estimation of ΔV g ,but:

[0029] V m =(V i +V i+1 ) / 2

[0030]

[0031] Where ΔH is the height increment;

[0032] The gravity loss ΔV g Substituting into the first transformation equation, the mass ratio estimation model of the rocket propulsion mode is obtained as follows:

[0033]

[0034] Where V i is the speed of the ith flight node, V i+1 is the speed of the i+1th flight node, H i is the height of the ith flight node, H i+1 is the height of the i+1th flight node.

[0035] In at least one embodiment of the present application, the mass ratio estimation model of the rocket propulsion mode is written in the following form:

[0036]

[0037] in,

[0038] ΔV gi =2g(H i+1 -H i ) / (V i +V i+1 )

[0039]

[0040] Where ΔV gi is the gravity loss of the ith mode, is the average effective specific impulse.

[0041] In at least one embodiment of the present application, in step 3, constructing the mass ratio estimation model of the air-breathing propulsion mode according to the general equation of the aircraft acceleration process and the mass ratio estimation model of the rocket propulsion mode includes:

[0042] Since the thrust-drag ratio of the aircraft is relatively small in the air-breathing propulsion mode, the drag loss cannot be ignored, and the engine specific impulse I sp It is no longer a constant, so an estimation method different from the rocket propulsion mode is used. The general relationship between the acceleration of a general aircraft and the propellant consumption rate is multiplied by V and divided by m to obtain the second transformation equation:

[0043]

[0044] Where TV is the engine power;

[0045]

[0046] Substituting the engine power TV into the second transformation equation and assuming that the thrust-to-drag ratio is a constant, the mass ratio estimation model of the air-breathing propulsion mode is obtained by integration:

[0047]

[0048] Where η is the thermal efficiency of the engine and q is the calorific value of the fuel.

[0049] In at least one embodiment of the present application, the mass ratio estimation model of the air-breathing propulsion mode is written in the following form:

[0050]

[0051] in,

[0052] ΔV gi =2g(H i+1 -H i ) / (V i +V i+1 )

[0053]

[0054] Where ΔV gi is the gravity loss of the ith mode, is the average effective specific impulse.

[0055] In at least one embodiment of the present application, in step 4, constructing a mass estimation model for a single-stage orbital aerospace vehicle according to the mass ratio estimation model of the rocket propulsion mode and the mass ratio estimation model of the air-breathing propulsion mode includes:

[0056] The mass ratio μ of each mode after the mission profile is determined according to the mass ratio estimation model of the rocket propulsion mode and the mass ratio estimation model of the air-breathing propulsion mode. i ;

[0057] The total mass ratio of all modes of single-stage orbital aerospace vehicle propulsion μ b for:

[0058]

[0059] Define the structural quality fraction σ b for:

[0060]

[0061] For a given task load m pl , propellant mass m bp , structure and equipment quality m bs They are:

[0062]

[0063]

[0064] The mass estimation model of a single-stage orbital aerospace vehicle is:

[0065]

[0066] In the formula, m to is the total mass of the single-stage orbital aerospace vehicle, m bp is the propellant mass, m bs is the mass of structure and equipment, m pl For mission load.

[0067] The invention has at least the following beneficial technical effects:

[0068] The aerospace vehicle mass estimation method of the present application completes the estimation of the total mass of a single-stage orbital aerospace vehicle, can quickly obtain the law between the dynamic mode transition point, the vehicle structure mass fraction, the orbital load and the total mass of the vehicle, and improves the speed of solution convergence. BRIEF DESCRIPTION OF THE DRAWINGS

[0069] Figure 1 It is a schematic diagram of a preliminary mission profile of a single-stage orbital aerospace vehicle according to one embodiment of the present application;

[0070] Figure 2 It is a schematic diagram of the change law between the mass of the orbital payload and the total mass of the aircraft in the air-breathing propulsion and rocket-powered modes in one embodiment of the present application;

[0071] Figure 3 is a schematic diagram of the variation law between the aircraft structure mass fraction and the total mass of the aircraft according to one embodiment of the present application;

[0072] Figure 4It is a schematic diagram of the change law between the flight Mach number and the total mass of the aircraft when the air-breathing power mode and the rocket power mode are switched according to an embodiment of the present application;

[0073] Figure 5 It is a schematic diagram of the change law between the flight altitude and the total mass of the aircraft when the air-breathing power mode and the rocket power mode are converted in one embodiment of the present application. DETAILED DESCRIPTION

[0074] In order to make the purpose, technical scheme and advantages of the implementation of this application clearer, the technical scheme in the embodiment of this application will be described in more detail below in conjunction with the drawings in the embodiment of this application. In the drawings, the same or similar reference numerals throughout represent the same or similar elements or elements with the same or similar functions. The described embodiments are part of the embodiments of this application, not all of them. The embodiments described below with reference to the drawings are exemplary and are intended to be used to explain this application, and should not be construed as limitations on this application. Based on the embodiments in this application, all other embodiments obtained by ordinary technicians in this field without making creative work are within the scope of protection of this application. The embodiments of this application are described in detail below in conjunction with the drawings.

[0075] In the description of the present application, it should be understood that the terms "center", "longitudinal", "lateral", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inside", "outside", etc., indicating orientations or positional relationships, are based on the orientations or positional relationships shown in the accompanying drawings, and are only for the convenience of describing the present application and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and therefore should not be understood as limiting the scope of protection of the present application.

[0076] The following is combined with Figures 1 to 5 This application is described in further detail.

[0077] The present application provides a method for estimating the mass of an aerospace vehicle, comprising the following steps:

[0078] S001. Obtain the general equation of the aircraft acceleration process;

[0079] S002. Construct a mass ratio estimation model of the rocket propulsion mode based on the general equation of the aircraft acceleration process;

[0080] S003. Constructing a mass ratio estimation model for an air-breathing propulsion mode according to a general equation for an aircraft acceleration process and a mass ratio estimation model for a rocket propulsion mode;

[0081] S004. Construct a mass estimation model for a single-stage orbital aerospace vehicle based on the mass ratio estimation model of the rocket propulsion mode and the mass ratio estimation model of the air-breathing propulsion mode.

[0082] The speed and altitude of each flight node in the mission profile of a single-stage orbital aerospace vehicle are (V i , H i ), (i=1, 2, ..., N), where (V i , H i )=(0,0) indicates the take-off state, (V N , H N ) are the orbital velocity and orbital altitude required by the mission.

[0083] Define Node(V i , H i ) to (V i+1 , H i+1 ) is the i-th mode. The mass of the aircraft at the i-th node is m i , define the mass ratio of the i-th mode as μ i =m i / m i+1 , the velocity increment is ΔV i =V i+1 -V i , the height increment is ΔH i =H i+1 -H i The following specifically establishes a mass estimation model for a single-stage orbital aerospace vehicle.

[0084] First, S001, obtain the general equation of the aircraft acceleration process including:

[0085] A general relationship between the acceleration and propellant consumption rate of a general aircraft is established, where Newton's second law is applied in the direction of flight speed and the thrust is expressed in terms of engine specific impulse:

[0086]

[0087] Define the effective specific impulse as:

[0088] I eff =I sp (1-D / T)

[0089] The general equation for the aircraft acceleration process is obtained as follows:

[0090]

[0091] In the formula, m is the mass of the aircraft, V is the speed of the aircraft, and I spis the engine specific impulse, T is the engine thrust, D is the vehicle drag, g is the acceleration due to gravity, and γ is the vehicle climb angle.

[0092] Compared with the engine specific impulse, the effective specific impulse that takes into account the flight resistance loss can better reflect the overall acceleration performance of the aircraft. The larger the thrust-to-drag ratio T / D of the aircraft, the closer the effective specific impulse is to the engine specific impulse.

[0093] S002. According to the general equation of the aircraft acceleration process, the mass ratio estimation model of the rocket propulsion mode is constructed, including:

[0094] Since the thrust-to-drag ratio of the aircraft is large in the rocket propulsion mode, it is considered that the effective specific impulse I eff Approximately equal to I sp , rocket engine I sp It is approximately constant during the flight process. The first transformation equation is obtained by integrating the general equation of the aircraft acceleration process (Formula (2)):

[0095]

[0096] Where ΔV i is the velocity increment of the ith mode, μ i is the mass ratio of the ith mode;

[0097] in,

[0098] ΔV i =V i+1 -V i

[0099] μ i =m i / m i+1

[0100]

[0101] Where ΔV g It is the gravity loss, which is generally small compared to the velocity increment ΔV and cannot be accurately obtained in the conceptual design stage, so it is often ignored in the preliminary estimation;

[0102] Since more flight phases are considered and a larger range of altitude changes are involved, the average speed V of the i-th mode is m Estimation of ΔV g ,but:

[0103] V m =(V i +V i+1 ) / 2

[0104]

[0105] Where ΔH is the height increment;

[0106] The gravity loss ΔV g Substituting the first transformation equation (substituting formula (4) into formula (3)), the mass ratio estimation model of the rocket propulsion mode is obtained as follows:

[0107]

[0108] Where V i is the speed of the ith flight node, V i+1 is the speed of the i+1th flight node, H i is the height of the ith flight node, H i+1 is the height of the i+1th flight node.

[0109] Further, S003, based on the general equation of the aircraft acceleration process and the mass ratio estimation model of the rocket propulsion mode, constructing the mass ratio estimation model of the air-breathing propulsion mode includes:

[0110] Since the thrust-drag ratio of the aircraft is relatively small in the air-breathing propulsion mode, the drag loss cannot be ignored, and the engine specific impulse I sp It is no longer a constant, so an estimation method different from the rocket propulsion mode is used. The general relationship between the acceleration of a general aircraft and the propellant consumption rate (Formula (1)) is multiplied by V and divided by m to obtain the second transformation equation:

[0111]

[0112] Where TV is the engine power;

[0113]

[0114] Substituting the engine power TV into the second transformation equation (substituting formula (7) into formula (6)), and assuming that the thrust-to-drag ratio is a constant, the mass ratio estimation model of the air-breathing propulsion mode is obtained by integration:

[0115]

[0116] Where η is the thermal efficiency of the engine and q is the calorific value of the fuel.

[0117] Advantageously, the mass ratio estimation model for the rocket propulsion mode (Formula (5)) and the mass ratio estimation model for the air-breathing propulsion mode (Formula (8)) can be finally written in a unified form:

[0118]

[0119] Among them, the gravity loss ΔV of the i-th mode is gi for:

[0120] ΔV gi =2g(H i+1 -H i ) / (V i +V i+1 )

[0121] The average effective specific impulse reflects the acceleration performance of the aircraft.

[0122] For rocket propulsion modes, the average effective specific impulse for:

[0123]

[0124] For the air-breathing propulsion mode, the average effective specific impulse for:

[0125]

[0126] Finally, S004, based on the mass ratio estimation model of the rocket propulsion mode and the mass ratio estimation model of the air-breathing propulsion mode, a mass estimation model of the single-stage orbital aerospace vehicle is constructed, including:

[0127] According to the mass ratio estimation model of rocket propulsion mode (Formula (5)) and the mass ratio estimation model of air-breathing propulsion mode (Formula (8)), the mass ratio μ of each mode after the mission profile is determined. i ;

[0128] Thus, the total mass ratio μ of all modes of propulsion of a single-stage orbital aerospace vehicle can be obtained: b for:

[0129]

[0130] Define the structural quality fraction σ b for:

[0131]

[0132] For a given task load m pl , combining equations (12) and (13), we can calculate the propellant mass m bp , structure and equipment quality m bs They are:

[0133]

[0134]

[0135] The mass estimation model of a single-stage orbital aerospace vehicle is:

[0136]

[0137] In the formula, m to is the total mass of the single-stage orbital aerospace vehicle, m bp is the propellant mass, m bs is the mass of structure and equipment, m pl For mission load.

[0138] Combined with the mass ratio μ of each mode i , the structure and equipment mass of the aircraft and the mass of propellant consumed in each mode can be calculated.

[0139] The aerospace vehicle mass estimation method of the present application derives the relationship between the vehicle mass and flight speed, climb angle, engine specific impulse and other parameters based on the dynamic equation of the vehicle climbing process; and calculates the mass ratio of each power mode based on parameters such as speed increment, altitude increment, thrust-to-drag ratio, mass acceleration and specific impulse. The mass ratio calculation formula of the rocket power mode is formula (5), and the mass ratio calculation formula of the air-breathing power mode is formula (8); obtains the total mass ratio of all modes of vehicle propulsion, and based on the developed single-stage orbital flight plan, inquires and gives the vehicle structure mass fraction, and obtains the total mass ratio μ b and the vehicle structure mass fraction σ b Based on the above, the mass of the aircraft propellant m is calculated according to formulas (14), (15) and (16) respectively: bp , structural mass m bs And the total takeoff mass of the aircraft m to .

[0140] The present application can deflect parameters such as the power mode transition point, the vehicle structure mass fraction, and the orbital load as needed, and quickly calculate the law between the power mode transition point, the vehicle structure mass fraction, the orbital load and the total mass of the vehicle.

[0141] In a preferred embodiment of the present application, Figure 2 The variation law between the orbital payload mass and the total mass of the vehicle under air-breathing propulsion and rocket propulsion modes is given when the dynamic mode transition point is at an altitude of 26 kilometers, a flight speed of 5 Mach, and the vehicle structure mass fraction is 0.19. Figure 3 The variation law between the vehicle structure mass fraction and the total mass of the vehicle is given when the orbital load is 3000 kg, the dynamic mode transition point is at an altitude of 26 km, and the speed is 5 Mach. Figure 4 The variation law between the flight Mach number and the total mass of the vehicle during the transition between the air-breathing power mode and the rocket power mode is given when the orbital load is 3000 kg, the power mode transition flight altitude is 26 km, the flight speed is 5 Mach, and the vehicle structure mass fraction is 0.19. Figure 5The change law between the flight altitude and the total mass of the vehicle when the air-breathing power mode and the rocket power mode are switched is given when the orbital load is 3000 kg, the vehicle structure mass fraction is 0.19, and the flight speed is 5 Mach. From the implementation effect, it can be obtained that the total takeoff mass of the single-stage orbital vehicle is m to Mass of the orbital load m pl The increase is large, and the change relationship is basically linear. The total takeoff mass is more sensitive to the change of the mass of the orbital payload. The total takeoff mass of the single-stage orbital vehicle is m to With the aircraft structure mass fraction σ b The aircraft structure quality fraction σ changes rapidly and greatly due to the change of b The total takeoff mass of a single-stage orbital vehicle is m to The impact is huge and positively correlated; the total takeoff mass of a single-stage orbital vehicle is m to The flight speed changes greatly with the power mode conversion, and changes in a negative correlation, mainly due to the fact that the rocket power mode needs to carry its own oxidizer, resulting in a lower specific impulse; the total takeoff mass of the single-stage orbital vehicle is m to The flight altitude does not change much with the power mode transition, and the flight altitude at the mode transition point is not an important factor affecting the total mass of the aircraft.

[0142] The aerospace vehicle mass estimation method of the present application, for a single-stage orbital aerospace vehicle, takes into account the changes in specific impulse of the engine in different modes and the differences in thrust-drag characteristics in different flight phases, completes a rapid estimation of the total mass of the single-stage orbital aerospace vehicle, and provides a sensitivity analysis of the main parameters.

[0143] The above is only a specific implementation of the present application, but the protection scope of the present application is not limited thereto. Any changes or substitutions that can be easily thought of by a person skilled in the art within the technical scope disclosed in the present application should be included in the protection scope of the present application. Therefore, the protection scope of the present application shall be based on the protection scope of the claims.

Claims

1. A method for estimating the mass of an aerospace vehicle, characterized in that: include: Step 1: Get the general equation of the aircraft acceleration process including: A general relationship between the acceleration and propellant consumption rate of a general aircraft is established, where Newton's second law is applied in the direction of flight speed and the thrust is expressed in terms of engine specific impulse: Define the effective specific impulse as: I eff =I sp (1-D / T) The general equation for the aircraft acceleration process is obtained as follows: In the formula, m is the mass of the aircraft, V is the speed of the aircraft, and I sp is the engine specific impulse, T is the engine thrust, D is the vehicle drag, g is the gravitational acceleration, and γ is the vehicle climb angle; Step 2: According to the general equation of the aircraft acceleration process, a mass ratio estimation model of the rocket propulsion mode is constructed, including: Since the thrust-to-drag ratio of the aircraft is large in the rocket propulsion mode, it is considered that the effective specific impulse I eff Approximately equal to I sp , rocket engine I sp It is approximately constant during the flight process. The first transformation equation is obtained by integrating the general equation of the aircraft acceleration process: Where, ΔV i is the velocity increment of the ith mode, μ i is the mass ratio of the ith mode; in, ΔV i =V i+1 -V i m i =m i / m i+1 Where, ΔV g It is the gravity loss, which is generally small compared to the velocity increment ΔV and is ignored in the initial estimation; The average velocity V of the i-th mode m Estimation of ΔV g ,but: V m =(V i +V i+1 ) / 2 Where ΔH is the height increment; The gravity loss ΔV g Substituting into the first transformation equation, the mass ratio estimation model of the rocket propulsion mode is obtained as follows: Where V i is the speed of the ith flight node, V i+1 is the speed of the i+1th flight node, H i is the height of the ith flight node, H i+1 is the height of the i+1th flight node; Step 3: According to the general equation of the aircraft acceleration process and the mass ratio estimation model of the rocket propulsion mode, a mass ratio estimation model of the air-breathing propulsion mode is constructed, which includes: Since the thrust-drag ratio of the aircraft is relatively small in the air-breathing propulsion mode, the drag loss cannot be ignored, and the engine specific impulse I sp It is no longer a constant, so an estimation method different from the rocket propulsion mode is used. The general relationship between the acceleration of a general aircraft and the propellant consumption rate is multiplied by V and divided by m to obtain the second transformation equation: Where TV is the engine power; Substituting the engine power TV into the second transformation equation and assuming that the thrust-to-drag ratio is a constant, the mass ratio estimation model of the air-breathing propulsion mode is obtained by integration: In the formula, η is the thermal efficiency of the engine, and q is the calorific value of the fuel; Step 4: constructing a mass estimation model for a single-stage orbital aerospace vehicle based on the mass ratio estimation model of the rocket propulsion mode and the mass ratio estimation model of the air-breathing propulsion mode includes: The mass ratio μ of each mode after the mission profile is determined according to the mass ratio estimation model of the rocket propulsion mode and the mass ratio estimation model of the air-breathing propulsion mode. i ; The total mass ratio of all modes of single-stage orbital aerospace vehicle propulsion μ b for: Define the structural quality fraction σ b for: For a given task load m pl , propellant mass m bp , structure and equipment quality m bs They are: The mass estimation model of a single-stage orbital aerospace vehicle is: In the formula, m to is the total mass of the single-stage orbital aerospace vehicle, m bp is the propellant mass, m bs is the mass of structure and equipment, m pl For mission load.

2. The method for estimating the mass of an aerospace vehicle according to claim 1, characterized in that: The mass ratio estimation model of the rocket propulsion mode is written as follows: in, ΔV gi =2g(H i+1 -H i ) / (V i +V i+1 ) Where, ΔV gi is the gravity loss of the ith mode, is the average effective specific impulse.

3. The method for estimating the mass of an aerospace vehicle according to claim 2, characterized in that: The mass ratio estimation model of the air-breathing propulsion mode is written as follows: in, ΔV gi =2g(H i+1 -H i ) / (V i +V i+1 ) Where, ΔV gi is the gravity loss of the ith mode, is the average effective specific impulse.

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