Method and system for simulating full flight process of reusable carrier rocket

Abstract

The invention discloses a method and a system for simulating the whole flight process of a reusable carrier rocket, and the method comprises the steps: dividing the whole flight process of the carrier rocket into seven stages; the method comprises the following steps of: performing an ascending stage before gravity steering, a first-stage ascending stage, a stage before entering the atmosphere, an engine swinging and overturning attitude stage, a stage after entering the atmosphere, a vertical attitude adjusting stage and a power descending stage, and performing iterative analysis on flight parameter variables in each stage according to a dynamic differential equation. The performance parameters of the reusable carrier rocket in different stages can be accurately and rapidly calculated and evaluated, the motion condition of the reusable carrier rocket in the whole flight process can be fully reflected, and the method is more suitable for the actual test run condition.

Description

Technical Field

[0001] The present invention belongs to the field of aerospace technology, and in particular relates to a method and system for simulating the entire flight process of a reusable launch vehicle. Background Art

[0002] With the frequent development of space activities, people have put forward an urgent demand for low-cost and high-efficiency launch vehicles. Reusable rockets have become an important research direction in the aerospace field because they can reduce launch costs and improve launch efficiency. At present, the application of reusable rocket technology has penetrated into many fields, among which the most eye-catching is SpaceX's Starlink project, which launches thousands of satellites through reusable Falcon 9 rockets to build global Internet coverage.

[0003] The development of reusable rocket technology faces many technical challenges, such as precise return flight and safe landing control technology, variable thrust reusable engine technology, and high-reliability landing buffer mechanism technology. The gradual maturity of these technologies is inseparable from a large number of experiments, but with the increase in the number of experiments, the simultaneous increase in the number of failures has also brought a lot of economic losses. For example, during a ground test in September 2016, the Falcon 9 rocket exploded, destroying the rocket and its communications satellite AMOS-6S. Therefore, it is crucial to simulate the full flight process of reusable rockets to reduce the cost of actual test runs and failures.

[0004] In order to achieve the performance prediction and optimization of reusable launch vehicles, it is necessary to perform precise computer programming and simulation analysis on the aerodynamic parameters of the rocket, including determining the parameters of the rocket body motion process. These parameters are numerous and involve a large number of theoretical formula calculations. Using computer programming to determine the parameters of the rocket body motion process is crucial for quickly completing the performance prediction and optimization of reusable rocket engines and effectively shortening the research and development cycle. This can not only improve design efficiency, but also test various extreme conditions of the rocket in a virtual environment, reduce the risk of actual test runs, reduce environmental pollution, and accelerate the development and iteration of technology.

[0005] The Chinese invention patent application with the announcement number CN106021628B proposed a design method for the vertical return trajectory of a launch vehicle. The method established a three-degree-of-freedom dynamic model of the vertical return of the first stage of the launch vehicle and considered the flight overload, dynamic pressure and heat flow peak constraints, and terminal constraints. However, the scheme lacked a corresponding description of the motion process of the rocket's ascent phase and could not objectively and comprehensively reflect the motion of the reusable launch vehicle during the entire flight process. Yu Guangxue et al. (Yu Guangxue, Li Zhaotang, Lin Ping. Mathematical Modeling of Reusable Launch Vehicle Reentry [J]. Chinese Space Science and Technology, 2014, 23(3): 24-31.) proposed a study on the conceptual design and control modeling of the reusable launch vehicle (RLV) reentry. The RLV aerodynamic model was established by establishing a reaction control system / aerodynamic surface composite control mathematical model and an engineering calculation method based on aerodynamic forces. However, the scheme lacked the description and calculation of the rocket body performance parameters and could not relatively accurately reflect the motion of the reusable launch vehicle during the entire flight process. Summary of the invention

[0006] In order to overcome the shortcomings of the above-mentioned prior art, the purpose of the present invention is to provide a method and system for simulating the entire flight process of a reusable launch vehicle. By subdividing the entire flight process of the launch vehicle into seven stages, the flight parameter variables in each stage are iteratively analyzed according to the dynamic differential equation. This method can not only accurately and quickly calculate and evaluate the performance parameters of the reusable launch vehicle at different stages, but also fully reflect the movement conditions of the reusable launch vehicle during the entire flight process.

[0007] In order to achieve the above object, the technical solution of the present invention is as follows:

[0008] A method for simulating the entire flight process of a reusable launch vehicle comprises the following steps:

[0009] Step 1: Set and initialize the flight parameter variables of the launch vehicle during the entire flight process, wherein the flight parameter variables include the flight speed v, the flight path inclination γ, the flight altitude h, the horizontal position x between the launch vehicle and the launch site, and the launch vehicle mass m;

[0010] Step 2: Divide the entire flight process of the carrier rocket into a launch phase and a recovery phase, wherein the launch phase includes an ascent phase before gravity turning, a first-stage ascent phase, and a pre-reentry phase; the recovery phase includes an engine swing and flip attitude phase, a reentry phase, a vertical attitude adjustment phase, and a powered descent phase; according to the flight parameter variables initialized in step 1, the flight parameter variables of the carrier rocket in each flight phase are iterated based on the dynamic differential equation to obtain the flight parameter variables of the carrier rocket in each phase, and solve the gravity y, flight Mach number η, and range R of the corresponding phase;

[0011] Step 3: Draw a graph of the launch vehicle's flight parameter variables, gravity y, flight Mach number η, and range R based on the flight parameter variables, gravity y, flight Mach number η, and range R of each stage obtained in step 2.

[0012] The process of iterating the flight parameter variables of the launch vehicle at all stages in step 2 is as follows:

[0013] Step 2.1: According to the flight parameter variables initialized in step 1, set the running time t of the ascent phase before gravity turn 1 According to the dynamic differential equation, the flight parameter variables of the launch vehicle in the ascent phase before gravity turning and the corresponding gravity y, Mach number η and range R of this phase are solved until the running time t of this phase is reached. 1 ;

[0014] Step 2.2: Use the flight parameter variables of the ascent phase before gravity is turned in step 2.1 as the starting variables of the first-stage ascent phase, iterate the flight parameter variables of the launch vehicle in the first-stage ascent phase until the fuel set for the first-stage ascent phase is exhausted, separate the excess mass of the launch vehicle, and obtain the flight parameter variables of the launch vehicle in the first-stage ascent phase and the gravity y, Mach number η, and range R corresponding to this phase;

[0015] Step 2.3: Use the flight parameter variables of the first stage ascent phase in step 2.2 as the starting variables of the pre-reentry phase and set the running time t of this phase. 3 , iterate the flight parameter variables of the launch vehicle in the pre-reentry stage until the running time t set for this stage is reached 3 , obtain the flight parameter variables of the launch vehicle in the pre-reentry stage and the gravity y, Mach number η and range R corresponding to this stage;

[0016] Step 2.4: Use the flight parameter variables of the pre-reentry stage in step 2.3 as the starting variables of the engine swing and flip attitude stage, iterate the flight parameter variables of the launch vehicle in the engine swing and flip attitude stage until the flight path inclination angle of this stage reaches the reentry angle, and obtain the flight parameter variables of the launch vehicle in the engine swing and flip attitude stage and the gravity y, Mach number η and range R corresponding to this stage;

[0017] Step 2.5: Use the flight parameter variables of the engine swing and flip attitude phase in step 2.4 as the starting variables of the re-entry phase, and set the running time t of the re-entry phase 5 , iterate the flight parameter variables of the launch vehicle in the re-entry phase until the running time t set for this phase is reached 5, obtain the flight parameter variables of the launch vehicle in the atmospheric reentry phase and the gravity y, Mach number η and range R corresponding to this phase;

[0018] Step 2.6: Use the flight parameter variables of the reentry phase in step 2.5 as the starting variables of the vertical attitude adjustment phase, and set the running time t of the vertical attitude adjustment phase. 6 , iterate the flight parameter variables of the launch vehicle in the vertical attitude adjustment phase until the running time t set for this phase is reached 6 , obtain the flight parameter variables of the launch vehicle in the vertical attitude adjustment stage and the gravity y, Mach number η and range R corresponding to this stage;

[0019] Step 2.7: Use the flight parameter variables of the vertical attitude adjustment phase in step 2.6 as the starting variables of the powered descent phase, and iterate the flight parameter variables of the launch vehicle in the powered descent phase until the flight altitude h≤0.

[0020] The iterative equations of the flight parameter variables at each stage of the launch vehicle's full flight process in step 2 are as follows:

[0021]

[0022] Among them, S1 represents the ascent stage before gravity turns, S2 represents the first-stage ascent stage, S3 represents the stage before re-entry, S4 represents the engine swing and flip attitude stage, S5 represents the re-entry stage, S6 represents the vertical attitude adjustment stage, and S7 represents the power descent stage; T is the thrust of the launch vehicle engine at the current stage, D is the aerodynamic drag of the launch vehicle at the current moment, m is the mass of the launch vehicle at the current moment, g is the gravity acceleration of the launch vehicle, and R e is the radius of the earth, h is the flight altitude of the launch vehicle at the current moment, V is the speed of the launch vehicle relative to the air at the current moment, γ i-1 is the flight path inclination angle at the previous moment, γ i is the inclination of the flight path at the current moment, v is the flight speed of the launch vehicle at the current moment, I sp is the specific impulse of the launch vehicle engine at the current stage, g 0 is the initial acceleration of the launch vehicle.

[0023] In step 2.2, the running time t of the first stage ascent phase is calculated according to the fuel flow rate of the launch vehicle. 2 , the expression is as follows:

[0024]

[0025] Where, mp is the total propellant mass of the launch vehicle before the first ascent phase begins, mf is the remaining propellant mass of the launch vehicle, g0 is the initial acceleration of the launch vehicle in the first stage of ascent, T is the thrust of the launch vehicle engine in the current stage, I sp It is the specific impulse of the launch vehicle engine at the current stage.

[0026] When the fuel set for the first stage ascent in step 2.2 is exhausted, the excess mass of the carrier rocket is separated, that is, the first stage of the carrier rocket is retained, and its second stage and third stage are separated. The mass m of the carrier rocket after separation is as follows:

[0027] m=m 1 +m 2

[0028] Among them, m 1 is the dry mass of the first stage rocket, m 2 is the weight of the remaining fuel.

[0029] The aerodynamic drag D of each stage during the entire flight of the launch vehicle is:

[0030] D=0.5·A·ρ·C d ·V 2

[0031] Among them, A is the windward area of ​​the launch vehicle at the current moment, ρ is the air density, C d is the air resistance coefficient, and V is the speed of the launch vehicle relative to the air at the current moment.

[0032] The thrust T of the launch vehicle engine at each stage during the entire flight of the launch vehicle is as follows:

[0033]

[0034] Among them, S1 represents the ascent stage before gravity turning, S2 represents the first-stage ascent stage, S3 represents the stage before re-entry, S4 represents the engine swing and flip attitude stage, S5 represents the re-entry stage, S6 represents the vertical attitude adjustment stage, and S7 represents the power descent stage; T 1 is the thrust of the launch vehicle engine, C t is the thrust coefficient; h is the height of the launch vehicle in meters; y is the gravity of the launch vehicle at the current moment.

[0035] The gravity y and the flight Mach number η at each stage of the whole flight process of the carrier rocket are:

[0036]

[0037] Among them, M eis the mass of the Earth, a is the speed of sound of the launch vehicle at the current flight altitude h, v is the flight speed of the launch vehicle at the current moment, G is the gravitational constant, R e is the radius of the earth, m is the mass of the launch vehicle at the current moment;

[0038] The range R of each stage during the entire flight of the launch vehicle is:

[0039]

[0040] Among them, l k The length of the flight time t vector solved for each stage, a k is the upper limit of the flight time t variable subscript in the function of the flight range R at the current moment, v i is the flight speed of the launch vehicle at the current moment, t i is the current moment, t i-1 For the previous moment.

[0041] The present invention also provides a full flight process simulation system for a reusable launch vehicle, comprising:

[0042] Initialization module: used to set and initialize the flight parameter variables of the launch vehicle during the entire flight process;

[0043] The calculation module is used to divide the whole flight process of the carrier rocket into a launch phase and a recovery phase, wherein the launch phase includes an ascent phase before gravity turning, a first-stage ascent phase, and a pre-reentry phase; the recovery phase includes an engine swing and flip attitude phase, a reentry phase, a vertical attitude adjustment phase, and a power descent phase; according to the initialized flight parameter variables, the flight parameter variables of the whole flight process of the carrier rocket are iterated based on the dynamic differential equation to obtain the flight parameter variables of the carrier rocket in each phase, and solve the gravity y, the flight Mach number η, and the range R of the corresponding phase;

[0044] Mapping module: Draw the flight parameter variables, gravity y, flight Mach number η and range R of the launch vehicle at each stage, and draw the map of the flight parameter variables, gravity y, flight Mach number η and range R of the launch vehicle.

[0045] Compared with the prior art, the present invention has the following beneficial effects:

[0046] 1. The present invention fully considers the engine swinging and flipping attitude stage during the entire flight process of the rocket, and at the same time iteratively analyzes the flight parameter variables in the attitude adjustment stage before re-entering the atmosphere, thereby reducing the error between the flight parameter variables obtained by simulation and the results obtained in the actual test run, making the simulation results more practical and efficient.

[0047] 2. The full-flight process simulation method of a reusable launch vehicle of the present invention establishes a reasonable dynamic simulation model and integrates the various processes of the launch vehicle's engine swinging, flipping and attitude adjustment before re-entering the atmosphere, making it more convenient to divide the stages of the launch vehicle in the return phase and effectively saving the time for large-scale iterative calculations of the attitude adjustment segment, deceleration turning segment and gliding segment of the launch vehicle in the attitude adjustment state.

[0048] 3. The method for simulating the entire flight process of a reusable launch vehicle of the present invention allows the user to freely set the thrust of the rocket body to calculate the flight parameters of the launch phase and the recovery phase, thereby simulating the launch and recovery effects of rockets of different magnitudes, so that the simulation results have the characteristics of a wide range of applications and a wide range of applicable rocket body types.

[0049] 4. The simulation method of the present invention performs iterative calculations through flight parameters, which can ensure that other parts remain unaffected after the sub-kinematic model of a certain process is optimized and modified. It has the characteristics of being easy to modify and having a wide range of adaptability to kinematic models.

[0050] 5. The simulation method of the present invention can simulate various flight conditions and environmental changes, including different atmospheric conditions, gravity effects, etc.; by simulating different flight conditions, the performance of the launch vehicle can be more comprehensively evaluated, the launch vehicle design can be optimized to adapt to various environments, and the reliability and performance of the launch vehicle can be improved.

[0051] 6. The simulation method of the present invention provides a large amount of flight parameter data, including flight altitude, flight speed, flight Mach number, etc., to provide data support for decision-making. By analyzing the simulation data, design decisions can be made more scientifically, thereby improving the accuracy and reliability of the design.

[0052] In summary, the present invention divides the entire flight process of a carrier rocket into seven stages, and iteratively analyzes the flight parameter variables in each stage according to the dynamic differential equation. This can not only accurately and quickly calculate and evaluate the performance parameters of a reusable carrier rocket at different stages, but also can more comprehensively and detailedly reflect the motion conditions of the entire process of launch and recovery of the carrier rocket. It has the characteristics of high precision of motion models and data, and is easy for users to distinguish different motion states of the rocket body and make corresponding decisions. At the same time, it reduces the number and cost of actual test runs, shortens the actual development cycle, and improves development efficiency. BRIEF DESCRIPTION OF THE DRAWINGS

[0053] Figure 1 A flow chart of a method for simulating the entire flight process of a reusable launch vehicle provided by the present invention.

[0054] Figure 2 A range-time diagram of the entire flight process of the carrier rocket provided by the present invention.

[0055] Figure 3 A gravity-time diagram during the entire flight process of the carrier rocket provided by the present invention.

[0056] Figure 4 A diagram of the horizontal distance from the launch site and the flight altitude during the entire flight of the carrier rocket provided by the present invention.

[0057] Figure 5 The present invention provides an aerodynamic drag-time diagram of the launch vehicle during the entire flight process.

[0058] Figure 6 The present invention provides a flight Mach number-time diagram of a carrier rocket during its entire flight process.

[0059] Figure 7 A flight speed-time diagram of the carrier rocket during the entire flight process provided by the present invention.

[0060] Figure 8 A first-stage mass-time diagram of a launch vehicle during the entire flight process provided by the present invention.

[0061] Fig. 9 The invention provides a flight path inclination-time diagram of the carrier rocket during the entire flight process.

[0062] Fig.10 A flight altitude-time diagram during the entire flight process of the carrier rocket provided by the present invention. DETAILED DESCRIPTION

[0063] The technical solution of the present invention is further described below in conjunction with the accompanying drawings.

[0064] like Figure 1 As shown, a method for simulating the entire flight process of a reusable launch vehicle comprises the following steps:

[0065] Step 1: Set and initialize the flight parameter variables of the launch vehicle during the entire flight process, wherein the flight parameter variables include the flight speed v, the flight path inclination γ, the flight altitude h, the horizontal position x between the launch vehicle and the ground, and the launch vehicle mass m;

[0066] In this embodiment, the initial overall mass of the launch vehicle is m 0 , flight path starting inclination Starting height h 0 =0, starting position x 0 =0, initial speed v 0 =0, start time t 0 =0;

[0067] Step 2: According to the flight parameter variables initialized in step 1, set the running time t of the ascent phase before gravity steering1 According to the dynamic differential equation, the flight parameter variables of the launch vehicle in the ascent phase before gravity turning and the corresponding gravity y, Mach number η and range R of this phase are solved until the running time t of this phase is reached. 1 ;

[0068] The ascending phase before the gravity turn is from the launch ignition of the carrier rocket to the start of the gravity turn, and its main purpose is to accelerate the carrier rocket to a sufficient speed and change its flight direction so that it enters the predetermined orbit;

[0069] First, the aerodynamic drag D of the launch vehicle in the ascent phase before gravity turns is solved as:

[0070] D=0.5·A·ρ·C d ·V 2

[0071] Among them, A is the windward area of ​​the launch vehicle at the current moment, ρ is the air density, C d is the air resistance coefficient, V is the speed of the launch vehicle relative to the air at the current moment;

[0072] The thrust of the launch vehicle engine in the ascending stage before gravity turns is T = T 1 , T 1 is the set launch vehicle engine thrust;

[0073] The aerodynamic drag D of the launch vehicle in the ascent phase before gravity is turned, and the running time t of the ascent phase before gravity is turned is set 1 and the launch vehicle engine thrust T = T 1 Substituting into the dynamic differential equation, the flight parameter variables of the launch vehicle in the ascent phase before gravity turns are calculated, and the expression is as follows:

[0074]

[0075] Among them, γ i-1 is the flight path inclination angle at the previous moment, γ i is the flight path inclination at the current moment, R e is the radius of the earth, T is the thrust of the launch vehicle engine at the current stage, I sp is the specific impulse of the launch vehicle engine at the current stage, g 0 is the initial acceleration of the launch vehicle at the current stage, g is the gravitational acceleration of the launch vehicle at the current moment, m is the mass of the launch vehicle at the current moment; D is the aerodynamic drag of the launch vehicle at the current moment, and V is the speed of the launch vehicle relative to the air at the previous moment;

[0076] In this embodiment, the operating time t of the carrier rocket in the ascending stage before gravity turns 1is 17.5s, and the thrust of the launch vehicle engine is set to T 1 is 2205000N; the flight path end inclination angle γ1 is 89°;

[0077] Step 3: Use the flight parameter variables of the ascent phase before the gravity is turned in step 2 as the starting variables of the first-stage ascent phase, iterate the flight parameter variables of the launch vehicle in the first-stage ascent phase, until the fuel set for the entire first-stage ascent phase is exhausted, separate the excess mass of the launch vehicle, and obtain the flight parameter variables of the launch vehicle in the first-stage ascent phase and the gravity y, Mach number η and range R corresponding to this phase;

[0078] The first-stage ascent phase is from the launch ignition of the carrier rocket to the shutdown and separation of the first-stage engine. After the carrier rocket is launched, the first-stage engine is ignited, and the carrier rocket begins to climb vertically and quickly pass through the dense lower atmosphere to reduce the impact of air resistance on the rocket. During this phase, the angle of attack of the carrier rocket is close to zero, and the vertical ascent is mainly achieved by relying on the thrust of the engine. As the carrier rocket increases in altitude and accumulates in speed, the carrier rocket gradually adjusts its flight attitude and transitions from vertical climb to horizontal flight to prepare for orbit entry. After reaching a certain altitude and speed, the carrier rocket begins to perform gravity steering, using the gravity of the earth to change the flight direction, from vertical or inclined ascent to horizontal flight, to enter the predetermined orbit. After completing the above phases, the first-stage engine will be shut down and then separated from the rest of the rocket.

[0079] First, the aerodynamic drag D of the launch vehicle in the first stage of ascent is solved as:

[0080] D=0.5·A·ρ·C d ·V 2

[0081] Among them, A is the windward area of ​​the launch vehicle at the current moment, ρ is the air density, C d is the air resistance coefficient, V is the speed of the launch vehicle relative to the air at the current moment;

[0082] The thrust of the launch vehicle engine in the first stage of ascent is T = T 1 , T 1 is the set launch vehicle engine thrust;

[0083] Calculate the running time t of the first stage ascent phase according to the fuel flow rate of the launch vehicle 2 , the expression is as follows:

[0084]

[0085] Where, mp is the total propellant mass of the launch vehicle before the first ascent phase begins, mf is the remaining propellant mass of the launch vehicle, g0 is the initial acceleration of the launch vehicle in the first stage of ascent, T is the thrust of the launch vehicle engine in the current stage, I sp is the specific impulse of the launch vehicle engine at the current stage;

[0086] The aerodynamic drag D of the launch vehicle in the first stage of ascent and the running time t of the first stage of ascent are set. 2 and the launch vehicle engine thrust T = T 1 Substituting into the dynamic differential equation, the flight parameter variables of the first-stage ascent launch vehicle are calculated, and the expression is as follows:

[0087]

[0088] When the fuel set for the first stage of ascent is exhausted, the excess mass of the launch vehicle is separated, that is, the first stage of the launch vehicle is retained, and its second stage and third stage are separated. The mass m of the launch vehicle after separation is as follows:

[0089] m=m 1 +m 2

[0090] Among them, m 1 is the dry mass of the first stage rocket, m 2 is the weight of the remaining fuel;

[0091] Step 4: Use the flight parameter variables of the first stage ascent phase in step 3 as the starting variables of the pre-reentry phase and set the running time t of this phase 3 , iterate the flight parameter variables of the launch vehicle in the pre-reentry stage until the running time t set for this stage is reached 3 , obtain the flight parameter variables of the launch vehicle in the pre-reentry stage and the gravity y, Mach number η and range R corresponding to this stage;

[0092] The pre-reentry stage is the process from the completion of the separation process of the first-stage rocket to the preparation for the engine swing and flip attitude adjustment. In this stage, the first-stage rocket continues to rise due to inertia, but due to the thrust adjustment, the speed of the carrier rocket decreases, thereby preparing for the engine swing and flip attitude adjustment.

[0093] First, the aerodynamic drag D of the launch vehicle before re-entry into the atmosphere is solved as:

[0094] D=0.5·A·ρ·C d ·V 2

[0095] Among them, A is the windward area of ​​the launch vehicle at the current moment, ρ is the air density, C d is the air resistance coefficient, V is the speed of the launch vehicle relative to the air at the current moment;

[0096] The thrust of the launch vehicle engine in the pre-reentry stage T 1 is the set launch vehicle engine thrust;

[0097] Set the aerodynamic drag D of the launch vehicle before re-entry and the running time t of the pre-re-entry stage 3 and launch vehicle engine thrust Substituting into the dynamic differential equation, the flight parameter variables of the launch vehicle before re-entry are calculated, and the expression is as follows:

[0098]

[0099] In this embodiment, the running time t before re-entry into the atmosphere 3 13s;

[0100] Step 5: The flight parameter variables in the pre-reentry stage in step 4 are used as the starting variables in the engine swing and flip attitude stage, and the flight parameter variables of the launch vehicle in the engine swing and flip attitude stage are iterated until the flight path inclination angle of this stage reaches the reentry angle, and the flight parameter variables of the launch vehicle in the engine swing and flip attitude stage and the gravity y, Mach number η and range R corresponding to this stage are obtained;

[0101] The engine swinging and flipping attitude stage is from the second start of the first-stage engine of the carrier rocket to the second shutdown of the engine. The main purpose is to change the speed of the rocket body in the X-axis direction by relying on the engine thrust, so that the rocket enters the dense atmosphere in a "head-up" attitude after the attitude adjustment;

[0102] First, the aerodynamic drag D of the launch vehicle during the engine swing and flip attitude stage is solved as:

[0103] D=0.5·A·ρ·C d ·V 2

[0104] Among them, A is the windward area of ​​the launch vehicle at the current moment, ρ is the air density, C d is the air resistance coefficient, V is the speed of the launch vehicle relative to the air at the current moment;

[0105] The thrust of the launch vehicle engine in the engine swinging and flipping posture stage is T=0;

[0106] The aerodynamic drag D of the launch vehicle in the engine swing and flip attitude stage is set, and the running time t in the engine swing and flip attitude stage is set 4Substitute the thrust of the launch vehicle engine T = 0 into the dynamic differential equation to calculate the flight parameter variables of the launch vehicle during the engine swing and flip attitude stage. The expression is as follows:

[0107]

[0108] The reentry angle γ 3 The pitch angle of the engine swing and flip attitude set before re-entry into the atmosphere; the flight parameter variables of the launch vehicle in this stage are iterated until the running time of this stage is reached t 4 And the flight path inclination reaches the reentry angle γ 3 ;

[0109] In this embodiment, the running time t of the engine swing and flip posture stage 4 is 6s; the reentry angle γ 3 for The initial flight path inclination angle γ at this stage 2 is the last flight path inclination angle iterated in the flight path inclination change rate formula in the pre-reentry stage;

[0110] Step 6: Use the flight parameter variables of the engine swing and flip attitude stage in step 5 as the starting variables of the re-entry stage, and set the running time t of the re-entry stage 5 , iterate the flight parameter variables of the launch vehicle in the re-entry phase until the running time t set for this phase is reached 5 , obtain the flight parameter variables of the launch vehicle in the atmospheric reentry phase and the gravity y, Mach number η and range R corresponding to this phase;

[0111] The re-entry phase is from the second shutdown of the first-stage engine of the launch vehicle to the third start-up of the engine. During this process, the rocket body re-enters the atmosphere in an unpowered state and is aerodynamically decelerated under the action of the dense atmosphere. At the same time, due to the action of the atmosphere, the inclination angle of the flight path of the launch vehicle increases.

[0112] The aerodynamic drag D of the launch vehicle during the re-entry phase is solved as:

[0113] D=0.5·A·ρ·C d ·V 2

[0114] Among them, A is the windward area of ​​the launch vehicle at the current moment, ρ is the air density, C d is the air resistance coefficient, V is the speed of the launch vehicle relative to the air at the current moment;

[0115] The thrust of the launch vehicle engine at the re-entry stage is T=0;

[0116] The aerodynamic drag D of the launch vehicle during the re-entry phase is set as the running time t 5 Substitute the thrust of the launch vehicle engine T = 0 into the dynamic differential equation to calculate the flight parameter variables of the launch vehicle during the re-entry phase. The expression is as follows:

[0117]

[0118] In this embodiment, the running time t used in the re-entry phase is 5 83s;

[0119] Step 7: Use the flight parameter variables of the reentry phase in step 6 as the starting variables of the vertical attitude adjustment phase and set the running time t of this phase. 6 , iterate the flight parameter variables of the launch vehicle in the vertical attitude adjustment phase until the running time t set for this phase is reached 6 , obtain the flight parameter variables of the launch vehicle in the vertical attitude adjustment stage and the gravity y, Mach number η and range R corresponding to this stage;

[0120] The vertical attitude adjustment phase is from the third start of the launch vehicle's first-stage engine to the reduction of the rocket's horizontal speed to 0. The main purpose is to rely on the engine thrust to change the rocket's X-axis speed and adjust the landing point of the first-stage so that the requirements for fixed-point landing and recovery are met after the attitude adjustment is completed.

[0121] First, the aerodynamic drag D of the launch vehicle in the vertical attitude adjustment phase is solved as:

[0122] D=0.5·A·ρ·C d ·V 2

[0123] Among them, A is the windward area of ​​the launch vehicle at the current moment, ρ is the air density, C d is the air resistance coefficient, V is the speed of the launch vehicle relative to the air at the current moment;

[0124] The thrust of the launch vehicle engine in the vertical attitude adjustment stage T 1 is the set launch vehicle engine thrust;

[0125] The aerodynamic drag D of the launch vehicle in the vertical attitude adjustment phase and the running time t in the vertical attitude adjustment phase are set. 6 and launch vehicle engine thrust Substituting into the dynamic differential equation, the flight parameter variables of the launch vehicle in the vertical attitude adjustment phase are calculated, and the expression is as follows:

[0126]

[0127] In this embodiment, the running time of the vertical posture adjustment stage is 10s;

[0128] Step 8: Use the flight parameter variables of the vertical attitude adjustment phase in step 7 as the starting variables of the powered descent phase, iterate the flight parameter variables of the launch vehicle in the powered descent phase until the flight altitude h≤0, and obtain the flight parameter variables of the launch vehicle in the powered descent phase and the gravity y, Mach number η and range R corresponding to this phase;

[0129] The power descent stage is from the rocket body horizontal speed decreasing to 0 to the first stage landing, flying at a prescribed 180° inclination angle according to the rocket reverse thrust deceleration requirements, and the engine performs reverse thrust in the Y-axis direction until the rocket body speed drops below the landing required speed. In this embodiment, the landing required speed is about 0;

[0130] The aerodynamic drag D of the launch vehicle during the powered descent phase is solved as:

[0131] D=0.5·A·ρ·C d ·V 2

[0132] Among them, A is the windward area of ​​the launch vehicle at the current moment, ρ is the air density, C d is the air resistance coefficient, V is the speed of the launch vehicle relative to the air at the current moment;

[0133] The thrust of the launch vehicle engine in the power descent phase Among them, C t is the thrust coefficient; h is the height of the launch vehicle at the current moment, in meters, and y is the gravity of the launch vehicle at the current moment; set the running time t of the powered descent phase 7 ;

[0134] The aerodynamic drag D of the launch vehicle in the powered descent phase and the running time t in the powered descent phase are set. 7 and launch vehicle engine thrust Substituting into the dynamic differential equation,

[0135] Calculate the flight parameter variables of the launch vehicle during the powered descent phase, and the expression is as follows:

[0136]

[0137] In this embodiment, the running time t of the power reduction stage 7 is 60s, thrust coefficient C t is 0.99999;

[0138] Step 9: Based on the flight parameter variables, gravity y and Mach number η obtained in each stage of steps 2-8, plot the relationship between the corresponding variables and the flight time, as shown in Figure 2-Figure 10 .

[0139] The expressions of gravity y and flight Mach number η at each stage during the entire flight of the launch vehicle are as follows:

[0140]

[0141] Among them, M e is the mass of the Earth, a is the speed of sound of the launch vehicle at the current flight altitude h, v is the flight speed of the launch vehicle, G is the gravitational constant, R e is the radius of the earth, m is the mass of the launch vehicle at the current moment;

[0142] The expression of the range R of each stage during the whole flight process of the launch vehicle is as follows:

[0143]

[0144] Among them, l k The length of the flight time t vector solved for each stage, a k is the upper limit of the running time t variable subscript in the function of the flight R at the current moment, v i is the speed of the launch vehicle at the current moment, t i is the current moment, t i-1 For the previous moment.

[0145] The entire flight process of the carrier rocket is divided into a launch phase and a recovery phase. The launch phase includes the above-mentioned ascent phase before gravity turning, the first-stage ascent phase and the pre-reentry phase. The recovery phase includes the above-mentioned engine swing and flip attitude phase, the reentry phase, the vertical attitude adjustment phase and the powered descent phase.

[0146] The present invention also provides a full flight process simulation system for a reusable launch vehicle, comprising:

[0147] Initialization module: used to set and initialize the flight parameter variables of the launch vehicle during the entire flight process;

[0148] The calculation module is used to divide the whole flight process of the carrier rocket into a launch phase and a recovery phase, wherein the launch phase includes an ascent phase before gravity turning, a first-stage ascent phase, and a pre-reentry phase; the recovery phase includes an engine swing and flip attitude phase, a reentry phase, a vertical attitude adjustment phase, and a power descent phase; according to the initialized flight parameter variables, the flight parameter variables of the whole flight process of the carrier rocket are iterated based on the dynamic differential equation to obtain the flight parameter variables of the carrier rocket in each phase, and solve the gravity y, the flight Mach number η, and the range R of the corresponding phase;

[0149] Mapping module: Draw the flight parameter variables, gravity y, flight Mach number η and range R of the launch vehicle at each stage, and draw the map of the flight parameter variables, gravity y, flight Mach number η and range R of the launch vehicle.

Claims

1. A method for simulating the entire flight process of a reusable launch vehicle, characterized in that: The steps include: Step 1: Set and initialize the flight parameter variables of the launch vehicle during the entire flight process, wherein the flight parameter variables include the flight speed v, the flight path inclination γ, the flight altitude h, the horizontal position x between the launch vehicle and the launch site, and the launch vehicle mass m; Step 2: Divide the entire flight process of the carrier rocket into a launch phase and a recovery phase, wherein the launch phase includes an ascent phase before gravity turning, a first-stage ascent phase, and a pre-reentry phase; the recovery phase includes an engine swing and flip attitude phase, a reentry phase, a vertical attitude adjustment phase, and a powered descent phase; according to the flight parameter variables initialized in step 1, the flight parameter variables of the carrier rocket in each flight phase are iterated based on the dynamic differential equation to obtain the flight parameter variables of the carrier rocket in each phase, and solve the gravity y, flight Mach number η, and range R of the corresponding phase; Step 3: Draw a graph of the launch vehicle's flight parameter variables, gravity y, flight Mach number η, and range R based on the flight parameter variables, gravity y, flight Mach number η, and range R of each stage obtained in step 2.

2. A method for simulating the full flight process of a reusable launch vehicle according to claim 1, characterized in that: The process of iterating the flight parameter variables of the launch vehicle at all stages in step 2 is as follows: Step 2.1: According to the flight parameter variables initialized in step 1, set the running time t1 of the ascent phase before gravity turning, and solve the flight parameter variables of the launch vehicle in the ascent phase before gravity turning and the gravity y, Mach number η and range R corresponding to this phase according to the dynamic differential equation until the running time t1 of this phase is reached; Step 2.2: Use the flight parameter variables of the ascent phase before gravity is turned in step 2.1 as the starting variables of the first-stage ascent phase, iterate the flight parameter variables of the launch vehicle in the first-stage ascent phase until the fuel set for the first-stage ascent phase is exhausted, separate the excess mass of the launch vehicle, and obtain the flight parameter variables of the launch vehicle in the first-stage ascent phase and the gravity y, Mach number η, and range R corresponding to this phase; Step 2.3: Use the flight parameter variables of the first-stage ascent stage in step 2.2 as the starting variables of the pre-reentry stage, set the running time t3 of this stage, iterate the flight parameter variables of the launch vehicle in the pre-reentry stage until the running time t3 set for this stage is reached, and obtain the flight parameter variables of the launch vehicle in the pre-reentry stage and the gravity y, Mach number η and range R corresponding to this stage; Step 2.4: Use the flight parameter variables of the pre-reentry stage in step 2.3 as the starting variables of the engine swing and flip attitude stage, iterate the flight parameter variables of the launch vehicle in the engine swing and flip attitude stage until the flight path inclination angle of this stage reaches the reentry angle, and obtain the flight parameter variables of the launch vehicle in the engine swing and flip attitude stage and the gravity y, Mach number η and range R corresponding to this stage; Step 2.5: Use the flight parameter variables of the engine swing and flip attitude stage in step 2.4 as the starting variables of the re-entry stage, set the running time t5 of the re-entry stage, iterate the flight parameter variables of the launch vehicle in the re-entry stage until the running time t5 set for this stage is reached, and obtain the flight parameter variables of the launch vehicle in the re-entry stage and the gravity y, Mach number η and range R corresponding to this stage; Step 2.6: Use the flight parameter variables of the reentry phase in step 2.5 as the starting variables of the vertical attitude adjustment phase, set the running time t6 of the vertical attitude adjustment phase, iterate the flight parameter variables of the launch vehicle in the vertical attitude adjustment phase until the running time t6 set for this phase is reached, and obtain the flight parameter variables of the launch vehicle in the vertical attitude adjustment phase and the gravity y, Mach number η and range R corresponding to this phase; Step 2.7: Use the flight parameter variables of the vertical attitude adjustment phase in step 2.6 as the starting variables of the powered descent phase, iterate the flight parameter variables of the launch vehicle in the powered descent phase until the flight altitude h≤0, and obtain the flight parameter variables of the launch vehicle in the powered descent phase and the gravity y, Mach number η and range R corresponding to this phase.

3. The method for simulating the full flight process of a reusable launch vehicle according to claim 2, characterized in that: The iterative equations of the flight parameter variables at each stage of the launch vehicle's full flight process in step 2 are as follows: Among them, S1 represents the ascent stage before gravity turns, S2 represents the first-stage ascent stage, S3 represents the stage before re-entry, S4 represents the engine swing and flip attitude stage, S5 represents the re-entry stage, S6 represents the vertical attitude adjustment stage, and S7 represents the power descent stage; T is the thrust of the launch vehicle engine at the current stage, D is the aerodynamic drag of the launch vehicle at the current moment, m is the mass of the launch vehicle at the current moment, g is the gravity acceleration of the launch vehicle, and R e is the radius of the earth, h is the flight altitude of the launch vehicle at the current moment, V is the speed of the launch vehicle relative to the air at the current moment, γ i-1 is the flight path inclination angle at the previous moment, γ i is the inclination of the flight path at the current moment, v is the flight speed of the launch vehicle at the current moment, I sp is the specific impulse of the launch vehicle engine at the current stage, and g0 is the initial acceleration of the launch vehicle.

4. The method for simulating the full flight process of a reusable launch vehicle according to claim 3, characterized in that: In step 2.2, the running time t2 of the first stage ascent phase is calculated according to the fuel flow rate of the launch vehicle, and the expression is as follows: Among them, mp is the total propellant mass of the launch vehicle before the start of the first stage ascent, mf is the remaining propellant mass of the launch vehicle, g0 is the initial acceleration of the launch vehicle in the first stage ascent, T is the thrust of the launch vehicle engine at the current stage, I sp It is the specific impulse of the launch vehicle engine at the current stage.

5. The method for simulating the full flight process of a reusable launch vehicle according to claim 3, characterized in that: When the fuel set for the first stage ascent in step 2.2 is exhausted, the excess mass of the carrier rocket is separated, that is, the first stage of the carrier rocket is retained, and its second stage and third stage are separated. The mass m of the carrier rocket after separation is as follows: m=m1+m2 Among them, m1 is the dry weight of the first-stage rocket and m2 is the weight of the remaining fuel.

6. The method for simulating the full flight process of a reusable launch vehicle according to claim 3, characterized in that: The aerodynamic drag D of each stage during the entire flight of the launch vehicle is: D=0.5·A·ρ·C d ·V 2 Among them, A is the windward area of ​​the launch vehicle at the current moment, ρ is the air density, C d is the air resistance coefficient, and V is the speed of the launch vehicle relative to the air at the current moment.

7. The method for simulating the full flight process of a reusable launch vehicle according to claim 3, characterized in that: The thrust T of the launch vehicle engine at each stage during the entire flight of the launch vehicle is as follows: Among them, S1 represents the ascent stage before gravity turning, S2 represents the first-stage ascent stage, S3 represents the stage before re-entry, S4 represents the engine swing and flip attitude stage, S5 represents the re-entry stage, S6 represents the vertical attitude adjustment stage, and S7 represents the power descent stage; T1 is the set launch vehicle engine thrust, C t is the thrust coefficient; h is the height of the launch vehicle in meters; y is the gravity of the launch vehicle at the current moment.

8. The method for simulating the full flight process of a reusable launch vehicle according to claim 1 or 2, characterized in that: The gravity y and the flight Mach number η at each stage of the whole flight process of the carrier rocket are: Among them, M e is the mass of the Earth, a is the speed of sound of the launch vehicle at the current flight altitude h, v is the flight speed of the launch vehicle at the current moment, G is the gravitational constant, R e is the radius of the earth, m is the mass of the launch vehicle at the current moment; The range R of each stage during the entire flight of the launch vehicle is: Among them, l k The length of the flight time t vector solved for each stage, a k is the upper limit of the flight time t variable subscript in the function of the flight range R at the current moment, v i is the flight speed of the launch vehicle at the current moment, t i is the current moment, t i-1 For the previous moment.

9. A full flight process simulation system for a reusable launch vehicle, characterized in that: include: Initialization module: used to set and initialize the flight parameter variables of the launch vehicle during the entire flight process; The calculation module is used to divide the whole flight process of the carrier rocket into a launch phase and a recovery phase, wherein the launch phase includes an ascent phase before gravity turning, a first-stage ascent phase, and a pre-reentry phase; the recovery phase includes an engine swing and flip attitude phase, a reentry phase, a vertical attitude adjustment phase, and a power descent phase; according to the initialized flight parameter variables, the flight parameter variables of the whole flight process of the carrier rocket are iterated based on the dynamic differential equation to obtain the flight parameter variables of the carrier rocket in each phase, and solve the gravity y, the flight Mach number η, and the range R of the corresponding phase; Mapping module: Draw the flight parameter variables, gravity y, flight Mach number η and range R of the launch vehicle at each stage, and draw the map of the flight parameter variables, gravity y, flight Mach number η and range R of the launch vehicle.

Images