A reusable launch vehicle full flight simulation method and system

Abstract

The application discloses a reusable carrier rocket full flight process simulation method and system, wherein the full flight process of the carrier rocket is subdivided into seven stages, namely, an ascending stage before gravity turning, a first-stage ascending stage, a pre-atmospheric reentry stage, an engine swing and roll attitude stage, an atmospheric reentry stage, a vertical attitude adjustment stage and a powered descent stage, then flight parameter variables in each stage are iteratively analyzed according to a dynamic differential equation, so that the performance parameters of the reusable carrier rocket in different stages can be accurately and quickly calculated and evaluated, and the motion of the full flight process of the reusable carrier rocket can be fully reflected, and the actual test driving condition is more fitted.

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 carrier rocket. Background Art

[0002] With the increasing frequency of space activities, there is an urgent need for low-cost, high-efficiency launch vehicles. Reusable rockets have become a key research direction in the aerospace field due to their ability to reduce launch costs and improve launch efficiency. Currently, the application of reusable rocket technology has penetrated into various fields, the most notable of which is SpaceX's Starlink program, which launches thousands of satellites using reusable Falcon 9 rockets to achieve global internet coverage.

[0003] The development of reusable rocket technology faces numerous technical challenges, including precise return flight and safe landing control, variable-thrust reusable engines, and highly reliable landing cushioning mechanisms. While the maturity of these technologies requires extensive experimentation, the increased number of tests also leads to a corresponding increase in failures, resulting in significant economic losses. For example, during a ground test in September 2016, a Falcon 9 rocket exploded, destroying both the rocket and the AMOS-6S communications satellite it was carrying. Therefore, simulating the entire flight process of reusable rockets is crucial to reducing the costs of actual test runs and failures.

[0004] Predicting and optimizing the performance of reusable launch vehicles requires precise computer programming and simulation analysis of the rocket's aerodynamic parameters. This involves determining the numerous parameters involved in the rocket's motion, which requires extensive theoretical calculations. Using computer programming to determine these parameters is crucial for rapidly predicting and optimizing the performance of reusable rocket engines, effectively shortening the development cycle. This not only improves design efficiency but also allows for testing various extreme conditions in a virtual environment, mitigating the risks of actual test runs, while also reducing environmental pollution and accelerating technological development and iteration.

[0005] The Chinese invention patent application with the publication number CN106021628B proposed a method for designing the vertical return trajectory of a carrier rocket. The method established a three-degree-of-freedom dynamic model of the vertical return of the first stage of the carrier rocket and considered the flight overload, dynamic pressure and heat flux peak constraints, and terminal constraints. However, the method 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 carrier rocket during the entire flight process. Yu Guangxue et al. (Yu Guangxue, Li Zhaotang, Lin Ping. Mathematical modeling of reusable carrier rocket 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 carrier rocket (RLV) reentry. The method established an RLV aerodynamic model by establishing a reaction control system / aerodynamic surface composite control mathematical model and an engineering calculation method based on aerodynamic forces. However, the method lacked the description and calculation of the rocket body performance parameters and could not relatively accurately reflect the motion of the reusable carrier rocket 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 carrier rocket. By subdividing the entire flight process of the carrier rocket into seven stages and iteratively analyzing the flight parameter variables in each stage according to the dynamic differential equation, not only can the performance parameters of the reusable carrier rocket at different stages be accurately and quickly calculated and evaluated, but also the motion conditions of the reusable carrier rocket during the entire flight process can be fully reflected.

[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. The flight parameter variables include flight speed v, flight path inclination γ, flight altitude h, horizontal position x between the launch vehicle and the launch site, and 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, an atmospheric 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: Based on the flight parameter variables, gravity y, flight Mach number η, and range R of the launch vehicle at each stage obtained in step 2, draw a graph of the flight parameter variables, gravity y, flight Mach number η, and range R of the launch vehicle.

[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: Based on the flight parameter variables initialized in step 1, set the running time t1 of the ascent phase before gravity reversal. Solve the flight parameter variables of the launch vehicle in the ascent phase before gravity reversal and the corresponding gravity y, Mach number η, and range R according to the dynamic differential equation until the running time t1 of this phase is reached;

[0014] Step 2.2: Use the flight parameter variables of the ascent phase before gravity shift 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 to obtain the flight parameter variables of the launch vehicle in the first-stage ascent phase, as well as 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, set the running time t3 of this phase, and iterate the flight parameter variables of the launch vehicle in the pre-reentry phase until the set running time t3 of this phase is reached. Then, obtain the flight parameter variables of the launch vehicle in the pre-reentry phase and the corresponding gravity y, Mach number η, and range R.

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

[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, set the running time t5 of the re-entry phase, and iterate the flight parameter variables of the launch vehicle in the re-entry phase until the set running time t5 is reached. Then, obtain the flight parameter variables of the launch vehicle in the re-entry phase and the corresponding gravity y, Mach number η, and range R;

[0018] Step 2.6: Use the flight parameter variables of the re-entry 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, and iterate the flight parameter variables of the launch vehicle in the vertical attitude adjustment phase until the set running time t6 is reached. Then, obtain the flight parameter variables of the launch vehicle in the vertical attitude adjustment phase and the corresponding gravity y, Mach number η, and range R;

[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 for 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 powered 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 gravitational acceleration of the launch vehicle, and R e is the radius of the earth, h is the flight altitude of the carrier rocket at the current moment, V is the speed of the carrier rocket 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 flight path inclination at the current moment, v is the flight speed of the carrier rocket 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.

[0023] In step 2.2, the running time t2 of the first stage ascent phase is calculated based on the fuel flow rate of the launch vehicle, and the expression is as follows:

[0024]

[0025] Where, mp is the total propellant mass of the carrier rocket before the start of the first-stage ascent phase, mf is the remaining propellant mass of the carrier rocket, g0 is the initial acceleration of the carrier rocket during the first-stage ascent phase, T is the thrust of the carrier rocket engine at the current phase, and 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 phase in step 2.2 is exhausted, the excess mass of the carrier rocket is separated, that is, the first-stage rocket of the carrier rocket is retained, and its second-stage rocket and third-stage rocket are separated. The mass m of the carrier rocket after separation is as follows:

[0027] m=m1+m2

[0028] Among them, m1 is the dry weight of the first-stage rocket and m2 is the weight of the remaining fuel.

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

[0030] D=0.5·A·ρ·C d ·V 2

[0031] Among them, A is the frontal 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 carrier rocket engine at each stage during the entire flight process of the carrier rocket 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 powered 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.

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

[0036]

[0037] Among them, M e is the mass of the Earth, a is the speed of sound of the carrier rocket at the current flight altitude h, v is the flight speed of the carrier rocket at the current moment, G is the gravitational constant, R e is the radius of the earth, and 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 carrier rocket is:

[0039]

[0040] Among them, lk 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 current range R, v i is the flight speed of the carrier rocket 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 carrier rocket, 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 entire flight process of the carrier rocket into a launch phase and a recovery phase. The launch phase includes an ascent phase before gravity diversion, 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 initialized flight parameter variables, the flight parameter variables of the carrier rocket throughout the entire flight process 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.

[0044] Mapping module: Draw a graph of the launch vehicle's flight parameter variables, gravity y, flight Mach number η, and range R based on the launch vehicle's flight parameter variables, gravity y, flight Mach number η, and range R at each stage.

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

[0046] 1. The present invention fully considers the engine swing and flip attitude stage during the entire rocket flight process, and at the same time iteratively analyzes the flight parameter variables in the attitude adjustment stage before re-entry into the atmosphere, 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 realistic and efficient.

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

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

[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 adaptable to a wide range of 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., which provides 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 not only can 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 The present invention provides a flow chart of the method for simulating the entire flight process of a reusable carrier rocket.

[0054] Figure 2 The present invention provides a range-time diagram of the entire flight process of the carrier rocket.

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

[0056] Figure 4 The present invention provides a diagram of the horizontal distance from the launch site to the flight altitude during the entire flight process of the carrier rocket.

[0057] Figure 5The aerodynamic drag-time diagram of the carrier rocket during the entire flight process provided by the present invention.

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

[0059] Figure 7 The present invention provides a flight speed-time diagram of the carrier rocket during its entire flight process.

[0060] Figure 8 The present invention provides a first-stage mass-time diagram of the carrier rocket during its entire flight process.

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

[0062] Figure 10 The present invention provides a flight altitude-time diagram of the carrier rocket during its entire flight process. DETAILED DESCRIPTION

[0063] The technical solution of the present invention will be further described below with reference to the accompanying drawings.

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

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

[0066] In this embodiment, the initial overall mass of the carrier rocket is m0, and the initial inclination angle of the flight path is Starting height h0 = 0, starting position x0 = 0, starting speed v0 = 0, starting time t0 = 0;

[0067] Step 2: Based on the flight parameter variables initialized in step 1, set the running time t1 of the ascent phase before gravity reversal, and solve the flight parameter variables of the launch vehicle in the ascent phase before gravity reversal and the corresponding gravity y, Mach number η, and range R according to the dynamic differential equation until the running time t1 of this phase is reached;

[0068] The ascent phase before gravity diversion is from the launch ignition of the carrier rocket to the start of gravity diversion, 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 during 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 frontal 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 carrier rocket engine during the ascent phase before gravity turns is T=T1, where T1 is the set thrust of the carrier rocket engine;

[0073] Substitute the aerodynamic drag D of the launch vehicle in the ascent phase before gravity turns, the assumed running time t1 of the ascent phase before gravity turns, and the launch vehicle engine thrust T=T1 into the dynamic differential equation to calculate the flight parameter variables of the launch vehicle in the ascent phase before gravity turns. 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, g0 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 t1 of the carrier rocket in the ascent phase before gravity diversion is 17.5s, the thrust T1 of the carrier rocket engine is set to 2205000N; the flight path end inclination angle γ1 is 89°;

[0077] Step 3: Use the flight parameter variables of the ascent phase before gravity is turned in step 2 as the starting variables of the first-stage ascent phase, iterate the flight parameter variables of the carrier rocket 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 carrier rocket, and obtain the flight parameter variables of the carrier rocket in the first-stage ascent phase, as well as 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, rapidly passing through the dense lower atmosphere to reduce the impact of air resistance on the rocket. During this phase, the carrier rocket's angle of attack is close to zero, and vertical ascent is achieved mainly 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, transitioning from vertical climb to horizontal flight in preparation for entering orbit. After reaching a certain altitude and speed, the carrier rocket begins to perform gravity turn, using the Earth's gravity to change its flight direction, transitioning from vertical or inclined ascent to horizontal flight, in order 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 during the first ascent phase is solved as follows:

[0080] D=0.5·A·ρ·C d ·V 2

[0081] Among them, A is the frontal 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 carrier rocket engine in the first stage of ascent is T=T1, where T1 is the set thrust of the carrier rocket engine;

[0083] The running time t2 of the first-stage ascent phase is calculated based on the fuel flow rate of the launch vehicle. The expression is as follows:

[0084]

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

[0086] Substitute the aerodynamic drag D of the launch vehicle in the first stage of ascent, the set running time t2 of the first stage of ascent, and the thrust of the launch vehicle engine T=T1 into the dynamic differential equation to calculate the flight parameter variables of the launch vehicle in the first stage of ascent. The expression is as follows:

[0087]

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

[0089] m=m1+m2

[0090] Where m1 is the dry weight of the first stage rocket and m2 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, set the running time t3 of this phase, iterate the flight parameter variables of the launch vehicle in the pre-reentry phase until the set running time t3 of this phase is reached, and obtain the flight parameter variables of the launch vehicle in the pre-reentry phase and the corresponding gravity y, Mach number η, and range R;

[0092] The pre-reentry phase 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. During this phase, the first-stage rocket continues to rise due to inertia, but due to 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 follows:

[0094] D=0.5·A·ρ·C d ·V 2

[0095] Among them, A is the frontal 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 T1 is the set launch vehicle engine thrust;

[0097] The aerodynamic drag D of the launch vehicle before re-entry, the running time t3 of the pre-re-entry stage and the thrust of the launch vehicle engine are set. Substituting into the dynamic differential equation, the flight parameter variables of the launch vehicle in the pre-reentry stage are calculated as follows:

[0098]

[0099] In this embodiment, the operating time t3 of the pre-reentry stage is 13s;

[0100] Step 5: Using the flight parameter variables in the pre-reentry phase in step 4 as the starting variables for the engine swing and flip attitude phase, iterate the flight parameter variables of the launch vehicle in the engine swing and flip attitude phase until the flight path inclination angle in this phase reaches the reentry angle, and obtain the flight parameter variables of the launch vehicle in the engine swing and flip attitude phase and the corresponding gravity y, Mach number η, and range R;

[0101] The engine swing and flip attitude phase is from the second start of the launch vehicle's first-stage engine to the second shutdown of the engine. The main purpose is to rely on the engine thrust to change the speed of the rocket body in the X-axis direction, 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 phase is solved as:

[0103] D=0.5·A·ρ·C d ·V 2

[0104] Among them, A is the frontal 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 carrier rocket engine in the engine swing and flip attitude stage is T=0;

[0106] Substitute the aerodynamic drag D of the launch vehicle in the engine swing and flip attitude stage, the set running time t4 in the engine swing and flip attitude stage, and the launch vehicle engine thrust T=0 into the dynamic differential equation to calculate the flight parameter variables of the launch vehicle in the engine swing and flip attitude stage. The expression is as follows:

[0107]

[0108] The reentry angle γ3 is the pitch angle of the engine swing and flip attitude set in the pre-reentry stage; the flight parameter variables of the launch vehicle in this stage are iterated until the running time t4 of this stage is reached and the flight path inclination angle reaches the reentry angle γ3;

[0109] In this embodiment, the running time t4 of the engine swing and flip attitude stage is 6s; the reentry angle γ3 is The initial flight path inclination angle γ2 in this stage is the last flight path inclination angle iterated in the flight path inclination rate of change 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, set the running time t5 of the re-entry stage, iterate the flight parameter variables of the carrier rocket in the re-entry stage until the running time t5 set for this stage is reached, and obtain the flight parameter variables of the carrier rocket in the re-entry stage and the gravity y, Mach number η, and range R corresponding to this stage;

[0111] The re-entry phase is from the second shutdown of the launch vehicle's first-stage engine to the third engine startup. During this process, the rocket body re-enters the atmosphere in an unpowered state and undergoes aerodynamic deceleration under the action of the dense atmosphere. At the same time, due to the action of the atmosphere, the inclination angle of the launch vehicle's flight path 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 frontal 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 during the re-entry phase is T=0;

[0116] Substitute the aerodynamic drag D of the launch vehicle during the re-entry phase, the assumed operating time t5 in the re-entry phase, and the launch vehicle engine thrust 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 t5 used in the atmospheric reentry phase is 83s;

[0119] Step 7: Use the flight parameter variables of the re-entry phase in step 6 as the starting variables of the vertical attitude adjustment phase, set the running time t6 of this 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;

[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 velocity to zero. The main purpose is to rely on the engine thrust to change the rocket's X-axis velocity 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 during the vertical attitude adjustment phase is solved as follows:

[0122] D=0.5·A·ρ·C d ·V 2

[0123] Among them, A is the frontal 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 carrier rocket engine in the vertical attitude adjustment stage T1 is the set launch vehicle engine thrust;

[0125] The aerodynamic drag D of the launch vehicle in the vertical attitude adjustment phase, the running time t6 in the vertical attitude adjustment phase, and the thrust of the launch vehicle engine are set to Substituting into the dynamic differential equation, the flight parameter variables of the launch vehicle during the vertical attitude adjustment phase are calculated 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 carrier rocket in the powered descent phase until the flight altitude h≤0, and obtain the flight parameter variables of the carrier rocket in the powered descent phase and the corresponding gravity y, Mach number η, and range R;

[0129] The powered descent phase is from the time when the horizontal velocity of the rocket body decreases to 0 to the landing of the first stage. The rocket flies 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 velocity drops below the required landing speed. In this embodiment, the required landing speed is approximately 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] Wherein, A is the current time of the launch vehicle windward area, p is the air density, C d is the air resistance coefficient, V is the current time of the launch vehicle relative to the air speed;

[0133] The launch vehicle engine thrust of the power descent phase Wherein, C t is the thrust coefficient; h is the current time of the launch vehicle height, unit is meter, y is the current time of the launch vehicle gravity; Set the running time t7 of the power descent phase;

[0134] The aerodynamic drag D of the launch vehicle in the power descent phase, the running time t7 set in the power descent phase and the launch vehicle engine thrust Substitute the dynamics differential equation,

[0135] The flight parameter variables of the launch vehicle in the power descent phase are calculated, and the expression is as follows:

[0136]

[0137] In this embodiment, the running time t7 of the power descent phase is 60s, the thrust coefficient C t is 0.99999;

[0138] Step 9: draw the corresponding variable and flight time graph by the flight parameter variables of each stage, gravity y and Mach number η obtained by step 2-8, as shown in Figure 2-Figure 10 .

[0139] The expression of the gravity y and the flight Mach number η of each stage of the launch vehicle in the whole flight process is as follows:

[0140]

[0141] Wherein, M e is the mass of the earth, a is the sound speed of the launch vehicle at the current flight height 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 time;

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

[0143]

[0144] Wherein, l k is the length of the flight time t vector solved in each stage, a k is the upper limit value of the running time t variable subscript in the function of the current time of the range R, v iV(t) is the velocity of the launch vehicle at the current time t i t is the current time i-1 t-1 is the previous time.

[0145] The full flight process of the launch vehicle is divided into a launch phase and a recovery phase, the launch phase includes the pre-gravity turn ascent phase, the first-stage ascent phase, and the pre-atmospheric reentry phase, and the recovery phase includes the engine swing flip attitude phase, the atmospheric reentry phase, the vertical attitude adjustment phase, and the powered descent phase.

[0146] The application further provides a full flight process simulation system of a reusable launch vehicle, comprising:

[0147] An initialization module is configured to set and initialize flight parameter variables in the full flight process of the launch vehicle.

[0148] A calculation module is configured to divide the full flight process of the launch vehicle into a launch phase and a recovery phase, the launch phase includes the pre-gravity turn ascent phase, the first-stage ascent phase, and the pre-atmospheric reentry phase, and the recovery phase includes the engine swing flip attitude phase, the atmospheric reentry phase, the vertical attitude adjustment phase, and the powered descent phase; according to the initialized flight parameter variables, the flight parameter variables in the full flight process of the launch vehicle are iterated based on a dynamic differential equation, flight parameter variables of the launch vehicle in each phase are obtained, and gravity y, flight Mach number η, and range R in the corresponding phase are solved.

[0149] A plotting module is configured to plot flight parameter variable, gravity y, flight Mach number η, and range R graphs of the launch vehicle according to the flight parameter variables, gravity y, flight Mach number η, and range R of the launch vehicle in each phase.

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. The flight parameter variables include flight speed v, flight path inclination γ, flight altitude h, horizontal position x between the launch vehicle and the launch site, and 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, an atmospheric 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: Based on the flight parameter variables, gravity y, flight Mach number η, and range R of the launch vehicle at each stage obtained in step 2, draw a graph of the flight parameter variables, gravity y, flight Mach number η, and range R of the launch vehicle.

2. The method for simulating the entire flight process of a reusable launch vehicle according to claim 1, wherein: The process of iterating the flight parameter variables of the launch vehicle at all stages in step 2 is as follows: Step 2.1: Based on the flight parameter variables initialized in step 1, set the running time t1 of the ascent phase before gravity reversal. Solve the flight parameter variables of the launch vehicle in the ascent phase before gravity reversal and the corresponding gravity y, Mach number η, and range R 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 shift 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 to obtain the flight parameter variables of the launch vehicle in the first-stage ascent phase, as well as 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 phase in step 2.2 as the starting variables of the pre-reentry phase, set the running time t3 of this phase, and iterate the flight parameter variables of the launch vehicle in the pre-reentry phase until the set running time t3 of this phase is reached. Then, obtain the flight parameter variables of the launch vehicle in the pre-reentry phase and the corresponding gravity y, Mach number η, and range R. Step 2.4: Use the flight parameter variables of the pre-reentry phase in step 2.3 as the starting variables of the engine swing and flip attitude phase, iterate the flight parameter variables of the launch vehicle in the engine swing and flip attitude phase until the flight path inclination angle of this phase reaches the reentry angle, and obtain the flight parameter variables of the launch vehicle in the engine swing and flip attitude phase and the corresponding gravity y, Mach number η, and range R; 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, set the running time t5 of the re-entry phase, and iterate the flight parameter variables of the launch vehicle in the re-entry phase until the set running time t5 is reached. Then, obtain the flight parameter variables of the launch vehicle in the re-entry phase and the corresponding gravity y, Mach number η, and range R; Step 2.6: Use the flight parameter variables of the re-entry 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, and iterate the flight parameter variables of the launch vehicle in the vertical attitude adjustment phase until the set running time t6 is reached. Then, obtain the flight parameter variables of the launch vehicle in the vertical attitude adjustment phase and the corresponding gravity y, Mach number η, and range R; 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. Obtain the flight parameter variables of the launch vehicle in the powered descent phase and the corresponding gravity y, Mach number η, and range R.

3. The method for simulating the full flight process of a reusable carrier rocket according to claim 2, characterized in that: The iterative equations for 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 powered 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 gravitational acceleration of the launch vehicle, and R e is the radius of the earth, h is the flight altitude of the carrier rocket at the current moment, V is the speed of the carrier rocket 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 flight path inclination at the current moment, v is the flight speed of the carrier rocket 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 carrier rocket according to claim 3, characterized in that: In step 2.2, the running time t2 of the first stage ascent phase is calculated based on the fuel flow rate of the launch vehicle, and the expression is as follows: Where, mp is the total propellant mass of the carrier rocket before the start of the first-stage ascent phase, mf is the remaining propellant mass of the carrier rocket, g0 is the initial acceleration of the carrier rocket during the first-stage ascent phase, T is the thrust of the carrier rocket engine at the current phase, and 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 carrier rocket according to claim 3, characterized in that: When the fuel set for the first-stage ascent phase in step 2.2 is exhausted, the excess mass of the carrier rocket is separated, that is, the first-stage rocket of the carrier rocket is retained, and its second-stage rocket and third-stage rocket 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 carrier rocket according to claim 3, characterized in that: The aerodynamic drag D of each stage during the entire flight of the carrier rocket is: D=0.5·A·ρ·C d ·V 2 Among them, A is the frontal 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 carrier rocket according to claim 3, characterized in that: The thrust T of the carrier rocket engine at each stage during the entire flight process of the carrier rocket 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 powered 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 entire flight process of a reusable carrier rocket according to claim 1 or 2, characterized in that: The gravity y and flight Mach number η at each stage of the entire 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 carrier rocket at the current flight altitude h, v is the flight speed of the carrier rocket at the current moment, G is the gravitational constant, R e is the radius of the earth, and m is the mass of the launch vehicle at the current moment; The range R of each stage during the entire flight of the carrier rocket 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 current range R, v i is the flight speed of the carrier rocket 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 entire flight process of the carrier rocket into a launch phase and a recovery phase. The launch phase includes an ascent phase before gravity diversion, 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 initialized flight parameter variables, the flight parameter variables of the carrier rocket throughout the entire flight process 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. Mapping module: Draw a graph of the launch vehicle's flight parameter variables, gravity y, flight Mach number η, and range R based on the launch vehicle's flight parameter variables, gravity y, flight Mach number η, and range R at each stage.

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