A Point-to-Point Transport Trajectory Design Method for a Two-Stage VTVL Launch Vehicle

Abstract

This invention discloses a point-to-point transport trajectory design method for a two-stage VTVL launch vehicle, comprising: using shift handover conditions as the terminal constraint of the powered ascent phase trajectory, determining the powered ascent phase trajectory, and obtaining the maximum effective payload mass satisfying the terminal constraint through powered ascent phase trajectory optimization; setting flight constraint conditions for the unpowered reentry phase; obtaining the upper and lower boundaries of the angle of attack profile based on the flight constraint conditions; and obtaining the maximum range S of the entire trajectory based on the upper and lower boundaries of the angle of attack profile and the powered ascent phase trajectory. max Determine the maximum range S of the entire ballistic trajectory. max With target range S target Does it satisfy 0≤S? max -S target If ≤ε is satisfied, the unpowered return trajectory is determined based on the upper and lower boundaries of the angle of attack profile, shift handover conditions, and maximum effective payload mass, resulting in a full-range trajectory with maximum carrying capacity. If not satisfied, the shift handover conditions are modified, and the full-range trajectory is iteratively optimized. This invention can obtain the optimal full-range trajectory with maximum carrying capacity.

Description

Technical Field

[0001] This invention belongs to the field of launch vehicle trajectory design, and specifically relates to a point-to-point transportation trajectory design method for a two-stage VTVL launch vehicle. Background Technology

[0002] With the rapid advancement of science and technology and the rapid development of the world's space transportation system, the world's space transportation system is moving towards flight-like operation. As an important part of flight-like operation, the global point-to-point high-speed transportation system is developing rapidly and has extremely strong commercial and military value.

[0003] The global point-to-point high-speed transportation system adopts a two-stage configuration, with both stages using vertical takeoff and vertical landing (VTVL). This differs from traditional launch vehicles and reentry vehicles such as the Space Shuttle and X-37B. Its ballistic design is mainly divided into a powered ascent stage and a powered return stage. The two flight stages are closely coupled. The terminal conditions of the powered ascent stage have a significant impact on the range, thermal / load environment, and propellant consumption of the reentry stage. The entire range of the transportation system consists of the ascent stage range and the return stage range, requiring joint optimization design. The number of optimization variables and constraints increases significantly. Therefore, a new two-stage VTVL launch vehicle point-to-point transportation full-range ballistic design method is needed. Summary of the Invention

[0004] The purpose of this invention is to overcome the above-mentioned defects and provide a point-to-point transportation trajectory design method for a two-stage VTVL launch vehicle. This method solves the technical problems of complex and low-accuracy existing trajectory design processes. Through full-course trajectory calculation, this invention ultimately obtains the optimal full-course trajectory with maximum carrying capacity.

[0005] To achieve the above-mentioned objectives, the present invention provides the following technical solution:

[0006] A method for designing the entire trajectory of a two-stage VTVL launch vehicle for point-to-point transport, comprising:

[0007] S1 divides the entire point-to-point transport trajectory of the two-stage VTVL launch vehicle into a powered ascent phase and a powerless return phase.

[0008] S2 uses the shift handover conditions as the terminal constraint of the powered ascent stage trajectory, and optimizes the powered ascent stage trajectory to obtain the maximum effective payload mass of the powered ascent stage terminal under the terminal constraint.

[0009] S3 sets flight constraints for the unpowered return phase;

[0010] S4 performs an angle-of-attack profile boundary analysis on the unpowered return phase based on the flight constraints, and obtains the upper and lower boundaries of the angle-of-attack profile.

[0011] S5 calculates the maximum range of the entire trajectory based on the upper and lower boundaries of the angle-of-attack profile and the powered ascent phase ballistics. max ;

[0012] S6 determines the maximum range of the entire ballistic trajectory. max With target range S target Does it meet the following requirements:

[0013] 0≤S max -S target ≤ε, where ε is a preset threshold;

[0014] If the condition is met, proceed to step S7; otherwise, proceed to step S8.

[0015] S7 determines the unpowered return trajectory based on the upper boundary of the angle of attack profile, the lower boundary of the angle of attack profile, shift handover conditions, and the maximum effective payload mass.

[0016] By integrating the powered ascent phase trajectory and the unpowered return phase trajectory into a complete trajectory, the optimal complete trajectory with the maximum payload capacity is obtained.

[0017] S8 adjusts the shift handover conditions and returns to step S2.

[0018] Furthermore, the handover conditions include the vehicle's altitude h, speed v, and trajectory inclination γ.

[0019] Furthermore, in step S2, the method for determining the powered ascent trajectory using shift handover conditions as the terminal constraint of the powered ascent trajectory includes:

[0020] S2.1 Establish a ballistic dynamics model for the powered ascent phase;

[0021] S2.2 Based on the ballistic dynamics model of the powered ascent phase, the shift handover conditions are used as the terminal constraints of the powered ascent phase trajectory to optimize the pitch program angle of the powered ascent phase, so as to maximize the effective payload mass at the terminal of the powered ascent phase.

[0022] Furthermore, in step S2.2, the method for optimizing the pitch program angle is as follows:

[0023] The first-level pitch program angle satisfies the following formula:

[0024]

[0025] Where γ represents the trajectory inclination angle, ω z The component of the Earth's rotational angular velocity on the Z-axis of the launch coordinate system is represented by t1, t2, and t3. The launch vehicle flies at an angle of attack close to zero during the time interval [t2, t3].

[0026] Angle of Attack Where α m This represents the maximum absolute value of the angle of attack in the subsonic segment during the power ascent phase, where 'a' is a constant coefficient.

[0027] The second level uses a linear pitch program angle that satisfies the following formula:

[0028]

[0029] Where k is the pitch program angle change rate, and t3 and t4 represent the pitch program angle change time intervals;

[0030] Newton's iteration method is used for iterative calculations to solve for the pitch program angle profile parameter α that satisfies the shift handover conditions. m The formulas k and k maximize the effective payload mass at the end of the powered ascent stage of the launch vehicle.

[0031] Furthermore, in step S3, the flight constraints for the unpowered return phase include thermal flow constraints, normal overload constraints, and dynamic pressure constraints:

[0032] The heat flow constraint is:

[0033]

[0034] in, K represents heat flux density. q R is the heat transfer coefficient. c The radius of the vehicle's nose stagnation point. To constrain the maximum heat flux at the stagnation point, Based on the material system of the thermal protection system, ρ is the atmospheric density, V is the velocity vector, and n and k are constants.

[0035] The normal overload constraint is:

[0036]

[0037] Where, n y For normal overload during flight, n ymax α is the maximum permissible normal overload during flight, m ​​is the mass of the launch vehicle, g is the acceleration due to gravity, and L and D are the lift and drag forces, respectively.

[0038] The dynamic pressure constraint is:

[0039] q=0.5ρv 2 ≤q max

[0040] Where q is the dynamic pressure during flight, q max The maximum permissible dynamic pressure, in N / m. 2 .

[0041] Furthermore, the unpowered return phase includes the free flight phase, the initial reentry phase, the gliding phase, and the vertical landing phase;

[0042] In step S4, the angle-of-attack profile boundary analysis for the unpowered return segment is performed based on the flight constraints. The method for obtaining the upper and lower boundaries of the angle-of-attack profile is as follows:

[0043] S4.1 Establish the relationship between angle of attack and velocity:

[0044]

[0045] Where α0 is the initial constant angle of attack of the launch vehicle during reentry, and M is the Mach number of the launch vehicle during flight;

[0046] S4.2 Iterate α0, obtain α based on α0, and fly with this α. n y When one of q reaches the maximum value of the constraint condition, and the other two terms are less than the maximum value of the constraint condition, the lower boundary of the angle of attack profile is obtained;

[0047] The boundary of the angle of attack profile is obtained by subtracting the predetermined safety margin from the stall angle of attack of the S4.3 launch vehicle.

[0048] Furthermore, in step S5, the maximum range S of the entire trajectory is obtained based on the upper boundary of the angle of attack profile, the lower boundary of the angle of attack profile, and the trajectory during the powered ascent phase. max The method is as follows:

[0049] S5.1 The shortest and longest gliding ranges are obtained from the lower boundary and the upper boundary of the angle of attack profile, respectively.

[0050] S5.2 derives the shortest range S of the entire trajectory based on the shortest and longest ranges of the powered ascent and gliding phases. min and longest range S max .

[0051] Furthermore, in step S7, the method for determining the trajectory of the unpowered return phase based on the upper boundary of the angle of attack profile, the lower boundary of the angle of attack profile, shift handover conditions, and the maximum effective payload mass at the end of the powered ascent phase is as follows:

[0052] S7.1 Establish a ballistic dynamics model for the unpowered return phase;

[0053] S7.2 Based on the ballistic dynamics model of the unpowered return phase, the angle of attack and tilt angle of the unpowered return phase are determined according to the upper boundary of the angle of attack profile, the lower boundary of the angle of attack profile, the shift handover conditions, and the maximum effective payload mass at the end of the powered ascent phase.

[0054] Furthermore, in step S7.2, the method for determining the angle of attack of the unpowered return trajectory is as follows:

[0055] Within the boundary range of the angle of attack profile, the angle of attack that satisfies the range requirement and the flight constraints of the unpowered return segment is obtained by iterating α0. The range requirement is satisfied when the flight range determined by the angle of attack is equal to the target range S. target ;

[0056] When the flight range determined based on the angle of attack is less than the target range S target Decrease the angle of attack, and vice versa.

[0057] Furthermore, in step S7.2, the method for determining the tilt angle of the unpowered return trajectory is as follows:

[0058] Free flight phase and initial descent phase:

[0059] The free flight phase and the initial descent phase use a fixed 0° roll angle:

[0060] σ=σ0=0

[0061] Gliding phase:

[0062] According to the equilibrium gliding condition, the gliding segment satisfies the rate of change of the trajectory inclination angle γ being equal to 0° / s. Assuming γ≈0, the inclination angle σ is obtained from the formula for the rate of change of the trajectory inclination angle:

[0063]

[0064] Where r is the position vector of the launch vehicle's center of mass, g is the gravitational acceleration, and L is the lift force.

[0065] Compared with the prior art, the present invention has the following advantages:

[0066] (1) This invention creatively proposes a point-to-point transportation full-course trajectory design method for a two-stage VTVL launch vehicle. The handover parameters are used as optimization variables for the full-course trajectory. Under the condition of satisfying various constraints, the carrying capacity is used as the objective function. The inner loop is optimized by segmenting the trajectory of the powered ascent stage and the unpowered return stage, thereby completing the full-course trajectory optimization of the inner and outer loops.

[0067] (2) The present invention uses Newton's iteration method to optimize the trajectory of the powered ascent phase. Under the condition of meeting the terminal constraint requirements of the ascent phase, the flight procedure angle is iteratively optimized to maximize the use of propellant and improve the carrying capacity.

[0068] (3) This invention proposes a multi-constraint trajectory optimization method for the unpowered return segment, which maximizes the reentry and return range under process constraints such as heat flow, overload, and dynamic pressure.

[0069] (4) The present invention can meet the requirements of full-range ballistic design, and by taking into account a variety of factors, it can effectively improve the accuracy of ballistic design and improve design efficiency. Attached Figure Description

[0070] Figure 1 A schematic diagram of a point-to-point transport flight profile of a two-stage VTVL launch vehicle.

[0071] Figure 2 This is a flowchart illustrating the entire ballistic design process of the present invention. Detailed Implementation

[0072] The features and advantages of the present invention will become clearer and more apparent from the following detailed description.

[0073] The term “exemplary” as used herein means “serving as an example, embodiment, or illustration.” Any embodiment illustrated herein as “exemplary” is not necessarily to be construed as superior to or better than other embodiments. Although various aspects of embodiments are shown in the accompanying drawings, the drawings are not necessarily drawn to scale unless specifically indicated otherwise.

[0074] This invention provides a novel point-to-point transport trajectory design method for a two-stage VTVL launch vehicle. It employs a decomposition-based trajectory design concept, dividing the trajectory design into a powered ascent stage and a depowered reentry stage. The powered ascent stage trajectory design references the launch vehicle ascent stage design method, considering the return of the first stage. The depowered reentry stage trajectory design draws heavily on the design methods of lifting reentry vehicles, designing the trajectory by defining the angle of attack and tilt angle profiles. Finally, based on the powered ascent and depowered reentry stage trajectory designs, a full-range trajectory design is carried out, using shift change parameters as optimization variables to establish an inner and outer loop full-range trajectory optimization method, ultimately obtaining the optimal full-range trajectory with maximum payload capacity.

[0075] The specific solution of this invention is as follows:

[0076] A method for designing the entire ballistic trajectory for point-to-point transport of a two-stage VTVL launch vehicle includes the following steps:

[0077] S1, designing a two-stage VTVL launch vehicle point-to-point transport flight profile;

[0078] S2, Establish a ballistic dynamics model for the powered ascent phase;

[0079] S3, Establish a ballistic design method for the powered ascent phase;

[0080] S4. Establish a ballistic dynamics model for the unpowered return segment;

[0081] S5, set flight constraints for the unpowered return phase;

[0082] S6, Conduct boundary analysis of the angle-of-attack profile of the unpowered return segment;

[0083] S7, Establish a ballistic design method for unpowered return phase;

[0084] S8, based on S1-S7, establishes a two-stage VTVL launch vehicle point-to-point transportation full-course trajectory design method.

[0085] Example:

[0086] This invention provides a method for designing the entire trajectory of a two-stage VTVL launch vehicle for point-to-point transport, comprising the following steps:

[0087] S1, designing a two-stage VTVL launch vehicle point-to-point transport flight profile;

[0088] The long-range high-speed transport system employs a two-stage vertical takeoff and vertical landing method, enabling it to deliver payloads to a designated location within one hour. The entire flight process is divided into two main phases: the powered ascent phase and the unpowered return phase. Figure 1 A schematic diagram of the flight trajectory is provided, where the powered ascent phase includes the first and second ascent phases; the unpowered return phase includes the free flight phase, initial reentry phase, gliding phase, and vertical landing phase. Because the vertical landing phase has a shorter range, significantly shorter than other flight phases, the trajectory design for the unpowered return phase primarily focuses on the free flight, initial reentry, and gliding phases. This invention employs a trajectory design method based on a decomposition strategy. Steps S2-S3 and S4-S7 design the powered ascent and unpowered return phases separately, and finally, step S8 combines the two flight phases for optimized design.

[0089] S2, Establish a ballistic dynamics model for the powered ascent phase;

[0090] The dynamic equations for the powered ascent phase are established in the launch inertial frame as follows:

[0091] 1) Equations of translational motion of the center of mass

[0092] In the launch coordinate system, the equation of motion for the translational center of mass is:

[0093]

[0094] Where m is the mass of the launch vehicle, g is the gravity vector, V is the velocity vector, P is the thrust vector of the launch vehicle, and R is the velocity vector. n It is aerodynamic.

[0095] 2) Kinematic equations of the center of mass

[0096] In the launch inertial coordinate system, the kinematic equation of the launch vehicle's center of mass is:

[0097]

[0098]

[0099]

[0100] x, y, z, V x V y V z These represent the position and velocity of the launch vehicle in the launch inertial coordinate system, respectively.

[0101] S3, conduct ballistic design for powered ascent phase;

[0102] Based on the ballistic dynamics model of the powered ascent phase established by S2, the ballistic trajectory design of the powered ascent phase is carried out. The ballistic trajectory design of the ascent phase mainly involves designing the flight procedure angle to meet the handover conditions required by the design, including the second-stage shutdown point altitude, velocity, and trajectory inclination angle, while designing the launch azimuth angle to ensure that the launch vehicle flies to the target point in the shortest distance.

[0103] Designing the flight procedure angle is an important part of the overall design of the launch vehicle. The flight performance of the launch vehicle (such as carrying capacity, interstage separation altitude, and substage landing position) is related to the trajectory shape determined by the flight procedure angle. Therefore, in addition to meeting the given terminal constraints (i.e., shift handover conditions), the flight procedure angle of the powered ascent phase of the launch vehicle must also take into account the above-mentioned flight performance.

[0104] Corresponding to the atmospheric flight performance requirements, the first-stage pitch program angle must meet the requirements of zero angle of attack and constant separation attitude. The first stage can be summarized as follows:

[0105]

[0106] In the formula, γ represents the trajectory inclination angle, ω z The component of the Earth's rotational angular velocity on the Z-axis of the launch coordinate system is represented by t1, t2, and t3. The launch vehicle flies at an angle of attack close to zero during the time interval [t2, t3].

[0107] The angle of attack varies exponentially, as shown in the following formula:

[0108] α(t)=4α m e -a(t-t1) (1-e -a(t-t1) )

[0109] In the formula, α m This represents the maximum absolute value of the angle of attack in the subsonic range, where 'a' is a constant coefficient.

[0110] During the second stage, which is in a vacuum phase, the pitch program angle is generally used in a linear form, as shown in the following formula:

[0111]

[0112] Where k is the pitch program angle change rate, and t3 and t4 represent the pitch program angle change time intervals.

[0113] Newton's iteration method is used for iterative calculations to solve for the pitch program angle profile parameter α that meets the handover conditions required by the design. m The design variables, such as k, are used to maximize the effective payload mass during the powered ascent phase of the launch vehicle.

[0114] Through the ballistic design of the powered ascent stage in step S3, the terminal state parameters of the powered ascent stage ballistics (mass m, height h, velocity v, trajectory inclination angle γ) are obtained, providing initial input for the ballistic design of the unpowered return stage.

[0115] S4. Establish a ballistic dynamics model for the unpowered return segment;

[0116] In the trajectory coordinate system, the dynamic equations for the entire reentry phase can be expressed as follows:

[0117]

[0118]

[0119]

[0120]

[0121]

[0122]

[0123] Where r is the launch vehicle's center of mass position vector, V is the velocity vector, θ is longitude, φ is latitude, σ is the roll angle, ψ is the trajectory azimuth angle, and γ is the ballistic inclination angle. e The value represents the Earth's rotational angular velocity; m is the mass of the launch vehicle; g is the acceleration due to gravity; L and D are the lift and drag forces, respectively.

[0124]

[0125]

[0126] ρ=ρ0e (-h / H)

[0127] h = rR e

[0128] In the formula, S ref For reference area; C L C DThese are the lift coefficient and drag coefficient, respectively, which are related to the angle of attack α; ρ is the atmospheric density. This invention uses an exponential atmospheric model, with the atmospheric density at sea level ρ0 = 1.226 kg / m³. 3 H is the reference altitude, taken as 7254.24m, and the Earth's radius R e =6371.2km.

[0129] S5, setting flight constraints for the unpowered return phase.

[0130] 1) Heat Flux Constraints: To ensure the safety of the launch vehicle, the heat flux density at the stagnation point (especially the nose cone) must be less than the maximum allowable flight value when designing the reentry trajectory.

[0131]

[0132] in K represents heat flux density. q R is the heat transfer coefficient. c The radius of the vehicle's nose stagnation point. The maximum heat flux constraint at the stagnation point is determined based on the material system of the thermal protection system. n and k are constants. For typical hypersonic reentry problems, we can take n = 3.15 and k = 0.5.

[0133] 2) The magnitude of the allowable normal overload during reentry is determined by the load, the structure of the aircraft, or the load-bearing capacity of its equipment. The normal overload constraint is set as follows:

[0134]

[0135] Where, n y For normal overload during flight, n ymax α represents the maximum permissible normal overload during flight, and α is the angle of attack during flight.

[0136] 3) Dynamic pressure constraint

[0137] For high-speed transport aircraft controlled by aerodynamic control surfaces, the hinge torque cannot be too large. Therefore, dynamic pressure needs to be limited to reduce the load on the actuators. Dynamic pressure constraints limit the flight trajectory of transatmospheric aircraft after entering the dense atmosphere.

[0138] q=0.5ρv 2 ≤q max

[0139] Where q is the dynamic pressure during flight, q max The maximum permissible dynamic pressure, in N / m. 2 .

[0140] S6, Conduct boundary analysis of the angle-of-attack profile of the unpowered return segment.

[0141] Based on the constraint conditions in step S5, the boundary of the angle of attack profile for the unpowered return segment is analyzed. Referring to the traditional reentry angle of attack profile, the angle of attack profile is set as a function of velocity, with the following form:

[0142]

[0143] Where α0 is the initial constant angle of attack of the aircraft during reentry, and M is the Mach number of the aircraft. The constant large angle of attack α0 is mainly used to suppress dynamic pressure and heat flow.

[0144] According to ballistic analysis, the angle of attack profile is inversely proportional to dynamic pressure, heat flux, overload, and range. Therefore, the minimum value of the initial constant angle of attack α0 corresponds to the maximum constraints of dynamic pressure, heat flux, and overload, while the maximum value corresponds to the minimum range constraint.

[0145] 1) Solving for the lower boundary of the angle of attack profile: Iterate through the initial constant angle of attack α0, satisfying the lower limit constraints of dynamic pressure, overload, heat flux, and maximum lift-to-drag ratio angle of attack. When any one of the constraints—dynamic pressure, heat flux, or overload—approaches the maximum value of the constraint condition, while the other two are less than the constraint values, the corresponding minimum angle of attack value corresponds to the aircraft's maximum range S. max .

[0146] 2) Solving for the upper boundary of the angle of attack profile: This is obtained by subtracting a certain safety margin from the stall angle of attack for different aircraft configurations. The larger the angle of attack profile, the higher the deceleration efficiency of the aircraft, the shorter the flight time, and the smaller the total heat absorption. Therefore, the upper boundary of the angle of attack profile corresponds to the minimum range S of the aircraft. min .

[0147] From the upper and lower boundaries of the angle of attack profile obtained in step S6, the shortest and longest ranges of the unpowered gliding phase can be obtained. Then, from the range of the powered ascent phase obtained in step S5, the shortest range S of the entire flight phase can be obtained. min and longest range S max Given S target If 0 ≤ S max -S target If the total flight range is ≤ε (ε is a fixed constant set to 100km), then start step S7, design the unpowered return trajectory to ensure the total flight range meets the target range; otherwise, skip to step S8 and adjust the handover parameters: altitude h, speed v, and trajectory inclination γ.

[0148] S7, conduct ballistic design for unpowered return phase;

[0149] Based on the unpowered reentry phase ballistic dynamics model established in step S4 and the angle-of-attack profile boundary obtained in step S6, the unpowered reentry phase ballistic design is carried out. The unpowered reentry phase ballistic design mainly includes two parts: angle-of-attack profile design and bank angle profile design. Based on the initial altitude and velocity, the angle of attack and bank angle are designed to meet constraints and range requirements. The angle of attack also uses the aforementioned angle-of-attack profile, where the initial constant angle of attack α0 for reentry needs to be obtained iteratively through range requirements. The bank angle design is determined according to the following steps:

[0150] 1) Free flight phase and initial descent phase

[0151] The free flight phase and the initial descent phase use a fixed 0° roll angle:

[0152] σ=σ0=0

[0153] 2) Gliding phase

[0154] The gliding phase employs a balanced gliding design, where the condition for balanced gliding is that the rate of change of the trajectory inclination is equal to 0° / s, i.e.:

[0155]

[0156] According to the formula for the rate of change of the trajectory inclination:

[0157]

[0158] For balanced gliding trajectories with relatively small inclination angles, γ≈0 can be set, and the roll angle can be calculated based on the balanced gliding conditions.

[0159]

[0160] After determining the roll angle based on the conditions for balanced gliding, the flight range is mainly determined by the angle of attack profile. Therefore, iterative design of the angle of attack profile is required. Within the angle of attack boundary, an angle of attack profile that satisfies both constraints and range requirements should be designed. If the range is less than the target range, the angle of attack should be reduced; conversely, the angle of attack should be increased.

[0161] S8, based on S1-S7, establishes a two-stage VTVL launch vehicle point-to-point transportation full-course trajectory design method.

[0162] Based on steps S2-S3 and S4-S7, an inner and outer loop full-range ballistic optimization method is established. Under the condition that the overall parameters and target range are determined, altitude h, velocity v, and ballistic inclination angle γ are used as the shift handover parameters as optimization variables for the entire ballistic trajectory. The shift handover parameters are used as the terminal constraint for the powered ascent phase. The maximum effective payload mass satisfying this constraint is obtained through ballistic optimization of the powered ascent phase and added to the shift handover parameter input for the unpowered return phase. An angle-of-attack profile boundary analysis is then performed to obtain the longest-range ballistic trajectory for the unpowered return phase. If 0 ≤ Smax -S target If the distance is less than or equal to ε, then proceed to step S7, designing the unpowered return trajectory, to ensure the flight range meets the target range requirement. If not, adjust the shift handover parameters and perform the full-range trajectory calculation again. This process is iterated until the optimal full-range trajectory for maximum payload capacity is obtained. The entire flowchart is as follows: Figure 2 As shown.

[0163] The present invention has been described in detail above with reference to specific embodiments and exemplary examples; however, these descriptions should not be construed as limiting the present invention. Those skilled in the art will understand that various equivalent substitutions, modifications, or improvements can be made to the technical solutions and embodiments of the present invention without departing from the spirit and scope of the invention, and all such modifications and improvements fall within the scope of the present invention. The scope of protection of the present invention is defined by the appended claims.

[0164] The contents not described in detail in this specification are common knowledge to those skilled in the art.

Claims

1. A method for designing the entire ballistic trajectory for point-to-point transport of a two-stage VTVL launch vehicle, characterized in that, include: S1 divides the entire point-to-point transport trajectory of the two-stage VTVL launch vehicle into a powered ascent phase and a powerless return phase. S2 uses the shift handover conditions as the terminal constraint of the powered ascent stage trajectory, and optimizes the powered ascent stage trajectory to obtain the maximum effective payload mass of the powered ascent stage terminal under the terminal constraint. S3 sets flight constraints for the unpowered return phase; S4 performs an angle-of-attack profile boundary analysis on the unpowered return phase based on the flight constraints, and obtains the upper and lower boundaries of the angle-of-attack profile. S5 calculates the maximum range of the entire trajectory based on the upper and lower boundaries of the angle-of-attack profile and the powered ascent phase ballistics. max ; S6 determines the maximum range of the entire ballistic trajectory. max With target range S target Does it meet the following requirements: 0≤S max -S target ≤ε, where ε is a preset threshold; If the condition is met, proceed to step S7; otherwise, proceed to step S8. S7 determines the unpowered return trajectory based on the upper boundary of the angle of attack profile, the lower boundary of the angle of attack profile, shift handover conditions, and the maximum effective payload mass. By integrating the powered ascent phase trajectory and the unpowered return phase trajectory into a complete trajectory, the optimal complete trajectory with the maximum payload capacity is obtained. S8 adjusts the shift handover conditions and returns to step S2.

2. The method for designing the entire trajectory of a two-stage VTVL launch vehicle for point-to-point transportation according to claim 1, characterized in that, The handover conditions include the vehicle's altitude h, speed v, and trajectory inclination γ.

3. The point-to-point transport trajectory design method for a two-stage VTVL launch vehicle according to claim 2, characterized in that, In step S2, the method for determining the trajectory of the powered ascent phase, using shift handover conditions as the terminal constraint, includes: S2.1 Establish a ballistic dynamics model for the powered ascent phase; S2.2 Based on the ballistic dynamics model of the powered ascent phase, the shift handover conditions are used as the terminal constraints of the powered ascent phase trajectory to optimize the pitch program angle of the powered ascent phase, so as to maximize the effective payload mass at the terminal of the powered ascent phase.

4. The point-to-point transport trajectory design method for a two-stage VTVL launch vehicle according to claim 3, characterized in that, In step S2.2, the method for optimizing the pitch program angle is as follows: The first-level pitch program angle satisfies the following formula: Where γ represents the trajectory inclination angle, ω z The component of the Earth's rotational angular velocity on the Z-axis of the launch coordinate system is represented by t1, t2, and t3. The launch vehicle flies at an angle of attack close to zero during the time interval [t2, t3]. Angle of Attack Where α m This represents the maximum absolute value of the angle of attack in the subsonic segment during the power ascent phase, where 'a' is a constant coefficient. The second level uses a linear pitch program angle that satisfies the following formula: Where k is the pitch program angle change rate, and t3 and t4 represent the pitch program angle change time intervals; Newton's iteration method is used for iterative calculations to solve for the pitch program angle profile parameter α that satisfies the shift handover conditions. m The formulas k and k maximize the effective payload mass at the end of the powered ascent stage of the launch vehicle.

5. The point-to-point transport trajectory design method for a two-stage VTVL launch vehicle according to claim 2, characterized in that, In step S3, the flight constraints for the unpowered return phase include thermal flow constraints, normal overload constraints, and dynamic pressure constraints: The heat flow constraint is: in, K represents heat flux density. q R is the heat transfer coefficient. c The radius of the vehicle's nose stagnation point. To constrain the maximum heat flux at the stagnation point, Based on the material system of the thermal protection system, ρ is the atmospheric density, V is the velocity vector, and n and k are constants. The normal overload constraint is: Where, n y For normal overload during flight, n ymax α is the maximum permissible normal overload during flight, m ​​is the mass of the launch vehicle, g is the acceleration due to gravity, and L and D are the lift and drag forces, respectively. The dynamic pressure constraint is: q=0.5ρv 2 ≤q max Where q is the dynamic pressure during flight, q max The maximum permissible dynamic pressure, in N / m. 2 .

6. The point-to-point transport trajectory design method for a two-stage VTVL launch vehicle according to claim 5, characterized in that, The unpowered return phase includes the free flight phase, the initial reentry phase, the glide phase, and the vertical landing phase. In step S4, the angle-of-attack profile boundary analysis for the unpowered return segment is performed based on the flight constraints. The method for obtaining the upper and lower boundaries of the angle-of-attack profile is as follows: S4.1 Establish the relationship between angle of attack and velocity: Where α0 is the initial constant angle of attack of the launch vehicle during reentry, and M is the Mach number of the launch vehicle during flight; S4.2 Iterate α0, obtain α based on α0, and fly with this α. n y When one of q reaches the maximum value of the constraint condition, and the other two terms are less than the maximum value of the constraint condition, the lower boundary of the angle of attack profile is obtained; The boundary of the angle of attack profile is obtained by subtracting the predetermined safety margin from the stall angle of attack of the S4.3 launch vehicle.

7. The point-to-point transport trajectory design method for a two-stage VTVL launch vehicle according to claim 6, characterized in that, In step S5, the maximum range S of the entire trajectory is obtained based on the upper boundary of the angle of attack profile, the lower boundary of the angle of attack profile, and the trajectory during the powered ascent phase. max The method is as follows: S5.1 The shortest and longest gliding ranges are obtained from the lower boundary and the upper boundary of the angle of attack profile, respectively. S5.2 derives the shortest range S of the entire trajectory based on the shortest and longest ranges of the powered ascent and gliding phases. min and longest range S max .

8. The point-to-point transport trajectory design method for a two-stage VTVL launch vehicle according to claim 7, characterized in that, In step S7, the method for determining the trajectory of the unpowered return phase based on the upper boundary of the angle of attack profile, the lower boundary of the angle of attack profile, shift handover conditions, and the maximum effective payload mass at the end of the powered ascent phase is as follows: S7.1 Establish a ballistic dynamics model for the unpowered return phase; S7.2 Based on the ballistic dynamics model of the unpowered return phase, the angle of attack and tilt angle of the unpowered return phase are determined according to the upper boundary of the angle of attack profile, the lower boundary of the angle of attack profile, the shift handover conditions, and the maximum effective payload mass at the end of the powered ascent phase.

9. The point-to-point transport trajectory design method for a two-stage VTVL launch vehicle according to claim 8, characterized in that, In step S7.2, the method for determining the angle of attack of the unpowered return trajectory is as follows: Within the boundary range of the angle of attack profile, the angle of attack that satisfies the range requirement and the flight constraints of the unpowered return segment is obtained by iterating α0. The range requirement is satisfied when the flight range determined by the angle of attack is equal to the target range S. target ; When the flight range determined based on the angle of attack is less than the target range S target Decrease the angle of attack, and vice versa.

10. The point-to-point transport trajectory design method for a two-stage VTVL launch vehicle according to claim 9, characterized in that, In step S7.2, the method for determining the tilt angle of the unpowered return trajectory is as follows: Free flight phase and initial descent phase: The free flight phase and the initial descent phase use a fixed 0° roll angle: σ=σ0=0 Gliding phase: According to the equilibrium gliding condition, the gliding segment satisfies the rate of change of the trajectory inclination angle γ being equal to 0° / s. Assuming γ≈0, the inclination angle σ is obtained from the formula for the rate of change of the trajectory inclination angle: Where r is the position vector of the launch vehicle's center of mass, g is the gravitational acceleration, and L is the lift force.

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