Reconfigurable hypersonic aircraft head and deformation control strategy thereof

Abstract

The invention discloses a reconfigurable hypersonic aircraft head and a deformation control strategy thereof. Relates to the technical field of resistance and heat reduction of aircrafts, a supporting frame is arranged in the end of the aircraft, a hollow rod is arranged in the supporting frame, and a shock wave needle is slidably connected into the hollow rod; a pneumatic assembly used for driving the shock wave needle to slide is further arranged in the aircraft end. A through hole is formed in the shock wave needle and communicates with a reverse jet flow assembly. The end of the shock wave needle is provided with a boss and an aircraft end concave cavity, and the aircraft end is further provided with an ejection assembly. Through the arrangement of the reconfigurable head, the aircraft can be matched with the flight incoming flow condition and change the shape of the head, so that the requirement of the real-time working condition is met, the end heat flow or the aerodynamic resistance of the aircraft is reduced, and the optimal flight effect is achieved; meanwhile, the invention further provides a control strategy of the structure. And carrying out collaborative optimization on the continuously changing head configuration and the incoming flow parameters updated in real time to ensure that an optimal head configuration switching strategy is obtained.

Description

Technical Field

[0001] This invention relates to the field of aircraft drag reduction and heat reduction technology, and in particular to a reconfigurable hypersonic aircraft nose and its deformation control strategy. Background Technology

[0002] Shock needles, reverse jets, and windward concave cavity technology are three active drag reduction and heat reduction technologies for hypersonic vehicles that are currently receiving more attention; however, each of them has its applicable conditions and different drawbacks.

[0003] However, shock needle technology has obvious drawbacks: 1. Even with the use of high-temperature resistant materials such as tungsten alloys, it is still prone to ablation during long-term flight at Mach 8 and above; 2. Ablation of the shock needle will cause changes in its aerodynamic shape, affecting the attitude stability of the aircraft; 3. It has a narrow range of adaptable Mach numbers. When the flight Mach number deviates from the design value, the oblique shock wave will switch between "detachment and attachment", which will lead to sudden changes in drag and heat flow.

[0004] The core shortcomings of reverse jet technology lie in propellant consumption and system complexity: 1. To maintain effective jet intensity during hypersonic flight, the propellant flow rate requirement is extremely high, and propellant storage significantly increases the aircraft's added mass, reducing the payload; 2. The interaction between the jet and the incoming flow is prone to generating unsteady flow field oscillations, and shock wave oscillations may occur during jet pressure regulation, leading to fluctuations in the force on the nose and affecting flight stability.

[0005] The limitations of the windward concave cavity technology are also significant: 1. It has a weak drag reduction and heat reduction capacity. Compared with shock needles and reverse jets, its protection of the head is limited, and the heat flux density may still exceed the tolerance limit of traditional thermal protection materials; 2. There are strong flow oscillations inside the concave cavity. The interaction between the recirculation zone and the shear layer will generate periodic pressure pulsations. These high-frequency pressure pulsations will cause fatigue damage to the head structure and induce attitude fluctuations in the aircraft.

[0006] In summary, all three types of single technologies have obvious adaptability deficiencies: the full profile of hypersonic flight covers a wide range of operating conditions from low Mach number acceleration to high Mach number cruise and then to low speed reentry, and a single technology cannot simultaneously meet the requirements of "low drag, low heat load, low mass, and high stability" under all operating conditions. Summary of the Invention

[0007] To meet the requirements of low drag and low heat load across the entire range of hypersonic vehicles, and to address the need for drag reduction and heat dissipation across the entire profile of the vehicle, this application provides a reconfigurable hypersonic vehicle nose and its deformation control strategy.

[0008] In a first aspect, this application provides a reconfigurable hypersonic vehicle head, employing the following technical solution:

[0009] A reconfigurable hypersonic vehicle head includes a vehicle end head, characterized in that: a support frame is provided inside the vehicle end head, a hollow rod is provided inside the support frame, a shock needle is slidably connected inside the hollow rod, and the shock needle slides to extend or retract into the vehicle end head; the vehicle end head also provides a pneumatic component for driving the shock needle to slide.

[0010] The shock needle has a through hole along its length inside, and the through hole is connected to the reverse jet component;

[0011] The shock needle has a boss at its end, and the aircraft end has a cavity that matches the boss. When the shock needle retracts into the aircraft end, the boss is embedded in the cavity. The aircraft end also has an ejection assembly for ejecting the shock needle.

[0012] Optionally, the jet assembly includes an air source box, a jet pipeline, an electric pump, a telescopic pipe, and a flow valve. The air source box is located inside the end of the aircraft, and its outlet is connected to the jet pipeline via the electric pump. The other end of the jet pipeline passes through the hollow rod and is connected to the telescopic pipe. The telescopic pipe is connected to the end of the shock needle and communicates with the through hole. The flow valve is located on the jet pipeline.

[0013] Optionally, the pneumatic assembly includes a pneumatic pipeline and a solenoid valve. One end of the pneumatic pipeline is connected to the hollow rod, and the other end is connected to the electric pump. The solenoid valve is installed on the pneumatic pipeline.

[0014] Optionally, the ejection assembly includes a fixed magnet and an electromagnet; the fixed magnet is disposed at the end of the shock needle, and the electromagnet is disposed at the front end of the jet pipe.

[0015] Secondly, this application provides a deformation control strategy for the head of a reconfigurable hypersonic vehicle, employing the following technical solution:

[0016] A control strategy for the nose of a multimodal reconfigurable hypersonic vehicle includes the following steps:

[0017] S1. Based on the flight mission objectives and combined with the full-aircraft aerodynamic dataset under the flight profile, a baseline flight profile is generated through a trajectory optimization algorithm.

[0018] S2. By combining full-profile numerical calculation with ground test correction at characteristic points, an accurate output model for six nose configuration states under the full flight profile is established, and the output model is trained to obtain a rapid evaluation model for drag reduction and heat reduction performance.

[0019] S3. Match the drag reduction and heat reduction rapid evaluation model with the benchmark flight profile, and obtain a preliminary nose deformation scheme through outer layer optimization.

[0020] S4. Perform inner-layer optimization on the preliminary flight plan, and perform global optimization by combining the optimization objectives, flight profile constraints, and the effects and costs of nose deformation;

[0021] S5. Generate a new aerodynamic dataset based on the global flight scheme optimization results, and obtain a better flight trajectory through trajectory optimization algorithm.

[0022] Optionally, the process in step S2 is described in detail below:

[0023] A. Spatial sampling design; using full factorial design, in N sample points are generated in space. Where Ma is the Mach number, α is the angle of attack, H is the altitude, and S is a discrete configuration variable, representing six different states of the aircraft's nose. , For continuously adjustable parameter vectors;

[0024] B. High-fidelity data acquisition; for each sample point The output data was obtained by solving the viscous compressible Navier-Stokes equations using computational fluid dynamics. ;

[0025] C. Proxy model construction and training: Gaussian process regression is selected for model training. The model form is as follows:

[0026]

[0027] in, This refers to the k-th output. Let Z(x) be a basis function with zero mean and covariance of . Gaussian random process; Representing the regression coefficients; by fitting the hyperparameters through maximum likelihood estimation, a series of fast evaluation models are finally obtained:

[0028]

[0029] Wherein, the drag coefficient is The lift coefficient is The heat flow at the characteristic points of the wall is And the wall temperature of key parts is .

[0030] Optionally, step S4 includes the following specific details:

[0031] a. Optimize variables

[0032] Discretize the total flight time For each consecutive time stage i, the state and control variables are described by the following vector:

[0033]

[0034] in, This represents the duration of the i-th flight phase, and the sum of the durations of all phases equals the total flight time. ; This represents the average or representative altitude of the i-th stage aircraft; This represents the average or representative Mach number of the i-th stage aircraft; This represents the average or representative angle of attack of the aircraft in the i-th stage; Represents the head configuration state, with values ​​belonging to a set. one of the; This represents the adjustment parameter vector for the corresponding configuration; T is the transpose of the matrix.

[0035] b. Construction of the objective function

[0036] The overall objective function J is a comprehensive cost function that quantifies the sum of multiple performance metrics minimized in the task; it is in weighted sum form.

[0037]

[0038] in, Indicates a non-negative weighting coefficient; Indicates the total flight time. It is the heat flux density at the i-th stage characteristic point predicted by the surrogate model, which is related to time. The product of these two factors is approximately equal to the heat accumulation during that stage. The drag coefficient for stage i predicted by the surrogate model. Indicates the resistance encountered during flight. This refers to the flow rate of the flow valve.

[0039] Optionally, the optimization process in step S4 must satisfy the following constraints:

[0040] A. Flight path constraints:

[0041] Thermal flow constraint: Material tolerance limit ,

[0042] in, The heat flow experienced by the material The limiting heat flux of the material;

[0043] Dynamic compressive constraint: structural strength limit. ,

[0044] in, For dynamic pressure, It is the structural strength limit;

[0045] Overload restraint: Occupant or structural safety, ,

[0046] in, For aircraft overload, Maximum overload;

[0047] Flight envelope altitude constraints ,

[0048] in, For flight altitude, Minimum flight altitude Maximum flight altitude;

[0049] B. Configuration and system-related constraints:

[0050] Shock needle extension / retraction travel constraint: ,

[0051] in, This refers to the extension / retraction length of the shock needle. Minimum travel of the shock needle This represents the maximum stroke of the shock needle;

[0052] Operating range constraints of the jet flow valve: ,

[0053] in, For the flow rate of the flow valve, This is the minimum flow rate of the flow valve. This represents the maximum flow rate of the flow valve.

[0054] Configuration switching direction constraints:

[0055] Configuration switching is not arbitrary but has a sequential and directional nature;

[0056] C. Terminal constraints:

[0057] ,

[0058] It is a preset set of terminal states, such as requiring the aircraft to arrive at the target area at a specific altitude, speed and track angle.

[0059] In summary, this application includes the following beneficial technical effects:

[0060] Integrating multiple drag reduction and heat dissipation technologies into the nose of a hypersonic vehicle and using them in combination / alternately through continuous state switching creates greater variability. This overcomes the limitations of traditional single-technology "local optima," achieving a globally optimal trade-off between drag reduction, heat dissipation, range, and maneuverability throughout the entire mission cycle. It can cope with the complex and ever-changing aerodynamic and thermal environment and uncertain mission changes during flight, possessing online adaptive reconfiguration capabilities. Multiple technology states serve as backups for each other; a failure of a single component (such as the shock needle) will not lead to complete system failure, improving mission reliability. The combination / alternate operation of multiple technologies significantly optimizes the vehicle's aerodynamic performance and reduces peak heat flux constraints, allowing mission planners to design flight profiles impossible with traditional vehicles, truly unleashing the potential of hypersonic vehicles. Attached Figure Description

[0061] Figure 1 This is a cross-sectional view of the overall structure of the head of a reconfigurable hypersonic vehicle according to this application;

[0062] Figure 2 This is a schematic diagram of the construction and training process of the rapid evaluation model in this application;

[0063] Figure 3 This is a schematic diagram of the hierarchical collaborative optimization process of the global flight plan in this application;

[0064] Figure 4 This is a schematic diagram showing the switching relationship of the various head states of the aircraft in this application.

[0065] Explanation of reference numerals in the attached figures:

[0066] 1. Aircraft nose; 2. Support frame; 3. Hollow rod; 4. Shock needle; 41. Fixed magnet; 5. Telescopic tube; 51. Electromagnet; 6. Jet pipe; 61. Flow valve; 7. Pneumatic pipe; 71. Solenoid valve; 8. Electric pump; 9. Air source box. Detailed Implementation

[0067] The following is in conjunction with the appendix Figure 1-3 This application will be described in further detail.

[0068] This application discloses a reconfigurable hypersonic vehicle head, which improves the front end. It has a support frame inside the vehicle head, with a hollow rod inside the support frame. A shock needle is slidably connected inside the hollow rod, and the shock needle slides to extend or retract into the vehicle head. The vehicle head also has a pneumatic assembly for driving the shock needle's sliding. The shock needle has a through hole along its length, which connects to a reverse jet assembly. The shock needle's end has a boss, and the vehicle head has a cavity matching the boss. When the shock needle retracts into the vehicle head, the boss embeds into the cavity. The vehicle head also has an ejection assembly for ejecting the shock needle.

[0069] The reverse jet assembly includes an air source box, a jet pipeline, an electric pump, a telescopic tube, and a flow valve. The air source box is located inside the end of the aircraft, and its outlet is connected to the jet pipeline via the electric pump. The other end of the jet pipeline passes through a hollow rod and connects to the telescopic tube, which is connected to the end of the shock needle and communicates with a through hole. The flow valve is located on the jet pipeline. The pneumatic assembly includes a pneumatic pipeline and a solenoid valve. One end of the pneumatic pipeline is connected to the hollow rod, and the other end is connected to the electric pump. The solenoid valve is located on the pneumatic pipeline.

[0070] In the initial state, the shock needle retracts into the hollow rod; when the shock needle needs to extend, the flow valve is closed and the solenoid valve is opened, and gas is injected into the tail of the hollow rod by the electric pump, thereby pushing the shock needle out by air pressure; when a reverse jet is needed, the flow valve is opened, and gas is ejected from the through hole on the shock needle by the electric pump.

[0071] Finally, to achieve the windward concave cavity state, a boss is provided at the front end of the shock needle, and a cavity matching the boss is provided at the end of the aircraft. When the windward concave cavity is needed, the shock needle can be ejected by the ejection assembly. Specifically, the ejection assembly includes a fixed magnet and an electromagnet. The fixed magnet is located at the end of the shock needle, and the electromagnet is located at the front end of the jet pipe. When ejection is needed, the polarity of the electromagnet can be changed.

[0072] The control schemes for each state of the invention are as follows:

[0073] A. Basic state (passivated head): The flow valve is de-energized and no jetting occurs; the solenoid valve is de-energized and the shock needle does not extend from the head.

[0074] B. Reverse jet flow state (shock needle not working, head jetting): The flow valve is energized, the check valve is working, the head jets in reverse, and the reverse jet mass flow rate can be changed; the solenoid valve is de-energized, and the shock needle does not extend out of the head.

[0075] C. Shock needle status (adjustable extension length): When the flow valve is de-energized, no flow is emitted; when the solenoid valve is energized, the upper position is connected, gas is input to the end of the hollow rod, and the shock needle extends; when the solenoid valve is energized, the lower position is connected, gas is extracted, and the electric-pneumatic combined drive mechanism for extension and retraction is activated, and the shock needle shortens.

[0076] D. Shock needle + reverse jet state: When the flow valve is energized, the check valve operates, and the head performs reverse jetting, which can change the mass flow rate of the reverse jetting; when the solenoid valve is energized, the upper position is connected, the gas input telescopic electric-pneumatic composite drive mechanism is activated, and the shock needle extends; when the solenoid valve is energized, the lower position is connected, the gas extraction telescopic electric-pneumatic composite drive mechanism is activated, and the shock needle shortens.

[0077] E. Windward concave cavity state: The flow valve is de-energized and no jet flow occurs.

[0078] F. Windward concave cavity + reverse jet flow state: When the flow valve is energized, the check valve works, the head performs reverse jet flow, and the mass flow rate of the reverse jet flow can be changed.

[0079] By extending and retracting the shock needle, starting and stopping the jet, and controlling the ejection of the shock needle, six technical states can be switched. The structural characteristics and switching logic of each state are as follows:

[0080] A. Basic State (Passivated Nose): The shock needle is fully retracted into the nose, the reverse jet system is closed, and the nose has a smooth, passivated configuration. In this state, the leading edge curvature of the nose is relatively large, which is suitable for low-speed takeoff, reentry, and landing at low Mach numbers. It can avoid the airflow separation problem of a small-curvature nose at low speeds and ensure attitude stability.

[0081] B. Reverse Jet Flow State: The gas source tank is connected to the pipeline supplying the reverse jet working fluid. The electric pump starts operating, and high-pressure fluid is directly ejected from a small orifice at the nose of the aircraft. This high-pressure, high-speed fluid is directly injected into the near-wall flow through the orifice, significantly altering the near-wall flow field structure at the nose, increasing the shock wave separation distance and weakening the shock wave intensity, thus improving the aerodynamic and thermal load environment on the wall. Simultaneously, the ejected low-temperature working fluid carries away heat from the wall, performing secondary cooling of the wall.

[0082] C. Shock Needle State (Adjustable Extension Length): The shock needle extends along its axis via a built-in electro-pneumatic hybrid drive mechanism in the nose section. The extension length is continuously adjustable within the range of 0-100mm, while the jet system is shut down. In this state, a portion of the original structure of the aircraft nose section transforms into the aerodynamic disk of the shock needle head, enhancing the shock needle's ability to reconstruct its flow field structure. Using pneumatic pressure as the extension control scheme fully utilizes the air source tank without requiring additional mechanical structures, reducing mechanical complexity and improving system reliability. The extension shock needle integrates a displacement sensor in the drive mechanism to monitor the shock needle's extension position in real time, matching the airflow environment and adjusting the pneumatic pump pressure to ensure the shock needle's extension length meets design requirements.

[0083] D. Shock Needle + Reverse Jet State: With the shock needle extended, the head jet system is activated. The head jet system sprays high-pressure working fluid from the center of the pneumatic disc at the shock needle head towards the opposite flow direction. The jet pressure and flow rate are adjusted in real time via an electric pump and solenoid valves. At this time, the shock needle retains its extension and retraction function, allowing adjustment of the shock needle extension distance as needed to achieve drag reduction and heat dissipation performance regulation.

[0084] E. Windward Concave Cavity State: When the shock needle faces the risk of ablation during flight or the aerodynamic thermal threat diminishes, requiring a switch to long-range energy-saving operation, the shock needle is extended to its limit along its axis via an electro-pneumatic composite drive mechanism. Subsequently, the direction of the electromagnet current is changed, adjusting the electromagnet's magnetism and switching the state between the fixed magnet at the shock needle's end and the electromagnet at the front of the reverse jet channel from attraction to repulsion. Under the action of electromagnetic repulsion, the shock needle is ejected, and the conical cavity left behind after the shock needle is ejected forms a windward concave cavity, shutting down the jet system.

[0085] F. Windward concave cavity + reverse jet flow state: In the windward concave cavity state, the jet flow system is turned on, and the jet outlet is located at the center of the bottom of the concave cavity, spraying the working medium in the windward direction to achieve synergistic drag reduction and heat reduction of the concave cavity and the jet flow.

[0086] Meanwhile, this application also proposes a control strategy for the head of the aforementioned multimodal reconfigurable hypersonic vehicle, aiming to resolve the contradiction between aerodynamic thermal protection and flight performance faced by hypersonic vehicles in the entire mission profile. By dynamically and collaboratively optimizing the active head configuration system and flight trajectory planning, the global optimization of multiple objectives such as drag reduction, heat reduction, range, and maneuverability can be achieved throughout the entire mission cycle.

[0087] Includes the following steps:

[0088] S1. Based on the flight mission objectives and combined with the full-aircraft aerodynamic dataset under the flight profile, a baseline flight profile is generated through a trajectory optimization algorithm.

[0089] S2. By combining full-profile numerical calculation with ground test correction at characteristic points, an accurate output model for six nose configuration states under the full flight profile is established, and the output model is trained to obtain a rapid evaluation model for drag reduction and heat reduction performance.

[0090] S3. Match the drag reduction and heat reduction rapid evaluation model with the benchmark flight profile, and obtain a preliminary nose deformation scheme through outer layer optimization.

[0091] S4. Perform inner-layer optimization on the preliminary flight plan, and perform global optimization by combining the optimization objectives, flight profile constraints, and the effects and costs of nose deformation;

[0092] S5. Generate a new aerodynamic dataset based on the global flight scheme optimization results, and obtain a better flight trajectory through trajectory optimization algorithm.

[0093] The process in step S2 is described in detail below:

[0094] A. Spatial sampling design; using full factorial design, in N sample points are generated in space. Where Ma is the Mach number, α is the angle of attack, H is the altitude, and S is the discrete configuration variable, representing six states respectively (A represents the reference state, B represents the reverse jet state, C represents the retractable shock needle state, D represents the shock needle + reverse jet state, E represents the windward concave cavity state, and F represents the windward concave cavity + reverse jet state). , To continuously adjust the parameter vector, for example, when for state C, For state D, ;

[0095] B. High-fidelity data acquisition; for each sample point The output data was obtained by solving the viscous compressible Navier-Stokes equations using computational fluid dynamics. ;

[0096] C. Proxy model construction and training: Gaussian process regression is selected for model training. The model form is as follows:

[0097]

[0098] in, This refers to the k-th output. Let Z(x) be a basis function with zero mean and covariance of . Gaussian random process; Representing the regression coefficients; by fitting the hyperparameters through maximum likelihood estimation, a series of fast evaluation models are finally obtained:

[0099]

[0100] Wherein, the drag coefficient is The lift coefficient is The heat flow at the characteristic points of the wall is And the wall temperature of key parts is ;

[0101] Step S4 contains the following details:

[0102] a. Optimize variables

[0103] Discretize the total flight time For each consecutive time stage i, the state and control variables are described by the following vector:

[0104]

[0105] in, This represents the duration of the i-th flight phase, and the sum of the durations of all phases equals the total flight time. ; This represents the average or representative altitude of the i-th stage aircraft; This represents the average or representative Mach number of the i-th stage aircraft; This represents the average or representative angle of attack of the aircraft in the i-th stage; Indicates the head configuration state; This represents the adjustment parameter vector for the corresponding configuration; T is the matrix transpose symbol.

[0106] b. Construction of the objective function

[0107] The overall objective function J is a comprehensive cost function that quantifies the sum of multiple performance metrics minimized in the task; it is in weighted sum form.

[0108]

[0109] in, Indicates a non-negative weighting coefficient; Indicates the total flight time. It is the heat flux density at the i-th stage characteristic point predicted by the surrogate model, which is related to time. The product of these two factors is approximately equal to the heat accumulation during that stage. The drag coefficient for stage i predicted by the surrogate model. Indicates the resistance encountered during flight. This refers to the flow rate of the flow valve.

[0110] The optimization process in step S4 must satisfy the following constraints:

[0111] A. Flight path constraints:

[0112] Thermal flow constraint: Material tolerance limit ,

[0113] in, The heat flow experienced by the material The limiting heat flux of the material;

[0114] Dynamic compressive constraint: structural strength limit. ,

[0115] in, For dynamic pressure, It is the structural strength limit;

[0116] Overload restraint: Occupant or structural safety, ,

[0117] in, For aircraft overload, Maximum overload;

[0118] Flight envelope altitude constraints ,

[0119] in, For flight altitude, Minimum flight altitude Maximum flight altitude;

[0120] B. Configuration and system-related constraints:

[0121] Shock needle extension / retraction travel constraint: ,

[0122] in, This refers to the extension / retraction length of the shock needle. Minimum travel of the shock needle This represents the maximum stroke of the shock needle;

[0123] Operating range constraints of the jet flow valve: ,

[0124] in, For the flow rate of the flow valve, This is the minimum flow rate of the flow valve. This represents the maximum flow rate of the flow valve.

[0125] Configuration switching direction constraints:

[0126] Configuration switching is not arbitrary but has a sequence and direction. Specifically, state A and state B can switch freely, state B and state C can switch freely, state B and state C can both be transformed into state D in one direction, and state D and state E can switch freely.

[0127] C. Terminal constraints:

[0128] ,

[0129] It is a preset set of terminal states, such as requiring the aircraft to arrive at the target area at a specific altitude, speed and track angle.

[0130] For detailed optimization methods, please refer to [link / reference]. Figure 3 ;

[0131] First, input the task objective and constraints, including path, system, and terminal constraints. The first stage decomposes and initializes the task. The first step is to divide the task into stages based on the input, clarifying the start and end times or distances of each stage; the second step is to initialize the dynamic programming (DP) based on the input endpoint state, obtaining the initialized DP cost table.

[0132] The second stage is the core loop of the task, which is responsible for performing reverse dynamic programming recursion. It includes step three, the outer loop processes the current stage index k, step four, the inner loop gives the head state currently under consideration, and step five generates the set of allowed transition states for the next stage, outputting the current state, the next state and the constraints.

[0133] The third stage is the inner-layer continuous optimization and evaluation, which includes step six calling the NPL solver to optimize the trajectory and parameters in the current stage, and step seven being responsible for calculating the corresponding transition cost and passing the current transition cost as output to the next stage.

[0134] The fourth stage is responsible for specifying the optimal state. Step eight compares all transition cost options in the current stage (ki, k-1i). Step nine records the optimal (k-1)i state and outputs k and (k-1)i. Step ten judges the value of k. If k ≠ 1, then k = k-1 and returns to the second stage. If k = 1, the loop is completed and the complete optimal decision table is output.

[0135] The fifth stage is scheme generation and output. Step eleven is forward backtracking to the optimal sequence. Step twelve is global trajectory optimization, which takes the optimal sequence as input and solves the trajectory and parameters of the whole process. Step thirteen outputs the final scheme, outputting the complete spatiotemporal trajectory and parameter curves, and supports the end of scheme planning.

[0136] The above are all preferred embodiments of this application, and are not intended to limit the scope of protection of this application. Therefore, all equivalent changes made in accordance with the structure, shape and principle of this application should be covered within the scope of protection of this application.

Claims

1. A reconfigurable hypersonic vehicle nose comprising a vehicle nose tip, characterized in that: The aircraft nose is internally provided with a support frame, the support frame is internally provided with a hollow rod, the hollow rod is internally slidably connected with a shock wave needle, the shock wave needle slides to extend out of or retract into the aircraft nose; the aircraft nose is further internally provided with a pneumatic assembly for driving the shock wave needle to slide; The shock wave needle is internally provided with a through hole along the length direction of the shock wave needle, the through hole is communicated with a reverse jet flow assembly; The end of the shock wave needle is provided with a boss, the aircraft nose is provided with a recess matching the boss, when the shock wave needle retracts into the aircraft nose, the boss is embedded into the recess, the aircraft nose is further provided with an ejection assembly for ejecting the shock wave needle.

2. The reconfigurable hypersonic vehicle nose of claim 1, wherein: The reverse jet flow assembly comprises a gas source box, a jet flow pipeline, an electric pump, an extension tube and a flow valve, the gas source box is arranged inside the aircraft nose, the outlet thereof is connected with the jet flow pipeline through the electric pump; the other end of the jet flow pipeline penetrates into the hollow rod and is connected with the extension tube, the extension tube is connected to the end of the shock wave needle and is communicated with the through hole, and the flow valve is arranged on the jet flow pipeline.

3. The reconfigurable hypersonic vehicle nose of claim 2, wherein: The pneumatic assembly comprises a pneumatic pipeline and an electromagnetic valve, one end of the pneumatic pipeline is communicated into the hollow rod, the other end is connected with the electric pump, and the electromagnetic valve is arranged on the pneumatic pipeline.

4. The reconfigurable hypersonic vehicle nose of claim 3, wherein: The ejection assembly comprises a permanent magnet and an electromagnet; the permanent magnet is arranged at the end of the shock wave needle, and the electromagnet is arranged at the front end of the jet flow pipeline.

5. A morphing control strategy for a reconfigurable hypersonic vehicle nose, applied to the reconfigurable hypersonic vehicle nose according to claim 4, characterized in that: The method comprises the following steps: S1. According to the flight task target, combining the full-aircraft aerodynamic data set under the flight profile, a reference flight profile is generated through a trajectory optimization algorithm; S2. An output model accurate in five head configuration states under the full flight profile is established through full-profile numerical calculation combined with a feature point ground test correction mode, and the output model is trained to obtain a rapid evaluation model of drag reduction and heat reduction performance; S3. The rapid evaluation model of drag reduction and heat reduction is matched with the reference flight profile, and a preliminary head deformation scheme is obtained through outer optimization; S4. The preliminary flight scheme is subjected to inner optimization, and global optimization is performed in combination with an optimization target, a flight profile constraint, and the effect and cost of head deformation; S5. A new aerodynamic data set is generated according to the global flight scheme optimization result, and a more optimal flight trajectory is obtained through a trajectory optimization algorithm.

6. The morphing control strategy for a reconfigurable hypersonic vehicle nose according to claim 5, wherein: The process in step S2 is described in detail as follows: A. Space sampling; Adopting full-factor design, in N sample points are generated in the space ; wherein Ma is Mach number, a is attack angle, H is height, S is discrete configuration variable, respectively representing six states of aircraft nose, , is a continuous adjustment parameter vector; B. High-fidelity data acquisition; for each sample point obtaining output data by solving the viscous compressible Navier-Stokes equations through computational fluid dynamics and constructing a training set G; C. Proxy model construction and training, a Gaussian process regression is selected for model training, and the model form is: where, is the kth output, is the basis function, Z(x) is a Gaussian random process with zero mean and covariance denotes the regression coefficients; by fitting the hyperparameters through maximum likelihood estimation, a series of fast evaluation models are finally obtained:​ wherein the drag coefficient is , the lift coefficient is , the wall feature point heat flow is , and the wall temperature of the key position is .

7. The morphing control strategy for a reconfigurable hypersonic vehicle nose according to claim 6, wherein: In step S4, the specific content is as follows: a. Optimization variables Discretizing the total flight time into a number of consecutive time phases, the state and control quantities of each phase i being described by the following vectors: wherein, denotes the time duration of the i-th flight phase, the sum of all phase durations equals the total flight time ; denotes the average or representative altitude of the aircraft in the i-th flight phase; denotes the average or representative Mach number of the aircraft in the i-th flight phase; denotes the average or representative angle of attack of the aircraft in the i-th flight phase; denotes the head configuration state; denotes the adjustment parameter vector for the corresponding configuration; T is the matrix transpose symbol; b. Objective function construction The total objective function J is a comprehensive cost function, which quantifies the sum of multiple performance indicators to be minimized in the task; a weighted sum form is adopted: wherein, represents a non-negative weight coefficient; represents the total flight time, is the i-th stage feature point heat flux density predicted by the surrogate model, which is approximately equal to the heat accumulation of the stage with the product of the time is the i-th stage drag coefficient predicted by the surrogate model, represents the resistance to flight, is the flow valve flow.​ 8. The morphing control strategy for a reconfigurable hypersonic vehicle nose according to claim 7, wherein: The optimization process in step S4 needs to meet the following constraint conditions: A. Flight path constraint; Heat flow constraint: material tolerance limit , wherein, q is the heat flux to the material, q is the heat flux to the material; Dynamic pressure constraint: structural strength limit, , wherein is the dynamic pressure, is the structural strength limit; Overload constraint: occupant or structural safety , wherein is the aircraft overload, is the maximum overload; flight envelope height constraint, , wherein is the flight altitude, is the minimum flight altitude, is the maximum flight altitude; B. Configuration and system related constraint: Shock wave needle stroke constraint: , wherein, is the shock wave needle stroke length, is the shock wave needle minimum stroke is the shock wave needle maximum stroke; Jet flow valve operating range constraints: , wherein, Q is the flow rate of the flow valve, Qmin is the minimum flow rate of the flow valve, Qmax is the maximum flow rate of the flow valve; Configuration switching direction constraint: Configuration switching is not arbitrary switching but has sequence and directionality; C. Terminal constraint; ， is a set of predetermined terminal states requiring the aircraft to arrive over the target area at a specific altitude, speed, and track angle.

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