129 results about "High-speed flight" patented technology

The invention relates to a high-speed aircraft transition identification method based on a flight test high-frequency vibration signal. The method comprises the following steps: selecting high-quality measurement point data suitable for analysis from flight test high-frequency vibration data as a vibration signal to be analyzed; performing time-frequency analysis on the to-be-analyzed vibration signal, outputting a time-frequency diagram, and representing the change condition of energy in a frequency domain by using a centroid frequency and an envelope difference of the centroid frequency; carrying out preliminary identification on the transition occurrence moment; decomposing a to-be-analyzed vibration signal into a plurality of intrinsic mode functions with different center frequencies through variational mode decomposition; quantizing the change condition of the energy of each modal component along with time; and aiming at the change condition of the energy of each modal component along with the time, the transition is identified according to the energy of the mode excited by the transition, and the energy sudden increase point is the moment when the transition begins to occur. According to the method, the transition occurrence time can be effectively identified under the condition that no hole is formed in the surface of the aircraft.

The embodiment of the invention relates to the technical field of aerospace safety, and provides an aircraft trajectory prediction method and system based on space gridding and heuristic search. According to the method, adaptive grid resolution adjustment is performed on high-speed aircraft trajectory prediction, and a possible high-probability trajectory is quickly calculated according to the type of the aircraft, so that the problems of inflexible calculation resource coordination, insensitive aircraft type perception and low calculation binding physical constraint in the current high-speed aircraft trajectory prediction are solved.

The invention relates to a method for intercepting a flying object through aerial high-speed net casting, which is used for solving the technical problem of low interception success rate caused by insufficient maneuverability of an interception means and lack of enough large and reliable interception area when approaching a target in the prior art, and comprises the following steps: controlling an unmanned aerial vehicle to fly to approach the target; an unlocking signal is sent to a locking mechanism arranged on the machine main body, so that the locking mechanism synchronously releases all the power systems; the released power system flies away by means of self-rotation inertia, and the catching net connected with the power system is pulled by the pulling rope, so that the catching net is unfolded to form an intercepting area in the flying-away process. The unmanned aerial vehicle keeps flying at a high speed in the pursuit stage, the unmanned aerial vehicle flies away through inertia of a plurality of power systems in the interception stage, the capture net is unfolded to form an interception area larger than the projection area of the vehicle body, and compared with a traditional interception method, the interception method has the higher interception speed and the higher interception success rate.

The invention discloses a high-speed aircraft aerodynamic configuration agile optimization method based on an artificial intelligence model. The method comprises the following steps: A1, establishing an aerodynamic force agent model; a2, performing a global optimization algorithm; a3, establishing an intelligent autonomous decision-making model; by constructing a multi-target adaptive weight distribution model, weight dynamic collaborative optimization of proxy model confidence, a constraint penalty function and target function key elements is achieved, the mathematical modeling problem of dynamic matching of weight parameters and optimization targets in the iteration process is solved, a closed-loop optimization system with the autonomous perception-decision-execution capacity is developed, and the optimization efficiency is improved. Therefore, the convergence speed is improved in a complex engineering optimization scene, local optimum is avoided, and the robustness of a global optimal solution is ensured; and A4 pneumatic optimization iteration: based on the proxy model for predicting indexes such as the lift-drag ratio, performing global optimization of design parameters by using an artificial intelligence optimization algorithm, performing verification through CFD, and after verification, putting a CFD calculation result into a training set of the model to re-train the model and continuing optimization until optimization design is converged.

The invention discloses a high-speed aerodynamic derivative analysis modeling method and system based on symbolic regression, and belongs to the technical field of aircraft aerodynamic modeling and artificial intelligence cross, and the method comprises the steps: obtaining an aerodynamic data set of a high-speed aircraft, and carrying out the preprocessing; defining grammar for generating candidate symbol expressions and constructing an expression tree; modeling based on the training set by adopting a layered symbol regression framework to obtain an analytical model; the prediction precision of the analytical model is verified on the test set, and physical consistency analysis is carried out; carrying out uncertainty quantification on the analytical model to obtain predicted uncertainty estimation of the analytical model; and finally, outputting an analytic model in a mathematical expression form for describing the relationship between the aerodynamic derivative and the flight state variable and uncertainty estimation of the analytic model. The method overcomes the defects of a traditional method in precision, generalization and interpretability, realizes full-automatic efficient modeling from data to the analytic model, and improves the modeling efficiency. The method is suitable for control design, stability analysis and real-time simulation of the high-speed aircraft.

The utility model discloses a bionic aircraft side wing, and belongs to the technical field of bionic equipment. The aircraft mainly comprises an aircraft framework and side wing parts arranged on the two sides of the aircraft framework, a fixing frame and a driving component installed on the fixing frame and used for driving the side wing parts to move are arranged at the front end of the aircraft framework, a battery pack and a microcontroller are arranged on the aircraft framework, and each side wing part comprises a front wing, a rear wing and a connecting block. The bionic aircraft side wing is composed of the front wing body, the rear wing body and the connecting block, the front wing body and the rear wing body are connected through the connecting block and provided with the outer contour frames and the wing faces, and the cross structures are arranged in the outer contour frames, so that the rigidity, stability and overall strength of the side wing can be effectively enhanced, and the excellent performance of an aircraft in high-speed flight or complex airflow is guaranteed; meanwhile, the airfoil is made of a thin film material with high tensile strength and elasticity, so that the structural weight is reduced, the aerodynamic performance is optimized, and the efficiency and reliability of the aircraft are improved.

The invention relates to the technical field of high-speed aircraft ground simulation and electromagnetic property calculation, and provides a high-speed aircraft electromagnetic calculation method and system based on a quantum classification algorithm. A variable component sub-classifier is used as an intelligent scheduling core to perform feature recognition and region division on a whole flow field, and a transmission matrix method and a WKB approximation method are adaptively called according to a division result to perform hybrid calculation, so that the technical problem that a traditional single calculation strategy is difficult to collaboratively optimize in precision and efficiency is solved; the method has the advantages that global computing resources can be distributed according to needs, and the overall solving speed is greatly increased while high simulation precision of the key area is guaranteed.

The invention discloses a fixed wing patrol aircraft flight attitude control method, and belongs to the technical field of unmanned aerial vehicles. Aiming at the technical contradiction that high-speed flight maneuverability and weapon shooting stability cannot be compatible, the method achieves breakthrough through the following innovative schemes that a four-wing propeller collaborative framework is designed, and each propeller is provided with a horizontal linear driver to achieve front-back telescopic displacement in the axial direction of a fuselage; in the horizontal flight stage, based on the feedback of a gyro stabilizer, pitching / rolling double-axis attitude accurate control is realized through the cooperative adjustment of propeller position stretching and rotation speed difference; triggering a layered compensation strategy by an ignition event, and synchronously adjusting the horizontal position and the rotating speed of a propeller according to the recoil force direction; and the tail vane intervenes in fine adjustment only when the yaw is out of tolerance. By combining the seamless conversion design of the vertical take-off and landing stage, the three technical problems of height loss without initial-speed deployment, low aerodynamic efficiency of high-speed flight and insufficient shooting precision are solved.

The invention relates to the technical field of high-speed aircraft health management, and particularly discloses an interpretable aircraft state sensing and predicting system and method. The method comprises the following steps: constructing and training a lightweight anomaly diagnosis network with an expansion perception layer; aircraft dynamics, kinematics and fault mechanism formulas are used as soft constraints to be embedded into the future state estimation network; an SHAP interpretation framework is used for sorting the importance of the input features, and a network structure is optimized accordingly; and finally, accurate diagnosis of abnormal types and positions and rapid prediction of aircraft states at a plurality of moments in the future are realized. According to the method, by introducing three innovation points of an expansion perception layer, physical information embedding and feature importance sorting, the diagnosis accuracy and the model interpretability are remarkably improved, and the high-performance calculation requirement in a scene with limited calculation power is met.

The invention discloses a bearing and wave-transmitting integrated design method and system for a variable-density honeycomb structure radome, and relates to the technical field of high-speed aircraft radome design. The method comprises the following steps: carrying out force / heat / electric performance characterization on a honeycomb cell element model; sensitivity analysis is conducted on the relation between the cell element structure parameters and the electromagnetic performance, and key design parameters are obtained; establishing an electrical performance calculation model of the two-dimensional honeycomb structure planar radome; determining an antenna-radome system, dispersing the three-dimensional radome in the axial direction to form a plurality of design points, and calculating a weighted average incident angle and force thermal load distribution of each design point; calculating an initial honeycomb relative density recommendation value at each design point; calculating the wave transmission rate and the aiming error of the formed variable-density radome; calculating the force thermal deformation and stress of the variable-density radome; and judging whether the scheme performance meets constraint conditions or not. By optimizing the distribution rule of the variable-density honeycomb structure, comprehensive optimization of the bearing performance and the wave-transparent performance of the radome is achieved.

The application discloses a hypersonic flow instability multi-field coupling prediction method and system, relates to the technical field of modern high-speed aircraft design, and comprises the following steps: acquiring the multi-physical field parameter distribution of a hypersonic flow field; based on a shock wave-boundary layer interaction mechanism, constructing an instability sensitive graph data structure by taking the multi-physical field parameters as input, setting graph nodes in the shock wave influence area, the boundary layer area and the intersection area thereof, and calculating the connection weight between the nodes; performing multi-scale decomposition on the instability sensitive graph data structure, and extracting instability modes of the shock wave scale, the vortex scale and the molecular scale; screening the instability modes by adopting a hypersonic stability criterion, and determining a dominant instability mode; and constructing an instability evolution prediction model based on the dominant instability mode, and outputting the spatiotemporal evolution law of the hypersonic flow instability. The application can solve the problem that the flow instability phenomenon under the complex multi-physical field environment of hypersonic speed is difficult to accurately predict.

The invention belongs to the technical field of high-speed wind tunnel tests, and particularly relates to a high-speed aircraft forward jet flow wind tunnel test device and a design method thereof. A thrust balance, a jet flow device and a ring type balance device are arranged in a jet flow test model of the test device, and the jet flow test model is externally connected with an air supply pipeline and a supporting device. The design method comprises the steps of selecting simulation parameters; carrying out similar simulation on the jet flow momentum; designing a jet flow test model; determining a jet flow test model supporting mode; designing a jet flow device; designing a jet thrust balance; carrying out ground jet flow debugging; and performing wind tunnel test verification. According to the test device and the design method thereof, direct measurement of jet flow thrust and direct measurement of aerodynamic characteristics of a jet flow test model are achieved, direct measurement of the jet flow thrust is achieved under the condition that the internal space of the model is not increased, and the jet flow thrust can be accurately obtained while the axial force of the model is obtained; therefore, the method has an engineering practical value.

An ultra-high-speed helicopter with stowed single-blade coaxial rotors is provided and relates to the technical field of aircrafts. The ultra-high-speed helicopter includes a fuselage. Wings and tails can be arranged on the fuselage, respectively. Tail blades can be arranged at a tail end of the fuselage. A rotor mechanism can be arranged on a top of the fuselage. The rotor mechanism includes rotors. The rotor mechanism can drive the rotors to be stored in a rotor storage cabin of the fuselage. When the helicopter flies at high speed, the rotors stop working, and tail blades provide propulsive force and the wings provide lift force, so that the defects in the aerodynamic characteristics of the rotors are avoided during high-speed flight. In addition, the rotors can be retracted and stowed in a fuselage to further reduce drag and increase the maximum flight speed.

The invention discloses a low-altitude high-speed flight manned flying saucer, and relates to the technical field of aviation aircrafts, the low-altitude high-speed flight manned flying saucer comprises a saucer-shaped body, a power device and an undercarriage, the saucer-shaped body is composed of a central bearing ball cabin and a peripheral reinforcing ring coaxially sleeved outside the central bearing ball cabin; the central bearing ball cabin is divided into a celestial navigation and communication cabin, a panoramic cockpit and an environment control and life support system cabin from top to bottom; the inner cabin section of the peripheral reinforcing ring forms a passenger cabin; the bearing strength and the space utilization rate of the machine body are effectively improved through the composite structure of the central spherical cabin and the peripheral reinforcing ring; due to the integrated design of the annular duct and the vector engine, the aerodynamic efficiency and the control flexibility in the vertical take-off and landing and high-speed flight states are remarkably improved; an independent environment control system ensures the safety and comfort of passengers in high-altitude flight; the overall layout gives consideration to functionality and maintenance convenience, and has high practical value.

The invention relates to the technical field of data link communication abnormal flow detection, in particular to a high-speed aircraft data link communication abnormal flow detection method and system, and the method comprises the steps: judging a communication message which accords with a preset data link communication protocol specification as a protocol layer compliance message, and carrying out the analysis to obtain a formation cooperation instruction; collecting multi-source task context information of a local high-speed aircraft to generate a task execution field snapshot; projecting the formation cooperation instruction to the task execution field snapshot for comparison and analysis to obtain a multi-dimensional semantic comparison result; if all the multi-dimensional semantic comparison results are passed, the corresponding communication flow is judged to be task-driven normal collaborative flow; and if the judgment result of the multi-dimensional semantic comparison result is that the flow does not pass, judging the corresponding communication flow as the intention camouflage type spoofing abnormal flow. The deception traffic of which the protocol layer is compliant but the tactical intention is abnormal is effectively identified, so that the accuracy of high-speed aircraft data link communication abnormal traffic detection is improved.

This invention relates to the field of target recognition technology, specifically providing a lightweight target recognition method and system for high-speed aircraft terminals. The method includes: extracting a first infrared image sequence when target occlusion is detected; calculating aerosol static distribution index and aerosol dynamic distribution index by analyzing the first infrared image sequence; fusing the two into an aerosol spatial information index; selecting a first target image sequence using a genetic algorithm; calculating optimized parameters for an infrared imaging system by analyzing the aerosol static distribution index and aerosol dynamic distribution index of the first target image sequence; obtaining a second target image sequence using the optimized parameters; and tracking and identifying the target based on the second target image sequence. This invention reduces computational load while maintaining target recognition accuracy, enabling efficient target recognition by a high-speed aircraft terminal in dusty aerosol environments.

This invention belongs to the field of plasma flow control technology for aircraft, specifically relating to an air-breathing plasma jet generator suitable for active flow control of aircraft. It is installed within an internal cavity of the aircraft, connected to an air intake and a nozzle. The air-breathing plasma jet generator generates orthogonal electric and magnetic fields. The electric field includes components in the direction of airflow and perpendicular to airflow. The magnetic field is used to confine electrons to gyroscopic motion and to ionize them through collisions with gas molecules. Furthermore, the magnetic field causes the ion current perpendicular to the airflow direction to generate an Ampere force in the direction of airflow under the influence of the magnetic field. This invention effectively solves the problem of exciter failure during high-speed flight by embedding the exciter entirely inside the aircraft and combining an innovative design of integrated excitation using magnetic fields, thermal energy, and acoustic waves, significantly improving the speed applicability range of plasma flow control technology.

This invention discloses an integrated heat dissipation and power generation device for high-speed aircraft, comprising: a heat dissipation pulsating heat pipe arranged in the nose cone of the high-speed aircraft; the heat dissipation pulsating heat pipe includes an annular pulsating heat pipe condensation section, several U-shaped pulsating heat pipe evaporation sections, and two intermediate sections connecting each U-shaped pulsating heat pipe evaporation section to the annular pulsating heat pipe condensation section; at least one branch pipe communicating with the intermediate section is arranged at intervals between the two intermediate sections; a heat extraction pipe for heat exchange with the annular pulsating heat pipe condensation section of the heat dissipation pulsating heat pipe; and a power generation module for converting the heat energy acquired by the heat extraction pipe into electrical energy. This integrated device provides thermal protection for the aircraft while converting high-temperature aerodynamic heat into electricity, alleviating the aircraft's demand for onboard battery power, reducing the need for carrying cooling fluid, and improving energy utilization efficiency. It is of great significance for aircraft operating in harsh thermal environments for extended periods.

A retractable propeller device for an airborne vehicle has a structure that allows the propeller to be exposed or concealed depending on whether the airborne vehicle is in high-speed flight or parked. Specifically, the retractable propeller device includes: a base having a mounting member with a space; a propeller unit disposed on the mounting member and configured to generate airflow; a cover configured to close the space and movable vertically relative to the mounting member; and a lifting unit disposed on the mounting member and connected to the cover, configured to move the cover. The cover selectively closes or opens the space based on the operation of the lifting unit, making the propeller unit operable.

The invention relates to the technical field of thermal protection of aerospace engines and high-speed aircrafts, and discloses a porous structure for sweating and cooling, the porous structure comprises a first subarea, a second subarea and a third subarea which are connected in sequence, each subarea comprises octahedral trusses which are arranged periodically, each octahedral truss comprises a regular octahedral frame located in the center and eight regular tetrahedral frames arranged on the periphery of the regular octahedral frame in a filling mode, the porosity of the first subarea, the porosity of the second subarea and the porosity of the third subarea are different, the porosity of the first subarea is larger than that of the third subarea, the first subarea is located in a high-heat-flow area, and the third subarea is located in a low-heat-flow area. According to the method, the using amount of a coolant can be remarkably reduced on the premise that the cooling efficiency and uniformity are guaranteed, the service life of the structure can be prolonged, the lightweight level of the structure can be improved, and finally an efficient and reliable advanced thermal protection solution with engineering implementation is provided for high-heat-flow components such as high-speed aircrafts, aero-engine combustion chambers and gas turbines.

The invention discloses a composite material and metal mixed connection structure and method, and belongs to the technical field of high-speed aircrafts, the structure comprises a heat-proof layer, a metal layer, an adhesive layer and a fastener; wherein the heat-proof layer is connected with the metal layer through the adhesive layer; and a part of the fastener is arranged in the heat-proof layer, and the other part of the fastener is connected with the metal layer. The structure reliability of the aircraft under the high dynamic pressure and high heat flow environment is guaranteed.

This invention discloses an integrated design method for a forebody / inlet based on longitudinal segmented multi-stage compression, relating to the field of high-speed aircraft forebody / inlet design. The method includes: designing a conical waverider forebody with at least two stages of compression and determining the incoming Mach number of the inlet; constructing an inner conical axisymmetric reference flow field using a characteristic line design method, and matching the axisymmetric reference flow field with the shock wave of the conical waverider forebody to form the inlet capture profile; obtaining an inward-rotating inlet using a forward and reverse tracking design fusion approach; and fusing the inward-rotating inlet with the conical waverider forebody to obtain an integrated forebody / inlet configuration. This invention, after employing forward and reverse fusion design of the inlet, further integrates it with the forebody. Compared to existing independent forebody and inlet designs, this method offers advantages such as improved inlet pre-compression performance, shortened aircraft forebody length, increased total pressure recovery and pressure ratio of the inlet, and enhanced inlet start-up performance.

This invention discloses a heat dissipation system and its control method, as well as a high-altitude, high-speed aircraft. It belongs to the fields of thermal protection and heat pipe technology. The heat dissipation system includes a single-phase heat storage circuit, a heat pump heat dissipation circuit, and an auxiliary heat sink module. The single-phase heat storage circuit includes a connected heat-absorbing cold plate and a liquid storage tank. The heat-absorbing cold plate absorbs heat emitted by the heat-generating equipment, and the liquid storage tank absorbs and stores heat from the medium in the first pipeline. The evaporator of the heat pump heat dissipation circuit absorbs heat from the medium in the first pipeline, and the condenser of the heat pump heat dissipation circuit releases heat from the heat pump heat dissipation circuit. The auxiliary heat sink module includes a liquid storage tank, an evaporation chamber, and a nozzle located in the evaporation chamber and connected to the liquid storage tank. Part of the piping of the heat pump heat dissipation circuit is located within the evaporation chamber, and the nozzle is used to spray working fluid into the evaporation chamber. The purpose of this invention is to provide a heat dissipation system and its control method, as well as a high-altitude, high-speed aircraft, with good heat dissipation performance, suitable for heat dissipation of high-power equipment.

This invention discloses a method and system for dynamic communication reconfiguration of high-speed aircraft. The method calculates plasma characteristic parameters by real-time acquisition of flight status parameters and communication link quality parameters, establishing a plasma sheath electron density distribution model; calculates plasma attenuation based on electromagnetic wave propagation theory, and predicts future link attenuation using a flight trajectory analytical extrapolation method; determines whether to initiate reconfiguration by integrating multi-dimensional triggering criteria of link attenuation, signal-to-noise ratio, and bit error rate; and jointly optimizes communication frequency, transmit power, and beam pointing angle using a multi-objective collaborative optimization method aimed at minimizing communication interruption probability and maximizing power efficiency. The optimal parameters are then distributed to execute reconfiguration and form a closed-loop control. This invention can dynamically adjust communication parameters according to real-time flight status and plasma characteristics, significantly improving the communication support capability and energy utilization efficiency of high-speed aircraft in a plasma sheath environment.

The invention discloses a high-speed aircraft aerodynamic topology design method based on a level set method. The method comprises the steps that 1, a flow field control equation containing a high-speed physical model and a structural analysis model are established; 2, defining an optimization problem taking a lift-drag ratio, a lift coefficient or a resistance coefficient as a target function and taking a volume or an area as a constraint condition; 3, implicitly describing the boundary of the aerodynamic contour of the aircraft by adopting a level set function, and converting an optimization problem into an evolution problem of the level set function; 4, efficiently calculating shape sensitivity information of the objective function about the level set function by solving the adjoint equation; and 5, based on the shape sensitivity information, constructing and solving a level set evolution equation, and driving the aerodynamic boundary to automatically evolve to a topological structure with better performance until convergence. The method solves the problem that an existing parameterization optimization method is difficult to realize automatic evolution of an aerodynamic configuration topological structure and accurate capture of shock waves in a high-speed flow field.

A device and method for measuring the gap between closed compartments is disclosed, belonging to the field of industrial automation production technology. This invention addresses the difficulty and complexity of measuring and calculating gap distances in existing technologies. The invention simulates the fitting process using a closed thermal protection model and a closed compartment model, employing multiple gap measuring instruments to simultaneously collect characteristic data of the fitting area. By analyzing this characteristic data, the current gap condition is determined, and the gap is adjusted to the target range. This invention is primarily used for measuring the gap between closed compartments in high-speed aircraft.

This invention relates to the field of thermal structure and thermal protection technology for high-speed aircraft, and particularly to a heat-dissipating structure, comprising: a high thermal conductivity leading edge, a rudder shell, a heat shield, a rudder handle, a power capillary, a drag-reducing capillary, and a dual working medium; the high thermal conductivity leading edge is fixedly connected to the front end stagnation point of the rudder shell; the rudder shell is a hollow cavity structure, which is fixedly connected to the rudder handle; the heat shield is embedded in the connection gap between the rudder shell and the rudder handle to block heat conduction from the rudder shell to the rudder handle; the power capillary and the drag-reducing capillary are respectively disposed inside the rudder shell; the dual working medium is uniformly adsorbed within the porous structure of the power capillary and the drag-reducing capillary to achieve heat conduction and transfer through phase change. This invention's heat-dissipating structure combines material conduction with structural conduction, which not only minimizes the operating temperature of the heat-dissipating structure but also reduces the structural deformation tendency, improving its safety and reliability.

This invention discloses a high-speed laminar flow airfoil, belonging to the field of aircraft aerodynamic design. The airfoil has a maximum relative thickness of 14%C, a maximum relative thickness location of 41.66%C, a maximum relative camber of 2.14%C, and a maximum relative camber location of 79.06%C, where C is the chord length. The airfoil has a flat upper surface and a concave structure at the trailing edge of the lower surface. Its operating Mach number range is 0.2–0.7, and its operating Reynolds number range is 6 × 10⁻⁶. 6 ~20×10 6 Under design conditions, the laminar flow coverage of the upper surface of the airfoil reaches 50% of the chord length, at Mach 0.7 and Reynolds number 20 × 10⁻⁶. 6 Under these conditions, the maximum lift-to-drag ratio reaches 150, at Mach 0.2 and Reynolds number 6 × 10⁻⁶. 6 Under these conditions, the maximum lift coefficient reaches 1.774, and the stall angle of attack reaches 17°. This invention overcomes the technical challenges of low critical Mach number in high-lift airfoils and easy laminar transition in traditional laminar airfoils at high Mach numbers, achieving a balance between high-speed cruise and low-speed takeoff and landing performance, and is suitable for high-speed aircraft such as turboprop regional aircraft.

Thermal protection systems for aerospace vehicles, such as hypersonic vehicles or space vehicles, are disclosed herein. In some embodiments, the disclosed thermal protection systems can be used to protect propellant tanks (e.g., composite propellant tanks) and vehicle primary structures (e.g., composite primary structures) from exceeding upper design temperature limits during vehicle ascent, reentry into Earth's atmosphere, and / or high velocity flight through an atmosphere. In some embodiments, a thermal protection system can include a framework of elongated composite members forming a plurality of cells. One or more layers of insulation material such as aerogel can be positioned in each of the cells. The cells can be covered by an external skin of the vehicle. In some embodiments, the framework is attached to a propellant tank and the insulation material is in contact with the external surfaced of the propellant tank.

The application discloses a rear-swept wing distributed non-equal-diameter rotorcraft and a control method thereof, which comprises a fuselage, a wing, a tilt-rotor group, a landing gear, a horizontal tail and a vertical tail; the tilt-rotor group comprises four tilt-rotors which are connected to the leading edge of the wing through a support rod; the rotor disc area near the wing tip is larger than the rotor disc area near the wing root, so as to balance the flight efficiency at low speed and the flight speed at high speed, and the rear sweep of the wing is beneficial to improving the stability of the tilt-rotor aircraft in the fixed-wing mode flight, and the environmental adaptability of the aircraft is improved. The application balances the efficiency of the tilt-rotor aircraft in the hovering and cruising states, the aircraft structure is relatively simple, is beneficial to the attitude control of the aircraft and the flight control software design.