406 results about "High-speed flight" patented technology

The invention relates to an intelligent control method for a high-speed aircraft, and the method comprises the steps: building an aircraft control architecture, carrying out the decision-making of a neural network model, training a decision-making neural network, and carrying out the real-time compensation of a control instruction according to the flight state of the aircraft, so as to achieve the control of a variable-configuration high-speed aircraft. An aircraft controller adopts a model + learning double-layer architecture, a decision neural network is designed to be of a Double Network structure, overload errors of pitching, yawing and rolling channels are considered in a reward function design, training of a decision neural network model is based on a double-delay deterministic strategy gradient algorithm, a training process is optimized by adopting a prior experience playback technology and a target strategy smoothing mechanism, and a decision neural network model is obtained. The control system can achieve real-time intelligent adjustment of control instructions, overload quantities and attitude angles of three channels of the aircraft and corresponding rudder deflection instructions output by the controller at the last moment are obtained from the environment, the intelligent body outputs control instructions of pitching, yawing and rolling channels, and therefore intelligent control over the variable-configuration high-speed aircraft is achieved.

The invention discloses a micro-channel device for phase change sweating and an application method, and belongs to the technical field of thermal protection of aircrafts. The structure sequentially comprises a solid inner wall, a transportation phase change layer and a sweating outer wall surface from inside to outside, and is mainly composed of a bottom plate, a porous rib wall, a solid rib wall, a solid outer wall, an inlet flow channel, a main flow channel, a branch flow channel and the like. The bottom plate is a solid inner wall and is used for supporting and conveying; the flow channel can be divided into an inlet flow channel, a main flow channel and a branch flow channel, and a cooling medium sequentially passes through the inlet flow channel, the main flow channel and the branch flow channel; a cooling medium is injected into an external flow field through the porous rib walls which are periodically arranged; the solid rib walls isolate the adjacent main runners, so that a cooling medium flows out through the porous rib walls; the outer surfaces of the porous rib walls of the solid outer wall are flush to form a sweating outer wall surface; the device has the advantages of being integrated in structure, small in thickness, uniform in sweating and efficient, and a new way can be provided for long-time cross-domain sailing thermal protection of high-speed aircrafts.

The invention provides a morphing wing integrating telescopic variable-elongation and rotary variable-sweepback functions and a morphing method. The morphing wing comprises an inner wing section, an outer wing section and an elastic skin, the inner wing section and the outer wing section are positioned in the same spanwise plane; the inner wing section stretches and retracts in the spanwise plane along the spanwise direction of the wing, so that the span length of the inner wing section is adjusted; the outer wing sections rotate in the chordwise direction relative to the inner wing sections in the spanwise plane, and then the sweepback angles of the outer wing sections are adjusted. By adjusting the extension length and the sweepback angle of the wings, the lift force of the aircraft during low-speed flight is improved, and the resistance of the aircraft during high-speed flight is reduced. Specifically, the extension length of the wings is increased during low-speed flight to improve lift force, and the extension length is reduced and the sweepback angle is increased during high-speed flight to reduce resistance. The optimal aerodynamic performance can be achieved no matter in a low-speed flight state or a high-speed flight state, and the adaptability and the flight performance of the aircraft are remarkably improved.

The invention discloses an active thermal protection system for a high-speed aircraft and a using method of the active thermal protection system. The system comprises a liquid storage tank, an active thermal protection plate, a gas high-pressure storage tank and a thrust engine. The liquid storage tank, the active thermal protection plate and the gas high-pressure storage tank are connected in sequence to form a loop; a first electromagnetic valve and a fourth electromagnetic valve are respectively arranged on a connecting pipeline between the liquid storage tank and the active thermal protection plate and a connecting pipeline between the gas high-pressure storage tank and the liquid storage tank; the gas high-pressure storage tank is further provided with two gas paths, one gas path is connected with the thrust engine through a third electromagnetic valve, and the other gas path is communicated with the outside through a second electromagnetic valve. The liquid storage tank is used for storing a liquid heat-absorbing working medium; and the active thermal protection plate is filled with a catalyst. The active thermal protection system is high in heat absorption capacity, high-temperature and high-pressure gas can be generated through aerodynamic heat to do work, waste heat utilization is achieved, the thickness of a thermal insulation layer of the aircraft is reduced, the effective load of the system is increased, and the maneuverability of the aircraft is effectively improved.

The invention provides a high-speed aircraft time cooperative guidance method and system based on sliding mode control, and belongs to the field of aircraft guidance and control. The problems that an existing guidance method is large in calculation amount, insufficient in real-time performance and poor in cooperation precision are solved. According to the method, the residual flight time of multiple aircrafts is quickly predicted through a centralized collaborative strategy, and the time deviation of the aircrafts is adjusted according to the principle that who follows the aircrafts slowly; designing a bottom layer time controllable guide law based on sliding mode control, and introducing a saturation function to improve a buffeting phenomenon and solve a singular point problem; realizing transverse trajectory bending of the aircraft through a three-dimensional offset proportional guidance law so as to adjust the flight time; the precision and efficiency of residual flight time calculation are improved by using a variable step size numerical integration method; and completing formation transformation of multiple aircrafts at the tail end of the gliding section to support a terminal guidance task. According to the method, the calculation complexity is reduced, meanwhile, the precision and robustness of time cooperation are remarkably improved, and reliable technical support is provided for a multi-aircraft cooperative guidance task.

The invention relates to an aerodynamic force-thermal-structure multi-field coupling intelligent prediction method for a high-speed aircraft, and belongs to the technical field of high-speed aircraft design. According to the flight path of the aircraft, sample parameters used for calculating aerodynamic force and aerodynamic heat states are extracted; modeling the aircraft, and carrying out analog calculation to obtain an aerodynamic force and aerodynamic heat data set; calculating an aircraft temperature field along the trajectory; training the aerodynamic force data set to obtain an aerodynamic force prediction model; calculating structural stress displacement field data of the aircraft model; extracting concerned aircraft feature point corresponding data from the aerodynamic data set, the aircraft temperature field and the structural stress displacement field data of the aircraft model to obtain a feature point data set; training the feature point data set to obtain an accurate prediction result; according to the method, after artificial intelligence algorithm offline learning modeling is carried out on the aerodynamic force-heat-structure coupling data of analog simulation, the flight path is input, and the data, changing along the flight path, of the feature point data of the output aircraft is rapidly predicted.

The invention belongs to the field of dynamics of high-speed rotor aircrafts, and particularly relates to a method for calculating rotation flutter damping of a tilt rotorcraft. The method comprises the following steps: acquiring a full-aircraft modal shape and modal frequency of the tilt-rotor helicopter; screening out a mode which can be excited by a load transmitted by a rotor wing from a full-machine mode vibration mode of the tilt rotor wing helicopter; extracting modal displacement vectors and angular displacement vectors of the centers of the left and right pairs of rotor hubs in respective modals from the modals excited by the loads transmitted by the rotors, and converting the modal displacement vectors and the angular displacement vectors to a fuselage coordinate system; a rotor dynamics simulation model is established in CAMRAD, and the modal displacement vector and the angular displacement vector which are converted to the fuselage coordinate system are used for replacing the fuselage modal; a fuselage horizontal tail and a wing aerodynamic force output interface are respectively connected with the centers of a left propeller hub and a right propeller hub by using a CAMRAD CORE command.

According to the transition prediction method suitable for the wide Mach domain of the high-speed aircraft provided by the invention, temperature and Mach number coupling correction is performed on key parameters in time scale modeling of a second mode of a transition mode, and simplified processing of a flow compression effect and a heat transfer process by traditional single parameter association is broken through; the influence of an incoming flow disturbance effect is embedded in the mode, and dynamic association between test environment parameters and transition triggering criteria is achieved; according to the improved mode, accurate prediction results are obtained on a plurality of examples with the Mach number range of 4.7-14, the research results provide a more reliable theoretical tool for aerodynamic configuration optimization of a high-speed aircraft with a wider speed range, and the method has important engineering guiding significance for thermal protection system design of a new-generation reusable aircraft.

The invention relates to an additive repairing method for a special-shaped curved surface of a large-size thin-wall component, belongs to the technical field of metal additive manufacturing, and solves the problems of low profile precision, poor aerodynamic layout and difficulty in meeting high-speed flight due to local deformation of the special-shaped curved surface of the large-size thin-wall SLM component. The invention discloses a large-size ultrathin component special-shaped curved surface additive repairing method. The method comprises the following steps that S1, a component profile is scanned and compared with a standard model; s2, polishing the surface of a product repair area; s3, repairing path planning; s4, preparing equipment and raw materials; s5, component laser additive repairing is conducted; s6, detecting a cladding layer; s7, heat treatment of the component; and S8, product detection. And on-demand repair and high-precision shape control of the SLM component in the complex geometrical shape are achieved. Local concave low-stress repair of the large-size thin-wall component is achieved, the method is suitable for large-scale engineering repair manufacturing of the component, and the problem of secondary deformation generated in the repair process of surfacing, traditional laser cladding and other technological methods is effectively solved.

The invention discloses a preparation method of an interface self-reaction strong chemical bonding radiant heat protection multilayer gradient coating, and belongs to the technical field of coating surface modification. The invention aims to solve the problems that the existing coating is high in preparation temperature, long in period, high in cost and not suitable for industrial batch production, the interface bonding between the coating and a nickel-based high-temperature alloy matrix is poor, the coating is easy to crack and even lose efficacy due to mismatching of thermal expansion under high-temperature service, and the thermal protection performance of the traditional coating is insufficient. The interface self-reaction strong-chemical-bonding radiant heat protection multilayer gradient coating and a nickel-based high-temperature substrate are subjected to self-reaction at the interface to form strong chemical bonding, the bonding strength of the coating reaches 15 MPa or above, the emissivity value of the coating reaches up to 0.9, the structure of the coating is kept complete and does not fall off after the coating is oxidized in 1250 DEG C high-temperature static air for 30 h, and the coating has the advantages of being simple in preparation process, low in cost and the like. The method has application potential in a nickel-based high-temperature alloy anti-oxidation radiation thermal protection multifunctional integrated metal thermal protection system for a high-speed aircraft.

The invention discloses a medium-high temperature heat-power conversion working medium for a high-speed aircraft as well as a preparation method and application of the medium-high temperature heat-power conversion working medium. The medium-high temperature heat-power conversion working medium comprises the following components in percentage by mass: 10-90% of ammonia, 10-70% of multi-component alcohol and the balance of water, the multi-component alcohol comprises at least two of ethanol, isopropanol, ethylene glycol, n-propyl alcohol and n-butyl alcohol; and the mass of each alcohol in the multi-component alcohol is 5-40% of the total mass of the medium-high temperature heat-work conversion working medium. Ammonia, multi-component alcohol and water are combined for the first time, a double-bed catalyst is combined, physical heat absorption can be utilized, heat absorption can be enhanced through chemical reaction, the heat sink can reach 7.0 MJ / kg or above, and the cooling agent can be used for pneumatic heat absorption of a high-speed aircraft with the speed higher than 5 Mach and is high in cooling capacity and low in cost. The method is suitable for a thermal protection system of a pneumatic heating component with higher heat flux and continuous long time; meanwhile, high-temperature and high-pressure small-molecule gas is generated, thrust can be generated to do work, and the maneuverability of the high-speed aircraft is effectively enhanced.

The invention discloses a hypersonic flow instability multi-field coupling prediction method and system, and relates to the technical field of modern high-speed aircraft design, and the method comprises the steps: obtaining 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 multi-physical field parameters as input, setting graph nodes in a shock wave influence region, a boundary layer region and an intersection region thereof, and calculating connection weights among the nodes; performing multi-scale decomposition on the instability sensitive graph data structure, and extracting instability modes of a shock wave scale, a vortex scale and a molecular scale; screening the instability modes by adopting a hypersonic velocity stability criterion, and determining a dominant instability mode; and constructing an instability evolution prediction model based on the dominant instability mode, and outputting a spatio-temporal evolution rule of hypersonic flow instability. According to the method, the problem that the flow instability phenomenon in the hypersonic velocity complex multi-physical field environment is difficult to accurately predict can be solved.

The invention discloses a high-thermal-conductivity foam metal supported catalyst as well as a preparation method and application thereof. The supported catalyst comprises a foam metal carrier, an active metal component and an auxiliary agent, wherein the active metal component and the auxiliary agent are supported on the foam metal carrier; the foam metal carrier comprises one of foam nickel, foam copper and foam titanium; the active metal components comprise at least two of Cu, Fe, Ni, Mo, Mn, Rh, Pt and Pd; the auxiliary agent comprises one or more than two of La, Mg, Zr and Ce. The high-thermal-conductivity foam metal supported catalyst disclosed by the invention can be used for a heat-work conversion reactor of a high-speed aircraft, and can be used for rapidly catalyzing a multi-component alcohol and water system to generate a heat-absorbing chemical reforming reaction by strengthening heat conduction and constructing an efficient active center; harmful aerodynamic heat is absorbed, high-temperature and high-pressure small-molecule gas (hydrogen, carbon monoxide and the like) is generated and used for doing work to generate thrust, and aerodynamic heat management and heat-work conversion of the high-speed aircraft are achieved.

The invention relates to the field of aircraft wing design, in particular to a multi-stage telescopic wing for improving the flight performance of an aircraft, the extension length and the area of the wing are changed through multi-stage stretching of the wing, and different requirements of the aircraft for the wing in low-speed and high-speed flight states are met. When the aircraft takes off, lands and flies at a low speed, the telescopic wings stretch out to increase the area of the wings and enable the aircraft to obtain enough lift force, and when the aircraft flies at a high speed, the telescopic wings retract to reduce the extension length and area of the wings and enable the aircraft to obtain enough lift force and reduce the resistance of the wings, so that the extension length and area of the wings are changed through extension of the wings; and the aircraft can simultaneously meet different requirements on the extension length and the area of the wings in low-speed and high-speed flight states.

The embodiment of the invention provides a modeling method and system for aerodynamic efficiency of a nonlinear control surface of a high-speed aircraft, and the method comprises the steps: carrying out the numerical simulation of aerodynamic data of the high-speed aircraft at different control surface deflection angles through employing a computational fluid dynamics (CFD) method, building a control surface aerodynamic efficiency data set in a high-speed flight scene, comprising a sample set and a verification set; taking the sample set as an input variable, and constructing a high-speed aircraft nonlinear control surface aerodynamic efficiency prediction model; and performing verification and error evaluation on the constructed nonlinear control surface efficiency aerodynamic model of the high-speed aircraft by adopting the verification set, and updating model parameters to obtain a final control surface aerodynamic efficiency model. According to the method, the aerodynamic efficiency of the rapid nonlinear control surface of the high-speed aircraft can be rapidly predicted, and a rapid control surface aerodynamic efficiency prediction means is provided for design of a control system in a high-speed flight scene.

The invention discloses a cross-domain variable configuration high-speed aircraft adaptive attitude control method and system based on neural network identification, and a storage medium, relates to the technical field of aircraft control, and aims to solve the problems that in an existing method, unified parametric modeling needs to be carried out on aircraft pneumatics in advance, and parameters need to be identified; the attitude control problem caused by aerodynamic uncertainty cannot be solved under the condition that the aerodynamic parameter model of the cross-domain variable-configuration high-speed aircraft is completely unknown. A three-channel neural network model is constructed and comprises a pitching channel neural network model, a yawing channel neural network model and a rolling channel neural network model; a loss function is constructed based on the influence sensitivity weighting of the rudder deviation error on the attitude change, and a three-channel rudder deviation solver neural network model is obtained through training; and an online transfer learning sample library is constructed for online transfer learning, aerodynamic torque is solved based on a backstepping method controller, and finally a yaw instruction is obtained through solving.

The invention discloses a high-speed aircraft folding rudder aerodynamic configuration design method based on aerodynamic thermal constraint, which comprises the following steps of: determining a rudder leading edge radius and a rudder leading edge sweepback angle by taking a current state nominal trajectory as a basis and taking a rudder leading edge thermal response temperature as a judgment standard; secondly, the cone angle of one conical surface of the folded rudder is determined through iterative calculation, so that the chordwise size of the high-temperature-resistant refractory metal material and the thickness of a skin and a heat-proof coating at the cone position are designed; and thirdly, through the design of the interference position of the arched shock wave of the heat-proof cup, the gap between the fixed rudder and the movable rudder avoids the interference area of the arched shock wave, so that the chordwise distance between the folding gap and the rudder shaft is determined. By means of the aerodynamic configuration design method, the radius of the front edge of the rudder can be as low as 3 mm, the taper angle is as low as 4 degrees, it can be guaranteed that ablation deformation does not occur when the maximum heat flow of the front edge of the rudder does not exceed 10 MW / m < 2 >, the resistance characteristic of a high-speed aircraft can be effectively reduced, the characteristic of near-zero ablation is achieved, and the good aerodynamic configuration is maintained.

The invention discloses a bimorphic high-speed unmanned aerial vehicle capable of taking off and landing on the water surface and flying in ground effect and a control method, and belongs to the technical field of unmanned aerial vehicles, the bimorphic high-speed unmanned aerial vehicle comprises a vehicle body and a control device, the vehicle body is provided with an adjustable turbine jet driving mechanism and telescopic folding wings, and the tail of the vehicle body is provided with an empennage and a tail turbine jet driving mechanism. Meanwhile, the invention discloses a control method based on the two-form aircraft capable of taking off and landing on the water surface and flying in the ground effect, and the two-form high-speed unmanned aerial vehicle capable of taking off and landing on the water surface and flying in the ground effect and the control method are adopted. By adjusting the inclination angle, the distance and the height of the two front turbojet engines, operation and conversion of various forms of water sailing, ground running, ground effect flight and height flight are achieved, the use scenes of the unmanned aerial vehicle are greatly enriched, and the requirement for high-speed flight is met by adopting the telescopic folding wings.

The utility model discloses a high-speed large-maneuvering target drone, which relates to the field of unmanned aerial vehicles and comprises a drone body, wings, a horizontal tail, a vertical tail and an air inlet channel. The fuselage can reduce resistance to a great extent to realize high-speed flight; the wings adopt sweepback upper single wings to reduce the resistance during high-speed flight, the problem of overlarge transverse stability caused by sweepback and high placement of the wings can be avoided through the downward reverse design, and meanwhile, the wings can be integrally disassembled and assembled and have better operability and convenience; the control efficiency of a full-motion horizontal tail is high, and the heading stability under a large-angle-of-attack attitude can be improved through the design of the inverted-V-shaped tail; the vertical fin without the control surface can improve the course stability and ensure the matching property of the transverse course stability, and is light in structural weight and small in flight resistance; the bearing separation channel type air inlet channel can avoid landing damage of the air inlet channel in the parachute landing recovery process, and meanwhile the requirements for high total pressure recovery, low air inlet distortion, radar stealth and the like are met.

The invention discloses a GAN-based high-speed aerodynamic heat flow field reconstruction method and system, and relates to the technical field of analogue simulation, and the method comprises the steps: collecting historical heat flow field data of a high-speed aircraft under different flight conditions, and constructing a real aerodynamic heat flow field data set; establishing a heat flow field generator and a discriminator through feature extraction and simulation discrimination training; performing grid initialization and adaptive adjustment according to the distribution change of the heat flow field, and dividing into N hierarchical reconstruction regions; alternately training the generator and the discriminator in each region by adopting a balance training strategy, and constructing N heat flow field reconstruction networks; and monitoring flight condition update data, optimizing by using each reconstruction network, and generating a simulated heat flow field. Therefore, the technical effects of simplifying parameter configuration, reducing process cost and improving timeliness are achieved.

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 invention relates to the technical field of man-machine interaction, and discloses a wearable multi-channel physiological parameter acquisition system which comprises wearable physiological parameter acquisition equipment and preprocessing software. The technical problems that most of existing physiological state collecting devices are used in a laboratory environment, the use preparation time is long, wearing is inconvenient, and the working requirement in the actual high-speed cabin environment cannot be met are solved, collecting of physiological data of a driver in the high-speed flight or ground driving environment is achieved, and the working efficiency is improved. And data input is provided for evaluating the fatigue state and the cognitive state of the driver in real time.

The invention provides a cluster path planning method and system based on an improved A-star algorithm and reinforcement learning, and belongs to the field of unmanned aerial vehicle cluster control. The problems of collision and failure in a high-dynamic environment caused by the fact that cluster volume is not considered in existing traditional path planning are solved. According to the invention, path planning is carried out based on an improved A * path planning algorithm, and an optimal path is generated by considering constraint conditions; by setting a local target point, the high-speed aircraft is guided to fly along the shortest path; constructing a reward function in combination with an SAC reinforcement learning algorithm to guide the high-speed aircraft to fly to the target; designing a target flight reward according to the constraint condition of the high-speed aircraft; and a hierarchical control framework based on improved A-star and deep reinforcement learning is adopted to further optimize a cluster flight path and obstacle avoidance capability. Accurate path planning and stable formation control of the cluster are ensured in the complex obstacle environment, and the obstacle avoidance capability and cooperative control of the unmanned aerial vehicle cluster in the complex environment are effectively improved.

The invention discloses a distributed active thermal protection and heat-work conversion system and a using method thereof. The system comprises a distributed liquid cooling plate, a heat exchanger, a supply electromagnetic valve, a heat-work conversion plate, a storage box, a propulsion electromagnetic valve, an oxygen supplementation electromagnetic valve, an oxygen supplementation system and a thrust engine. The distributed liquid cooling plate is filled with a high-density liquid cooling working medium; the heat-work conversion plate is filled with a catalyst; an outlet of the distributed liquid cooling plate is connected with an inlet of the heat exchanger, an outlet of the heat exchanger is connected with an inlet of the heat-work conversion plate through the supply electromagnetic valve, an outlet of the heat-work conversion plate is connected with an inlet of the storage box, and an outlet of the storage box is connected with the first end of the tee joint through the propelling electromagnetic valve. The second end of the tee joint is connected with an oxygen supplementing system through an oxygen supplementing electromagnetic valve, the third end of the tee joint is connected with a thrust engine, and an outlet of the thrust engine is communicated with the outside. The heat absorption capacity of the thermal protection system is remarkably improved, thrust can be generated to do work, and the thermal protection system has wide application prospects in high-speed aircrafts.

The embodiment of the invention discloses a high-speed flight-oriented extraterrestrial planet atmosphere collection control device and method, and relates to the technical field of control or adjustment of non-electrical variables in aerospace equipment. A sampling port is connected with an air inlet channel, and the air inlet channel is connected with each air storage channel; the gas storage valve is mounted at the front end of the gas storage channel and used for controlling opening and closing of the gas storage channel; the air storage channels are connected with the corresponding air storage tanks, the six air storage channels and the six air storage tanks are arranged, and the adjusting devices are arranged in the air storage tanks and conduct control based on non-electric quantities such as temperature, pressure and air speed. The device is suitable for an extraterrestrial planet atmosphere collection task in a high-speed flight state, the planet spacecraft invades the planet atmosphere for multiple times, different target gases are collected in multiple times and stored and packaged respectively, extreme temperature and pressure caused by hypersonic flight can be dealt with, the state during gas collection can be kept, and the collection efficiency is improved. And the gas can be conveniently detected and utilized after being brought back to the ground.

The present invention discloses a surface-reinforced aerogel composite material and a preparation method thereof. The prepared aerogel has the characteristics of light weight, high temperature resistance, low thermal conductivity, high strength, simple structure, etc., can withstand high temperatures, resist high-speed airflow scouring and exert excellent thermal insulation performance, and can be used as a thermal protection material in the aerospace field and other related fields. It can solve the problem that the structural strength of aerogel composite materials in the prior art is usually 2 to 3 orders of magnitude different from that of corresponding dense materials, and they cannot withstand high-speed airflow scouring on the surface of the fuselage of high-speed aircraft, and therefore cannot be used as heat-proof materials.

The invention discloses a multifunctional sensor of an FADS and a total temperature estimation method. The multifunctional sensor comprises a pressure sensor, an adapter, a temperature sensor and a heat flow sensor. The air entraining cover is arranged at the upper end of the base, and an inner cavity is formed between the air entraining cover and the base; a pressure measuring hole is formed in the outer surface of the air entraining cover and is communicated with the inner cavity; a T-shaped air guide hole is formed in the base; an air entraining channel is arranged in the adapter, and the upper end of the air entraining channel is connected with the lower end of the air entraining hole; the lower end of the air entraining channel is connected with an external air path and is used for transmitting pressure; the sensing ends of the temperature sensor and the heat flow sensor are located in the inner cavity and used for measuring the regional temperature in the inner cavity and measuring the heat flow of the pressure sensor. The device has a dustproof and anti-blocking structural design, and integrates the functions of measuring pressure, total temperature, heat flow and static temperature. The device is simple in structure, convenient to install and low in manufacturing cost, and can be applied to an embedded atmosphere data system of a high-speed 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 the technical field of aircrafts, in particular to an air-breathing power system and method for a Mars surface aircraft, and the air-breathing power system comprises an air inlet channel, an isolation section, a combustion chamber, a cooling channel, a spray pipe, a heat exchanger and a fuel device, the fuel device comprises a liquid fuel tank, a high-pressure pump, a magnesium powder storage tank, a powder feeder and a gasification nozzle. The air-breathing power system of the aircraft is designed aiming at the fact that gas on the surface of Mars is mainly composed of CO2, the air-breathing power system is provided with magnesium powder capable of reacting with CO2 and methane or hydrogen carrying the magnesium powder to be sprayed into a combustion chamber, CO2 needed by combustion is provided through compressed air in the high-speed flight process, and the combustion efficiency is improved. The air-breathing type power system does not need to carry a large amount of fuel and oxidizing agents like a common rocket engine, and the specific impulse of the engine of the air-breathing type power system is far higher than that of the engine of a common rocket.

The invention provides a wave rider capable of vertically ascending and descending, and belongs to the technical field of wide-speed-range aircrafts, the wave rider specifically comprises a front edge shock wave generating surface, a central fuselage and sweepback wings located on the two sides of the central fuselage, the two sweepback wings are provided with mounting holes penetrating through the upper surface and the lower surface of the wave rider, and lift fans are mounted in the mounting holes; the lift fans are embedded into the mounting holes and used for providing vertical lift force for the wave rider, the axes of the lift fans are parallel to the vertical axis of the wave rider, and the lift fans on the two sweepback wings are symmetrically distributed on the two sides of the central fuselage. According to the treatment scheme, the vertical lifting capacity is provided for the wave rider, and the high-speed flight aerodynamic performance is considered.