446 results about "High-speed flight" patented technology

The invention relates to a data processor system for space-based measurement and control of a high-speed aircraft. The data processor system comprises a power supply module, an interface module, a control module, an encryption module, a forward processing module and a reverse processing module, wherein the power supply module is used for providing the secondary power source for the other modules; the control module is used for receiving control instructions, and sending the control instructions to the other modules; the encryption module is used for encrypting the reverse data stream and decrypting the forward data stream respectively; the forward processing module is used for receiving the mid-frequency signal, completing the receiving and processing link, recovering forward information, and outputting the forward information to an external device through the interface module; the reverse processing module is used for receiving image data, flight state information and remote measurement information, generating the mid-frequency signal after completing the launching and processing link to output the mid-frequency signal to an external radio frequency front end, and eventually sending the mid-frequency signal to a ground system. The data processor system supports the carrier high-dynamic measurement and control of the aircraft, is large in region coverage, supports multiple users, can realize real-time two-way communication with the aircraft, and has certain interference capability.

The invention discloses a high-speed aircraft large-area interlayer heat protection and insulation structure and a forming method thereof, belonging to the field of aerospace aircraft surface layer thermal-protection structures. The interlayer heat protection and insulation structure comprises a grid rib support layer, a heat insulation layer and an outer skin heat protection layer which are sequentially laid from bottom to top on the upper surface of a base of a metal supporting structural part, wherein the grid rib support layer is composed of an inner skin and a plurality of grids; the rib height of the grids is 10mm-12mm, and the sectional width of ribs is 2mm; the heat insulation layer is composed of an aerogel composite material laid in the grid rib support layer; the surface of the heat insulation layer is coated with an adhesive, and then an upper skin heat-protection layer is laid on the surface of the heat insulation layer. The heat protection and insulation structure is applicable to a large-area interlayer between the metal base and a surface layer material of a body of a control cabinet, a load cabinet or an instrument cabinet of a high-speed aircraft, and has the properties of ablation resistance, high-temperature airflow scouring resistance, low thermal conduction, light weight and the like, can meet long-time heat protection and insulation requirements, has good application prospect, and is beneficial to popularization and implementation.

The modular aircraft system includes a single fuselage having a permanently installed empennage and plural sets of wing modules and engine modules, with each wing and engine module optimized for different flight conditions and missions. The fuselage and each of the modules are configured for rapid removal and installation of the modules to minimize downtime for the aircraft. Short wings having relatively low aspect ratio are provided for relatively high speed flight when great endurance and/or weight carrying capacity are not of great concern. Long wings having high aspect ratio are provided for longer range and endurance flights where speed is not absolutely vital. A medium span wing module is also provided. Turboprop, single turbojet, and dual turbojet engine modules are provided for installation depending upon mission requirements for any given flight. The aircraft is primarily adapted for use as an autonomously operated or remotely operated unmanned aerial vehicle.

The present invention relates to an express delivery method and a system based on an unmanned aerial vehicle. The express delivery method based on an unmanned aerial vehicle comprises the following steps: (1) user coordinates and identity identification information are obtained by the unmanned aerial vehicle; and (2) the user identity is identified by the unmanned aerial vehicle, and a control relation with a sender is established after the verification of the identity identification information is legal. The unmanned aerial vehicle is controlled by an express delivery, a receiver or a sender in different times through identifying the user identity, wherein the control authority is divided according to respective identity. The unmanned aerial vehicle may remind a target user while flying close to the circumference of a destination. According to the express delivery method and the system based on the unmanned aerial vehicle provided by the invention, the high-speed flight ability and the cost advantage of the unmanned aerial vehicle may be fully utilized, and a plurality of technical obstacles required to be overcome are avoided, wherein the technical obstacles are caused by directly converting a manual service to a full-automatic service.

A helicopter capable of preventing an increase in size of an airframe and achieving high-speed flight is provided. Provided is an airframe for supporting a main rotor so as to be rotatable, a propeller having a plane of rotation intersecting a plane of rotation of the main rotor, a propeller supporting portion that supports the propeller so as to be movable between positions behind and at the side of the airframe, and a tail disposed on the airframe, having a surface intersecting the plane of rotation of the main rotor.

The invention relates to a high-speed aircraft lifting surface aerodynamic heating structure multidisciplinary optimization design platform belonging to the technical field of modern high-speed aircraft designs. In the invention, by aiming at the demands of lifting surface heat vibration analysis in an aerodynamic heating environment when a high-speed aircraft lifting surface is designed and considering the positive correlation among aerodynamic heating structures, an aerodynamic heating structure integrated analysis method including aerodynamic heating calculation, transient heat conduction analysis, structural heat model analysis and non-constant aerodynamic force and heat vibration analysis is researched, and the aerodynamic heating structure multidisciplinary optimization design taking the most dangerous heat vibration speed as the constraint and taking the lifting surface structural quality as a target function is researched to achieve the goal of the optimization design on the high-speed aircraft lifting surface. The invention breaks through the problem that the multidisciplinary design optimization is carried out without combining with the influence of three aspects of aerodynamic heating on the structure in the traditional high-speed aircraft design, provides a set of high-speed aircraft lifting surface aerodynamic heating structure optimization design methods and waysand has the advantages of simple scheme and high efficiency.

The invention provides a high-speed aircraft having tilting propellers and being capable of taking off and landing vertically and a flight control method of such high-speed aircraft. The high-speed aircraft is characterized in that canard wings are mounted at the head of a fuselage of the high-speed aircraft, wings and a vertical fin are mounted in the front of and above the tail of the fuselage respectively, the tilting propellers are mounted at the outer ends of two tilting arms, the two tilting arms are symmetrically mounted on two sides of the front middle of the fuselage, each tilting arm, along with the corresponding tilting propeller, can be tilted synchronously around a tilting arm axis, and a tail rotor driven by an independent motor is mounted at the tail of the fuselage and can deflect leftwards and rightwards around a vertical fuselage axis. The high-speed aircraft having tilting propellers and being capable of taking off and landing vertically has the advantages that the tilting propellers are mounted at the middle of the fuselage between the canard wings and forward-swept main wings, and in the vertical takeoff and landing stage, downwash airflow of the propellers does not flow through any fixed airfoils with no blockage, so that hovering efficiency of the whole aircraft can be increased; smooth transition of the aircraft from the vertical takeoff and landing to high-speed flying can be realized and control of flight conversion is easier due to small aerodynamic interference between the downwash airflow of the propellers and the fixed airfoils in a tilting process.

The invention discloses a high-speed axisymmetric aircraft composite control method. The high-speed axisymmetric aircraft composite control method is used for solving the technical problem of poor stability in an existing interceptor missile direct force and aerodynamic force composite control method. According to the technical scheme, a varying centroid and pneumatic rudder composite control are adopted, based on defining a varying centroid control mode, an instruct allocation strategy under a composite control scheme is formulated; a dynamical model for control is established through force analysis of a high-speed aircraft; an overall scheme of a varying centroid mechanism is further determined; then ballistic characteristic points are taken to finish design of a control system. According to the high-speed axisymmetric aircraft composite control method, the varying centroid mechanism is introduced, so that the technical problem of poor stability of a background technology control method is effectively avoided, larger control torque is obtained while energy consumption is reduced as much as possible, and reentry flight of the high-speed aircraft for a long time in a large airspace is facilitated; a composite mode with the pneumatic rudder is adopted, so that work pattern of a sliding block executing mechanism is simplified, and system stability is improved.

The invention relates to a dishing vertical takeoff and landing aircraft, namely a dishing magnetic suspension aircraft. The aircraft has an overall layout according to a dishing outline. An upper magnetic suspension ring-wing and a lower magnetic suspension ring-wing are suspended between an engine room 700 and a fixed ring 100 by a magnetic suspension driving device. The elevation angles of an upper ring-wing blade and a lower ring-wing blade are opposite to each other, and the driving forces of ring-wing power-drive magnetic rings (blocks) (2200 and 2600) are opposite to each other. The magnetic suspension ring-wing structure makes use of an electromotive power magnetic suspension driving principle to ensure that upper and lower magnetic suspension ring-wings are not only uprising and hovering rotor wings, but also an effective extension of fixed wings during forward flight. When the 2200 and the 2600 are loaded with electric power, a magnetic force drives the upper and lower ring-wings to rotate reversely to generate a lift force. The steering of the aircraft is also adjusted by the rotating speed of the upper and lower ring-wings. The forward flight is carried out by a driving device 800. The aircraft integrates the advantages of the fixed wing, a helicopter and a motor-less rotorcraft, can take off and land vertically, hover, cruise at a low altitude at a low speed, and can fly at a high speed. The dishing magnetic suspension aircraft overcomes the defects of the aircraft, is not provided with a mechanical device, is simple in control, and is high in safety.

The invention relates to a canard forward-sweep telescoping wing aerodynamic configuration with variable span wing area. The canard forward-sweep telescoping wing aerodynamic configuration comprises a fuselage, canards, wings, and a vertical fin; the wings comprise forward-sweep inner wings and forward-sweep telescopic outer wings having internal connections with the forward-sweep inner wings by a telescoping mechanism; and the telescoping mechanism extends to allow the forward-sweep telescopic outer wings to extend to the outside of the forward-sweep inner wings when the mach number (Ma) of an unmanned aerial vehicle is 0.2, and the telescoping mechanism contracts to allow the forward-sweep telescopic outer wings to contract to the inside of the forward-sweep inner wings when the Ma of the unmanned aerial vehicle is 0.4, wherein the area ratio of the forward-sweep telescopic outer wings to the forward-sweep inner wings is 0.25-0.45. The telescoping wing configuration with variable span wing area adopted in the invention allows the unmanned aerial vehicle to have good aerodynamic performances in scopes of different spatial domains and different speeds and the mobility and the flexibility of the unmanned aerial vehicle to be improved. According to the canard forward-sweep telescoping wing aerodynamic configuration of the invention, when Ma is equal to 0.2 or 0.4, the cruise lift-to-drag ratio of the high aspect ratio aerodynamic configuration is over 20% higher than the cruise lift-to-drag ratio of the low aspect ratio aerodynamic configuration when the unmanned aerial vehicle flies at a low speed, and the cruise lift-to-drag ratio of the low aspect ratio aerodynamic configuration is about 15% higher than the cruise lift-to-drag ratio of the high aspect ratio aerodynamic configuration when the unmanned aerial vehicle flies at a high speed, so the capacities of the cruise flight with the Ma of 0.2 and 0.4 are possessed.

A fuze for a submunition comprises a fuze housing with a stabilizer ribbon for aerodynamic orientation, a fuze slider released by tension on the stabilizer ribbon, an air-powered electric generator extended into the airstream by the fuze slider and powered in flight by high-speed airflow, a MEMS safety and arming device, a fuze circuit board including an explosive fireset, and an electrically initiated firetrain. The fuze is fixed to and communicates explosively with the end of a grenade warhead.

The invention discloses a space arbitrary configuration distributed SAR moving target parameter estimation method based on a cancellation accumulation space-time spectrum and relates to the distributed SAR ground moving target parameter estimation method. The method is used for solving the problems of great impacts caused by signal-to-noise ratio and waste of information when the existing SAR multi-channel processing method carries out parameter estimation, and the specific process of the method is as follows: 1) establishing an extended space-time model of space arbitrary configuration distributed synthetic aperture radar echo signals according to the super-sparse distribution and the high-speed flight characteristic of a spaceborne distributed SAR; 2) taking a conjugate of the extended space-time model obtained in the step 1) as a cancellation factor, and utilizing the space-time accumulation method for establishing an objective function; and 3) utilizing the objective function obtained in the step 2) to carry out the moving target parameter estimation on multi-source signals of the space arbitrary configuration distributed SAR. The method is applicable to the distributed SAR moving target parameter estimation.

The invention provides a new concept high-speed aerocraft propulsion system layout method. Two cycles are built in a propulsion system, namely a Brayton cycle with the air as the working medium and a closed cycle with supercritical state fluid as the working medium. The two cycles are coupled through a supercritical microscopic scale heat exchange technology, a supercritical state fluid turbine and compressor power balance. By adjusting related valves, the propulsion system can be in a turbofan engine model when taking off or flying at a low speed and in a turbine rocket engine model when flying at a high Mach number, so that it is guaranteed that the aerocraft can effectively cruise for a long time in both a subsonic state and a supersonic state. Through the supercritical microscopic scale heat exchange technology, the gas flow temperature at an inlet of a compressor can be effectively reduced when the propulsion system is flying at a high speed, and when the supercritical microscopic scale heat exchange technology is applied in combination with a closed cycle technology, optical distribution of energy of the propulsion system can be achieved. By means of the method, the defects of the propulsion system of an existing high-speed aerocraft are overcome, and working performance of the high-speed aerocraft propulsion system is remarkably improved when Ma ranges from 0 to 5.

The invention discloses a film heat flux sensor and a preparation method thereof. The film heat flux sensor is characterized in that a high temperature-resistant film thermocouple array (thermopile) is prepared on the surface of a substrate by a micromachining technology; a thick heat barrier layer and a thin heat barrier layer are arranged on the high temperature-resistant film thermocouple array; and through sensing the difference between temperatures under the thick heat barrier layer and the thin heat barrier layer, the film heat flux sensor can measure a heat flux according to the height difference of the thick heat barrier layer and the thin heat barrier layer. The film heat flux sensor can measure heat fluxes of middle and outer layer materials of a high-speed aircraft in flight and provides data reference for a heat protection design. The preparation method has simple processes and a reliable structure.