47 results about "Flight vehicle" patented technology

A system and method is provided for thruster control in a flight vehicle. The system and method uses a proportional derivative matrix control technique to determine thrust commands in a Divert and Attitude control (DAC) system. The proportional derivative matrix control system is configured to receive pitch, roll and yaw commands as inputs, and generate thrust commands as outputs. The performance of the proportional derivative matrix controller is such that the thrust commands can achieve the desired attitude angles quickly and with reduced fuel consumption. The matrix control system can efficiently control a variety of thrusters, including asymmetric thrusters having different moment arms. The matrix control system is particularly suitable to the control of DAC systems with asymmetric thrusters that are configured for attitude control.

The invention discloses a vehicle-mounted aircraft taking-off and landing device. The device comprises an airborne device and a vehicle-mounted device, wherein the airborne device is arranged at the bottom of an aircraft, and comprises a landing convex base, a traction device, a traction ball, an electric winch and a taking-off and landing controller; the vehicle-mounted device is fixed at the top of a vehicle, and comprises a landing concave lawn, a ball dropping hole, a locking device and a detecting feedback device; and the landing convex base of the airborne device and the landing concave lawn of the vehicle-mounted device are funnel-shaped with upward openings, and can fit in an embedding manner. The vehicle-mounted aircraft taking-off and landing device can realize safe and quick alignment, landing and locking of the aircraft on a moving landing platform under the condition of no person intervention, and has the automatic transportation locking and random taking-off capacity to guarantee the reliability of taking-off/landing and transportation fixation in the movement of the aircraft.

The invention discloses an amphibious power propulsion device suitable for sea and air, comprising a frame type retainer, a tilting motor, a power motor fixedly arranged at the middle part of the retainer, as well as propellers for air and propellers for sea, which are coaxially and rotatably arranged at a certain interval, wherein the propellers for air are in driving connection with the power motor through a unidirectional rotation mechanism, the propellers for sea are in driving connection with the power motor, a rotating shaft of the tilting motor is fixed at the outer side of the retainer through a connecting plate, and the rotating axis of the tilting motor is perpendicular to the rotating axis of the propellers for air and the propellers for sea. The invention also provides an amphibious multi-axis aircraft suitable for sea and air. According to the amphibious power propulsion device and the amphibious multi-axis aircraft, the same motor is adopted to simultaneously adapt to two different working environments, namely in the air and under the water, thereby reducing the weight of the amphibious aircraft for sea and air and simplifying the structure of the aircraft; and meanwhile, the tilting motor is used for controlling the direction of power output, thereby increasing the maneuverability of the amphibious aircraft and strengthening the capability of the amphibious aircraft for finishing complex tasks.

The invention discloses a gliding aircraft three-dimensional reentry guidance method based on analytical prediction. The method comprises the following implementation steps of: constructing a three-dimensional flight corridor taking a longitudinal lift-drag ratio, a lateral lift-drag ratio and energy as a standard frame by considering constraint conditions, directly designing a flight profile in the three-dimensional flight corridor based on a prediction model, designing a reasonable heeling overturning strategy, updating the flight profile, and tracking the flight profile to generate a heeling angle instruction and an attack angle instruction. The method is based on a high-precision analysis model; the aircraft target accessibility and the lateral maneuvering task requirements are comprehensively considered; A flight profile and an overturning strategy are directly designed in a three-dimensional flight corridor constructed by considering a plurality of constraint conditions; An attack angle instruction and a roll angle instruction are generated through profile tracking, so that the design of the three-dimensional reentry guidance method is completed, the integrity of dynamic information can be ensured, and the task adaptive capacity of the gliding aircraft under the long-distance large-range maneuvering condition is fully exerted.

The embodiment of the application provides an aircraft guidance control simulation test method and device. The method comprises the steps: respectively building subsystem models of all controlled objects corresponding to a preset guidance control system model of an aircraft based on historical simulation engineering data of the aircraft; establishing a guidance control simulation system of the aircraft in a virtual simulation environment according to the guidance control system model of the aircraft and each subsystem model; performing simulation verification on the operation state of the guidance control simulation system so as to debug operation parameters of the guidance control simulation system according to a simulation verification result; and accessing preset hardware equipment intothe guidance control simulation system to carry out a semi-physical simulation test for the aircraft. According to the application, the integrated simulation test between the guidance control systemand the controlled subsystem of the aircraft can be effectively realized, and the accuracy and reliability of the aircraft guidance control simulation test can be improved by providing an open and extensible virtual integrated simulation and test environment.

A system and method for providing remote access to an on-board aircraft system provides establishing communication between an aircraft server that is located within the aircraft and a remote computer that is located outside of the aircraft. In the aircraft, a request is received to establish a remote session between the remote computer and the aircraft server. A passcode is generated in the aircraft, and is transmitted from a first communication device that is collocated with the aircraft to a second communication device that is collocated with the remote computer. The passcode is then transmitted from the remote computer to the aircraft server and, upon receipt of the passcode by the aircraft server, the remote computer is provided with remote access to the on-board aircraft system.

The invention discloses a boundary layer transition array plane dynamic evolution process determination method based on flight test data. The method comprises the following steps: (1), converting an original measurement result output by a sensor mounted on a surface measurement point of a hypersonic aircraft into heat flow or temperature information at the position of the surface measurement pointof the aircraft, and filtering abnormal measurement point information to obtain available heat flow or temperature information at the surface measurement point of the aircraft; (2), according to theheat flow or temperature information of the available aircraft surface measuring points, obtaining the transition moment of each measuring point; (3), for any moment, judging whether transition occursat each measuring point at the moment or not according to the obtained moment when transition occurs at each measuring point; (4), in a transition measurement time window, selecting a plurality of time points, and for each time point, obtaining a transition array plane image at the time; and (5), binding the transition array plane images at each moment obtained in the step (3) into an animation according to a flight time sequence, and obtaining a transition array plane dynamic evolution process, thereby obtaining a transition area on the surface of the aircraft at each moment.

The invention provides a phase changing flapping mechanism and a dragonfly-imitated flapping wing aircraft. The phase changing flapping mechanism comprises a body bracket, a screw mechanism fixed on the body bracket and a rotary driving mechanism, wherein the screw mechanism comprises a screw, a nut matched with one end of the screw, a sliding rail fixedly connected to one side of the nut, a sliding block matched with the sliding rail and a sliding driving mechanism driving the nut to move back and forth along the sliding rail; and the rotary driving mechanism comprises a driving motor and a transmission mechanism connecting the driving motor and the sliding block. The phase changing flapping mechanism and the dragonfly-imitated flapping wing aircraft are simple in structure, convenient tomanufacture and low in cost, and flapping parameters of the flapping-wing aircraft can be changed in the flight process, so that the flight posture of the aircraft is changed, different flapping functions are realized, the flapping-wing aircraft is enabled to hover for flight, the flexibility is enhanced, and the application value of the dragonfly-imitated flapping-wing aircraft in the fields ofmilitary reconnaissance, complex terrain observation and the like is improved.

The invention provides a four-shaft rotary-wing aircraft. A power device of the four-shaft rotary-wing aircraft is four symmetrically-distributed cycloidal propellers, and each cycloidal propeller isdriven by a separate motor. The rotation direction of the four cycloidal propellers is shown in the figure, and the center of gravity of the four-shaft rotary-wing aircraft is located in the same horizontal plane as the center of the four cycloidal propellers. The four-shaft rotary-wing aircraft a novel aircraft which can vertically take-off and landing and fly forward at a high speed, in addition, the four-shaft rotary-wing aircraft further has the characteristics of high efficiency, high mobility, low noise and so on.

The invention discloses a multi-rotor unmanned plane, and belongs to the technical field of an unmanned plane. The multi-rotor unmanned plane comprises a plane body, a plane arm, a motor, a rotor wing, and a foot stool; the plane arm is connected to the plane body, and the motor is installed on the plane arm; the rotor wing is connected to a rotary shaft of the motor, and the foot stool is arranged beneath the plane body; besides, a suspending rod is further arranged on the plane body, the upper end of the suspending rod is provided with a magnetic absorbing head; a suspending hook is further arranged on the suspending rod. In use, the multi-rotor unmanned plane can be absorbed and suspended on an iron object at a high place through the magnetic absorbing head on the suspending rod, or suspended on an electric wire in the air, branches and other objects through the suspending hook. The lifting force of the multi-rotor unmanned machine in a hanging state is not provided by power energy of the battery, thus the working time of an onboard working device can be prolonged greatly in a rising and landing process. The multi-rotor unmanned plane can be widely applied to overhead monitoring at a fixed point, monitoring, live telecast, environmental protection monitoring, case detection, urban administration, courtyard safeguard, vehicle anti-theft, express delivery, and other domains.

The invention provides a full-frequency wave-absorbing high-temperature heat-insulating stealth integrated paint. The paint comprises: high temperature-resistant reinforced chopped fiber, pressure-resistant hollow microspheres, a wave-absorbing agent, resin and a solvent. According to the full-frequency wave-absorbing high-temperature heat-insulating stealth integrated paint, carbon nanotubes anda magnetic absorbent are effectively compounded and combined, so that the reflection and absorption performance of a coating can be improved. The paint coating has high-temperature resistance, heat insulation and full-frequency wave absorption, can be used on the outer surfaces of weaponry and aircrafts, achieves effective thermal protection on the weaponry and the aircrafts, and also has a full-frequency band stealth effect at the same time. The paint can be applied to various types of airplanes, missiles and the like which have requirements for high-temperature resistance, heat insulation and invisibility, and has important significance for improving the autonomous ability of national defense and military industry production in China and enhancing the comprehensive strength of China.

The invention provides an aircraft aerodynamic layout method and device capable of meeting a wide-speed-domain waverider performance and a medium. The layout method comprises the following steps: establishing a basic shape, and drawing a grid for a computational domain of the basic shape; performing flow field calculation on the calculation domain of the basic shape by adopting a shock wave assembly method, and respectively performing calculation according to the state parameters of the ballistic points to obtain a flow field stability solution and a stable shock wave surface; for shock wave surfaces obtained by calculating flight Mach numbers from large to small, selecting a cutting position from the windward sides in the flow direction of the shock wave surfaces respectively, cutting theshock wave surfaces, taking obtained cutting lines as leading edge lines of the windward sides of the waverider, taking points on the cutting lines as starting points, and obtaining the waverider surfaces through a streamline tracking method. The length and the width of each cutting line are multiplied by a coefficient according to the actual length requirement of the wave rider design, the actual length requirement and the actual width requirement are met, and then the leeside of the wave rider is obtained according to the filling ratio requirement of the wave rider design.

The invention relates to a wind field gradient measuring method. A lifting rope is suspended at the lower end of an aerostat, and a hanging ball is suspended at the tail end of the lifting rope. The change of the lifting rope along with time is not considered, and the measuring method comprises the following steps: 1, the hanging ball moving under a wind field load; meanwhile, the lifting rope being driven to move, taking any point U on the lifting rope; establishing a rectangular plane coordinate system by taking the fixed end point of the lifting rope as an original point 0, the initial balance position of the lifting rope as a Y axis, the vertical upward direction as a positive direction, the direction vertical to the Y axis as an X axis and the horizontal rightward direction as a positive direction, and detecting an included angle theta between the lifting rope at a U point and an OX axis or a tensile force T1 borne by the lifting rope at the U point. The method does not need a large amount of electric energy consumption, replaces a traditional wind speed detection instrument with a model calculation mode, is high in precision and reliability, can effectively measure the gradient of the high-altitude wind field in a low-energy-consumption, long-term and high-precision mode, and can be used for controlling the airship to stay at a low-wind-speed height and reducing energy consumption.

The invention discloses a multi-rotor aircraft with an electric wire supplying power and application of the multi-rotor aircraft, belongs to the technical field of aviation, and particularly relates to a combined-type aircraft. The combined-type aircraft comprises a fixed wing aircraft, the multi-rotor aircraft and a ground power supply device; the structure of the fixed wing aircraft is similar to that of a regular fixed wing aircraft, and the fixed wing aircraft is further provided with a connecting end; the multi-rotor aircraft comprises a machine body, a connecting device, a radar detecting device and a power supply distributor; the ground power supply device comprises a controller, an electric wire collecting and putting implement and the electric wire; one end of the electric wire is fixedly connected to the power supply distributor, the other end of the electric wire is connected to a power source and is twined on the electric wire collecting and putting implement, and the electric wire transmits power supply for the multi-rotor aircraft; the multi-rotor aircraft flies in the sky with the electric wire, and the connecting device joints the connecting end so as to make the multi-rotor aircraft and the fixed wing aircraft connected and combined to load or carry the fixed wing aircraft. According to the multi-rotor aircraft with the electric wire supplying power and the application of the multi-rotor aircraft, the power source is provided for the multi-rotor aircraft through the electric wire, the combination of the multi-rotor aircraft and the fixed wing aircraft is utilized, and a new scheme is provided for the taking off and landing of the fixed wing aircraft.

The invention discloses a reusable two-stage injection aircraft with pneumatic combined structures connected in parallel, which comprises a booster-stage aircraft and a track-stage aircraft, a guide plane for mounting the track-stage aircraft is arranged on the upper surface of the booster-stage aircraft, and the guide plane is arranged along the length direction of booster-stage flight. And the guide plane is seamlessly attached to the lower surface of the orbit-level aircraft. According to the aerodynamic combined structure of the two-stage injection aircraft, the boosting-stage waverider base body and the orbit-stage waverider are constructed in a two-stage injection horizontal separation mode, and the aerodynamic layout of the deformable wings and the vertical tail rudder is added to the boosting-stage waverider base body, so that the aerodynamic combined structure of the two-stage injection aircraft has good wide-speed-range aerodynamic performance and stability performance; the technical risk of interstage separation is effectively reduced, and the method can be used as a scheme for transporting a reciprocating aircraft in the field of aerospace and can be repeatedly used.

The invention belongs to the technical field of aircrafts, in particular relates to a bi-directional powerful aircraft. The bi-directional powerful aircraft comprises an aircraft body, an aircraft frame, a left motor fixing device, a medium motor fixing device, a right motor fixing device, a shaft column, rotor wings and motors. The inner portion of the aircraft body is matched with a control system and airborne equipment. Seven motors are arranged on the aircraft body. Seven rotor wings are arranged on the aircraft body. The left motor fixing device is fixedly installed on the aircraft body.The medium motor fixing device is fixedly installed on the aircraft body. The right motor fixing device is fixedly installed on the aircraft body. Three motors and three rotor wings are arranged on the left motor fixing device. One motor and one rotor wing are arranged on the medium motor fixing device. Three motors and three rotor wings are arranged on the right motor fixing device. The shaft column penetrates through the left motor fixing device, the medium motor fixing device and the right motor fixing device and is fixed on the aircraft body. The bi-directional powerful aircraft is safe and stable, and can accelerate rapidly.

The invention provides a carrying device. The carrying device is used for being mounted on an aircraft body of an aircraft and comprises a gimbal on which an obstacle-avoidance positioning assembly and an inertial measuring assembly of the aircraft are mounted and a shock absorbing assembly mounted on the gimbal, wherein the gimbal is mounted on the aircraft body through the shock absorbing assembly. The obstacle-avoidance positioning assembly and the inertial measuring assembly are rigidly connected through the gimbal, and the obstacle-avoidance positioning assembly and the inertial measuringassembly use the same shock absorbing assembly for shock avoidance, so as to ensure that the data fed back from the obstacle-avoidance positioning assembly and the data fed back from the inertial measuring assembly can be synchronized accurately, to effectively solve the problem that the feedback data of the obstacle-avoidance positioning assembly and the feedback data of the inertial measuring assembly are inconsistent, to ensure the flight attitude of the aircraft, to enable the aircraft to fly smoothly and to improve the satisfaction of users during use. The invention also provides the aircraft.

The invention discloses goods transfer method and system suitable for an area where vehicles are not easy to travel. The method comprises the steps of, controlling a lifting mechanism on an unmanned chassis to rise when it is detected that the unmanned chassis is located below an aircraft, so as to convey goods to the aircraft or receive the goods conveyed by the aircraft, and controlling an adsorption mechanism at the bottom of the aircraft to correspondingly adsorb or release the goods, so that goods handover is completed; and controlling a lifting mechanism on the unmanned chassis to descend when it is detected that the goods handover is finished, and controlling the unmanned chassis and the aircraft to respectively drive to target positions so as to complete goods transfer. According to the method, the influence of the ground effect on the aircraft can be overcome, automatic loading, unloading and transferring of goods can be achieved, and then seamless logistics transferring service is achieved; and meanwhile, the flight advantage of an unmanned aerial vehicle in the low-altitude field is utilized, so that the goods carrying efficiency can be greatly improved.

The invention belongs to the technical field of attitude control, and relates to an aircraft attitude adjustment path optimization method based on rotation. The invention provides an inertia/starlightcombined attitude adjustment star observation sequence aerial autonomous setting method. Rotation freedom degree constraints are divided into three types, namely single-freedom-degree posture adjustment, two-freedom-degree posture adjustment and three-freedom-degree posture adjustment, and aiming at the three types of freedom degree constraints, posture adjustment parameters based on the optimalplatform error angle observation matrix D, namely the optimal design criterion of the optimal rotation angle, are given respectively.

The invention discloses an aircraft fuselage assembly and the aircraft. The fuselage assembly comprises a fuselage, a plurality of vehicle arms and a plurality of supporting arms. The multiple vehiclearms are distributed in the circumferential direction of the vehicle body, each vehicle arm comprises a first arm section and a second arm section, one end of each first arm section is connected withthe vehicle body and fixed relative to the vehicle body, the other end of each first arm section is rotatably connected with the corresponding second arm section, the number of the supporting arms isthe same as that of the vehicle arms, and the supporting arms correspond to the vehicle arms in a one-to-one mode, and each supporting arm is adjacent to the corresponding vehicle arm. One end of each supporting arm is rotatably connected with the vehicle body, the other end of each supporting arm is detachably connected with the second arm section, when the vehicle arms are in the unfolded state, the second arm sections are located on the axial sides of the first arm sections, the second arm sections are connected with the corresponding supporting arms, and when the vehicle arms are in the folded state, the second arm sections are separated from the corresponding supporting arms. The second arm section is located on the radial side of the first arm section. According to the fuselage assembly, the vehicle arms can be folded, and the folding structure is reliable.