83 results about "Forward flight" patented technology

The invention discloses a distributed jet active flow control-based rotor wind tunnel test method, which belongs to the technical field of aviation aerodynamic wind tunnel tests, and comprises the steps of rotor test bed dynamic characteristic test, manipulation matrix calibration, preposed preparation of rotor model dynamic balance and common taper adjustment, and dynamic balance and common taper adjustment. And then hovering and forward flight wind tunnel tests of distributed jet active flow control are carried out in sequence, test data are collected and processed, and the aerodynamic performance of the rotor wing under the condition that jet control exists or not is compared and analyzed. According to the invention, the industrial blank of the lack of a distributed jet active flow control rotor wind tunnel test method is filled, and key technical support is provided for theoretical research, technical optimization and engineering application of the rotor distributed jet active flow control method.

The invention provides a micro-miniature semi-tilting unmanned aerial vehicle and a control method thereof.The unmanned aerial vehicle comprises a main body structure, a coaxial double-paddle system, a tilting double-rotor system, a full-motion control surface assembly, a flight control system and a power supply assembly; the main body structure is composed of a carbon fiber carbon plate, a carbon tube and a 3D printing photosensitive resin part, and light weight and rigid supporting are achieved; the coaxial double-propeller system provides main lift force and counteracts reverse torsional moment; the double-rotor system is tilted, and a lift force / thrust mode is switched; the pneumatic efficiency of the tilting rotor wing is improved by the full-motion control surface assembly; the control method comprises a low-speed mode and a high-speed forward flight mode, the vertical take-off and landing, hovering and high-speed cruising are integrated, the structure is simplified, the stability is improved, and the cost is low.

A vertical take-off and landing aircraft which uses fixed rotors for both VTOL and forward flight operations. Some of the rotors are forward canted, while some of the rotors are forward facing. The forward canted rotors are tilted forward slightly from vertical and provide some forward propulsion during horizontal flight. The forward facing rotors may be fixed in a forward facing configuration. Forward swept wings allow for rotor placement which brings efficiency during hover operations.

The invention belongs to the field of aviation fuel systems, and provides an integrated multifunctional fuel feed pump with pressure self-sensing and mode self-adaption functions, which comprises an upper pump assembly, a lower pump assembly, a motor controller assembly and a fuel discharge switch assembly, and the upper pump assembly and the lower pump assembly are respectively mounted on the upper side and the lower side of the motor controller assembly; the motor controller assembly provides a power source for the upper pump assembly and the lower pump assembly, the oil drainage switch assembly is installed on a flange plate at the bottom of the lower pump assembly, and the flange plate of the lower pump assembly is installed at the bottom of an oil tank. When the oil liquid completely submerges the oil supply pump, the upper pump assembly and the lower pump assembly simultaneously participate in pressurization work in a self-adaptive manner; when the airplane flies forward and the lower pump assembly is only submerged by oil in the oil tank, the lower pump assembly participates in pressurization work in a self-adaptive manner; when the airplane flies upside down and oil in the oil tank only submerges the upper pump assembly, the upper pump assembly participates in pressurization work in a self-adaptive mode. By means of the built-in pressure sensor, high-precision pressure self-sensing is achieved, and oil supply modes of forward flight and reverse flight of the aircraft are switched and adapted in a self-adaptive mode.

The application discloses a flight control method and device of an aircraft and a medium, and is suitable for the technical field of the aircraft. The current sinking rate and forward flight speed of the aircraft are used to complete control on the lift value and attitude trim moment of the aircraft, the lift value and attitude trim moment are superimposed and distributed to obtain total pitch values of power systems of each propeller, then when the aircraft height satisfies a preset condition, each total pitch value is adjusted to reduce the vertical speed, and when the speed and height satisfy certain conditions, the aircraft attitude is adjusted to a target attitude to complete the aircraft attitude preparation before landing for landing. After the power system of the aircraft fails, stable glide is maintained, the propeller rotating speed in the unpowered state is adjusted through variable total pitch, and then the pulling force of the propeller is changed, so that the sinking rate and attitude of the aircraft in the power failure state are controlled, the disadvantages that the glide handling strategy cannot be implemented in the low-speed and stall states are made up, and additional emergency handling devices such as parachutes are not needed.

The invention relates to the technical field of aircrafts, and discloses an aircraft flight control device which comprises a longitudinal arm and a vertical empennage, the vertical empennage is arranged on the lower side of the tail of the longitudinal arm, and a plurality of lift propellers are further arranged on the upper side of the longitudinal arm in the front-back direction. The longitudinal arms are arranged in the front-back direction of the aircraft; when the aircraft flies forward and lowers the head, the cross section of the longitudinal arm and the horizontal plane is a standard airfoil profile, the projection of the standard airfoil profile on the cross section is the cross section of the arm, and the connecting surface of the vertical empennage and the longitudinal arm is in a vertical empennage type. The aircraft has the advantages of reducing the resistance of the longitudinal arms during forward flight, reducing the course control burden of a power system during high-speed forward flight and the like, and solves the problems that the course control capability is weak, course control is easy to oscillate or diverge, and the maximum forward flight speed is limited.

Disclosed is an electric aircraft, which includes a fuselage and at least one wing coupled to the fuselage. The electric aircraft includes a plurality of tilting fans coupled to the at least one wing, the plurality of tilting fans being configured to move between a vertical lift position and a forward flight position. The electric aircraft includes a plurality of tilting mechanisms coupled with at least one tilting fan. The electric aircraft includes a first actuator coupled to a first subset of the plurality of tilting mechanisms. The first subset of tilting mechanisms are identified among the plurality of tilting mechanisms according to a coupling scheme. The first actuator tilts a first group of tilting fans coupled to the first subset of the plurality of tilting mechanisms simultaneously.

This utility model discloses a waterborne vertical takeoff and landing (VTOL) manned aircraft, comprising a fuselage, wings, a vertical tail, tail floats, outer wing floats, and several rotor propulsion units. This waterborne VTOL manned aircraft is a lightweight seaplane capable of vertical takeoff and landing on water at any time, providing convenience for applications such as water rescue and sightseeing. The multi-float design, combining wing floats and tail floats, ensures the aircraft's stability on water. The propellers are entirely enclosed within duct openings, increasing propeller efficiency while significantly enhancing safety and reducing the risk of injury to personnel on the ground. Furthermore, the multi-rotor and wing-based structure allows both the multi-rotor and wings to provide lift during forward flight, ensuring a certain level of safety while increasing cruising efficiency. The motors employ a dual-winding configuration (or coaxial dual propellers), ensuring a stable landing even if a single winding fails.

Embodiments provide an aircraft having one or more tilting fan assemblies configured to tilt between a forward flight position and a vertical lift position. The aircraft may also include a plurality of lift fan assemblies for vertical movement. The tilt fan assembly may be coupled to a fuselage or wing of an aircraft via one or more tilt mechanisms. A control system coupled to the aircraft may control the one or more tilt mechanisms to move the tilt fan assembly between a forward flight position and a vertical lift position. The tilt fan assembly may be coupled to one or more support structures that couple a fuselage or wings of the aircraft.

Embodiments provide an aircraft with one or more tilting fan assemblies that are configured to tilt between a forward flight position and a vertical lift position. The aircraft may also include a plurality of lift fan assemblies for vertical movement. The tilting fan assemblies may be coupled to the fuselage or wings of the aircraft via one or more tilting mechanisms. A control system coupled to the aircraft may control the one or more tilting mechanisms to move the tilting fan assemblies between the forward flight position and the vertical lift position. The tilting fan assemblies may be coupled to one or more support structures that are coupled the fuselage or wings of the aircraft.

The aircraft can include: an airframe, a tilt mechanism, a payload housing, and can optionally include an impact attenuator, a set of ground support members (e.g., struts), a set of power sources, and a set of control elements. The airframe can include: a set of rotors and a set of support members. By utilizing a larger rotor blade area (and / or larger rotor disc area) and adjusting the blade pitch and RPM, the rotors can augment the lift generated by the aerodynamic profile of the aircraft in the forward flight mode in addition to providing forward thrust. Variants generating lift with the rotors can reduce or eliminate additional control surfaces (e.g., wing flaps, ailerons, ruddervators, elevators, rudder, etc.) on the aircraft since the thrust and motor torque is controllable (thereby indirectly controlling lift) at each rotor, thereby enabling pitch, yaw, and / or roll control during forward flight.

The invention belongs to the field of pneumatic and flight control of tilting rotorcrafts, and relates to a vortex ring boundary protection control method of a tilting rotorcraft. Comprising the steps that firstly, according to the maximum takeoff weight M and the rotor wing radius R of the tilt rotorcraft, the induced speed V1 of rotor wing downwash during hovering of the tilt rotorcraft is determined; 2, according to the induced velocity V1 of rotor wing downwash, the vortex ring boundary of a single rotor wing is determined, and the boundary is a binary quadratic curve composed of a forward flight velocity VX and a vertical velocity Vy; 3, the vortex ring boundary of the whole tilt rotorcraft is determined according to the vortex ring boundary of the single rotor and the flight characteristics of the tilt rotorcraft in the flight process, and the flight characteristics of the tilt rotorcraft in the flight process are that the upper half part of the aircraft entering the binary quadratic curve indicates that the aircraft enters the vortex ring boundary; therefore, whether the airplane enters the upper half part of the binary quadratic curve or not is considered in practice; and 4, reserving a preset safety margin on the basis of the boundary of the vortex ring of the whole tilt rotorcraft, and taking the preset safety margin as an early warning boundary of the vortex ring in the flight process.

The utility model discloses a variable-speed transmission system and a crossed double-rotor helicopter, the variable-speed transmission system comprises a first input shaft, a first clutch and a first output shaft, and the first input shaft is connected with the first output shaft through the first clutch; the speed changing device comprises a first gear train mechanism, a second gear train mechanism and a second clutch, the first gear train mechanism is connected with the second gear train mechanism through the second clutch, the input end of the first gear train mechanism is connected with the first input shaft, and the output end of the second gear train mechanism is connected with the first output shaft; and at least one of the first gear train mechanism and the second gear train mechanism is a speed reducing mechanism. The power output path of the first input shaft is changed through connection or disconnection of the first clutch and the second clutch, and the rotating speeds output by the first output shaft are different. According to the high-speed helicopter adopting the variable-speed transmission system, the purpose of reducing the rotating speed of the rotor wing through the variable-speed transmission system is achieved, and the influence of blade shock waves on the forward side of the rotor wing is reduced.

A hydrogen fuel cell powered electric vertical take-off and landing (eVTOL) aircraft with a high efficiency hydrogen fuel system. The eVTOL aircraft may utilized tilt-up rotors for hover flight, which then transition to a forward facing forward flight configuration. The fuel cell system may use one or more compressors to compress air to a sufficiently high pressure for the fuel cell. Liquid hydrogen may be compressed and then utilized in heat exchangers to cool the compressed air, maintaining the air at a temperature low enough for the fuel cell. The hydrogen may also be used to cool the fuel cell as it is also depressurized prior to its entry in the fuel cell cycle.

The utility model provides a tandem type double-rotor structure which comprises a first rotor assembly and a second rotor assembly, the first rotor assembly is arranged at the front end of an unmanned helicopter body, and the second rotor assembly is arranged at the rear end of the unmanned helicopter body; the bottom ends of rotor shafts of the first rotor assembly and the second rotor assembly are connected with the corresponding mounting bases correspondingly, the mounting bases are connected with the fuselage, and the mounting bases have preset forward inclination angles relative to the fuselage. The front rotor wing and the rear rotor wing incline forwards by a certain angle at the same time, so that when the unmanned helicopter flies forwards at a high speed, the helicopter body can be kept horizontal as much as possible, only the front propeller disc and the rear propeller disc incline forwards to generate forward propelling component force, and therefore the windward area of the helicopter body can still be kept minimum; according to the unmanned helicopter, the resistance of the helicopter body during forward flight can be reduced, the windward area of the propeller disc can be smaller, the forward flight resistance can be further reduced, the aerodynamic performance of the unmanned helicopter is improved, and meanwhile the applicability of the unmanned helicopter is improved.

The invention discloses an aircraft additional air injection structure capable of adjusting the lift ratio and an adjusting method, and belongs to the field of aerospace, the aircraft additional air injection structure comprises an air injection unit, when an aircraft is a complete machine body integrating the wing function, the air injection unit is arranged at the top or the bottom of the complete machine body, and when the aircraft is provided with independent wings, the air injection unit is arranged at the top or the bottom of the complete machine body. The air injection units are arranged at the tops or the bottoms of a fuselage and wings of the aircraft, the installation positions of the air injection units are designed according to the aerodynamic basis of the fuselage and the wings of the whole aircraft, and the air injection units do not interfere with an air injection channel of an original engine. According to the aircraft additional air injection structure with the adjustable lift ratio and the adjusting method, high-speed airflow is injected through the angle-adjustable air injection units installed on the top or the bottom of the whole fuselage or the wings, and the angle and the strength of the injected high-speed airflow are used for assisting and controlling the aircraft to increase the forward flight bearing capacity; the aircraft can also fly upside down, roll horizontally, roll in a barrel, rush upwards, dive and other maneuvers at the bottom of the aircraft.

The application discloses an optimization method of a rolling wing blade connecting rod mechanism, and belongs to the field of rolling wing aircrafts.The blade swing angle is represented by structural parameters of the blade connecting rod mechanism, and the structural parameters of the blade connecting rod mechanism, including the lengths of the connecting rods, the coordinates of key hinge points and configuration bias, are defined as a variable vector; multiple target functions are constructed for multiple key working conditions such as hovering, forward flight and transition, the multiple target functions not only quantify the performance of the connecting rod mechanism itself, such as kinematic fitting error with a target motion law, dynamic stability index and force transmission efficiency index, but also cover aerodynamic efficiency index and aerodynamic acoustic index which are directly related to the performance of the whole machine; and the optimal structural parameters of the blade connecting rod mechanism are obtained by solving the multiple target functions under constraints, so that the global optimization of the structural parameters of the blade connecting rod mechanism under multiple working conditions, multiple targets and multiple constraints can be realized.

This invention discloses a method for calculating the takeoff trajectory of a helicopter in Category A flight. The key technical points include the following calculation steps: S1: Statistically analyze the parameters of a twin-engine helicopter and the parameters of the engines it uses; S2: Based on the statistically analyzed parameters, calculate the power required for forward flight of the twin-engine helicopter and plot N. xu -V0 curve; S3: Calculate the safe takeoff speed of the twin-engine helicopter based on the power required for forward flight and the statistical power of a single helicopter engine; S4: Calculate the lift and forward thrust provided by the helicopter rotor based on the statistical parameters; S5: Calculate the critical takeoff decision point of the twin-engine helicopter and plot the critical takeoff decision point curve; S6: Based on the calculated critical takeoff decision point of the twin-engine helicopter, calculate the takeoff trajectory of the twin-engine helicopter in Category A flight according to the location of the failure point, and plot the Category A takeoff trajectory curve of the helicopter.

The utility model discloses a single-blade rotor type vertical take-off and landing fixed-wing aircraft which comprises a single-blade rotor assembly which penetrates through an aircraft body, is connected with a power system and is used for achieving the functions of vertical take-off and landing and hovering of the aircraft. The rotor wing clasping clamping device is mounted on a vertical tail wing framework of the fuselage main body and is used for clasping or loosening the single-blade rotor wing assembly; the homing driving device is mounted on a middle framework of the fuselage main body and is used for driving the single-blade rotor wing assembly to return to the initial position; the aileron driving devices are mounted on the tail and the side wings of the fuselage main body and are used for assisting the aircraft in completing flight actions such as turning and climbing; the power system is mounted on the fuselage main body and used for driving the single-blade rotor assemblies to move; the propelling assembly is installed on the fuselage and used for providing forward flight thrust. Compared with the prior art, the advantages of long endurance of a fixed-wing aircraft and vertical take-off and landing of a traditional helicopter are combined, and fuel economy and high flight speed are achieved in the cruising process.

A spinner is described that can provide aerodynamic advantages for vertical takeoff and landing (VTOL) aircraft. The spinner can be shaped with multiple faces divided by chines. The chines can run down the edge of the spinner to fairings that can cover rotor blades and / or couplings between rotor blades and the rotor hubs of the aircraft. During forward flight the rotor blades can be folded down into recesses on the spinner. The spinner and fairings can provide aerodynamic advantages in addition to protection for sensitive components within the rotor and blades apparatus. This improves efficiency and allows the aircraft to fly faster.

An aircraft structure comprising: a fuselage; first and second front wings mounted to and / or extending from opposite sides of the fuselage; a continuous rear span defining first and second rear wings and a central stationary connecting portion; a first wing connection member extending between the first front wing and the first rear wing; a second wing connection member extending between the second front wing and the second rear wing; wherein the aft span is supported by a centrally positioned V-tail junction, the V-tail junction defined by a first angled angled arm and a second angled angled arm; a first electric motor and a second electric motor each having a rotor are mounted to each wing, each rotor pivoting between a first configuration for vertical flight and a second configuration for forward flight.

A convertible aircraft is configured to fly in a hover mode and a forward flight mode, the aircraft including a body defining a longitudinal axis and a thruster system coupled to the body and configured to apply forces to the body to fly the aircraft. The thruster system includes a first thruster configured to create a first thrust profile having a fixed direction and a variable amount. The thruster system includes a second thruster configured to create a second thrust profile having a variable direction and a variable amount. The thruster system is configured to operate in a first mode and a second mode. In the first mode, the first thrust profile and the second thrust profile are positive, and in the second mode, the first thrust profile is minimized and the second thrust profile is positive.

The invention discloses a variable rotor duct system and a tilt rotor aircraft, and relates to the technical field of aircrafts, the variable rotor duct system comprises a transmission shaft, a rotor assembly and a duct assembly; the transmission shaft is in transmission connection with an aircraft body; the rotor assemblies are connected to the transmission shaft; the duct assembly comprises a deformation shell and a first driving mechanism, the first driving mechanism is connected to the transmission shaft, and the deformation shell is arranged outside the rotor wing assembly in a surrounding mode in the circumferential direction; the first driving mechanism is in transmission connection with the deformation shell, the first driving mechanism can drive the deformation shell to circumferentially expand in the direction away from the transmission shaft so as to form a duct configuration, and the first driving mechanism can further drive the deformation shell to circumferentially contract in the direction close to the transmission shaft so as to form a fairing configuration. According to the variable rotor duct system and the tilt rotor aircraft provided by the invention, the rotor duct system gives consideration to high hovering efficiency and low forward flight resistance, and the overall aerodynamic efficiency and engineering application value of the tilt rotor aircraft are improved.

The application discloses a control method of a tiltable three-rotor unmanned aerial vehicle, and relates to the technical field of unmanned aerial vehicles. The control method comprises the following steps: acquiring real-time state data of the unmanned aerial vehicle; acquiring a task instruction of the unmanned aerial vehicle; performing double-loop control calculation based on the real-time state data and the task instruction to generate a pseudo control instruction; and performing control distribution calculation based on the pseudo control instruction to generate a motor throttle and a rudder deflection angle instruction of the three-rotor unmanned aerial vehicle. The control method generates the motor throttle and the rudder deflection angle instruction of the three-rotor unmanned aerial vehicle through distribution calculation after double-loop control calculation based on state data and a task instruction, realizes decoupling of pitch attitude control and forward flight control and a control mode of fixed-inclination forward flight of the unmanned aerial vehicle, and solves the problem of single control mode and limited adjustment effect of a conventional multi-rotor unmanned aerial vehicle flight control method.

A vertical takeoff and landing aircraft uses fixed rotors for both VTOL and forward flight operations. The rotors form a synthetic wing and are positioned to achieve high wing span efficiency. The rotors are positioned to shed lift evenly across the synthetic wing span. The synthetic wing can also have narrow forward and aft airfoils that can provide structural support as well as provide lift during forward flight. The wing rotors are tilted forward and provide some forward propulsion during horizontal flight.

The invention discloses an optimization method of a rolling wing blade connecting rod mechanism, and belongs to the field of rolling wing aircrafts, the method comprises the following steps: representing a blade swing angle through structural parameters of the blade connecting rod mechanism, and defining the structural parameters of the blade connecting rod mechanism including the length of each connecting rod, coordinates of a key hinge point and configuration offset as variable vectors; for various key working conditions such as hovering, forward flight and transition, a multi-objective function is constructed, and the multi-objective function not only quantifies the performance of the connecting rod mechanism, such as kinematics fitting errors, dynamic stability indexes and force transmission efficiency indexes of a target motion law, but also covers aerodynamic efficiency indexes and aerodynamic acoustic indexes directly related to the performance of the whole machine. The optimal structural parameters of the blade connecting rod mechanism are obtained by solving the multi-objective function under the constraint set, and global optimization of the structural parameters of the rolling wing blade connecting rod mechanism under the multi-working-condition, multi-objective and multi-constraint conditions can be achieved.

The invention belongs to the technical field of unmanned aerial vehicles, and discloses an oil-driven multi-rotor carrying unmanned aerial vehicle based on vibration reduction and energy absorption and a control method.The oil-driven multi-rotor carrying unmanned aerial vehicle adopts a vertical coaxial double-paddle single-rotor high-low staggered layer layout, and is matched with multiple sets of ball cage supports and engine cabins at the front end and the rear end of the central axis of a fuselage to form a multi-rotor power system; the vertical coaxial double propellers can counteract reaction torque, vibration is reduced from a power source, vibration reduction and energy absorption are achieved through ball cage support wrapping type installation, and the influence of vibration of the oil-driven engine on a fuselage and a carrying structure is effectively relieved. Meanwhile, fixed wings are arranged on the two sides of an engine cabin, a horizontal coaxial double-paddle single rotor wing is inversely installed in front of the fuselage, a power framework with vertical take-off and landing and level flight separated is formed, the vertical rotor wing is responsible for take-off and landing, the horizontal rotor wing provides forward flight thrust, the fixed wings improve the level flight lift force efficiency, and the advantages of large load and long voyage of oil power are considered; vertical take-off and landing and efficient level flight are both considered, and therefore the technical problems existing in the prior art are solved.

The utility model discloses a vertical fixed wing unmanned aerial vehicle which comprises a rack and vertical propellers, four sets of vertical propellers are installed on the side wall of the top end of the rack at equal intervals, a rocket booster is installed at the center position of the top end of the rack, forward propellers are installed on the side wall of the rack, and the rocket booster is installed on the side wall of the rack. Cabins are installed on the two sides of the rack correspondingly, side plates are arranged outside the cabins correspondingly, and rotating sleeves are installed at the ends, close to the cabins, of the side plates correspondingly. According to the vertical fixed-wing unmanned aerial vehicle, the unmanned aerial vehicle is conveniently driven to fly forwards after the vertical fixed-wing unmanned aerial vehicle takes off vertically, the initial speed during forward flying is increased, the undercarriage is conveniently and conveniently folded and unfolded, elastic buffering protection during landing is facilitated, the vibration amplitude during landing is reduced, and the safety of the unmanned aerial vehicle is improved. And the service life of the vertical fixed-wing unmanned aerial vehicle is prolonged.

An aircraft that closely integrates thrust and aerodynamics to achieve VTOL flight, high speed forward flight, and smooth transitions from VTOL to forward flight. The invention combines Electric Ducted Rotors, Pivoting Angled Ducts, Ducted Column Assemblies, and Ventral Control Door assemblies for VTOL and high speed forward flight of an aircraft. In forward flight, the concept uses a plurality of Ducted Rotors arranged in columns to not only provide thrust, but also enhance aerodynamic lift. In VTOL flight and transitioning to forward flight, the PADs in a single Ducted Column Assembly rotate in unison in the pitch axis to create a single thrust vector through the Ventral Control Door Assembly, providing smooth power, controllability, and aircraft orientation throughout transition. Throughout all phases of flight, differential actuation of Electric Ducted Rotors and conventional flight controls provide control.