51 results about "Gliding flight" patented technology

The invention belongs to the field of craft design, and relates to a craft which can vertically take off and land and can also carry out a gliding flight by using fixed wings. According to the invention, a craft which can concurrently have the advantages of flexibility and convenience of multi-rotor flight and the advantages of large load and long endurance of fixed wing aircrafts is designed. According to the invention, a multi-rotor craft is taken as a basic form, and fixed wings are taken as auxiliary parts. The fixed wings can be collapsed and expanded, folded and flatted, or overturned for resetting as required. The craft can be powered by multiple rotors or power mechanisms, and rotor parts can also be overturned or covered so as to provide a better aerodynamic performance and provide forward or rising power. In the process of rising or environmental constraint, the craft is mainly powered by rotor mechanisms until flying to a stable height, the fixed wings can be expanded in proper condition, the craft is in a gliding state, and the rotors (or other power supply devices) can be overturned for propelling. After flying to a certain height, the craft carries out gliding by the fixed wings.

The invention discloses a thermal environment design method of space between control rudders of a hypersonic flight vehicle. The method comprises the steps of on account of a simplified appearance of the hypersonic flight vehicle, adopting a prediction method of aerodynamic heat engineering to conduct a prediction for aerodynamic-heat thermal environment and obtain a prediction result of aerodynamic heat engineering; according to the prediction result of aerodynamic heat engineering, confirming flow pattern changes of a rudder shaft section of a control rudder along a trajectory, and conducting numerical calculation of thermal environment of the flight vehicle on multiple groups of typical trajectories under a condition of different flow patterns according to the trajectory time period of flow pattern changes, then obtaining heat flow distribution of the space between the control rudders of the flight vehicle; choosing a laminar flow state to conduct numerical calculation of thermal environment of the space between the control rudders, and modifying the prediction result of aerodynamic heat engineering according to the calculation result; designing the thermal environment of the space between the control rudders along a trajectory according to the modified result. The thermal environment design method of space between control rudders of a hypersonic flight vehicle resolves the problems that the flow patterns of space between control rudders of a hypersonic gliding flight vehicle are too complex to predict under a ballistic condition, and the thermal environment is so tough that a part of heat-proofing risks are hard to be estimated.

A high-altitude long-endurance airship with a top surface that is highly emissive of infrared radiation and a bottom surface that is highly absorptive of infrared radiation. Movable displacer blankets inside the airship separate the upper and lower portions of the airship. Lifting gas in the airship is warmed by radiation from the earth when the displacer blankets are in their upper position. Lifting gas is cooled by radiation to space when the displacer blankets are in their lower position. The whole airship is a heat engine. By expanding the volume of lifting gas when it is relatively warm and compressing the lifting gas when it is relatively cold, net power output can be recovered in the form of electric power. The overall configuration of the preferred airship is a variable-thickness flying wing. If the whole airship is alternately expanded and compressed, imbalance between its weight and buoyancy allow gliding flight.

The invention relates to an aircraft landing monitoring guiding system. According to the aircraft landing monitoring guiding system, firstly, an ideal landing point, a stern line and a runway finishing line are set on a runway; secondly, a camera (6) is selected and is mounted outside the runway at one side of a transverse line of the runway finishing line, a view center of the third camera (6) aligns a far-end point of an ideal glide path; an optical landing assisting system (5) is mounted outside the runway at a same side as the camera (6); a commander (2) and a monitor (3) are at same side of the runway as the camera (6) and the optical landing assisting system (5); the camera (6) is connected with a control computer (4) and a monitor (3) through signal wires; a whole flight process of an aircraft (1) along the ideal glide path is observed by the commander (2), images of the aircraft (1) during gliding flight are acquired from the monitor (3); high-low and left-right deviation of the aircraft (1) is determined by the commander (2), and a pilot is prompted to land the aircraft (1).

A micro aerial vehicle capable of controlled transitory or sustained gliding flight. The vehicle includes a fuselage. A pair of articulated wings are forward of a center of gravity of the vehicle, the wings being articulated and having trailing edge flaps, and having actuators for controlling the dihedral angles of the wings and the flaps for effective yaw control across the flight envelope. The dihedral angles can be varied symmetrically on both wings to control the aircraft speed independently of the angle of attack and flight-path angle, while an asymmetric dihedral setting can be used to control yaw and the actuators control the dihedral settings of each wing independently. The aircraft lacks a vertical tail or other vertical stabilizer.

The invention discloses a troposphere intermittent gliding flight control method and belongs to the flight control technical field. The troposphere intermittent gliding flight control method is particularly applied to an aircraft to increase flight time or range of the aircraft by reasonably using updrafts in a troposphere region. The method includes the following steps that: step one, heat updraft modeling and judgment are carried out; step two, aircraft modeling is carried out, and a control law when updrafts do not exist is designed; and step three, intermittent gliding flight control is carried out. According to the control method of the invention, atmospheric energy is utilized to perform intermittent gliding flight, and therefore, the flight time and range of the aircraft can be improved, and the control method has important significance for the aircraft to realize long-flight time, long-range and attack patrol operational goals. The control method of the invention can be widely applied to small-sized unmanned aerial vehicles and small-sized cruise missiles, and can effectively improve the combat effectiveness of the unmanned aerial vehicles and cruise missiles.

A high-altitude long-endurance airship with a top surface that is highly emissive of infrared radiation and a bottom surface that is highly absorptive of infrared radiation. Movable displacer blankets inside the airship separate the upper and lower portions of the airship. Lifting gas in the airship is warmed by radiation from the earth when the displacer blankets are in their upper position. Lifting gas is cooled by radiation to space when the displacer blankets are in their lower position. The whole airship is a heat engine. By expanding the volume of lifting gas when it is relatively warm and compressing the lifting gas when it is relatively cold, net power output can be recovered in the form of electric power. The overall configuration of the preferred airship is a variable-thickness flying wing. If the whole airship is alternately expanded and compressed, imbalance between its weight and buoyancy allow gliding flight.

The invention relates to an auto-rotation gliding control method of an unmanned helicopter. The method includes design of an auto-rotation gliding flight section, design of a flight control law and simulation verification in a condition with engine failure and scull failure of the unmanned helicopter. According to the invention, control of safe landing of the unmanned helicopter is simulated in the condition with engine failure and scull failure. Through analyzing the auto-rotation flight feature of the unmanned helicopter, the flight section in an engine failure state is designed and a motion command is generated. By adopting a classic control method, control law parameter regulation is performed according to the auto-rotation flight features and accurate tracking of the motion command is realized. The method is tested and verified in a simulation environment. The landing state meets safety requirements and the validity of the control method is proved.

Characterized in that in a first stage are loaded or produced on board in a mother aircraft at least a product or mixed of products to act on forest fires, pests or atmospheric phenomena; in a second stage the said product is stored in a glider container with gliding flight ability and equipped with a precision guidance system; in a third stage the glider container is dropped from the mother aircraft and guided towards the release point of the load; in a fourth stage the glider container releases of its interior the load of product on the release point, and in a fifth stage the glider container is recovered for a following reusing.

A propulsion module configured to launch a craft into space including a recoverable module and a non-recovered part, which is secured to the recoverable module at launch. The recoverable module includes a propulsive system to launch the craft, systems for command and control of the propulsive system, a subsonic flight propulsion motor, airfoils for the subsonic flight, a landing gear, and a braking parachute. The recoverable module is installed in a lower position of the craft when the craft is in the launch position. The non-recovered part includes at least one tank to feed the propulsive system. The recoverable module and the non-recovered part are configured to be separated when the propulsion module reaches a given altitude, and the recoverable module is capable of landing in a controlled fashion after a coasting flight, for example for a return to the launch site.

Characterized in that in a first stage are loaded or produced on board in a mother aircraft at least a product or mixed of products to act on forest fires, pests or atmospheric phenomena; in a second stage the said product is stored in a glider container with gliding flight ability and equipped with a precision guidance system; in a third stage the glider container is dropped from the mother aircraft and guided towards the release point of the load; in a fourth stage the glider container releases of its interior the load of product on the release point, and in a fifth stage the glider container is recovered for a following reusing.

A rocket-powered spacecraft having a wing which has hinged aft portions which can be elevated about a hinge line. Tail booms extend rearwardly from the outer ends of the aft wing portions, and rudders are mounted at the aft ends of the booms. Each tail boom supports a horizontal tail with an elevon at its trailing edge. In normal flight, the wing aft portions are not elevated, and the wing has a normal airfoil shape. During atmosphere reentry, the wing aft portions are steeply elevated to provide a stable high-drag altitude for the spaceship for speed reduction at low thermal and structural loading. After reentry, the aft wing is returned to an unelevated position which enables gliding flight to a horizontal-runway landing.

An auxiliary propeller folding mechanism of a model airplane glider comprises a steering engine and an auxiliary propeller, wherein the steering engine comprises an oscillating arm, the auxiliary propeller is connected with an output shaft of a power element, the power element is connected with a machine body through a parallel four-connection-bar mechanism, the steering engine is fixed in the machine body, and the oscillating arm is connected with any one connection bar in the parallel four-connection-bar mechanism. A position sensor is arranged at the tail end of the stroke of the power element, the first connection bar or the second connection bar and is connected with the control end of a controller or the steering engine through a signal line. The steering engine drives the parallel four-connection-bar mechanism, and the auxiliary propeller and the power element can contract in the machine body or protrude out of the machine body. In the powerless gliding flight process, the auxiliary propeller and the power element contract in the machine body, and accordingly the aerodynamic performance of the machine body can be remarkably improved, resistance can be reduced and the glider hang time can be prolonged.

The invention discloses a hypersonic aircraft airborne trajectory planning method and device, an aircraft and a medium. A reentry gliding section motion equation of a hypersonic aircraft is used as a training environment of a trajectory planning problem of the hypersonic aircraft; constructing a state space, a strategy action space, an environment dynamic transfer function, a reward function and a discount factor of the trajectory planning problem; and training the neural network of the trajectory planning problem by using a deep reinforcement learning algorithm. And inputting the state space obtained in the flight process into the trained Actor network to obtain a strategy action space, so that the hypersonic aircraft flies to a predetermined target area according to the strategy action space to complete the flight task of reentering the gliding section. According to the method, airborne real-time trajectory planning of the hypersonic flight vehicle in the gliding section is realized by using an artificial intelligence algorithm, the task adaptability of the flight vehicle is improved, and the problem that the trajectory of the hypersonic deformation flight vehicle like gliding flight is difficult to plan on line is effectively solved.

A propulsion module configured to launch a craft into space including a recoverable module and a non-recovered part, which is secured to the recoverable module at launch. The recoverable module includes a propulsive system to launch the craft, systems for command and control of the propulsive system, a subsonic flight propulsion motor, airfoils for the subsonic flight, a landing gear, and a braking parachute. The recoverable module is installed in a lower position of the craft when the craft is in the launch position. The non-recovered part includes at least one tank to feed the propulsive system. The recoverable module and the non-recovered part are configured to be separated when the propulsion module reaches a given altitude, and the recoverable module is capable of landing in a controlled fashion after a coasting flight, for example for a return to the launch site.

This aircraft comprises both thrusters (7, 16, 17) for conventional aircraft flight through the atmosphere and for high-altitude use as a rocket. It is divided into a payload compartment (8) and a compartment (13) chiefly containing the rocket motor propellants. It comprises a long, low-backsweep transverse wing (3) near the rear to provide better wing lift in the dense strata of the atmosphere and allow the aircraft to climb to high altitudes at subsonic speed before the rocket motors are used. Return to earth is in gliding flight or flight controlled in the same way as a conventional aircraft.

The invention belongs to the technical field of flight control, and discloses a control method for the fixed-point landing of a small fixed-wing unmanned aerial vehicle. Compared with conventional automatic landing control methods, the control method mainly fixes the glide pitch angle of unmanned aerial vehicle landing through an elevator; and the opening degree of an engine throttle can be realized by controlling the elevating rates during glide landing. Thus, through the adoption of the control method, the problems of the strong coupling of altitude trajectory and pitch attitude during the glide landing of the fixed-wing unmanned aerial vehicle can be solved, and the controlling of fixed-point landing can be realized; airspeed self-balancing during glide flight can be realized under the circumstance of inaccurate flight performance data of the unmanned aerial vehicle; and the correctness and feasibility of the control method are verified in actual flight of a 40 kg-grade small unmanned aerial vehicle, and high-precision fixed-point landing control effects can be realized.

A concealed autonomous flying miniature unmanned aerial vehicle (UAV) is provided, which comprises a wing, an electronic assembly, a fuselage, a rudder surface, a tail, a steering gear, a connecting rod and a battery. The wing is made of PCB. An electronic assembly is welded on the lower surface of the wing, and the electronic assembly is uniformly distributed on both sides of the wing central axis; The fuselage cover is arranged on the electronic assembly and is connected with the lower surface of the wing; The rudder surface is connected to the tail of the wing; The tail wing is connected with the tail of the upper surface of the wing; The steering gear is embedded at the mounting holes reserved on both sides of the fuselage; The first end of the connecting rod is connected with the steering gear, and the second end of the connecting rod is connected with the rudder surface; The battery is located inside the fuselage. As that pow device is not provided with the design of high-altitude delivery and autonomous glide flight, the unmanned aerial vehicle can work for longer time and the fault tolerance rate of task execution is increased; At the same time, due to the advantages of material and size, it can also keep good concealment, which greatly increases the practical value and tactical value.

The invention discloses a dual-mode unmanned aerial vehicle working method and a dual-mode unmanned aerial vehicle. The two-mode unmanned aerial vehicle is mainly composed of an unmanned aerial vehicle body, propellers, folding main wings, a folding tail wing and a shock absorbing frame. According to the dual-mode unmanned aerial vehicle working method, the dual-mode unmanned aerial vehicle can realize vertical lifting and lowering, horizontal driving flight or horizontal gliding flight, and has the characteristics that the dual-mode unmanned aerial vehicle is easy to lift and lower, is less affected by wind direction and has long flight time and long flight distance.

The invention discloses a four-stage solid carrier rocket trajectory optimization method based on a particle swarm algorithm, and relates to a trajectory optimization control method of a four-stage solid commercial carrier rocket platform. The method aims at a four-stage solid rocket trajectory flight section. The maximum flight attack angle of the first-stage power flight section, the pitching program angular rate of the second-stage power flight section, the gliding flight time of the rocket after second-stage depletion, the pitching program angular rate of the third-stage power flight section, the pitching program angular rate of the fourth-stage gliding flight section, the pitching program angular rate of the fourth-stage power flight section and the pitching program angular rate of the fourth-stage power flight section are changed into the maximum flight attack angle; the flight time of the four-stage sliding section serves as a trajectory design variable needing to be optimized, global optimization is conducted on the seven variables according to a particle swarm optimization algorithm, and global optimal parameters which can meet the injection conditions of the four-stage solid rocket and have the minimum performance index are sought. After initial values and variation ranges of the seven variables are selected, global optimization is carried out on the seven variables by applying a particle swarm optimization algorithm, and global optimal parameters which can meet four-stage solid rocket orbit injection conditions and have minimum performance indexes are sought.

A micro aerial vehicle capable of controlled transitory or sustained gliding flight. The vehicle includes a fuselage. A pair of articulated wings are forward of a center of gravity of the vehicle, the wings being articulated and having trailing edge flaps, and having actuators for controlling the dihedral angles of the wings and the flaps for effective yaw control across the flight envelope. The dihedral angles can be varied symmetrically on both wings to control the aircraft speed independently of the angle of attack and flight-path angle, while an asymmetric dihedral setting can be used to control yaw and the actuators control the dihedral settings of each wing independently. The aircraft lacks a vertical tail or other vertical stabilizer.

A rocket-powered spacecraft having a wing which has hinged aft portions which can be elevated about a hinge line. Tail booms extend rearwardly from the outer ends of the aft wing portions, and rudders are mounted at the aft ends of the booms. Each tail boom supports a horizontal tail with an elevon at its trailing edge. In normal flight, the wing aft portions are not elevated, and the wing has a normal airfoil shape. During atmosphere reentry, the wing aft portions are steeply elevated to provide a stable high-drag altitude for the spaceship for speed reduction at low thermal and structural loading. After reentry, the aft wing is returned to an unelevated position which enables gliding flight to a horizontal-runway landing.

The invention belongs to the technical field of helicopter application, and particularly relates to a design method of a helicopter exhaust pipe shaping spoiler aiming at the problems that in a gliding flight state, when wake flow acts on a vertical fin, the vertical tail pneumatic environment is unstable, oscillating transverse force is generated, and tail screening occurs. The shaping spoiler structure meets the requirement that the leeward area of the exhaust pipe is in smooth transition to the machine body or the leeward face is reduced. The method specifically comprises the following steps: S1, establishing a second-order unsteady display flow field solving model and carrying out gliding state simulation, S2, carrying out mechanism analysis, and determining the reason of tail screening, S3, determining a reference state of design scheme evaluation, S4, forming a design scheme, and S5, carrying out design scheme verification. By means of the design method, the effects of stabilizing wake flow and restraining tail screening can be achieved.

The invention discloses a fixed wing and flapping wing composite aircraft. The aircraft is characterized in that a whole aircraft skin adopts a flexible material, wings with an inner wing structure and an outer wing structure are designed, in the takeoff process, the inner section wings provide stable lifting force, and the outer section wings generate all the thrust and partial lifting force through large flapping and chord direction twisting; and in the flight process, gliding flight is mainly used, the lifting force is provided through the inner section wings and the outer section wings, flapping of the outer wings with the lower flapping frequency is adopted for assistance, and the flight stability of the aircraft is improved through tail wings. According to the fixed wing and flappingwing composite aircraft, the flight attitude of the aircraft is adjusted by changing the center of gravity of a suspension object below the airfoil; and in the emergency state, the outer section flapping wings can stop flapping and are locked in the designated positions and are parallel to the fixed wings to form a fixed wing layout and realize gliding flight. The aircraft has higher stability, safety and comfort, and can improve the lifting force characteristic of the whole miniature aircraft.

The invention discloses a method for controlling glide flying ballistic curve damping. The method comprises the following three major steps: step 1: taking into consideration influence of earth curvature, ignoring influence of earth rotation, studying three degree-of-freedom kinematical equation of gliding of an aircraft, and acquiring condition of furthest distance of powerless gliding of the aircraft; step 2: based on the atmosphere model and the aircraft aerodynamic force model already known, performing force analysis on the longitudinal plane of the aircraft under stable gliding condition, and acquiring stable gliding ballistic curve angle of inclination gamma*; step 3: adding feedback to the control instruction, introducing ballistic curve damping control. The method is a guidance method enabling stable gliding of the aircraft, and is applicable to aircrafts having a glide ballistic curve, including a cannonball of a low speed and a short range as well as a long-distance, reentry, and hypersonic-velocity aircraft.

An electromagnetic gun for warship missiles. The gliding missile is driven in a slender multi-coil gun barrel, and is accelerated by the relay of multiple electromagnetic coils in a long stroke and time. It does not need a lot of electric power to increase the speed of the gliding missile. To the limit, adjust the direction of warship travel and the elevation angle of the multi-coil barrel, and the warship missile electromagnetic gun can aim and shoot at the target. The highest speed of the gliding missile exiting the barrel tube is supersonic speed, which is the initial velocity of the gliding missile. The gliding missile relies on The inertial motion and direction-controllable gliding flight hit the target. The camera of the gliding missile searches and locks the target based on the image information of the target, the infrared sensor based on the heat source information of the target, and the left and right laser radars jointly search and lock the target according to the reflection information of the target. The missile intelligent control box controls the vertical wing to change the left and right flight direction of the missile, and controls the left and right horizontal wings to change the missile's up and down flight direction, so that the gliding missile can automatically track and strike the target.

The present application discloses a method and system for forced landing of a fixed-wing UAV. The method includes: obtaining the current airspeed of the fixed-wing UAV when the fixed-wing UAV enters the forced landing procedure; Adjust the current airspeed to the gliding flight airspeed; obtain the ground terrain information corresponding to the current airspace of the fixed-wing UAV; select the forced landing area to perform the forced landing and glide mission according to the ground terrain information; during the execution of the forced landing and glide mission, collect the area information of the forced landing area , regional ground wetness and softness information, and regional terrain trend information; according to the regional area information, regional ground wetness and softness information, and regional topographic trend information, a landing grounding plan for the fixed-wing UAV is generated; Landing and grounding scheme, control the fixed-wing UAV to land and ground. The application improves the safety of the fixed-wing unmanned aerial vehicle when it is forced to land in the forced landing area, and reduces the damage rate of the fixed-wing unmanned aerial vehicle when it is forced to land.

The invention discloses a multi-flapping-wing micro air vehicle designed by imitating wild goose array formation and gliding flight. The multi-flapping-wing micro air vehicle is in a six-flapping-wing layout form and is composed of a vehicle body and three pairs of flapping wings, the three pairs of flapping wings are in bilateral symmetry and are arranged on the two sides of the vehicle body in a stepped mode, and wing tips of each pair of front flapping wings are flush with the half-span positions of the adjacent rear flapping wings. Wild goose array formation and gliding flight are simulated, and when the aircraft flies at high altitude and high speed, the flapping wing attack angle is fixed for gliding flight; and when high lift force and high thrust are needed, the flapping wings flap obliquely in the same phase to generate lift force and thrust. According to the multi-flapping-wing micro air vehicle imitating wild goose array formation and gliding flight design, the interference effect of wingtip vortexes generated by the front flapping wings on the rear flapping wings is effectively utilized, and the problems that a traditional flapping-wing air vehicle is insufficient in lifting force and low in aerodynamic efficiency are solved.