35results about How to "Slow down the landing speed" patented technology

A jet engine nozzle having a pair of pivoting extension arms and a pair of pivoting shells connected to a jet pipe is provided. The first extension arm is pivotally connected to the jet pipe at a first location, while the second extension arm is pivotally connected to the jet pipe at a second location. Each shell is pivotally connected at both the first and second locations. An axis of rotation associated with the first location is coaxial to an axis of rotation associated with the second location. Independent actuation means are provided to rotate the extension arms and the shells, thereby enabling adjustment of the thrust-vector angle or the exhaust area of the nozzle.

The invention discloses a bionic wing-flapping and bouncing multi-mode movement robot which comprises a rack, a bouncing mechanism, a wing-flapping mechanism and a driving mechanism, wherein the bouncing mechanism, the wing-flapping mechanism and the driving mechanism are arranged on the rack. The bouncing mechanism comprises feet, a folding mechanism, a torsional spring frame and a winch wheel. The wing-flapping mechanism comprises two wing mechanisms and two tail wing mechanisms, wherein the two wing mechanisms and the two tail wing mechanisms are respectively symmetrical in a left-right mode. The driving mechanism comprises a driving motor, a gear transmission mechanism and a ratchet mechanism. The gear transmission mechanism comprises an input gear and an output gear meshed with the input gear. The output gear comprises a tooth-incomplete gear, a crank gear connected with a crank, and a winch wheel gear connected with the winch wheel. The crank gear is meshed with the half side, with complete teeth, of the tooth-incomplete gear, and the winch wheel gear is meshed with the half side, with incomplete teeth, of the tooth-incomplete gear. The driving mechanism with the single motor drives the wing mechanisms and the bouncing mechanism to work, the complexity and the whole mass of the robot are effectively reduced, and the feasibility of the multi-mode movement robot is improved.

Described herein are features for a riser release flaring system for parafoils and other descent flight vehicles for controlled descent and landing of the flight vehicle. The descent flight vehicle may have a payload suspended by a canopy. The descent flight vehicle may be released from a high altitude lighter-than-air (LTA) system, from another system, or may not be associated with any other flight system. The riser release auto flare system is used with the descent system, such as the parafoil, for controlled and safe landing of the payload. Riser lines are released at a controlled rate and for a fixed distance to automatically cause the payload to pull control lines to flare the parafoil and slow a descent and / or forward speed of the vehicle. The riser lines may be released in response to the descent system satisfying a landing criterion, such as altitude.

The invention relates to a recovery system of a helicopter. The recovery system comprises a parachute bay, a switching device and a control module; the parachute bay is fixedly connected with a propeller spindle of the helicopter through a switching device, and is arranged above a rotor wing, wherein the parachute bay comprises a parachute bay cover, a parachute bag, an ejector and a bottom plate of the parachute bay; the parachute bay cover is in a circular or streamline shape; the parachute bay cover is provided with a semi-circular or streamline-shaped spherical shell, and is buckled on the bottom plate of the parachute bay; the parachute bag is stacked in a parachute bay cavity which is formed by the parachute bay cover and the bottom plate of the parachute bay; the ejector is arranged inside the parachute bay cavity, and is fixedly arranged in the center of the bottom plate of the parachute bay; a separation part of the ejector is fixedly connected onto the parachute bay cover; the control module is loaded on the helicopter, and is used for starting the recovery system of the helicopter. The system is used for reducing the landing speed of the helicopter when the helicopter has an accident, so that the landing time is prolonged, personnel on the ground can evacuate conveniently, and the damage of the helicopter when the helicopter hits the ground is reduced.

The invention discloses a high-performance hybrid hullform for an amphibious aircraft. The high-performance hybrid hullform comprises a multi-bilge-line hull, two ground-effect wings and two force-bearing buoys, wherein the hull is tapering in front and rear; the ground-effect wings are symmetrically arranged at the two sides of the middle part of the hull, and the force-bearing buoys are fixed at the outer end. A non-step-fault design is adopted at the middle part of the hull, the hull is provided with three bilge lines, an upper hull between the upper bilge line and the middle bilge line and a middle hull between the middle bilge line and the lower bilge line are in the shape of a linear ramp with a bilge turn respectively, the bilge turns are arranged at the outer sides of the transverse sections, and the ramp angles of the upper hull and the middle hull range from 65 degrees to 75 degrees; and the transverse section of a lower hull is in the shape of a curve ramp. According to the high-performance hybrid hullform disclosed by the invention, the motion response of the amphibious aircraft in a wave is reduced through using the hull with the large ramp angle and multiple bilge lines; and meanwhile, an aerodynamic lift is remarkably increased and a water take-off and landing speed is reduced through using the ground-effect wings and a ground-effect lift-enhancing air cavity, and then the motion response of the amphibious aircraft in the wave is reduced and the pitching motion time of the amphibious aircraft during take-off and landing in the wave is reduced, and the anti-wave capacity of the amphibious aircraft is remarkably improved.

The invention relates to a power-forced parachute opening device of a rocket, which comprises a rocket (1) and a traction belt (3), wherein the rocket (1) is arranged outside a parachute pack (2) and connected with a parachute (4) in the parachute pack (2) by the traction belt (3). Compared with the prior art, the invention has the advantages that the rocket can be used for rapidly pulling the parachute in the preset direction after the rocket is started, thereby greatly shortening the straightening time of the parachute, improving the parachute opening speed, greatly reducing height loss; and the parachute can be rapidly completely opened and work, a plurality of problems of abnormal parachute opening (such as parachute inside out, and the like) can be solved, and the aim of lifesaving is achieved finally, therefore, the reliability of parachute use is improved.

The invention relates to a control method for an aircraft carbon-ceramic wheel braking system with adjustable maximum braking pressure. The braking pressure of the carbon-ceramic wheel braking system is controlled through a drogue chute signal, wherein when the drogue chute signal is 1, the friction coefficient of a carbon-ceramic composite brake pad is normal, and an aircraft is braked and stopped with rated braking pressure through the carbon-ceramic wheel braking system; and when the drogue chute signal is 0, the friction coefficient of the carbon-ceramic composite brake pad becomes lower, the maximum braking pressure braking instruction of the carbon-ceramic wheel braking system is output by an anti-skid control box, and the carbon-ceramic wheel braking system multiplies the rated braking pressure according to a preset proportionality coefficient so that the braking pressure can be maximum. A method for increasing the braking pressure of the carbon-ceramic wheel braking system is used for eliminating the influences caused by decrease of the friction coefficient of the carbon-ceramic composite brake pad on braking torque and braking efficiency, and the problem that the braking efficiency of the carbon-ceramic wheel braking system is low when the landing speed of the aircraft is high or the drogue chute signal is 0 is solved. The braking efficiency and the reliability in the landing process of the aircraft are guaranteed.

In a hydraulic actuator comprising a hydraulic piston, a hydraulic cylinder and a hydraulic pressure source wherein said hydraulic piston has a larger area exposed to pressure on the side of the control chamber than on the side of oil supply chamber of said hydraulic cylinder, wherein said hydraulic piston has, in its outer peripheral part, a slot connecting said oil supply chamber side with the control chamber side in the direction of shaft and the cross sectional area of the slot in the direction of said shaft changes according to the displacement of the hydraulic piston in the direction of its shaft, and the control chamber has a pilot valve, the suction and exhaust valve in a variable valve system of the internal combustion engine are opened and closed by said hydraulic actuator through a coupling and springs for the valve, for variably controlling the opening and closing timing, the opening and closing time, and the lift of the valve, whereby it is possible to provide a variable valve system for internal combustion engine that is a cam-less and sensor-less, and a hydraulic actuator that enable to reduce power consumption and to improve response speed, and to control landing speed so that a smooth landing may be realized.

In a hydraulic actuator comprising a hydraulic piston, a hydraulic cylinder and a hydraulic pressure source wherein said hydraulic piston has a larger area exposed to pressure on the side of the control chamber than on the side of oil supply chamber of said hydraulic cylinder, wherein said hydraulic piston has, in its outer peripheral part, a slot connecting said oil supply chamber side with the control chamber side in the direction of shaft and the cross sectional area of the slot in the direction of said shaft changes according to the displacement of the hydraulic piston in the direction of its shaft, and the control chamber has a pilot valve, the suction and exhaust valve in a variable valve system of the internal combustion engine are opened and closed by said hydraulic actuator through a coupling and springs for the valve, for variably controlling the opening and closing timing, the opening and closing time, and the lift of the valve, whereby it is possible to provide a variable valve system for internal combustion engine that is a cam-less and sensor-less, and a hydraulic actuator that enable to reduce power consumption and to improve response speed, and to control landing speed so that a smooth landing may be realized.

The invention provides a double-suction ring broad-width precision sowing apparatus. The apparatus comprises a seed containing box, a seed discharging device and a seed discharging tube, the seed discharging device comprises a shell, a sealing cover and an inner core, the seed discharging device is a double-suction ring air suction type seed discharging device, the bottom of the seed discharging device is fixedly provided with a hollow baffle block, the seed discharging tube is arc-shaped, and the arc angle theta is not smaller than 40DEG and not greater than 50DEG. When the apparatus works, seeds fall into the seed discharging device from the seed containing box, the inner wall rotates with a driving shaft, the seed discharging device adopts a double-suction ring structure to increase the suction force, the seed containing box is divided into two parts, a negative pressure is generated in the seed discharging device through a high speed fan, seeds adsorbed in recessed round pits rotate with the inner core, the suction force to the seeds decreases and even disappears when the seeds rotate to the fixed hollow baffle block under the seed discharging device, the seeds fall into a seed outlet under the action of a seed scraper and enter the seed discharging tube, and the seed discharging tube is arc-shaped, has a certain angle to the ground and is not vertical to the ground, so the falling speed of the seeds is reduced, deviation from sowing points due to bouncing after seed falling to the ground is avoided, and the sowing precision is improved.

The invention provides a suction wheel type broad-width precious speed distribution apparatus. The apparatus comprises a seed containing box, a seed discharging device and a seed discharging tube, the seed discharging device is a suction wheel type seed discharging device and comprises a box body, a sealing cover, an inner core and a seed discharging opening, the bottom of the seed discharging device is fixedly provided with a baffle block, the inner core is cylindrical in shape, a driving shaft traverses through the inner core, recessed round pits are uniformly distributed on the external surface of the inner core, the recessed round pits are distributed in a five-in-line and six-in-line interval manner, the seed discharging tube is arc-shaped, and the arc angle is not smaller than 40DEG and not greater than 50DEG. When the apparatus works, a negative pressure is generated in the seed discharging device through a high speed fan, seeds are adsorbed in recessed round pits under the action of below air suction holes and rotate with the inner core, the suction force to the seeds decreases and even disappears when the seeds rotate to the fixed hollow baffle block under the seed discharging device, the seeds depart from the recessed round pits under the action of a seed scraper and enter the seed discharging tube, and the seed discharging tube is arc-shaped and has a certain angle to the ground, so the falling speed of the seeds is reduced, deviation from sowing points due to bouncing after seed falling to the ground is avoided, and the sowing precision is improved.

The invention relates to a novel carrier landing method applied to a carrier-based aircraft. The novel carrier landing method applied to the carrier-based aircraft comprises the steps of arranging a buffering track capable of doing mechanical rotation within an appropriate gliding area, used for the carrier-based aircraft to land, on an aircraft carrier deck; before the carrier-based aircraft begins to land, making the buffering track to begin to rotate, wherein the rotation speed of the buffering track is close to the landing speed of the carrier-based aircraft, and the rotation direction ofthe buffering track is reverse to the landing motion direction of the carrier-based aircraft; making the carrier-based aircraft to glide on the deck during just landing on the aircraft carrier; makingthe carrier-based aircraft to be prepared to enter the area of the buffering track; after the carrier-based aircraft enters the area of the buffering area, making the carrier-based aircraft to beginto do decelerated motion on the buffering track; and meanwhile, making the buffering track and the carrier-based aircraft to do deceleration motions in a synchronous manner so that the speed of the buffering track is always close to the speed of the carrier-based aircraft; making the carrier-based aircraft to run out of the buffering track at an extremely small controllable speed so as to finish the decelerating process; and finally, making the carrier-based aircraft to finish landing on the deck. According to the novel carrier landing method applied to the carrier-based aircraft, the airfieldlanding runway can be reduced to about 300 m, and the accident rate of a small airport can be reduced.

The invention provides a spacecraft recycling linear electromagnetic damping device, relates to the field of spacecraft recycling, and aims at solving disadvantages of parachute landing systems that are conventionally used as spacecraft recycling devices, like large diameter and large weight of the parachute surface, and occupying spacecraft load resources. When preparing recycling, a plurality ofevenly arranged support frames support recycling nets, and net corners of the recycling nets are respectively connected with linear electromagnetic damper movers of the support frames. When the spacecraft falls into the recycling nets, the linear electromagnetic damper movers are pulled downward; meanwhile, permanent magnets on stators cut third conductor plates on the movers to generate magneticfields; and braking forces are generated to buffer the descent process of the spacecraft falling into the recycling nets and control the falling speed of the spacecraft. After the recycling process is completed, the movers and the recycling nets are lifted by drive coils integrated in the linear electromagnetic dampers to prepare for a next spacecraft recycling. The device is used for recycling aspacecraft.

The invention discloses a portable single-person reverse spraying type soft landing safety protection device and a using method thereof. A method of combining a parachute, an inflatable coat and an air injection device is adopted, the parachute is opened at a low altitude, and reverse air injection is conducted when the parachute approaches the ground, so that the landing speed is reduced to a large extent; meanwhile, the wearable coat is inflated by part of gas, so that landing impact force is reduced. The device and method are high in reliability, low in operation difficulty, free of professional training and high in generalization performance and realizability; the landing speed and the controllable flight attitude can be obviously reduced, and the parachuting risk is reduced; the device and method are wide in future application prospect, can be popularized in a large range, is safe and environment-friendly, and can be applied to general aviation equipment, high-rise rescue equipment, safety protection devices and the like.

The invention relates to a carrier rocket return body horizontal position control method, which comprises the following steps of: before a sublevel enters a landing section, interpolating horizontal direction positions Xfscx and Zfscx and horizontal direction speeds Vxfscx and Vzfscx under a drop point coordinate system according to a standard trajectory; carrying out filtering processing on horizontal direction positions Xfs and Zfs and horizontal direction speeds Vxfs and Vzfs which are measured and output by the rocket navigation system; calculating an acceleration instruction controlled in the horizontal direction by using the measurement information Xfslb, Zfslb, Vxfslb and Vzfslb after filtering processing; calculating an attitude angle control instruction of the rocket according to the pitch direction acceleration instruction ay and the yaw direction acceleration instruction az of the sublevel; and by using the obtained pitch attitude angle instruction and the yaw attitude angle instruction psi cx, controlling the flight of the first sub-stage according to an attitude control method until the engine of the first sub-stage is landed and shut down. The method has the following predictable technical effects that the control instruction calculation amount is small, and the guidance loop stability and the engineering application value are high.

The invention discloses a high-performance hybrid hullform for an amphibious aircraft. The high-performance hybrid hullform comprises a multi-bilge-line hull, two ground-effect wings and two force-bearing buoys, wherein the hull is tapering in front and rear; the ground-effect wings are symmetrically arranged at the two sides of the middle part of the hull, and the force-bearing buoys are fixed at the outer end. A non-step-fault design is adopted at the middle part of the hull, the hull is provided with three bilge lines, an upper hull between the upper bilge line and the middle bilge line and a middle hull between the middle bilge line and the lower bilge line are in the shape of a linear ramp with a bilge turn respectively, the bilge turns are arranged at the outer sides of the transverse sections, and the ramp angles of the upper hull and the middle hull range from 65 degrees to 75 degrees; and the transverse section of a lower hull is in the shape of a curve ramp. According to the high-performance hybrid hullform disclosed by the invention, the motion response of the amphibious aircraft in a wave is reduced through using the hull with the large ramp angle and multiple bilge lines; and meanwhile, an aerodynamic lift is remarkably increased and a water take-off and landing speed is reduced through using the ground-effect wings and a ground-effect lift-enhancing air cavity, and then the motion response of the amphibious aircraft in the wave is reduced and the pitching motion time of the amphibious aircraft during take-off and landing in the wave is reduced, and the anti-wave capacity of the amphibious aircraft is remarkably improved.

The utility model relates to a linear electromagnetic damping device for spacecraft recovery, relating to the field of spacecraft recovery. The purpose of the present invention is to solve the problem that the parachute landing system, which is widely used in the existing spacecraft recovery device, has the disadvantages of large diameter and heavy weight of the umbrella surface, and the problem of occupying the load resources of the spacecraft. When it is ready to be recycled, multiple evenly arranged support frames support the recovery net, and the net angles of the recovery net are respectively connected with the linear electromagnetic damper movers of the multiple support frames. When the spacecraft falls into the recovery net, it will pull the mover of the linear electromagnetic damper to move downward. At this time, the permanent magnet on the stator cuts the No. 3 conductor plate on the mover to generate a magnetic field, which will generate a braking force against falling into the recovery net. The descending speed of the inner spacecraft is buffered to control the descending speed of the spacecraft. After the recovery is completed, the drive coil integrated in the linear electromagnetic damping is used to lift the mover and the recovery net to prepare for the next spacecraft recovery. It is used to recover spacecraft.

The invention relates to the technical field of construction equipment, and discloses an inverted arch synchronous backfilling device for tunnel construction. The inverted arch synchronous backfilling device comprises an outer bracket, a fixed pile is arranged at the left end of the outer bracket, a traction piece is movably connected to the left side of the fixed pile, an auxiliary seat is arranged in the middle of the outer bracket, a filling opening is fixedly connected to the upper end of the left side of the auxiliary seat, a filling truck is arranged at the position, under the auxiliary seat, of the outer bracket, a chain groove is formed in the outer side of the filling truck, and a driving shaft is arranged at the upper end of the left side of the filling truck. According to the inverted arch synchronous backfilling device for tunnel construction, filling materials are firstly smashed through a stirring shaft, then are filled into empty grooves in the surface of the chain groove and are filled into an inverted arch along the obliquely-downward filling truck, so that the situation that the filling materials directly impact the inverted arch to cause sediment splashing can be avoided, and material waste is prevented; and meanwhile, the construction risk can be reduced, the landing speed of the filling materials is reduced, material gaps formed during filling can be effectively controlled, and convenience is provided for follow-up flattening.

The invention discloses a safety auxiliary device for airplane forced landing. The safety auxiliary device comprises a sliding track, a freight trolley and a monitor. The monitor is used for being arranged in an aircraft warehouse and connected with a cockpit. The sliding track is used for being arranged in an aircraft warehouse, a rack is arranged on the sliding track, the cargo conveying trolley is arranged above the sliding track and used for moving on the rack, limiting holes are formed in the two sides of the sliding track respectively, two grooves are formed in the limiting holes, and the grooves communicate with the limiting holes. A metal sliding block and a spring are installed in the groove, one end of the spring is fixedly connected with the bottom of the groove, the other end of the spring is fixedly connected with the metal sliding block, the metal sliding block is in sliding fit with the groove, and a first electromagnet is further arranged at the bottom of the groove. The invention aims to solve the problem that in the prior art, the dead weight of an airplane is too large, so that impact force generated by acceleration and inertia during forced landing of the airplane is too large, and safe forced landing cannot be achieved.

The invention provides a wheel shaft type broad width fertilizing apparatus. The apparatus comprises a box body, an inner core, a fertilizer discharging opening and a fertilizer distribution tube, the inner core is cylindrical and is positioned in the box body, a driving shaft traverses through the inner core, two ends of the inner core are connected with the driving shaft through bearings, the external surface of the inner core is provided with uniformly distributed recessed round pits, the recessed round pits are distributed in a five-in-line and six-in-line interval manner, a row of brushes is arranged in the box body, the fertilizer discharging opening is positioned in the bottom of the box body, the fertilizer discharging tube is positioned under a fertilizer discharging device and is communicated with the fertilizer discharging opening, the fertilizer distribution tube is arc-shaped, and the arc angle theta is not smaller than 40DEG and not greater than 50DEG. When the apparatus works, a fertilizer enters the box body, the recessed round pits on the inner core complete 11-line fertilizer distribution, the superfluous fertilizer outside the recessed round pits is brushed off by the brushes in the box body, and a certain amount of the fertilizer is contained in each of the recessed round pits, rotates with the driving shaft and enters the arc-shaped fertilizer distribution tube through the fertilizer discharging opening, so the falling speed of the fertilizer is reduced, bouncing after fertilizer falling to the ground is avoided, and the fertilizing precision is improved.

The invention discloses a helicopter and an ejection escape buffering device, belongs to the technical field of helicopter escape devices, and aims to solve the problems of complex ejection processes,low safety coefficient and poor protection performance during crashing of helicopters in the prior art. The helicopter comprises a fuselage, a cockpit detachably connected with the fuselage, a connection protecting device which is arranged in the connected position of the fuselage and the cockpit to realize connection or separation of the fuselage and the cockpit as well as a separation bufferingdevice arranged in the position, close to one end of the connected position of the fuselage and the cockpit, of the cockpit; and the falling and landing speeds of the cockpit are reduced under the action of rocket catapults and parachutes of the separation buffering device.

The invention relates to a recovery system of a helicopter. The recovery system comprises a parachute bay, a switching device and a control module; the parachute bay is fixedly connected with a propeller spindle of the helicopter through a switching device, and is arranged above a rotor wing, wherein the parachute bay comprises a parachute bay cover, a parachute bag, an ejector and a bottom plate of the parachute bay; the parachute bay cover is in a circular or streamline shape; the parachute bay cover is provided with a semi-circular or streamline-shaped spherical shell, and is buckled on the bottom plate of the parachute bay; the parachute bag is stacked in a parachute bay cavity which is formed by the parachute bay cover and the bottom plate of the parachute bay; the ejector is arranged inside the parachute bay cavity, and is fixedly arranged in the center of the bottom plate of the parachute bay; a separation part of the ejector is fixedly connected onto the parachute bay cover; the control module is loaded on the helicopter, and is used for starting the recovery system of the helicopter. The system is used for reducing the landing speed of the helicopter when the helicopter has an accident, so that the landing time is prolonged, personnel on the ground can evacuate conveniently, and the damage of the helicopter when the helicopter hits the ground is reduced.

The invention provides an electromagnetic damping device for spacecraft recovery, and belongs to the field of the spacecraft recovery. The electromagnetic damping device for the spacecraft recovery isused for solving the problems that a parachute landing system is mainly adopted in an existing spacecraft recovery device, the parachute landing system has the disadvantages of large umbrella fabricdiameter, large weight and the like, and load resources of a spacecraft are occupied. Stators are a plurality of permanent magnet discs, rotors are a plurality of conductor discs, the permanent magnetdiscs and the conductor discs are distributed alternately to form a multilayer disc type structure, electromagnetic damper and the rotors and steel rope brackets of steel rope unrolling motors all penetrate through and are coaxially fixed to rotary shafts, steel ropes wind the steel rope brackets, mounting plates are fixed to the side walls of the top ends of supporting frames, the multiple permanent magnet discs of the electromagnetic dampers and the stators of the steel rope unrolling motors are all fixed to the mounting plates, and net angles of a recovery net are connected with the head ends of the steel ropes. The electromagnetic damping device for the spacecraft recovery is used for recovering the spacecrafts.