372results about How to "Overcome or mitigate technical deficiencies" patented technology

The invention belongs to the field of gas compressor stator blade angle adjustment design, and particularly relates to a stator casing structure. The stator casing structure is characterized by comprising a stator casing, a plurality of stator blades, a stator blade angle adjusting mechanism and a linkage ring; a plurality of mounting holes distributed in the circumferential direction are formed in the stator casing; the plurality of stator blades are distributed in the stator casing in the circumferential direction; the shaft diameter of each stator blade correspondingly extends out of one mounting hole; the stator blade angle adjusting mechanism comprises a plurality of rocker arms, and one end of each rocker arm is correspondingly connected with one shaft diameter; the linkage ring sleeves the periphery of the stator casing and is hinged to the other end of each rocker arm; each rocker arm can rotate relative to the linkage ring and can slide in the circumferential direction of thelinkage ring; and the linkage ring can move in the axial direction of the stator casing so as to drive all the rocker arms to move, and therefore the stator blades can rotate. In addition, the invention relates to the stator blade angle adjusting mechanism, and the stator blade angle adjusting mechanism has all characteristics of the stator blade angle adjusting mechanism in the stator casing structure.

The invention belongs to the technical field of two-dimensional thrust vectoring nozzle outer cover design, and particularly relates to a two-dimensional thrust vectoring nozzle outer cover structure. The two-dimensional thrust vectoring nozzle outer cover structure comprises a rear-section annular cover, a middle-section annular cover and a front-section annular cover, wherein the rear-section annular cover sleeves the periphery of a two-dimensional thrust vectoring nozzle, the tail end is connected with the outlet end of the expanding section of the two-dimensional thrust vectoring nozzle, and a set angle alpha is formed between the tail end and the expanding section. The middle-section annular cover sleeves the periphery of the two-dimensional thrust vectoring nozzle, the tail end of the middle-section annular cover is rotationally connected with the front end of the rear-section annular cover, and the inner side of the middle-section annular cover is rotationally connected with the expanding section. The front-section annular cover sleeves the periphery of the two-dimensional thrust vectoring nozzle, the tail end of the front-section annular cover is partially overlapped with the front end of the middle-section annular cover and is in sliding connection with the middle-section annular cover, and the inner side of the front-section annular cover is rotationally connected with the round-to-square section of the two-dimensional thrust vectoring nozzle. In addition, the invention relates to a design method of the two-dimensional thrust vectoring nozzle outer cover structure.

The invention belongs to the technical field of durability tests of stator blade angle adjusting mechanisms of engines, and particularly relates to a durability test structure of a stator blade angleadjusting mechanism. The durability test structure comprises: a compressor casing; a plurality of elastic pieces which are connected with the compressor casing and distributed in the circumferential direction of the compressor casing; a plurality of rocker arms, wherein one end of each rocker arm is correspondingly connected with one elastic piece; a linkage ring which is arranged on the compressor casing in a sleeving manner and is connected with the other end of each rocker arm; and an actuating cylinder which is arranged on the compressor casing, wherein the piston rod of the actuating cylinder is connected with the linkage ring so as to drive the linkage ring to rotate in the circumferential direction of the compressor casing, and therefore the rocker arms are driven to swing synchronously, and the elastic pieces are deformed.

The invention belongs to the technical field of gas guide design of aircraft engines, and particularly relates to a gas guide structure of a gas compressor. The gas guide structure comprises a statorouter ring, a stator casing and a gas guide pipe, wherein the stator outer ring is provided with a gas guide hole, the inner side of the stator outer ring is provided with a flow guide groove, one endof the flow guide groove extends to the gas guide hole, the other end of the flow guide groove obliquely extends towards an incoming flow direction of an airflow channel in the stator outer ring, thestator outer ring is arranged in the stator casing, an air collection cavity communicating with the gas guide hole is formed between the stator casing and the stator outer ring, an air outlet hole communicating with the air collection cavity is formed in the stator casing, and an inlet of the gas guide pipe communicates with the air outlet hole. In addition, the invention further relates to a gascompressor flow guide groove which has all characteristics of the flow guide groove in the gas compressor gas guide structure.

The invention belongs to the technical field of jet pipe structure design, and particularly relates to an adjustable jet pipe structure. The adjustable jet pipe structure comprises a jet pipe connecting cylinder, an adjusting sheet and actuating cylinders; the outer wall of the jet pipe connecting cylinder is provided with a reverse thrust exhaust port, the reverse thrust exhaust port is inclined towards the inlet end of the jet pipe connecting cylinder, and the adjusting sheet is arranged at the outlet end of the jet pipe connecting cylinder and is in sliding connection with the jet pipe connecting cylinder; the actuating cylinders are hinged between the jet pipe connecting cylinder and the adjusting sheet so as to drive the adjusting sheet to slide relative to the jet pipe connecting cylinder, and then the actuating cylinders have a forward thrust state and a reverse thrust state; in the forward thrust state, the adjusting sheet moves to slide relative to the jet pipe connecting cylinder, the effective area for shielding the outlet end of the jet pipe connecting cylinder is gradually reduced, the effective area for shielding the reverse thrust exhaust port is gradually increased, and finally, the effective area of the reverse thrust exhaust port can be completely shielded; and in the reverse thrust state, the adjusting sheet moves to slide relative to the jet pipe connecting cylinder, the effective area for shielding the reverse thrust exhaust port is gradually reduced, the effective area for shielding the outlet end of the jet pipe connecting cylinder is gradually increased, and finally, the effective area of the outlet end of the jet pipe connecting cylinder can be completely shielded.

The invention belongs to the technical field of bearing cooling design, and particularly relates to a bearing cooling structure. The bearing cooling structure comprises a shaft neck, a bearing and a bearing cooling oil spray nozzle, an annular shaft neck oil collecting groove is formed in the inner wall of the shaft neck, a plurality of shaft neck oil throwing holes distributed in the circumferential direction are formed in the outer wall of the shaft neck, and each shaft neck oil slinging hole communicates with the annular shaft neck oil collecting grooves; the outer ring of the bearing is connected to the shaft neck in a sleeving mode, a plurality of bearing oil inlet holes are formed in the outer ring, each bearing oil inlet hole correspondingly communicates with one shaft neck oil throwing hole; and the bearing cooling oil spray nozzle is arranged in the shaft neck so as to spray lubricating oil into the annular shaft neck oil collecting groove.

The invention belongs to the technical field of engine inlet casing assembly design and specifically relates to an engine inlet casing structure. The engine inlet casing structure comprises an annularengine inlet casing, a plurality of support plates and an inner ring, wherein the engine inlet casing is provided with limiting grooves distributed circumferentially; the plurality of support platesare arranged in the engine inlet casing; one end of each support plate serves as an outer end, and the other end of each support plate serves as an inner end; the outer end of each support plate is inserted onto one limiting groove correspondingly; the inner ends of the support plates are spliced mutually to form a support plate ring; a contact face formed by splicing the inner end of one supportplate and the inner ends of the two adjacent support plates is parallel to a symmetry plane of the engine inlet casing; and the inner ring is arranged in the support plate ring and connected with thesupport plate ring. In addition, the invention further relates to an assembly method of the engine inlet casing structure. The assembly method of the engine inlet casing structure is used for realizing assembly of the engine inlet casing structure.

The application belongs to the technical field of engine rotor testing, and particularly relates to an engine rotor testing lead structure which comprises a rotor journal, an outer lead cylinder arranged on the inner side of the rotor journal along the axial direction and connected with the rotor journal, an the inner lead cylinder arranged on the inner side of the outer lead cylinder in the axialdirection and connected with the outer lead cylinder, and a plurality of test leads. A plurality of lead channels extending in the axial direction are formed between the inner lead cylinder and the outer lead cylinder; and each test lead correspondingly penetrates through one lead channel.

The invention belongs to the technical field of aircraft structure fatigue tests, and particularly relates to a load loading device, which comprises a support beam provided with a groove, strip-shapedholes are formed in two opposite side walls of the groove; an actuator cylinder, a piston rod of the actuator cylinder is used for being connected to an aircraft structure so as to apply a load to the aircraft structure; a sliding rod which is connected with a cylinder body of the actuating cylinder, one end of the sliding rod extends into one strip-shaped hole, the other end of the sliding rod extends into the other strip-shaped hole, and the sliding rod can slide in the extending direction of the strip-shaped holes, so that the actuating cylinder can be driven to move or driven by the actuating cylinder to move, and the direction of load applied to the aircraft structure by the actuating cylinder can be adjusted.

The invention belongs to the technical field of high and low temperature impact test design, and particularly relates to a high and low temperature impact test device. The device comprises: a supporting cylinder; a heating source which is arranged on the outer side of the supporting cylinder; two end covers, wherein each end cover is correspondingly connected to one end of the supporting cylinder, a plurality of impact holes are formed in one end cover, and each impact hole is communicated with the interior of the supporting cylinder; a cover plate, wherein a liquid nitrogen inflow hole and a high-pressure air inflow hole are formed in the cover plate, the cover plate is connected to the outer wall of the end cover with the multiple impact holes, a mixing cavity is formed between the cover plate and the end cover, and the mixing cavity is communicated with the impact holes, the liquid nitrogen inflow hole and the high-pressure air inflow hole.

The invention belongs to the technical field of gas compressor stator blade rotating angle measuring, and particularly relates to a stator blade rotating angle adjusting and measuring structure. The stator blade rotating angle adjusting and measuring structure comprises multiple rocker arms, a linkage ring, a pointer and a graduated disc, wherein one end of each rocker arm is correspondingly connected with a part, extending of a gas compressor case, of one stator blade, the linkage ring is connected with the other ends of the rocker arms, is mounted on the gas compressor case in a sleeving manner, can rotate in the circumferential direction of the gas compressor case and drives the rocker arms to swing synchronously, therefore, all stator blades can synchronously rotate, accordingly, angles of the stator blades can be synchronously adjusted, the pointer is connected with one rocker arm, and can swing along with the rocker arm, the graduated disc is mounted on the gas compressor case soas to mark the swing amplitude of the pointer, and the rotating angles of the stator blades can be obtained. In addition, the invention further relates to a calibration method of the stator blade rotating angle adjusting and measuring structure.

The invention belongs to the technical field of radar test equipment design, and particularly relates to radar test equipment, which comprises a general storage module used for storing radar record data including target echo information, interference information, clutter signals, SAR information and other radar radiation information, a baseband signal processing module performing digital up-conversion and playback based on the radar record data to generate a digital intermediate frequency playback signal, a DAC module used for carrying out digital-to-analog conversion on the digital intermediate frequency playback signal to generate an analog intermediate frequency playback signal, a broadband up-conversion module used for carrying out up-conversion on the analog intermediate frequency playback signal to generate a radio frequency playback signal, and two broadband transmitting channels used for amplifying and outputting the radio frequency playback signal, so that a target echo signal, an interference signal, a clutter signal, an SAR signal and other radar radiation signals based on radar record data can be simulated.

The invention belongs to the technical field of structural designs of rotor and stator structures of compressors, and particularly relates to a casing structure; the casing structure comprises an outer casing and an inner casing, wherein the inner casing is arranged at the inner side of the outer casing and located at the rear section of the outer casing; and an annular air gap is formed between the inner casing and the outer casing. In addition, the invention relates to a rotor and stator structure comprising the casing structure.

The invention belongs to the technical field of variable-camber wing leading edge functional strength tests, and particularly relates to a variable-camber wing leading edge functional strength test load loading device which comprises: a supporting frame used for being connected to a variable-camber wing leading edge; a deformation driving motor which is connected to the supporting frame and used for being connected to a driving connecting rod mechanism at the root of the front edge of the variable-camber wing so as to apply a deformation load to the front edge of the variable-camber wing; an upper airfoil aerodynamic load loading actuator cylinder, wherein the cylinder body of theupper airfoil aerodynamic load loading actuator cylinder is connected to the supporting frame in a sliding mode, and a piston rod of the upper airfoil aerodynamic load loading actuator cylinder is used for being connected to the upper airfoil of the front edge of the variable-camber wing so that aerodynamic loads can be applied in the normal direction of the upper airfoil of the front edge of the variable-camber wing; a lower airfoil aerodynamic load loading actuator cylinder, wherein a cylinder body of the lower airfoil aerodynamic load loading actuator cylinder is connected to the lower portion of the supporting frame in a sliding mode, and a piston rod of the lower airfoil aerodynamic load loading actuator cylinder is used for being connected to the lower airfoil of the front edge of the variable-camber wing so that aerodynamic loads can be applied in the normal direction of the lower airfoil of the front edge of the variable-camber wing.

The invention relates to an intermediate case structure which comprises an intermediate case, a front shunting ring, an outer ring, an inner ring, a plurality of support plates, a rear shunting ring,a plurality of duct converting valves, a plurality of rotary shafts and an actuating mechanism; the front shunting ring is arranged in the intermediate case; the outer ring is in butt joint with the front shunting ring; the inner ring is arranged in the outer ring; one end of each of the support plates is connected with the intermediate case, the other end thereof is connected with the inner ring,and a gap is formed in the side, back on to the inlet end of the intermediate case; the rear shunting ring is clamped in all gaps and is positioned between the outer ring and the inner ring; each ofthe duct converting valves is hinged to the front shunting ring, and conical gears are arranged on the duct converting valves; one conical gear is arranged at one end of each of the rotary shafts; theconical gear on each of the rotary shafts is meshed with the conical gear on the corresponding duct converting valve; the actuating mechanism is arranged on the intermediate case and is connected with the other end of each of the rotary shafts to drive all rotary shafts to rotate, so that the duct converting valves rotate, and the intermediate case structure has the outer duct closing state and the outer duct opening state; when the intermediate case structure is in the outer duct closing state, the duct converting valves are in lap joint with the rear shunting ring; and when the intermediatecase structure is in the outer duct opening state, the duct converting valves are separated from the rear shunting ring.

The invention relates to a flap movement mechanism fatigue test device. The device comprises a support; a flap movement mechanism connected to the support; a flap connected to the flap movement mechanism so as to be driven by the flap movement mechanism to move; loading actuating cylinders, wherein piston rods are connected to the flap; and a follow-up supporting plate connected with the base of each loading actuating cylinder and capable of moving along with the flap, so that each loading actuating cylinder is kept vertical to the flap. In addition, the invention relates to a flap movement mechanism fatigue test method which is implemented based on the flap movement mechanism fatigue test device.

The invention relates to an intermediate case assembly. The intermediate case assembly comprises an intermediate case, a front shunt ring, an outer ring, an inner ring, a plurality of support plates,a plurality of duct conversion valves, a plurality of adjusting rods and an actuating mechanism, wherein the front shunt ring is arranged in the intermediate case; the outer ring is connected with thefront shunt ring in a butt joint mode; the inner ring is arranged in the outer ring; one end of each support plate is connected with the intermediate case, the other end of each support plate is connected with the inner ring, and a notch is formed in one side, face back onto the inlet end of the intermediate case, of each support plate; a rear shunt ring is clamped into the notches; the pluralityof duct conversion valves are hinged to the front shunt ring; the two sides of each duct conversion valve are in sealing contact with the corresponding two adjacent support plates; one end of each adjusting rod is hinged to the corresponding duct conversion valve; the actuating mechanism is arranged on the intermediate case and hinged with the other end of each adjusting rod so as to drive each adjusting rod to move and make each duct conversion valve to rotate, so that the intermediate case assembly is provided with an outer duct closing state and an outer duct opening state; in the outer duct closing state, each duct conversion valve is connected to the rear shunt ring in an overlap joint mode; and in the outer duct opening state, each duct conversion valve is separated from the rear shunt ring.

The invention belongs to the technical field of aero-engine jet pipe structure design, and particularly relates to an expansion piece pull rod structure of an aero-engine jet pipe. The expansion piece pull rod structure comprises: a first connecting rod, wherein a connecting hole is formed in one end of the first connecting rod; a second connecting rod, wherein one end of the second connecting rod is inserted into the connecting hole and can axially move relative to the first connecting rod, and an annular clamping groove is formed between the second connecting rod and the connecting hole; and an elastic ring, wherein the elastic ring is clamped in the annular clamping groove and can generate friction force between the second connecting rod and the first connecting rod so as to prevent the second connecting rod from moving relative to the first connecting rod in the axial direction. In addition, the invention relates to the aero-engine jet pipe structure comprising the expansion piece pull rod structure of the aero-engine jet pipe.

The invention belongs to the technical field of disassembly and assembly design of an engine rotor and stator structure, and particularly relates to an engine rotor and stator structure assembling anddisassembling device. The engine rotor and stator structure assembling and disassembling device comprises a support seat, a positioning shaft, a positioning cylinder, a support and a rotating device,wherein one end of the positioning shaft is connected with the support seat, one end of the positioning cylinder is connected with the support seat, the positioning cylinder and the positioning shaftare coaxially arranged, and the rotating device is arranged on the support and is connected with the support seat or the positioning cylinder so as to drive the positioning cylinder and the positioning shaft to turn over.

The invention belongs to the technical field of engine air intake frame design, and particularly relates to an engine air intake frame, which comprises an inlet casing, an inner ring and a plurality of support plates, wherein the inlet casing is provided with a plurality of through holes distributed in a circumferential direction, and a plurality of flanges are arranged on the outer wall of the inlet casing; each flange is correspondingly close to one through hole; the inner ring is arranged in the inlet casing; the plurality of support plates are arranged between the inlet casing and the inner ring; one end of each support plate is connected to the inner ring; and the other end of each support plate correspondingly extends out from one through hole and is connected to the corresponding flange through a bolt.

The invention belongs to the technical field of engine rotor wheel disk assembly design, and particularly relates to a rotor wheel disk assembly structure. The assembly structure comprises a primary rotor wheel disk, a plurality of connecting bolts, a plurality of clamping rings, a secondary rotor wheel disk and multiple connecting nuts, wherein the primary rotor wheel disk is provided with a primary drum, the primary drum is provided with a primary connecting edge, and the primary connecting edge is provided with a plurality of primary connecting holes formed in the circumferential direction;each connecting bolt correspondingly penetrate through one primary connecting hole, the head portions of the connecting bolts are located on the inner side of the primary drum, and annular clamping grooves are formed in the connecting bolts; each clamping ring is correspondingly clamped into one annular clamping groove; the secondary rotor wheel disk is provided with a plurality of secondary connecting holes in the circumferential direction; the portion, extending out of the corresponding primary connecting hole, of each connecting bolt correspondingly penetrates through one secondary connecting hole; each clamping ring is located between the primary connecting edge and the secondary rotor wheel disk; and each nut is correspondingly in threaded connection with one connecting bolt.

The invention belongs to the technical field of aero-engine nozzle adjusting mechanism design, and particularly relates to an aero-engine nozzle adjusting mechanism. The aero-engine nozzle adjusting mechanism comprises multiple rocker arms, pull rods, multiple actuating cylinders, a supporting ring and multiple connecting rod structures; the rocker arms are hinged to the outer wall of a nozzle connecting cylinder and distributed in the circumferential direction; one end of each pull rod is correspondingly hinged to one rocker arm, and the other end is correspondingly hinged to the outer wall of one convergence piece; a cylinder body of each actuating cylinder is hinged to the outer wall of the nozzle connecting cylinder; each piston rod is correspondingly hinged to one rocker arm to drive the corresponding rocker arm to rotate around the hinged part of the corresponding rocker arm and the nozzle connecting cylinder, so that the corresponding convergence piece is driven to move through the corresponding pull rod; the periphery of the nozzle connecting cylinder is sleeved with the supporting ring; one end of each connecting rod structure is hinged to the supporting ring, the other end is correspondingly hinged to the outer wall of one expansion piece, and the axial effective length can be adaptively adjusted in the process that the corresponding expansion piece moves along with the corresponding convergence piece.

The invention belongs to the technical field of assembly design of a stator of a gas compressor, and particularly relates to a stator structure. The stator structure comprises a stator casing, multiple stator blades and multiple positioning blocks, wherein multiple limit holes distributed in the circumferential direction are formed in the stator casing; the multiple stator blades are arranged in the stator casing; an upper journal is arranged on the upper end surface of each stator blade, and a lower journal is arranged on the lower end surface of each stator blade; each upper journal correspondingly extends into a limit hole; each positioning block is provided with a positioning surface and an arc surface opposite to the positioning surface; a positioning hole is formed in each positioning surface; each lower journal is correspondingly inserted into one positioning hole, and the corresponding lower end surface is attached to the corresponding positioning surface; and the arc surfacesare combined to form an annular surface. Besides, the invention further relates to an assembly method of the stator structure.

An intermediate case test load loading device comprises: an adapter cylinder, wherein one end of the adapter cylinder is in butt joint with an intermediate case, and the outer wall of the end is connected with a section torque loading connector and a section radial loading connector so as to apply section torque and radial load to the intermediate case; a switching frame connected with the end face of the other end of the switching cylinder, wherein two section bending moment and axial loading joints are connected to the switching frame so as to apply section bending moment and axial load to the intermediate case; a bearing frame connected to the switching frame, wherein a fulcrum axial load actuator is positioned and mounted on the bearing frame; a switching shaft, one end of the switching shaft is connected with a fulcrum part in the intermediate case, the other end of the switching shaft extends out of the switching cylinder, and the outer wall of the end is connected with a fulcrum radial loading joint so as to apply a fulcrum radial load to the intermediate case; and a switching rack arranged on the periphery of the switching shaft in a sleeving manner, connected with the fulcrum part in the intermediate case, ans connected with a fulcrum axial loading joint. The fulcrum axial loading joint is connected with the fulcrum axial load actuator so as to apply a fulcrum axial load to the intermediate case.

The invention belongs to the technical field of high-load fan and compressor blade profile design, and particularly relates to a blade profile mean camber line design method. The method comprises thefollowing steps: constructing a blade profile mean camber line corner distribution curve by using three sections of three-order Bezier curves; constructing a limiting function of the blade profile mean camber line corner distribution to limit the upper limit and the lower limit of the mean camber line corner distribution freedom degree; and the incidence relation between the blade profile mean camber line corner distribution control points and the blade profile load is constructed, so that the blade profile load is associated with the mean camber line control parameters, and the degree of freedom of blade profile mean camber line corner distribution is further limited. In addition, the invention relates to a blade profile design method which comprises the following steps: obtaining a bladeprofile mean camber line based on the blade profile mean camber line design method; and the obtained blade profile mean camber lines are subjected to superposition thickness distribution to obtain the blade profile.

The invention belongs to the technical field of flexible vibration damping connection design of a structure, and particularly relates to a flexible vibration damping connection structure. The flexiblevibration damping connection structure comprises a first supporting plate, a second supporting plate, multiple first supporting strips, multiple second supporting strips, multiple groups of first elastic pieces and multiple groups of second elastic pieces, wherein the second supporting plate is opposite to the first supporting plate; the multiple first supporting strips are arranged between the first supporting plate and the second supporting plate, and one end of each first supporting strip is connected with the first supporting plate; the multiple second supporting strips are arranged between the first supporting plate and the second supporting plate, one end of each second supporting strip is connected with the second supporting plate, and the second supporting strips and the first supporting strips are in cross distribution; each first elastic piece in each group of first elastic pieces is correspondingly arranged on the corresponding first supporting strip in the axial direction;each second elastic piece in each group of second elastic pieces is correspondingly arranged on the corresponding second supporting strip in the axial direction; and the elastic pieces on each firstsupporting strip are in cross distribution with the elastic pieces on the second supporting strip adjacent to the first supporting strip.

The invention belongs to the technical field of compressor assembly design, and particularly relates to a stator structure. The stator structure comprises a stator casing, two semi-annular edge rings,two semi-annular stator inner rings and two sets of stator blades. The inner wall of the stator casing is provided with an annular limit groove. The two semi-annular edge rings are arranged inside the stator casing and are butt each other to form an edge ring. One end of the edge ring is clamped into the annular limit groove. The two semi-annular stator inner rings are butt each other to form a stator inner ring. Each semi-annular stator inner ring is correspondingly arranged inside a semi-annular edge ring. Each set of the stator blades are arranged in a circumferential direction between thesemi-annular edge ring and the semi-annular stator inner ring corresponding to the semi-annular edge ring. The tip of each stator blade is formed on the inner wall of the corresponding semi-annular edge ring. A blade root is formed on the outer wall of the corresponding semi-annular stator inner ring. In addition, the invention also relates to a rotor stator assembly structure with the stator structure and an assembly method of a rotor stator structure.

The invention belongs to the technical field of aircraft driving, and particularly relates to an aircraft auxiliary driving method and system, and the aircraft auxiliary driving method comprises the steps: judging the flight stage of an aircraft according to flight parameters; judging the working state of the airborne system according to the airborne system information; according to the flight stage and the working state of the airborne system, obtaining the flight state of the aircraft; whether the aircraft control instruction meets the aircraft flight state safety requirement, the operation standard of the flight stage corresponding to the aircraft flight state and the aircraft flight state abnormity processing specification or not is judged, a corresponding auxiliary driving control instruction, a control suggestion and an alarm signal are generated according to the judgment result, and the aircraft is controlled through the auxiliary driving control instruction.