100 results about "Rolling angle" patented technology

The invention discloses a roll angle measuring method for a continuous variable roll test of a transient wind tunnel, and belongs to the technical field of wind tunnel test. Comprising the following steps: step S1, early-stage preparation: selecting two angle measurement sensors to carry out measurement work, and installing the two angle measurement sensors and an installation platform according to a preset angle; s2, carrying out field calibration operation; carrying out field calibration on the coefficient of the angle measurement sensor by adopting a fixed attack angle and variable roll angle mode; and S3, data interception, discretization and application during a step variable attack angle continuous variable roll angle test: in a wind tunnel test process, continuously collecting output voltage values of the two angle measurement sensors, calculating roll angles corresponding to the output voltage values, intercepting roll angle range values corresponding to the two angle measurement sensors according to needs, and calculating the roll angle range values of the two angle measurement sensors according to the intercepted roll angle range values. Splicing into a target measurement range; compared with a traditional step variable rolling and fixed rolling variable attack angle test mode, the scheme has the advantage that the test efficiency is greatly improved.

The invention belongs to the field of aviation structure strength tests, and particularly relates to a testing machine pitching and rolling angle adjusting method and system for a full-aircraft drop test, and the method comprises the steps: adjusting the lengths of a front lifting rope l1 and a rear lifting rope l2 through a first servo motor; the lengths of the adjusted front lifting rope l1 and the adjusted rear lifting rope l2 are calculated through the cosine law according to the geometrical relationship; the lengths of the left lifting rope 13 and the right lifting rope 14 are adjusted through a second servo motor; and the lengths of the adjusted left lifting rope 13 and the right lifting rope 14 can be calculated through the cosine law according to the geometrical relationship. The pitching and rolling postures can be automatically adjusted, and the problems that an existing method is low in efficiency and poor in precision are solved. According to test experience, according to an existing posture adjusting method, replacement needs about half a day to one day every time, the number of times of putting every day does not exceed ten at most, and manual operation is needed on wings or a fuselage every time. And after automatic adjustment is achieved, the target can be achieved in several minutes through adjustment of the motor, efficiency is high, and safety is higher.

The invention relates to the technical field of aircrafts, in particular to a sensor parameter adjustment method and device, equipment and a medium, and the method comprises the steps: obtaining the acceleration data of an unmanned plane, the acceleration data of a holder, and the sensing data; and acquiring a detected pitch angle and a detected roll angle based on the sensor parameters and the sensing data. The pitching angle and the rolling angle of the holder are obtained, and the pitching angle and the rolling angle of the unmanned aerial vehicle are obtained. And obtaining an estimated pitch angle according to the pitch angle of the holder and the pitch angle of the unmanned aerial vehicle, and obtaining an estimated roll angle according to the roll angle of the holder and the roll angle of the unmanned aerial vehicle. If the angle difference between the estimated pitch angle and the detected pitch angle is larger than or equal to an angle difference threshold value, or the angle difference between the estimated roll angle and the detected roll angle is larger than or equal to the angle difference threshold value, sensor parameters are adjusted. According to the scheme, the parameters of the sensor can be adjusted in a wide range of scenes, normal use of the holder is ensured, and the user experience is improved.

The application discloses a launching projectile support for irregular projectile body self-defined rolling angle and an assembling method thereof, which comprises a support body, the support body comprising an upper shell, a lower shell, positioning columns and a lever rod, the upper shell and the lower shell being coaxially and threadedly connected at the ends, a cavity with a closed end and an open end being formed between the upper shell and the lower shell, a projectile body being arranged in the middle of the cavity, the positioning columns filling the remaining space of the cavity and supporting the projectile body, the lever rod comprising a lifting rod, a flat plate lever and a supporting column, the flat plate lever being horizontally arranged on the end face of the lower shell, the lifting rod being arranged in the upper shell along the axial direction, one end of the lifting rod being in contact with one end of the flat plate lever, and the supporting column being arranged in the lower shell along the axial direction, one end of the supporting column being in contact with the other end of the flat plate lever. The application can well position various irregular projectile bodies, thereby meeting the demand for various specifications of projectile bodies in experiments, and the application is convenient for experiments, saves manpower and material resources, and guarantees the accuracy of experimental results and the effectiveness of data.

The application designs a high-speed aircraft cluster cooperative guidance method based on space-time synchronous attraction domain. An artificial potential field is designed to keep a proper distance between aircrafts, avoid collision while gathering; a speed matching item is designed based on consistency theory to realize dynamic consistency of aircraft speed; a guidance law based on resistance acceleration-speed profile is designed to meet the cooperative requirements of terminal range, height, speed and other aircrafts, and the combination of the three realizes efficient autonomous clustering and guidance of the aircraft. The application can make the aircraft cluster reach the target point with high precision under the conditions of meeting complex constraints such as heat flow, overload, dynamic pressure, collision avoidance, clustering, dynamic consistency of speed and the like through the design of attraction domain, angle of attack and roll angle control.

The invention provides a high-precision roll angle measurement method and system, and relates to the technical field of optical measurement, and the method comprises the following steps: respectively measuring corresponding first roll angle data, second roll angle data and environment data; performing preprocessing and error estimation on the first roll angle data and the second roll angle data in combination with the environmental parameters, and generating a corresponding to-be-processed sequence according to a data processing result; constructing a roll angle measurement model, and inputting the to-be-processed sequence into the roll angle measurement model for processing to obtain a corresponding roll angle measurement result; according to the invention, through error compensation and data fusion, part of high-precision and high-cost hardware sensor requirements are replaced, and the measurement precision can be effectively improved.

The application discloses a roll aircraft control method based on a two-loop roll attitude pilot, in which a roll angle error and a roll angle rate are obtained in real time, a roll rudder deflection control instruction corresponding to a unit roll angle error and a roll attitude pilot damping coefficient are further weighted, so that a rudder deflection control instruction is obtained and is transmitted to a rudder, the rudder works based on the rudder deflection control instruction, and the roll aircraft is controlled to fly to a target.

The invention provides an aircraft roll angle reconstruction method and system, and belongs to the technical field of aircraft control, the method comprises the steps that aircraft state signals are acquired, and the state signals comprise a ground speed signal Vg, a vacuum speed signal Vt, a yaw angle rate signal R, a pitch angle rate signal Q and a pitch angle signal Pitch; and under the condition that the roll angle signal fails, according to a lift force and gravity balance principle in a rolling state of the airplane, an approximate roll angle reconstruction signal is constructed based on a ground speed signal or a vacuum speed signal, a yaw angle rate signal, a pitch angle rate signal and a pitch angle signal.

This invention provides a high-dynamic space-based cooperative guidance method and system for multi-level virtual target points, relating to the field of aircraft control technology. It involves acquiring initial flight state data of each aircraft and position data of the target point; generating at least one virtual target point based on the initial flight state data and position data, using a zero heading angle constraint; generating roll angle control command data for each aircraft based on the initial flight state data and the virtual target point, using a two-sided azimuth error corridor; and adjusting the attitude of each aircraft according to the roll angle control command data to ensure that each aircraft meets the terminal heading angle constraint and reaches the target point. This method enables precise control of the terminal heading angle of high-dynamic aircraft in complex dynamic environments, thereby enabling multiple aircraft to perform cooperative strike missions from different spatial orientations.

The application discloses a kind of double hulls lift-float integrated aircraft roll adjusting method, it is related to aircraft design technical field, using the form of auxiliary air bag control, main air bag control and full-motion horizontal tail control combination to carry out double hulls lift-float integrated aircraft roll adjusting, using the mode that auxiliary air bag is connected with atmosphere environment and air is pumped out, main air bag is connected and helium is pumped out, full-motion horizontal tail deflection, changes the lift distribution of aircraft, buoyancy distribution, so that aircraft generates roll torque, adjusts aircraft roll angle.The double hulls lift-float integrated aircraft roll adjusting method that the application adopts has significantly improved the aircraft roll adjusting efficiency, is conducive to improving the maneuverability and safety of aircraft, compared with the existing aircraft attitude adjusting mode, combines the characteristics of double hulls lift-float integrated aircraft, aircraft roll adjusting mechanism work can generate greater moment change, aircraft has very high adjusting capacity to attitude.

The application discloses a kind of high-efficiency omni-directional decoupling annular vector tilt-rotor aircraft and control method thereof, aircraft includes fuselage, fuselage driving gear set, four tilting mechanisms, four annular vector tilt-rotor power groups and electronic equipment etc.Parts.The aircraft described in the application can realize six degrees of freedom omni-directional pose decoupling flight, the attitude angle is omni-directional arbitrary angle in it, including roll angle, pitch angle and yaw angle, can realize arbitrary attitude hovering and flight of aircraft.The aircraft described in the application solves the controllable attitude angle limit of conventional variable tilt-rotor mechanism from structure, improves the ability of complex space activity of aircraft, and fundamentally solves the problem of power offset and power redundancy, reduces energy loss, improves power efficiency, prolongs endurance time, and is suitable for application scenarios requiring strong anti-interference and air physical interaction capability, and has important significance and value in the future.

Systems and methods are provided for detecting and warning of aircraft rollover conditions. The systems include one or more sensors onboard the aircraft and a controller onboard the aircraft that is in operable communication with the one or more sensors. The controller is configured to, by one or more processors: receive input data from the one or more sensors and, optionally, one or more other sources, process the input data to dynamically determine a roll angle of the aircraft, an angular roll rate of the aircraft, and a critical rollover angle of the aircraft, determine whether current or potential static or dynamic rollover conditions of the aircraft exist based, at least in part, on the roll angle, the angular roll rate, and the critical rollover angle, and generate an alert in response to determining that the current or potential static or dynamic rollover conditions exist.

The application relates to the technical field of aero-engines, in particular to a lubricating oil tank liquid level detection method, device, equipment, medium and program product. The method comprises the following steps: acquiring a measured liquid level of oil liquid in a lubricating oil tank of an aircraft, and acquiring current attitude information of the aircraft; the attitude information comprises a pitch angle, a yaw angle and a roll angle; based on a standard corresponding relationship, a target standard liquid level corresponding to the current attitude information is determined; the standard corresponding relationship is a corresponding relationship between the attitude information of the aircraft and the liquid level when the oil liquid in the lubricating oil tank is in a normal state; and based on the size relationship between the target standard liquid level and the measured liquid level, a detection result is determined. The method can improve the accuracy of the liquid level measurement of the oil liquid in the lubricating oil tank.

The application discloses a roll angle measurement method for continuous variable roll test of a temporary flush wind tunnel, and belongs to the technical field of wind tunnel testing. The method comprises the following steps: step S1, preliminary preparation, two angle measurement sensors are selected for measurement, and the two angle measurement sensors are installed at a predetermined angle with a mounting platform; step S2, on-site calibration operation; on-site calibration of the angle measurement sensor coefficient is performed in a mode of constant attack angle and variable roll angle; step S3, data interception, discretization and application during step-by-step variable attack angle and continuous variable roll angle test, output voltage values of the two angle measurement sensors are continuously collected during the wind tunnel test, and corresponding roll angles are calculated, and the corresponding roll angle range values of the two angle measurement sensors are intercepted and spliced into a target measurement range as required; compared with a traditional step variable roll and constant roll variable attack angle test mode, the scheme greatly improves test efficiency.

The invention belongs to the field of aircraft drop simulation and test, and particularly relates to an aircraft sinking speed determination method and system under the condition of no lift force and no mass reduction, and the method comprises the steps: determining parameters needed by the dynamics simulation of an aircraft landing gear according to a test working condition; determining the mechanical relationship of the undercarriage according to the parameters required by simulation, and building a full-aircraft landing simulation model; performing lift simulation on the full-aircraft landing simulation model, and determining a vertical ground load of the aircraft landing gear; extracting a vertical ground load of an aircraft landing gear, performing lift-free drop simulation on the full-aircraft landing simulation model, and iteratively determining a putting height according to the ground load; and calculating the sinking speed of the full-aircraft drop test according to the putting height. Compared with the prior art, the method has the advantages that lift force simulating loading equipment is omitted, initial launching attitudes such as the pitch angle, the roll angle and the yaw angle of the aircraft are fully considered, the method can be used for guiding initial launching of a drop test, the large overload risk of free falling impact is reduced, and the number of times of adjusting the launching height can be reduced.

The application is a kind of lateral asymmetry online trimming control method of hypersonic aircraft, which comprises the following steps: firstly, constructing a lateral asymmetry hypersonic aircraft model; then, setting a sideslip angle and a roll angle judgment threshold of the hypersonic aircraft in the lateral asymmetry hypersonic aircraft model, detecting the lateral asymmetry of the hypersonic aircraft, judging whether the lateral asymmetry exists, if the lateral asymmetry exists, judging whether the hypersonic aircraft is controllable according to the detected abnormal information, if the hypersonic aircraft is uncontrollable, dynamically controlling the rudder of the hypersonic aircraft through control law reconstruction, and giving a systematic solution of the lateral asymmetry of the hypersonic aircraft from the aspects of asymmetry model construction, detection method, controllability analysis and automatic trimming control law design, so that the roll and sideslip caused by the asymmetry can be corrected without increasing additional trimming surfaces, and the flight safety of the hypersonic aircraft is ensured.

The application belongs to the field of aircraft flight control system, and particularly relates to an anti-collision control method based on artificial potential field, collision conditions of two aircrafts are set according to relative motion conditions and current flight data of the two aircrafts; the collision conditions of the two aircrafts are obtained, a potential field function is set, and a distance between the two aircrafts is designed as a virtual repulsive potential field; whether there is a collision risk is judged through the potential field function, if yes, a negative gradient in the potential field function is calculated, is converted into repulsive force between the two aircrafts, the repulsive force between the two aircrafts is solved into an aircraft roll angle instruction and a speed increment instruction, an anti-collision instruction is formed to control the two aircrafts; a relative position vector and a closest time of the two aircrafts are obtained, whether the anti-collision condition is met is judged, and the anti-collision is exited after the condition is met; by adopting the potential field function and performing instruction tracking, it is ensured that the two aircrafts can keep a certain relative distance and do not collide.

The application discloses a kind of for unmanned aerial vehicle roll angle's stabilizing control method, belong to the field of quadrotor aircraft attitude angle stabilizing control, comprising: constructing three-order linearized dynamics model;Three-order linearized dynamics model is handled using pole assignment technique, to obtain stable state matrix;Jordan canonical form is constructed based on the eigenvalue component of stable state matrix, to obtain positive eigenvalue matrix and real matrix, and deduce high dynamic output matrix;High dynamic roll angle stabilizing control signal is designed based on positive eigenvalue matrix and high dynamic output matrix;High dynamic roll angle stabilizing control signal is used to solve the command roll torque;The command roll torque is input coupling roll torque output characteristic model, to obtain unmanned aerial vehicle roll channel actual roll torque value.The application improves the high dynamic convergence ability of roll angle, so that unmanned aerial vehicle can be faster, more stable to realize the stabilizing control of roll angle under complex environment.

The invention discloses a roll attitude control method for a large aircraft. The roll attitude control method comprises the following steps: S1, instruction generation: converting the manipulated variable of a pilot into an expected roll angle rate instruction; s2, feedforward control: directly calculating a feedforward control surface instruction according to the expected roll angle rate instruction for accelerating the initial response of the aircraft and ensuring the roll performance of the aircraft; s3, angular rate correction: correcting the roll angle rate measured in real time according to the current pitch angle, roll angle, yaw angle rate and pitch angle rate of the aircraft to obtain a calibrated roll angle rate; s4, proportional control: generating a proportional control quantity according to the deviation between the calibration roll angle rate and the expected roll angle rate instruction, and improving the roll control bandwidth of the large aircraft; and S5, integral control is carried out, an integral control quantity is generated according to the deviation between the calibration roll angle rate and the expected roll angle rate instruction, and the roll control precision of the large aircraft is improved.

This invention relates to a 4D millimeter-wave radar roll angle calibration system. A roll angle calibration value is configured by placing left and right corner reflectors on the upper and lower sides of an assumed horizontal plane of the millimeter-wave radar, respectively. Then, a host computer triggers a calibration program. Upon receiving a command, a slave computer controls the millimeter-wave radar to send test signals to the left and right corner reflectors, thereby detecting their position information. Based on the position information of the left and right corner reflectors, the roll angle measurement value of the millimeter-wave radar can be further calculated. Finally, the roll angle calibration value and the roll angle measurement value are used to calibrate the roll angle of the millimeter-wave radar. This application demonstrates that the host computer controls the radar to automatically complete the roll angle calibration function and derives a method for calculating the radar roll angle based on the angular position information measured by the radar.

A method for performing gyroscopic azimuth angle measurements includes estimating a pitch angle and a roll angle of a gyroscopic inclinometer in a wellbore; determining a measurement duration for each of the plurality of gyroscope measurements from the estimated pitch and roll angles; making each of the plurality of gyroscopic measurements at the determined measurement duration when the gyroscopes are disposed at a corresponding plurality of rotational positions in the tool housing; and calculating an azimuth angle of the wellbore from the plurality of gyroscopic measurements.

This invention provides a method for automatically leveling and attitude measurement of a three-degree-of-freedom support system in a low-speed wind tunnel. The method includes: installing a dual-axis accelerometer on the support system's strut, such that its first sensitive axis is parallel to the pitch axis of the aircraft model and its second sensitive axis is parallel to the roll axis of the model; controlling the roll motor and angle-of-attack motor of the support system so that the output values ​​of both the first and second sensitive axes are zero, thus achieving automatic leveling of the model; during wind tunnel testing, comparing the roll angle measured by a first encoder (detecting the roll angle) and the angle of attack measured by a second encoder (detecting the angle of attack), using a first threshold and a second threshold, selecting the roll angle detected by the first sensitive axis or the measurement value of the first encoder as the actual roll angle of the model, and selecting the angle of attack detected by the second sensitive axis or the measurement value of the second encoder as the actual angle of attack of the model, thereby improving the measurement accuracy of the aircraft model's attitude angle during wind tunnel dynamic testing.

The application discloses a method for calibrating a roll angle of an emergency anti-solar time of a satellite on the earth, which is beneficial to reestablishing a three-axis attitude of the satellite on the earth. Firstly, an angle between a projection of a sun vector on an orbit plane and an orbit system-X axis is calculated; then, whether a current orbit position meets a requirement is determined according to the angle between the projection of the sun vector on the orbit plane and the orbit system-X axis, and whether an absolute attitude reference is available at present is determined, if the current orbit position meets the requirement and the absolute attitude reference is available at present, the roll angle is calibrated at a current control beat; otherwise, the roll angle is not calibrated at the current control beat; if the roll angle is calibrated at the current control beat, the roll angle is calculated based on the absolute attitude reference; if the roll angle is not calibrated at the current control beat, the roll angle is calculated by using gyro inertia angular velocity integration. The application can provide high-precision and reliable attitude information for reestablishing the attitude of the satellite on the earth.

The invention provides a method for measuring and calculating a roll angle of an aircraft by using a forward dual-beam Doppler navigation radar. The method is mainly characterized in that two coordinate systems are respectively used for analytical calculation. Firstly, a drift angle and an axial speed are solved in an airframe coordinate system. Then, in horizontal plane coordinate systems, a Doppler frequency shift equation when the aircraft is only in a rolling state is considered, and at the moment, the plane included angle between the horizontal planes of the two coordinate systems is the rolling angle of the aircraft. The analytic solution of the roll angle can be obtained by utilizing the relationship between the beam deflection angles of the two coordinate systems in the inclined vertical plane where the beam is located and performing reasonable approximate processing. At present, only a roll angle calculation method when no pitch angle exists in the side view surface of the aircraft is given. When the pitch angle exists in the side view surface of the aircraft, due to the fact that many unknown variables are involved, analysis and calculation of the roll angle of the aircraft need to be further analyzed and studied. However, the existing analysis result lays a reliable foundation for subsequent in-depth research.

This application discloses a trajectory planning method and apparatus based on artificial potential field theory. The trajectory planning method includes obtaining the longitudinal trajectory of a glider within its altitude-velocity profile and the roll angle value at each time point; during the glider's flight based on the longitudinal trajectory within the altitude-velocity profile, determining whether the current heading angle is within the angle coverage range of the threat zone; if so, controlling the roll angle sign based on the heading angle coverage range of the portion of the threat zone on both sides of the reference plane relative to the glider; if not, controlling the roll angle sign based on the heading angle deviation threshold, the tangent heading angle when the current heading is tangent to the edge of the threat zone, the target heading angle, and the current heading angle; and planning the lateral maneuver trajectory of the glider online based on the roll angle value, heading angle, and roll angle sign at each time point. This application solves the problems of complex implementation and local optima in existing artificial potential field methods.

The invention relates to a flight control and servo combined control method suitable for low-inertia strong control, and the method comprises the steps: enabling an airborne computer to transmit a dynamic pressure, an attack angle, a sideslip angle and a roll angle to an electric steering engine in real time based on the real-time state of a flight control system; the electric steering engine performs segmentation design on PD gains to obtain small-deviation PD gains, medium-deviation PD gains and large-deviation PD gains according to deviation fed back by a rudder instruction and a rudder surface position; the small-deviation PD gain is adaptively adjusted according to the dynamic pressure, the attack angle, the sideslip angle and the roll angle; and according to the medium deviation PD gain, the large deviation PD gain and the adjusted small deviation PD gain, the steering engine control system adopts a PD control algorithm of a rudder instruction and rudder feedback deviation to calculate a control voltage and convert the control voltage into a system duty ratio, and closed-loop control of a steering engine position loop is realized. According to the invention, the problem of large roll angular velocity jitter caused by poor small rudder deflection tracking performance of a small-size aircraft under the low-inertia strong-manipulation working condition is solved.

The application provides a threat area fly-around prediction-correction guidance method and system. The method comprises the following steps: constructing a reentry motion model and determining an initial state of an aircraft; determining a state value of a transition point between an initial descent segment and a glide segment according to a nominal attack angle and a constant roll angle acting on the model based on the initial state of the aircraft; in the longitudinal guidance design, taking the state value of the transition point as the initial state of the glide segment, determining a five-segment resistance acceleration profile according to a reentry range requirement, generating a longitudinal range analytical solution, determining a roll angle amplitude based on the resistance acceleration profile and a controller requirement, and inputting the roll angle amplitude into the lateral guidance; in the lateral guidance design, updating a heading angle corridor according to the state value of the transition point and the roll angle amplitude, determining a roll angle reverse logic for flying around the forbidden flight area, and outputting a roll angle sign. The application can improve the range calculation speed and adapt to different configurations and multiple threat areas.

It is provided a method of manufacturing a propeller having a plurality of blades, each of the plurality of blades having an intake portion, an exhaust portion, and a tip portion extending from the intake portion to the exhaust portion, the method comprising: defining a plurality of parameter sections by selecting some of the parameters including skew angle, roll angle, rake, radius, pitch angle, vertical angle values defining a parameter section at the transition from the intake portion to the tip portion by parameters to cause the amount of non-axial lift in the tip portion to be greater than the axial lift in the tip portion defining parameter sections to include a roll value of 90 degrees in the tip portion and extrapolating between parameter sections to form smooth lines to form a blade configured to form a loop when attached to a hub.

The present application relates to the technical field of aviation aircraft, in particular to a kind of longitudinal relaxation static stability aircraft no angle of attack stability augmentation control method, equipment and medium, comprising the following steps: obtaining the normal overload of aircraft by onboard inertial sensor;Dynamic pressure is obtained by onboard atmospheric data sensor;The weight of aircraft is calculated according to fuel quantity sensor information;Roll angle target value is provided by aileron channel control law solution;Stability augmentation control gain is determined;Each value obtained is substituted into formula, and elevator control command is calculated, and then aircraft is controlled to realize no angle of attack stability augmentation control.The present application proposes a kind of longitudinal relaxation static stability aircraft no angle of attack stability augmentation control method, and normal overload, dynamic pressure, aircraft weight and other information are used as control quantity, and normal overload control target is constructed using roll angle target value, so as to realize stability augmentation control.