70 results about "Aircraft flight mechanics" patented technology

An aircraft display and control system generally includes a processor, a cursor control and selection device, a multifunction keyboard, an aeronautical information database, a geographic database, and a plurality of display devices. Users, such as an aircraft pilot and copilot, can perform flight plan entry and modification by manipulating graphical and textual information on the display devices using the cursor control device and the multifunction keyboard. In one embodiment, the present invention provides an interface to flight deck information systems that is intuitive and easy to use, and which enables flight crew members to easily and effectively enter and modify data displayed by the aircraft flight deck, substantially centralizes avionics controls within one controller, and decreases flight crew “heads-down” time.

An aircraft flight risk measuring system for analyzing risks related to a flight of an aircraft. A user of the risk measuring system can be a flight dispatcher, an owner / operator, a pilot and other interested parties. The risk measuring system includes a risk management server system computer. The system computer has a two-way communication with a user computer operated by the user. An accident history database is connected to the system computer for providing accident reports related to the aircraft and other accident data. Also, a navigation database is connected to said system computer for providing airspace data, radio navigation aids, preferred routes, elevation data, geographic data and information related to a destination airport. Further, a non-static database is connected to the system computer for providing live information related to weather forecasts and data related to the aircraft's flight. As an option, a two-way communication between said system computer and an aircraft computer on board the aircraft can be included. The two-way communication used for receiving and transmitting encoded data from the aircraft when the flight is in progress.

The invention provides a rotor wing dynamic test device which comprises a model rotor wing system, a pitching motion simulation mechanism, a rolling motion simulation mechanism, a course rotation simulation mechanism, an x-y-z axis horizontal moving device and a test accessory device. Seven independent control servo motors are used for the rotor wing dynamic test device and used for respectively driving longitudinal, lateral and transverse linear motions, pitching, rolling and course rotation motions and rotation of a model rotor wing. The seven servo motors are respectively provided with an independent control system, six-freedom-degree motion laws and the rotor wing rotating speed can be dynamically combined to be controlled under control of a central industrial personal computer, model test study on influences of dynamic maneuvering flight of helicopters on the aspects such as rotor wing pneumatic and flight mechanics and dynamics can be conducted, and motion combination tests among different degrees of freedom can be carried out according to needs.

The invention discloses a calculation method of a tilt corridor of a tilt four-rotor aircraft with periodic variable pitch at constant rotational speed, which comprises the following steps: establishing flight mechanics models of various parts of the aircraft, converting all forces and moments into a body shaft system, and establishing a six-degree-of-freedom balance equation; determining the tiltcontrol mode; solving the upper and lower boundaries of the pitch angle by taking the lift characteristics of the wing as the limiting factor; obtaining the control variables and attitude angles under different state conditions by optimal trimming; finding the corresponding wing lift characteristic limit minimum speed boundary and wing lift characteristic limit high speed boundary; obtaining theavailable power limiting high-speed boundary by solving the required power at different state points. The invention is based on a flight mechanics model, takes into account the parameters of each aspect of the aircraft design, and reflects the specific attributes of each feasible flight point. The invention is also applicable to the tilt corridor calculation of tilt four-rotor aircraft under different control modes.

The invention discloses a dynamics modeling and analyzing method for an aerospace vehicle and belongs to the technical field of flight dynamics modeling and simulation analysis of aircrafts. According to the dynamics modeling and analyzing method, the variable-span and variable-sweepback aerospace vehicle is simplified and is composed of a fuselage, a left wing inner side, a left wing outer side, a right wing inner side and a right wing outer side, wherein the fuselage, the left wing inner side, the left wing outer side, the right wing inner side and the right wing outer side are independent rigid bodies. The deformation movement of wings relative to the fuselage is represented through a set of constraint equations, so that the freedom degree of a dynamics model is reduced; the aerodynamic force exerted on the aerospace vehicle is simplified to act on the fuselage only, and the dynamics model is established through the Kane method; related items derived from the deformation movement are extracted from the dynamics model to form additional force and additional force moment so that the influence on the dynamics characteristics of the aerospace vehicle by the inertia force and inertia moment caused by the deformation moment can be represented. The dynamics model of the aerospace vehicle is simplified through the dynamics modeling and analyzing method, and integral items and derivation terms of the inertia moment do not appear in the equations. In this way, dynamics modeling and simulating can be easily performed on the aerospace vehicle, and meanwhile accuracy is high.

The invention provides a flight control method of a deformable unmanned aerial vehicle. The deformable unmanned aerial vehicle has a canard type telescopic forward-swept wing and single vertical fin aerodynamic arrangement; according to the adjustment and analysis of a root-locus method, a high-low speed deformable unmanned aerial vehicle can perform the adjustment on parameters of a controller through a design method of a low-speed controller, and therefore the design requirement on the controllers under two flight speeds can be met. Matching the deformation time and the control time of an engine is part of a control strategy of the deformable unmanned aerial vehicle; and as shown in the analysis of the flight mechanics design, the thrust of the engine is changed prior to implementing the process and the strategy under the deformation time.

The invention provides an aircraft flight attitude visual image recognition method comprising the steps that an aircraft flight real-time image or an offline single frame / sequence image is acquired toact as an input image, and the image is denoised by using adaptive median filtering; the denoised image is binarized by using by using a moving mean adaptive threshold method; the foreground is processed by using morphological operation and the holes of the area are filled so as to acquire each connected area, and preliminary screening is performed according to the aircraft feature similarity index so as to acquire a suspected target set; the concave angle points and the convex angle points of each target in the suspected target set are detected by using the multi-scale interior angle point operator, and the concave and convex angle points of each target are connected so as to form the contour of each target; the multi-feature fusion matching index is established according to the contourfeatures of the target to be detected, and the contour of the aircraft to be detected is recognized from the suspected target set by using the index; the angle code template image most similar to thecontour of the target aircraft is acquired from a simulation template library by using the local predictive searching and template matching strategy; and the flight attitude parameters of the currentframe of aircraft are solved from the matched angle code image.

A cockpit for an aircraft includes a windscreen through which light may pass, at least one seat spaced from and facing the windscreen, a flight deck having at least a portion disposed below the windscreen and having at least one head down display having an adjustable brightness that may be set by a brightness signal, a camera having a field of view including at least a portion of the at least one seat and outputting an image signal indicative of luminance information within the field of view, and a processor operably coupled to the camera and the head down display and configured to receive the image signal, determine a luminance of at least a portion of the field of view, determine a brightness for the head down display based on the determined luminance, and outputting to the head down display a brightness signal corresponding to the determined brightness.

The invention belongs to the technical field of high-speed wind tunnel experiments and flight mechanics, and particularly relates to a device for testing the external store separation performance of an aircraft in a high-speed wind tunnel. The device comprises a wind tunnel experiment section, a wind tunnel diffusion section, a main machine model, an external store model, a tail supporting rod andsix digital displacement cylinders; a multi-component balance is arranged in the external store model; a rolling driving device is arranged in the tail supporting rod; an output shaft of the rollingdriving device is connected with the multi-component balance; the rear ends of the digital displacement cylinders are connected to the rear part of the wind tunnel experiment section or the front partof the wind tunnel diffusion section through rear ball hinges; the front ends of the digital displacement cylinders are connected to the tail supporting rod through front ball hinges; the displacements of the digital displacement cylinders can achieve the movement, pitching and yaw of the external store model; and the rolling driving device can drive the external store model to roll. The invention further provides a method for testing the external store separation performance of the aircraft. The sufficient rigidity of the tail supporting rod can be ensured, and the blockage degree of the wind tunnel can be reduced as much as possible.

Fluid Mechanics of Inertia incorporates all the variables of flight mechanics, interplanetary ballistics, perpetual motion, human powered transportation, and all aspects of fluid mechanics involving matter, energy, and motion, and includes elastic fluids into the definition of “fluids”. The “Mechanics” in Fluid Mechanics, may mistakenly be construed to be dynamics. Fluid Mechanics of Inertia is simplified calculus and converts calculus to trigonometry, logarithms, geometry, algebra, and basic functions. Fluid Mechanics of Inertia is a three dimensional math occurring in a time frame of no elapsed time although the locomotive animation of life is at full locomotive performance, the “snap-shot” elapsed time is zero, except where elapsed time is derived to elapse a time interval of a prescribed time allowance duration which is described, and there is no motion, thus “dynamics” does not occur in these problems since time is zero.

The invention discloses an airplane gravity center envelope calculation method. The method comprises the steps that original parameters of an airplane are input, and the loading condition and the initial gravity center position are given; according to the aircraft aerodynamic model, the undercarriage model, the engine model and the full-aircraft flight mechanical model, calculating full-aircraft aerodynamic force and aerodynamic moment of the aircraft, and finishing trimming; the load weight and the gravity center position are dispersed; judging whether the aircraft can normally cruise and flyand normally take off and land or not; and then introducing static stability margin constraint and wing and horizontal tail limit load constraint under the condition of large aircraft load under thecondition of minimum horizontal flight speed to obtain a relation graph of front and rear gravity center limits and aircraft flight weight. According to the method, discretization processing is carried out on the load gravity center coordinate point, the flight process and the take-off and landing process of the aircraft are simulated, the gravity center of the aircraft is accurately calculated, the process is simple, and calculation is convenient.

According to the invention, the system for determining, in real time, flight parameters of an aircraft, in the course of a flight of the latter, comprises an extended Kalman filter (10) which is configured on the basis of flight mechanics equations establishing a dependency relationship between at least two preselected flight parameters of said aircraft to be estimated and which is formed so as to deliver, during a flight of said aircraft, joint estimations of said selected flight parameters.

A method of managing a flight procedure of an aircraft flight plan, which has at least one segment, includes a step in which an additional attribute of the segment representative of a compulsory or optional nature of the flight of said segment is dynamically selected.

Disclosed are a method and system for improving real-time determination of flight parameters of an aircraft during flight. A selection module selects and delivers measured values of input parameters to an estimation unit, and the estimation unit uses the input parameters to estimate at least two selected flight parameters. The input parameters are received by the estimation unit on the basis of identified flight mechanics equations associated with the selected flight parameters. The estimation unit includes an extended Kalman filter, which is configured to receive the values of the input parameters and to output joint estimations of the selected flight parameters based on the flight mechanics equations, during the flight of the aircraft.

The invention provides a trajectory capture test system for simultaneous separation simulation of double external stores, and belongs to the field of high-speed wind tunnel test technology and aircraft flight mechanics. The trajectory capture test system comprises a half-arm attack angle mechanism installed on a wind tunnel test section, and an aerial carrier model is fixedly connected with the half-arm attack angle mechanism through an aerial carrier supporting rod; a missile-loaded base is connected with the wall surface of the wind tunnel test section through a vertical movement mechanism, parallel bottom plates are connected with the missile-loaded base through rotating shafts, three-degree-of-freedom parallel mechanisms are installed on the parallel bottom plates respectively, the other ends of the three-degree-of-freedom parallel mechanisms are connected with a movable platform, and a rolling driving mechanism is arranged in the movable platform; and the movable platform is connected with missile-loaded balances and missile-loaded models through missile-loaded support rods. According to the invention, the two missile-loaded balances are used for respectively measuring the aerodynamic force and torque applied to the missile-loaded models, the motion modes of the two missile-loaded models are inversely solved, the motion of each driving device is controlled, the multi-degree-of-freedom position and attitude change of the two missile-loaded models is realized in a wind tunnel test section, and the simulation of the motion trajectory of the double-external store model is realized.

A cockpit for an aircraft includes a windscreen through which light may pass, at least one seat spaced from and facing the windscreen, a flight deck having at least a portion disposed below the windscreen and having at least one head down display having an adjustable brightness that may be set by a brightness signal, a camera having a field of view including at least a portion of the at least one seat and outputting an image signal indicative of luminance information within the field of view, and a processor operably coupled to the camera and the head down display and configured to receive the image signal, determine a luminance of at least a portion of the field of view, determine a brightness for the head down display based on the determined luminance, and outputting to the head down display a brightness signal corresponding to the determined brightness.

The invention discloses a space stability analysis method of an aircraft control system. The method comprises the following steps: 1), acquiring aerodynamic data; 2), solving aerodynamic derivatives on the basis of the aerodynamic data, and establishing a coupled aircraft motion equation between channels; 3),performing linearization analysis; 4),testing linearization analysis performance and improving results. Overall assessment and comprehensive design of aircraft flight tests can be achieved with the method, cause analysis and design improvement direction are provided for various instabilities of the aircraft under specific conditions, algorithm design work can also be promoted, various coupling flight states can be taken into account more comprehensively, and the method has obvious significance for accelerating the development progress of aircrafts.

The invention discloses an aeromodelling aircraft, which includes a fuselage, a first wing surface, and a second wing surface. The fuselage is provided with a power system, the power system is connected to the first wing surface and the second wing surface through a transmission device so as to drive synchronous flapping of the first wing surface and the second wing surface under the action of the power system. The first wing surface and the second wing surface both have peak valley joints, the peak valley joints are disposed on the wing surfaces so as to facilitate formation of peak valleys during flapping of the wing surfaces. The invention also discloses a remote control aeromodelling flight system, which is characterized in that the aircraft includes an aeromodelling aircraft and a corresponding remote control device. The aeromodelling aircraft provided by the invention can simulate the flight action of organisms, not only enriches aeromodelling aircraft flight modes, but also plays a promoting role in development of artificial flight technologies.

The invention relates to a flight dynamics modeling and analysis method for a wingtip hinge combined type flight platform, and belongs to the technical field of aircraft flight mechanics modeling andsimulation analysis. According to the method disclosed in the invention, the sub-aircrafts in the wingtip hinged combined type flying platform are treated as rigid bodies; the degree of freedom of motion is expressed in the form of absolute coordinates relative to an inertial coordinate system, a multi-body dynamics theoretical analysis method is introduced, a constraint equation is established incombination with a Lagrange multiplier method to express an articulation relationship, and a Newton-Euler equation is utilized to establish a flight mechanical equation of the flight platform. The invention provides a method and process for analyzing the flight mechanical characteristics of the wingtip hinge combined type flight platform. Aircraft trimming and flight mechanical stability analysiscan be carried out.

The invention relates to a method for detecting and eliminating "outlier values" of flight data based on aircraft motion equations, belonging to the technical field of flight mechanics and information processing, and used for checking whether there are "outlier values" in flight data that are greatly different from other measured values. In order to overcome the misjudgment of "outlier value" caused by the prior art method only relying on its own measurement data, the present invention proposes a method for judging and eliminating the threshold value of "outlier value" based on the aircraft motion equation. The speed of change is to predict the flight state and parameters step by step according to the effectively simplified output response, and judge and eliminate the "outlier value" in the measurement according to the set threshold value at each step; due to the possibility of "outlier value" in the fast-changing state or parameters in the actual measurement The characteristic is larger than the slow variable, and the output response has engineering precision. The method proposed by the invention is more effective for eliminating flight data "outliers".

An aircraft standby display device includes a set of data sources, a data processing unit, and a display. The display presents altitude control indicators representing speed, altitude, attitude, and a heading scale that includes a characteristic symbol indicating a next route point of an aircraft flight plan. The heading scale is independent of the altitude control indicators. The display also presents a quantitative value that: (1) indicates the remaining distance of flight of the aircraft to reach the next route point and (2) is associated with the characteristic symbol.

A method of detecting the misalignment of a first point (1) in space is disclosed. The method includes a trajectory of an aircraft flight plan, wherein the flight plan has two points (2, 3) in succession and preceding the first point (1) and two points (4, 5) in succession and subsequent to the first point (1), A first acute angle (θ1) is calculated which is formed by a first segment (LEG-1) joining the points (2, 3) and a second segment joining the third (3) and the fourth (4) points. A second acute angle (θ2) is calculated which is formed by the second segment and a third segment (LEG-4) joining the fourth (4) and the fifth (5) points. A polygon is calculated which is dependent on a first distance (DH) defined between the fourth point (4) and the straight line including the first segment (LEG-1) and on a second distance (DD) defined between the projection (4′) of the fourth point (4) on the straight line comprising the first segment (LEG-1) and the third point (3).

The invention belongs to the technical field of helicopter flight tests, and discloses a civil helicopter marine lifesaving performance verification test flight method. The method comprises the following steps: formulating test flight requirements and test requirements of risk subjects such as offshore lifesaving according to regulations such as CCAR29R1 in Airworthiness Specification of Transportation Rotorcrafts, evaluating the influence of offshore turbulence on a helicopter by adopting a helicopter flight mechanics theoretical model, and designing test flight subjects and a test flight method according to an analysis result of the theoretical model; pre-test flight is carried out by using a simulator, a test flight process is verified, so that a test flight risk is controlled, and a complete data processing method is given according to test data in combination with test flight unit evaluation; and a civil helicopter marine lifesaving performance verification subject test flight technology is formed.

The invention relates to the field of aircraft high-lift system design and flight mechanics, in particular to a method for predicting a test-flight fault load, and aims to solve the problems that the existing method for predicting test-flight fault load is long in consumed time and low in calculation accuracy. According to the method for predicting the test-flight fault load of the invention, based on the relationship between the aircraft high-lift system test-flight fault load and two weight data of an aircraft, a test-flight load of a fault location can be calculated through the relatively easily obtained two weight data of the aircraft, so as to determine the relationship between the test-flight load of the fault location and a design required load. The method can provide a reliable technical support for the high-lift system design, the fault location and the like, can reduce the uncertainty of original method, and can shorten the troubleshooting and designing time.