109 results about "Aircraft dynamics" patented technology

Dynamic weather model systems and methods are operable to assess weather in proximity to an airborne aircraft. An exemplary embodiment receives a radar return from the weather, determines reflectivity information from the received radar return, retrieves a weather model from a weather model data base, compares the weather with the retrieved weather model and the determined reflectivity information, predicts a characteristic of the weather based upon the comparison of the weather and the weather model, and determines if the predicted characteristic is potentially hazardous to the airborne aircraft. The weather model is defined by at least one weather modeling algorithm, and is defined by at least one of a parameter and a variable parameter range residing in a weather characteristics database.

The invention discloses a small-size unmanned rotary wing aircraft dynamic model identification method based on an adaptive genetic algorithm, which relates to flight status data acquisition and optimization, dynamic model building and parameter identification, and parameter optimization validation. Firstly, status data and control data when a small-size unmanned rotary wing aircraft executes standard actions are acquired through a data acquisition system, and smoothing and filtration are conducted to eliminate wild values; then aiming at the operating characteristics of the small-size unmanned rotary wing aircraft at autonomous takeoff and landing stages, a small-size unmanned rotary wing aircraft dynamic model is built through a balance point linearization method and model parameters are identified through the adaptive genetic algorithm; and finally intelligent parameter evaluation indexes are built and the effectiveness of the model parameters is evaluated and judged through a one-step predication method. The small-size unmanned rotary wing aircraft dynamic model identification method based on the adaptive genetic algorithm solves the problem in the dynamic model identificationof the small-size unmanned rotary wing aircraft, can realize the high-accuracy control of the small-size unmanned rotary wing aircraft, and has the advantages of low testing cost, short cycle, simplecalculation, high dynamic model accuracy and weak dependence on initial values.

The invention provides a guidance law of a multi-constraint aircraft based on a Bezier curve, and belongs to the technical field of aircraft dynamics and guidance. According to a specified aircraft motion model and a three-order Bezier curve equation simulative ballistic trajectory, an attack angle and a heeling angle of the aircraft are determined by four controlling points with two middle controlling points described by parameters k1 and k2; a range of the parameters k1 and k2 meeting requirements of miss distance, collision angle and terminal attack angle is acquired through optimization; values of k1 and k2 corresponding to terminal speed maximum and terminal speed minimum are figured out; and according to a given terminal speed, the parameters of k1 and k2 are determined. Based on inverse dynamics, the terminal collision angle, the attack angle and speed constraint are all considered, requirements on the collision angle, the terminal attack angle and terminal collision speed can be met, and accurate attacks on fixed targets are achieved.

The invention provides an airplane continuously descending approach locus obtaining method and device; the method comprises the following steps: building an airplane descending locus model, an airplane dynamics model and an airplane oil consumption model, wherein the airplane descending locus model is used for representing an airplane descending process having three phases; building a first objective function based on fuel quantity according to the airplane descending locus model, the airplane dynamics model and the airplane oil consumption model; discretizing the first objective function so as to obtain a second objective function; using a difference evolution algorithm and sequence secondary planning algorithm combined method to solve the second objective function, thus obtaining the airplane continuously descending approach locus corresponding to the second objective function optimal solution. The airplane continuously descending approach locus obtaining method and device can obtain the airplane continuously descending approach locus with high accuracy and efficiency.

The invention discloses a flight simulation original state compound genetic balancing method. The flight simulation original state compound genetic balancing method comprises the following steps of initializing an aircraft constant straight line flat flight kinetic model and a constraint equation; step 2, obtaining an initial elevator deflection angle, motor thrust, an aircraft attack angle and an aircraft pitch angle; step 3 enabling a formula balancing result to be served as an initial value of a genetic algorithm, executing a genetic optimization algorithm and obtaining an optimal elevator deflection angle, the motor thrust, the aircraft attack angle and the pitch angle; step four solving the optimal elevator deflection angel and the motor thrust according to a mixed genetic optimization algorithm, reasoning backward to solve the displacement of a steering column and a throttle lever, configuring the steering column and the throttle lever and assigning the obtained aircraft attack angle and the pitch angle to the aircraft kinetic simulation model.

The invention discloses a carrier aircraft landing guiding half-physical emulating system, which comprises an automatic carrier landing system emulating computer with the function of go-around decision, and a radar emluator, an animation demonstrating computer and an steering engine system which are connected respectively with the emulating computer; the automatic carrier landing system emulating computer is responsible for three-dimensional carrier landing guiding law, flying control lay, aircraft dynamics, kinematic real-time computation and dynamic demonstration of a longitudinal lateral carrier landing track, respectively transmits inertial spatial aircraft track information, position coordinates of an aircraft and a carrier, and attitude information obtained by computation in real time to the radar emluator and the animation demonstrating computer, and exchanges data with the steering engine system; after the inertial spatial aircraft track information is processed by the radar emluator, the information is returned to the automatic carrier landing system emulating computer for go-around decision; and the animation demonstrating computer demonstrates the carrier landing process of the carrier aircraft according to the received position coordinates of the aircraft and the carrier and the attitude information in real time. The system of the invention has the advantages of convenient utilization, clear interface and the like.

The invention discloses an iterative propulsion disturbance domain updating method for aircraft dynamic aerodynamic characteristic simulation. Aiming at the problem that a large amount of invalid calculation exists in an existing aircraft dynamic aerodynamic characteristic numerical simulation method, three kinds of dynamic calculation domains of unsteady, convection and viscosity are adopted, anda solving thought that only non-convergent unsteady units are solved and the viscosity effect is only considered in a local area is provided, wherein the unsteady dynamic calculation domain only comprises grid units which play a leading role in the unsteady effect at each physical moment; the convection dynamic calculation domain only comprises grid units which are disturbed in each iteration step and are not converged yet, the viscosity dynamic calculation domain only comprises grid units which play a leading role in the viscosity effect in the convection dynamic calculation domain, and thethree types of dynamic calculation domains can be updated in a self-adaptive mode in each iteration step. By avoiding invalid calculation in unsteady numerical simulation and reducing the number of iteration steps required for inner iteration to reach a convergence state, the calculation efficiency of aircraft dynamic aerodynamic characteristic numerical simulation can be remarkably improved.

The invention provides a method for determining aircraft dynamic weight characteristic, which belongs to weight design in the field of aircraft overall design and is used for analyzing whole aircraft dynamic weight characteristic under overload condition. The method comprises the following steps of: S1, establishing an oil tank model and determining a clapboard model; S2, determining oil mass and liquid level position in the oil tank in the oil tank model established in the S1; S3, analyzing aircraft overload information, and determining load on fuel oil, empty aircraft usage and external load weight characteristic information; S4, performing discretization for the oil tank model established in the S1 according to the oil mass and the liquid level position in the S2; S5, in combination with the load on fuel oil determined in the S3, performing simulation for a finite element model obtained in the S4; S6, according to the fuel oil dynamic weight characteristic obtained in the S5, and in combination with the S3, obtaining whole aircraft dynamic weight characteristic data. The method of determining aircraft dynamic weight characteristic, provided by the invention, is a method capable of accurately tracing variation conditions of free surface, and calculating the dynamic weight characteristic of the fuel oil and even the whole aircraft.

The invention relates to an aircraft dynamic scale model landing gear bumper spring design method. The method comprises the following steps of 1, estimating a ground load and calculating landing gear loads, wherein the landing gear loads comprise a nose landing gear load Fn and a main landing gear load Fm; 2, scaling the size of a landing gear bumper according to a predetermined proportion; 3, fitting a static pressure curve of the landing gear bumper into a linear curve; 4, determining bumper spring design constraint parameters according to a predetermined proportion; 5, establishing a bumper spring effective parameter calculation formula; 6, iteratively computing bumper spring effective parameters; and 7, selecting an optimum parameter group. According to the aircraft dynamic scale model landing gear bumper spring design method, landing gear bumpers can be simplified to the greatest extent under the premise of ensuring the mechanical properties and model sizes of the landing gear bumpers, aircraft dynamic scale model landing gear bumpers are designed most really, the design and processing costs are reduced, and the dynamic requirement and similarity requirement are satisfied.

The invention discloses a four-rotor aircraft dynamics model/airborne sensor combination navigation method, which belongs to the field of combination navigation. The four-rotor aircraft dynamics modelis combined with an airborne sensor, angular velocity, attitude, speed and positional information are estimated, in the invention, the dynamics model comprises a lift model, a resistance model, a torque model, a roll moment model, and a pitching moment model, and the employed airborne sensor comprises a rotor rotating speed sensor, a magnetic sensor, a satellite navigation system, and an air pressure height meter. In the method, a common inertia sensor such as a gyroscope and an accelerometer in the current four-rotor aircrafts are not employed, the acceleration and the angular acceleration information are output by the dynamics model of the four-rotor aircraft, and the inertia sensor can be substituted. The method can be used for solving the navigation problem of the four-rotor aircraftwhen the inertia sensor is failed.

The invention relates to an aircraft dynamics monitoring method based on frequency response analysis, and belongs to the technical field of aeroelastic test. The aircraft dynamics monitoring method based on frequency response analysis comprises the following steps: step one, acquiring the reference frequency fri and the inspection frequency fti of an aircraft structure in an ith-order mode; and step two, acquiring frequency weigh errors of an aircraft flutter characteristic according to a weight error formula. According to the external field dynamics monitoring method based on frequency response analysis, dynamics monitoring can be carried out on an external field by simple frequency examination through application frequency response function analysis and mode weight error analysis, thus,the dynamic characteristic of an aircraft in a using process is obtained and is applied to assessment of the flutter characteristic of the aircraft, and finally, the use envelope of the aircraft is given to guarantee flight safety of the aircraft.

The invention discloses a gravity anomaly extended interpolation method of a nearspace large-range maneuverable trajectory.The method includes the following steps that firstly, a mathematical model of envelope lateral width and longitudinal range of the trajectory is built; secondly, a pole changing coordinate system is built; thirdly, an aircraft dynamic model in the pole changing coordinate system is built; fourthly, a trajectory envelope of the pole changing coordinate system is calculated; fifthly, whole airspace and spatial domain subdivision is conducted on the pole changing coordinate system; sixthly, a common coordinate system gravity anomaly whole airspace reconstructed model is built; seventhly, a gravity anomaly partial reconstructed model along the trajectory is built; eighthly, a subdomain partial curve coordinate system is built; ninthly, gravity anomaly extended interpolation calculation is conducted in the flight process.The method has the advantages of being high in interpolation precision, high in calculation speed and small in data storage amount, can meet the requirements for gravity anomaly quick calculation along any nearspace large-range maneuverable trajectory, and has the capacity of quick flight task responding and temporary changing.

A method and apparatus is disclosed for displaying a dynamic parameter of an aircraft, the apparatus comprises a display unit receiving a display signal and displaying a scale that changes dynamically and non-linearly in accordance with a selected display algorithm, the display unit further displaying a pointer pointing to said scale in accordance with a reading of said dynamic parameter, thereby emphasizing a range of said reading of said dynamic parameter.