58 results about "True airspeed" patented technology

A system and method are provided for calculating Mach number and true airspeed without reference to data from a pitot static sensor. The true airspeed and Mach number are calculated using the altitude information from GPS, IRS, Radio Altimeter and other onboard sensors other than the air data computer (ADC). The computed true airspeed or Mach number could be used to confirm the ADC information or in lieu of the ADC information when the ADC information is unreliable or unavailable.

The object of the present invention is to provide an air data sensor that does not require an external input of a reference velocity as a Doppler LIDAR, has a function of autonomously determining the absolute airspeed, and has no position error. The optical air data sensor in accordance with the present invention is an optical air data sensor, mounted on an aircraft, for emitting a laser light as a transmission light into atmosphere, and then receiving a laser scattered light produced by scattering of the laser light by aerosol present in the atmosphere as a reception light, thereby to measure an airspeed and a wind velocity of airflow in a distant region on the basis of a Doppler shift amount between the transmission light and the reception light, wherein a true airspeed is autonomously measured, without setting a reference velocity, by successively sweeping a frequency offset corresponding to a reference velocity for providing an offset to a measurement frequency, and performing this sweeping within a frequency range in which the Doppler shift amount is measured.

The invention discloses an estimation method of an atmosphere angle of attack and an angle of sideslip in a high-angle-of-attack flight status. Under the premise that the structure of an aircraft is not changed and additional measuring devices and hardware devices are not added, the parameters output by the existing onboard strapdown inertial navigation system, a GPS (Global Position System) system, a flight control system and a fuel measuring system are fully used for resolving the angular acceleration, the rotational inertia, the product of inertia, the coefficient of a moment equation set and the total mass of the aircraft; based on a flight dynamics model, by establishing an extended Kalman filter, and integrating the processes of resolving the flight dynamics model and the status estimation together, accurate estimation of the current angle of attack, the angle of sideslip and the true airspeed is realized, and then the current air speed and the change rate of the air speed are resolved and input into the extended Kalman filter as the feedback information so as to finish the real-time accurate estimation of the angle of attack and the angle of sideslip of next time; the real-time accurate estimation of the angle of attack and the angle of sideslip is realized by a recursive resolving mode, measurement range and measurement precision of an atmosphere data system are improved, and adaptability to complex flight environment is also improved.

A navigation and guidance computer system using navigation sensors to determine aircraft position, and automatically control the aircraft to fly from an initial position to intercept and follow a crew-selected rendezvous orbit in accordance with Air Traffic Control and military specified rules. It creates maneuvers that result in smooth, efficient motion of the aircraft from an initial position to the execution of a rendezvous orbit. It controls the aircraft by applying commands to the aircraft's autopilot. The system creates the smooth and coordinated maneuvers by considering and anticipating all constraints on the aircraft flight path by creating a single maneuver that is a re-entry maneuver, an over fly maneuver, and a turning maneuver all as a single turn that is contained entirely within the allowable airspace. All maneuvers are created and executed recognizing the turning capability of the aircraft considering allowable bank angles, current wind conditions and true airspeed.

A system and method are provided for calculating Mach number and true airspeed without reference to data from a pitot static sensor. The true airspeed and Mach number are calculated using the altitude information from GPS, IRS, Radio Altimeter and other onboard sensors other than the air data computer (ADC). The computed true airspeed or Mach number could be used to confirm the ADC information or in lieu of the ADC information when the ADC information is unreliable or unavailable.

A system for calculating airspeed and dynamic pressure comprises a system body, an internal accelerometer, located within the system body, an internal pressure sensor, located in the system body, the internal pressure sensor being not hermetically sealed within the system body and capable of measuring the static pressure of the ambient atmosphere, and a processor in reception of the internal accelerometer, and the internal pressure sensor, capable of calculating Mach number via an axial acceleration, and capable of calculating a dynamic pressure and a true airspeed via the Mach number.

The object of the present invention is to provide an air data sensor that does not require an external input of a reference velocity as a Doppler LIDAR, has a function of autonomously determining the absolute airspeed, and has no position error. The optical air data sensor in accordance with the present invention is an optical air data sensor, mounted on an aircraft, for emitting a laser light as a transmission light into atmosphere, and then receiving a laser scattered light produced by scattering of the laser light by aerosol present in the atmosphere as a reception light, thereby to measure an airspeed and a wind velocity of airflow in a distant region on the basis of a Doppler shift amount between the transmission light and the reception light, wherein a true airspeed is autonomously measured, without setting a reference velocity, by successively sweeping a frequency offset corresponding to a reference velocity for providing an offset to a measurement frequency, and performing this sweeping within a frequency range in which the Doppler shift amount is measured.

A method and aircraft system are provided for determining an optimal gliding speed that maximizes gliding distance of the aircraft upon engine failure, and comprises calculating the optimal gliding speed from a speed polar curve, true airspeed, heading, and vertical wind speed. The polar curve may be constructed with consideration of the aircraft center of gravity and weight. The calculated gliding speed, which may be restricted to a threshold, is provided to the pilot and optionally to an autopilot. The method is repeated periodically, or if the wind vertical or horizontal velocities exceed a threshold for a predetermined period of time.

An inertial navigation system which includes a first control system and a second control system is described. The first control system is configured to estimate a vertical velocity based in part on received vertical acceleration data. The second control system is configured to receive the estimated vertical velocity from the first control system and determine a compensated vertical velocity utilizing the estimated vertical velocity and airspeed.

The invention relates to a bionic polarization / inertial / atmospheric data integrated navigation system. The system consists of a navigation sensor, a solving module, an interface circuit, a microprocessor, a storage module and a power supply module. The sensor is composed of a biomimetic polarization sensor, an inertial measurement unit and an atmospheric data system. Specifically, the bionic polarization sensor includes an optical camera, a motor, a polarizing device and other components, and the atmospheric data system is composed of a pitot-static sensor, a pitot-static pipeline and a totaltemperature circuit. The work flow includes that: the sensor module collects the polarized light information, atmospheric static pressure, full pressure and other data in the atmospheric environment,and transmits the data to the solving module, and the processor determines whether the sensor works reliably, selects a working mode, and works out the course angle, true airspeed, height and other information for compensation correction of the attitude and speed of the navigation system, and stores and outputs navigation parameters. The integrated navigation system provided by the invention has the advantages of flexible navigation mode, strong autonomy, small volume and the like, and has interference suppression, fault detection and system reconstruction functions.

A navigation and guidance computer system using navigation sensors to determine aircraft position, and automatically control the aircraft to fly from an initial position to intercept and follow a crew-selected rendezvous orbit in accordance with Air Traffic Control and military specified rules. It creates maneuvers that result in smooth, efficient motion of the aircraft from an initial position to the execution of a rendezvous orbit. It controls the aircraft by applying commands to the aircraft's autopilot. The system creates the smooth and coordinated maneuvers by considering and anticipating all constraints on the aircraft flight path by creating a single maneuver that is a re-entry maneuver, an over fly maneuver, and a turning maneuver all as a single turn that is contained entirely within the allowable airspace. All maneuvers are created and executed recognizing the turning capability of the aircraft considering allowable bank angles, current wind conditions and true airspeed.

A flight display for an aircraft includes a virtual ice accretion meter having a liquid water content portion and an ice thickness portion. A method for determining ice accretion includes integrating over a time period a constant K multiplied by a liquid water content (LWC) and true airspeed (TAS).

An inertial navigation system which includes a first control system and a second control system is described. The first control system is configured to estimate a vertical velocity based in part on received vertical acceleration data. The second control system is configured to receive the estimated vertical velocity from the first control system and determine a compensated vertical velocity utilizing the estimated vertical velocity and airspeed.

The invention belongs to the aviation meteorology field and discloses an updraft measurement method of an airplane used in weather modification. By using a traditional method, updraft can not be settled. In the invention, the above problem is solved. The method comprises the following steps of 101, acquiring the true airspeed Vt, the angle of attack alpha and the sideslip angle beta of fixed wingairplane flight; 102, converting a vector VtSb=(VtX, VtY, VtZ) into a ground coordinate system from a body coordinate system; 103, acquiring the three-dimensional ground speeds VgXg, VgYg and VgZg, the pitch angle theta, the roll angle phi and the course angle psi of the fixed wing airplane flight; and 104, in the ground coordinate system, subtracting a three-dimensional airspeed from a three-dimensional ground speed and acquiring the updraft. In the invention, the updraft can be accurately measured; the measurement of the updraft is performed in an artificial precipitation field, which can beused to determine an airflow direction and a size in cloud and can be used for precipitation work scheme making; and catalytic efficiency is increased.

The invention discloses a flight record system for a plane and a data processing method thereof, relates to the technical field of aircrafts, and aims to solve the technical problem involved in monitoring of plane working conditions. The system comprises a ground communication terminal as well as a self-contained power supply, an onboard satellite communication terminal and two trigger switches which are arranged on the plane, wherein the first trigger switch is a mechanical micro switch; the second trigger switch is a bipolar latch Hall switch; the onboard satellite communication terminal is connected with the ground communication terminal through a satellite; the power supply port of the onboard satellite communication terminal is connected to an onboard power supply and the self-contained power supply respectively through the two trigger switches; the first trigger switch is arranged in the undercarriage cabin of the plane, and is triggered to be switched on / off through an undercarriage; the second trigger switch is arranged in a true-airspeed meter, and is triggered to be switched on / off through two magnets arranged on the inner wall of a dial plate; flight working condition data can be transmitted to the ground communication terminal in real time through the onboard satellite communication terminal. The system and the method contribute to ground monitoring of the plane working condition.

The invention discloses a high-sensitivity optical-type airspeed measuring device which comprises an airspeed pipe, a shell body and an opening sylphon. An optical curve sensing unit is contained inside the shell body, wherein the optical curve sensing unit comprises a curve-type test channel and a measuring unit, the measuring unit is connected with a disposing unit, and the disposing unit is connected with a display unit. The curve-type test channel comprises a curve-type support, multiple side A transformation teeth and multiple side B transformation teeth. The multiple side A transformation teeth and the multiple side B transformation teeth are arranged in a crossing mode, and the curve-type test channel which is provided for one or more signal optical fibers passing is formed among heads of the multiple side A transformation teeth and the multiple side B transformation teeth. The multiple side A transformation teeth and the multiple side B transformation teeth are arranged on two sides of the signal optical fibers correspondingly. A movable point of the opening sylphon is fixedly connected with one end of the curve-type support, and the other end of the curve-type support is fixed on the inner wall of the shell body. The high-sensitivity optical-type airspeed measuring device is simple in structure, reasonable in design, convenient to process and manufacture, low in cost and flexible in using manner, high in sensitivity and good in using effect.

An Energy State Awareness System that alerts the pilot of an aircraft during the climb, cruise and descent phases of a flight of low airspeed conditions, and during the Approach / Landing and Go-Around phases of flight when airspeed has deviated from a pre-determined reference airspeed that is considered “Stable” for the flight conditions. In order to determine the degree of deviation, the system monitors some of various readily available signals representative of flight parameters and of the aircraft's configuration such as true airspeed, weight, flap position, center of gravity and normal load factor—nz. Some signals are used directly, others are used as inputs to internal algorithms that compute the data necessary to determine the magnitude of the deviations.

The present invention provides a method and system for optimizing rotor flapping, which utilizes the collective stick position and true airspeed to determine a commanded longitudinal stick position.

A method for measuring a windspeed vector is described. A true airspeed vector of a flying machine is measured while the machine is in flight using one or more nanowires on the flying machine. Each nanowire is configured to measure a value of local air velocity relative to the flying machine. A velocity of the flying machine relative to the ground is measured while the machine is in flight, and then (a) the true airspeed vector is subtracted from (b) the velocity of the flying machine relative to the ground. Other applications are also described.

The invention provides an atmosphere data ground testing system. The invention aims to provide a testing system which can test the dynamic characteristics of optical atmosphere speed sensors in a ground environment and has low requirements for environments. The technical solution of the invention is characterized in that an atmosphere data computer, a GPS / Beidou resolver and an inertial navigationapparatus are integrated in a case; a combined wind speed measurement apparatus measures the wind speed and direction of relative airflow in the horizontal direction and the vertical direction at measurement points to provide benchmark data of the true airspeed, angle of attack and angle of sideslip for an optical atmosphere speed sensor; a rotating platform changes the included angle between theoptical atmosphere speed sensor and the relative airflow by rotating a stipulated angle; the optical atmosphere speed sensor makes emitted laser beams converge at a focus, and the projection speed ofthe macroscopic atmosphere relative movement speed Vc in the axis direction of each laser beam at the focus is detected according to the principle of laser scattering; the projection speeds of more than three different directions are combined to obtain testing data.

The invention provides a true airspeed reconstruction strategy for an unmanned aerial vehicle. Under two conditions that a mach number fails and a total temperature sensor fails, true airspeed reconstruction is performed in different ways; when the mach number fails, the mach number reconstruction is performed according to the eastward speed, the northward speed and the skyward speed of a GPS and the GPS height; when the total temperature sensor fails, the true airspeed is directly obtained by look-up of a mach number and air pressure altimeter or the true airspeed is reconstructed according to the eastward speed, the northward speed and the skyward speed of the GPS. The true airspeed information validity is ensured, and the true airspeed is calculated by fully utilizing atmosphere data information and GPS data information.

A system for calculating airspeed and dynamic pressure comprises a system body, an internal accelerometer, located within the system body, an internal pressure sensor, located in the system body, the internal pressure sensor being not hermetically sealed within the system body and capable of measuring the static pressure of the ambient atmosphere, and a processor in reception of the internal accelerometer, and the internal pressure sensor, capable of calculating Mach number via an axial acceleration, and capable of calculating a dynamic pressure and a true airspeed via the Mach number.

A system may include a static air temperature probe attached to an aircraft, an electronic flight instrument system, and a processor. The processor may be configured to measure a static air temperature at the aircraft using the static air temperature probe. The processor may further be configured to calculate a Mach number associated with the aircraft based at least partially on the static air temperature. The processor may also be configured to calculate a true air speed of the aircraft based on the Mach number. The processor may display an indication of the true air speed using the electronic flight instrument system. The processor may also be configured to calculate the speed of sound based at least partially on the static air temperature.

The invention belongs to the helicopter atmosphere system field, particularly to an air speed correction method for helicopter maneuvering flight, and solves the problem of air speed change in reverse direction when a helicopter performs large maneuvering movements like changing from rapid downslide to rapid climbing. An inertia sensor on the helicopter measures a vertical acceleration a; the vertical acceleration a opens a mobile points window for low pass filtering processing; the vertical acceleration is corrected after filtering processing to obtain a correction air speed that is close to a real air speed. <{EN3}>The correction of the air speed that is has a local distortion benefits controlling the flight stability by the automatic stationary system, reducing pilot's control burden and improving flight comfort level.

The invention discloses an estimation method of an atmosphere angle of attack and an angle of sideslip in a high-angle-of-attack flight status. Under the premise that the structure of an aircraft is not changed and additional measuring devices and hardware devices are not added, the parameters output by the existing onboard strapdown inertial navigation system, a GPS (Global Position System) system, a flight control system and a fuel measuring system are fully used for resolving the angular acceleration, the rotational inertia, the product of inertia, the coefficient of a moment equation set and the total mass of the aircraft; based on a flight dynamics model, by establishing an extended Kalman filter, and integrating the processes of resolving the flight dynamics model and the status estimation together, accurate estimation of the current angle of attack, the angle of sideslip and the true airspeed is realized, and then the current air speed and the change rate of the air speed are resolved and input into the extended Kalman filter as the feedback information so as to finish the real-time accurate estimation of the angle of attack and the angle of sideslip of next time; the real-time accurate estimation of the angle of attack and the angle of sideslip is realized by a recursive resolving mode, measurement range and measurement precision of an atmosphere data system are improved, and adaptability to complex flight environment is also improved.

The invention provides a method and device for calculating rudder deviation required by unit normal overload of a statically unstable aircraft, and the method comprises the steps: selecting a plurality of calculation state points, and obtaining a small disturbance linearization equation at each calculation state point; a stability augmentation control law is constructed, the pitching moment characteristic of the aircraft is changed by feeding back the attack angle, the pitching damping characteristic of the aircraft is improved by feeding back the pitching angle rate, and therefore the aircraft is not divergent under the control of the stability augmentation control law and has the expected stability and damping characteristic; establishing a calculation model of rudder deflection required by unit normal overload based on a stability augmentation control law and a small disturbance linearization equation at a calculation state point, calculating normal overload by the calculation model through a pitch angle rate, an attack angle and a vacuum speed, and processing the calculation model to obtain a transfer function from a control channel elevator instruction to the normal overload; and calculating a normal overload value generated by unit elevator skewness based on the transfer function, wherein the reciprocal of the normal overload value is the rudder skewness required by unit normal overload.

This invention relates to the field of aviation safety monitoring technology, and discloses an aircraft overload prediction method, system, electronic equipment, and storage medium. The method includes: calculating air density based on flight environment parameters using the wet air state equation; calculating the true airspeed based on flight environment parameters and positioning parameters using vector synthesis; calculating the standard engine thrust based on EDR parameters; calculating the actual engine thrust using an environmental correction model; calculating lift and drag using an aerodynamic model; calculating the instantaneous overload value based on the balance relationship between the synthesized total normal force and gravity; predicting future overload sequences using a deep learning model based on an input sequence composed of historical EDR parameters, flight environment parameters, and instantaneous overload values; counting the number of times the instantaneous overload value exceeds the overload threshold within a set time window, calculating the overload frequency per unit time, and outputting a turbulence intensity index. This invention can improve the conversion accuracy, prediction efficiency, and reliability of aircraft overload prediction.

The application belongs to the technical field of aircraft flight load design, and particularly relates to a correction method of an aircraft aerodynamic roll moment, initialization of a calculation parameter, calculation of an aircraft true airspeed, equivalent speed and aircraft speed pressure based on the calculation parameter, calculation of a motion parameter based on a dynamic response of a maneuver simulation, calculation of a full-aircraft target aerodynamic roll moment based on the motion parameter, given of an unbalanced moment tolerance, solving of the unbalanced moment based on the full-aircraft target roll moment, comparison and analysis of a difference value between the unbalanced moment and the unbalanced moment tolerance, and end of the calculation if an absolute value of the unbalanced moment is less than the unbalanced moment tolerance. By adjusting a local angle of attack of each profile of a wing component in a span direction, a wing spanwise pressure distribution characteristic is reconstructed, an unbalanced aerodynamic roll moment of the aircraft caused by the pressure distribution characteristic is eliminated, a load balance purpose is achieved, and a design requirement in engineering is met.