56 results about "Pressure altitude" patented technology

A cabin pressure altitude monitor and warning system provides a warning when a detected cabin pressure altitude has reached a predetermined level. The system is preferably embodied in a portable, pager-sized device that can be carried or worn by an individual. A microprocessor calculates the pressure altitude from signals generated by a calibrated pressure transducer and a temperature sensor that compensates for temperature variations in the signals generated by the pressure transducer. The microprocessor is programmed to generate a warning or alarm if a cabin pressure altitude exceeding a predetermined threshold is detected. Preferably, the microprocessor generates two different types of warning or alarm outputs, a first early warning or alert when a first pressure altitude is exceeded, and a second more serious alarm condition when either a second, higher pressure altitude is exceeded, or when the first pressure altitude has been exceeded for a predetermined period of time. Multiple types of alarm condition indicators are preferably provided, including visual, audible and tactile. The system is also preferably designed to detect gas concentrations and other ambient conditions, and thus incorporates other sensors, such as oxygen, relative humidity, carbon dioxide, carbon monoxide and ammonia sensors, to provide a more complete characterization and monitoring of the local environment.

The invention provides an aviation emergency instrument which comprises an atmospheric pressure measurement unit, an inertial measurement unit, a double-antenna GPS satellite receiver connected with a satellite signal receiving antenna, a graphic display unit, a secondary power supply circuit, and a comprehensive processing and calculation unit, wherein the comprehensive processing and calculation unit receives the information from the individual sensors, calculates the navigation parameters indicative of air speed, pressure altitude, posture, flight direction, longitude and latitude by use of a sensor information fusion algorithm, and sends the calculation result to the high-density liquid crystal graphic display unit. According to the invention, since the aviation emergency instrument can provide the navigation parameters indicative of air speed, pressure altitude, posture (pitch angle and roll angle), flight direction, longitude and latitude, the aviation emergency instrument can not only provide reference information to a pilot who operates the main flight instrument in the normal state but also serve as an emergency instrument in case that the main flight instrument is in the failure state. When the main flight instrument completely fails to work, the pilot can operate the aircraft to return safely by means of the navigation information provided by the emergency instrument.

The invention discloses an inertia altitude channel damping Kalman filtering method based on atmosphere assistance. The method comprises the following steps that firstly, an atmosphere data system pressure altitude error model in the moving process of an unmanned helicopter is established; then, based on the pressure altitude error model and a traditional inertia/atmosphere damping loop, a state equation and a measuring equation of a six-state damping Kalman filter including pressure altitude errors of an atmosphere data system and sensor errors of an inertia navigation system are established; finally, inertia/atmosphere navigation system damping Kalman filtering is carried out on the altitude error and the vertical speed error generated in the flying process of the unmanned helicopter, and a combined navigation system navigation result obtained after damping Kalman filtering is obtained. By means of the inertia altitude channel damping Kalman filtering method based on atmosphere assistance, tracking and compensation for atmosphere altitude measuring errors can be achieved in the flying process of the unmanned helicopter so as to improve the altitude precision in the atmosphere-assisted inertia navigation altitude channel, and the method is suitable for engineering application.

The invention provides an unmanned plane pressure altitude change rate measurement system. The system comprises an atmosphere static pressure transducer, a temperature transducer, a signal processing circuit, an A/D converter and a micro controller unit; an acquired atmospheric pressure signal and an acquired temperature value signal are transmitted to the signal processing circuit through the atmosphere static pressure transducer and the temperature transducer arranged on an unmanned plane, a signal processed by the signal processing circuit is transmitted to the A/D converter, and a signal through the AD conversion is sent to the micro controller unit; the uncorrected pressure altitude change rate is fast resolved by using linear interpolation, and is corrected by using a software compensating method so as to obtain the final pressure altitude change rate. The atmosphere static pressure transducer is used for measuring the atmosphere static pressure change rate to obtain the value of the pressure altitude change rate, the measurement speed of the system is increased, the measurement precision of the system and the accuracy of the unmanned plane pressure altitude change rate are obviously improved; meanwhile the security of the unmanned plane is improved.

A flight control system for a helicopter having a main rotor and a tail rotor is provided. The flight control system includes an automatic rotor speed control being configured to transition the main rotor and the tail rotor between a high speed and a low speed based upon a plurality of received information without requiring any pilot action, the plurality of received information comprising height above ground level, knot indicated air speed, outside air temperature, barometric altitude, pressure altitude, density altitude, and an engine operational status.

The invention discloses a multi-degree of freedom inertial sensor four-axis unmanned aerial vehicle autonomous navigation flight controller, which comprises a DSP arranged inside the unmanned aerial vehicle. A power supply module provides power for the DSP; the DSP is directly connected with the inertial sensor and an air pressure altitude sensor via a system bus to integrate the inertial sensor and the air pressure altitude sensor as a whole; the DSP is also connected with a wireless data transmission radio station and can communicate wirelessly with a ground station control system via the wireless data transmission radio station; the information output end of the DSP is connected with an electron speed regulator; and the electron speed regulator is connected with the motor of the unmanned aerial vehicle and can control the speed of the motor of the unmanned aerial vehicle. The multi-degree of freedom inertial sensor four-axis unmanned aerial vehicle autonomous navigation flight controller has the advantages that the processor is integrated with the inertial navigation sensor and an air data sensor and can perform data exchange; and due to the integrated design, the system-level serial communication is reduced, the inertial data updating rate is improved, the reliability of the system is enhanced, and the cost, the volume, the weight and the power consumption are reduced.

A system for identifying human motion of a subject may comprise: a sensor platform attached to the subject; an attitude filter that receives sensor data and calculates the attitude and heading of the platform; a feature construction and aggregation unit that receives the sensor data and the attitude and heading to calculate a feature vector for a human motion model, an evaluation unit that receives feature vectors and calculates the probability that an instance of a specific human motion occurred; wherein the feature vector comprises: one or more feature constructions, X, such as: forces along and perpendicular to the gravity field direction; a magnitude of rotations perpendicular to the gravity field direction; an adjusted pressure altitude; one or more feature constructions which are time derivatives dX/dt; and one or more feature constructions which are derivatives V(dX/dt), wherein: V(x)=2*H(x)−1, H(x)=+1 (x>0), 0 (x<0), and x is dX/dt.

The invention discloses a general aviation onboard monitoring system. The general aviation onboard monitoring system comprises a pressure altitude measurement module, a positioning module, a Beidou processing unit, a central processing unit and a battery module, wherein the pressure altitude measurement module is used for measuring the pressure altitude information of a general aircraft, and sending the pressure altitude information to the central processing unit, the positioning module is used for providing the positioning information of the general aircraft in real time, and sending the positioning information to the central processing unit, the Beidou processing unit is connected with a Beidou antenna, and is used for receiving and dispatching communication by utilizing Beidou satellites, the Beidou processing unit is used for processing received external command information, sending the processed command information to the central processing unit, and sending the processed positioning information obtained from the central processing unit by the Beidou antenna, the central processing unit is used for processing positioning information required for each module, and managing the running condition of whole onboard monitoring system, and the battery module is used for providing a power supply. With the adoption of the general aviation onboard monitoring system, the monitoring capability of the general aircraft can be effectively improved, and a feasible and effective monitoring means can be provided for a general aircraft owner and a low-altitude airspace manager.

The invention relates to a method for measuring the deviation rate of an airplane relative to a side direction of a runway, and belongs to the technical field of avionics display control. The method comprises the steps: firstly collecting the glide deviation of the airplane relative to a standard glide slope, the pressure altitude and ground velocity of the airplane relative to the runway, the length of the runway and a set glide slope angle, and solving the distance between the airplane and a heading beacon station according to the above collected information; secondly solving the acceleration of the airplane in a side direction of the runway according to a roll angle, a pitch angle, a yaw angle, a selected course, and the X-axis, Y-axis and Z-axis accelerations of an airplane body coordinate system; finally obtaining the deviation rate of the airplane relative to the side direction of the runway through employing a complementary filtering algorithm. According to the invention, the method inhibits the high-frequency noise in heading beacon data and the low-frequency noise in inertia data while making the comprehensive use of the heading beacon data and inertia data, is higher in precision of output data, and can support a head-up flight guide function.

The invention discloses a tumble detection system. The system comprises an air pressure sensor, a six-axis sensor, a photoelectric type pulse sensor, a main control unit (MCU), a server and a mobile terminal, wherein the air pressure sensor is used for acquiring air pressure altitude data of a sporter; the six-axis sensor is used for acquiring accelerated speed data and angular speed data generated by the sporter; the photoelectric type pulse sensor is used for acquiring user pulse data; the MCU is used for performing feature processing on the acquired air pressure altitude data, the accelerated speed data, the angular speed data and the user pulse data to acquire feature data, judging whether the feature data is tumble data or not, and generating tumble information if the feature data is the tumble data and transmitting the tumble information to the server through the GPRS module; the MCU comprises the GPRS module; the server is used for transmitting the tumble information to the mobile terminal; and the mobile terminal is used for receiving the tumble information transmitted by the server side and displaying the tumble information. The system can improve the detection accuracy.

The invention relates to a pressure altitude simulator and belongs to the technical field of structure design of pressure altitude simulators. A static port, a total port and plug-in holes are formed in the outer end of a simulator body. A pump assembly, a liquid crystal display module, a sensor assembly, a power source assembly, a volume assembly and a deflation valve assembly are arranged inside the simulator body. A fixing plate is further arranged inside the simulator body, and a several-piece assembly and a valve assembly are arranged on the fixing plate. Under the condition that the function and performance of equipment are not damaged, the functions of all units inside the pressure altitude simulator and the structure of the pressure altitude simulator are reasonably adjusted, so that the equipment is reduced to 4U at present from original 8U, and the miniaturization design of the pressure altitude simulator is achieved; meanwhile, due to the fact that functional modules of the simulator are integrated and the structure is simpler after adjustment, production and maintenance cost is correspondingly reduced, and the demands of users for miniaturization and low cost of test equipment are met.

A computer system 20 provided in an aircraft 1 includes, as a module of a computer program to be executed, an approach determination section 31 configured to determine whether the aircraft 1 is approaching a landing site. The approach determination section 31 includes, as a condition for determination of approaching, establishment (C1) of one or both of a condition (C11) that a first absolute altitude A1 obtained by an radio altimeter 21 mounted on the aircraft 1 is low relative to a first approach altitude AA1, and a condition (C12) that a second absolute altitude A2 that is obtained by subtracting an altitude A_L of the landing site from a pressure altitude Ap determined by a barometric altimeter 22 mounted on the aircraft 1 is low relative to a predetermined second approach altitude AA2.

The invention relates to a pressure altitude generation device. The pressure altitude generation device comprises a pressure altitude adjusting unit, a computation unit and a display unit. Compared with the existing altitude measurement system, an operator can acquire the QFE altitude without querying the corresponding technical manual; and any value of the standard pressure altitude (QNE altitude), the amended sea level pressure altitude (QNH altitude) and the airdrome pressure altitude (QFE) needed in the flight process of an aircraft can be acquired through simple operation.

The invention discloses a vehicle-mounted inertial navigation system. The vehicle-mounted inertial navigation system comprises a pressure altitude sensor, a milemeter, an acceleration sensor, an image acquisition module, a control module, a touch display module, an audio module and a communication module, wherein a plurality of interrupt response modules, an I2C bus interface module, a serial port communication interface module and a data input-output module are arranged inside the control module; the milemeter is connected with the interrupt response modules of the control module; the pressure altitude sensor is connected with the control module by virtue of the I2C bus interface module; the communication module is connected with the control module by virtue of the serial port communication interface module; and the acceleration sensor, the image acquisition module, the touch display module and the audio module are connected with the control module by virtue of the data input-output module. The vehicle-mounted inertial navigation system is capable of providing three-dimensional position, speed and track data and has the assistance of the characteristics of driver eye-closing caution, impact caution, border crossing caution, automatic cooperation of near interest points, and the like.

The invention relates to a system and method for correcting the reference value of pressure conversion altitude. The system is used in a navigational aid and comprises a storage unit, a judgment unit, a sensing unit and a computing unit, the storage unit is used for storing contour line geographic information; the judgment unit is used for judging whether the geographic position information positioned by the navigational aid is on the contour lines; the sensing unit is used for measuring the atmospheric pressure around the navigational aid and obtaining an atmospheric pressure value; and the computing unit is used for computing a reference value according to the contour line information and the atmospheric pressure value, and carrying out conversion computation among the contour line information, atmospheric pressure value and reference value. By using the system and method, a user does not need to mind the problem of setting of the reference value, and the GPS device can automatically correct the conversion reference value of the pressure altitude at any time and enable the altimeter to always keep high measurement accuracy, thereby saving trouble for the user and facilitating applications of the user.

The invention provides methods and systems for determining airspeed of an aircraft. The methods and systems allow for calculation of airspeed in near-ground and on-ground aircraft operation. A GPS altitude and a vertical acceleration of the aircraft are obtained for a current time frame. A geometric altitude for the previous time frame is determined, and the difference between the GPS altitude andgeometric altitude are combined with the vertical acceleration to calculate a geometric altitude rate of change. The geometric altitude rate of change is used to calculate a pressure altitude rate ofchange, which is used to calculate a pressure altitude for the aircraft. A static pressure is calculated from the pressure altitude, and the airspeed is calculated using the static pressure.