Reliable Dual-redundant Sensor Failure Detection and Identification for the NASA F-8 DFBW Aircraft |
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acceleration accelerometer addition aircraft alpha vane altimeter angle approximately assume attitude gyro average axes axis BFM Bias biases calculations choose chosen compensated computed crosses defined detection determine direct redundancy trigger discussed DR SPRT effect equations estimate examines expected failed instrument failure failure identification Figure filter flight gain given gives identification indicates initialization instrument type longitudinal m/secē Mach meter magnitude mean measurement negative noise normal accelerometer observed obtained outer loop output parameters Performance pitch rate gyro positive possible present QSPRT ramp rate gyro region remains residuals respectively RK residuals roll rate gyro sample seconds Segment Segment 4A sensor shear simulation SPRT SPRT's Table term translational kinematics turn error unfailed variable variance vector velocity vertical wind worst worst-case yaw rate gyro zero