The corresponding corrected method is proposed for the INS ( INS-Inertial Navigation System ) accumulated error of large transport aircraft. System errors contain aircraft position error, altitude error and speed erro...The corresponding corrected method is proposed for the INS ( INS-Inertial Navigation System ) accumulated error of large transport aircraft. System errors contain aircraft position error, altitude error and speed error,one is increasing the accuracy of hardw are; the other is development of low cost softw are algorithms. Because of improving hardw are is more difficult in my country at present, developing softw are algorithms is essential w ay,w hich have been validated in my types of airplane. The combined heuristic algorithms ( ABPNN,Advanced Back-propagation neural netw orks algorithm and LSM -least square method) are presented,w hich incorporates the effects of flight region and measured terrain height data by radar and barometer. Based on this algorithm,the appropriate match region w as gotten by recognition of fiducial digital map in real time online. In process of w ork,the minimum of position error as a cost function and the constraint conditions are gave,the flight positions are recognized in real time and continuously,least sum of square is calculated based on LSM ,in other w ords,the optimized result is obtained. The simulation case demonstrate that the method is very successful,the correct rate of recognition is more 90 percent. In w ords,the algorithm presented is economical,validation and effective.展开更多
Missiles may be damaged when aeroservoelastic problem occurs,which is caused by the interaction of structure flexibility and flight control system.Because of the limit of wind tunnel test condition,numerical methods a...Missiles may be damaged when aeroservoelastic problem occurs,which is caused by the interaction of structure flexibility and flight control system.Because of the limit of wind tunnel test condition,numerical methods are mostly used in previous aeroservoelastic studies.However,series of assumptions and simplification on structures,aerodynamics and flight control systems are unavoidably introduced,and various nonlinear factors are also ignored,therefore,they result in considerable errors.A novel method called aeroservoelasticity semi-physical simulation test is proposed in this paper,which takes the flexible missile with control system as the test object.Vibration signals at several locations of the missile are measured by accelerometers,then corresponding unsteady aerodynamics is computed based on the fact that airflow at high Mach is nearly quasi-steady,and finally unsteady aerodynamics is exerted simultaneously by shakers at certain locations of the missile.The aeroservoelasticity semi-physical simulation test system can be constructed after the control system is closed.Open loop transfer function test and closed loop stability test are carried out in sequence.The test principle and method proposed in this paper are verified by the concordance between the results of numerical simulation and experiment.展开更多
Time-domain approaches are presented for analysis of the dynamic response of aeroservoelastic systems to atmospheric gust excitations. The continuous and discrete gust inputs are defined in the time domain. The time-d...Time-domain approaches are presented for analysis of the dynamic response of aeroservoelastic systems to atmospheric gust excitations. The continuous and discrete gust inputs are defined in the time domain. The time-domain approach to continuous gust response uses a state-space formulation that requires the frequency-dependent aerodynamic coefficients to be approximated with the rational function of a Laplace variable. A hybrid method which combines the Fourier transform and time-domain approaches is used to calculate discrete gust response. The purpose of this approach is to obtain a time-domain state-space model without using rational function approximation of the gust columns. Three control schemes are designed for gust alleviation on an elastic aircraft, and three control surfaces are used: aileron, elevator and spoiler. The signals from the rate of pitch angle gyroscope or angle of attack sensor are sent to the elevator while the signals from accelerometers at the wing tip and center of gravity of the aircraft are sent to the aileron and spoiler, respectively. All the control laws are based on classical control theory. The results show that acceleration at the center of gravity of the aircraft and bending-moment at the wing-root section are mainly excited by rigid modes of the aircraft and the accelerations at the wing-tip are mainly excited by elastic modes of the aircraft. All the three control schemes can be used to alleviate the wing-root moments and the accelerations. The gust response can be alleviated using control scheme 3, in which the spoiler is used as a control surface, but the effects are not as good as those of control schemes 1 and 2.展开更多
文摘The corresponding corrected method is proposed for the INS ( INS-Inertial Navigation System ) accumulated error of large transport aircraft. System errors contain aircraft position error, altitude error and speed error,one is increasing the accuracy of hardw are; the other is development of low cost softw are algorithms. Because of improving hardw are is more difficult in my country at present, developing softw are algorithms is essential w ay,w hich have been validated in my types of airplane. The combined heuristic algorithms ( ABPNN,Advanced Back-propagation neural netw orks algorithm and LSM -least square method) are presented,w hich incorporates the effects of flight region and measured terrain height data by radar and barometer. Based on this algorithm,the appropriate match region w as gotten by recognition of fiducial digital map in real time online. In process of w ork,the minimum of position error as a cost function and the constraint conditions are gave,the flight positions are recognized in real time and continuously,least sum of square is calculated based on LSM ,in other w ords,the optimized result is obtained. The simulation case demonstrate that the method is very successful,the correct rate of recognition is more 90 percent. In w ords,the algorithm presented is economical,validation and effective.
基金supported by the National Natural Science Foundation of China (Grant Nos. 90716006,10902006)
文摘Missiles may be damaged when aeroservoelastic problem occurs,which is caused by the interaction of structure flexibility and flight control system.Because of the limit of wind tunnel test condition,numerical methods are mostly used in previous aeroservoelastic studies.However,series of assumptions and simplification on structures,aerodynamics and flight control systems are unavoidably introduced,and various nonlinear factors are also ignored,therefore,they result in considerable errors.A novel method called aeroservoelasticity semi-physical simulation test is proposed in this paper,which takes the flexible missile with control system as the test object.Vibration signals at several locations of the missile are measured by accelerometers,then corresponding unsteady aerodynamics is computed based on the fact that airflow at high Mach is nearly quasi-steady,and finally unsteady aerodynamics is exerted simultaneously by shakers at certain locations of the missile.The aeroservoelasticity semi-physical simulation test system can be constructed after the control system is closed.Open loop transfer function test and closed loop stability test are carried out in sequence.The test principle and method proposed in this paper are verified by the concordance between the results of numerical simulation and experiment.
文摘Time-domain approaches are presented for analysis of the dynamic response of aeroservoelastic systems to atmospheric gust excitations. The continuous and discrete gust inputs are defined in the time domain. The time-domain approach to continuous gust response uses a state-space formulation that requires the frequency-dependent aerodynamic coefficients to be approximated with the rational function of a Laplace variable. A hybrid method which combines the Fourier transform and time-domain approaches is used to calculate discrete gust response. The purpose of this approach is to obtain a time-domain state-space model without using rational function approximation of the gust columns. Three control schemes are designed for gust alleviation on an elastic aircraft, and three control surfaces are used: aileron, elevator and spoiler. The signals from the rate of pitch angle gyroscope or angle of attack sensor are sent to the elevator while the signals from accelerometers at the wing tip and center of gravity of the aircraft are sent to the aileron and spoiler, respectively. All the control laws are based on classical control theory. The results show that acceleration at the center of gravity of the aircraft and bending-moment at the wing-root section are mainly excited by rigid modes of the aircraft and the accelerations at the wing-tip are mainly excited by elastic modes of the aircraft. All the three control schemes can be used to alleviate the wing-root moments and the accelerations. The gust response can be alleviated using control scheme 3, in which the spoiler is used as a control surface, but the effects are not as good as those of control schemes 1 and 2.
