Transactions of Nanjing University of Aeronautics & Astronautics
Volume 34,Issue 6,2017 Table of Contents
2017, 34(6):593-599.
DOI: 10.16356/j.1005-1120.2017.06.593
Abstract:
The control problem for underactuated reentry vehicle like HTV2 is considered with small angle of attack. The control strategy for an aircraft with positive lateral control departure parameter relies on strong lateral stability, which declines with the decrease of the angle of attack. Thus, to control the lateraldirectional motion in a stable state is hard and even impossible in some scenarios where the underactuated reentry vehicle, like HTV2, flies in a low angle of attack. To address this problem, the lateraldirectional openloop motion characteristics are analyzed. The results show that in an uncontrolled state, the lateraldirectional motion can automatically converge to stabilization thanks to the aerodynamic damping effect. Therefore, a method of turningoff the lateraldirectional control and inviting aerodynamic damping to control can achieve stability. The sixdegreeoffreedom simulation show that the lateraldirectional motion can be stabilized by the aerodynamic damping, and the lateral position error caused by the bank angle deviation is limited near the zerorise angle of attack. The control strategy is effective.
The control problem for underactuated reentry vehicle like HTV2 is considered with small angle of attack. The control strategy for an aircraft with positive lateral control departure parameter relies on strong lateral stability, which declines with the decrease of the angle of attack. Thus, to control the lateraldirectional motion in a stable state is hard and even impossible in some scenarios where the underactuated reentry vehicle, like HTV2, flies in a low angle of attack. To address this problem, the lateraldirectional openloop motion characteristics are analyzed. The results show that in an uncontrolled state, the lateraldirectional motion can automatically converge to stabilization thanks to the aerodynamic damping effect. Therefore, a method of turningoff the lateraldirectional control and inviting aerodynamic damping to control can achieve stability. The sixdegreeoffreedom simulation show that the lateraldirectional motion can be stabilized by the aerodynamic damping, and the lateral position error caused by the bank angle deviation is limited near the zerorise angle of attack. The control strategy is effective.
2017, 34(6):600-608.
DOI: 10.16356/j.1005-1120.2017.06.600
Abstract:
We summarize the guidance and control techniques of automatic carrier landing for carrierbased aircraft. First, we analyze the carrier landing operations of the manned fixed-wing aircraft, unmanned fixed-wing aircraft and helicopters. Second, we look into the navigation and guidance system and the flight control methods for current different aircraft. Finally, we draw several conclusions of the development prospects for aircraft carrier landing, including the precision landing control techniques, precision approach and landing guidance techniques, and adaptive, reconfigurable and intelligent flight control techniques.
We summarize the guidance and control techniques of automatic carrier landing for carrierbased aircraft. First, we analyze the carrier landing operations of the manned fixed-wing aircraft, unmanned fixed-wing aircraft and helicopters. Second, we look into the navigation and guidance system and the flight control methods for current different aircraft. Finally, we draw several conclusions of the development prospects for aircraft carrier landing, including the precision landing control techniques, precision approach and landing guidance techniques, and adaptive, reconfigurable and intelligent flight control techniques.
2017, 34(6):609-616.
DOI: 10.16356/j.1005-1120.2017.06.609
Abstract:
We studied carrier landing robust control based on longitudinal decoupling. Firstly, due to the relative strong coupling between the tangential and the normal directions, the height and the velocity channels were decoupled by using the exact linearization method, so that controllers for the two channels could be designed seperately. In the height control, recursive dynamic surface was used to accelerate the convergence of the height control and eliminate ″the explosion of complexity″. The radial basis fuction(RBF) neural network was designed by using the minimum learning parameter method to compensate the uncertainty. A kind of surface with nonsingular fast terminal sliding mode and its reaching law were developed to ensure finite time convergence and to avoid singularity. The controller for the velocity was designed by using supertwisting secondorder sliding mode control. The stability of the proposed system was validated by Lyapunov method. The results showed that the Levant′s robust differential observer was improved and used for the observation of the required higher order differential of signals in the controller. The response of aircraft carrier landing under the complex disturbance is simulated and the results verified the approach.
We studied carrier landing robust control based on longitudinal decoupling. Firstly, due to the relative strong coupling between the tangential and the normal directions, the height and the velocity channels were decoupled by using the exact linearization method, so that controllers for the two channels could be designed seperately. In the height control, recursive dynamic surface was used to accelerate the convergence of the height control and eliminate ″the explosion of complexity″. The radial basis fuction(RBF) neural network was designed by using the minimum learning parameter method to compensate the uncertainty. A kind of surface with nonsingular fast terminal sliding mode and its reaching law were developed to ensure finite time convergence and to avoid singularity. The controller for the velocity was designed by using supertwisting secondorder sliding mode control. The stability of the proposed system was validated by Lyapunov method. The results showed that the Levant′s robust differential observer was improved and used for the observation of the required higher order differential of signals in the controller. The response of aircraft carrier landing under the complex disturbance is simulated and the results verified the approach.
2017, 34(6):617-626.
DOI: 10.16356/j.1005-1120.2017.06.617
Abstract:
Multiple unmanned aerial vehicles (UAVs) cooperative operation is the main form for UAVs fighting in battlefield, and multi-UAV mission rendezvous is the premise of cooperative reconnaissance and attack missions. We propose a rendezvous control strategy, which divides the rendezvous process into two parts: The loose formation rendezvous and the close formation rendezvous. In the first stage, UAVs are supposed to reach the specific target locations simultaneously and form a loose formation. A distributed control strategy based on first-order consensus algorithm is presented to achieve this goal. Then the second stage is designed based on the second-order consensus algorithm to complete the transition from the loose formation to the close formation.This process needs the speeds and heading angles of UAVs to reach an agreement. Besides, control algorithms with a virtual leader are proposed, by which the formation states can reach a specific value. Fi nally, simulation results show that the control algorithms are capable of realizing the mission rendezvous of multi-UAV and the consistence of UAVs′ final states, which verify the effectiveness and feasibility of the designed control strategy.
Multiple unmanned aerial vehicles (UAVs) cooperative operation is the main form for UAVs fighting in battlefield, and multi-UAV mission rendezvous is the premise of cooperative reconnaissance and attack missions. We propose a rendezvous control strategy, which divides the rendezvous process into two parts: The loose formation rendezvous and the close formation rendezvous. In the first stage, UAVs are supposed to reach the specific target locations simultaneously and form a loose formation. A distributed control strategy based on first-order consensus algorithm is presented to achieve this goal. Then the second stage is designed based on the second-order consensus algorithm to complete the transition from the loose formation to the close formation.This process needs the speeds and heading angles of UAVs to reach an agreement. Besides, control algorithms with a virtual leader are proposed, by which the formation states can reach a specific value. Fi nally, simulation results show that the control algorithms are capable of realizing the mission rendezvous of multi-UAV and the consistence of UAVs′ final states, which verify the effectiveness and feasibility of the designed control strategy.
2017, 34(6):627-636.
DOI: 10.16356/j.1005-1120.2017.06.627
Abstract:
The mathematical model of quadcopterunmanned aerial vehicle (UAV) is derived by using two approaches: One is the NewtonEuler approach which is formulated using classical mechanics; and other is the EulerLagrange approach which describes the model in terms of kinetic (translational and rotational) and potential energy. The proposed quadcopter′s non-linear model is incorporated with aero-dynamical forces generated by air resistance, which helps aircraft to exhibits more realistic behavior while hovering. Based on the obtained model, the suitable control strategy is developed, under which two effective flight control systems are developed. Each control system is created by cascading the proportional-derivative(PD) and T-S fuzzy controllers that are equipped with six and twelve feedback signals individually respectively to ensure better tracking, stabilization, and response. Both proposed flight control designs are then implemented with the quadcopter model respectively and multitudinous simulations are conducted using MATLAB/Simulink to analyze the tracking performance of the quadcopter model at various reference inputs and trajectories.
The mathematical model of quadcopterunmanned aerial vehicle (UAV) is derived by using two approaches: One is the NewtonEuler approach which is formulated using classical mechanics; and other is the EulerLagrange approach which describes the model in terms of kinetic (translational and rotational) and potential energy. The proposed quadcopter′s non-linear model is incorporated with aero-dynamical forces generated by air resistance, which helps aircraft to exhibits more realistic behavior while hovering. Based on the obtained model, the suitable control strategy is developed, under which two effective flight control systems are developed. Each control system is created by cascading the proportional-derivative(PD) and T-S fuzzy controllers that are equipped with six and twelve feedback signals individually respectively to ensure better tracking, stabilization, and response. Both proposed flight control designs are then implemented with the quadcopter model respectively and multitudinous simulations are conducted using MATLAB/Simulink to analyze the tracking performance of the quadcopter model at various reference inputs and trajectories.
2017, 34(6):637-646.
DOI: 10.16356/j.1005-1120.2017.06.637
Abstract:
Disturbance rejection algorithm based on model reference adaptive control (MRAC) augmentation is investigated for uncertain turbulence disturbances. A stable adaptive control scheme is developed based on lower diagonal upper (LDU) decomposition of the high frequency gain matrix, which ensures closedloop stability and asymptotic output tracking. Under the proposed control techniques, the bounded stability is achieved and the controller is able to remain within tight bounds on the matched and unmatched uncertainties. Finally, simulation studies of a linearized lateraldirectional dynamics model are conducted to demonstrate the performance of the adaptive scheme.
Disturbance rejection algorithm based on model reference adaptive control (MRAC) augmentation is investigated for uncertain turbulence disturbances. A stable adaptive control scheme is developed based on lower diagonal upper (LDU) decomposition of the high frequency gain matrix, which ensures closedloop stability and asymptotic output tracking. Under the proposed control techniques, the bounded stability is achieved and the controller is able to remain within tight bounds on the matched and unmatched uncertainties. Finally, simulation studies of a linearized lateraldirectional dynamics model are conducted to demonstrate the performance of the adaptive scheme.
2017, 34(6):647-658.
DOI: 10.16356/j.1005-1120.2017.06.647
Abstract:
Since the aerodynamic center moving backward sharply in hypersonic flight, the stability margin of the hypersonic vehicle increases largely while the maneuverability decreases. We proposed a novel method to solve this contradiction. We used relaxed static stability (RSS) to improve the maneuverability in hypersonic flight, and designed the stability augmentation system (SAS) to ensure the stability in subsonic flight. Therefore, the relationship between static stability and maneuverability was quantitatively analyzed in the first step, and the numerical value of RSS was obtained on the premise of good maneuverability. Secondly, the relationship between static stability and aerodynamic parameters was quantitatively analyzed. We properly adjusted aerodynamic parameters based on the quantitative relationship to achieve the specific static stability set in the first step, and therefore provided the engineering realization methods. The vehicle will be statically unstable in subsonic flight with the specific static stability. Lastly, SAS was needed to ensure the stability of the vehicle in subsonic flight. Simulation studies were conducted by comparing the linear SAS to the nonlinear SAS, and the results showed that the nonlinear dynamicinversion controller can synthesize with proportionalintegrallderivative(PID) controller robustly and stabilize the hypersonic vehicle.
Since the aerodynamic center moving backward sharply in hypersonic flight, the stability margin of the hypersonic vehicle increases largely while the maneuverability decreases. We proposed a novel method to solve this contradiction. We used relaxed static stability (RSS) to improve the maneuverability in hypersonic flight, and designed the stability augmentation system (SAS) to ensure the stability in subsonic flight. Therefore, the relationship between static stability and maneuverability was quantitatively analyzed in the first step, and the numerical value of RSS was obtained on the premise of good maneuverability. Secondly, the relationship between static stability and aerodynamic parameters was quantitatively analyzed. We properly adjusted aerodynamic parameters based on the quantitative relationship to achieve the specific static stability set in the first step, and therefore provided the engineering realization methods. The vehicle will be statically unstable in subsonic flight with the specific static stability. Lastly, SAS was needed to ensure the stability of the vehicle in subsonic flight. Simulation studies were conducted by comparing the linear SAS to the nonlinear SAS, and the results showed that the nonlinear dynamicinversion controller can synthesize with proportionalintegrallderivative(PID) controller robustly and stabilize the hypersonic vehicle.
2017, 34(6):659-668.
DOI: 10.16356/j.1005-1120.2017.06.659
Abstract:
Carrierbased aircraft carrier landing is a special kind of tracking control problem and not suitable for classical control methods, which may miss the desired performance or result in overdesign. Therefore, we present an optimal preview control for automatic carrier landing system (ACLS) by using state information of system, as well as future reference information, which can avoid the shortcomings of classical control methods. Since the flight performance of carrierbased aircraft is disturbed by air wake when the aircraft flies near the area of carrier stern, we design a disturbance rejection strategy to ensure that aircraft track the glide path with high precision and robustness. Further, carrier-based aircraft is a complex nonlinear system. However, the nonlinear model of carrier-based aircraft can be linearized at equilibrium landing state and decoupled into the longitudinal model and the lateral model. Therefore, an optimal preview control system is designed. The simulation results of a carrierbased aircraft show that the optimal preview control system can effectively suppress air wake. Tracking accuracy of optimal preview controller is higher than that of the proportional integral differential (PID) control system.
Carrierbased aircraft carrier landing is a special kind of tracking control problem and not suitable for classical control methods, which may miss the desired performance or result in overdesign. Therefore, we present an optimal preview control for automatic carrier landing system (ACLS) by using state information of system, as well as future reference information, which can avoid the shortcomings of classical control methods. Since the flight performance of carrierbased aircraft is disturbed by air wake when the aircraft flies near the area of carrier stern, we design a disturbance rejection strategy to ensure that aircraft track the glide path with high precision and robustness. Further, carrier-based aircraft is a complex nonlinear system. However, the nonlinear model of carrier-based aircraft can be linearized at equilibrium landing state and decoupled into the longitudinal model and the lateral model. Therefore, an optimal preview control system is designed. The simulation results of a carrierbased aircraft show that the optimal preview control system can effectively suppress air wake. Tracking accuracy of optimal preview controller is higher than that of the proportional integral differential (PID) control system.
2017, 34(6):669-679.
DOI: 10.16356/j.1005-1120.2017.06.669
Abstract:
In order to achieve the goal that unmanned aerial vehicle(UAV) automatically positioning during power inspection, a visual positioning method which utilizes encoded sign as cooperative target is proposed. Firstly, we discuss how to design the encoded sign and propose a robust decoding algorithm based on contour. Secondly, the Adaboost algorithm is used to train a classifier which can detect the encoded sign from image. Lastly, the position of UAV can be calculated by using the projective relation between the object points and their corresponding image points. Experiment includes two parts. First, simulated video data is used to verify the feasibility of the proposed method, and the results show that the average absolute error in each direction is below 0.02 m. Second, a video, acquired from an actual UAV flight, is used to calculate the position of UAV. The results show that the calculated trajectory is consistent with the actual flight path. The method runs at a speed of 0.153 s per frame.
In order to achieve the goal that unmanned aerial vehicle(UAV) automatically positioning during power inspection, a visual positioning method which utilizes encoded sign as cooperative target is proposed. Firstly, we discuss how to design the encoded sign and propose a robust decoding algorithm based on contour. Secondly, the Adaboost algorithm is used to train a classifier which can detect the encoded sign from image. Lastly, the position of UAV can be calculated by using the projective relation between the object points and their corresponding image points. Experiment includes two parts. First, simulated video data is used to verify the feasibility of the proposed method, and the results show that the average absolute error in each direction is below 0.02 m. Second, a video, acquired from an actual UAV flight, is used to calculate the position of UAV. The results show that the calculated trajectory is consistent with the actual flight path. The method runs at a speed of 0.153 s per frame.
2017, 34(6):680-687.
DOI: 10.16356/j.1005-1120.2017.06.680
Abstract:
A gradient descent algorithm with adjustable parameter for attitude estimation is developed, aiming at the attitude measurement for small unmanned aerial vehicle (UAV) in real -time flight conditions. The accelerometer and magnetometer are introduced to construct an error equation with the gyros, thus the drifting characteristics of gyroscope can be compensated by solving the error equation utilized by the gradient descent algorithm. Performance of the presented algorithm is evaluated using a self-proposed micro-electro-mechanical system (MEMS) based attitude heading reference system which is mounted on a tri-axis turntable. The on-ground, turntable and flight experiments indicate that the estimation attitude has a good accuracy. Also, the presented system is compared with an open-source flight control system which runs extended Kalman filter (EKF), and the results show that the attitude control system using the gradient descent method can estimate the attitudes for UAV effectively.
A gradient descent algorithm with adjustable parameter for attitude estimation is developed, aiming at the attitude measurement for small unmanned aerial vehicle (UAV) in real -time flight conditions. The accelerometer and magnetometer are introduced to construct an error equation with the gyros, thus the drifting characteristics of gyroscope can be compensated by solving the error equation utilized by the gradient descent algorithm. Performance of the presented algorithm is evaluated using a self-proposed micro-electro-mechanical system (MEMS) based attitude heading reference system which is mounted on a tri-axis turntable. The on-ground, turntable and flight experiments indicate that the estimation attitude has a good accuracy. Also, the presented system is compared with an open-source flight control system which runs extended Kalman filter (EKF), and the results show that the attitude control system using the gradient descent method can estimate the attitudes for UAV effectively.
2017, 34(6):688-699.
DOI: 10.16356/j.1005-1120.2017.06.688
Abstract:
A coalition formation algorithm is presented with limited communication ranges and delays in unknown environment, for the performance of multiple heterogeneous unmanned aerial vehicles (UAVs) in cooperative search and attack missions. The mathematic model of coalition formation is built on basis of the minimum attacking time and the minimum coalition size with satisfying resources and simultaneous strikes requirements. A communication protocol based on maximum number of hops is developed to determine the potential coalition members in dynamic network. A multistage sub-optimal coalition formation algorithm (MSOCFA) with polynomial time is established. The performances of MSOCFA and particle swarm optimization (PSO) algorithms are compared in terms of complexity, mission performance and computational time. A complex scenario is deployed to illustrate how the coalitions are formed and validate the feasibility of the MSOCFA. The effect of communication constraints (hop delay and max-hops) on mission performance is studied. The results show that it is beneficial to determine potential coalition members in a wide and deep range over the network in the presenc e of less delay. However, when the delays are significant, it is more advantageous to determine coalitions from among the immediate neighbors.
A coalition formation algorithm is presented with limited communication ranges and delays in unknown environment, for the performance of multiple heterogeneous unmanned aerial vehicles (UAVs) in cooperative search and attack missions. The mathematic model of coalition formation is built on basis of the minimum attacking time and the minimum coalition size with satisfying resources and simultaneous strikes requirements. A communication protocol based on maximum number of hops is developed to determine the potential coalition members in dynamic network. A multistage sub-optimal coalition formation algorithm (MSOCFA) with polynomial time is established. The performances of MSOCFA and particle swarm optimization (PSO) algorithms are compared in terms of complexity, mission performance and computational time. A complex scenario is deployed to illustrate how the coalitions are formed and validate the feasibility of the MSOCFA. The effect of communication constraints (hop delay and max-hops) on mission performance is studied. The results show that it is beneficial to determine potential coalition members in a wide and deep range over the network in the presenc e of less delay. However, when the delays are significant, it is more advantageous to determine coalitions from among the immediate neighbors.
2017, 34(6):700-709.
DOI: 10.16356/j.1005-1120.2017.06.700
Abstract:
CubeSats have evolved from purely educational tools to a standard platform for technology demonstration, scientific instrumentation and application in less than a decade. They open the door to new challenges and interplanetary missions which lead to the direct realization of autonomous orbit determination (AOD) which has been investigated before with different integrated sensors combined with various filters. Mostly these studies were carried out for larger satellites with more accurate sensors. Magnetometer and sun sensor combined with extended Kalman filter (EKF) are chosen to complete AOD task considering their light weight. For the purpose of AOD and the computational cost requirements imposed on CubeSats, it is important to develop and apply low cost onboard models. In this perspective, a magnetic model based on a table look up is proposed to generate the reference magnetic field with a low computational burden. In current article the simulations through Matlab and Satellites Tool Kit (STK) especially focus on the accuracy of the AOD system provided by this model. For analysis three EKFs are carried out with different calculation models and data types. The system based on the proposed model is fully autonomous, low-cost and has moderateaccuracy required by most CubeSats missions. The AOD system can be applied as main or backup system depending on the space missions′ demands.
CubeSats have evolved from purely educational tools to a standard platform for technology demonstration, scientific instrumentation and application in less than a decade. They open the door to new challenges and interplanetary missions which lead to the direct realization of autonomous orbit determination (AOD) which has been investigated before with different integrated sensors combined with various filters. Mostly these studies were carried out for larger satellites with more accurate sensors. Magnetometer and sun sensor combined with extended Kalman filter (EKF) are chosen to complete AOD task considering their light weight. For the purpose of AOD and the computational cost requirements imposed on CubeSats, it is important to develop and apply low cost onboard models. In this perspective, a magnetic model based on a table look up is proposed to generate the reference magnetic field with a low computational burden. In current article the simulations through Matlab and Satellites Tool Kit (STK) especially focus on the accuracy of the AOD system provided by this model. For analysis three EKFs are carried out with different calculation models and data types. The system based on the proposed model is fully autonomous, low-cost and has moderateaccuracy required by most CubeSats missions. The AOD system can be applied as main or backup system depending on the space missions′ demands.
2017, 34(6):710-719.
DOI: 10.16356/j.1005-1120.2017.06.710
Abstract:
Aiming at the noise of helicopter scissors tailrotor, an advanced numerical method is established by combining computational fluid dynamics (CFD) model with Farassat 1A (F1A) formula. In this method, Naviertokes (N-S) equations are used as governing equations, and the flow field is solved at quasi-steady and unsteady states in hover and forward fight, respectively, based on two different types of embedded grid systems. A simple and effective solution approach is provided for the generation difficulty of donor cells caused by the close gap among scissors tail-rotor blades. Using the CFD calculation results as input, the thickness noise, loading noise and total noise of tail-rotor are calculated by F1A formula. By the method, numerical examples on rotor flowfield and noise are performed and the results are compared with available data. Then, aerodynamic and acoustic characteristics of scissors tail-rotor are emphatically calculated in both hover and forward flight. Furthermore, the research on the effects of blade-tip shape parameters on scissors tailrotor noise is carried out. Also, the scissors tailrotor is compared with the conventional tailrotor, and the results show that in hover, the noise of a scissors tailrotor is not always the smaller one.
Aiming at the noise of helicopter scissors tailrotor, an advanced numerical method is established by combining computational fluid dynamics (CFD) model with Farassat 1A (F1A) formula. In this method, Naviertokes (N-S) equations are used as governing equations, and the flow field is solved at quasi-steady and unsteady states in hover and forward fight, respectively, based on two different types of embedded grid systems. A simple and effective solution approach is provided for the generation difficulty of donor cells caused by the close gap among scissors tail-rotor blades. Using the CFD calculation results as input, the thickness noise, loading noise and total noise of tail-rotor are calculated by F1A formula. By the method, numerical examples on rotor flowfield and noise are performed and the results are compared with available data. Then, aerodynamic and acoustic characteristics of scissors tail-rotor are emphatically calculated in both hover and forward flight. Furthermore, the research on the effects of blade-tip shape parameters on scissors tailrotor noise is carried out. Also, the scissors tailrotor is compared with the conventional tailrotor, and the results show that in hover, the noise of a scissors tailrotor is not always the smaller one.
2017, 34(6):720-732.
DOI: 10.16356/j.1005-1120.2017.06.720
Abstract:
To understand the effect of attitude angle on helicopter water ditching, a finite element model (FEM) of helicopter water ditching was established via the nonlinear finite element method. Based on the LS-DYNA software, the finite element model of the helicopter was established. Different attitude of helicopter water ditching was simulated and analysed. Three pressure measuring points were arranged in the aircraft belly area. The pressure curves were provided seperately. Finally, the effects of attitude on helicopter water ditching was discussed, particularly the stability and the overload of helicopter. Through the contours of stress, it can be figured out that the maximum pressure locates at the position of aircraft belly where stiffness is larger. Then simulation and experimental results were compared. The difference of peak acceleration along the Z direction was small. Therefore, this method is feasible and helpful for the design of helicopter water ditching.
To understand the effect of attitude angle on helicopter water ditching, a finite element model (FEM) of helicopter water ditching was established via the nonlinear finite element method. Based on the LS-DYNA software, the finite element model of the helicopter was established. Different attitude of helicopter water ditching was simulated and analysed. Three pressure measuring points were arranged in the aircraft belly area. The pressure curves were provided seperately. Finally, the effects of attitude on helicopter water ditching was discussed, particularly the stability and the overload of helicopter. Through the contours of stress, it can be figured out that the maximum pressure locates at the position of aircraft belly where stiffness is larger. Then simulation and experimental results were compared. The difference of peak acceleration along the Z direction was small. Therefore, this method is feasible and helpful for the design of helicopter water ditching.
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