Record Details

Autonomous Landing Of Unmanned Aerial Vehicles

Electronic Theses of Indian Institute of Science

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Title Autonomous Landing Of Unmanned Aerial Vehicles
 
Creator Singh, Shashiprakash
 
Subject Unmanned Space Craft
Spacecraft - Landing
Aerial Vehicles
Unmanned Aerial Vehicles
Aerial Vehicles - Guidance And Control
Aerial Vehicles - State Estimation
Aerial Vehicles - Mathematical Modeling
AE-2 (All Electric Airplane-2)
UAVs
Astronautics
 
Description In this thesis the problem of autonomous landing of an unmanned aerial vehicle named AE-2 is addressed. The guidance and control technique is developed and demonstrated through numerical simulation results. The complete work includes Mathematical modeling, Control design, Guidance and State estimation for AE-2, which is a fixed wing vehicle with 2m wing span and 6kg weight.
The aerodynamic data for AE-2 is available from static wind tunnel tests. Functional fit is done on the wind tunnel data with least squares method to find static aerodynamic coefficients. The aerodynamic forces and moment coefficients are highly nonlinear some of them are partitioned in two zones based on the angle of attack. The dynamic derivatives are found with Athena Vortex Lattice software. For the validation of vortex lattice method the static derivatives obtained by the wind tunnel tests and vortex lattice method, are compared before finding dynamic derivatives. The dynamics of the servo actuators for the aerodynamic control surfaces is incorporated in the simulation.
The nonlinear dynamic inversion technique has been used for the guidance and control design. The control is structured in two loops, outer and inner loop. The goal of outer loop is to track the guidance commands of altitude, roll angle and yaw angle by converting them into body rate commands through dynamic inversion. The inner loop than tracks these commanded roll rate, pitch rate and yaw rate by finding the required deflection of control surfaces. The forward velocity of the vehicle is controlled by varying the throttle. A controller for actuator is also designed to reduce the lag.
The guidance for landing consists of three phases approach, glideslope and flare. During approach the vehicle is aligned with the runway and guided to a specified height from where the glideslope can begin. The glideslope is straight line path specified by a flight path angle which is restricted between 3 to 4 degree. At the end of glideslope which is marked by flare altitude the flare maneuver begins which is an exponential curve. The problem of transition between the glideslope and flare has addressed by ensuring continuity and smoothness at transition. The exponential curve of flare is designed to end below the ground so that it intersects the ground at a prespecified point. The sink rate at touchdown is also controlled along with the location of touchdown point.
The state estimation has been done with Extended Kalman Filter in continuous discrete formulation. The external disturbances like wind shear and wind gust are accounted by appending them in state variables. Further the control design with guidance is tested from various initial conditions, in presence of wind disturbances. The designed filter has also been tested for parameter uncertainty.
 
Contributor Padhi, Radhakant
 
Date 2010-03-26T11:06:23Z
2010-03-26T11:06:23Z
2010-03-26T11:06:23Z
2009-02
 
Type Thesis
 
Identifier http://etd.iisc.ernet.in/handle/2005/662
 
Language en_US
 
Relation G22963