Design of an Unmanned Aerial Vehicle for Ecological Conservation

50th AIAA Aerospace Sciences Meeting Including the New Horizons Forum and Aerospace Exposition

Celi

2012

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The objective of the present work is to design the control systems required to hold altitude and heading in severe atmospheric disturbances in cruise flight for the Unmanned Airplane for Ecological Conservation using modern control design techniques. The airplane mathematical model in open-loop was defined by the linearized longitudinal and lateraldirectional equations of motion. Different control systems were designed based on modern control design techniques. These are the eigenstructure assignment or polo-placement technique, linear quadratic design with full state feedback, and linear quadratic Gaussian design. To verify the design of control systems, simulations of the open-loop and closed-loop systems were performed, and each control system was tested. The atmospheric disturbances considered were the gust disturbance, which was idealized by the one-minus-cosine gust profile, and the atmospheric turbulence, which was modeled by Dryden's function. The control systems designed to hold in-flight altitude and heading achieved their purpose. In any simulation, the maximum load factor, the stall speed and the maximum dive speed were not achieved. Nomenclature A, B, C = state, control, and gust disturbance matrices a = coefficients of open-loop plant matrix characteristic equation a = coefficients of desired closed-loop characteristic equation d = gust diameter g = constant vector H = flight altitude K= feedback gain L k = steady-state Kalman gain L T = correlation length of turbulence n y , n z = load factors in y and z axes Q, R = weighting matrices for linear quadratic regulator problem Q n , R n = weighting matrices in the algebraic filter Riccati equation P = Riccati matrix p, q, r = roll, pitch, and yaw rates p g , q g , r g = rotary gust velocities about x, y, and z axes t = time u, v, w = longitudinal, lateral and vertical components of velocity u e , v e , w e = longitudinal, lateral and vertical components of gust velocity respect to earth axes u g , v g , w g = longitudinal, lateral and vertical components of gust velocity respect to wind airframe axes u o = initial forward speed v go , w go = maximum lateral and vertical gust velocities w A , w B = weighting factors

Section: Introductionmentioning

confidence: 99%

Design of Control Systems to Hold Altitude and Heading in Severe Atmospheric Disturbances for an Unmanned Airplane

50th AIAA Aerospace Sciences Meeting Including the New Horizons Forum and Aerospace Exposition

Celi

2012

Self Cite

“…The idea is to produce as little structural modifications as possible, because the wing structural design describes a secondary spar located at 80% of the chord. 2 Successively, to estimate the value of the deflection, the NACA 4415 airfoil was tested with flap up, which occupied 20% of the chord at different angles of deflection. This was performed until one of the angles of deflection generated the same lift of the airfoil at optimum twist angle without changing the angle of attack.…”

Section: A Aerodynamic Analysis Methods Verificationmentioning

confidence: 99%

“…11 The induced drag coefficient is equal to the product of the square lift coefficient and the lift dependent drag factor or induced drag factor. Equation (2) shows that a lift dependent drag factor is linked to the maximum lift-drag ratio. 8 …”

Section: Drag Induced Drag and Maximum Lift-drag Ratiomentioning

confidence: 99%

Induced Drag Reduction of an Airplane Using Local Twist, Panel Method Verification

50th AIAA Aerospace Sciences Meeting Including the New Horizons Forum and Aerospace Exposition

Amerio

2012

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The objective of this work is to modify the wing of the Unmanned Airplane of Ecological Conservation adding wing twist in order to reduce induced drag at a design lift coefficient using a method based on the lifting-line theory, and test such modification using two different three-dimensional panel method codes. The aircraft is a small airplane designed to make daily patrols over petroleum extraction zones. A linear washout distribution between the boom and the wing tip was chosen for the ANCE wing. The optimum twist angle at a design lift coefficient was calculated, resulting in 5.4119 deg. The resultant induced drag reduction for the wing at design lift coefficient is equal to 6.425%. The panel method codes PAN AIR and CMARC were used to study the inviscid aerodynamic flowfield around the aircraft. The results show that the addition of wing twist increases the lift-drag ratio 2.03% and 1.53% respect to the original, when it is computed using PAN AIR and CMARC, respectively, assuming that the wing twist does not change the viscous drag. The lift coefficient at zero angle of attack was reduced by wing twist addition and the lift slope is independent of twist. Nomenclature a m = planform contribution to the coefficients in the infinite series solution to the lifting-line equation b = wingspan b m = washout contribution to the coefficients in the infinite series solution to the lifting-line equation= lift coefficient at zero angle of attack C Lα = lift slope C lα = airfoil section lift slope c = mean aerodynamic chord c c = airfoil section chord e = Oswald efficiency factor e n = two-dimensional free transition criterion k = lift dependent drag factor or induced drag factor, k = (π·e·R A ) -1 (L/D) max = maximum lift-drag ratio n = fixed parameter for two-dimensional free transition criterion R A = wing aspect ratio y to = distance between wing root and the twist start line y, z = spanwise and height coordinates α = angle of attack α 0 = airfoil section zero-lift angle of attack ε Ω = 2 θ = change of variables for the spanwise coordinate κ D = planform contribution to the induced drag factor κ DL = lift-washout contribution to the induced drag factor κ Do = optimum induced drag factor κ DΩ = washout contribution to the induced drag factor Ω = twist angle Ω opt = optimum twist angle ω = normalized washout distribution function

“…3 This is a complex reticular type structure. 2 The wing has a main spar, a rear spar and six stringers joined with sixteen ribs, eight in each semi-span. The horizontal tail structure is composed by a main spar, a rear spar and five ribs.…”

Section: Introductionmentioning

confidence: 99%

Static-Stability Analysis of an Unmanned Airplane as a Flexible-Body

González

AIAA Atmospheric Flight Mechanics Conference

et al. 2010

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The aerodynamic and static stability analysis of the Unmanned Airplane for Ecological Conservation as a flexible-body is the aim of present paper. The aerodynamic and structural analyses were performed using the ASWING code. This is a program for the aerodynamic, structural, and control-response analysis of airplanes with flexible components. The aerodynamic data obtained were compared against data obtained by an analytical-empirical method and potential flow codes. A good correlation is achieved for the longitudinal stability coefficients, but for lateral and directional coefficients there is a poor correlation. The distributed aerodynamic loads on surfaces (wing and horizontal tail) produce variations in the wing beam, changing the elastic twist distribution and reducing the lift-slope respect to the rigid-body analysis. In general, the comparison of flexible-body data to rigid-body results shows slight differences. Nomenclature C Do= viscous drag coefficient or parasite drag coefficient C Lo , C Mo = lift and pitching moment coefficient at zero angle of attack C Lq , C Mq = variation of lift, and pitching moment coefficients with pitch rate C Lu , C Du , C Mu = variation of lift, drag, and pitching moment coefficients with non-dimensional speed C Lα , C Dα , C Mα = lift, drag, and pitching moment slopes C Lδe , C Mδe = variation of lift, and pitching moment coefficients with elevator deflection C ℓp , C np , C Yp = variation of rolling, yawing and side force coefficients with roll rate C ℓr , C nr , C Yr = variation of rolling, yawing and side force coefficients with yaw rate C ℓβ , C nβ , C Yβ = variation of rolling, yawing and side force coefficients with sideslip angle C ℓδr , C nδr , C Yδr = variation of rolling, yawing and side force coefficients with rudder deflection C ℓδa , C nδa , C Yδa = variation of rolling, yawing and side force coefficients with aileron deflection c = medium chord e n = two-dimensional free transition criterion h o = stick fixed neutral point k = lift dependent drag factor or induced drag factor n = fixed parameter for two-dimensional free transition criterion α = angle of attack