Design, experimental investigation, and nonlinear flight dynamics with atmospheric disturbances of a fixed-wing micro air vehicle

Performance analysis, dynamic simulation and control of mass-actuated airplane

Development and Flight Test of Moving-mass Actuated Unmanned Aerial Vehicle Sampath Reddy Vengate, M.S. The University of Texas at Arlington, 2016 Supervising Professors: Dr. Atilla Dogan Conventional airplane control is achieved by aerodynamic control surfaces by generating moments around all the three axes of the aircraft. Deflections of the control surfaces have some disadvantages such as induced drag, increase in radar signature, and exposure to high temperature in high speed applications. As an alternative moment generation mechanism, prior research proposed internal mass-actuation, which is to generate gravitational moment by changing the center of gravity of the aircraft through motion of internal masses within the aircraft. Prior research investigated the feasibility and benefit of internal mass-actuation in airplane control based on simulation analysis. The main focus of this research is to design, build and flight test a UAV (Unmanned Aerial Vehicle) with internal mass-actuation, as a proof-of-concept. Specifically, this effort has built and flight tested a small electric powered UAV with an internal mass within each wing to generate rolling moment instead of aerodynamic rolling moment by ailerons. The internal structure of each wing is specifically designed to place a linear electric actuator that moves the internal mass. The aircraft is also equipped with all three conventional control surfaces. Most parts of the airplane were laser cut based on 3D CAD designs. The airplane is also equipped with a data ac-

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Development and Flight Test of Moving-mass Actuated Unmanned Aerial Vehicle

In this research effect of propeller induced flow on aerodynamic characteristics of low aspect ratio flying wing micro aerial vehicle has been investigated experimentally in subsonic wind tunnel. Left turning tendencies of right-handed propellers have been discussed in literature, but not much work has been done to quantify them. In this research, we have quantified these tendencies as a change in aerodynamic coefficient with a change in advance ratio at a longitudinal trim angle of attack using subsonic wind tunnel. For experimental testing, three fixed pitch propeller diameters (5 inch, 6 inch and 7 inch), three propeller rotational speeds (7800, 10800 and 12300 RPMs) and three wind tunnel speeds (10, 15 and 20 m/s) have been considered to form up 27 advance ratios. Additionally, wind tunnel tests of 9 wind mill cases were conducted and considered as baseline. Experimental uncertainty assessment for measurement of forces and moments was carried out before conduct of wind tunnel tests. Large variation in lift, drag, yawing moment and rolling moment was captured at low advance ratios, which indicated their significance at high propeller rotational speeds and large propeller diameters. Side force and pitching moment did not reflect any significant change. $\frac {L}{D}$ at trim point was found a nonlinear function of propeller diameter to wingspan ratio $\frac {D}{b}$ , and propeller rotational speed. Rate and control derivatives were obtained using unsteady vortex lattice method with propeller induced flow effect modeled by Helical Vortex Modeling approach. In this research, we have proposed improved 6-DOF equations of motion, with a contribution of advance ratio $J$ . It is concluded, that propeller induced flow effects have a significant contribution in flight dynamic modeling for vehicles with large propeller diameter to wingspan ratio, $\frac {D}{b}$ of 22% or more.

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Experimental Investigation of Propeller Induced Flow on Flying Wing Micro Aerial Vehicle for Improved 6DOF Modeling

This paper presents the nonlinear six degrees of freedom dynamic modeling of a fixed wing micro air vehicle. The static derivatives of the micro air vehicle are obtained through the wind tunnel testing. The propeller effects on the lift, drag, pitching moment and side force are quantified through wind tunnel testing. The dynamic derivatives are obtained through empirical relations available in the literature. The trim conditions are computed for a straight and constant altitude flight condition. The linearized longitudinal and lateral state space models are obtained about trim conditions. The variations in short period mode, phugoid mode, Dutch roll mode, roll subsidence mode and spiral mode with respect to different trim operating conditions is presented. A stabilizing static output feedback controller is designed using the obtained model. Successful closed loop flight trials are conducted with the static output feedback controller.

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Modeling and Closed Loop Flight Testing of a Fixed Wing Micro Air Vehicle

An experimental study has been carried out on a typical Micro Air Vehicle of span 300mm having inverse Zimmerman planform. The objective is to get i) the aerodynamic characteristics of the vehicle in the range of incidence and sideslip angle the vehicle expected to encounter during its flight; ii) an understanding of the propeller effect on the aerodynamic data and iii) the control surface (elevon) effectiveness with incidence. Tests were carried out in a low speed wind tunnel at a freestream velocity of 8 m/s and 12 m/s corresponding to a test Reynolds number based on chord of about 120000 and 180000 respectively. Analysis of the aerodynamic data showed significant effect of propeller flow field on the lift, stall angle and drag of the vehicle. The propeller induced flow is seen to increase the lift coefficient at higher angle of attack and delay the stall. Nonlinear variation is observed in the rolling moment indicating the onset of asymmetric flow field at higher incidence. The effectiveness of the elevon is observed to increase linearly with incidence.

https://www.researchgate.net/profile/Suraj_C_S/publication/258424535_Experimental_Studies_on_a_Propelled_Micro_Air_Vehicle/links/0deec5283010e53072000000.pdf

Experimental Studies on a Propelled Micro Air Vehicle

This paper attempts to bring out the challenges associated with the design of a high altitude mini UAV especially from the aerodynamics perspective. The mini UAV under consideration is a 2 kg class conventional, high wing and T-Tail configuration. A comparative study of various high lift airfoils has been done to illustrate that the selection of a suitable airfoil for high altitude applications is indeed an important part of the design activity and it shows that the wing loading of a UAV designed for high altitudes does not depend on the changes in air density alone. The chosen high lift low Reynolds number airfoil is found to have a minor effect on the aerodynamic parameters (Cl and Cd) with changes in Reynolds number. This paper also addresses the performance variation due to operation at off design condition such as at sea level.

https://nal-ir.nal.res.in/12109/1/cp135.pdf

Design of a High Altitude Fixed Wing Mini UAV - Aerodynamic Challenges

This paper deals with an experimental investigation of a family of cone-cylinder bodies having circular and equivalent elliptical (a/b = 2.0) cross sections with variation in nose and afterbody fineness ratios at Mach number 2.0. All elliptical bodies are tested at roll orientations of 0 deg (semimajor axis normal to cross flow velocity) and 90 deg (semiminor axis normal to crossflow velocity). The nose fineness ratios for these bodies are 1.0, 2.0, and 3.0, and the afterbody fineness ratios are 5.0 and 7.0. The effect of changing the cross section shape from circular to elliptical, the roll orientation, and the variation in nose and afterbody fineness ratios on aerodynamic characteristics are discussed. (Author)

Experimental investigation on cone cylinder bodies of elliptical and circular cross sections at M = 2.0

A wind-tunnel test program was conducted to generate a systematic aerodynamic database for airbreathing vehicles. Generic models consisting of tangent ogival nose, cylindrical body with cruciform intakes or twin intakes were tested at freestream Mach numbers ranging from 0.5 to 3.0. The length and span of intakes were varied. The intakes were two-dimensional with blocked entry. Normal force and pitching moment were nondimensionalized using planform area and distance of centroid (from nose tip) of the planform of the model rather than body cross-sectional area and body diameter, which are traditionally used. When normal-force and pitching-moment coefficients nondimensionalized this way are plotted against angle of incidence, the coefficients of different configurations coalesce for zero roll. In addition, data for different roll angles are found to coalesce when an empirical function of roll angle is introduced in the nondimensionalizing. A prediction method was developed to estimate the normal force and pitching moment of similar body-intake configurations based on this trend. Nomenclature A = cross-sectional area A B = body cross-sectional area of the configuration, = πr 2 A P = planform area of configuration APB = planform area of body API = planform area of intakes alone, A P − APB A R = reference area (equal to A P unless otherwise specified) Cdn = crossflow drag coefficient of circular cylindrical section Cm = pitching-moment coefficient about nose, M p/ qA R X CmNL = nonlinear component of pitching-moment coefficient about nose C N = normal-force coefficient, = N/ qA R C N NL = nonlinear component of normal-force coefficient cn = local normal-force coefficient per unit length d = body diameter H = height of air intake l = length of model li = length of air intake M = freestream Mach number M p = pitching moment about nose N = normal force q = freestream dynamic pressure r = body radius s = total span of body-intake configuration W = width of air intake

Effect of Intake Geometry on Longitudinal Aerodynamics of Airbreathing Vehicles

KNOWLEDGE of intake drag is important in estimating the
performance of the aerospace vehicle. The contribution to the intake drag is mainly due to the external and internal flow through the intake and termed as external and internal drag. External drag is mainly composed of drag due to cowl, diverter, and spillage of flow (pre-entry drag), while the internal drag is due to the wall skin friction and pressure distribution on the walls of the duct. Various components
of the drag associated with a typical intake of an air-breathing vehicle are shown in Fig. 1. It is possible to estimate the internal drag in the absence of any flow separation while the estimation of external drag is difficult particularly when the intake is operating in subcritical condition. At this operating condition of intake, experimental method is the only source to determine the intake drag. Hitherto, drag due to intake alone is measured having the intake specially mounted to the forebody of the fuselage and front portion of intake forming a metric part with a balance.Available test techniques demand a special balance and suitable modifications to the intake module [1]. This is cumbersome and involves more time and additional cost towards design, fabrication and calibration of new balance. As an alternative, a test technique developed and patented
for the direct measurement of after body drag in the base flow wind tunnel [2–4] is used to demonstrate the usefulness of this technique for the measurement of intake drag in a limited Mach number range. An important feature of this test technique (Fig. 2) is that it facilitates
simultaneous measurements of total pressure recovery profiles of the duct flow as well as intake drag for the different mass flow conditions.

Sajeer Ahmed

Papers

Experimental Studies on a Propelled Micro Air Vehicle

Design of a High Altitude Fixed Wing Mini UAV - Aerodynamic Challenges

Experimental investigation on cone cylinder bodies of elliptical and circular cross sections at M = 2.0

Effect of Intake Geometry on Longitudinal Aerodynamics of Airbreathing Vehicles

Test Technique for the Measurement of Intake Drag of Air-Breathing Vehicles