Vibration Analysis of an Embedded Actuator Based UAVpaper.uscip.us/jvamc/jvamc.2015.1010.pdf · Journal of Vibration Analysis ... compared to experimentally measured frequencies from

Columbia International Publishing Journal of Vibration Analysis, Measurement, and Control (2015) Vol. 3 No. 3 pp. 174-193 doi:10.7726/jvamc.2015.1010

Research Article

______________________________________________________________________________________________________________________________ *Corresponding e-mail: [email protected] 1 Mechanical Engineering, California State Polytechnic University at Pomona, USA 2 Aerospace Engineering, California State Polytechnic University at Pomona, USA

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Vibration Analysis of an Embedded Actuator Based UAV

Kevin R. Anderson1*, Sukwinder Singh Sandhu1, Donald Edberg2, and Steven K. Dobbs2

Received 3 July 2015; Returned for revision 10 July 2015; Received in revised form 12 July 2015; Accepted: 29 November 2015; Published online 12 December 2015 © The author(s) 2015. Published with open access at www.uscip.us

Abstract This paper presents the results of a Finite Element based vibration analysis of a solar powered Unmanned Aerial Vehicle (UAV). The purpose of this paper was to quantify the free vibration, forced vibration response due to differing point inputs in order to predict the relative response magnitudes and frequencies at various wing locations of vibration induced power generators (magnet in coil) excited by gust and/or control surface pulse-decays used to help power the flight of the electric UAV. A Fluid Structure Interaction (FSI) study was performed in order to ascertain pertinent design stresses and deflections as well as aerodynamic parameters of the UAV airfoil. The 3.048 m (10 ft) span airfoil is modeled using Mylar as the primary material. Results show that the free mode in bending is 3.0 Hz while the first forced bending mode is on range of 16.2 to 16.7 Hz depending on the location of excitation. The free torsional bending mode is 28.3 Hz, and the first forced torsional mode is range of 26.4 to 27.8 Hz, depending on the location of excitation. A case study regarding the optimal placement of the embedded actuators was also carried out and results are summarized herein. A Fluid Structure Interaction (FSI) analysis was also carried out which showed the coefficients of aerodynamic drag and lift of to be 0.0052 and 0.077, respectively. Keywords: Vibration; UAV; Modal Analysis; Actuators; Structural Dynamics

1. Introduction This paper describes the vibration analysis and fluid-structure-interaction of an electric motor Unmanned Air Vehicle (UAV) which employs solar cells and magnet in coil vibration response power generators as embedded actuators to induce vibration to aid in flight. Applications for UAVs capable of extreme endurance and loitering times are becoming more important. An integration of multidisciplinary technologies including autonomous flight controls, UAV electric powered aircraft, solar cells and a new manufacturing approach for fabricating Graphene super capacitors that are significantly more efficient than current batteries are key for enabling 24/7 flight times. The UAV analyzed herein is shown in Figure 1.

Kevin Anderson, Sukwinder Singh, Donald Edberg, and Steven Dobbs / Journal of Vibration Analysis, Measurement, and Control (2015) Vol. 3 No. 3 pp. 174-193

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Fig. 1. UAV analyzed in the present study The vibration generators can be positioned inside the wing at various locations to be excited by gusts and control surface pulses to produce structural vibrations to produce power to the aircraft storage devices. In order to aid the further design of UAV with embedded actuators, a FEM based Vibration/FSI study was performed in order to: i) optimize the locations of the magnet in coil generators. Since the generators are weight intensive, they affect the frequencies and mode shapes of the wing. The end goal is to find the wing span and cord locations where the vibration modes offer the highest response magnitudes to excite the vibration generators, ii) identify the wing frequencies so the magnet-spring-coil natural oscillation frequency can be tuned by selecting the springs stiffness that suspend the magnet to match the wing frequency. This tuning will be used to amplify the motion of the magnet oscillation due to a wing vibration, Results for drag, lift, von Mises stresses, and critical modes of vibration will presented. Results of the FSI used to design the solar powered aircraft will be summarized herein.

2. Literature Review Recent years have seen an increased focus on the research and development of UAV system which can fly for a substantial period of time in order to realize specific missions. The work of Lee et al. (2014) gives results of a flight test and power simulations of an UAV powered by solar cells, a fuel cell and batteries. Here the mid-class UAV system used in this study consists of three types of power sources operating simultaneously. These power sources are designed and constructed to share the same operation voltage range and connect to the power bus without additional converters or controllers. The flight test of the target UAV system was conducted for 22.13 hours. Power limitations are the current bottleneck of solar / battery powered UAVs. Several research efforts have focuses on high-altitude, long endurance (HALE) UAV missions. Noteworthy studies in this regard include the work of Romeo et al. (2006), Colozza & Dolce (2005) and Nickol et al. (2007). The previous studies of solar-powered HALE UAVs have shown that their development is connected with the critical aspect of generation and storage of solar power. The study of Krings et al. (2013) presents the design and development of the so-called Unmanned Low-cost Testing and Research Aircraft (ULTRA) project. The work of Tuzcu et al. (2007) presents the stability and control of a HALE UAV. The more recent work of Gao et al. (2013) presents a method to determine the design


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parameters for a solar powered HALE UAV. Bahrami (2011) presents the design and fabrication of a solar power system for as experimental UAV. The work of Jasnani et al. (2013) presents sizing and preliminary hardware testing of a solar powered UAV. The work of Suzuki et al. (2012) presents an automatic battery replacement system for UAVs. Chang and Yu (2015) investigate improving electric powered UAV endurance by incorporating battery dumping. Here extensions to electric powered UAVs’ endurance are realized by physically dumping exhausted batteries out of the aircraft while in flight. The UAV propulsion system is considered in the study of Lindhal et al. (2012), whereby simulation, design, and validation of an UAV solid oxide fuel cell (SOFC) propulsion system is presented. Trends toward the future of UAVS are studied in small UAV automation using MEMS in the work of Jang et al. (2007). In support of UAV development work, finite element based modeling studies have been used to guide the design of the UAV system. From the modeling perspective, the work of Oliver et al. (2007) details the implementation of a meta-model based correlation technique on a composite UAV wing test piece and associated finite element model. In the work of Oliver et al. (2007), NASTRAN finite element models of the four structural components are correlated independently, using modal frequencies as correlation features, before being joined together into the assembled structure and compared to experimentally measured frequencies from the assembled wing in a cantilever configuration. Results show that significant improvements can be made to the assembled model fidelity. The study of Haghighat et al. (2012) presents modeling and Simulation of Flexible UAVs with large aspect ratio wherein the full non-linear equations of motion for a flexible UAV are developed. The developed model is applied to a generic UAV with large aspect ratio and the wing tip acceleration is compared to the wing tip acceleration of the rigid aircraft. The study of Ahmad and Rahman (2012) presents results for flutter analysis of the X-HALE UAV-A test bed for aeroelastic results validation. The X-HALE UAV of Ahmad and Rahman (2012) is a test bed exhibiting large structural deformation. Equipped with strain gauges and other measuring sensors, it will provide experimental data which can then be used for nonlinear aeroelastic analyses for other such kind of structures. This papers deals with the linear aeroelastic analysis of this type of aircraft. The study of Waisman (2004) provides open-loop flutter analysis of a composite UAV model using the active stiffening effect. In the more recent work by Di Palma, et al. (2014) the Italian Aerospace Research Center (CIRA) is currently designing an unmanned aerial research system that is lightweight and has high-structural flexibility, code named the High Altitude Performance Demonstrator (HAPD) was analyzed with attention focused on primary structure stiffness distributions (fuselage, wings, and vertical tail), compatibly with the absence of any aeroelastic instability, and structural failure under operative loads. In the studies of Ahmad and Rahman (2013) and Balakhrisnan et al. (2014) an aeroelastic analysis of a high aspect ratio wing in subsonic flow was carried out. Here it is noted that High altitude long endurance (HALE) UAVs have very flexible wing because of mission requirements. High lifts to drag ratio and more surveillance time in air requirements make their structure highly flexible and large tip deflections can occur which can be as large as 30% of wing semi span. Linear theory fails to accurately analyze such deformation and the changes in the structural and aerodynamic characteristics of the wing accompanying such deformation. Low aspect ratio wing and stiff structures can be best simulated by normal modes in aeroelastic analysis. However high aspect ratio restricts the use of reduce order model based on linear normal modes in aeroelastic


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analysis. The reason not to use the linear elastic normal modes in high AR wings is the stiffening effects because of large deflections. These effects cause the normal modes to fail as a good basis set. In such situation one needs to use the modified basis function. The computational model used in this work consists of a modified modal based reduced order nonlinear structural dynamics model coupled to a higher-order panel method for the aerodynamic forces. In the study of Kotkkalpudi et al. (2014) a nonlinear simulation of a flexible UAV is presented for a high aspect ratio UAV. Here, the aerodynamic forces and moments are obtained using a doublet lattice panel method. The interaction between structural dynamics and aerodynamics is given special attention from a modeling standpoint. The simulation is finally integrated into the existing simulation infrastructure maintained by the research group. A modular approach is emphasized in the simulation build-up, which allows for easy switching between models. From the materials and manufacturing standpoint, the work of Grodzki and Lukaszewicz (2015) details the design and manufacture of UAV wing structure using numerical analysis of composite materials. Based upon the literature review presented above, the present vibrational analysis of a UAV employing embedded actuators appears to be warranted. The following sections of this paper outline the finite element modeling and results of our investigation.

3. UAV Airfoil Geometry and Materials 3.1 Airfoil Geometry

For this study, ANSYS workbench was employed. The ANSYS workbench is a quick turn-around turn-key development environment which allows the engineer to quickly analyze complicated structures. The use of ANSYS workbench is proliferate in academia and industry as given by the study of Snyder and Mo (2014). The primary challenge with the project is the non-linear geometry of the aircraft wing. The AG-25 is the airfoil profile which was used for the designing of the airfoil, there are 3 different sections for the wing. There is a uniform spar in the wings for the linear section to improve the rigidity of the airfoil. The first section which is attached to the fuselage is at an angel of 2° whereas the range of the airfoil scale remains constant for the section as shown in the Figure 2 from the ANSYS workbench environment.

Fig. 2. First section of the left wing

The second and third sections are attached to the first and second sections at an angle of 10° and 15° with the horizontal plane respectively with the airfoil profile scale decreasing constantly till the


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outer tip of wing as shown in the Figure 3 and Figure 4.

Fig. 3. Second section of the wing

Fig. 4. Third section extended at an angle of 15°

3.2 Materials Mylar is one of the high performance plastic, since it is very light weight and can retain physical properties for wide temperature range it is the best option as a stressed skin for airfoil. Figure 5 shows the linear isotropic elastic properties and physical properties of Mylar and assigning Mylar as the default material for the airfoil. Table 1 lists the mechanical properties used for the modeling of the Mylar.

Table 1 Elastic properties for Mylar

Property Value Density 1.39 g/cm3 Young’s Modulus 2800 MPa Poisson’s Ratio 0.37 Bulk Modulus 3.5897109 Pa Shear Modulus 1.0219109 Pa

4. Finite Element Modeling 4.1 Mesh Since nonlinearities of material exist, the quality of the mesh plays a significant role since as the finer the mesh becomes, the larger does the stiffness matrix, which could potentially lead the


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solution to diverged solutions because of unbalanced forces. In contrast, for the coarser stiffness matrices one may encounter singular matrices for the faces which would lead to the divergence of the solution. Here, the finite element mesh is constructed using tetrahedron cells which work well for both fluid flow systems as well as for mechanical vibrations problems. The mesh is kept at a medium quality with edge sizing for all the edges of the surface and by introducing the so-called ANSYS “sphere of influence “local mesh refinement for all of the sensitive regions of the airfoil. Figure 5 shows the meshing of the airfoil as taken from ANSYS. Since our primary objective was not to capture or analyze the vortex or downwash wake formation of the airfoil a discrete mesh size for the leading and trailing edge of the airfoil was not formulated.

Fig. 5. Finite element mesh

5. Results 5.1 Free Modal Analysis of UAV Airfoil Figure 6 through Figure 11 document the results for the vibration modal analysis. Figure 6 shows the first bending mode shape for the airfoil without a lumped mass (free vibration) which shows a natural frequency in bending of 3.0 Hz. Figure 6 shows the un-deformed shape and the maximum deformation for the airfoil. The use of point masses for excitation is shown in Figure 7 and Figure 8.

Here, a lumped mass is placed near the airfoil edge in order to alter the frequency mkn / by

increasing the mass, m. Figure 7 shows the first bending mode shape for the airfoil with a lumped mass at location 1 which results in a maximum frequency of 16.729 Hz. Figure 8 gives the first bending mode shape for the airfoil with a lumped mass placed at the outer edge giving a maximum frequency of 16.616 Hz. From Figure 8, since the mass is moved further close to the outer edge of the wing, the deformation of the wing is increased from 89.5 mm to 90 mm.


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Fig. 6. Bending mode for free vibration

Fig. 7. 1st Bending mode shape with lumped mass at location 1

Fig. 8. Bending mode shape with lumped mass at the location 2

For this particular project we are using subspace method for computing the eigenvalues and eigenvectors of the vibrations. Figure 9 shows the second bending mode shape without lumped mass corresponding to a frequency of 17. 719 Hz. Figure 10 shows the second bending mode shape with lumped mass at location 1 giving a frequency of 16.729 Hz. Figure 11 indicates the second bending mode shape with the lumped mass at location 2 giving a frequency of 16.616 Hz. Figure 12 shows the torsional modal shape for free vibrations with maximum deformation of 99.174 mm and frequency of 28.306 Hz. Figure 13 shows the torsional modal shape for the airfoil with lumped mass at location 1 with maximum frequency of 26.433 Hz at the outer edge and maximum deformation of 103.1mm. Figure 14 shows the torsional modal shape with lumped mass at spot 2 with the maximum frequency of 27.814 Hz at the outer edge and maximum deformation of 100.56 mm. The above results are found to be in qualitative agreement with the findings of Di Palma et al. (2014).


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Fig. 9. 2nd Bending mode shape without lumped mass

Fig. 10. 2nd bending mode shape with lumped mass at spot 1

Fig. 11. 2ndbending mode shape with lumped mass at location 2

Fig. 12. Torsional mode shape for free vibrations


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Fig. 13. Torsional Modal Shape with lumped mass at location 1

Fig. 14. Torsional Modal Shape with lumped mass at location 2

5.2 Embedded Actuator Forced Vibration Modal Analysis Here the vibration excitation due to embedded actuators is simulated by placing several actuators in the UAV airfoil, and then applied a load at each actuator location, thus leading to forced vibration. The simulation of the embedded actuators was carried out in order to ascertain the proper control scenario in which to assist the UAV during flight. Various case studies of placing the embedded actuators at various spatial locations on the left and right wing of the UAV airfoil were simulated. Here, we present the results for three such configuration case studies. For performing pre-stress analysis the static structural system is coupled to the dynamical the modal system a frequency responses is obtained. Th deformation and modal shapes for the loaded wing are carried out for 10 modes.

For the optimum numerical solution, ANSYS employs the iterative method to solve for the eigenvalues. Although it takes much more computational time than the direct method of solution, the iterative method of solution is much more accurate. For the modal settings the damping of the airfoil is turned on with the proportional coefficient to be 0.03 < < 0.05. Next, results for three (3) different case studies regarding the placement of the embedded actuators are presented.

For the first asymmetric case study, five actuators (generators) were located on the leading edge of the left wing and five generators were placed on the trailing edge of the right wing of the airfoil. The actuator excitation forces used in the simulation were 4 N (0.9 lbf) on the leading edge and tip of


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the left wing, and 2 N (0.45 lbf) on the leading edge and tip of the right wing. Figure 15 shows the FEM mesh used for this control configuration.

Fig. 15. Location of generators for first asymmetric excitation case study

Table 1 lists the frequencies corresponding to the control configuration of Figure 15. The deformation and modal shape for the 5th torsional, and 6th torsional modes corresponding to the control strategy of Figure 15 are shown in Figure 16 and Figure 17, respectively. Table 1 Asymmetric excitation first case study

Mode Frequency (Hz) Max. Deflection (mm)

1 11.81 62.853 2 11.822 62.896 3 58.045 78.066 4 58.161 78.289 5 93.707 71.617 6 95.121 72.751 7 149.18 85.073 8 149.6 86.696 9 233.59 108.38 10 239.9 109.27

Fig. 16. Deformation and modal shape for 5th torsional mode


For the second asymmetric case study, five actuators were staggered spatially on the left wing and five generators were staggered spatially on right wing of the airfoil. The actuator excitation forces


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used in the simulation were 4 N on the leading edge and tip of the left wing, and 2 N on the leading edge and tip of the right wing. Figure 18 shows the FEM mesh used for this control configuration.


Table 2 lists the frequencies corresponding to the control configuration of Figure 18. The deformation and modal shape for the 5th torsional, and 6th torsional modes corresponding to the control strategy of Figure 18 are shown in Figure 19 and Figure 20, respectively. Table 2 Asymmetric excitation second case study


1 11.756 37.328 2 11.762 37.333 3 57.834 46.537 4 57.84 46.274 5 94.109 42.878 6 94.419 42.784 7 148.61 52.168 8 148.72 49.879 9 231.09 66.448 10 231.35 72.89




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The final case study of asymmetric modal analysis using embedded actuators used , five actuators were staggered arranged spatially concentrated near the outboard region on the left wing and five generators were arranged spatially concentrated in the vicinity of the outboard area of the right wing of the airfoil. The actuator excitation forces used in the simulation were 4 N on the leading edge and tip of the left wing, and 2 N on the leading edge and tip of the right wing. Figure 21 shows the FEM mesh used for this control configuration.


Table 3 lists the frequencies corresponding to the control configuration of Figure 21. The deformation and modal shape for the 5th torsional, and 6th torsional modes corresponding to the control strategy of Figure 21 are shown in Figure 22 and Figure 23, respectively. Table 3 Asymmetric excitation second case study


1 11.538 61.575 2 11.539 61.567 3 57.395 74.413 4 57.457 74.181 5 92.282 70.407 6 92.514 70.099 7 146.28 80.522 8 146.64 79.700 9 222.67 115.63 10 227.11 121.51



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The results above from the asymmetrically embedded actuator forced vibration analysis indicate that for the first asymmetric loading case in which five actuators (generators) were located on the leading edge of the left wing and five generators were placed on the trailing edge of the right wing of the airfoil, the first, second and third modal frequencies are 11.81 Hz, 11.822 Hz, and 58.045 Hz corresponding to maximum deflections of 62.853 mm, 62.896 mm, and 78.066 mm, respectively. For the second asymmetric loading case whereby five actuators were staggered spatially on the left wing and five generators were staggered spatially on right wing of the airfoil, the first, second and third modal frequencies are 11.756 Hz, 11.762 Hz, and 57.834 Hz corresponding to maximum deflections of 37.328 mm, 37.333 mm, and 46.537 mm, respectively. For the third asymmetric loading case where five actuators were staggered arranged spatially concentrated near the outboard region on the left wing and five generators were arranged spatially concentrated in the vicinity of the outboard area of the right wing of the airfoil, the first three modal frequencies are 11.538 Hz, 11.539 Hz, and 57.395 Hz corresponding to maximum deflections of 61.575 mm, 61.567 mm and 74.413 mm, respectively. Hence, it is clear that the architectural layout and placement of the embedded actuators has a profound effect on the vibrational characteristics of the UAV airfoil. 5.3 Fluid-Structure-Interaction Analysis The first step for the FSI setup is to figure out whether to use two way coupling or one way coupling method. In this project the more direct approach of one-way coupling was adopted in order obtain reliable and accurate results with a quick-turn-around time. By using one-way coupling the output variable of pressure is derived from ANSYS CFD FLUENT a software package and is transferred as an input parameter into Mechanical System as shown in the Figure 24 taken from ANSYS Workbench. The workflow tree of this coupled FSI analysis within ANSYS as shown in Figure 25. The FSI analysis was carried out by meshing a large “wind tunnel” around the airfoil situated at zero-angle of attach. This allowed the implementation of a free-stream velocity boundary condition to be input and the resulting one-way pressure coupling from the CFD package to the mechanical vibrations package of ANSYS to be achieved in a reasonable time fashion.


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Fig. 24. FSI analysis set-up

Fig. 25. Mechanical Model Outline

The derived Reynolds number Re = ρVD/μ was found to exceed Re = 500,000 so the flow is turbulent in nature. Hence the k- turbulence model was used to model the turbulence. The first 150 iterations were run using first order upwind followed by 850 iterations using second order upwind in order to accurately resolve the gradual change in the values for the lift coefficient Cl and drag coefficient Cd residuals. The solution was seen to converged for first order upwind scheme at the 141st iteration and the 851st iteration for the second order upwind scheme, respectively. Using the first order upwind scheme enables the elimination of any irregularities in the values for aerodynamic forces, Cl and Cd. Figure 27 shows the equivalent elastic stain a.k.a. von Mises Strain with the contours showing the maximum strain for the wing where it is attached to fuselage. From Figure 26, the maximum value for non Mises strain =0.000101 mm/mm.


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Fig. 26. Equivalent elastic stain

Figure 27 indicates the von Mises Stress 0.00037 < < 0.282 MPa with the maximum value oriented at the center of the airfoil where it engages the fuselage, as expected.

Fig. 27. von Mises Stresses

Figure 28 shows the total deformation of the airfoil due to the aerodynamic forces ranging from 0 < w < 3.4 mm.

Fig. 28. Total deformation


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Figure 29 shows the dynamic pressure contours over the airfoil with pressure ranging from -1.44 kPa < p < 1.23 kPa. Figure 30 indicates the velocity vector field over the airfoil with a magnitude ranging from 21.7 < U < 66 m/s. Figure 31 shows the overall force reaction due to the aerodynamic forces acting on the airfoil with the reaction centered near the airfoil as expected. Figure 32 gives the overall moment reaction for the airfoil.

Fig. 29. Pressure contours

Fig. 30. Velocity vector field

Fig. 31. Force reaction


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Fig. 32. Overall moment reaction

The aerodynamic derivatives of drag Cd and lift Cl are plotted in Figure 33 and Figure 34, respectively.

Fig. 33. Coefficient of drag

Fig. 34. Coefficient of lift

A sanity check was performed on the coefficient of drag and coefficient of lift by performing back-of-the-envelope calculations. These hand-calculated values of Cd =0.005167, and Cl=0.07698 shown in Table 4 together with other miscellaneous aerodynamic force parameters. The hand compute values of Table 4 are found to be in reasonable agreement with the numerical FSI results as shown the asymptotes of Figure 33 and Figure 34, respectively.


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Table 4 Aerodynamic force results from FSI analysis

Parameter Vale Lift coeff. 0.077 Drag coeff. 0.0052 Lift /Drag coeff. 14.9 Drag force 11.74 N Lift force 47.83 N Center of press. (-6.8, -10.09,1.669)

Press. moment -50.96 in +x-dir. Max. press. coeff. 0.98

6. Conclusions

This paper has presented the results of a Finite Element based vibration analysis of a solar powered Unmanned Aerial Vehicle (UAV) with a wing span of 3.048 m. The wing is modeled as Mylar being the primary material. The solar battery powered UAV uses vibration induced magnet-in-coil power generators embedded in its airfoil and excited by gusts and control surface pulse-decays to augment the power required for the electric motor flight. Using FEM based Vibration Analysis affords the design team a better indication of where to locate these actuators. This paper has presented results for i) free modal analysis of the UAV airfoil, ii) forced vibration analysis of the UAV airfoil to quantify the cause and effect of asymmetrically embedded actuators on the vibration of the UAV airfoil and finally iii) a Fluid Structure Interaction (FSI) for the UAV. Results from i) free modal analysis of the UAV airfoil show that the free mode in bending is 3.0 Hz while the first forced bending mode falls in the range of 16.2 to 16.7 Hz depending on the location of excitation. The free torsional bending mode is 28.3 Hz, and the first forced torsional mode is range of 26.4 to 27.8 Hz, depending on the location of excitation. Results for the forced vibration analysis of the UAV airfoil show that for the first asymmetric loading case in which five actuators (generators) were located on the leading edge of the left wing and five generators were placed on the trailing edge of the right wing of the airfoil, the first modal frequency is 11.81 Hz with a maximum deflection of 62.853 mm. For the second asymmetric loading case whereby five actuators were staggered spatially on the left wing and five generators were staggered spatially on right wing of the airfoil, the first modal frequency is 11.756 Hz, with corresponding maximum deflection of 37.328 mm. For the third asymmetric loading case where five actuators were staggered arranged spatially concentrated near the outboard region on the left wing and five generators were arranged spatially concentrated in the vicinity of the outboard area of the right wing of the airfoil, the first modal frequency is 11.538 Hz and a maximum deflection of 61.575 mm. Thus, the physical layout of how the embedded sensors are arranged in the UAV airfoil is seen to have a profound effect on the vibrational characteristics of the structure. The results of the FSI analysis afforded coefficients of aerodynamic drag and lift of to be 0.0052 and 0.077, respectively which were found to agree with hand calculations. Future work will encompass a flutter analysis of the UAV studied herein.

Funding Source None


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Conflict of Interest None

References Ahmad, K.W. & Rahman, H. (2012). Flutter analysis of X-HALE UAV-A test bed for aeroelastic results

validation. Applied Mechanics and Materials. 245(303). http://dx.doi.org/10.4028/www.scientific.net/AMM.245.303

Ahmad, K.W. & Rahman, H. (2013). Aeroelastic Analysis of High Aspect Ratio Wing in Subsonic Flow. Proceedings of 2013 10th International Bhurban Conference on Applied Sciences & Technology (IBCAST). 219-223. http://dx.doi.org/10.1109/IBCAST.2013.6512157

Bahrami T.H. (2011) Solar Power System for Experimental Unmanned Aerial Vehicle (UAV); Design and Fabrication. 2nd Power Electronics, Drive Systems and Technologies Conference,129-134. http://dx.doi.org/10.1109/pedstc.2011.5742404

Balakrishnan, A.V., Tuffaha, A.M., Patino, I. & Melnikov, O. (2014). Flutter Analysis of an Articulated High Aspect Ratio Wing in Subsonic Airflow. Journal of the Franklin Institute-engineering and Applied Mathematics. 351(8), 4230-4250. http://dx.doi.org/10.1016/j.jfranklin.2014.04.010

Chang, T., & Hu Y. (2015). Improving electric powered UAVs' endurance by incorporating battery dumping concept. Procedia Engineering, 99,168-179. http://dx.doi.org/10.1016/j.proeng.2014.12.522

Colozza, A., & Dolce, J. L. (2005). "High-altitude, long-endurance airships for coastal surveillance." NASA Technical Memorandum, TM 2005-213427, National Aeronautics and Space Administration (NASA) Glenn Research Center, Cleveland, OH.

Di Palma, L., Belardo, M., Pecora, M. & Paletta N. (2014). Structural and Aeroelastic Design of a Joined-Wing UAV. Journal of Aerospace Engineering. 27(1): 93-111. http://dx.doi.org/10.1061/(ASCE)AS.1943-5525.0000251

Gao, X.Z., Hou, Z.X., Zheng, G., Zhu, X.F., Liu, J.X. & Chen, X.Q. (2013). Parameters Determination for Concept Design of Solar-powered, High-altitude Long-endurance UAV. Aircraft Engineering and Aerospace Technology. 85(4) 293-303. http://dx.doi.org/10.1108/AEAT-Jan-2012-0011

Grodzki, W. & Łukaszewicz, A. (2015). Design and manufacture of Unmanned Aerial Vehicles (UAV) wing structure using composite materials. Materialwissenschaft Und Werkstofftechnik. 46(3), 269-278. http://dx.doi.org/10.1002/mawe.201500351

Jang, J.S., & Liccardo, D. (2007). Small UAV automation using MEMS. IEEE Aerospace and Electronic Systems Magazine, 22(5), 30-34. http://dx.doi.org/10.1109/MAES.2007.365332

Jashnani, S., T.R Nada, M Ishfaq, A Khamker, and P Shaholia. (2013). Sizing and preliminary hardware testing of solar powered UAV. The Egyptian Journal of Remote Sensing and Space Science. http://dx.doi.org/10.1016/j.ejrs.2013.05.002

Kotikalpudi, A., Moreno, C., Taylor, B., Pfifer, H. & Balas, G. (2014). Low Cost Development of a Nonlinear Simulation for a Flexible Uninhabited Air Vehicle. 2014 American Control Conference. 2029-2034. http://dx.doi.org/10.1109/acc.2014.6859455

Krings, M., Annighofer, B. & Thieleckey, F. (2013). ULTRA - Unmanned Low-cost Testing Research Aircraft. 2013 American Control Conference. 1472-1477.

http://dx.doi.org/10.4028/www.scientific.net/AMM.245.303

http://dx.doi.org/10.1109/IBCAST.2013.6512157

http://dx.doi.org/10.1109/pedstc.2011.5742404

http://dx.doi.org/10.1016/j.jfranklin.2014.04.010

http://dx.doi.org/10.1016/j.proeng.2014.12.522

http://dx.doi.org/10.1061/%28ASCE%29AS.1943-5525.0000251

http://dx.doi.org/10.1108/AEAT-Jan-2012-0011

http://dx.doi.org/10.1002/mawe.201500351

http://dx.doi.org/10.1109/MAES.2007.365332

http://dx.doi.org/10.1016/j.ejrs.2013.05.002

http://dx.doi.org/10.1109/acc.2014.6859455


193


Lee, B., Poomin P., Keunbae K., & Sejin K. (2014). The flight test and power simulations of a UAV powered by solar cells, a fuel cell and batteries. Journal of Mechanical Science and Technology, 28 (1), 399-405. http://dx.doi.org/10.1007/s12206-013-0936-7

Lindahl, P., Moog, E. & Shaw, S.R., (2012). Simulation, design, and validation of an UAV SOFC propulsion system. IEEE Transactions on Aerospace and Electronic Systems, 48(3), 2582-2593. http://dx.doi.org/10.1109/TAES.2012.6237610

Nickol, C. L., Guynn, M. D., Kohout, L. L. & Ozoroski, T. A. (2007). "High altitude long endurance air vehicle analysis of alternatives and technology requirements development." Proc., 45th AIAA Aerospace Sciences Meeting and Exhibit, American Institute of Aeronautics and Astronautics, Reston, VA. http://dx.doi.org/10.2514/6.2007-1050

Oliver, J.A., Kosmatka, J.B., Hemez, F.M., & Farrar, C.R. (2007). Finite element model correlation of a composite UAV wing using modal frequencies. Proceedings of SPIE. 6532(1). http://dx.doi.org/10.1117/12.717456

Romeo, G., Frulla, G., Cestino, E., and Corsino, G. (2004). "HELIPLAT:Design, aerodynamic, structural analysis of long-endurance solar powered stratospheric platform." J. Aircr., 41(6), 1505-1520. http://dx.doi.org/10.2514/1.2723

Snyder, J. & Mo, C. (2014). A study of vibration control of a critical lift tower" Journal of Vibration Analysis, Measurement, and Control, 2(1), 41-56.

Sohrab, H., Haghighat, S., Liu, H., Martins, J. & Martins, J. (2008). Modeling and simulation of flexible UAVs with large aspect ratio. 2008 Asia Simulation Conference - 7th International Conference on System Simulation and Scientific Computing. 504-509.

Suzuki, K., Paulo K.F., & Morrison, J. (2012). Automatic battery replacement system for UAVs: analysis and design. Journal of Intelligent & Robotic Systems, 65(1), 563-586. http://dx.doi.org/10.1007/s10846-011-9616-y

Tuzcu, I., Marzocca, P., Cestino, E., Romeo, G. & Frulla, G. (2007). Stability and Control of a High-Altitude, Long-Endurance UAV. Journal of Guidance, Control, and Dynamics. 30(3),713-721. http://dx.doi.org/10.2514/1.25814

Waisman, H. (2004). Open-loop flutter analysis of a composite UAV model using the active stiffening effect. Finite Elements in Analysis and Design. 40(11),1283-1295. http://dx.doi.org/10.1016/j.finel.2003.06.004


http://dx.doi.org/10.1007/s12206-013-0936-7

http://dx.doi.org/10.1109/TAES.2012.6237610

http://dx.doi.org/10.2514/6.2007-1050

http://dx.doi.org/10.1117/12.717456

http://dx.doi.org/10.2514/1.2723

http://dx.doi.org/10.1007/s10846-011-9616-y

http://dx.doi.org/10.2514/1.25814

http://dx.doi.org/10.1016/j.finel.2003.06.004

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Vibration Analysis of an Embedded Actuator Based UAVpaper.uscip.us/jvamc/jvamc.2015.1010.pdf · Journal of Vibration Analysis ... compared to experimentally measured frequencies from