The helicopter possesses the unrivaled capacity for vertical takeoff and landing which has made the helicopter suitable for numerous tasks such as carrying passengers and equipment, providing air medical services, firefighting, and other military and civil tasks. The nature of the aerodynamic environment surrounding the helicopter gives rise to a significant amount of vibration to its whole body. Among different sources of vibrations, the main rotor blade is the major contributor. The dynamic characteristics of the hingeless rotor consisting of elastic blades are of particular interest because of the strongly coupled equations of motion. The elastic rotor blades are subjected to coupled flapping, lead-lag, and torsional (triply coupled) deflections. Once these deflections exceed the maximum allowable level, the structural integrity of the rotor blade is affected leading to the ultimate failure. The maximum deflection that a blade can undergo for a specific operating condition needs to be estimated. Therefore, in this study, the triply coupled free and forced response of the Bo 105 hingeless, composite helicopter rotor blade is investigated at hovering and forward flights. At first, a model of the composite cross-section of the rotor blade is proposed for which a semi-analytical procedure is developed to estimate the sectional properties. These properties are used in the mathematical model of the free vibration of the rotor blade having the proposed cross-section to solve for the natural frequencies and the mode shapes. The aerodynamic loadings from the strip theory are used to estimate the time-varying forced response of the rotor blade for hovering and forward flights. The large flapping and inflow angles are introduced in the mathematical model of the forward flight and the corresponding nonlinear mathematical model requires a numerical solution technique. Therefore, a generalization of the method of lines is performed to develop a robust numerical solution in terms of time-varying deflections and velocities. The effect of the unsteady aerodynamics at the forward flight is included in the mathematical model to estimate the corresponding dynamic response. Both the analytical and the numerical models are validated by finite element results and the convergence study for the free vibration is performed.

Les essais en environnement libre et en soufflerie ont été effectués pour étudier la performance propulsive et la déformation de pales de référence et de pales souples. La poussée et le couple ont été évalués par deux méthodes: une mesure directe par balance et une estimation indirecte par bilan de quantité de mouvement, les deux méthodes ayant leurs avantages et limitations respectifs. La méthode indirecte s’est construite sur l’acquisition de champs de vitesse obtenus par PIV et s’appuie sur une estimation de la pression par mise en œuvre de l’équation de Poisson. En vol stationnaire, les pales flexibles ne peuvent pas aider à l’amélioration du rendement en mode rotor (FM), à chargement faible, puisque la distribution de vrillage est sans doute assez éloignée de l’optimal de vol stationnaire. En vol avancé, le rendement propulsif des pales flexibles est la plupart du temps plus élevé que l’hélice rigide de référence en raison de la torsion bénéfique généré en rotation. Dans le cas des pales flexibles, la vitesse axiale se trouve être inférieure au cas rigide, à même station aval; ceci correspondant à la la déformation de vrillage négatif. Pour les deux pales, la différence de poussée entre celle déduite du champ PIV test 2et celle obtenue avec la balance est plus grande que la différence entre les valeurs déduites du champ PIV test 1 et de la mesure directe. La technique de mesure laser pour les déplacements(LDS) a été utilisée pour mesurer la déformation stationnaire des pales lors de leur rotation. Par analyse du nuage de points mesurés par la LDS, la flexion et la torsion de la lame en rotation ont été identifiées à l’aide des régressions multiples. / The wind tunnel tests were conducted to explore the performance difference caused by the potential twist deformation between baseline blades and flexible blades. The balance was built in SaBre wind tunnel for measuring the thrust and torque of blades. The BEMT predictions of blades with varied twist were also performed in hover and forward flight, respectively. In hover,flexible blades cannot help in improving the FM at light disk loading since the twist generated on flexible blades is probably beyond the ideal hover twist. In forward flight, the propulsive efficiency η of flexible blades is mostly higher than baseline blades due to the beneficial twist generated in rotation. A Particle Image Velocimetry (PIV) approach of loads determination was developed based on control volume method to obtain thrust and torque of small-scale proprotor,especially for off-optimum conditions. The pressure Poisson equation was implemented for the pressure estimation based on the PIV velocity data. The axial velocity of flexible blades is found to be lower than baseline blades on the same station at downstream. This corresponds to the lower inflow ratio distribution along flexible blade, which results from the negative twist deformation. For both baseline blades and flexible blades, the thrust differences between PIV test 2 and balance are larger when compared to the differences between PIV test 1 based on nearfield and balance. The Laser Displacement Sensor (LDS) technique was employed for measuring the stationary deformation of rotating flexible blades. By obtaining the LDS point cloud, the bending and torsion of the rotating blade were identified using the multiple regressions.

The unsteady aerodynamic and aeroelastic analysis of flapping flight under various kinematics and flow parameters is presented in this dissertation. The main motivation for this study arises from the challenges facing the development of micro air vehicles. Micro air vehicles by requirement are compact with dimensions less than 15-20 cm and flight speeds of around 10-15 m/s. These vehicles operate in low Reynolds number range of 10,000 to 100,000. At these low Reynolds numbers, the aerodynamic efficiency of conventional fixed airfoils significantly deteriorates. On the other hand, flapping flight employed by birds and insects whose flight regime coincides with that of micro air vehicles offers a viable alternate solution. For the analysis of flapping flight, a boundary fitted moving grid algorithm is implemented in a flow solver, GenIDLEST. The dynamic movement of the grid is achieved using a combination of spring analogy and trans-finite interpolation on displacements. The additional conservation equation of space required for moving grid is satisfied. The solver is validated with well known flow problems such as forced oscillation of a cylinder, a heaving airfoil, a moving indentation channel, and a hovering fruitfly. The performance of flapping flight is analyzed using Large Eddy Simulation (LES) for a wide range of Reynolds numbers and under various kinematic parameters. A spiral Leading Edge Vortex (LEV) forms during the downstroke due to the high angle of attack, which results in high force production. A strong spanwise flow of the order of the flapping velocity is observed along the core of the LEV. In addition, the formation of a negative spanwise flow is observed due to the tip vortex, which slows down the removal of vorticity from the LEV. This leads to the instability of the LEV at around mid-downstroke. Analysis with different rotation kinematics shows that a continuous rotation results in better propulsive efficiency as it generates thrust during the entire flapping cycle. Analysis with different angles of attack shows that a moderate angle of attack which results in complete shedding of the LEV offers high propulsive efficiency. The analysis of flapping flight at Reynolds numbers ranging from 100 to 100,000 shows that higher lift and thrust values are obtained for Re?100. The critical reasons are that at higher Reynolds numbers, the LEV is closer to the surface and as it sheds and convects it covers most of the upper surface. However, the Reynolds number has no or little effect on the lift and thrust as identical values are obtained for Re=10,000 and 100,000. The analysis with different tip shapes shows that tip shapes do not have a significant effect on the performance. Introduction of stroke deviation to kinematics leads to drop in average lift as wing interacts with the LEV shed during the downstroke. A linear elastic membrane model with applied aerodynamic load is developed for aeroelastic analysis. Analysis with different wing stiffnesses shows that the membrane wing outperforms the rigid wing in terms of lift, thrust and propulsive efficiency. The main reason for the increase in force production is attributed to the gliding of the LEV along the camber, which results in a high pressure difference across the surface. In addition, a high stiffness along the spanwise direction and low stiffness along the chordwise direction results in a uniform camber and high lift and thrust production. / Ph. D.

This thesis describes the development of a series of codes for trim and longitudinal stability analysis of a helicopter in forward flight. In general, particular use of these codes can be made for parametric investigation of the effects of the external and internal systems integrated to UH-60 helicopters. However, in this thesis the trim analysis results are obtained for a clean UH-60 configuration and the results are compared with the flight test data that were acquired by ASELSAN, Inc. The first of the developed trim codes, called TRIM-CF, is based on closedform equations which give the opportunity of having quick results. The second code stems from the trim code of Prouty. That code is modified and improved during the course of this study based on the theories outlined in [3], and the resultant code is named TRIM-BE. These two trim codes are verified by solving the trim conditions of the example helicopter of [3]. Since it is simpler and requires fewer input parameters, it is more often more convenient to use the TRIM-CF code. This code is also verified by analyzing the Bo105 helicopter with the specifications given in [2]. The results are compared with the Helisim results and flight test data given in this reference. The trim analysis results of UH-60 helicopter are obtained by the TRIM-CF code and compared with flight test data. A forward flight longitudinal dynamic stability code, called DYNA-STAB, is also developed in the thesis. This code also uses the methods presented in [3]. It solves the longitudinal part of the whole coupled matrix of equations of motion of a helicopter in forward flight. The coupling is eliminated by linearization. The trim analysis results are used as inputs to the dynamic stability code and the dynamic stability characteristics of a forward flight trim case of the example helicopter [3] are analyzed. The forward flight stability code is applied to UH-60 helicopter. The codes are easily applicable to a helicopter equipped with external stores. The application procedures are also explained in this thesis.

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