Parawing vehicles may have unusual values of many of the mass and aerodynamic factors affecting dynamic lateral stability and control. These unusual characteristics are due in large part to the fact that the center of gravity of parawing vehicles is located far below the parawing, whereas conventional aircraft usually have the vertical center-of-gravity location near the plane of the wings. The present thesis is an analytical investigation of the dynamic lateral stability and control of a typical parawing vehicle. The analysis was made using three-degree-of-freedom, rigid body equations of motion. Stability derivatives used in the calculations were obtained from static and dynamic force tests of a parawing model with rigid leading-edge and keel members. The analysis is treated mainly in terms of the effects of vertical center-of-gravity position, since this was found to be the most significant factor affecting the lateral stability and control of the hypothetical vehicle. / Master of Science

The physics of flapping is very important in the design of MAVs. As MAVs cannot have an engine that produces the amount of thrust required for forward flight, and yet be light weight, harnessing thrust and lift from flapping is imperative for its design and development. In this thesis, aerodynamics of pitch and plunge are simulated using a 3-D, free wake, vortex lattice method (VLM), and structural characteristics of the wing are simulated as a membrane supported by a rigid frame. The aerodynamics is validated by comparing the results from the VLM model for constant angle of attack flight, pitching flight and plunging flight with analytical results, existing 2-D VLM and a doublet lattice method. The aeroelasticity is studied by varying parameters affecting the flow as well as parameters affecting the structure. The parametric studies are performed for cases of constant angle of attack, plunge and, pitch and plunge. The response of the aeroelastic model to the changes in the parameters are analyzed and documented. The results show that the aerodynamic loads increase for increased deformation, and vice-versa. For a wing with rigid boundaries supporting a membranous structure with a step change in angle of attack, the membrane oscillates about the steady state deformation and influence the loads. For prescribed oscillations in pitch and plunge, the membrane deformations and loads transition into a periodic steady state. / Master of Science

A general nonlinear, nonplanar unsteady vortex panel method for potential-flow is developed. The surface is modeled as a collection of triangular elements on which the vorticity vector is piecewise linearly varying. The wake emanates from the sides and trailing edges of the thin lilting surfaces and is modeled as a progressively formed collection of vortex filaments. This model provides a continuous pressure distribution on the surface while allowing the wake to roll up as tightly as needed. The wake position is determined as part of the solution and no prior knowledge of the position or strength is assumed. An adaptive grid technique is used to redistribute the circulation of the vortex filaments of the wake as the wake sheet spreads. The aerodynamic model is coupled with dynamic equations of motion. Forced oscillation tests are conducted on flat rectangular and delta wings. Dynamic tests are performed to predict wing rock of a slender delta wing restricted to one degree of freedom in roll. The aerodynamic/dynamic model is coupled with control laws that govern the motion of flaperons so that a prescribed pitch motion is executed and wing rock is suppressed. ( 300 pages, 107 figures ) / Ph. D.

Several procedures for the simultaneous aerodynamic-structural design optimization of aircraft wings are investigated. These procedures include efficient methods for optimization and sensitivity calculations that are applied to two specific design examples. The first is a subsonic transport aircraft with a composite forwardswept wing. The aerodynamic modeling for this case is provided by vortex-lattice theory and the structural model initially utilizes finite-element analyses. Even with efficient sensitivity methods, the approximate optimization problem still requires a large computational effort. To reduce this cost, a variable-complexity model for the structural analyses is introduced. First, an algebraic equation model for wing weight is used in the optimization procedure to obtain an aerodynamic design that approximately accounts for the effects of wing geometry on wing weight. Then this design is refined by simultaneous aerodynamic-structural optimization based on the finite-element analysis. The net effect of this dual structural model is a substantial reduction in optimization costs. The second example is the wing design of a supersonic High-Speed Civil Transport (HSCT). For this case, the simple wing-weight equations for structures are retained. For the aerodynamics, a variable-complexity model was introduced with the complex models provided by volumetric wave drag analysis and panel methods. In addition, simple algebraic models for wave and drag due to lift provide inexpensive approximations during most of the optimization cycles. With the minimization of the costly complex sensitivity calculations, a reduction in optimization costs is realized. / Ph. D.

In this study the primary objective was to determine the effect of aspect ratio in particular and in general the effect of three dimensionality on the flow around wings at low Reynolds numbers. It was seen that the effects observed at high Reynolds number are also present in this Re range. There is the usual increase in lift slope and this increase can even be predicted with reasonable accuracy using Prandtl's lifting line theory. In addition to the change in lift slope the zero lift angle of attack was also influenced by the aspect ratio. Through flow visualization it was ascertained that the wingtips have a rather restricted effect on the laminar separation bubble. The disappearance of the bubble extends only for a small distance inboard from the tips. The size of the hysteresis loop and the Reynolds number at which hysteresis starts was found to be influenced by the aspect ratio. The momentum deficit method was used to validate the data obtained by the strain gauge method and there was adequate agreement between the values found through the two methods. From the measurements of pressure done around the airfoil contour one could determine both the location of the laminar separation bubble and the regions were flow is separated. The pressure taps themselves were found to influence measurements somewhat in certain regions of angle of attack and Reynolds number. In the future it would be beneficial to continue strain gauge measurements on this airfoil with flaps and control surfaces to determine their effect on the formation of the laminar separation bubble. Also measurements on other shapes would give more insight into the phenomena occurring here. The effects of turbulence and noise will have to be investigated in detail to determine what performance to expect from an actual aircraft. Finally detailed measurements on boundary layer stability and its effect on the occurrence of reattachment should be studied in detail to gain insight into the reasons for the presence of a hysteresis loop in stall at these Reynolds numbers. / Ph. D.

Laser Doppler velocimetry, a non-intrusive flow measurement technique, was employed to experimentally investigate two-dimensional airfoil-vortex interaction. Vortices were generated by sinusoidally oscillating a NACA 0012 airfoil about its quarter-chord at a reduced frequency of k = 2.05 and an amplitude of ±10° angle of attack. The target airfoil, a NACA 63₂A015, was immersed in the wake, two chord lengths downstream of the vortex generators trailing edge. Phase-averaged velocity measurements of the flow around the target airfoil were made with the airfoil at angles of attack of α = 0° and α = 10°. A close encounter with a counterclockwise rotating vortex was observed for both angles of attack, and a head-on collision, which split the counterclockwise rotating vortex in two, was observed for α = 10°. Vorticity fields were constructed from the velocity measurements and the circulation of the vortex was evaluated throughout the interaction. The surface pressure fluctuations on the airfoil were determined by substituting the measured velocities into the Navier-Stokes equations and numerically integrating the resulting pressure gradients. Furthermore, an extensive investigation of the undisturbed wake of the oscillating airfoil was performed in order to determine the effect of oscillation frequency and amplitude on the wake development. / Ph. D.

A new flying wing design has been developed at the Unmanned Systems Lab (USL) at Virginia Tech to serve delivery and remote sensing applications in the developing world. The fully autonomous unmanned aerial vehicle (UAV), named EcoSoar, was designed with the goal of creating a business opportunity for local entrepreneurs in low-resource communities. The system was developed in such a way that local fabrication, operation, and maintenance of the aircraft are all possible. In order to present a competitive financial model for sustained drone services, EcoSoar is made with reliable low-cost materials and electronics. This paper lays out the rapid prototyping and flight experiment efforts that went into polishing the design, test results from an EcoSoar centered drone workshop in Kasungu, Malawi, and finally a range optimization study with flight test validation. / Master of Science / A new humanitarian drone has been developed at the Unmanned Systems Lab (USL) at Virginia Tech. The unmanned aerial vehicle (UAV), named EcoSoar, was designed with the goal of creating a business opportunity for local entrepreneurs in low-resource communities. In order to be a viable solution in the developing world EcoSoar utilizes customizable 3D-printed parts and wings made from cheap materials like posterboard and packing tape. In addition, tools for building the drone have been developed in such a way that anyone can learn to construct and operate EcoSoar regardless of experience. This paper lays out the engineering efforts that went into the design, lessons learned from an EcoSoar-centered workshop in Kasungu, Malawi, and finally offers an upgraded design.

Delaying stall is always an attractive option in the aerospace industry. The major benefit of delaying stall is increased lift during takeoff and landings as well as during high angle of attack situations. Devices, such as fluidic oscillators, can be integrated into wing flaps to help delay the occurrence of stall by adding energized air to the airflow on the upper surface of the wing flap. The energized air from the oscillator allows the airflow to remain attached to the upper surface of the wing flap. The fluidic oscillator being integrated in this thesis is an active flow control device (AFC). One common method for integrating any device into a wing flap is to remove a section of the flap and mechanically secure the device. A current trend in the aerospace industry is the increased use of fiber-reinforced composites to replace traditional metal components on aircraft. The traditional methods of device integration cause additional complications when applied to composite components as compared to metal components. This thesis proposes an alternative method for integration of the AFC devices, which occurs before the fabrication of wing flaps is completed and they are attached to the aircraft wing. Seven design concepts are created to reduce the complications from using current methods of integration on composite wing flaps. The concepts are based on four design requirements: aerodynamics, manufacturing, maintenance, and structure. Four of the design concepts created are external designs, which place the AFC on the exterior surface of the wing flap in two types of grooved channels. The other three designs place the AFC inside the wing flap skin and are categorized as internal designs. In order for the air exiting the AFC to reach the upper surface of the wing flap, slots are created in the wing flap skin for the internal designs. Within each of the seven design concepts two design variants are created based on foam or ribbed core types. Prototypes were created for all of the external design AFC devices and the side inserted AFC and retaining pieces. Wing flap prototypes were created for the rounded groove straight AFC design, the semi-circular groove with straight AFC, and the side inserted AFC designs. The wing flaps were created using the VARTM process with a vertical layup for the external designs. The rounded groove and semi-circular groove prototypes each went through three generations of prototypes until an acceptable wing flap was created. The side inserted design utilized the lessons learned through each generation of the external design prototypes eliminating the need for multiple generations. The lessons learned through the prototyping process helped refine the designs and determine the ease of manufacturing to be used in the design evaluation. The evaluation of the designs is based on the four design requirements stated above. The assessment of the designs uses two levels of evaluation matrices to determine the most fitting design concept. As a result of the evaluation, all four of the external designs and one of the internal designs are eliminated. The two remaining internal designs' foam core and ribbed variants are compared to establish the final design selection. The vertically inserted AFC foam core design is the most fitting design concept for the integration of an AFC device into a composite wing flap.

No description available.

Multidisciplinary design optimization (MDO) requires many designs to be evaluated while searching for an optimum. As a result, the calculations done to evaluate the designs must be quick and simple to have a reasonable turn-around time. This makes aerodynamic calculations in the transonic regime difficult. Running computational fluid dynamics (CFD) calculations within the MDO code would be too computationally expensive. Instead, CFD is used outside the MDO to find two-dimensional aerodynamic properties of a chosen airfoil shape, BACJ, at a number of points over a range of thickness-to-chord ratios, free-stream Mach numbers, and lift coefficients. These points are used to generate surrogate models which can be used for the two-dimensional aerodynamic calculations required by the MDO computational design environment. Strip theory is used to relate these two-dimensional results to the three-dimensional wing. Models are developed for the center of pressure location, the lift curve slope, the wave drag, and the maximum allowable lift coefficient before buffet. These models have good agreement with the original CFD results for the airfoil. The models are integrated into the aerodynamic and aeroelastic sections of the MDO code. / Master of Science