Distributed electric propulsion is a promising technology currently being considered for gen- eral aviation-class aircraft that has the potential to increase range and performance without sacrificing low-speed flight characteristics. However, the high-aspect ratio wings enabled by distributed electric propulsion make these designs more susceptible to adverse aeroe- lastic phenomena. This thesis describes the development of a gradient-based optimization framework for aircraft with distributed electric propulsion using structural and aeroelastic constraints. The governing equations for the coupled aeroelastic system form the basis of the static aeroelastic and flutter analysis. In this work, the Doublet-Lattice method is used to evaluate the aerodynamic forces exerted on the wing surface. In order to consider the impact of propeller-induced flow on aerodynamic loading, a one-way propeller-wing coupling is com- puted by superposition of the propeller induced velocity profile calculated using actuator disk theory and the wing flow field. The structural finite-element analysis is performed using the Toolkit for the Analysis of Composite Structures (TACS). The infinite-plate spline method is used to perform load and displacement transfer between the aerodynamic surface and the structural model. Instead of utilizing a conventional flutter analysis, the Jacobi-Davidson method is used to solve the governing eigenvalue problem without a reduction to the lowest structural modes, facilitating the evaluation of the gradient for design optimization. This framework is applied to different configurations with distributed electric propulsion to minimize structural weight subject to structural and aeroelastic constraints. The effect of flutter constraints, wing aspect ratio, and electric propeller quantity are compared through a series of design optimization studies. The results show that larger aspect ratio wings and more electric motors lead to heavier wings that are more susceptible to flutter. This framework can be used to develop lighter aircraft with distributed electric propulsion configuration that satisfy strength and flutter requirements.

Proefschrift - Technische Hoogeschool Delft. / "Stellingen": 4 p. inserted. Bibliography: p. [99]-100.

Thesis (M.S.)--Virginia Polytechnic Institute and State University. / Vita. Abstract. Includes bibliographical references (leaves 34-35). Also available via the Internet

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The need for more efficient and environmentally sustainable aircraft has been a rapidly increasing topic for research and development over the last few decades. Within this area of research, boundary layer ingestion (BLI) concepts have been developed which integrates the airframe and propulsion system of an aircraft. In turn, BLI increases the fuel efficiency and decreases emissions by reducing the overall drag and reenergizing the aircraft wake. However, the boundary layer flow of an airframe or duct can impose undesired flow conditions, such as swirl and pressure distortions, at the inlet of a jet engine. Therefore, efficient research and testing capabilities are essential to advance the development of these integrated systems. The StreamVane swirl distortion generator was developed by Virginia Tech to provide cost and time efficient ground testing methods for BLI research. StreamVanes are constructed of unique vane packs that are specifically tailored to generate a desired swirl distortion profile. To maximize efficiency, StreamVanes are additive manufactured which cause geometry limitations to the overall vane design. Due to these restrictions, as well as the complexity of the vane pack, unwanted dynamic responses and unsteady flows can be generated. In order to predict both of these phenomena before testing, two different computational methodologies were developed and investigated on a StreamVane and its airfoil parameters. First, a one-way fluid-structure interaction methodology was developed to predict flutter mconditions of the vanes within StreamVanes. The presented methodology includes steady and unsteady computational fluid dynamics (CFD) as well as linear structural and modal finite element analysis (FEA) simulations. A simplified StreamVane model was designed as a testcase for the methodology, and it was found that two unique vane shapes did not undergo flutter conditions at three different operating points. The results provided a linear analysis method to compute the aerodynamic damping, which gave insight on how different vane shapes respond dynamically. Secondly, a parameter study was conducted to predict the vortex shedding from the modified NACA 63-series airfoil profile used within StreamVane design. The effects of the airfoil turning angle and trailing edge thickness on the vortex shedding frequency were computationally predicted using the unsteady Reynolds averaged Navier-Stokes equations (URANS) and shear stress transport (SST) turbulence model. In turn, the shedding frequencies for each parameter were recorded, and more intuition was gained on the TE flow field in correspondence to different airfoil specifications. Overall, the two sets of methodologies and results can be used to efficiently reduce failure uncertainties in future StreamVane designs. / Master of Science / The need for more efficient and environmentally sustainable aircraft has been a rapidly increasing topic for research and development over the last few decades. Within this area of research, boundary layer ingestion (BLI) concepts have been developed to advance the fuel efficiency in future aircraft designs. However, unlike traditional tube and wing aircraft, BLI produces nonuniform flow at the engine inlet, reducing the performance and durability of jet engine components. Therefore, more efficient research and testing capabilities are essential to advance the development of BLI aircraft. The StreamVane swirl distortion generator was developed by Virginia Tech to provide cost and time efficient ground testing methods for BLI research. These devices can be secured upstream of a test engine, and their complex vane pack can produce the same nonuniform flow found at the inlet of BLI aircraft engines. To further increase efficiency, StreamVanes are additive manufactured which causes geometry limitations to the overall vane design. Due to these restrictions, as well as the complexity of the vane pack, unwanted dynamic responses and unsteady flows can be generated. In order to predict both of these phenomena before testing, two different computational methodologies were developed and investigated on a StreamVane and its airfoil parameters. The first methodology was developed to compute the fluid dynamics and structural response of a simplified StreamVane model at different operating conditions. The results provided insight on how different vanes react dynamically to the surrounding flow field. The second methodology included a parameter study to predict the frequencies generated from the StreamVane airfoils. With these frequencies, more intuition was gained on how the overall fluid-structure system would behave. Overall, both methodologies and results can be used to efficiently reduce failure uncertainties in future StreamVane designs.

With the advent of smart materials it is becoming possible to alter the structural characteristics of turbomachine airfoils. This change in structural characteristics can include, but is not limited to, changes in the shape (morphing) of the airfoil. Through changes in the airfoil shape, aerodynamic performance can be improved. Moreover, this technique has the potential to act as a flutter suppressant. In this investigation changes in the airfoil front camber while maintaining the airfoil thickness distribution are made to increase airfoil stability. The airfoil section is representative of current low aspect ratio fan blade tip sections. To assess the influence of the change in airfoil shape on stability the work-per-cycle was evaluated for torsion mode oscillations around the mid-chord at an inlet Mach number of 0.5 with an interblade phase angle of 180 degree Cchordal incidence angles of both 0 degree and 10 degree, and a reduced frequency of 0.4.

Neste trabalho é desenvolvido um sistema de controle ativo para supressão de flutter de uma asa utilizando-se sensores de pressão em pontos estratégicos de sua superfície. O flutter é um fenômeno aeroelástico que caracteriza um acoplamento instável entre estrutura flexível e escoamento aerodinâmico não estacionário. Quando a modificação da estrutura ou da aerodinâmica da asa não é viável, o uso de sistemas de controle passa a ser uma boa opção. Para o desenvolvimento do sistema de controle proposto, é primeiramente desenvolvido um modelo numérico de asa flexível. Com esse modelo numérico e a pressão na superfície da asa medida em certos pontos e realimentada ao sistema controlador, são determinadas correções no ângulo de uma superfície de controle no bordo de fuga. A tentativa de se utilizar um sistema de controle bem simples, com o uso de um único sensor de pressão, mostra a viabilidade de se implementar um sistema deste tipo em aeronaves reais. Esse sistema pode tornar-se uma alternativa aos desenvolvidos até então com o uso de acelerômetros, além de ser útil em sistemas onde se procura prever o estol e observar o comportamento da distribuição de pressão sobre a asa em vôo. / In this work, a wing flutter suppression active control system using pressure sensors in strategic points is developed. Flutter is an aeroelastic phenomenon characterized by an unstable coupling of a flexible structure and a non-stationary aerodynamic flow. When changes of the wing structure or of the aerodynamics are not viable, the use of automatic control systems becomes a good option. For the developing of the suggested control system, a numeric model of a finite flexible wing is firstly done. With this model and the pressure over the wing surface read in certain points and fedback to the control system, changes of the control surface angle on the trailing edge are determined. The attempt to use a simple control system, with a unique pressure sensor shows the viability of implanting this kind of system in real aircrafts. This system may become an alternative to those developed until now, using accelerometers. Yet, it could be useful, in systems where it is necessary to predict stall and observe the pressure load behavior over the wing in flight.

An experimental and computational study, motivated by the need to improve current understanding of blade flutter in turbomachinery and provide 3D test data for the validation of advanced computational methods for the prediction of this aeroelastic phenomenon, is presented. A new, low speed flutter test facility has been developed to facilitate a detailed investigation into the unsteady aerodynamic response of a turbine blade oscillating in a three dimensional bending mode. The facility employs an unusual configuration in which a single turbine blade is mounted in a profiled duct and harmonically driven. At some cost in terms of modelling a realistic turbomachinery configuration, this offers an important benefit of clearly defined boundary conditions, which has proved troublesome in previous work performed in oscillating cascade experiments. Detailed measurement of the unsteady blade surface pressure response is enabled through the use of externally mounted pressure transducers, and an examination of the techniques adopted and experimental error indicate a good level of accuracy and repeatability to be attained in the measurement of unsteady pressure. A detailed set of steady flow and unsteady pressure measurements, obtained from five spanwise sections of tappings between 10% and 90% span, are presented for a range of reduced frequency. The steady flow measurements demonstrate a predominant two-dimensional steady flow, whilst the blade surface unsteady pressure measurements reveal a consistent three dimensional behaviour of the unsteady aerodynamics. This is most especially evident in the measured amplitude of blade surface unsteady pressure which is largely insensitive to the local bending amplitude. An experimental assessment of linearity also indicates a linear behaviour of the unsteady aerodynamic response of the oscillating turbine blade. These measurements provide the first three dimensional test data of their kind, which may be exploited towards the validation of advanced flutter prediction methods. A three dimensional time-marching Euler method for the prediction of unsteady flows around oscillating turbomachinery blades is described along with the modifications required for simulation of the experimental test configuration. Computationalsolutions obtained from this method, which are the first to be supported by 3D test data, are observed to exhibit a consistently high level of agreement with the experimental test data. This clearly demonstrates the ability of the computational method to predict the relevant unsteady aerodynamic phenomenon and indicates the unsteady aerodynamic response to be largely governed by inviscid flow mechanisms. Additional solutions, obtained from a quasi-3D version of the computational method, highlight the strong three dimensional behaviour of the unsteady aerodynamics and demonstrate the apparent inadequacies of the conventional quasi-3D strip methodology. A further experimental investigation was performed in order to make a preliminary assessment of the previously unknown influence of tip leakage flow on the unsteady aerodynamic response of oscillating turbomachinery blades. This was achievedthrough the acquisition of a comprehensive set of steady flow and unsteady pressure measurements at three different settings of tip clearance. The steady flow measurements indicate a characteristic behaviour of the tip leakage flow throughout the range of tip clearance examined, thereby demonstrating that despite the unusual configuration, the test facility provides a suitable vehicle for the investigation undertaken. The unsteady pressure data show the blade surface unsteady pressure response between 10% and 90% span to be largely unaffected by the variation in tip clearance. Although close examination of the unsteady pressure measurements reveal subtle trends in the first harmonic pressure response at 90% span, which are observed to coincide with localised regions where the tip leakage flow has a discernible impact on the steady flow blade loading characteristic. Finally, some recommendations for further work are proposed

Thesis (M.S.)--Virginia Polytechnic Institute and State University, 1991. / Vita. Abstract. Includes bibliographical references (leaves 102-107). Also available via the Internet.

Thesis (Ph. D.)--Duke University, 2008.