With the pursuit of more advanced and environmentally-friendly technologies of today’s society, the airline industry has been pushed further to investigate solutions that will reduce airport noise and congestion, cut down on emissions, and improve the overall performance of aircraft. These items directly influence airport size (runway length), flight patterns in the community surrounding the airport, cruise speed, and many other aircraft design considerations which are setting the requirements for next generation aircraft. Leading the research in this movement is NASA, which has set specific goals for the next generation regional airliners and has categorized the designs that meet the criteria as Cruise Efficient Short Takeoff and Land (CESTOL) aircraft. With circulation control (CC) technology addressing most of the next generation requirements listed above, it has recently been gaining more interest, thus the basis of this research. CC is an active flow control method that uses a thin sheet of high momentum jet flow ejected over a curved trailing edge surface and in turn utilizes Coanda effect to increase the airfoil’s circulation, augmenting lift, drag, and pitching moment. The technology has been around for more than 75 years, but is now gaining more momentum for further development due to its significant payoffs in both performance and system complexity. The goal of this research was to explore the design of the CC flap shape and how it influences the local flow field of the system, in attempt to improve the performance of existing CC flap configurations and provide insight into the aerodynamic characteristics of the geometric parameters that make up the CC flap. Multiple dual radius flaps and alternative flap geometry, prescribed radius, flaps were developed by varying specific flap parameters from a baseline dual radius flap configuration that had been previously developed and researched. The aerodynamics of the various flap geometries were analyzed at three different flight conditions using two-dimensional CFD. The flight conditions examined include two low airspeed cases with blown flaps at 60° and 90° of deflection, and a transonic cruise case with no blowing and 0° of flap deflection. Results showed that the shorter flaps of both flap configurations augmented greater lift for the low airspeed cases, with the dual radius flaps producing more lift than the corresponding length prescribed radius. The large lift generation of these flaps was accompanied by significant drag and negative pitching moments. The incremental lift per drag and moment produced was best achieved by the longer flap lengths, with the prescribed radius flaps out-performing each corresponding dual radius. Longer flap configurations also upheld the better cruise performance with the least amount of low airspeed flow, drag, and required angle of attack for a given cruise lift coefficient. The prescribed radius flaps also presented a favorable trait of keeping a more continuous skin friction distribution over the flap when the flaps were deflected, where all dual radius configurations experienced a distinct fluctuation at the location where the surface curvature changes between its two radii. The prescribed radius flaps displayed a similar behavior when the flaps were not deflected, during the cruise conditions analyzed. Performance trends for the different flap configurations, at all three flight conditions, are presented at the end of each respective section to provide guidance into the design of CC geometry. The results of the presented research show promise in modifying geometric surface parameters to yield improved aerodynamics and performance.

This work presents the experimental study of the flow generated in the wakes of three three-dimensional bumps in the Virginia Polytechnic Institute and State University Boundary Layer Wind Tunnel. The three bumps examined are named bump 1, small bump 3, and large bump 3, and are the same test cases studied by Byun et al. (2004) and Ma and Simpson (2004) with a LDV system and a quad-wire hot-wire probe, respectively. Various experimental methods are used in this work: For measuring the mean velocity component in the planes examined, a seven-hole pressure probe is used with the data reduction algorithm developed by Johansen et al. (2001). A sixteen-hole pressure rake is used for boundary layer data on the sidewalls and ceiling of the test section and a Pitot-static probe is used to obtain mean velocity magnitude in the centerline of the test section. Specific techniques are developed to minimize the uncertainties due to the apparatus used, and an uncertainty analysis is used to confirm the efficiency of these techniques. Measurements in the wake of bump 1 reveal a strong streamwise vorticity creating large amounts of high moment fluid entrained close to the wall. In the wake of small bump 3, the amount of high momentum fluid entrained close to the wall is small as well as the streamwise vorticity. The flow in the wake of large bump 3 incorporate the characteristics of the two previous bumps by having a relatively large entrainment of high momentum fluid close to the wall and a low generation of streamwise vorticity. In the wakes of the three bumps, a pair of counter rotating vortices is created. The influence of large bump 3 on the incoming flow-field is found to be significant and induces an increase of the boundary layer thickness. By comparing LDV data and quad-wire hot-wire data with seven-hole probe data in the wakes of the bumps at the same locations, it is shown that uncertainties defined for a quasi-steady, non-turbulent flow-field without velocity gradient are bad indicators of the magnitude of the uncertainties in a more complex flow-field. A theoretical framework is discussed to understand the effects of the velocity gradient and of turbulence on the pressures measured by the seven-hole probe. In this fashion, a model is proposed and validated to explain these effects. It is observed that the main contribution to the uncertainties in seven-hole probe measurements due to the velocity gradient and to the turbulence comes from the velocity gradient. To correct for the effects of the velocity gradient on seven-hole probe measurements in an unknown flow-field, a technique is proposed. Using an estimation of the velocity gradient calculated from the seven-hole probe, the proposed model could be used to re-evaluate non-dimensional pressure coefficients used in the data reduction algorithm therefore correcting for the effects of the velocity gradient on seven-hole probe measurements. / Master of Science

A drag study was conducted on an oblate ellipsoid body in the Virginia Tech Stability Wind Tunnel. Two-dimensional wake surveys were taken with a seven-hole probe and an integral momentum method was applied to the results to calculate the drag on the body. Several different model configurations were tested; these included the model oriented at a 0° and 10° angle of attack with respect to the oncoming flow. For both angles, the model was tested with and without flow trip strips. At the 0° angle of attack orientation, data were taken at a speed of 44 m/s. Data with the model at a 10° angle of attack were taken at 44 m/s and 16 m/s. The high speed flow corresponded to a length-based Reynolds number of about 4.3 million; the low speed flow gave a Reynolds number of about 1.6 million. The results indicated that the length-squared drag coefficients ranged from around 0.0026 for the 0° angle of attack test cases and 0.0035 for the 10° angle of attack test cases. The 10° angle of attack cases had higher drag due to the increase in the frontal profile area of the model and the addition of induced drag. The flow trip strips appeared to have a tiny effect on the drag; a slight increase in drag coefficient was seen by their application but it was not outside of the uncertainty in the calculation. At the lower speed, uncertainties in the calculation were so high that the drag results could not be considered with much confidence, but the drag coefficient did decrease from the higher Reynolds number cases. Uncertainty in the drag calculations derived primarily from spatial fluctuations of the mean velocity and total pressure in the wake profile; uncertainty was estimated to be about 16% or less for the 44 m/s test cases. / Master of Science

Wall shear stress is one of the most essential scaling parameters used in fluid dynamics. It is significant because it helps us compare results in different experimental studies. The accurate measurements of wall shear stress will be instrumental in improving the existing empirical models and validating CFD models. Wall shear stress is also vital in improving fuel efficiency, heat transfer efficiency, and aerodynamic efficiency in real-world applications. This work discusses the design and implementation of a prototype floating element balance — a direct method of wall shear measurement. The direct measurement methods are robust and can significantly improve the validity of experimentation when perfected. In this work, a prototype floating element balance is designed and developed to estimate the wall shear stress in a smooth wall pilot facility to assess its feasibility for large-scale development. The floating element balance utilizes a strain gauge to estimate the wall shear stress. The preliminary tests show promising results, revealing potential design improvements. A strain measurement study is conducted to investigate the force-strain relationship and the reliability of the balance, which highlights the long-term stability and consistency in the strain measurement. However, further investigations are required into the drift response of the floating element balance. The strain measurements are also employed to calibrate the balance using a linear curve fit with a coefficient of determination of R^2 = 0.99, indicating a satisfactory linear estimation. / Master of Science / Drag is a phenomenon that occurs when the surface of a solid body interacts with a fluid. In fluid mechanics, there are two fundamental types of drag forces: pressure and skin friction. Pressure drag occurs due to the shape of the body, creating a pressure difference across the body, while skin friction drag arises due to the dominant viscous nature of the fluid. Understanding these forces is vital in improving the aerodynamic efficiency of various devices. Also, these forces play an essential role in the fuel efficiency and performance of the vehicles. The measurement of pressure drag is relatively straightforward compared to the skin friction drag. However, the measurement of skin friction drag can pose a challenge due to its smaller magnitude than the pressure drag. The importance of skin friction is due to its physical properties, which allow us to compare different experimental results and understand details about the turbulence in the flow. Also, accurate information on skin friction would improve existing relationships in fluid mechanics, and this information is also utilized to validate mathematical models in fluid mechanics. This work presents the design and implementation of a prototype used to estimate skin friction in a smooth wall facility using a novel and robust measurement known as floating element skin friction balance. Preliminary tests are conducted to assess the viability of the floating element balance for a large-scale development, which shows promising results while underlining some inherent limitations in the design and performance.

Grid fins (lattice fins) are used as a lifting and control surface for highly manoeuvrable missiles. Grid fins also find their applications for air-launched submunitions. The main advantages are its low hinge moment requirement and good high angle of attack performance characteristics. Two dimensional analysis has been carried out using linear and shock-expansion theo- ries. The results indicate that above certain depth-to-height ratio, (called critical depth-to-height ratio,) the local normal force becomes negative due to shock reflection from the opposite side. Hence, depth (chord) for grid fin cell should not exceed a critical value. A prediction method has been developed for the estimation of aerodynamic character- istics of grid fin-body combinations at supersonic Mach numbers based on shock-expansion theory. Body upwash theory has been used for the effect of body; method of images has been used for carry-over forces onto the body. Empirical relation has been used for the modelling of separated body vortices and their effect on the leeward side fins. The method has been validated with experimental results for three configurations. The comparison is good for individual fin characteristics as well as overall characteristics for all the cases at higher supersonic Mach numbers. For lower supersonic Mach numbers at higher angles of attack, the prediction deviates from experiment. The reason for the deviation is due to shock detachment and shock reflection from opposite side, which is not modelled in the present method. Vortex lattice method has been used for prediction of linear aerodynamic character- istics of grid-fins at subsonic Mach numbers. Empirical relation based on trends from available experimental data has been used for the non-linear effect. The method has been validated with experimental results for several configurations without and with control surface deflections. The predicted aerodynamic characteristics compare well with experimental results for all the cases and the difference is within 15%. Based on the subsonic and supersonic analytical methods, a prediction code for the aerodynamic analysis of configuration with grid fins has been developed. Flow field computations inside isolated cells have been carried out using CFD code, PARAS-3D. Effects of depth-to-height ratio, web thickness, web leading edge angle and cell width-to-height ratio have been studied. Increase in thickness reduces the critical depth and increases the normal force. This increment in normal force is due to shock wave formation at the expansion side and its interaction with the opposite side. Effect of cell cross sectional shape has been studied using inviscid computation over isolated cells. Square, right triangular, equilateral triangular and hexagonal cross sections have been considered for this study. The normal force for square cell at zero roll is higher compared to 45 deg roll (diamond shape). Triangular cells show large variation in normal force with roll orientation due to large variation in projected area with roll angle. To compare the characteristics of different cross sectional cells, the normal force is normalised with respect to total internal web area. The comparison shows that the hexagonal cell gives maximum normal force and right triangular cell gives the minimum. Packaging efficiency of different cross sections is analysed by normalising the normal force with frontal area. The results show that triangular cells are preferred for packaging efficiency. Viscous flow computations over complete configuration have been carried out using FLUENT. GAMBIT has been used for geometry definition and grid generation. Hexahedral finite volumes are used to generate the grids including the nose region. Flow computations have been carried out at supersonic Mach numbers. To reduce the compu- tational time, Flow computations upto 0.5 calibre ahead of grid fin have been carried out with body-alone configuration. Flow over the fin-body section has been computed sep- arately taking the inlet pressure condition from the body-alone computed results. This procedure has reduced the grid size to around 1/5th and the computations converged faster due to imposition of converged solution at the pressure inlet. The computed results on the body show that the Flow separation occurs on the lee- ward side of the body and formation of separated vortices. The comparison of pressure distribution on the body with experiment is good. Flow computations over the fin-body section have been carried out at different Mach numbers and angles of attack. The computed normal force coefficient on the horizontal fin compares well with experimental data. Computations with fin deflection of -15 deg have also been carried out and the computed results are within 10% of the experimental data. Flow computations over another grid fin configuration have been carried out at dif- ferent roll angles. The comparison of individual fin force and overall normal force and pitching moment coefficients with experiment is good. The comparison demonstrates the capability of prediction methods as well as CFD in analysing aerodynamic performance of grid fin configurations.

No description available.

<p> This study deals with applied research in the field of environmentai noise problems, specifically the measuring of noise patterns near Mount Hope Airport originating from subsonic jet aircraft using the present runway facilities. Based on actual measurements the results have been analyzed and reduced to simple contour lines. </p> <p> An attempt has been made in this study to relate the concept of community noise in the vicinity of the airport to specific runway configurations, traffic density and patterns, and to provide a comparison between the noise levels due to the existing operations and those which may result due to the proposed expansion. Versatile computer programs have been developed in this study to simulate an airport model, compute and construct the noise contours for any combination of design requirements such as runway orientation, flight procedure, type of aircraft, etc </p> / Thesis / Master of Engineering (MEngr)

An obvious reason for studying unsteady flows is the prediction of the effect of unsteady aerodynamic forces on a flight vehicle, since these effects tend to increase the likelihood of aeroelastic instabilities. This is a major concern in aerodynamic design of aircraft that operate in transonic regime, where the flows are characterised by the presence of adjacent regions of subsonic and supersonic flow, usually accompanied by weak shocks. It has been a common expectation that the numerical approach as an alternative to wind tunnel experiments would become more economical as computers became less expensive and more powerful. However even with all the expected future advances in computer technology, the cost of a numerical flutter analysis (computational aeroelasticity) for a transonic flight remains prohibitively high. Hence it is vitally important to develop an efficient, cheaper (in the sense of computational cost) and physically accurate flutter simulation tech nique which is capable of reproducing the data, which would otherwise be obtained from wind tunnel tests, at least to some acceptable engineering accuracy, and that it is essentially appropriate for industrial applications. This need motivated the present research work on exploring and developing efficient and physically accurate computational techniques for steady, unsteady and time-linearised calculations of transonic flows over an aircraft wing with moving shocks. This dissertation is subdivided into eight chapters, seven appendices and a bibliography listing all the reference materials used in the research work. The research work initially starts with a literature survey in unsteady transonic flow theory and calculations, in which emphasis is placed upon the developments in these areas in the last three decades. Chapter 3 presents the small disturbance theory for potential flows in the subsonic, transonic and supersonic regimes, including the required boundary conditions and shock jump conditions. The flow is assumed irrotational and inviscid, so that the equation of state, continuity equation and Bernoulli's equation formulated in Appendices A and B can be employed to formulate the governing fluid equation in terms of total velocity potential. Furthermore for transonic flow with free-stream Mach number close to unity, we show in Appendix C that the shocks that appear are weak enough to allow us to neglect the flow rotationality. The formulations are based on the main assumption that aerofoil slopes are everywhere small, and the flow quantities are small perturbations about their free-stream values. In Chapter 4, we developed an improved approximate factorisation algorithm that solves the two-dimensional steady subsonic small disturbance equation with nonreflecting far-field boundary conditions. The finite difference formulation for the improved algorithm is presented in Appendix D, with the description of the solver used for solving the system of difference equations described in Appendix E. The calculation of steady and unsteady nonlinear transonic flows over a realistic aerofoil are considered in Chapter 5. Numerical solution methods, based on the finite difference approach, for solving the two-dimensional steady and unsteady, general-frequency transonic small disturbance equations are presented, with the corresponding finite difference formulation described in Appendix F. The theories and solution methods for the time-linearised calculations, in the frequency and time domains, for the problem of unsteady transonic flow over a thin planar wing undergoing harmonic oscillation are presented in Chapters 6 and 7, respectively. The time-linearised calculations include the periodic shock motion via the shock jump correction procedure. This procedure corrects the solution values behind the shock, to accommodate the effect of shock motion, and consequently, the solution method will produce a more accurate time-linearised solution for supercritical flow. Appendix G presents the finite difference formulation of these time-linearised solution methods. The aim is to develop an efficient computational method for calculating oscillatory transonic aerodynamic quantities efficiently for use in flutter analyses of both two- and three-dimensional wings with lifting surfaces. Chapter 8 closes the dissertation with concluding remarks and future prospects on the current research work.

The current limits of supersonic parachute technology may constrain the ability to safely land future robotic assets on the surface of Mars. This constraint has led to a renewed interest in supersonic inflatable aerodynamic decelerator (IAD) technology, which offers performance advantages over the DGB parachute. Two supersonic IAD designs of interest include the isotensoid and tension cone, named for their respective formative structural theories. Although these concepts have been the subject of various tests and analyses in the 1960s, 1970s, and 2000s, significant work remains to advance supersonic IADs to a technology readiness level that will enable their use on future flight missions. In particular, a review of the literature revealed a deficiency in adequate aerodynamic and aeroelastic data for these two IAD configurations at transonic and subsonic speeds. The first portion of this research amended this deficiency by testing flexible IAD articles at relevant transonic and subsonic conditions. The data obtained from these tests showed that the tension cone has superior drag performance with respect to the isotensoid, but that the isotensoid may demonstrate more favorable aeroelastic qualities than the tension cone. Additionally, despite the best efforts in test article design, there remains ambiguity regarding the accuracy of the observed subscale behavior for flight scale IADs. Due to the expense and complexity of large-scale testing, computational fluid-structure interaction (FSI) analyses will play an increasingly significant role in qualifying flight scale IADs for mission readiness. The second portion of this research involved the verification and validation of finite element analysis (FEA) and computational fluid dynamic (CFD) codes for use within an FSI framework. These verification and validation exercises lend credence to subsequent coupled FSI analyses involving more complex geometries and models. The third portion of this research used this FSI framework to predict the static aeroelastic response of a tension cone IAD in supersonic flow. Computational models were constructed to mimic the wind tunnel test articles and flow conditions. Converged FSI responses computed for the tension cone agreed reasonably well with wind tunnel data when orthotropic material models were used and indicated that current material models may require unrealistic input parameters in order to recover realistic deformations. These FSI analyses are among the first results published that present an extensive comparison between FSI computational models and wind tunnel data for a supersonic IAD.

This thesis presents the results of steady, Reynolds-averaged Navier-Stokes (RANS) computations for jet flow emanating from a generic turbofan engine exhaust. All computations were performed with commercial solver FLUENT v6.2.16. Different turbulence models were evaluated. In addition to turbulence modeling issues, a parametric study was considered. Different modeling approaches for turbulent jet flows were explained in brief, with specific attention given to the Reynolds-averaged Navier-Stokes (RANS) method used for the calculations. First, a 2D ejector problem was solved to find out the most appropriate turbulence model and solver settings for the jet flow problem under consideration. Results of one equation Spalart-Allmaras, two-equation standart k-&amp / #949 / , realizable k-&amp / #949 / , k-&amp / #969 / and SST k-&amp / #969 / turbulence models were compared with the experimental data provided and also with the results of Yoder [21]. The results of SST k-&amp / #969 / and Spalart-Allmaras turbulence models show the best agreement with the experimental data. Discrepancy with the experimental data was observed at the initial growth region of the jet, but further downstream calculated results were closer to the measurements. Comparing the flow fields for these different turbulence models, it is seen that close to the onset of mixing section, turbulence dissipation was high for models other than SST k-&amp / #969 / and Spalart-Allmaras turbulence models. Higher levels of turbulent kinetic energy were present in the SST k-&amp / #969 / and Spalart-Allmaras turbulence models which yield better results compared to other turbulence models. The results of 2D ejector problem showed that turbulence model plays an important role to define the real physics of the problem. In the second study, analyses for a generic, subsonic, axisymmetric turbofan engine exhaust were performed. A grid sensitivity study with three different grid levels was done to determine grid dimensions of which solution does not change for the parametric study. Another turbulence model sensitivity study was performed for turbofan engine exhaust analysis to have a better understanding. In order to evaluate the results of different turbulence models, both turbulent and mean flow variables were compared. Even though turbulence models produced much different results for turbulent quantities, their effects on the mean flow field were not that much significant. For the parametric study, SST k-&amp / #969 / turbulence model was used. It is seen that boundary layer thickness effect becomes important in the jet flow close to the lips of the nozzles. At far downstream regions, it does not affect the flow field. For different turbulent intensities, no significant change occurred in both mean and turbulent flow fields.