In the past, 2D CFD analysis of Circulation Control technology have shown poor comparison with experimental results. In Circulation Control experiments, typical results show a relationship between lift coefficient, CL, vs blowing momentum coefficient, Cμ. CFD analysis tend to over-predict values of CL due to gridding issues and/or turbulence model selection. This thesis attempted to address both issues by performing Richardson’s Extrpolation method to determine an acceptable mesh size and by using FLUENT’s 2-equation turbulence models. The experimental results and CAD geometry were obtained from Georgia Tech Research Institute for comparison with the CFD analysis. The study showed that 3D CFD analysis of circulation control showed similar results of over-predicting CL, which can also be attributed to gridding issues and turbulence model selection. When compared to the experimental results, the k − ω turbulence model produced the lowest errors in CL of approximately 15-17%. The other turbulence models produced errors within 5% of k − ω. A fully unstructured volume mesh with prismatic cells on the surfaces was used as the grid. The CCW con- figuration was analyzed with and without wind tunnel walls present, which produced errors of 20% and 15% in CL, respectively, when compared to experimental results. Despite the large errors in CL, CFD was able to capture the trend of increasing CL as Cμ was increased. Results reported in this thesis can be further calibrated to allow the CFD model to be used as a predictive tool for other CCW applications.

The design process and construction of the initial modular ramjet attachment to the Cal Poly supersonic wind tunnel is presented. The design of a modular inlet, combustor, and nozzle are studied in depth with the intentions of testing in the modular ramjet. The efforts undertaken to characterize the Cal Poly supersonic wind tunnel and the individual component testing of this attachment are also discussed. The data gathered will be used as a base model for future expansion of the ramjet facility and eventual hot fire testing of the initial components. Modularity of the inlet, combustion chamber, and nozzle will allow for easier modification of the initial design and the designs ability to incorporate clear walls will allow for flow and combustion visualization once the performance of the hot flow ramjet is determined. The testing of the blank ramjet duct resulted in an error of less than 10% from predicted results. The duct was also tested with the modular inlet installed and resulted in between a 13-30% error based on the predicted results. Hot flow characteristics of the ramjet were not achieved, and the final cold flow test with the nozzle installed was a failure due to improper configuration of the nozzle. The errors associated with this testing can largely be placed on the poor performance of the Cal Poly supersonic wind tunnel and the alterations made to the testing in an attempt to accommodate these flaws. The final tests were halted for safety concerns and could continue after a thorough safety review.

This report describes the development and application of a validated Computational Fluid Dynamics (CFD) modelling approach for internal cooling passages in rotating turbomachinery. A CFD Modelling approach and accompanying assumptions are tuned and validated against academically available experimental results for various serpentine passages. Criteria of the CFD modelling approach selected for investigation into advanced internal cooling flows include accuracy, robustness, industry familiarity, and computational cost. Experimental data from NASA HOST (HOt Section Technology), Texas A&M, and University of Manchester tests are compared to RANS CFD results generated using Fluent v14.5 in order to benchmark a CFD modelling approach. Capability of various turbulence models in the representation of cooling physics is evaluated against experimental data. Model sensitivity to boundary conditions and mesh density is also evaluated. The development of a validated computational model of internal turbine cooling channels with bounded error allows for the identification of particular shortcomings of heat transfer correlations and provides a baseline for future CFD based exploration of internal turbine cooling concepts.

The Goldschmied Propulsor is a concept that was introduced in mid 1950's by Fabio Goldschmied. The concept combines boundary layer suction and boundary layer ingestion technologies to reduce drag and increase propulsor efficiency. The most recent testing, done in 1982, left questions concerning the validity of the results. To answer these questions a 38.5in Goldschmied Propulsor was constructed and tested in Cal Poly's 3x4ft wind tunnel. The focus of their wind tunnel investigation was to replicate Goldschmied's original testing and increase the knowledge base on the subject. The goal of this research was to create a computational fluid dynamics (CFD) model to help visualize the flow phenomenon and see how well CFD was able to replicate Cal Poly’s wind tunnel results. CFD cases were run to get a comparison of the computational model and the wind tunnel results. For the straight tunnel geometry for the 0.385” slot and cusp A we found a body, pressure and friction drag, fan off CD of 0.0526 and a fan on at 500 Pascals with a CD of 0.0545. This is similar to the wind tunnel results but because of large errors in measuring overall drag we are not able to directly compare to the wind tunnel results. Overall we see that the trends match, mainly that the fan does not decrease the total pressure drag. This was a result of poor geometry and high fan speeds needed for attachment. The tested geometry is less than ideal and has a long way to go before it is of a shape that would have the potential to reduce the pressure drag as much as Goldschmied claimed. Future efforts should be put forth optimizing the aft body to reduce the low pressure in front of the slot and improving aft entrance of the slot to allow for a smoother flow.

Analysis of the aerodynamic effects of a propeller flowfield on bodies downstream of the propeller is a complex task. These interaction effects can have serious repercussions for many aspects of the vehicle, including drag changes resulting in larger power requirements, stability changes resulting in adjustments to stabilizer sizing, and lift changes requiring wing planform adjustments. Historically it has been difficult to accurately account for these effects at any stage during the design process. More recently methods using Euler solvers have been developed that capture interference effects well, although they don't provide an ideal tool for early stages of aircraft design, due to computational cost and the time and expense of setting up complex volume grids. This research proposes a method to fill the void of an interference model useful to the aircraft conceptual and preliminary designer. The proposed method combines a flexible and adaptable tool already familiar to the conceptual designer in the aerodynamic panel code, with a pseudo-steady slipstream model wherein rotational effects are discretized onto vortex particle point elements. The method maintains a freedom from volume grids that are so often necessary in the existing interference models. In addition to the lack of a volume grid, the relative computational simplicity allows the aircraft designer the freedom to rapidly test radically different configurations, including more unconventional designs like the channel wing, thereby providing a much broader design space than otherwise possible. Throughout the course of the research, verification and validation studies were conducted to ensure the most accurate model possible was being applied. Once the vortex particle scheme had been verified, and the ability to model an actuator disk with vortex particles had been validated, the overall product was compared against propeller-wing wind tunnel results conducted specifically as benchmarks for numerical methods. The method discussed in this work provides a glimpse into the possibility of pseudo-steady interference modeling using vortex particles. A great groundwork has been laid that already provides reasonable results, and many areas of interest have been discovered where future work could improve the method further. The current state of the method is demonstrated through simulations of several configurations including a wing and nacelle and a channel wing.

The use of Computational Fluid Dynamics (CFD) tools throughout the engineering industry has become standard. Simulations are used during nearly all steps throughout the life cycle of products including design, production, and testing. Due to their wide range of use, industrial CFD codes are becoming more flexible and easier to use. These commercial codes require robustness, reliability, and efficiency. Consequently, linear eddy viscosity models (LEVM) are used to model turbulence for an increasing number of flow types. LEVM such as k-ε and k-ω provide modeling with little loss of computational efficiency and have proven to be robust. The LEVM that are most common in CFD tools, however, are not adequate for accurate prediction of complex flows. This includes flows with high streamline curvature, strong rotation and separation regions. Unfortunately, due to their ease of use in the commercial CFD tools, the models are used frequently for complex flows. Modifications have been made to LEVM such as k-ε in order to improve modeling, but generally, the modifications have only improved modeling of less complex flows. More advanced LEVM models have been developed using elliptic relaxation equations to help resolve these issues. The ν2-f model was developed to better capture flow physics for complex flows while being applicable to general flows. It is generally considered one of the most accurate LEVMs. It does, however, have issues with stability and robustness. Several improvements have been proposed. One of the most notable is its reformulation into the ζ-f model which offers several improvements while maintaining accurate flow prediction. The model improvement is still limited by being a LEVM. While models, such as differential Reynolds stress models, do exist which are able to capture relevant flow physics in complex flows, modeling difficulties make them impractical for use in a commercial CFD code. Algebraic Reynolds stress models have attempted to bridge this gap with varying levels of success. The models express the Reynolds stress tensor as a function of different higher level tensors. This is the same process used to derive non-linear eddy-viscosity models which add extra high-order terms to the Boussinesq approximation. According to Kassinos and Reynolds, however, this technique is fundamentally flawed. These models fail to capture all relevant information about the turbulence structure. The Reynolds stresses capture information regarding the turbulent componentiality, i.e. velocity components of turbulence. The dimensionality, which carries information regarding the direction of turbulent eddies, is not modeled, however. Kassinos and Reynolds constructed a structure-based model which attempts to capture turbulent componentiality and dimensionality by expressing the Reynolds stress tensor as a function of one-point turbulence structure tensors. Their original model introduced hypothetical turbulence eddies which could be averaged and then used to relate the eddy-axis transport equation to the proper structure tensors. The ideas behind this model were adapted into several different models including the R-D model and the Q-model. These formulations were able to accurately capture the flow physics for many complex flow types especially those with mean rotation. These resulting models, however, were overly complicated for application in commercial CFD codes. These structure-based models later resulted in the development of the algebraic structure based model (ASBM). The ASBM was developed in order to ensure computational efficiency while capturing relevant turbulence physics. The ASBM uses an algebraic model for the eddy statistics which is constructed from the local mean deformation and two turbulent scales. The original turbulent scales used were the turbulent kinetic energy and the large scale vorticity. Although the model was calibrated specifically for use with the turbulent kinetic energy and large scale vorticity transport equations, the algebraic model can be used in conjunction with any scalar transport equations as long as the field distribution of turbulent kinetic energy and turbulence time scale can be obtained. Based on its formulation, the ASBM, used in combination with any scalar transport equations, should be applicable to most commercial CFD codes. The objective of this work was to implement the ζ-f model and ASBM, coupled with k-ε and v2-f, in the commercial CFD solver FLUENT and validate its performance for canonical turbulent flows including a subsonic turbulent flat-plate, S3H4 2D hill, and backward-facing step. Each turbulent flow was evaluated using various turbulence models including Spalart-Allmaras, k-ε, k-ω, k-ω-SST, v2-f, ζ-f and two ASBM formulations and compared against experimental results. The ζ-f model produced improved results for both the flat plate and backward facing step as compared to all two-equation or less turbulence models and showed similar predictive capabilities to the v2-f model. It had difficulties predicting attached flow past the S3H4 2D hill just as the v2-f model. This, however, was expected due to its basis on the v2-f model. The model was also more stable than the v2-f model during calculation of the turbulent flat plate but showed no improvement in robustness for the more complex backward facing step. The semicoupled (linear eddy viscosity model based) v2-f-ASBM’s predictive capabilities were comparable to the two equation models for the turbulent flat plate case. It performed surprisingly well for the backward facing step and matched the experimental data within experimental uncertainty. The model did, however, have problems predicting the S3H4 2D hill just as the with the v2-f model.

The utility of computational fluid dynamics (CFD) for solving problems of engineering interest has experienced rapid growth due to the improvements in both memory capacity and processing speed of computers. While the capability now exists for the solution of the Navier-Stokes equations about complex and complete aircraft configurations, the bottleneck within the process is the time consuming task of properly generating a mesh that can accurately solve the governing partial differential equations (PDEs). This thesis explored two numerical techniques that attempt to circumvent the difficulty associated with the meshing process by solving a simplified form of the continuity equation within a meshless framework. The continuity equation reduces to the full potential equation by assuming irrotational flow. It is a nonlinear PDE that can describe flows for a wide spectrum of Mach numbers that do not exhibit discontinuities. It may not be an adequate model for the detailed analysis of a complex flowfield since viscous effects are not captured by this equation, but it is an appealing alternative for the aircraft designer because it can provide a quick and simple to implement estimate of the aerodynamic characteristics during the conceptual design phase. The two meshless methods explored in this thesis are the Dual Reciprocity Method (DRM) and the Generalized Finite Difference Method (GFD). The Dual Reciprocity Method was shown to have the capability to solve for the two-dimensional subcritical compressible flow over a Circular Cylinder and the non-lifting flow for a NACA 0012 airfoil. Unfortunately these solutions were obtained with the requirement of a priori knowledge of the solution to tune a parameter necessary for proper convergence of the algorithm. Due to the shortcomings of applying the Dual Reciprocity Method, the Generalized Finite Difference Method was also investigated. The GFD method solves a PDE in differential form and can be thought of as a meshless form of a standard finite difference scheme. This method proved to be an accurate and general technique for solving the previously mentioned cases along with the lifting flow about a NACA 0012 airfoil. It was also demonstrated that the GFD method could be formulated to discretize the full potential equation with second order accuracy. Both solution methods offer their own set of unique advantages and challenges, but it was determined that the GFD Method possessed the flexibility necessary for a meshless technique to become a viable aerodynamic design tool.

Recent unexpected experimental observations of the emptying of inverted bottles with perforations has generated interest in modeling and simulation of this phenomenon. It was observed that as a perforation, i.e., a small hole at the "top" of the inverted bottle, is added and enlarged, the overall emptying time first increases to a maximum value and then decreases until it reaches a lower limit. The change in emptying time is associated with a transition from jetting, where only water exits the neck, to glugging, a competition between air and water flows at the neck of the bottle. This paper develops analytical and numerical models to predict emptying time and liquid height as a function of time which capture the jetting-to-glugging transition. When qualitatively compared to experimental data using a bottle with neck diameters of 12.7 mm, 25.4 mm, and 38.1 mm and bottle diameter of approximately 355 mm (equating to several hundred to several thousand seconds to drain) a favorable agreement is observed. These models attempt to explain the transition in terms of a competition between liquid and bubble velocities at the bottle neck and build on an existing model of glugging available in the literature. The paper also explores the first steps taken toward simulation of bottle emptying using a commercial CFD package (Fluent) to simulate draining for a smaller bottle of neck diameter 21.6 mm and bottle diameter of 62.2 mm. The Fluent simulations are used to further elucidate the jetting-to-glugging transition mechanism by simulating emptying with and without perforations. CFD results reported are limited to a few select large perforation diameters. Specifically, a 4 mm perforation taking 15 hours to simulate and 6 mm perforation taking 5 hours to simulate. Despite the lengthy simulation times, both capture only the approximate 2 seconds required to drain the bottle, but demonstrate the effect of the perforation on emptying time. Smaller perforations on the order of 1 mm, which would align with the experimentally determined maximum emptying time would require unfeasibly long simulations for present resources as dictated by required low Courant numbers. Future work with greater computational capability will further expand upon the simulations conducted in this work.

Fundamental intuition of aerodynamics begins with understanding steady flow, a time- independent flow state. A fluid region undergoing steady flow consists of constant properties such as pressure and velocity at different positions in the flow field. This time-independent principle is crucial for beginning a foundation of understanding aerodynamics; however, analyzing this state of flow was beyond the limit at my university's Fundamentals of Aerodynamics course. There was minimal education on time-dependent unsteady flow, which created a vacuum on my understanding of how flow can be analyzed with time. The purpose of writing this thesis is to create a framework for aspiring learners of aerodynamics to better comprehend unsteady flow, including myself. The basis for developing an understanding of unsteady flow is accomplished by analyzing the aerodynamics of a simple two-dimensional zero-thickness flat plate, using a numerical method called Discrete Vortex Method under steady and unsteady conditions. Constructing a numerical method for steady and unsteady flow requires a software to compute enormous quantities of linear equations, therefore a combination of numerous arguments, functions, and loops were developed on MATLAB written in the C/C++ languages. Results from the numerical methods will be compared with the experimental and theoretical results from Katz & Plotkin (2001). The Steady Discrete Vortex Method was a basis for calculating the circulation of the flat plate at varying angles of attack and freestream velocities. The Unsteady Discrete Vortex Method derived much of the self-induced calculations in the body-fixed coordinate system. At the same time, a time-stepping method was developed to calculate the coordinates as the flat plate and shed vortices translated from the origin of an additional frame of reference called the inertial coordinate system. A wake vortex is shed from the trailing-edge of the flat plate at each time step iv to model vorticity shed from a body in motion. The flat plate undergoes sudden acceleration and plunging maneuvers to demonstrate further effects of unsteady aerodynamic conditions. The results from the flat plate undergoing sudden acceleration with a Reynolds number of 68,435.8 was an increasing proportionality between the lift and circulation of the steady and unsteady case until reaching a constant trend as time increases, demonstrating the nature of low-speed flow reaching a steady state after a given period. The results from the flat plate undergoing plunging with a Reynolds number of 106,759.8 demonstrate a sinusoidal trend in the normal force experienced as the flat plate traverses in its sinusoidal plunging translation like that observed in the theoretical results. This thesis intends to expand on the understanding of unsteady aerodynamics by developing a numerical method that can alter its dependent factors to visualize the effects of changing specific parameters on pressure and force acting on the two-dimensional body.

The high-pressure shock tube facility (HiPER-STAR) at the University of Central Florida (UCF) is analyzed experimentally to demonstrate the practicality of hypersonic aero-optical testing in an impulse facility without the use of an expansion nozzle or acceleration tube. The investigation analyses driver gas blending with helium and hydrogen to raise the speed of sound ratio in an attempt to increase the Mach number for aero-optics testing. HiPER-STAR has a unique ability to withstand pressures up to 1000 atm and run in a double diaphragm configuration allowing for a significant pressure differential to be created between the driver and driven sections. Results from this study show that hydrogen and helium blending can drastically increase the maximum Mach number of HiPER-STAR; Mach numbers up to 15 were generated at a variety of altitudes. Experiment test time varied on shock velocity but was purely dependent on the arrival of the reflected shock wave to measurement locations. The aero-optics data that was collected and visually captured with a high-speed camera clearly shows beam aberration due to density gradients and a diminishing light intensity indicating that hypersonic aero-optical phenomenon can be captured reliably and repeatedly with a shock tube.