**FULL TITLE** ASP Conference Series, Vol. **VOLUME**, **YEAR OF PUBLICATION** **NAMES OF EDITORS** Visualization of Scalar Adaptive Mesh Refinement Data

Abstract. Adaptive Mesh Refinement (AMR) is a highly effective computation method for simulations that span a large range of spatiotemporal scales, such as astrophysical simulations, which must accommodate ranges from interstellar to sub-planetary. Most mainstream visualization tools still lack support for AMR grids as a first class data type and AMR code teams use custom built applications for AMR visualization. The Department of Energy's (DOE's) Science Discovery through Advanced Computing (SciDAC) Visualization and Analytics Center for Enabling Technologies (VACET) is currently working on extending VisIt, which is an open source visualization tool that accommodates AMR as a first-class data type. These efforts will bridge the gap between generalpurpose visualization applications and highly specialized AMR visual analysis applications. Here, we give an overview of the state of the art in AMR scalar data visualization research

Visualization of Scalar Adaptive Mesh Refinement Data

Adaptive Mesh Refinement (AMR) is a highly effective computation method for simulations that span a large range of spatiotemporal scales, such as astrophysical simulations, which must accommodate ranges from interstellar to sub-planetary. Most mainstream visualization tools still lack support for AMR grids as a first class data type and AMR code teams use custom built applications for AMR visualization. The Department of Energy's (DOE's) Science Discovery through Advanced Computing (SciDAC) Visualization and Analytics Center for Enabling Technologies (VACET) is currently working on extending VisIt, which is an open source visualization tool that accommodates AMR as a first-class data type. These efforts will bridge the gap between general-purpose visualization applications and highly specialized AMR visual analysis applications. Here, we give an overview of the state of the art in AMR scalar data visualization research

Sensitivity Analysis and Optimization of Aerodynamic Configurations With Blend Surfaces

A novel (geometrical) parametrization procedure using solutions to a suitably chosen fourth order partial differential equation is used to define a class of airplane configurations. Inclusive in this definition are surface grids, volume grids, and grid sensitivity. The general airplane configuration has wing, fuselage, vertical tail and horizontal tail. The design variables are incorporated into the boundary conditions, and the solution is expressed as a Fourier series. The fuselage has circular cross section, and the radius is an algebraic function of four design parameters and an independent computational variable. Volume grids are obtained through an application of the Control Point Form method. A graphic interface software is developed which dynamically changes the surface of the airplane configuration with the change in input design variable. The software is made user friendly and is targeted towards the initial conceptual development of any aerodynamic configurations. Grid sensitivity with respect to surface design parameters and aerodynamic sensitivity coefficients based on potential flow is obtained using an Automatic Differentiation precompiler software tool ADIFOR. Aerodynamic shape optimization of the complete aircraft with twenty four design variables is performed. Unstructured and structured volume grids and Euler solutions are obtained with standard software to demonstrate the feasibility of the new surface definition

Sensitivity Analysis and Optimization of Aerodynamic Configurations with Blend Surfaces

A novel (geometrical) parametrization procedure using solutions to a suitably chosen fourth order partial differential equation is used to define a class of airplane configurations. Inclusive in this definition are surface grids, volume grids, and grid sensitivity. The general airplane configuration has wing, fuselage, vertical tail and horizontal tail. The design variables are incorporated into the boundary conditions, and the solution is expressed as a Fourier series. The fuselage has circular cross section, and the radius is an algebraic function of four design parameters and an independent computational variable. Volume grids are obtained through an application of the Control Point Form method. A graphic interface software is developed which dynamically changes the surface of the airplane configuration with the change in input design variable. The software is made user friendly and is targeted towards the initial conceptual development of any aerodynamic configurations. Grid sensitivity with respect to surface design parameters and aerodynamic sensitivity coefficients based on potential flow is obtained using an Automatic Differentiation precompiler software tool ADIFOR. Aerodynamic shape optimization of the complete aircraft with twenty four design variables is performed. Unstructured and structured volume grids and Euler solutions are obtained with standard software to demonstrate the feasibility of the new surface definition

Application of the multiblock method in computational aerodynamics

The main challenge in computational aerodynamics is to provide practical, credible, cost and schedule effective methods for routine design application and for full integration of these methods into the design cycle. Although advances in physical modelling and solution algorithms are continuing requirements of the aerospace industry, other more practical difficulties also impede the full realisation of the potential of existing methods. The contribution of this thesis is to examine and tackle several of these issues and to evaluate computational aerodynamics as a tool for engineering design and scientific enquiry. An advanced computational aerodynamics method is evaluated as an engineering tool for axisymmetric forebody and base flow problems. First the adaption of an existing two-dimensional flow solver to axisymmetric flow is described, then specific test cases are considered. The motivation for creating an axisymmetric flow solver is the considerable performance improvements compared to a fully three-dimensional method. The accuracy and robustness of the method are very good for forebody problems. For base flow problems accuracy and robustness are less satisfactory, although the performance of other prediction methods is also poorer for this more demanding problem. For both problem types the speed of the flow solver, the required computing resource and the time and effort necessary for pre- and post-processing are all satisfactory for routine calculation in an engineering environment. Shock reflection hysteresis and plume structure in a low density, axisymmetric highly underexpanded air jet is examined using a Navier-Stokes flow solver. This type of jet is found in a number of applications e.g. rocket exhausts and fuel injectors. The plume structure is complex, involving the interaction of several flow features, making this a demanding problem. Two types of shock reflection appear to occur in the plume, regular and Mach, depending on the jet pressure ratio. The existence of a dual solution domain where either type may occur has been predicted, in agreement with experiment where the same phenomenon has been observed for a nitrogen jet. There is a hysteresis in the shock reflection type; the reflection type observed in the dual solution domain depends on the time history of the plume development. A quasi-steady approach is employed in order to calculate the entire hysteresis loop. The results of the computational study are used to examine the structure of the plume, and are compared with experimental data where possible. Some flow features not initially recognised from experiment have been identified, notably curvature of the Mach disc, recirculation behind the Mach disc and the 'regular' reflection having Mach reflection characteristics. Included in the study is a review of the two dimensional shock reflection hysteresis problem to establish a theoretical background. The value of CFD as a tool for scientific investigation is clearly demonstrated by this study. The need for automation of the multiblock grid generation process is discussed. A new approach to automatically process a multiblock topology in order to prepare it for the grid generation process is described. The method is based on a cost function which attempts to model the objectives of the skilled grid generation software user who at present performs the task of block positioning and shaping in an interactive manner. A number of test cases are examined. It is also suggested that an existing unstructured mesh generation method could be adopted as an initial topology generation tool. Further work towards creating a fully automatic grid generation tool and extension into three dimensions are discussed. The parallel execution of an aerodynamic simulation code on a non-dedicated, heterogeneous cluster of workstations is examined. This type of facility is commonly available to CFD developers and users in academia, industry and government laboratories and is attractive in terms of cost for CFD simulations. However, practical considerations appear at present to discourage widespread adoption of this technology. The main obstacles to achieving an efficient, robust parallel CFD capability in a demanding multi-user environment are investigated. A static load-balancing method is described which takes account of varying processor speeds. A dynamic re-allocation method to account for varying processor loads has been developed. Use of proprietary software has facilitated the implementation of the method

Numerical investigation of film cooling fluid flow and heat transfer using large eddy simulations

Large eddy simulations of film cooling from discrete holes inclined at 35° with a feeding plenum chamber are performed at a density ratio of 2 and blowing ratios from 0.5 to 2.0 in order to gauge the suitability and performance of different hole shapes. Cylindrical holes at length to diameter ratios of 1.75 and 3.5 as well as shaped holes (laterally diffused and console holes) at a length to diameter ratio of 3.5 are simulated issuing into a laminar crossflow at a Reynolds number of approximately 16,000 based on freestream velocity and hole diameter. The domain extends 15 hole diameters downstream of a single coolant hole, and periodic boundary conditions on the lateral faces of the domain are used. The results are validated in terms of the flow field and surface adiabatic effectiveness to experiments for cylindrical hole cases. Horseshoe vortices, DSSN vortices, and hairpin vortices are resolved and isolated. Jetting is found to have significant effects on effectiveness in cylindrical hole cases (with less jetting at the exit plane and better cooling performance from the longer holes) and shaped hole cases (with a laterally split jetting action occurring around a central recirculation region). The performance of the shaped holes is dramatically better than the performance of the cylindrical holes in terms of surface adiabatic effectiveness, with the console holes performing slightly better than the laterally diffused holes. In terms of aerodynamic loss, the console and cylindrical hole far outperformed the laterally diffused hole

A numerical investigation of s-duct flows with boundary-layer ingestion

The essence of this paper is to report the computational research conducted to further NASA\u27s study of active flow control systems for a flush­-mounted inlet with significant boundary-­layer ingestion (BLI). In conjunction with a NASA-­sponsored research grant, the aim is to further accumulate knowledge and insight on the effectiveness of flow control devices in reducing circumferential distortion. Using Computational Fluid Dynamics (CFD), this study seeks to validate wind tunnel results recorded by the NASA Langley Research Center. After numerically reproducing experimental data, the future goal is use CFD to simulate the interaction between the inlet and turbofan stage. This study focused on NASA\u27s air jet flow control devices, specifically a 16-­jet version referred to as “configuration­-10”. The accuracy and validity of the numerical solutions was dependent on their ability to match the following experimental values: mass flow rate through the S-­duct, mass flow ratio, pressure measurements along the center line, and distortion coefficient at the aerodynamic interface plane. Three computational grids, containing 6.5, 30.5, and 36.9 million nodes, were constructed for the numerical simulations. The three grids varied in solid modeling techniques and grid element packing. Time limitations prevented the inclusion of results for the largest computational grid. Yet for the two smaller computational grids, the “Tenasi” flow solver was able to calculate valid numerical solutions for the “baseline” case (no jet flow) as well as low jet mass flow rate cases. High jet mass flow rate cases noticeably strayed from the experimental distortion measurements. Even though the numerical solutions did not replicate the experimental values for all mass flow ratio cases, this research significantly contributed to the knowledge and understanding of vorticity, flow characteristics, and distortion reduction for the BLI flush­mounted inlet. The CFD results provide visual representations of the changing flow characteristics as the jet mass flow rate is increased, in addition to flow interaction with intrusive experimental measuring devices. It is felt that better agreement with high­mass flow rate experimental cases can be calculated by numerically refining the jet boundary conditions and jet flow parameters. Overall, this paper reveals the steps taken in achieving early success in computationally verifying experimental data, and it discusses current actions being made towards further validating the numerical solutions