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The primary contribution of this research is to clarify the feasibility of a novel lightweight turbine blade with internal and external cooling, which is invented, aiming at drastic reduction in weight. With a considerably thinner airfoil, an extensive separation bubble is formed on the pressure side, and film cooling performance in such a flow field has to be investigated. Experimental results with a curved duct setup, which simulates the flow field around the proposed airfoil, show that a film cooling is still an effective measure of cooling even in the vastly separated region, and it behaves quite similarly to the conventional correlation, except for lower blowing ratios, where the thermal field is strongly affected by the intense recirculation flow. Comparisons between the experimental and numerical results verify that an affordable Reynolds-averaged Navier--Stokes simulation is useful to investigate the detailed physics of this flow field. With the numerical modeling, a cooling performance of the proposed blade under a typical engine operating condition is simulated, and the metal temperatures of the blade are also predicted with a fluid-solid conjugate calculation. The resultant thermal distribution in the airfoil suggests that the trailing edge portion is inevitably most critical in the temperature, and also a considerable thermal gradient across the blade is induced. Thermal profile, however, is partly recovered with some of the film coolant being bypassed from the pressure side to the suction side. [DOI: 10.1115/1.3144165]

Accurate and numerically efficient modeling of low-to-moderate Reynolds number nozzle flow expansions to vacuum can be difficult due to the presence of multiple flow length scales. Such simulations are important for the prediction of propulsive thrust as well as spacecra R contamination, both of which can he difficult to measure in ground-based facilities. To that end, conical nozzle flows were studied for Reynolds numbers of 1230 and 12,300 using the direct simulation Monte Carlo method, Navier-Stokes with velocity slip and temperature jump boundary conditions, and statistical and deterministic approaches to the solution of the Bhatnagar--Gross-Krook and ellipsoidal-statistical Bhatnagar--Gross--Krook equations. The deterministic and statistical solutions of the Bhatuagar--Gross--Krook equation were found to be in good agreement with the benchmark direct simulation Monte Carlo results. Statistical Bhatnagar--Gross--Krook and ellipsoidal-statistical Bhatnagar--Gross--Krook methods were also found to be more efficient methods than direct simulation Monte Carlo in the continuum and near-continuum regime, and more accurate than the Navier--Stokes equations in the portions of the flow with rarefaction, such as the boundary layer and the flow around the nozzle lip. DOI: 10.2514/1.J050208

The joint velocity-scalar filtered mass density function methodology is employed for large eddy simulation of Sandia National Laboratories' flame D. This is a turbulent piloted nonpremixed methane jet flame. In velocity--scalar filtered mass density function, the effects of the subgrid-scale chemical reaction and convection appear in closed forms. The modeled transport equation for the velocity-scalar filtered mass density function is solved by a hybrid finite difference/Monte Carlo scheme. For this flame, which exhibits little local extinction, a flamelet model is employed to relate the instantaneous composition to mixture fraction. The simulated results are assessed via comparison with laboratory data and show favorable agreements. DOI: 10.2514/1.J050154

The noise source distribution of coaxial jets with a diameter ratio of 1.6 and variable velocity ratio is investigated with a small-aperture microphone phased array. The array design enables differentiation of noise emitted by largescale and fine-scale turbulence structures, which have different directivities. The acoustic data are complemented by pitot surveys of the mean flow, which yield measurements of the primary and secondary cores of the jet. For zero velocity ratio (single-stream jet), the region near the nozzle emits strong high-frequency noise. Increasing the secondary-to-primary velocity ratio suppresses the near-nozzle noise and extends the location of the peak noise downstream, which increases moderately. The axial location of peak noise is approximately situated at the end of the primary core. The suppression of high-frequency noise is explained by the creation and elongation of the secondary core as the velocity ratio increases. The acoustic trends with velocity ratio are similar for small and large array observation angles from which large-scale and fine-scale turbulence noise, respectively, have been shown to radiate. However, the increase in peak noise is more pronounced for the large-scale noise. DOI: 10.2514/1.J050140

This paper discusses the effects of a thin flexible rectangular polymer film (fin) attached to the upper surface of a NACA0012 airfoil model on the natural low-frequency oscillation in poststall flows. The flexible fin passively manipulates flow structures, which leads to suppression of the low-frequency oscillation and reduction of drag. Detailed measurements of the velocity fields and kinematics of fin deformation show the significant effects of a flexible fin on the development of the flow structures in the fully-separated-flow region. The generating and controlling mechanisms of the low-frequency oscillation are explored. DOI: 10.2514/1.J050205

Measurements and predictions are made of a short-cowl coflowing jet with a bypass ratio of 8:1. The Reynolds number is 300,000, and the inlet Mach numbers are representative of aeroengine conditions. The low Reynolds number of the measurements makes the case well suited to the assessment of large-eddy-simulation-related strategies. The nozzle concentricity is carefully controlled to deal with the emerging metastability issues of jets with coflow. Measurements of mean quantities and turbulence statistics are made using both laser Doppler anemometry and particle image velocimetry. The simulations are completed on 6 x [10.sup.6], 12 x [10.sup.6], and 50 x [10.sup.6] cell meshes. To overcome near-wall modeling problems, a hybrid large-eddy-simulation--Reynolds-averaged-Navier--Stokes-related method is used. The near-wall Reynolds-averaged-Navier--Stokes layer is helpful in preventing nonphysical separation from the nozzle wall. DOI: 10.2514/1.44534

This work presents an immersed-boundary technique for compressible, turbulent flows and applies the technique to simulate the effects of micro vortex generators in controlling oblique-shock/turbulent boundary-layer interactions. The Reynolds-averaged Navier--Stokes equations, closed using the Menter k-[omega]) turbulence model, are solved in conjunction with the immersed-boundary technique. The approach is validated by comparing solutions obtained using the immersed-boundary technique with solutions obtained on a body-fitted mesh and with experimental laser Doppler anemometry data collected at Cambridge University for Mach 2.5 flow over single micro vortex generators. Simulations of an impinging oblique-shock boundary-layer interaction at Mach 2.5 with and without micro vortex-generator flow control are also performed, considering the development of the flow in the entire wind tunnel. Comparisons are made with experimental laser Doppler anemometry data and surface-pressure measurements from Cambridge University and an analysis of the flow structure is performed. The results show that three dimensional effects initiated by the interaction of the oblique shock with the sidewall boundary layers significantly influence the flow patterns in the actual experiment. The general features of the interactions with and without the micro vortex-generator array are predicted to good accord by the Reynolds-averaged Navier-Stokes/ immersed-boundary model. DOI: 10.2514/1.40049

Icing alters the shape and surface characteristics of aircraft components, which results in altered aerodynamic forces and moments caused by air flow over those iced components. The typical effects of icing are increased drag, reduced stall angle of attack, and reduced maximum lift. In addition to the performance changes, icing can also affect control surface effectiveness, hinge moments, and damping. These effects result in altered aircraft stability and control and flying qualities. Over the past 80 years, methods have been developed to understand how icing affects performance, stability, and control. Emphasis has been on wind-tunnel testing of two-dimensional subscale airfoils with various ice shapes to understand their effect on the flowfield and ultimately the aerodynamics. This research has led to wind-tunnel testing of subscale complete aircraft models to identify the integrated effects of icing on the aircraft system in terms of performance, stability, and control. Data sets of this nature enable pilot-in-the-loop simulations to be performed for pilot training or engineering evaluation of system failure impacts or control system design. DOI: 10.2514/1.44650

The main goal of this paper is to document a comparative study of different computational-fluid-dynamics-based optimization techniques applied to the solution of a three-dimensional wing drag minimization problem. To achieve this objective, three optimization tools were used: SYNI07 (Intelligent Aerodynamics International), MDOPT (The Boeing Company), and OPTIMAS (Israel Aerospace Industries). The first tool employs gradient-based search techniques using the continuous adjoint equation, the second one is a response-surface method, and the last one uses a floating-point genetic algorithm as its search engine. As the starting geometry, the public domain DPW-W1 wing (a test case for the Third AIAA Drag Prediction Workshop) was used. The comparisons included herein are provided in three stages: cross analysis of the initial geometry by the computational fluid dynamics tools employed in the optimizations, optimization of the initial geometry to minimum drag, and cross analysis of optimal shapes achieved by the optimization tools using all computational fluid dynamics tools employed. The cross analysis also includes results from an independent computational fluid dynamics method that was not used in any of the optimization efforts. These results help quantify the level of variation that is inherent in, and can be expected from, application of the current state-of-the-art aerodynamic optimization methods. The present work may be regarded as a move toward the construction of reliable test cases for an aerodynamic shape optimization problem. Another goal of this collaborative investigation is to collect lessons learned from this pilot project to help develop a model for an aerodynamic optimization workshop.

A comprehensive strategy for developing and implementing discrete adjoint methods for aerodynamic shape optimization problems is presented. By linearizing each procedure in the entire optimization problem, transposing each linearization, and reversing the sequential order of operations, the adjoint of the complete optimization problem, including flow equations and mesh motion equations is constructed in a modular and verifiable fashion. This construction is also shown to produce minimal memory overheads and retain the same convergence characteristics of the original analysis problem in the sensitivity analysis. These techniques are implemented in a three-dimensional unstructured multigrid Navier--Stokes solver and demonstrated on a transonic drag reduction problem for a wing body configuration. DOI: 10.2514/1.22743

Design optimization using high-fidelity computational fluid dynamics simulations is becoming increasingly popular, sustaining the desire to make these methods more computationally efficient. A reduction in problem dimensions as a result of improved parameterization techniques is a common contributor to this efficiency. The focus of this paper is on the high-fidelity aerodynamic design of airfoil shapes. A multifidelity design search method is presented which uses a parameterization of the airfoil pressure distribution followed by inverse design, giving a reduction in the number of design variables used in optimization. Although an expensive analysis code is used in evaluating airfoil performance, computational cost is reduced by using a low-fidelity code in the inverse design process. This method is run side by side with a method which is considered to be a current benchmark in design optimization. The two methods are described in detail, and their relative performance is compared and discussed. The newly presented method is found to converge towards the optimum design significantly more quickly than the benchmark method, providing designs with greater performance for a given computational expense.

An investigation is performed to assess the importance of the nonlinear effects on the dynamic behavior of a profile in a transonic flow. The analysis is done by applying a system identification approach to a computational fluid dynamics (CFD) code, considered as a black-box system. The CFD code is based on the Euler equations. Both frequency and time domain approaches are used to evaluate the nonlinear effects on the response of different airfoils. The results show that for weak shocks the aerodynamic operator describing the dynamics of the profile around a steady (transonic) flow condition is linear. For strong shocks results obtained with linear models appear to be conservative.

Past research on airfoil aerodynamics iii icing are reviewed. This review emphasizes the time period after the 1978 NASA Lewis workshop that initiated the modern icing research program at NASA and the current period after the 1994 ATR accident where aerodynamics research has been more aircraft safety focused. Research pre-1978 is also briefly reviewed. Following this review, our current knowledge of iced airfoil aerodynamics is presented from a flowfield-physics perspective. This article identifies four classes of ice accretions: roughness, horn ice, streamwise ice, and spanwise-ridge ice. For each class, the key flowfield features such as flowfield separation and reattachment are discussed and how these contribute to the known aerodynamic effects of these ice shapes. Finally Reynolds number and Mach number effects on iced-airfoil aerodynamics are summarized. [c] 2005 Elsevier Ltd. All rights reserved.

Several theoretical and computational methods for modeling the unsteady flows about turbomachineries are studied. These methods have been formulated primarily for analyzing blade vibrations and noise generation that occur within an axial-flow turbomachines. They are also capable of predicting the aerodynamic responses of a blade row subjected to external aerodynamic excitations. The study indicates that further research is still needed to extend the applications of these traditional methods.

A three-dimensional methodology was developed for unsteady flows past bodies that are in relative motion and where the trajectory of the motion was determined from the instantaneous aerodynamic field. This method coupled the equations of fluid flow and those of rigid-body dynamics and then captured the unsteady aerodynamic interference between the stationary and moving boundaries. The time-dependent, compressible Euler equations were solved on dynamic, unstructured meshes by an explicit, finite volume, and upwind method. The grid adaptation was performed within a window placed around a moving body. The Euler equations of dynamics were then solved by a Runge--Kutta integration scheme. The flow solver and the adaptation scheme were validated by simulating the transonic, unsteady flow around a wing undergoing a forced, periodic pitching motion and then comparing the results with the experimental data. Finally, the overall methodology was demonstrated by simulating the unsteady flowfield and the trajectory of a store dropped from a wing. The methodology, with its computational cost notwithstanding, has proven to be accurate, automated, easy for dynamic gridding, and relatively efficient for the required work hours.