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Flame propagation in channels and cracks is a problem of considerable interest with applications in many practical combustion devices, in fire hazard scenarios, and in the emerging micropropulsion technologies. Understanding the dynamics and Stability characteristics of flame propagation in channels is, therefore, important both for fundamental research as well as for practical applications. In this work, we examine the propagation of a premixed flame front in a two-dimensional channel in the presence of a Poiseuille flow. Our primary objective is to determine within the flammable regions the structure of the flame front and the conditions that result in steady propagation and those leading to a pulsating mode of propagation. Special attention is given to the difference between propagation in narrow and wide channels, heat losses to the channel's walls, and an imposed flow that either supports or opposes the propagation. In general, flame oscillations are found to occur in mixtures for which the effective Lewis number is sufficiently large. They are more likely to occur in narrow or wide channels and particularly at near-extinction conditions where the critical Lewis number is reduced to physically accessible values.

Economic competitiveness is driving industry to new frontiers of engineering design. Robustness and reliability-bused design, multidisciplinary-simulation-based design, increased complexity and sophistication of our design, and optimization-aided design are four such areas that are seriously challenging our ability to keep pace with the need to adequately model the systems we seek to design--in spite of the exponential growth of computing power. An emerging consensus within the community is that the effective development of computationally benign models (metamodels) will help us navigate the challenging road ahead. The process of constructing metamodels for computationally expensive models, such as finite element, aerodynamic, and heat-transfer models, is representative of the tasks we must address. Among the available metamodeling techniques, radial basis functions (RBFs) have recently generated much interest for their effectiveness and versatility. Radial basis functions offer numerous advantages over the traditional response surface methodology, including their ability to effectively generate multidimensional interpolative approximations. However, we show how the typical RBF approach lacks the critical flexibility required to handle the wide variety of complex models arising from the use of advanced techniques, such as uncertainty handling and multiobjective optimization, often encountered in modern design. Furthermore, in this paper we propose a novel approach---the extended radial basis function (E-RBF) approach--that provides the designer with significant flexibility and freedom in the metamodeling process, compared to conventional RBFs. Examples are provided that demonstrate the effectiveness of the new approach and explore its potential superiority to traditional RBF and response surface methodologies. Initial investigation indicates that the E-RBF possesses unique and novel properties not available in any other single method.

The flow over two-dimensional geometries is studied via unsteady numerical simulations that are three dimensional, with periodic conditions applied along the spanwise coordinate. This framework is well accepted for direct numerical simulations (DNS), large-eddy simulations, and detached-eddy simulations (DES), but is here combined with standard Reynolds-averaged Navier-Stokes (RANS) turbulence models. This strategy, which is not new, is referred to as unsteady RANS (URANS). Limited previous evidence suggested that, in URANS, three dimensionality is suppressed by high eddy-viscosity levels. However, three dimensionality proves fairly easy to sustain with adequate initial conditions, in all three cases studied here: stalled airfoil, circular cylinder, and a rounded square, except that for one case three dimensionality failed to last from random-based initial perturbations and was sustained only when using a DES field as initial condition. It is much less fine grained and chaotic than in the classical turbulence-resolving methods (from DNS to DES). Three-dimensional URANS gives clear improvements over two-dimensional URANS. It is less costly than DES, but is not as accurate. URANS also displays a troublesome sensitivity to the spanwise period and to the turbulence model. The approach is interesting and will appear spontaneously in many applications, but remains only partly understood.

The flow around infinite swept wings is computed using a Reynolds-averaged Navier-Stokes method coupled to a boundary layer and a transition prediction method based on the [e.sup.N] approach. This approach applies the two N factor [e.sup.N] strategy, which allows the differentiation between excited Tollmien-Schlichting waves of constant frequency and stationary crossflow waves of constant spanwise wave number. The applicability is documented by comparing computations with infinite swept-wing experiments on four configurations, where one configuration was tested in two different wind tunnels. The limiting N factors for both types of waves are a priori unknown. Hence, firstly the limiting N factors, which represent the stability limit, are deduced from the measured transition locations. Subsequently, iterative computations are executed to determine the transition location using the coupled system, Navier-Stokes, boundary-layer, and two N factor [e.sup.N] method in combination with the stability limit. The correlation between experimentally observed and computed transition locations is shown to be good for all considered wind-tunnel studies.

Expressions to describe the correlation length scales of turbulent inflow to an aerodynamic body are derived as functions of the classic integral length scale and anisotropy correction factors. These one-point parameters are significantly easier to determine experimentally than traditional correlation-scale measurement techniques, which involve multiple probes at multiple locations. As such correlation scales are necessary to properly estimate the aeroacoustic response of the body, such a technique could have substantial benefit in a wide variety of applications. The approach is applied to a recent experimental study examining the response of a stator downstream of a propeller that is itself ingesting broadband turbulence. Results suggest that the derived expressions not only accurately represent correlation length scales, but also enable the accurate prediction of the acoustic output of the stator.

Within a parametric study on brush-type trailing-edge extensions, the noise reduction potential of several design concepts was determined. The obtained database represents the first phase of an ongoing project with the long-term objective to develop scaling laws for a future application of such devices as add-on solutions for today's aircraft components. The experiments comprised both acoustic and aerodynamic measurements on a zero-lift generic plate model (Re=2.1 x [10.sup.6] to 7.9 x [10.sup.6]) in DLR's open jet Aeroaconstic Wind Tunnel Braunschweig. Noise data were taken by means of a directional microphone system. Measurement results indicate a significant source noise reduction potential in excess of 10 dB, depending on the configuration. Two relevant noise reduction mechanisms were identified: 1) the suppression of narrowband bluntness noise, as well as 2) the reduction of broadband turbulent boundary-layer trailing-edge noise.

Instability of the three-dimensional boundary layer on a yawed circular cylinder placed in a uniform flow is investigated experimentally by introducing acoustic disturbances from a point near the attachment line. To exemplify the flow dominated by streamline-curvature instability rather than crossflow instability, which has been often observed in many swept-wing flows, is the aim here. In upstream regions of the disturbance wedge originating from the point source, both streamline-curvature and crossflow disturbances are superposed on each other and yield complicated amplitude distributions. A newly proposed method enables the decomposition of the distorted amplitude distribution into contributions from the two instability modes. Detailed observations, however, show that the crossflow mode decays with the distance from the source much faster than the streamline-curvature mode and allows the latter to be dominant in a region further downstream. A fundamental characteristic of the streamline-curvature instability wave is confirmed by examining its phase distribution in the spanwise and normal directions. Wave numbers and spatial growth rates are in good agreement with theoretical predictions.

The Harten, Lax, and van Leer with contact restoration (HLLC) scheme has been modified and extended in conjunction with time-derivative preconditioning to compute flow problems at all speeds. It is found that a simple modification of signal velocities in the HLLC scheme based on the eigenvalues of the preconditioned system is only needed to reduce excessive numerical diffusion at the low Mach number. The modified scheme has been implemented and used to compute a variety of flow problems in both two and three dimensions on unstructured grids. Numerical results obtained indicate that the modified HLLC scheme is accurate, robust, and efficient for flow calculations across the Mach-number range.

The objective of this study is to develop and employ a numerical simulation strategy for predicting the airflow from the spray bars to the test section of the NASA John H. Glenn Icing Research Tunnel (IRT). In particular, predictions of the mean velocity and turbulence distributions were desired throughout this flow domain to later investigate droplet dispersion. Computational airflow results were produced using the WIND code (developed by NPARC) with a second-order accurate finite difference scheme and the shear stress transport k-[OMEGA] urbulence model. The inflow conditions for the flow domain were derived from the IRT measurements just upstream of the spray bars, which reflected the contributions to turbulence from the upstream heat exchanger wake. It was found that inclusion of the spray bar wakes and the air jets (of the spray nozzles) were required to describe the wind-tunnel turbulence distribution. Because it was impractical to simultaneously resolve the overall flow domain (60 ft long), along with the detailed flow around the 10 spray bars and the flow within a hundred air jets (issuing from 1/8-in. nozzle diameters), these features were simulated individually and then algebraically combined together to give an approximate solution. The results of the spray bar wake combined with the heat exchanger flow yielded good prediction of test section mean velocity and turbulence for the jets-off condition. Inclusion of all of the individual air jets also yielded reasonable resulting predictions of mean velocity and turbulence in the test section.

The results of proper orthogonal decomposition analyses on C[O.sub.2] number density and vorticity magnitude data from reacting rectangular jet simulations are presented. The resulting proper orthogonal decomposition eigenfunctions are used to develop physical insight of the vortex formations and dynamics of these jets and their related mixing and spreading characteristics. It is seen that different vortex structures are captured in the eigenfunctions and that C[O.sub.2] and vorticity eigenfunctions are very similar indicating that vortex-driven mixing dominates in these jets. The eigenvalue spectra associated with these eigenfunctions are used to evaluate the information content of the eigenfunctions and the potential for reduced-order models. Using subsets of eigenfunctions with high information content, C[O.sub.2] and vorticity magnitude distributions can be represented with relatively few eigenfunctions. However, as the flows develop downstream, more eigenfunctions are needed to represent them to the same level of accuracy. The potential for reduced-order modeling of each field is approximately the same for the jets of aspect ratios 1, 2, and 3; however, there is stronger potential for reduced-order modeling of the C[O.sub.2] field than of the vorticity field.

The development of new, high-performance aerospace structures requires the availability of powerful and efficient methods for dynamic testing. Also, the European aircraft industry is calling for a reduction of the testing times for prototypes without diminishing the accuracy of the test results. As a consequence, substantial improvements in the test strategy of ground vibration and modal survey tests have been developed and applied during the last years. In addition, space agencies like the ESA are interested in advanced measurement techniques, for example, when a spacecraft has to be qualified by a boosted modal survey test. Very high loads are introduced in the structure during these tests, and significant nonlinear behavior of the spacecraft can sometimes be observed. A test strategy is presented that can be used to identify and characterize nonlinear, structural behavior during modal testing. The method assumes a weak non-linear behavior and operates in modal space. The approach can replace the common analysis by linearity plots that are utilized during modal survey testing of aerospace structures. The proposed test strategy supplies much more information on the nonlinear behavior than the common procedure and allows for a reduction in the test duration.

Recently, an algorithm for dynamic response optimization transforming dynamic loads into equivalent static loads has been proposed. In later research, it was proved that the solution obtained by the algorithm satisfies the Karush-Kuhn-Tucker necessary conditions. In the present research, the proposed algorithm is applied to the optimization of flexible multibody dynamic systems. The equivalent static load is derived from the equations of motion for a flexible multibody dynamic system. The equivalent load is utilized in sequential static response optimization of the flexible multibody dynamic system. In the end, the converged solution of the sequential static response optimization is the solution of the original dynamic response optimization. Some standard examples arc solved to show the feasibility and efficiency of the proposed method. The control arm of an automobile suspension system is optimized as a practical problem. The results are discussed regarding the application of the proposed algorithm to flexible multibody dynamic systems.

A numerical investigation of the adiabatic and condensing flowfield around the ONERA M6 and F-16 Falcon wing is performed. For the ONERA M6 wing, three different mean aerodynamic chord [c.sub.mac] lengths, 0.5, 1.0, and 2.0 m, three angles of attack, 1.07, 3.06, and 6.06 deg, and three different number densities of particles, [10.sup.6], [10.sup.12], and [10.sup.16] [m.sup.-3], are analyzed. At [10.sup.6] [m.sup.-3], the flow is dominated by homogeneous condensation, and due to the large chord length, the effects of condensation are small. As the particle density increases, a mixture of homogeneous/heterogeneous condensation occurs at [10.sup.12] [m.sup.-3], and at [10.sup.16] [m.sup.-3], only heterogeneous condensation occurs. For the F-16 Falcon wing, the size of the wing and the number density of particles [10.sup.12] [m.sup.-3] are kept constant. The angle of attack and freestream Mach number are changed, where a decrease in the lift-to-drag ratio reached 20%. In general, if the wing is experiencing separation, condensation can improve the lift-to-drag ratio, whereas for attached flow with small flight angles, condensation reduces the lift-to-drag ratio.

The morphing concept for unmanned aerial vehicles is a topic of current research interest in aerospace engineering. One concept of morphing is to change the wing configuration during flight to allow for multiple flight regimes. A particular approach to planform morphing is a variable-span morphing wing to increase wingspan to reduce induced drag and increase range and endurance. The wing area and the aspect ratio of the variable-span morphing wing increase as the wingspan increases. This means that the total lift increases while the induced drag is reduced, whereas the wing-root bending moment increases, thus, requiring a larger bending stiffness of the wing structure. Therefore, a study of the variable-span morphing wing requires not only aerodynamic analysis but also an investigation of the aeroelastic characteristics of the wing. The aerodynamic characteristics of the variable-span morphing wing are investigated, and a static aeroelastic analysis is performed.

Research shows that charge-coupled camera devices can resolve pitch and yaw angles of registration targets in luminescent paint images to within 0.0005 deg rms in wind tunnel models. However, differences between camera-based angles and those from sting-mounted balance can be as large as (plus or minus) 0.05 deg.

Research identifies two preferred detonation techniques, a downstream and an upstream mode with the former mode performing better. Data indicate that detonation method is useful to study various gasdynamics problems and they are simpler to operate than the free piston tunnels.

An analytical model is presented that accounts for the effects of self-illumination of a diffusely reflecting test article in luminescent-paint measurements. Contributions from multiple reflections of excitation and emission light are summed to infinite order in closed form. Also included are the effects of spectral leakage resulting from an imperfectly filtered source and imperfectly rejected reflected excitation light. It is shown that the conventional method of determining flow variables from the ratio of flow-on and flow-off signals is adversely affected by self-illumination and spectral leakage. An explicit solution is derived for the associated inversion problem. Sample calculations are presented. Key words: Luminescent paint, pressure-sensitive paint, self-illumination, spectral leakage, aerodynamic measurement technology, inverse problems, computational flow imaging.