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The note demonstrates the application of a pure Lagrangian numerical method to dam-break flows by solving the St. Venant equations. The method is developed based on the smoothed particle hydrodynamics. It is easy to apply and is shown to be capable of providing accurate simulations for mixed flow regimes with strong shocks.

The Spalart-Allmaras one-equation turbulence model is effective for general Navier-Stokes computations of high-lift aerodynamic flows, and the Menter two-equation turbulence model is accurate for separated flow regions. The Mentor model is useful for single-element case and results in better approximation with the experimental velocity and pressure distribution profiles. The Spalart-Allmaras model provides more accurate prediction of turbulent shear stresses than the Menter model. The performance of the models in some cases of multi-element airfoil geometries is discussed.

Time-dependent models based on recursive neural networks accurately predict unsteady boundary-layer development, separation, dynamic stall and dynamic reattachment. These models can also be used for the development of adaptive control systems. The models integrate the properties of the unsteady boundary layer or of the unsteady flowfield dominated by a large energetic vortex. Results demonstrating the effectiveness of the models for three-dimensional unsteady surface pressures and two-dimensional unsteady shear-stress measurements are presented.

An analysis of the effects of film cooling on the secondary flow in a turbine nozzle helps explain the phenomena of vortex generation and development in flow regions, where direct measurements cannot be taken. The greatest changes in the secondary flow are found near the tip for increase in the blowing ratio. Cooling air ejection with low momentum does not affect the streamlines at the suction side significantly. The position of secondary vortices remains unaffected, when compared to the conditions without film cooling, for blowing ratio M1 = 0.5, but changes for M1 = 1.4.

A symmetry-breaking bifurcation occurs at low Reynolds numbers in flows in two-dimensional channels with a sudden symmetric expansion. This bifurcation represents transition from a symmetric to an asymmetric developing jet. The critical Reynolds number is found to decrease with an increase in the channel expansion ratio. For a given supercritical Reynolds number, the position of jet impingement on the channel wall grows with the expansion ratio. For a given expansion ratio, the number of attachment positions for an asymmetric jet increases with an increase in the Reynolds number.

A two-dimensional Navier-Stokes solver for incompressible flows allows numerical simulation of the unsteady flow over a hydrofoil. The scheme is time-accurate and semi-implicit pressure-based. The method involves the use of a low-Reynolds-number form of the two-equation turbulence model for studying turbulence effects. The predictions of characteristics, such as velocity and phase changes, of unsteady boundary layers along the foil surfaces are accurate.

An explicit algebraic stress model (EASM), based on a pressure-strain model, is able to accurately predict nonequilibrium flows over an airfoil and a wing. The ISAAC Navier-Stokes code allows the simulation of the flows. The EASM gives a realistic representation of the anisotropy of turbulence. It is also able to predict shock locations and surface pressure distributions. The failure of dissipation rate equation to respond properly to adverse pressure gradients puts a limitation to the efficiency of the model.

The properties of shock turbulent mixing interaction of isoconcentration surfaces depend on the degree of mixing upstream from the compression. Direct numerical simulation employs various scalar injections to monitor the modifications in the in-plane strain rate localized in the shock zone. The laminar injection reveals the sensitivity of the curvature of the scalar field to the presence of pressure jump. The blob type injection shows the localization of the stretch within the shock zone because the stretch is more in control by the strain rate than by curvature.

The application of algebraic turbulence models to aerodynamic and propulsion flowfields reveals the prominent role of a velocity-based strain rate and a production limiter in improving the accuracy of the flowfield predictions. The flowfields pertain to turbulent kinetic energy at low Reynolds numbers, and the algebraic models are unable to capture the essence of these complex flowfields for the internal flow applications. The Spalart-Allmaras one-equation model seems the optimum choice of turbulence models with respect to design cycle time considerations.

A nonlinear disturbance survey of a two-dimensional, incompressible, spatially developing monochromatically forced flat-plate wake produces evidence of harmonical forcing of the wake at the inflow boundary. This wake occurs at the maximum amplification frequency given by linear stability theory, and produces a disturbance component at the forcing frequency that shows a rapid increase before saturating a short distance downstream. The disturbance energy is observed prominently in the fundamental disturbance, the second harmonic, and the mean-flow distortion component.

A time-accurate double thin layer Navier-Stokes computation for an unsteady supersonic open-cavity determines the flow-physics mechanisms that bring about the cavity oscillation cycle. A shed vortex collides with the cavity aft lip to form a pressure pulse that augments an internal upstream moving wave at vortex shedding frequency. This moving wave then reflects off the cavity forward wall and shedding of a new vortex occurs. The reflected wave dissipates before reaching the aft wall and a second wave formed beneath the shed vortex reflects from the aft corner to complete the cycle.

Aeroacoustic inversion for flat-plate airfoils depends not only on an optimal choice of the regularization parameter but also on the input data locations. A conscious avoidance of the input data present directly upstream/downstream of the airfoil controls the sensitivity of the inverse solution to errors in the input data. Sensitivity analysis of the effect of mean-flow parameters reveals the dependence of inverse process on Mach number and reduced frequency. The use of input data from the midfield improves the reconstruction.

Statistical analysis of large image ensembles helps identify and characterize the major features of large structures associated with compressible mixing layers. Laser sheet Mie scattering techniques image the large structures, with condensed ethanol droplets as the scattering media. Digital image ensembles for three flow conditions reveal structures related to spanwise and streamline vortices. Transverse image covariance field show elongated and compressed structures toward the streamline flow direction. The effects of compressibility on the shapes of the structures are studied.

An overview of two-phase gas-liquid flow with respect to flooding and disturbance waves in annular flow stresses the importance of better experimental design and instrumentation models, and availability of computational fluid dynamics (CFD) techniques. Flooding phenomenon is observed from visualization studies, droplet entrainment experiments, correlation methods based on Kutateladse number and CFD channel geometries. The high-speed and axial view photography with hot-film probes, fluorescence methods, and CFD techniques are analyzed for the disturbance wave studies.

A study of the flowfield and noise properties of a supersonic diamond-shaped jet reveals many structures in the shear layers that are not detected in these layers of expanded axisymmetric and rectangular jets. These structures express the streamwise vorticity in the layers, which may not affect the overall mixing and noise fields in these jets. Narrow-band spectral studies of the diamond jet demonstrate a moderate reduction of the mixing noise as compared to that of an equivalent axisymmetric jet. The shear layer global growth rates are similar for different jet geometries.

A reduced-order model based on eigenmodes facilitates the calculation of the unsteady aerodynamic response of a system over a range of reduced frequencies. The Lanczos algorithm helps to solve a large eigenvalue problem that forms the base for the computation of turbulent flow about airfoils, wings and cascades. The model is also applicable in the active control of aeroelastic and aeroacoustic problems, and flutter or gust response analysis.

The measurements of Mach number profiles for two stream conditions provide insights into the instabilities present in supersonic shear layers. An artificial excitation for two-dimensional (2-D) waves produces 2-D instabilities in both the cases. The excited three-dimensional waves create oblique instability waves for different excitation angles. The growth rate of waves for low Mach number conditions remain unaffected by excitation, whereas enhanced mixing is observed for the higher Mach number profiles.

A software program predicts plume velocity contours to map jet plume zones for a jet aircraft during airport ground operation. The velocity is expressed as a function of the aircraft geometry and engine power setting. The velocity contours, determined using a standard normalized velocity contour plot, are superimposed on a predigitized airport map. A computer graphics program allows the display of an airport plume zone map on a computer screen.

The maximum entropy direct simulation Monte Carlo (ME-DSMC) code reliably predicts flowfields in the rarefied hypersonic range. Its application to a blunted cone demonstrates agreement with experimental data for a wide range of Knudsen numbers. The ME-DSMC code is also able to provide accurate predictions for the surface and flowfield characteristics of the wake and reattachment flows. The ME-DSMC code uses the maximum entropy method for modeling chemically active flowfields but employs the generalized Borgnakke and Larsen method for nonreacting flows.

Flow visualization using laser Doppler velocimetry in the near wake of a cylindrical afterbody shows that the bleed flow displaces the primary recirculation region downstream of the base plane. The size and strength of the region reduces as the bleed flow provides most of the fluid needed for shear layer entrainment. High base pressure, absence of primary recirculation region and low turbulence levels in the near-wake flowfield enhance the benefits of base bleed.

The vortical structures in the wakes of starting flow behind regular polygonal plates interact and show convoluted deformations. The wake deformation and breakdown of the vortex structure are due to the longitudinal component of the vorticity. The arched vortex structure in the initial wake gives rise to the vorticity component. The wake formation behind octagonal plates is similar to that behind a disk. Sharp corners give rise to prominent vortex deformations.

Planar supersonic base flow visualizations show that large-scale turbulent structures exist in compressible shear layers. The structures interact vigorously with recirculation region. The shear layer gives rise to Mach and shock waves. An eddy shocklet formation may be due to the shock waves. The structures are elliptical and have an inclination towards mean flow direction. The coherent eddies cause shear layer formation, recompression and reattachment processes which lead to a flattening and tilting in the streamwise direction.

The high-enthalpy, hypersonic compression corner flow for a flat plate is laminar and frozen. The flow behavior is similar to that of a perfect gas. For compression corner flow, increasing wedge angle increases the heat transfer and pressure on the face of the ramp. The increase in upstream influence is similar to that in perfect gas flows for both the high- and low-enthalpy conditions. A study shows that the real gas effects are negligible for the experimental conditions possible with the models used.

A time-domain aeroelastic analysis code provides flow solutions for flutter of single airfoils and two-foil systems in incompressible, inviscid flow. A time-stepping panel code provides flow solutions and a two-degree-of-freedom spring/mass model computes airfoil motions. The computed results compare well with the frequency domain results. Three applications, two of which are simulations, of the aeroelastic code are discussed.

Numerical simulations help in explaining the role of paddles in enhancing the acoustic noise and mixing of supersonic rectangular jets. Paddles strengthen a periodic flapping mode in the minor axis direction. The FAST3D program, using flux-corrected transport algorithms and the virtual-cell embedding technique, performs the simulations. The flapping frequency estimations involve pitot pressure distributions and Fourier analysis. The thrust loss caused by the paddles agrees well with the experimentally determined value.

Low stagnation temperature of flow allows the most efficient control through the secondary nozzle of a hypersonic planar twin-nozzle system. A cold secondary jet, which seems to act like a gasdynamic wall due to its locally higher acoustic impedance, controls the primary jet channel shaping. Subsonic gas injection helps in stabilizing the flowfield, when pressure ratio is 10 or less. The study is conducted under systematic variations of the stagnation state and the dimensions of the secondary nozzle.

A hierarchically structured, Cartesian cell-based approach provides adaptively refined solutions of the Euler and Navier-Stokes equations in two dimensions for viscous and inviscid flows. The automated grid generation procedure helps generate grids about complicated geometries with minimum user intervention. The grid generation involves recursive subdivision of a single Cartesian cell containing the whole flow domain. The hierarchy of the grid generation process is stored in a binary tree.

The full Navier-Stokes calculations of forebody and fuselage surface static pressures in the F/A-18A High Alpha Research Vehicle inlet are in good agreement with flight data. The calculations consider different angles of attack and freestream Mach numbers. The streamwise flow separation at the bottom of the inlet lip is unknown. In this region of the inlet, the numerical calculations do not agree with the flight data at 30- and 60-degree angle of attack.

The numerical predictions of forebody, leading-edge extension, inlet lip and duct surface pressures in F/A-18A inlet flow compare well with the flight data. A significant degree of flow unsteadiness exists at the engine face. The inlet flow distortions and total pressure recovery at the engine face are at the upper and lower range of test data. The calculated total pressure contours are within the excursion range of the minimum total pressure indicated by the flight data.

Placement of vortex generators directly above the first and fifth ribs in a ribbed duct air flow affects the flow and heat transfer behavior. Flow temperature measurements and laser-Doppler flow measurements at the smallest spacing show that the generator-rib pairs behave as a single unit with an increased effective height. At the larger spacing, the flow expands in the cross-stream direction.

A study presents flow visualization plus force and moment information for a double delta wing model of 80 degrees/60 degrees/0.6at fixed angles of attack and for sinusoidal pitching motions. The results explain the dynamic effects of the double delta wing. The effects of varying the reduced frequencies of the motions and the angle of attack range are studied.

Two studies evaluated the ability to forecast complicated three-dimensional hypersonic configuration flowfields through computation in simulated environments. This technique can supply accurate moment and aftbody force results.

Vortex liftoff has a significant role in generating the maximum wing rock amplitude. Vortex breakdown restricts the magnitude of the maximum limit cycle amplitude, although the limit-cycle motion can be set up without vortex breakdown. Roll-rate induced camber effects can be a significant factor in the high alpha region mainly when vortex breakdown is the flow mechanism restricting the amplitude buildup.

A study of nonequilibrium-dissociated stagnation boundary-layer flow on an arbitrarily catalytic swept wing shows that the overall effects of sweep angle reduce the dissociation effects in the frozen gas-phase regime. The dissociation effects are cancelled for sweep angles above 65 degrees. This creates perturbations in the functioning of a noncatalytic wall in reducing nonequilibrium heat transfer. The study is conducted in the high-altitude hypersonic flight regime.

A numerical analysis shows that the distributions of specie concentrations of turbulent reacting flow in gas turbine combustors are sensitive to the different models used for observations. The mean velocity fields, the unburnt fuel and the temperature distributions in the flowfields in gas turbine combustors are relatively less sensitive. The flowfield is also sensitive to different model parameters. The sensitivity increases with an increase in turbulence.

A study on rolled delta wing flows shows that effective sweep and attack angles of the leading edge affect vortex behavior on both sides of a delta wing. Yawed and unyawed delta wings show similar results for effective sweep angles below 70 degrees in empirical calculations for burst point. The calculations are done using a curve drawn to compare yawed and unyawed data. The study is conducted on suction surface of a 65-degree swept delta wing at a 30-degree incidence angle.

Smoke-wire flow visualization helps analyze flow structures in the near wake region of unducted double concentric jets at low Reynolds numbers. The different jet velocities have different flow regimes in the near disk region. Experiments help analyze flow pattern, recirculation length and vortex shedding in each flow regime. Frequencies exhibit a correlation with the Strouhal number. Central-to-annular jet velocity ratio and various processes for vortex shedding are explained.

Flow visualization and laser Doppler estimations provide elaborate data on the influence of impingement edge geometry on the features of a cavity flow that exhibits organized oscillations. The nose-shape impingement edge shows a weakening of fluctuation peak levels, which results from the most frequent escape of the separated shear-layer vortices nearing the edge. Use of three different cavity edge geometries for the estimations does not alter the appropriate Strouhal number values.

Surface pressure measurements and flow visualization studies characterize the unsteady, vortex-dominated flow developed over a straight wing and the circular splitter plate it is attached to. The leading-edge vortex undergoes significant three-dimensional variations to cross section and convection over the entire wing span and this effect is correlated with spanwise alterations in force coefficients. Specific physical mechanisms are involved in three-dimensional fluctuations in unsteady forces and surface pressures.

The usage of Euler and Navier-Stokes equation solutions for a vortex on a highly-swept delta wing yields a Rossby number-based criterion for vortex breakdown. The Rossby number identifies vortex breakdown locations along a chord and yields results that agree well with experimental data. The experiment yields a critical Rossby number range of 0.9 to 1.4, beyond which vortex burst occurs.

An analysis of high pressure fluctuations on aircraft wing-body junctions reveals that pressure fluctuations and separation line locations vary at the wing nose, based on wing geometry. Pressure-fluctuation data yield bimodal flow formation locations. Microphones measure pressure fluctuations, while wire anemometry-generated velocity profiles depict incoming flow boundary layers.

Unsteady pressure estimates were conducted along the top and around the circumference of a body in low-speed forward flight to analyze rotor tip vortex interactions with the body. These interactions can be classified into three components, close tip vortex/body interaction, postvortex/surface and vortex/surface impingement. A high degree of responsiveness to tip vortex convection speed and location was evident.

A new topological structure for flow over a delta wing undergoing transient pitching maneuvers at high angle of attack is not similar to the traditional topology for stationary wings at low angle of attack. An outward spiral motion is exhibited by the leading edge vortex corresponding to a significant change in the entire crossflow topology. Due to this major change in the nature of the leading-edge vortex, an unstable focus is induced, possibly by the rapid, transient pitch-up motion of the wing which increases the compression of the core vortex filament.

An innovative indicial/gust response measurement technique for airfoils provides naturally decoupled loading parameters while overcoming the problems associated with step changes at the boundaries. An implementation of the technique into a Euler/Navier-Stokes solver enables direct calculation of the indicial responses as a function of step changes in angle of attack. The use of computational fluid dynamics gives a grid-velocity framework that is as accurate as known analytical results for linear flowfields. However, the agreement turns worse for increasingly nonlinear flowfields.

The discrete vortex method is an efficient tool for numerical simulation of external unsteady aerodynamic flows separated at low speeds. The method designs the flow around an airfoil with an extended spoiler having realistic features such as viscous diffusion of vorticity and zonal decomposition. The flow patterns, pressure distribution on the airfoil and time history of the lift agree with the computational results.

A strained coordinate slipline perturbation method provides a linearized, potential flow solution for freejet, axisymmetric and two-dimensional flows. The solution is verified using freejet experimental data and by comparing with other classical solutions. The strained coordinate solution is an able predictor of freejet critical locations over a large range of Mach numbers and pressure ratios. The solution can be directly applied to supersonic ejector flows and preliminary design problems.

The vortex-wake decay and trajectory characteristics of a supersonic transport planform in a Mach 2.5 stream aid in the development of vortex-wake models for vortex flow visualization. Shadowgraphy suggests three distinct structures in the vortex wake, the wing tip vortex, the leading-edge vortex and the swirling wavelike structure. Pitot pressure surveys provide decay curves and trajectory information for the structures. The diffusion behavior of these vortex flows is significant for supersonic transport operations and may affect aircraft design.

A mass-loss model based on fundamental boundary-layer theory helps in predicting shock tube test time and shock speed. A quasi-one-dimensional simulation code employs the model to design shock and expansion tubes. A study demonstrates that the model can provide accurate test time predictions for all turbulent flow shock tubes. The model uses a computationally less expensive method to provide design information. The formulations, however, do not provide as much information as two-dimensional viscous codes.

Artificial neural networks can act as components for the automation of wind-tunnel operations. A continuous source of light and a charge-coupled device camera near the Mach 2.0 nozzle exit help in recording shadowgraph patterns. The shadowgraphs give an estimation of the sensor readings for both the training and test sets. The manually controlled tunnel operator records the sensor values and the flow visualization patterns, which leads to shadowgraph-to-sensor sequences.

Particle image velocimetry and flow visualization are used to study a laminar wall jet, disturbed from an external source. Under external disturbance, vortical structure in outer shear layer is improved. Collection and ejection of wall vorticity into outer flow is caused by convecting shear layer vortical structure. Wall jet flow is totally thrown out due to strong interaction between inner and outer vortex structures, and the flow becomes turbulent.