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The theoretical groundwork for the detection of any aircraft from its unique turbulent wake signature will be discussed, and practical methods to implement this theoretical result will be recommended. During the mid-1990s, NASA conducted research on the detection of atmospheric turbulence using light detection and ranging equipment. This light detection and ranging technology was able to detect turbulent disturbances by the Doppler shift in the frequency of laser-emitted energy that is scattered from atmospheric aerosols. Using this technology, the detection of turbulent disturbances physically created by the passage of aircraft through the atmosphere can be developed. Traditional aircraft detection methods include pulse-Doppler radar systems that measure the frequency shift of radio frequency signals reflected off the skin of an aircraft. This method is currently being defeated by stealth technology. In a military application, light detection and ranging turbulent detection would be able to spot the unconcealable turbulent disturbances caused by aircraft, thereby defeating current stealth technology. In addition, research suggests that turbulent wake generators (i.e., aircraft) have unique turbulent energy signatures. Using the principles of turbulent self-preservation, models can be created to predict both the Reynolds stress and momentum-defect signatures of aircraft as a function of their altitude and airspeed. By using statistical analysis of the turbulent wake, the Reynolds stress or momentum defect can be compared with computed models to determine the identity of the aircraft being targeted. The theoretical and practical manner in which this can be achieved is discussed, and examples of aircraft ranging from the F-15, F-16, F-18, and B-52 are provided.

The wake vortex development in the near field and extended near field behind a four engined large transport aircraft model fitted with an active winglet is presented. A detailed wind-tunnel investigation is conducted using a half-model focusing on the high-lift ease of a typical approach configuration at a Reynolds number of 0.5 x [10.sup.6] based on the wing mean aerodynamic chord and an angle of attack of 6.5 deg. The flowfield is observed using advanced hotwire anemometry mainly focusing on the crossflow plane at 5.6 spans downstream of the model. Based on the time-dependent velocity components the wake flowfield is analyzed by distributions of mean vorticity, turbulence intensities, and spectral densities. Seven main vortical structures dominating the near-field wing vortex sheet roll up and merge to form the remaining trailing vortex in the extended near field. By the use of oscillating winglet flaps the velocity fluctuations at the core region of the remaining vortex are significantly influenced. Distinct narrowband concentrations of turbulent kinetic energy can be found for the farthest downstream plane documenting the presence of the disturbances generated by the winglet flaps which may result in an amplification of inherent far-field instabilities. The frequencies at which these narrowband energy concentrations occur are, on the one hand, dependent on the oscillation frequency of the winglet flaps; on the other hand, there are also independent energy concentrations within the frequency bands associated with wake instabilities.

This paper is prompted by a recent numerical study (Lewin, G. C., and Haj-Hariri, H., 'Modelling Thrust Generation of a Two-Dimensional Heaving Airfoil in Viscous Flow,' Journal of Fluid Mechanics, Vol. 492, Oct. 2003, pp. 339-362) that shows that for a two-dimensional (2-D) elliptic airfoil undergoing prescribed heaving motion in a viscous fluid, both leading-edge vortices and trailing-edge vortices contributed to the formation of the wake structures. However, an earlier dye-visualization study (Lai, J. C. S., and Platzer, M. F., 'Jet Characteristics of a Plunging Airfoil,' AIAA Journal, Vol. 37, No. 12, 1999, pp. 1529-1537) on a heaving NACA 0012 airfoil appears to show that the wake structures were derived from trailing-edge vortices only. The dissimilarity in the two studies remains unclear because there is no corresponding experimental data on a 2-D heaving elliptic airfoil. In this study, digital particle image velocimetry technique was used to investigate the wake-structure formation of a 2-D elliptic airfoil undergoing simple harmonic heaving motion. For the range of flow conditions investigated here, our results show that the type of wake structures produced is controlled by when and how the leading-edge vortices interact with the trailing-edge vortices.

The effect of acoustic excitation amplitude on boundary layer and wake development for a NACA 0025 airfoil was studied experimentally at low Reynolds numbers. Flow characteristics were investigated with hot-wire anemometry, surface pressure measurements, and flow visualization. A laminar boundary layer separation occurs on the upper surface of the airfoil, forming a separated shear layer, for all situations examined. When the flow is excited at the frequency matching the frequency of the most amplified disturbance in the separated shear layer, natural shear layer disturbances lock onto the excitation frequency and transition is promoted. In the case when the separated shear layer fails to reattach, an increase of the excitation amplitude above a minimum threshold eventually results in shear layer reattachment. The results suggest that an increase of the excitation amplitude not only advances the location of reattachment but also delays boundary layer separation, thereby reducing the extent of the separation region. As a consequence, excitation results in narrowing of the wake and diminishment of the length scales and coherence of the organized wake structures. However, this effect is eventually limited, and a maximum effective excitation amplitude can be identified. It is also shown that an increase of the excitation amplitude has a broadly similar effect on flow patterns to an increase of the chord Reynolds number. DOI: 10.2514/1.25439

A numerical study of the wake structure behind a three-dimensional hill is presented. Large eddy simulation is used to investigate the source of the wake vorticity and the variations of the wake structure with [U.sub.[infinity]]/Nh. Here, [U.sub.[infinity]] is the freestream velocity, N is the Brunt-Vaisala frequency, and h is the hill height. The simulation shows that when [U.sub.[infinity]]/Nh ~ 1, the wake vorticity originates from baroclinic processes at the initial stage of the wake development. Even though contributions from friction effects increase during the later stage of the wake development, the production of the wake vorticity from baroclinicity remains significant. The wake consists of a pair of vertically oriented vortices, which form vertical lateral boundaries for the wake. When [U.sub.[infinity]]/Nh [less than or equal to] 0.2, the vortical structure of the wake has a horseshoe shape, and the change of the wake width with height reflects the change of the radius of the hill with height. All evidence indicates that the wake at [U.sub.[infinity]] [less than or equal to] 0.2 follows the two-dimensional wake assumption, and thus the wake vorticity comes mainly from surface friction instead of baroclinicity.

An experimental program whose objective was an improved understanding of the effects of intermittent turbulence shed by a propeller on the aerodynamics of heat-exchanger installations at low Reynolds numbers is reported. The primary application is unmanned high-altitude atmospheric science aircraft, forming part of Old Dominion University's effort to develop technology applicable to ERAST (Environmental Research Aircraft and Sensor Technology) class aircraft. Wind-tunnel results indicate that the propeller wake affects, but does not eliminate, laminar separation bubbles on a representative wing. Rather, the bubbles reform between wake passages, limiting the drag penalty caused by increased skin friction. Further, there appears to be a possibility of an enhancement in performance of suitably designed ducted heat-exchanger installations.

The off-surface aerodynamic characteristics of a wing in ground effect are investigated using a number of methods including laser Doppler anemometry and particle image velocimetry. The study focuses on two aspects of the flow: turbulent wake and edge vortex. These features are closely associated with the behavior of the aerodynamic force in ground effect. The size of the wake increases in proximity to the ground. A downward shift of the path of the wake is also observed. Discrete vortex shedding is seen to occur behind the wing. As the wing height is reduced, separation occurred on the suction surface of the wing, and the spanwise vortex shedding is found to couple with a flapping motion of the wake in the transverse direction. An edge vortex is also observed off the edge of the end plate of the wing, which contributes to force enhancement and helps to define the force behavior in the force enhancement region. The rate of change in the downforce vs height curve is linked to the strength of the edge vortex. The vortex breakdown signals a slowdown in the force enhancement. When the maximum downforce height is reached, the edge vortex breaks down completely.

The vortices that trail from the wingtips of a large aircraft provide a significant hazard to an aircraft that follows in its wake. The objective is to contribute to the understanding of these vortices by identifying the conditions where an axial velocity in excess of the freestream value will be generated in the core of a trailing vortex. The axial velocity near the core of a trailing vortex was measured using a triple-sensor hot-wire probe and compared with measured values of vortex circulation strength. The vortex was generated in a wind tunnel using a NACA 0015 wing model with a semispan aspect ratio of 0.80. A linear relationship between the axial velocity and a nondimensional circulation parameter is indicated. For small values of the circulation parameter, the axial velocity shows a velocity deficit. An increase in circulation strength results in a progressive increase in axial velocity, which can exceed the freestream velocity by as much as 70%. The results indicate that the magnitude of the axial velocity is sensitive to the two end-cap configurations, flat and rounded, that were tested. It was found that scaling the circulation parameter by the vortex radius collapses the data from both configurations to a single linear curve.

Inviscid, three-dimensional, unsteady flow computational results for the aerodynamic interaction between a rotor and a hemispherical-cylinder airframe are presented. A multiblock strategy is used, with separate grids for the fuselage and the rotor blades. An overset grid methodology is used to communicate the flow information between the grids. In each grid a second-order temporally accurate, fifth-order spatially accurate, implicit finite volume scheme is employed to solve the Euler equations. The vortical wake generated by the blade is captured from first principles. The vorticity patterns that impinge on the airframe are analyzed, and the surface-pressure distributions on the airframe are compared with experimental values.

A numerical inviscid vortex method is employed to study the unsteady, two-dimensional, incompressible flow that occurs during an airfoil--vortex interaction in the vicinity of a ground plane. The airfoil bound vorticity is modeled using a panel method with linear piecewise-continuous vorticity distribution. The vortex is considered to be a point vortex, and the ground effect is obtained using the method of images. Point vortices are also generated at the airfoil trailing edge to ensure that circulation is conserved and that the Kutta condition is satisfied. All vortices are convected using a first-order Lagrangian time-marching scheme. After code validation, numerical results for the airfoil--vortex interaction in ground effect reveal that the loading on the airfoil is affected by the thickness, angle of attack, and height of the airfoil above the ground, as well as the vortex strength, direction of rotation, and distance to the airfoil. The wake evolution is nonlinear and strongly influenced by the interaction, whereas the pressure distribution on the ground presents steep adverse gradients.

Flight procedures that will decrease the following distance between successive aircraft have been considered as a means of alleviating pressures on airport capacity while still conforming to conventional air traffic management rules. A major limitation to the implementation of such procedures is perceived to be an associated increase in the possible severity of encounters with the wakes of nearby aircraft. Modeling efforts have focused on quantifying such interactions in terms of structural loading or degradation in handling and have established that helicopters respond in a fundamentally different way than fixed-wing aircraft to encounters with wake vortices. This paper suggests that, particularly at the low forward speeds typical of terminal phase operations, the mutually induced distortion of the helicopter's own wake and the wake of the interacting aircraft strongly influences the aerodynamics of the helicopter rotor. Results from numerical simulations suggest that neglecting this aspect of the interaction, as has commonly been done in the past, can indeed be valid at high helicopter forward speeds, but can lead to significant misrepresentation of the severity of wake encounters under terminal flight conditions.

The effects of an above-exit circular disk on the flow structures, turbulence properties, and mixing characteristics of swirling jets separated by a large diameter blockage disk at jet exit are studied by using the smoke-wire flow-visualization technique and a two-component laser Doppler velocimeter. A carbon-dioxide tracer-gas experiment is conducted to examine the mixing capability of the flows modulated by the control disk. The smoke-streak flow patterns, characteristic lengths, velocity vectors, streamlines, velocity distributions, turbulence properties, and concentration distributions of the carbon dioxide are presented and discussed. Topological techniques are applied to analyze and interpret the flow patterns. With the variation of the distance between the control and blockage disks and the central jet Reynolds number, the bubble structures in the near field display complex behaviors of extension, escape, and encapsulation. Under the influence of the control disk, the central jet deflects radially, merges with the swirling annular flow, then turns downstream to form a recirculation bubble in the control-disk wake. This arrangement creates significant momentum exchange and drastic turbulence fluctuations. The drastic increase of the turbulence fluctuations expedites mixing of the fluids issued from the central and annular jets, particularly in the bubble encapsulation and gap flow regimes.

A new model of the drag force generated by a freely decelerating bluff body is presented. The model is based, mainly, on the premise that the wake of an object accelerating downwind in a moving fluid is identical to that of the same object decelerating in the fluid at rest. After arguing for the drag of a wind drifter to depend only on a power function of its speed relative to the wind, a Galilean transformation is used to provide a formula for decelerating body drag of the form [F.sub.D] ~ [V.sup.[beta]]. The value of exponent [beta] is dependent on the amount of external force applied to the body, as well as on its initial and final drag coefficients and its initial speed. By implication, this exponent depends on the specific history of the motion. Applications to powered and unpowered vehicles trailing a parachute or any other high-drag devices are presented and discussed. The model is also shown to compare favorably with parachute-test drop data.

A New parametric wake vortex transport and decay model is proposed that predicts probabilistic wake vortex behavior as a function of aircraft and environmental parameters in real time. The probabilistic two-phase wake vortex decay model (P2P) accounts for the effects of wind, turbulence, stable stratification, and ground proximity. The model equations are derived from the analytical solution of the spatiotemporal circulation evolution of the decaying potential vortex and are adapted to wake vortex behavior as observed in large-eddy simulations. Vortex decay progresses in two phases, a diffusion phase followed by rapid decay. Vortex descent is a nonlinear function of vortex strength. Probabilistic components account for deviations from deterministic vortex behavior inherently caused by the stochastic nature of turbulence, vortex instabilities, and deformations, as well as uncertainties and fluctuations that arise from environmental and aircraft parameters. The output of P2P consists of confidence intervals for vortex position and strength. To assign a defined degree of probability to the predictions reliably, the model design allows for the continuous adjustment of decay parameters and uncertainty allowances, based on a growing amount of data. The application of a deterministic version of P2P to the Memphis wake vortex database yields favorable agreement with measurements.

The theoretical basis and the numerical implementation of free-vortex filament methods are reviewed for application to the prediction and analysis of helicopter rotor wakes. The governing equations for the problem are described, with a discussion of finite difference approximations to these equations and various numerical solution techniques. Both relaxation and time-marching wake solution techniques are reviewed. It is emphasized how the careful consideration of stability and convergence (grid-independent behavior) are important to ensure a physically correct wake solution. The implementation of viscous diffusion and filament straining effects are also discussed. The need for boundary condition corrections to compensate for the inevitable wake truncation are described. Algorithms to accelerate the wake solution using velocity field interpolation are shown to reduce computational costs without a loss of accuracy. Several challenging examples of the application of free-vortex filament methods to helicopter rotor problems are shown, including multirotor configurations, flight near the ground, maneuvering flight conditions, and descending flight through the vortex ring state.

The origin of the tip vortex is described and important properties, such as core radius and circulation, as a function of these parameters, are calculated. Despite the fact that a large amount of computational and experimental work on the rotor wake has been published, little of a quantitative nature is known about the origin of the main component of the rotor wake, the tip vortex, as a function of the rotor speed, rotor blade geometry, and angle of attack. Experimental data by several workers have revealed little dependence on the Reynolds number. This lack of dependence on Reynolds number will hold only if the flow remains substantially unseparated, although it may be noted that drag coefficients of bluff bodies are remarkably constant over wide range of Reynolds number. The rotor blade is assumed to be of large aspect ratio, that is, to leading order in the blade aspect ratio, it appears infinite in length. In this limit case, the analytical solution for the bound circulation on a fixed wing is shown to be extendable to the rotary wing case when the blade is viewed as a lifting line. Lifting-surface results are obtained also, and these compare well with the lifting-line results for the bound circulation.

An accurate simulation of rotor aerodynamics has been a challenging problem for the computational fluid dynamics community, largely because numerical diffusion in the typically coarse grids away from the blades seriously degrades the prediction of the returning wakes. A numerical scheme is proposed to improve the reconstruction stage of a MUSCL approach in the framework of a finite volume formulation. A fourth-order upwind scheme, based on the interpolation of the local characteristic variables in the reconstruction stage, is applied instead of the conventional third-order upwind-biased scheme. Different slope limiters are applied to different characteristic fields to reduce the numerical dissipation of entropy and shear waves. We have implemented the proposed scheme into the TURNS code. Numerical results for a hover case show a significant improvement in the wake capturing capability over the baseline TURNS code.

Computational fluid dynamics (CFD) is used to create an extensive aerodynamic ship-airwake database for a ship-specific flight simulator to train helicopter pilots for landing and takeoff from various points on a ship. The technology has been embedded into CAE's Merlin simulator and applied to six different ships of the Royal Navy. In achieving this, important scientific and engineering milestones are crossed, including high geometric fidelity through OEM's (Original Equipment Manufacturer) CAD description of the ships, three-dimensional CFD calculations, comparison to experimental data, and, finally, integration, testing, and acceptance into the flight simulator. The database is first built by solving the three-dimensional inviscid (Euler) flow around the complete ship, using Finite Element Navier-Stokes Analysis Package, a proprietary code. And a number of wind speeds and directions, corresponding to realistic approach scenarios, are analyzed. The CFD results are then compared to a set of experiments carried out for airwakes around the Canadian Patrol Frigate and made available only after the calculations were done. The calculations prove to be in excellent agreement with the data throughout its range. For the purpose of integration into the simulator, the CFD solutions are interpolated onto a Cartesian grid, which is then implemented in a look-up table for the Merlin CDS, after being enhanced with turbulence models. The paper then closes with the acceptance procedure the simulator is passed through.

Steady-state Reynolds-averaged Navier--Stokes computations are presented for a range of UH-60A model-rotor test cases in hover. The computations are designed to assess grid-related effects on the numerical results and employ 1) structured overset grids with high resolution on the rotor blades, 2) a systematic variation of grid resolution in the rotor wake, and 3) a systematic variation of outer-boundary locations. Computed rotor performance values agree very well with experimental measurements and show little sensitivity to either grid resolution or outer-boundary locations. However, the computations uniformly overpredict the blade sectional thrust near the rotor tip. This overprediction of blade tip thrust is explained by an analysis of the circulation distribution in the computed rotor wake system.

The use of high-order, first-principles-based Euler/Navier-Stokes methodologies for rotary-wing wake capturing is addressed. Fifth/seventh-order spatially accurate essentially nonoscillatory methodologies were implemented and their benefits in capturing wake vortices were assessed. The use of high-order solution reconstruction, in conjunction with the use of overset grids to resolve selectively regions of interest such as concentrated vorticity, was also explored. The tip vortex off of a NACA0015 wing was captured using this methodology, and the tip vortex characteristics are compared with experiments. The relative merits of fifth- and seventh-order schemes are assessed for capturing wake vorticity.

A numerical method has been developed for predicting the complex vortex wake for a helicopter rotor in hover and in forward flight. The method is based on the solution of the three-dimensional, compressible Euler equations expressed in an arbitrary Lagrangian Eulerian reference frame. A second-order accurate discontinuous Galerkin finite element method is used to discretize the governing equations on a hexahedral mesh. Unstructured, local mesh refinement is performed to enable prediction of the structure of the vortex wake. The capabilities of this computational fluid dynamics method are demonstrated by simulations of the flow around the Caradonna-Tung helicopter rotor in hover and simulations of the flow around the Operational Loads Survey helicopter rotor in forward flight. Accuracy of the method is assessed through comparison with wind-tunnel data, if available. Special attention is given to a rotor trimming procedure developed at the National Aerospace Laboratory/NLR and to the grid adaptation algorithm.

The regions over which steady and unsteady rotor wake patterns exist have been examined numerically for one-and two-bladed rotors. A lifting-line theory is used to model each rotor blade, and a time-stepping vortex method is used to calculate the tip-vortex motion. Tip-vortex geometries for a rotor in hover and vertical climb have been computed. The velocities within the first few turns of the wake have been obtained, and the results indicate a periodic nature of the wake in both time and space. Temporal aperiodicity is observed beyond the first few turns of the tip vortex for hovering rotors, whereas for rotors in sufficient climb the entire wake is steady relative to the blades and both spatially and temporally periodic. Both Euler and Adams--Moulton schemes have been tested for accuracy. The numerical results are shown to be independent of both time step and spatial grid below clearly defined values. Rotor thrust calculations have been performed and the effect of the lower wake on rotor thrust determined. The computed thrust values show some variation because the lower wake remains unsteady. The results for the wake geometry are analyzed and compared to previous predictions and two distinct sets of experimental results.

This paper discusses the problem of wake vortices shed by commercial aircraft. It presents a consolidated European view on the current status of knowledge of the nature and characteristics of aircraft wakes and of technical and operational procedures of minimizing and predicting the vortex strength and avoiding wake encounters. Methodological aspects of data evaluation and interpretation, like the description of wake ages, the characterization of wake vortices, and the proper evaluation of wake data from measurement and simulation, are addressed in the first part. In the second part an inventory of our knowledge is given on vortex characterization and control, prediction and monitoring of vortex decay, vortex detection and warning, vortex encounter models, and wake-vortex safety assessment. Each section is concluded by a list of questions and required actions which may help to guide further research activities. The primary objective of the joint international efforts in wake-vortex research is to avoid potentially hazardous wake encounters for aircraft. Shortened aircraft separations under appropriate meteorological conditions, whilst keeping or even increasing the safety level, is the ultimate goal. Reduced time delays on the tactical side and increased airport capacities on the strategic side will be the benefits of these ambitious ventures for the air transportation industry and services.

The three-dimensional asymmetric turbulent near-wake behind an infinitely swept wing with GAW(2) airfoil has been investigated at low speeds. The near-wake in the present study is asymmetric because the boundary layers on the top and bottom surfaces of the model develop under different streamwise pressure gradients. Distributions of mean velocity, three turbulent normal stresses, and two important Reynolds shear stresses have been measured using hot-wire anemometry. The profiles of mean velocity and Reynolds shear stress exhibit asymmetry near the trailing edge and seem to have become symmetric within a short distance of 60 trailing edge momentum thicknesses. Results of computation using K-[epsilon] turbulence model with a simple scheme to predict the near-wake behind the swept wing have also been presented and compared with the experimental data. The agreement of the predicted mean flow development with the experiment is fair considering the simplicity of the scheme.

Results demonstrate that ambient temperature has a significant role in influencing the most amplified wavelength of the Crow instability and the wake voices lifetime. Data further point out that the instability of most amplified wavelength decreases as the turbulence intensity increases.

Results indicate that under enhanced quasi-two-dimensional features, the large-scale flow organization in the cylinder wake plays a dominant role in the production of kinetic energy and transport. Data further point out that wake/boundary-layer interaction process affects vortices and kinetic energy production in the lower-half wake.

The authors describe preliminary experiments measuring the near wake effects on a axisymmetric bodies. Topics include turbulence analysis, drag wakes, and propeller-flow.

Unsteady wake effects on detailed heat transfer coefficient and film cooling effectiveness distributions from a gas turbine blade with film cooling are obtained using a transient liquid crystal technique. Tests were performed on a five-blade linear cascade at a axial chord Reynolds number of 5.3 x [10.sup.5] at cascade exit. Upstream unsteady wakes are simulated using a spoke-wheel type wake generator. The test blade has three rows of film holes on the leading edge and two rows each on the pressure and suction surfaces. Air and C[O.sub.2] were used as coolants to simulate different coolant-to-mainstream density ratio effect. Coolant blowing ratio for air injection is varied from 0.8 to 1.2 and is varied from 0.4 to 1.2 for C[O.sub.2]. Results show that Nusselt numbers for a film-cooled blade are much higher compared to a blade without film injection. Particularly, film injection causes earlier boundary layer transition on the suction surface. Unsteady wakes slightly enhance Nusselt numbers but significantly reduce film cooling effectiveness on a film-cooled blade compared with a film-cooled blade without wakes. Nusselt numbers increase slightly but film cooling effectiveness increases significantly with an increase in blowing ratio for C[O.sub.2] injection. Higher density coolant (C[O.sub.2]) provides higher effectiveness at higher blowing ratios (M = 1.2) whereas lower density coolant (Air) provides higher effectiveness at lower blowing ratios (M = 0.8).

The interaction between wakes of an adjacent rotor-stator or stator-rotor blade row pair in an axial turbomachine is known to produce regular spatial variations in both the time-mean and unsteady flow fields in a frame relative to the upstream member of the pair. This paper examines the influence of such changes in the free-stream disturbance field on the viscous losses of a following blade row. Hot-wire measurements are carried out downstream of the outlet stator in a 1.5-stage axial compressor having equal blade numbers in the inlet guide vane (IGV) and stator rows. Clocking of the IGV row is used to vary the disturbance field experienced by the stator blades; the influence on stator wake properties is evaluated. The magnitude of periodic fluctuations in ensemble-averaged stator wake thickness is significantly influenced by IGV wake-rotor wake interaction effects. The changes in time-mean stator losses appear marginal.

The interaction of wakes shed by a moving blade row with a downstream blade row causes unsteady flow. The meaning of the free-stream stagnation pressure and stagnation enthalpy in these circumstances has been examined using simple analyses, measurements, and CFD. The unsteady flow in question arises from the behavior of the wakes as so-called negative jets. The interactions of the negative jets with the downstream blades lead to fluctuations in static pressure, which in turn generate fluctuations in the stagnation pressure and stagnation enthalpy. It is shown that the fluctuations of the stagnation quantities created by unsteady effects within the blade row are far greater than those within the incoming wake. The time-mean exit profiles of the stagnation pressure and stagnation enthalpy are affected by these large fluctuations. This phenomenon of energy separation is much more significant than the distortion of the time-mean exit profiles that is caused directly by the cross-passage transport associated with the negative jet, as described by Kerrebrock and Mikolajczak. Finally, it is shown that if only time-averaged values of loss are required across a blade row, it is nevertheless sufficient to determine the time-mean exit stagnation pressure.

The structure and behavior of wakes from a transonic compressor rotor and their effect on the loading and performance of the downstream stator have been investigated experimentally. The rotor was 23.25 inches in diameter with a measured tip Mach number of 1.23 and a pressure ratio of 1.66. Time and space-resolved measurements have been completed of the rotor and stator outflow, as well of the pressure distribution on the surface of the stator blades. It was found that the wakes from this rotor have large flow angle and flow Mach number variations from the mean flow, significant pressure fluctuations, and a large degree of variation from hub to tip. There was a significant total pressure defect and practically no static pressure variation associated with the stator wakes. Wakes from the rotor exist nearly undiminished in the exit flow of the stator and decay in the annular duct behind the stator. The pressure at all points along the chord over each of the stator blades' surfaces fluctuated nearly in phase in response to the rotor wakes, that is the unsteady chordwise pressure distribution is determined mainly by the change in angle of incidence to the blade and not by the local velocity fluctuations within the passage. The unsteady forces on the stator blades, induced by the rotor wakes, were as high as 25 percent of the steady forces, and lagged the incidence of the wakes on the leading edge by approximately 180 deg at most radii.

Measurements of the rotor aerodynamics, far-field acoustics, wake geometry and blade motion, which were obtained from the higher harmonic control aeroacoustic rotor test done at Duits-Nederlandse Wind Tunnel, were used to validate and enhance the prediction of the blade-vortex interaction (BVI) noise. This has resulted to a number of improvements for the BVI model, which include the restructuring of the wake, the modification of the core size and the integration of measured blade motion to the calculations.

The development of the unsteady suction side boundary layer of a highly loaded LP turbine blade has been investigated in a rectilinear cascade experiment. Upstream rotor wakes were simulated with a moving-bar wake generator. A variety of cases with different wake-passing frequencies, different wake strength, and different Reynolds numbers were tested. Boundary layer surveys have been obtained with a single hotwire probe. Wall shear stress has been investigated with surface-mounted hot-film gages. Losses have been measured. The suction surface boundary layer development of a modem highly loaded LP turbine blade is shown to be dominated by effects associated with unsteady wake-passing. Whereas without wakes the boundary layer features a large separation bubble at a typical cruise Reynolds number, the bubble was largely suppressed if subjected to unsteady wake-passing at a typical frequency and wake strength. Transitional patches and becalmed regions, induced by the wake, dominated the boundary layer development. The becalmed regions inhibited transition and separation and are shown to reduce the loss of the wake-affected boundary layer. An optimum wake-passing frequency exists at cruise Reynolds numbers. For a selected wake-passing frequency and wake strength, the profile loss is almost independent of Reynolds number. This demonstrates a potential to design highly loaded LP turbine profiles without suffering large losses at low Reynolds numbers.

Thermal and chemical nonequilibrium effects in near-wake ionizing flows were investigated. The impact of real-gas effects on the structure of near wakes was studied and an idealized-gas model that considers real-gas effects was developed. The study was limited to flows around simple sphere-cone geometries and Apollo-like re-entry capsules. Findings showed that near-wake flows can be assumed to be both thermally and chemically frozen.

Direct numerical simulations of a vortex pair embedded in a stable atmospheric boundary layer are shown. The contributions of different crosswind conditions are investigated. These calculations illustrate the formation of secondary vortices for the range of Reynolds number (3.77 X 10(super 3)

Analytical/numerical matching (ANM) is a boundary element method that combines a low-resolution global numerical solution with a high-resolution local solution to offer a unified aerodynamic solution for lifting-surface problems. ANM facilitates the calculation of aerodynamic loading and the related circulation distribution. It uses overlapping smoothed doublets and local corrections to estimate the doublet strength distribution along the airfoil chord. High-resolution calculations, rapid numerical convergence, and flexibility in terms of wake geometry are some of its advantages.

An efficient high quality solution-adapted unstructured grid resolves shock waves and stagnation points in a viscous flowfield solution. Tracking selected streamlines resolve detached viscous features like wakes by treating them as embedded boundaries. This gives a natural framework for directional refinement and automatic alignment with the wake. Two dimensional cases with multiple viscous wakes and strong inviscid features show improved resolution of the flowfields.

Extensive wake surveys of the peculiar drag characteristics of low Reynolds number airfoil flows reveal the prominent role of the laminar separation bubble and airfoil geometry in spanwise drag variation. The amplitudes of the oscillations decrease with the increase in Reynolds number, and the spanwise locations of the oscillations shift with changes in the angle of attack. The largest fluctuations are observed for low angles of attack and at the lowest point of the low drag range. A weak agreement is observed in profile drag measurements between different wind-tunnel facilities.

The vortex lattice method is used to analyze the coupled flap-lag-torsion aeroelastic behavior of multibladed hingeless rotors in the hovering flight condition. A finite element beam model is employed to to take into account the various aerodynamic parameters, such as centrifugal forces and aerodynamic loads. The results indicate that the steady equilibrium and aerodynamic stability of hovering rotors are significantly affected by the three-dimensional tip and realistic wake effects.

Development of steady and periodic unsteady wake flows downstream of stationary and rotating cylindrical rods within a curved channel under zero longitudinal pressure gradient is theoretically and experimentally investigated. Wake quantities such as the mean velocity and turbulent fluctuations in longitudinal and lateral directions, as well as the turbulent shear stress, are measured. For the nondimensionalized velocity defect, affine profiles are observed throughout the flow regime. Based on these observations and using the transformed equations of motion and continuity, a theoretical frame work is established that generally describes the two-dimensional curvilinear wake flow. To confirm the theory, development of steady and periodic unsteady wakes in the above curved channel are experimentally investigated. The detailed comparison between the measurement and the theory indicates that the complex steady and unsteady wake flows are very well predicted.

The kappa-epsilon-gamma model resolves the two-dimensional momentumless wake anomaly, in which a jet and a wake coexist. This model finds the free shear flows by the standard kappa-epsilon model. The turbulent structure is satisfactorily analysed, as shown by computational results. The prediction performance of the model is good in comparison with the kappa-epsilon model.

Use of high-image-density particle image velocimetry (PIV) allows characterization of the instantaneous structure of wake and wake-blade interactions in a simulated rotating machine. The distribution of vorticity over an entire plane within the pumping system is related to the instantaneous pressure source terms in the wake of the impeller. Comparison of instantaneous and ensemble-averaged vorticity contours shows that limited ensemble-averaging can produce a substantial reduction in vorticity levels associated with the instantaneous pressure source terms. When the wake from the impeller interacts with a stationary diffuser blade, the instantaneous processes of flow separation and reattachment can be effectively characterized using combinations of instantaneous streamline patterns and contours of constant vorticity. Moreover, active control of the inflow into the pumping system allows substantial modification of these vorticity distributions.

A unique, actively controlled pumping system allows independent control of inflow and impeller perturbations, as well as the phase shift between them. The basic configurations of an impeller and an impeller-diffuser blade system have been investigated with the objective of manipulating the spectral content of the unsteadiness of the near-wake at the impeller discharge. Substantial alteration of the discrete spectral components can be attained. A central feature is the generation of a number of nonlinear interaction components, corresponding to sum and difference frequencies, of the forcing- and blade passing-frequencies and their harmonics. With proper choice of perturbation conditions, it is possible to attenuate the inflow perturbation, as well as to alter the magnitudes of the blade passing component and its harmonics.

An analysis of an elliptic cross-section tip reveals its similarity to a right circular cone at the point where the polynomial is tangential to the generator. The side force coefficient displays square waves caused by vortex breakaway, which permits vortex crossover and controls vortex wake asymmetry. The tip roll angle which varies smoothly with vortex wake asymmetry can be used as a control device in lighter aircraft.

A nonintrusive method requires near wake data to study the fluid dynamics involved in supersonic base flowfields. The method is a supersonic flow-aligned circular cylinder as the test model and yields pressure distributions, average velocities and Reynold's shear stresses for a given flowfield. The test results imply that, for a near wake flowfield, the mean static pressure pattern is nearly the same across the base.

Boundary layers on turbomachinery blades develop in a flow that is periodically disturbed by the wakes of upstream blade cascades. These wakes have a significant effect upon laminar-turbulent boundary-layer transition. In order to study these effects, detailed velocity measurements using hot-wire probes were performed within the boundary-layer of a plate in flow periodically disturbed by wakes produced by bars moving transversely to the flow. The measurements were evaluated using the ensemble-averaging technique. The results show how the wake disturbance enters the boundary-layer and leads to a turbulent patch, which grows and is carried downstream. In favorable pressure gradients, transition due to wake turbulence occurred much earlier than predicted by linear stability theory. Between two wakes, laminar becalmed regions were observed far beyond the point at which the undisturbed boundary-layer was already turbulent.

The rotor wake/stator interaction noise prediction method compares and calibrates a semiempirical two-dimensional fan wake model with two fan rigs and a full scale engine. This model has an analytical stator response function and noise calculation. The stator response function is divided into two-dimensional steady flow calculated along streamlines linked to a noise calculation in a constant area annular duct with mean axial flow. This is compared against data from an advanced ducted propeller rig, and predictions are calibrated using its data.

A finite state aeroelastic model issued for rotor design optimization is derived by coupling a generalized dynamic wake with blade finite elements. This coupled system of ordinary differential equations is used to directly compute dynamic response, and design derivatives and aerodynamic change in a rotorcraft. Data is verified by comparison with previously acquired data.

A general unsteady subsonic method, developed for computation of wing body aerodynamics and including wake effects on the body, simulates the wing-body configuration to represent realistic aircraft with various external store combinations for aeroelastic applications. This method applies correct unsteady boundary conditions on body surfaces, adequately accounts for body wake effect and accurately assesses wing-body influence. Code verification is done by comparison to existing computed and measured data.