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Showing papers in "AIAA Journal in 2009"


Journal ArticleDOI
TL;DR: The main goal of the present paper is to publish the full model and release it to the research community so that it can continue to be further validated and possibly extended or improved.
Abstract: A new correlation-based transition model has been developed, which is built strictly on local variables. As a result, the transition model is compatible with modern computational fluid dynamics techniques such as unstructured grids and massively parallel execution. The model is based on two transport equations, one for intermittency and one for a transition onset criterion in terms of momentum-thickness Reynolds number. A number of validation papers have been published on the basic formulation of the model. However, until now the full model correlations have not been published. The main goal of the present paper is to publish the full model and release it to the research community so that it can continue to be further validated and possibly extended or improved. Included in this paper are a number of test cases that can be used to validate the implementation of the model in a given computational fluid dynamics code. The authors believe that the current formulation is a significant step forward in engineering transition modeling, as it allows the combination of transition correlations with general-purpose computational fluid dynamics codes. There is a strong potential that the model will allow the first-order effects of transition to be included in everyday industrial computational fluid dynamics simulations.

1,073 citations



Journal ArticleDOI
TL;DR: In this paper, the authors present the results of a parametric experimental investigation aimed at optimizing the body force produced by single dielectric barrier discharge plasma actuators used for aerodynamic flow control.
Abstract: This paper presents the results of a parametric experimental investigation aimed at optimizing the body force produced by single dielectric barrier discharge plasma actuators used for aerodynamic flow control. A primary goal of the study is the improvement of actuator authority for flow control applications at higher Reynolds number than previously possible. The study examines the effects of dielectric material and thickness, applied voltage amplitude and frequency, voltage waveform, exposed electrode geometry, covered electrode width, and multiple actuator arrays. The metric used to evaluate the performance of the actuator in each case is the measured actuator-induced thrust which is proportional to the total body force. It is demonstrated that actuators constructed with thick dielectric material of low dielectric constant produce a body force that is an order of magnitude larger than that obtained by the Kapton-based actuators used in many previous plasma flow control studies. These actuators allow operation at much higher applied voltages without the formation of discrete streamers which lead to body force saturation.

459 citations


Journal ArticleDOI
TL;DR: In this paper, the authors used a large horizontal microphone array, positioned at a distance of about one rotor diameter from the turbine, to locate and quantify the noise sources in the rotor plane and on individual blades.
Abstract: Acoustic field measurements were carried out on a 94-m-diam three-bladed wind turbine with one standard blade, one blade with trailing-edge serrations, and one blade with an optimized airfoil shape. A large horizontal microphone array, positioned at a distance of about one rotor diameter from the turbine, was used to locate and quantify the noise sources in the rotor plane and on the individual blades. The acoustic source maps show that for an observer at the array position, the dominant source for the baseline blade is trailing-edge noise from the blade outboard region. Because of convective amplification and directivity, practically all of this noise is produced during the downward movement of the blade, which causes the typical swishing noise during the passage of the blades. Both modified blades show a significant trailing-edge noise reduction at low frequencies, which is more prominent for the serrated blade. However, the modified blades also show tip noise at high frequencies, which is mainly radiated during the upward part of the revolution and is most important at low wind speeds due to high tip loading. Nevertheless, average overall noise reductions of 0.5 and 3.2 dB are obtained for the optimized blade and the serrated blade, respectively.

375 citations


Journal ArticleDOI
TL;DR: In this paper, the potential of microramp sub-boundary-layer vortex generators for flow control in supersonic engine inlets is investigated, focusing on the ability of these devices to beneficially affect oblique shockwave/ boundary-layer interactions.
Abstract: The potential of microramp sub-boundary-layer vortex generators for flow control in supersonic engine inlets is investigated. In particular, the study focuses on the ability of these devices to beneficially affect oblique shock-wave/ boundary-layer interactions. Experiments have been conducted at Mach 2.5 to determine the nature of flow controlled by microramps and to investigate their ability to delay separation in a reflected shock interaction. Various ramp heights between 30 and 90% of the boundary-layer thickness were investigated. The details of the vortical flow generated by such devices were identified. The general flow features were found to scale with device height and it is suggested that smaller devices need to be placed closer to the expected adverse pressure gradients. When applied to a separated oblique shock-wave/boundary-layer interaction generated with a 7 degree wedge, microramps were not able to completely eliminate flow separation, although they were shown to break up separated regions. Other performance indicators across the shock-wave/boundary-layer interaction were also improved through the application of the devices.

303 citations


Journal ArticleDOI
TL;DR: In this article, the dynamics of unstart of a floor-mounted inlet/isolator model in a Mach 5 flow were investigated experimentally, where the inlet section contains a 6-deg compression ramp, and the isolator is a rectangular straight duct that is 25.4mm high by 50.8mm wide by 242.3mm long.
Abstract: DOI: 10.2514/1.40966 The dynamics of unstart of a floor-mounted inlet/isolator model in a Mach 5 flow are investigated experimentally. The inlet section contains a 6-deg compression ramp, and the isolator is a rectangular straight duct that is 25.4-mm high by 50.8-mm wide by 242.3-mm long. Measurements made include 8-kHz schlieren imaging and simultaneous fast-response wall pressures along the length of the inlet/isolator. Unstart is initiated by deflecting a flap at the downstream end of the isolator. The shock system, induced by unstart, initially propagates upstream through the isolator at a velocity of about 35 m=s (in the laboratory frame of reference), then decelerates to about 20 m=s near the isolator entrance, and then accelerates to a velocity of about 74 m=s within the inlet. Throughout the isolator, unstart is seen to be strongly associated with boundary-layer separation. Once the inlet has unstarted, a highamplitudeoscillatory(periodic)unstarted flowensues,forwhichtheoscillationfrequencyisabout124Hz.However, under some conditions, an 84-Hz oscillatory unstarted flow mode, with lower pressure fluctuations, is observed. Under other conditions, a nonoscillatory unstarted flow, with much lower pressure fluctuations, is observed.

264 citations


Journal ArticleDOI
TL;DR: In this paper, the authors describe a new experiment executed in the ONERA S3Ch transonic wind tunnel on shock oscillations over the OAT15A supercritical profile, which has allowed the precise definition of the conditions for buffet onset and the characterization of the properties of the periodic motion from unsteady surface pressure measurements.
Abstract: Shock wave/turbulent boundary-layer interaction and flow separation may induce self-sustained large-scale oscillations on a profile at transonic Mach number. This phenomenon, known as transonic buffet, is at the origin of intense pressure fluctuations which can have detrimental effects, both in external and internal aerodynamics. The present paper describes a new experiment executed in the ONERA S3Ch transonic wind tunnel on shock oscillations over the OAT15A supercritical profile. These experiments have allowed the precise definition of the conditions for buffet onset and the characterization of the properties of the periodic motion from unsteady surface pressure measurements. The flowfield behavior has been described in great detail thanks to high-speed schlieren cinematography and surveys with a two-component laser Doppler velocimetry along with a conditional sampling technique. The first aim of this study was to provide the computational fluid dynamics community with well-documented test cases to validate advanced computing methods. Concerning the physics of the phenomenon, it is suggested that it is mediated by acoustic waves which are produced at the trailing edge and which travel on the two sides of the airfoil. Also, the experimental results strongly suggest that the phenomenon is essentially two-dimensional, even if three-dimensional effects are also detected.

235 citations


Journal ArticleDOI
TL;DR: In this article, a control volume analysis of the compressible viscous flow about an aircraft is performed, including integrated propulsors and flow-control systems, and the result is a clear identification and quantification of all the power sources, power sinks, and their interactions.
Abstract: A control volume analysis of the compressible viscous flow about an aircraft is performed, including integrated propulsors and flow-control systems. In contrast to most past analyses that have focused on forces and momentum flow, in particular thrust and drag, the present analysis focuses on mechanical power and kinetic energy flow. The result is a clear identification and quantification of all the power sources, power sinks, and their interactions, which are present in any aerodynamic flow. The formulation does not require any separate definitions of thrust and drag, and hence it is especially useful for analysis and optimization of aerodynamic configurations that have tightly integrated propulsion and boundary-layer control systems. Nomenclature b, c = wingspan and chord CD = dissipation coefficient Cf = skin friction coefficient Di = induced drag Dp = profile drag Dw = wave drag dS = surface element of control volume dV = volume element of control volume _ Ea = axial kinetic energy deposition rate _ Ep = pressure-work deposition rate _ Ev = transverse (vortex) kinetic energy deposition rate _ Ew = lateral wave-outflow energy deposition rate Fn = streamwise force from lateral outflow velocity Vn Fu = streamwise force from axial velocity u Fv = streamwise force from transverse velocities v, w Fx, Fz = total streamwise, normal aerodynamic forces

211 citations


Journal ArticleDOI
TL;DR: In this article, the simulation of the unsteady separated flows encountered by a plunging airfoil under low-Reynolds-number conditions (Rec 6 ◊ 10 4 ).
Abstract: This investigation addresses the simulation of the unsteady separated flows encountered by a plunging airfoil under low-Reynolds-number conditions (Rec 6 ◊ 10 4 ). The flow fields are computed employing a previously developed and extensively validated high-fidelity implicit large-eddy simulation (ILES) approach. In order to permit comparison with available experimental measurements, calculations are performed first for an SD7003 airfoil section at an angle of attack o = 4 plunging with reduced frequency k = 3.93 and nondimensional amplitude ho = 0.05. Under these conditions, it is demonstrated that for Rec = 10 4 , transitional effects are not significant and that the dynamic-stall vortices remain fairly coherent as they propagate along the airfoil. For Rec = 4 ◊ 10 4 , the dynamic-stall vortex system is laminar at is inception, however shortly afterwards, it experiences an abrupt breakdown associated with the onset of spanwise instability effects. A detailed description of this transition process near the leading edge is provided. The computed phased-averaged structures for both values of Reynolds number are found to be in good agreement with the experimental data. As a second example, the suppression of static stall at high angle of attack ( o = 14 ) is investigated using high-frequency small-amplitude vibrations (k = 10,ho = 0.005). At Rec = 6 ◊ 10 4 , separation is completely eliminated in a time-averaged sense, and the mean drag is reduced by approximately 40%. The instantaneous flow is characterized by the periodic generation of dynamic-stall vortices near the leading edge and by their subsequent transition as they convect close to the airfoil. For Rec = 10 4 , significant reduction of the timeaveraged separation region is still possible with transitional effects present in the aft-portion of the airfoil. For larger forcing amplitude (ho = 0.04,Rec = 10 4 ), a very intriguing regime emerges. The dynamic stall vortex moves around and in front of the leading edge and experiences a dramatic breakdown as it impinges against the airfoil. As a result, the phased-averaged flow displays no coherent vortices propagating along the airfoil upper surface. This new flow structure is also characterized in the mean by the existence of a strong jet in the near wake which manifests in a high value of net thrust. The present study demonstrates the importance of transitional effects for low-Reynolds-number maneuvering airfoils, as well as the suitability of the ILES approch for exploring such flow regime.

172 citations


Journal ArticleDOI
TL;DR: In this article, the authors studied the weights stability and accuracy of the implicit fifth-order weighted essentially nonoscillatory finite difference scheme and proposed an increased e value of 10 -2 for the weighted essentially nonsmoothness factors, which removed the weights oscillation and significantly improved the accuracy.
Abstract: This paper studies the weights stability and accuracy of the implicit fifth-order weighted essentially nonoscillatory finite difference scheme. It is observed that the weights of the Jiang-Shu weighted essentially nonoscillatory scheme oscillate even for smooth flows. An increased e value of 10 -2 is suggested for the weighted essentially nonoscillatory smoothness factors, which removes the weights oscillation and significantly improves the accuracy of the weights and solution convergence. With the improved e value, the weights achieve the optimum value with minimum numerical dissipation in smooth regions and maintain the sensitivity to capture nonoscillatory shock profiles for the transonic flows. The theoretical justification of this treatment is given in the paper. The wall surface boundary condition uses a half-point mesh so that the conservative differencing can be enforced. A third-order accurate finite difference scheme is given to treat wall boundary conditions. The implicit time-marching method with unfactored Gauss-Seidel line relaxation is used with the high-order schemes to achieve a high convergence rate. Several transonic cases are calculated to demonstrate the robustness, efficiency, and accuracy of the methodology.

141 citations


Journal ArticleDOI
TL;DR: In this paper, the authors investigated the mechanism by which large turbulence structures radiate noise and found that the mechanism is Mach wave radiation, and theoretical model results and physical reasoning are presented to support the Mach wave mechanism.
Abstract: Extensive experimental evidence is now available to support the observation that there are two components of jet mixing noise. They are the fine-scale turbulence noise and the noise from the large turbulence structures of the jet flow. The large turbulence structure’s noise radiates primarily in directions with a large inlet angle around the downstream axis of the jet. The fine-scale turbulence noise dominates in the sideline and upstream directions. This study investigates the mechanism by which large turbulence structures radiate noise. It is believed that the mechanism is Mach wave radiation. Theoretical model results and physical reasoning are presented to support the Mach wave mechanism. They are further supported by experimental measurements both in the far field and in the near acoustic field. These measurements include peak noise direction, noise-source distribution along the jet column, and near-field pressure-contour pattern. A signature pattern of the near-field pressure contours associated with Mach wave radiation is identified.

Journal ArticleDOI
TL;DR: In this article, an inverse method based on a system-identification technique for identifying impact events on a complex structure with built-in sensors is presented, which uses the transfer functions in the system identification technique to identify the location and force time history of an impact event on a structure without the need of constructing a full-scale accurate structural model or of acquiring excessive training data on the structure, such as neural-network techniques.
Abstract: An inverse method based on a system-identification technique for identifying impact events on a complex structure with built-in sensors is presented. The method using the transfer functions in the system-identification technique identifies the location and force time history of an impact event on a structure without the need of constructing a full- scale accurate structural model or of acquiring excessive training data on the structure, such as neural-network techniques. The system transfer functions for the entire structure are constructed by two sequential procedures: 1) limited impact tests at selected points to establish the system transfer functions from the selected points to a sensor on the structure and 2) an interpolation function approach based on a linear finite element to approximate the system transfer functions from a point inside four neighboring selected points to the sensor. Comprehensive tests with various impact situations verified the accuracy of load and position predictions by the proposed method.

Journal ArticleDOI
TL;DR: In this article, an aerodynamic shape optimization code coupled with a structural morphing model is used to obtain a set of optimal wing shapes for minimum drag at different flight speeds, based on changes in wing-planform shape and wing-section shape achieved by extending spars and telescopic ribs.
Abstract: This paper presents the work done in designing a morphing wing concept for a small experimental unmanned aerial vehicle to improve the vehicle's performance over its intended speed range. The wing is designed with a multidisciplinary design optimization tool, in which an aerodynamic shape optimization code coupled with a structural morphing model is used to obtain a set of optimal wing shapes for minimum drag at different flight speeds. The optimization procedure is described as well as the structural model. The aerodynamic shape optimization code, that uses a viscous two-dimensional panel method formulation coupled with a nonlinear lifting-line algorithm and a sequential quadratic programming optimization algorithm, is suitable for preliminary wing design optimization tasks. The morphing concept, based on changes in wing-planform shape and wing-section shape achieved by extending spars and telescopic ribs, is explained in detail. Comparisons between optimized fixed wing performance, optimal morphing wing performance, and the performance of the wing obtained from the coupled aerodynamic-structural solution are presented. Estimates for the performance enhancements achieved by the unmanned aerial vehicles when fitted with this new morphing wing are also presented. Some conclusions on this concept are addressed with comments on the benefits and drawbacks of the morphing mechanism design.

Journal ArticleDOI
TL;DR: AEROSPACE LETTERS as mentioned in this paper are brief communications (approximately 2000 words) that describe new and potentially important ideas or results, including critical analytical or experimental observations that justify rapid publication.
Abstract: AEROSPACE LETTERS are brief communications (approximately 2000 words) that describe new and potentially important ideas or results, including critical analytical or experimental observations that justify rapid publication. They are stringently prescreened, and only a few are selected for rapid review by an Editor. They are published as soon as possible electronically and then appear in the print version of the journal.

Journal ArticleDOI
TL;DR: In this article, a set of experiments conducted on a NACA0012 airfoil undergoing stall flutter oscillations in a low-speed wind tunnel is presented, with the objective of characterizing the local bifurcation behavior of the system.
Abstract: Stall flutter is a nonlinear aeroelastic phenomenon that can affect several types of aeroelastic systems such as helicopter rotor blades, wind turbine blades, and highly flexible wings. Although the related aerodynamic phenomenon of dynamic stall has been the subject of many experimental studies, stall flutter itself has rarely been investigated. This paper presents a set of experiments conducted on a NACA0012 airfoil undergoing stall flutter oscillations in a low-speed wind tunnel. The aeroelastic responses are analyzed with the objective of characterizing the local bifurcation behavior of the system. It is shown that symmetric stall flutter oscillations are encountered as a result of a subcritical Hopf bifurcation, followed by a fold bifurcation. The cause of these bifurcations is the occurrence of dynamic stall, which allows the transfer of energy from the freestream to the wing. A second bifurcation occurs at the system's static divergence airspeed. As a consequence, the wing starts to undergo asymmetric stall flutter bifurcations at only positive (or only negative) pitch angles. The dynamic stall mechanism itself does not change but the flow only separates on one side of the wing.

Journal ArticleDOI
TL;DR: In this paper, the authors describe an active flow control concept that uses counterflowing jets to significantly modify external flowfields and strongly disperse the shock waves of supersonic and hypersonic vehicles to reduce aerothermal loads and wave drag.
Abstract: This study describes an active flow control concept that uses counterflowing jets to significantly modify external flowfields and strongly disperse the shock waves of supersonic and hypersonic vehicles to reduce aerothermal loads and wave drag. The potential aerothermal and aerodynamic benefits of the concepts were investigated by conducting experiments on a 2.6%-scale Apollo capsule model in Mach 3.48 and 4.0 freestreams in a trisonic blowdown wind tunnel, as well as pretest computational fluid dynamics analyses of the flowfields, with and without counterflowing jets. The model employed three sonic and two supersonic (with design Mach numbers of 2.44 and 2.94) jet nozzles with exit diameters ranging from 0.25 to 0.5 in. The schlieren images were consistent with the pretest computational fluid dynamics predictions, showing a long penetration mode jet interaction at low jet flow rates of 0.05 and 0.1 Ib m /s, whereas a short penetration mode jet was revealed at higher flow rates. The long penetration mode jet appeared to be almost fully expanded and was unsteady, with the bow shock becoming so dispersed that it was no longer discernible. High-speed camera schlieren data revealed the bow shock to be dispersed into striations of compression waves, which suddenly coalesced to a weaker bow shock with a larger standoff distance as the flow rate reached a critical value. Heat transfer results showed a significant reduction in heat flux, even giving negative heat flux for some short penetration mode interactions, indicating that the flow wetting the model had a cooling effect, instead of heating, which could significantly impact thermal protection system requirements and design. The findings suggest that high-speed vehicle design and performance can benefit from the application ofcounterflowing jets as an active flow control.

Journal ArticleDOI
TL;DR: In this paper, a simple burner model is used to scale both the direct and indirect noise in aeroengines, and the analytical relations for the combustion and the nozzle provide simple scaling laws for direct combustion noise ratio as a function of the Mach number.
Abstract: Core noise in aeroengines is due to two main mechanisms: direct combustion noise, which is generated by the unsteady expansion of burning gases, and indirect combustion noise, which is due to the acceleration of entropy waves (temperature fluctuations generated by unsteady combustion) within the turbine stages. This paper shows how a simple burner model (a flame in a combustion chamber terminated by a nozzle) can be used to scale direct and indirect noise. An analytical formulation is used for waves generated by combustion. The transmission and generation of waves through the nozzle is calculated using both the analytical results of Marble and Candel (Marble, F. E., and Candel, S., "Acoustic Disturbances from Gas Nonuniformities Convected Through a Nozzle," Journal of Sound and Vibration, Vol. 55, 1977, pp. 225-243.) and a numerical tool. Numerical results for the nozzle verify and extend the analytical approach. The analytical relations for the combustion and the nozzle provide simple scaling laws for direct and indirect noise ratio as a function of the Mach number in the combustion chamber and at the nozzle outlet.

Journal ArticleDOI
TL;DR: A computational framework for simulating structural models of varied fidelity and a Navier-Stokes solver, aimed at simulating flapping and flexible wings, and implications of fluid density on aerodynamic loading are explored.
Abstract: Because of their small size and flight regime, coupling of aerodynamics, structural dynamics, and flight dynamics are critical for micro aerial vehicles This paper presents a computational framework for simulating structural models of varied fidelity and a Navier-Stokes solver, aimed at simulating flapping and flexible wings The structural model uses either 1) the in-house developed UM/NLABS, which decomposes the equations of 3-D elasticity into cross-sectional and spanwise analyses for slender wings, or 2) MSCMarc, which is a commercial finite-element solver capable of modeling geometrically nonlinear structures of arbitrary geometry The flow solver employs a well-tested pressure-based algorithm implemented in STREAM A NACA0012 cross-sectional rectangular wing of aspect ratio 3, chord Reynolds number of 3 x 10 4 , and reduced frequency varying from 04 to 182, with prescribed pure plunge motion is investigated Both rigid and flexible wing results are presented, and good agreement between experiment and computation are shown regarding tip displacement and thrust coefficient Issues related to coupling strategies, fluid physics associated with rigid and flexible wings, and implications of fluid density on aerodynamic loading are also explored in this paper

Journal ArticleDOI
TL;DR: This paper presents matrix-free methods for the stability analysis and control design of high-dimensional systems arising from the discretized linearized Navier-Stokes equations.
Abstract: This paper presents matrix-free methods for the stability analysis and control design of high-dimensional systems arising from the discretized linearized Navier-Stokes equations. The methods are ap ...

Journal ArticleDOI
TL;DR: In this paper, a survey of numerical experiments from 12 different flux functions in one-and two-dimensional contexts is presented, and it is found that there are at least two kinds of shock instabilities: one is one-dimensional and the other is multidimensional.
Abstract: Shock-capturing finitevolumeschemesoftengiverisetoanomalousresultsinhypersonic flow.Wepresentawideranging survey of numerical experiments from 12 different flux functions in one- and two-dimensional contexts. Included is a recently developed function that satisfies the second law of thermodynamics. It is found here that there are at least two kinds of shock instabilities: one is one-dimensional and the other is multidimensional. According to the results, the former does not appear if a flux function satisfies the second law of thermodynamics, and the latter is suppressed by an additional dissipation with a multidimensional character. However, such dissipation has no effect on the one-dimensional mode. Among the flux functions investigated, no universally stable schemes are found that arefreefrombothone-andmultidimensionalshockinstabilities.Theappearanceoftheseinstabilitiesdependsonthe relative positioning of the shock on the grid.

Journal ArticleDOI
TL;DR: In this paper, large-eddy simulations of imperfectly expanded jet flows from a convergent-divergent nozzle with a sharp contraction at the nozzle throat have been carried out, and the flowfield and near-field acoustics for various total pressure ratios from overexpanded to underexpanded jet flow conditions have been investigated.
Abstract: Large-eddy simulations of imperfectly expanded jet flows from a convergent-divergent nozzle with a sharp contraction at the nozzle throat have been carried out. The flowfield and near-field acoustics for various total pressure ratios from overexpanded to underexpanded jet flow conditions have been investigated. The location and spacing of the shock cells are in good agreement with experimental data and previous theoretical results. The velocity profiles are also in good agreement with data from experimental measurements. A Mach disk is observed immediately downstream of the nozzle exit for overexpanded jet conditions with nozzle pressure ratios much lower than the fully expanded value. It is found that this type of nozzle with a sharp turning throat does not have a shock-free condition for supersonic jet flows. The near-field intensities of pressure fluctuations show wavy structures for cases in which screech tones are observed. The large-eddy simulations predictions of the near-field noise intensities show good agreement with those obtained from experimental measurements. This good agreement shows that large-eddy simulations and measurements can play complementary roles in the investigation of the noise generation from supersonic jet flows.


Journal ArticleDOI
TL;DR: In this article, the authors report on the simulation of the near nozzle region of a moderate Reynolds number cold jet flow exhausting from an achevron nozzle using a high-order accurate, multiblock, large-eddy simulation code with overset grid capability.
Abstract: This paper reports on the simulation of the near-nozzle region of a moderate Reynolds number cold jet flow exhaustingfrom achevron nozzle.Thechevron nozzleconsideredinthis studyisthe SMC001nozzleexperimentally studied by researchers at the NASA John H. Glenn Research Center. This nozzle design contains six symmetric chevrons that have a 5-deg penetration angle. The flow inside the chevron nozzle and the free jet flow outside are computed simultaneously by a high-order accurate, multiblock, large-eddy simulation code with overset grid capability.Theresolutionofthesimulationisabout100milliongridpoints.Themainemphasisofthesimulationisto capture the enhanced shear-layer mixing due to the chevrons and the consequent noise generation that occurs in the mixing layers of the jet within the first few diameters downstream of the nozzle exit. Details of the computational methodology are presented together with an analysis of the simulation results. The simulation data are compared with available experimental measurements of the flowfield and the noise spectrum in the sideline direction. Overall, thesimulationresultsareveryencouraginganddemonstratethefeasibilityofchevronnozzlejetcomputationsusing our simulation methodology. Nomenclature c

Journal ArticleDOI
TL;DR: In this article, the advantages of using multiple surrogates for approximation and reduction of helicopter vibration are studied, and the optimized designs are compared with a baseline design resembling a Messerschmitt-Bolkow-Blohm BO-105 blade.
Abstract: The advantages of using multiple surrogates for approximation andreduction of helicopter vibration are studied. Multiple approximation methods, including a weighted-average approach, are considered so that pitfalls associated with only using a single best surrogate for the rotor blade vibration-reduction problem are avoided. A vibration objective function corresponding to a flight condition in which blade–vortex interaction causes high levels of vibration is considered. The design variables consist of cross-sectional dimensions of the structural member of the blade and nonstructural masses. The optimized designs are compared with a baseline design resembling a Messerschmitt–Bolkow–Blohm BO-105 blade. The results indicate that at relatively little additional cost compared with optimizing a single surrogate, multiple surrogates can be used to locate various reduced-vibration designs that wouldbeoverlookedifonly asingleapproximationmethodwasemployed,andthemostaccurate surrogatemaynot lead to the best design.

Journal ArticleDOI
TL;DR: In this article, a two-dimensional Shack-Hartmann wavefront sensor is used to study aero-optic distortion in turbulent boundary layers at transonic and hypersonic speeds, with and without gas injection.
Abstract: A two-dimensional Shack-Hartmann wave-front sensor is used to study aero-optic distortion in turbulent boundary layers at transonic and hypersonic speeds, with and without gas injection. The large-scale motions in the outer layer, of the order of the boundary-layer thickness in size, are shown to dominate the aero-optic distortion. Gas injection always reduced the Strehl ratio, with helium injection generally giving lower Strehl ratios than nitrogen injection. The large aperture approximation is shown to be accurate for a wide variety of aberrations regardless of Mach number and gas injection. A new scaling argument for the root-mean-square phase distortion is proposed that appears to collapse the data better than previous models.

Journal ArticleDOI
TL;DR: In this paper, a local-pistons theory was proposed for the prediction of inviscid unsteady pressure loads at supersonic and hypersonic speeds, and the results of two-and three-dimensional air loads and flutter predictions were compared with those obtained by the classical piston theory and an unstrainedy Euler method to assess the accuracy and validity.
Abstract: DOI: 10.2514/1.37750 A highly efficient local-piston theory is presented for the prediction of inviscid unsteady pressure loads at supersonic and hypersonic speeds. A steady mean flow solution is first obtained by an Euler method. The classical pistontheoryismodifiedtoapplylocallyateachpointontheairfoilsurfaceontopofthelocalmean flowtoobtainthe unsteadypressureperturbationscausedbythedeviationoftheairfoilsurfacefromitsmeanlocationwithouttheneed of performing unsteady Euler computations. Results of two- and three-dimensional unsteady air loads and flutter predictions are compared with those obtained by the classical piston theory and an unsteady Euler method to assess theaccuracyandvalidityrangeinairfoilthickness, flightMachnumber,andangleofattackandwiththepresenceof blunt leading edges. The local-piston theory is found to offer superior accuracy and much wider validity range compared with the classical piston theory, with the cost of only a fraction of the computational time needed by an unsteady Euler method.

Journal ArticleDOI
TL;DR: A novel method of guided wave-based structural health monitoring is developed in which no direct baseline data are required to identify structural damage and this new method accomplishes reference-free damage detection by acquiring what is referred to as an instantaneous baseline measurement for analysis.
Abstract: A novel method of guided wave-based structural health monitoring is developed in which no direct baseline data are required to identify structural damage. Conventional wave propagation structural health monitoring techniques involve the comparison of structural response data to a prerecorded baseline or reference measurement taken while the structure is in pristine condition. The need to compare new data to a prerecorded baseline can present several complications, including data management issues and difficulty in accommodating the effects of varying environmental and operational conditions on the data. To address the complications associated with baseline comparison, this new method accomplishes reference-free damage detection by acquiring what is referred to as an instantaneous baseline measurement for analysis. The instantaneous baseline technique is validated through both analytical and experimental testing. Analytical tests show that the instantaneous baseline method is able to correctly identify simulated damage. It is found experimentally that nonpermanent damage in the form of removable putty as well as permanent damage in the form of corrosion and cuts are all identifiable in thin aluminum plate test structures without direct comparison to baseline data when implementing the instantaneous baseline method.

Journal ArticleDOI
TL;DR: In this article, the concept of electrospun polymer nanofiber fabric interleaving to enhance dynamic properties, impact damage resistance, fracture toughness and resistance, and delamination onset life was evaluated.
Abstract: The concept of electrospun polymer nanofiber fabric interleaving to enhance dynamic properties, impact damage resistance, fracture toughness and resistance, and delamination onset life was evaluated. Polymer nanofabric interleaving increased the laminate thickness and weight by an order of 1%, and its impact on in-plane mechanical properties of the composite laminate would be statistically zero. On the other hand, its influence on interlaminar fracture toughness and resistance, impact damage resistance, and damping is substantial. Results of this study showed that interleaving AS4/3501-6 composite laminate increased the damping by 13%, reduced the impact damage size to one-third, increased fracture toughness and resistance by 1.5 times and one-third, respectively, significantly increased delamination onset life, and increased the fatigue threshold energy release rate by two-thirds. These improvements are comparable to that of the commercial T800H/3900-2 composite but with no thickness increase penalty, loss of in-plane properties, or multiple glass transition temperatures.

Journal ArticleDOI
TL;DR: The ghost fluid method is found to be stable, accurate, and robust for wide range problems involving strong shocks interacting with embedded solid objects, particularly when the embedded interface is retained as a sharp entity.
Abstract: Numerical simulation of shock waves interacting with multimaterial interface is immensely challenging, particularly when the embedded interface is retained as a sharp entity. The challenge lies in accurately capturing and representing the interface dynamics and the wave patterns at the interface. In this regard, the ghost fluid method has been successfully used to capture the interface conditions for both fluid-fluid and solid-fluid interfaces. However, the ghost fluid method results in over/underheating errors when shocks impact interfaces, and hence must be supplemented with numerical corrective measures to mitigate these errors. Such corrections typically fail for strong shock applications. Variants and extensions ofthe ghost fluid method have been proposed to remedy its shortcomings with mixed success. In this paper, the performance of approaches based on the ghost fluid method, in the case of strong shocks impinging on immersed solid boundaries in compressible flows, is evaluated. It is found that (from the viewpoint of simplicity, robustness, and accuracy) a reflective boundary condition used in conjunction with a local Riemann solver at the interface proves to be a good choice. The method is found to be stable, accurate, and robust for wide range problems involving strong shocks interacting with embedded solid objects.

Journal ArticleDOI
TL;DR: In this paper, a numerical approach to study the turbulent flow of supercritical fluids is presented and validated by comparison with experimental data in an asymmetrically heated three-dimensional channel with a high-aspect ratio (channel height-to-width ratio).
Abstract: The knowledge of the flow behavior inside asymmetrically heated channels is of great importance to improve design and performance of regeneratively cooled rocket engines. The modeling of the coolant flow is a challenging task because of its particular features, such as the high wall temperature gradient, the high Reynolds number, the three-dimensional geometry of the passages, and the possible supercritical conditions of the fluid. In the present work, a numerical approach to study the turbulent flow of supercritical fluids is presented and validated by comparison with experimental data. Solutions of the supercritical nitrogen flowfield in an asymmetrically heated three-dimensional channel with a high-aspect ratio (channel height-to-width ratio) are presented and discussed. Emphasis is given to the analysis of the peculiar behavior and cooling performance of the supercritical fluid as compared with perfect gas. In particular, a long channel is considered, such that entrance effects are negligible, to analyze in detail wall heat-flux evolution throughout the channel.