Influence of wing tip morphology on vortex dynamics of flapping flight

NASA Astrophysics Data System (ADS)

Krishna, Swathi; Mulleners, Karen

2013-11-01

The mechanism of flapping wing flight provides insects with extraordinary flight capabilities. The uniquely shaped wing tips give insects an edge in flight performance and the interaction between the leading edge vortices and wing tip vortices enhance their propelling efficiencies and manoeuvrability. These are qualities that are sought after in current-day Micro Air Vehicles. A detailed understanding of the vortex dynamics of flapping flight and the influence of the wing tip planform is imperative for technical application. An experimental study is conducted to investigate the effects of different wing tip planforms on the formation, evolution and interaction of vortical structures. We thereby focus on the interaction between the coherent structures evolving from the leading edge and the wing tip during pitching and flapping motions.The spatial and temporal evolution of the three-dimensional flow structures are determined using Scanning (Stereo) Particle Image Velocimetry and an in-depth coherent structure analysis. By comparing the vortex dynamics, the aerodynamic performance of various wing tip planforms are evaluated.

Numerical simulation of tip vortices of wings in subsonic and transonic flows

NASA Technical Reports Server (NTRS)

Srinivasan, G. R.; Mccroskey, W. J.; Baeder, J. D.; Edwards, T. A.

1986-01-01

A multi block zonal algorithm which solves the thin-layer Navier-Stokes and the Euler equations is used to numerically simulate the formation and roll-up of the tip vortex in both subsonic and transonic flows. Four test cases which used small and large aspect ratio wings have been considered to examine the influence of the tip-cap shape, the tip planform and the free-stream Mach number. It appears that both the tip-planform and the tip-cap shape have some influence on the formation of the tip vortex, but its subsequent roll-up seems to be more influenced by the tip-planform shape. In general, a good definition of the formation and the roll-up of the tip vortex has been observed for all the cases considered here. Comparions of the numerical results with the limited, available experimental data show good agreement with both the surface pressures and the tip-vortex strength.

Numerical Simulation of Tip Vortices of Wings in Subsonic and Transonic Flows,

DTIC Science & Technology

1986-01-01

roll-up of the tip vor- rv : dimensionless strength of tip vortex " tex in both subsonic and transonic flows. Four test cases which used small and large...of their po- tion and the roll-up of the tip vortex has been observed for tential hazard to aircraft that encounter them in flight. To all the cases...such flows encompassing large air- tip- vortex strength. craft wakes (see for example Refs. 1-2). In spite of this, the present understanding of such

Experimental and Theoretical Study of a Rectangular Wing in a Vortical Wake at Low Speed

NASA Technical Reports Server (NTRS)

Smith, Willard G.; Lazzeroni, Frank A.

1960-01-01

A systematic study has been made, experimentally and theoretically, of the effects of a vortical wake on the aerodynamic characteristics of a rectangular wing at subsonic speed. The vortex generator and wing were mounted on a reflection plane to avoid body-wing interference. Vortex position, relative to the wing, was varied both in the spanwise direction and normal to the wing. Angle of attack of the wing was varied from -40 to +60. Both chordwise and spanwise pressure distributions were obtained with the wing in uniform and vortical flow fields. Stream surveys were made to determine the flow characteristics in the vortical wake. The vortex-induced lift was calculated by several theoretical methods including strip theory, reverse-flow theory, and reverse-flow theory including a finite vortex core. In addition, the Prandtl lifting-line theory and the Weissinger theory were used to calculate the spanwise distribution of vortex-induced loads. With reverse-flow theory, predictions of the interference lift were generally good, and with Weissinger's theory the agreement between the theoretical spanwise variation of induced load and the experimental variation was good. Results of the stream survey show that the vortex generated by a lifting surface of rectangular plan form tends to trail back streamwise from the tip and does not approach the theoretical location, or centroid of circulation, given by theory. This discrepancy introduced errors in the prediction of vortex interference, especially when the vortex core passed immediately outboard of the wing tip. The wake produced by the vortex generator in these tests was not fully rolled up into a circular vortex, and so lacked symmetry in the vertical direction of the transverse plane. It was found that the direction of circulation affected the induced loads on the wing either when the wing was at angle of attack or when the vortex was some distance away from the plane of the wing.

Numerical Capture of Wing-tip Vortex Using Vorticity Confinement

NASA Astrophysics Data System (ADS)

Zhang, Baili; Lou, Jing; Kang, Chang Wei; Wilson, Alexander; Lundberg, Johan; Bensow, Rickard

2012-11-01

Tracking vortices accurately over large distances is very important in many areas of engineering, for instance flow over rotating helicopter blades, ship propeller blades and aircraft wings. However, due to the inherent numerical dissipation in the advection step of flow simulation, current Euler and RANS field solvers tend to damp these vortices too fast. One possible solution to reduce the unphysical decay of these vortices is the application of vorticity confinement methods. In this study, a vorticity confinement term is added to the momentum conservation equations which is a function of the local element size, the vorticity and the gradient of the absolute value of vorticity. The approach has been evaluated by a systematic numerical study on the tip vortex trailing from a rectangular NACA0012 half-wing. The simulated structure and development of the wing-tip vortex agree well with experiments both qualitatively and quantitatively without any adverse effects on the global flow field. It is shown that vorticity confinement can negate the effect of numerical dissipation, leading to a more or less constant vortex strength. This is an approximate method in that genuine viscous diffusion of the vortex is not modeled, but it can be appropriate for vortex dominant flows over short to medium length scales where viscous diffusion can be neglected.

3D visualization of unsteady 2D airplane wake vortices

NASA Technical Reports Server (NTRS)

Ma, Kwan-Liu; Zheng, Z. C.

1994-01-01

Air flowing around the wing tips of an airplane forms horizontal tornado-like vortices that can be dangerous to following aircraft. The dynamics of such vortices, including ground and atmospheric effects, can be predicted by numerical simulation, allowing the safety and capacity of airports to be improved. In this paper, we introduce three-dimensional techniques for visualizing time-dependent, two-dimensional wake vortex computations, and the hazard strength of such vortices near the ground. We describe a vortex core tracing algorithm and a local tiling method to visualize the vortex evolution. The tiling method converts time-dependent, two-dimensional vortex cores into three-dimensional vortex tubes. Finally, a novel approach calculates the induced rolling moment on the following airplane at each grid point within a region near the vortex tubes and thus allows three-dimensional visualization of the hazard strength of the vortices. We also suggest ways of combining multiple visualization methods to present more information simultaneously.

Study on flow over finite wing with respect to F-22 raptor, Supermarine Spitfire, F-7 BG aircraft wing and analyze its stability performance and experimental values

NASA Astrophysics Data System (ADS)

Ali, Md. Nesar; Alam, Mahbubul

2017-06-01

A finite wing is a three-dimensional body, and consequently the flow over the finite wing is three-dimensional; that is, there is a component of flow in the span wise direction. The physical mechanism for generating lift on the wing is the existence of a high pressure on the bottom surface and a low pressure on the top surface. The net imbalance of the pressure distribution creates the lift. As a by-product of this pressure imbalance, the flow near the wing tips tends to curl around the tips, being forced from the high-pressure region just underneath the tips to the low-pressure region on top. This flow around the wing tips is shown in the front view of the wing. As a result, on the top surface of the wing, there is generally a span wise component of flow from the tip toward the wing root, causing the streamlines over the top surface to bend toward the root. On the bottom surface of the wing, there is generally a span wise component of flow from the root toward the tip, causing the streamlines over the bottom surface to bend toward the tip. Clearly, the flow over the finite wing is three-dimensional, and therefore we would expect the overall aerodynamic properties of such a wing to differ from those of its airfoil sections. The tendency for the flow to "leak" around the wing tips has another important effect on the aerodynamics of the wing. This flow establishes a circulatory motion that trails downstream of the wing; that is, a trailing vortex is created at each wing tip. The aerodynamics of finite wings is analyzed using the classical lifting line model. This simple model allows a closed-form solution that captures most of the physical effects applicable to finite wings. The model is based on the horseshoe-shaped vortex that introduces the concept of a vortex wake and wing tip vortices. The downwash induced by the wake creates an induced drag that did not exist in the two-dimensional analysis. Furthermore, as wingspan is reduced, the wing lift slope decreases, and the induced drag increases, reducing overall efficiency. To complement the high aspect ratio wing case, a slender wing model is formulated so that the lift and drag can be estimated for this limiting case as well. We analyze the stability performance of F-22 raptor, Supermarine Spitfire, F-7 BG Aircraft wing by using experimental method and simulation software. The experimental method includes fabrication of F-22 raptor, Supermarine Spitfire, F-7 BG Aircraft wing which making material is Gamahr wood. Testing this model wing in wind tunnel test and after getting expected data we also compared this value with analyzing software data for furthermore experiment.

An experimental study of the nonlinear dynamic phenomenon known as wing rock

NASA Technical Reports Server (NTRS)

Arena, A. S., Jr.; Nelson, R. C.; Schiff, L. B.

1990-01-01

An experimental investigation into the physical phenomena associated with limit cycle wing rock on slender delta wings has been conducted. The model used was a slender flat plate delta wing with 80-deg leading edge sweep. The investigation concentrated on three main areas: motion characteristics obtained from time history plots, static and dynamic flow visualization of vortex position, and static and dynamic flow visualization of vortex breakdown. The flow visualization studies are correlated with model motion to determine the relationship between vortex position and vortex breakdown with the dynamic rolling moments. Dynamic roll moment coefficient curves reveal rate-dependent hysteresis, which drives the motion. Vortex position correlated with time and model motion show a time lag in the normal position of the upward moving wing vortex. This time lag may be the mechanism responsible for the hysteresis. Vortex breakdown is shown to have a damping effect on the motion.

Flow visualization study of the HiMAT RPRV

NASA Technical Reports Server (NTRS)

Lorincz, D. J.

1980-01-01

Water tunnel studies were performed to qualitatively define the flow field of the highly maneuverable aircraft technology remotely piloted research vehicle (HiMAT RPRV). Particular emphasis was placed on defining the vortex flows generated at high angles of attack. The flow visualization tests were conducted in the Northrop water tunnel using a 1/15 scale model of the HiMAT RPRV. Flow visualization photographs were obtained for angles of attack up to 40 deg and sideslip angles up to 5 deg. The HiMAT model was investigated in detail to determine the canard and wing vortex flow field development, vortex paths, and vortex breakdown characteristics as a function of angle of attack and sideslip. The presence of the canard caused the wing vortex to form further outboard and delayed the breakdown of the wing vortex to higher angles of attack. An increase in leading edge camber of the maneuver configuration delayed both the formation and the breakdown of the wing and canard vortices. Additional tests showed that the canard vortex was sensitive to variations in inlet mass flow ratio and canard flap deflection angle.

Flow visualizations of perpendicular blade vortex interactions

NASA Technical Reports Server (NTRS)

Rife, Michael C.; Davenport, William J.

1992-01-01

Helium bubble flow visualizations have been performed to study perpendicular interaction of a turbulent trailing vortex and a rectangular wing in the Virginia Tech Stability Tunnel. Many combinations of vortex strength, vortex-blade separation (Z(sub s)) and blade angle of attack were studied. Photographs of representative cases are presented. A range of phenomena were observed. For Z(sub s) greater than a few percent chord the vortex is deflected as it passes the blade under the influence of the local streamline curvature and its image in the blade. Initially the interaction appears to have no influence on the core. Downstream, however, the vortex core begins to diffuse and grow, presumably as a consequence of its interaction with the blade wake. The magnitude of these effects increases with reduction in Z(sub s). For Z(sub s) near zero the form of the interaction changes and becomes dependent on the vortex strength. For lower strengths the vortex appears to split into two filaments on the leading edge of the blade, one passing on the pressure and one passing on the suction side. At higher strengths the vortex bursts in the vicinity of the leading edge. In either case the core of its remnants then rapidly diffuse with distance downstream. Increase in Reynolds number did not qualitatively affect the flow apart from decreasing the amplitude of the small low-frequency wandering motions of the vortex. Changes in wing tip geometry and boundary layer trip had very little effect.

Vortices and turbulence (The 23rd Lanchester Memorial Lecture)

NASA Astrophysics Data System (ADS)

Lilley, G. M.

1983-12-01

A comprehensive discussion is presented concerning the phenomena characteristically treated in vortex and turbulence theory, as well as the degree of success achieved by various computation and visualization methods and theoretical models developed for vortex flow behavior prediction. Note is taken of the pioneering research conducted by F. W. Lanchester in 1893-1907, and attention is given to vortex tip and edge generation by rectangular and delta wings, the cool core effect of the Ranque-Hilsch vortex tube, the modeling of shear flows by means of vortex array methods, the classification and modelling of turbulent flows (together with a taxonomy of their calculation methods), and NASA ILLIAC IV computations of two-dimensional channel flow. Also noted are recent results concerning the boundary layer coherent structure of a flat plate at zero pressure gradient, including the regeneration structure and flow distortion and breakdown of a turbulent boundary layer.

An experimental investigation of delta wing vortex flow with and without external jet blowing

NASA Technical Reports Server (NTRS)

Iwanski, Kenneth P.; Ng, T. Terry; Nelson, Robert C.

1989-01-01

A visual and quantitative study of the vortex flow field over a 70-deg delta wing with an external jet blowing parallel to and at the leading edge was conducted. In the experiment, the vortex core was visually marked with TiCl4, and LDA was used to measure the velocity parallel and normal to the wing surface. It is found that jet blowing moved vortex breakdown farther downstream from its natural position and influenced the breakdown characteristics.

A viscous flow analysis for the tip vortex generation process

NASA Technical Reports Server (NTRS)

Shamroth, S. J.; Briley, W. R.

1979-01-01

A three dimensional, forward-marching, viscous flow analysis is applied to the tip vortex generation problem. The equations include a streamwise momentum equation, a streamwise vorticity equation, a continuity equation, and a secondary flow stream function equation. The numerical method used combines a consistently split linearized scheme for parabolic equations with a scalar iterative ADI scheme for elliptic equations. The analysis is used to identify the source of the tip vortex generation process, as well as to obtain detailed flow results for a rectangular planform wing immersed in a high Reynolds number free stream at 6 degree incidence.

The inviscid pressure field on the tip of a semi-infinite wing and its application to the formation of a tip vortex

NASA Technical Reports Server (NTRS)

Hall, G. F.; Shamroth, S. J.; Mcdonald, H.; Briley, W. R.

1976-01-01

A method was developed for determining the aerodynamic loads on the tip of an infinitely thin, swept, cambered semi-infinite wing at an angle of attack which is operating subsonically in an inviscid medium and is subjected to a sinusoidal gust. Under the assumption of linearized aerodynamics, the loads on the tip are obtained by superposition of the steady aerodynamic results for angle of attack and camber, and the unsteady results for the response to the sinusoidal gust. The near field disturbance pressures in the fluid surrounding the tip are obtained by assuming a dipole representation for the loading on the tip and calculating the pressures accordingly. The near field pressures are used to drive a reduced form of the Navier-Stokes equations which yield the tip vortex formation. The combined viscid-inviscid analysis is applied to determining the pressures and examining the vortex rollup in the vicinity of an unswept, uncambered wing moving steadily at a Mach number of 0.2 at an angle of attack of 0.1 rad. The viscous tip flow calculation shows features expected in the tip flow such as the qualitatively proper development of boundary layers on both the upper and lower airfoil surfaces. In addition, application of the viscous solution leads to the generation of a circular type flow pattern above the airfoil suction surface.

Leading-edge flow reattachment and the lateral static stability of low-aspect-ratio rectangular wings

NASA Astrophysics Data System (ADS)

Linehan, Thomas; Mohseni, Kamran

2017-11-01

The relationship between lateral static stability derivative, Clβ,lift coefficient, CL, and angle of attack was investigated for rectangular wings of aspect ratio A R =0.75 ,1 ,1.5 , and 3 using Stereo-Digital Particle Image Velocimetry (S-DPIV) and direct force and moment measurements. When the product Cl βA R is plotted with respect to CL, the lateral stability curves of each wing collapse to a single line for CL<0.7 . For CL>0.7 , the linearity and scaling of Clβwith respect to CL is lost. S-DPIV is used to elucidate the flow physics in this nonlinear regime. At α =10∘ , the leading-edge separation region emerges on the leeward portion of the sideslipped wing by means of vortex shedding. For the A R ≤1.5 wings at α >15∘ , the tip vortex downwash is sufficient to restrict the shedding of leading-edge vorticity thereby sustaining the lift of the leading-edge separation region at high angles of attack. Concurrently, the windward tip vortex grows in size and strength with increasing angle of attack, displacing the leading-edge separation region further toward the leeward wing. This reorganization of lift-generating vorticity results in the initial nonlinearities between Cl β and CL at angles of attack for which CL is still increasing. At angles of attack near that of maximum lift for the A R ≤1 wings, the windward tip vortex lifts off the wing, decreasing the lateral static stability of the wing prior to lift stall. For the A R =3 wing at α >10∘ , nonlinear trends in Cl β versus CL occur due to the spanwise evolution of stalled flow.

Experimental study of the velocity field on a delta wing

NASA Technical Reports Server (NTRS)

Payne, F. M.; Ng, T. T.; Nelson, R. C.

1987-01-01

An experimental study of the leading edge vortices on delta wings at large angles of incidence is presented. A combination of flow visualization, seven-hole pressure probe surveys and laser velocimeter measurements were used to study the leading edge vortex formation and breakdown for a set of delta wings. The delta wing models were thin flat plates with sharp leading edges having sweep angles of 70, 75, 80, and 85 degrees. The flow structure was examined for angles of incidence from 10 to 40 degrees and chord Reynolds numbers from 85,000 to 640,000. Vortex breakdown was observed on all the wings tested. Both bubble and spiral modes of breakdown were observed. The visualization and wake survey data shows that when vortex breakdown occurs the core flow transforms abruptly from a jet-like flow to a wake-like flow. The result also revealed that probe induced vortex breakdown was more steady than the natural breakdown.

Measurement of circulation around wing-tip vortices and estimation of lift forces using stereo PIV

NASA Astrophysics Data System (ADS)

Asano, Shinichiro; Sato, Haru; Sakakibara, Jun

2017-11-01

Applying the flapping flight to the development of an aircraft as Mars space probe and a small aircraft called MAV (Micro Air Vehicle) is considered. This is because Reynolds number assumed as the condition of these aircrafts is low and similar to of insects and small birds flapping on the earth. However, it is difficult to measure the flow around the airfoil in flapping flight directly because of its three-dimensional and unsteady characteristics. Hence, there is an attempt to estimate the flow field and aerodynamics by measuring the wake of the airfoil using PIV, for example the lift estimation method based on a wing-tip vortex. In this study, at the angle of attack including the angle after stall, we measured the wing-tip vortex of a NACA 0015 cross-sectional and rectangular planform airfoil using stereo PIV. The circulation of the wing-tip vortex was calculated from the obtained velocity field, and the lift force was estimated based on Kutta-Joukowski theorem. Then, the validity of this estimation method was examined by comparing the estimated lift force and the force balance data at various angles of attack. The experiment results are going to be presented in the conference.

Wake structure and wing motion in bat flight

NASA Astrophysics Data System (ADS)

Hubel, Tatjana; Breuer, Kenneth; Swartz, Sharon

2008-11-01

We report on experiments concerning the wake structure and kinematics of bat flight, conducted in a low-speed wind tunnel using time-resolved PIV (200Hz) and 4 high-speed cameras to capture wake and wing motion simultaneously. 16 Lesser dog-faced fruit bats (C. brachyotis) were trained to fly in the wind tunnel at 3-6.5m/s. The PIV recordings perpendicular to the flow stream allowed observing the development of the tip vortex and circulation over the wing beat cycle. Each PIV acquisition sequence is correlated with the respective kinematic history. Circulation within wing beat cycles were often quite repeatable, however variations due to maneuvering of the bat are clearly visible. While no distinct vortex structure was observed at the upper reversal point (defined according the vertical motion of the wrist) a tip vortex was observed to develop in the first third of the downstroke, growing in strength, and persisting during much of the upstroke. Correlated to the presence of a strong tip vortex the circulation has almost constant strength over the middle half of the wing beat. At relatively low flight speeds (3.4 m/s), a closed vortex structure behind the bat is postulated.

Mechanisms of Wing Beat Sound in Flapping Wings of Beetles

NASA Astrophysics Data System (ADS)

Allen, John

2017-11-01

While the aerodynamic aspects of insect flight have received recent attention, the mechanisms of sound production by flapping wings is not well understood. Though the harmonic structure of wing beat frequency modulation has been reported with respect to biological implications, few studies have rigorously quantified it with respect directionality, phase coupling and vortex tip scattering. Moreover, the acoustic detection and classification of invasive species is both of practical as well scientific interest. In this study, the acoustics of the tethered flight of the Coconut Rhinoceros Beetle (Oryctes rhinoceros) is investigated with four element microphone array in conjunction with complementary optical sensors and high speed video. The different experimental methods for wing beat determination are compared in both the time and frequency domain. Flow visualization is used to examine the vortex and sound generation due to the torsional mode of the wing rotation. Results are compared with related experimental studies of the Oriental Flower Beetle. USDA, State of Hawaii.

Aerodynamic Comparison of Hyper-Elliptic Cambered Span (HECS) Wings with Conventional Configurations

NASA Technical Reports Server (NTRS)

Lazos, Barry S.; Visser, Kenneth D.

2006-01-01

An experimental study was conducted to examine the aerodynamic and flow field characteristics of hyper-elliptic cambered span (HECS) wings and compare results with more conventional configurations used for induced drag reduction. Previous preliminary studies, indicating improved L/D characteristics when compared to an elliptical planform prompted this more detailed experimental investigation. Balance data were acquired on a series of swept and un-swept HECS wings, a baseline elliptic planform, two winglet designs and a raked tip configuration. Seven-hole probe wake surveys were also conducted downstream of a number of the configurations. Wind tunnel results indicated aerodynamic performance levels of all but one of the HECS wings exceeded that of the other configurations. The flow field data surveys indicate the HECS configurations displaced the tip vortex farther outboard of the wing than the Baseline configuration. Minimum drag was observed on the raked tip configuration and it was noted that the winglet wake lacked the cohesive vortex structure present in the wakes of the other configurations.

Topological structures of vortex flow on a flying wing aircraft, controlled by a nanosecond pulse discharge plasma actuator

NASA Astrophysics Data System (ADS)

Du, Hai; Shi, Zhiwei; Cheng, Keming; Wei, Dechen; Li, Zheng; Zhou, Danjie; He, Haibo; Yao, Junkai; He, Chengjun

2016-06-01

Vortex control is a thriving research area, particularly in relation to flying wing or delta wing aircraft. This paper presents the topological structures of vortex flow on a flying wing aircraft controlled by a nanosecond plasma dielectric barrier discharge actuator. Experiments, including oil flow visualization and two-dimensional particle image velocimetry (PIV), were conducted in a wind tunnel with a Reynolds number of 0.5 × 106. Both oil and PIV results show that the vortex can be controlled. Oil topological structures on the aircraft surface coincide with spatial PIV flow structures. Both indicate vortex convergence and enhancement when the plasma discharge is switched on, leading to a reduced region of separated flow.

Rolling moments in a trailing vortex flow field

NASA Technical Reports Server (NTRS)

Mcmillan, O. J.; Schwind, R. G.; Nielsen, J. N.; Dillenius, M. F. E.

1977-01-01

Pressure distributions are presented which were measured on a wing in close proximity to a tip vortex of known structure generated by a larger, upstream semispan wing. Overall loads calculated by integration of these pressures are checked by independent measurements made with an identical model mounted on a force balance. Several conventional methods of wing analysis are used to predict the loads on the following wing. Strip theory is shown to give uniformly poor results for loading distribution, although predictions of overall lift and rolling moment are sometimes acceptable. Good results are obtained for overall coefficients and loading distribution by using linearized pressures in vortex-lattice theory in conjunction with a rectilinear vortex. The equivalent relation from reverse-flow theory that can be used to give economic predictions for overall loads is presented.

Lee-side flow over delta wings at supersonic speeds

NASA Technical Reports Server (NTRS)

Miller, D. S.; Wood, R. M.

1985-01-01

An experimental investigation of the lee-side flow on sharp leading-edge delta wings at supersonic speeds has been conducted. Pressure data were obtained at Mach numbers from 1.5 to 2.8, and three types of flow-visualization data (oil-flow, tuft, and vapor-screen) were obtained at Mach numbers from 1.7 to 2.8 for wing leading-edge sweep angles from 52.5 deg to 75 deg. From the flow-visualization data, the lee-side flows were classified into seven distinct types and a chart was developed that defines the flow mechanism as a function of the conditions normal to the wing leading edge, specifically, angle of attack and Mach number. Pressure data obtained experimentally and by a semiempirical prediction method were employed to investigate the effects of angle of attack, leading-edge sweep, and Mach number on vortex strength and vortex position. In general, the predicted and measured values of vortex-induced normal force and vortex position obtained from experimental data have the same trends with angle of attack, Mach number, and leading-edge sweep; however, the vortex-induced normal force is underpredicted by 15 to 30 percent, and the vortex spanwise location is overpredicted by approximately 15 percent.

Downwash in Vortex Region Behind Rectangular Half-wing at Mach Number 1.91

NASA Technical Reports Server (NTRS)

Cummings, John L; Haefeli, Rudolph C

1950-01-01

Results of an experimental investigation to determine downwash and wake characteristics in region of trailing vortex system behind a rectangular half-wing at Mach number 1.91 are presented. The wing had a 5-percent thick symmetric diamond cross section beveled to a knife edge at the tip. At small angles of attack, downwash angles were in close agreement with predictions of linearized theory based on the assumption of an undistorted vortex sheet. At higher angles of attack, the flow was greatly influenced by the rolling up of the vortex sheet.

Predicting aerodynamic characteristics of vortical flows on three-dimensional configurations using a surface-singularity panel method

NASA Technical Reports Server (NTRS)

Maskew, B.

1983-01-01

A general low-order surface-singularity panel method is used to predict the aerodynamic characteristics of a problem where a wing-tip vortex from one wing closely interacts with an aft mounted wing in a low Reynolds Number flow; i.e., 125,000. Nonlinear effects due to wake roll-up and the influence of the wings on the vortex path are included in the calculation by using a coupled iterative wake relaxation scheme. The interaction also affects the wing pressures and boundary layer characteristics: these effects are also considered using coupled integral boundary layer codes and preliminary calculations using free vortex sheet separation modelling are included. Calculated results are compared with water tunnel experimental data with generally remarkably good agreement.

Theoretical and Experimental Investigation of the Subsonic-Flow Fields Beneath Swept and Unswept Wings with Tables or Vortex-induced Velocities

NASA Technical Reports Server (NTRS)

Alford, William J , Jr

1957-01-01

The flow-field characteristics beneath swept and unswept wings as determined by potential-flow theory are compared with the experimentally determined flow fields beneath swept and unswept wing-fuselage combinations. The potential-flow theory utilized considered both spanwise and chordwise distributions of vorticity as well as the wing-thickness effects. The perturbation velocities induced by a unit horseshoe vortex are included in tabular form. The theoretical predictions of the flow-field characteristics were qualitatively correct in all cases considered, although there were indications that the magnitudes of the downwash angles tended to be overpredicted as the tip of the swept wing was approached and that the sidewash angles ahead of the unswept wing were underpredicted. The calculated effects of compressibility indicated that significant increases in the chordwise variation of flow angles and dynamic-pressure ratios should be expected in going from low to high subsonic speeds.

Aerodynamic forces and flow structures of the leading edge vortex on a flapping wing considering ground effect.

PubMed

Van Truong, Tien; Byun, Doyoung; Kim, Min Jun; Yoon, Kwang Joon; Park, Hoon Cheol

2013-09-01

The aim of this work is to provide an insight into the aerodynamic performance of the beetle during takeoff, which has been estimated in previous investigations. We employed a scaled-up electromechanical model flapping wing to measure the aerodynamic forces and the three-dimensional flow structures on the flapping wing. The ground effect on the unsteady forces and flow structures were also characterized. The dynamically scaled wing model could replicate the general stroke pattern of the beetle's hind wing kinematics during takeoff flight. Two wing kinematic models have been studied to examine the influences of wing kinematics on unsteady aerodynamic forces. In the first model, the angle of attack is asymmetric and varies during the translational motion, which is the flapping motion of the beetle's hind wing. In the second model, the angle of attack is constant during the translational motion. The instantaneous aerodynamic forces were measured for four strokes during the beetle's takeoff by the force sensor attached at the wing base. Flow visualization provided a general picture of the evolution of the three-dimensional leading edge vortex (LEV) on the beetle hind wing model. The LEV is stable during each stroke, and increases radically from the root to the tip, forming a leading-edge spiral vortex. The force measurement results show that the vertical force generated by the hind wing is large enough to lift the beetle. For the beetle hind wing kinematics, the total vertical force production increases 18.4% and 8.6% for the first and second strokes, respectively, due to the ground effect. However, for the model with a constant angle of attack during translation, the vertical force is reduced during the first stroke. During the third and fourth strokes, the ground effect is negligible for both wing kinematic patterns. This finding suggests that the beetle's flapping mechanism induces a ground effect that can efficiently lift its body from the ground during takeoff.

Investigation of Turbulent Tip Leakage Vortex in an Axial Water Jet Pump with Large Eddy Simulation

NASA Technical Reports Server (NTRS)

Hah, Chunill; Katz, Joseph

2012-01-01

Detailed steady and unsteady numerical studies were performed to investigate tip clearance flow in an axial water jet pump. The primary objective is to understand physics of unsteady tip clearance flow, unsteady tip leakage vortex, and cavitation inception in an axial water jet pump. Steady pressure field and resulting steady tip leakage vortex from a steady flow analysis do not seem to explain measured cavitation inception correctly. The measured flow field near the tip is unsteady and measured cavitation inception is highly transient. Flow visualization with cavitation bubbles shows that the leakage vortex is oscillating significantly and many intermittent vortex ropes are present between the suction side of the blade and the tip leakage core vortex. Although the flow field is highly transient, the overall flow structure is stable and a characteristic frequency seems to exist. To capture relevant flow physics as much as possible, a Reynolds-averaged Navier-Stokes (RANS) calculation and a Large Eddy Simulation (LES) were applied for the current investigation. The present study reveals that several vortices from the tip leakage vortex system cross the tip gap of the adjacent blade periodically. Sudden changes in local pressure field inside tip gap due to these vortices create vortex ropes. The instantaneous pressure filed inside the tip gap is drastically different from that of the steady flow simulation. Unsteady flow simulation which can calculate unsteady vortex motion is necessary to calculate cavitation inception accurately even at design flow condition in such a water jet pump.

In-flight flow visualization results from the X-29A aircraft at high angles of attack

NASA Technical Reports Server (NTRS)

Delfrate, John H.; Saltzman, John A.

1992-01-01

Flow visualization techniques were used on the X-29A aircraft at high angles of attack to study the vortical flow off the forebody and the surface flow on the wing and tail. The forebody vortex system was studied because asymmetries in the vortex system were suspected of inducing uncommanded yawing moments at zero sideslip. Smoke enabled visualization of the vortex system and correlation of its orientation with flight yawing moment data. Good agreement was found between vortex system asymmetries and the occurrence of yawing moments. Surface flow on the forward-swept wing of the X-29A was studied using tufts and flow cones. As angle of attack increased, separated flow initiated at the root and spread outboard encompassing the full wing by 30 deg angle of attack. In general, the progression of the separated flow correlated well with subscale model lift data. Surface flow on the vertical tail was also studied using tufts and flow cones. As angle of attack increased, separated flow initiated at the root and spread upward. The area of separated flow on the vertical tail at angles of attack greater than 20 deg correlated well with the marked decrease in aircraft directional stability.

Vortex Flow Correlation

DTIC Science & Technology

1981-01-01

vorticity model used on the wing as well as on the leading-edge vortex sheet. Since the trailing-edge wake vorti- city does not have the close...z SECTION B-B ( WAKE ) FIGURE 11. FLOW PAST A SLENDER WING WITH LEADING-EDGE VORTEX FLOW 49 * -- A water tunnel is useful in visualizing the reversed...on fighter aircraft which generate strong vortical flows. The differences in apparent mass between a model in air and a model in water require analysis

Experimental and numerical study of the British Experimental Rotor Programme blade

NASA Technical Reports Server (NTRS)

Brocklehurst, Alan; Duque, Earl P. N.

1990-01-01

Wind-tunnel tests on the British Experimental Rotor Programme (BERP) tip are described, and the results are compared with computational fluid dynamics (CFD) results. The test model was molded using the Lynx-BERP blade tooling to provide a semispan, cantilever wing comprising the outboard 30 percent of the rotor blade. The tests included both surface-pressure measurements and flow visualization to obtain detailed information of the flow over the BERP tip for a range of angles of attack. It was observed that, outboard of the notch, favorable pressure gradients exist which ensure attached flow, and that the tip vortex also remains stable to large angles of attack. On the rotor, these features yield a very gradual break in control loads when the retreating-blade limit is eventually reached. Computational and experimental results were generally found to be in good agreement.

Effect of wing flexibility in dragonfly hovering flight

NASA Astrophysics Data System (ADS)

Naidu, Vishal; Young, John; Lai, Joseph

2011-11-01

Dragonflies have two pairs of tandem wings, which can be operated independently. Most studies on tandem wings are based on rigid wings, which is in strong contradiction to the natural, flexible dragonfly wings. The effect of wing flexibility in tandem wings is little known. We carry out a comparative, computational study between rigid and flexible, dragonfly shaped wings for hovering flight. In rigid wings during downstroke, a leading edge vortex (LEV) is formed on the upper surface, which forms a low pressure zone. This conical LEV joins the tip vortex and shortly after the mid downstroke when the wing starts to rotate, these vortices are gradually shed resulting in a drop in lift. The vortex system creates a net downwards momentum in the form of a jet. The flexible wings while in motion deform due to aerodynamic and inertial forces. Since there is a strong interaction between wing deformation and air flow around the deformed wings, flexible wing simulations are carried out using a two way fluid structure interaction. The effect of wing flexibility on the flow structure and the subsequent effect on the aerodynamic forces will be studied and presented.

Control of Flap Vortices

NASA Technical Reports Server (NTRS)

Greenblatt, David

2005-01-01

A wind tunnel investigation was carried out on a semi-span wing model to assess the feasibility of controlling vortices emanating from outboard flaps and tip-flaps by actively varying the degree of boundary layer separation. Separation was varied by means of perturbations produced from segmented zero-efflux oscillatory blowing slots, while estimates of span loadings and vortex sheet strengths were obtained by integrating wing surface pressures. These estimates were used as input to inviscid rollup relations as a means of predicting changes to the vortex characteristics resulting from the perturbations. Surveys of flow in the wake of the outboard and tip-flaps were made using a seven-hole probe, from which the vortex characteristics were directly deduced. Varying the degree of separation had a marked effect on vortex location, strength, tangential velocity, axial velocity and size for both outboard and tip-flaps. Qualitative changes in vortex characteristics were well predicted by the inviscid rollup relations, while the failure to account for viscosity was presumed to be the main reason for observed discrepancies. Introducing perturbations near the outboard flap-edges or on the tip-flap exerted significant control over vortices while producing negligible lift excursions.

Attenuation of the tip vortex flow using a flexible thread

NASA Astrophysics Data System (ADS)

Lee, Seung-Jae; Shin, Jin-Woo; Arndt, Roger E. A.; Suh, Jung-Chun

2018-01-01

Tip vortex cavitation (TVC) is important in a number of practical engineering applications. The onset of TVC is a critical concern for navy surface ships and submarines that aim to increase their capability to evade detection. A flexible thread attachment at blade tips was recently suggested as a new method to delay the onset of TVC. Although the occurrence of TVC can be reduced using a flexible thread, no scientific investigation focusing on its mechanisms has been undertaken. Thus, herein, we experimentally investigated the use of the flexible thread to suppress TVC from an elliptical wing. These investigations were performed in a cavitation tunnel and involved an observation of TVC using high-speed cameras, motion tracking of the thread using image-processing techniques, and near-field flow measurements performed using stereoscopic particle image velocimetry. The experimental data suggested that the flexible thread affects the axial velocity field more than the circumferential velocity field around the TVC axis. Furthermore, we observed no clear dependence of the vortex core size, circulation, and flow unsteadiness on TVC suppression. However, the presence of the thread at the wing tip led to a notable reduction in the streamwise velocity field, thereby alleviating TVC.

Short revolving wings enable hovering animals to avoid stall and reduce drag

NASA Astrophysics Data System (ADS)

Lentink, David; Kruyt, Jan W.; Heijst, Gertjan F.; Altshuler, Douglas L.

2014-11-01

Long and slender wings reduce the drag of airplanes, helicopters, and gliding animals, which operate at low angle of attack (incidence). Remarkably, there is no evidence for such influence of wing aspect ratio on the energetics of hovering animals that operate their wings at much higher incidence. High incidence causes aircraft wings to stall, hovering animals avoid stall by generating an attached vortex along the leading edge of their wings that elevates lift. Hypotheses that explain this capability include the necessity for a short radial distance between the shoulder joint and wing tip, measured in chord lengths, instead of the long tip-to-tip distance that elevates aircraft performance. This stems from how hovering animals revolve their wings around a joint, a condition for which the precise effect of aspect ratio on stall performance is unknown. Here we show that the attachment of the leading edge vortex is determined by wing aspect ratio with respect to the center of rotation-for a suite of aspect ratios that represent both animal and aircraft wings. The vortex remains attached when the local radius is shorter than 4 chord lengths, and separates outboard on more slender wings. Like most other hovering animals, hummingbirds have wing aspect ratios between 3 and 4, much stubbier than helicopters. Our results show this makes their wings robust against flow separation, which reduces drag below values obtained with more slender wings. This revises our understanding of how aspect ratio improves performance at low Reynolds numbers.

Wind Tunnel Investigation of the Effects of Surface Porosity and Vertical Tail Placement on Slender Wing Vortex Flow Aerodynamics at Supersonic Speeds

NASA Technical Reports Server (NTRS)

Erickson, Gary E.

2007-01-01

A wind tunnel experiment was conducted in the NASA Langley Research Center (LaRC) Unitary Plan Wind Tunnel (UPWT) to determine the effects of passive surface porosity and vertical tail placement on vortex flow development and interactions about a general research fighter configuration at supersonic speeds. Optical flow measurement and flow visualization techniques were used that featured pressure sensitive paint (PSP), laser vapor screen (LVS), and schlieren, These techniques were combined with conventional electronically-scanned pressure (ESP) and six-component force and moment measurements to quantify and to visualize the effects of flow-through porosity applied to a wing leading edge extension (LEX) and the placement of centerline and twin vertical tails on the vortex-dominated flow field of a 65 cropped delta wing model. Test results were obtained at free-stream Mach numbers of 1.6, 1.8, and 2.1 and a Reynolds number per foot of 2.0 million. LEX porosity promoted a wing vortex-dominated flow field as a result of a diffusion and weakening of the LEX vortex. The redistribution of the vortex-induced suction pressures contributed to large nose-down pitching moment increments but did not significantly affect the vortex-induced lift. The trends associated with LEX porosity were unaffected by vertical tail placement. The centerline tail configuration generally provided more stable rolling moments and yawing moments compared to the twin wing-mounted vertical tails. The strength of a complex system of shock waves between the twin tails was reduced by LEX porosity.

Application of the vortex-lattice technique to the analysis of thin wings with vortex separation and thick multi-element wings

NASA Technical Reports Server (NTRS)

Smith, C. W.; Bhateley, I. C.

1976-01-01

Two techniques for extending the range of applicability of the basic vortex-lattice method are discussed. The first improves the computation of aerodynamic forces on thin, low-aspect-ratio wings of arbitrary planforms at subsonic Mach numbers by including the effects of leading-edge and tip vortex separation, characteristic of this type wing, through use of the well-known suction-analogy method of E. C. Polhamus. Comparisons with experimental data for a variety of planforms are presented. The second consists of the use of the vortex-lattice method to predict pressure distributions over thick multi-element wings (wings with leading- and trailing-edge devices). A method of laying out the lattice is described which gives accurate pressures on the top and part of the bottom surface of the wing. Limited comparisons between the result predicted by this method, the conventional lattice arrangement method, experimental data, and 2-D potential flow analysis techniques are presented.

Helical vortices generated by flapping wings of bumblebees

NASA Astrophysics Data System (ADS)

Farge, Marie; Engels, Thomas; Kolomenskiy, Dmitry; Schneider, Kai; Lehmann, Fritz; Sesterhenn, Jörn

2016-11-01

We analyze high resolution numerical simulation data of a bumblebee with fixed body and prescribed wing motion, flying in a numerical wind tunnel, presented in. The inflow condition of the tunnel varies from unperturbed laminar to strongly turbulent. The flow generated by the flapping wings indicates the important role of the leading edge vortex (LEV), responsible for elevated lift production and which is not significantly altered by the inflow turbulence. The LEV has a conical structure due to the three-dimensional motion of the wings. This flow configuration produces strong vorticity on the sharp leading edge and the outwards velocity (from the root to the tip of the wing) in the spanwise direction. Flow visualizations show that the generated vortical structures are characterized by a strong helicity. We study the evolution of the mean helicity for each wing and analyze the impact of turbulent inflow. We thankfully acknowledge financial support from the French-German AIFIT project funded by DFG and ANR (Grant 15-CE40-0019). DK gratefully acknowledges financial support from the JSPS postdoctoral fellowship.

A Visualization Study of Secondary Flows in Cascades

NASA Technical Reports Server (NTRS)

Herzig, Howard Z; Hansen, Arthur G; Costello, George R

1954-01-01

Flow-visualization techniques are employed to ascertain the streamline patterns of the nonpotential secondary flows in the boundary layers of cascades, and thereby to provide a basis for more extended analyses in turbomachines. The three-dimensional deflection of the end-wall boundary layer results in the formation of a vortex within each cascade passage. The size and tightness of the vortex generated depend upon the main-flow turning in the cascade passage. Once formed, a vortex resists turning in subsequent blade rows, with consequent unfavorable angles of attack and possible flow disturbances on the pressure surfaces of subsequent blade rows when the vortices impinge on these surfaces. Two major tip-clearance effects are observed, the formation of a tip-clearance vortex and the scraping effect of a blade with relative motion past the wall boundary layer. The flow patterns indicate methods for improving the blade tip-loading characteristics of compressors and of low- and high-speed turbulence.

Theoretical and Experimental Investigation of the Subsonic-Flow Fields Beneath Swept and Unswept Wings with Tables of Vortex-Induced Velocities

NASA Technical Reports Server (NTRS)

Alford, William J., Jr.

1956-01-01

The flow-field characteristics beneath swept and unswept wings as determined by potential-flow theory are compared with the experimentally determined flow fields beneath swept and unswept wing-fuselage combinations. The potential-flow theory utilized considered both spanwise and chordwise distributions of vorticity as well as the wing-thickness effects. The perturbation velocities induced by a unit horseshoe vortex are included in tabular form. The results indicated that significant chordwise flow gradients existed beneath both swept and unswept wings at zero lift and throughout the lift range. The theoretical predictions of the flow-field characteristics were qualitatively correct in all cases considered, although there were indications that the magnitudes of the downwash angles tended to be overpredicted as the tip of the swept wing was approached and that the sidewash angles ahead of the unswept wing were underpredicted. The calculated effects of compressibility indicated that significant increases in the chordwise variation of flow angles and dynamic-pressure ratios should be expected in going from low to high subsonic speeds.

Longitudinal vortex control - Techniques and applications (The 32nd Lanchester Lecture)

NASA Technical Reports Server (NTRS)

Bushnell, D. M.

1992-01-01

A summary is presented of vortex control applications and current techniques for the control of longitudinal vortices produced by bodies, leading edges, tips and intersections. Vortex control has up till now been performed by many approaches in an empirical fashion, assisted by the essentially inviscid nature of much of longitudinal vortex behavior. Attention is given to Reynolds number sensitivities, vortex breakdown and interactions, vortex control on highly swept wings, and vortex control in juncture flows.

A study of high alpha dynamics and flow visualization for a 2.5-percent model of the F-18 HARV undergoing wing rock

NASA Technical Reports Server (NTRS)

Quast, Thomas; Nelson, Robert C.; Fisher, David F.

1991-01-01

Free-to-roll experiments and flow visualization studies have been conducted for a 2.5-percent model of the F-18 undergoing unsteady wing rock oscillations. Data have been acquired in the form of roll angle time histories as well as video recordings and 35 mm photography of the forebody and leading edge extension vortices. The time histories were differentiated to produce angular velocity and angular acceleration. From this the roll moment as a function of time and/or roll angle could be estimated. A thorough analysis of the data has revealed a genuine wing-rock phenomenon. Off-surface flow visualization was used to identiify the forebody and LEX vortex core positions and their interaction in both static and dynamic configurations. A direct correlation between the dynamic data and visualized vortex activity during the wing-rock motion has been made.

Visualization and Quantification of Rotor Tip Vortices in Helicopter Flows

NASA Technical Reports Server (NTRS)

Kao, David L.; Ahmad, Jasim U.; Holst, Terry L.

2015-01-01

This paper presents an automated approach for effective extraction, visualization, and quantification of vortex core radii from the Navier-Stokes simulations of a UH-60A rotor in forward flight. We adopt a scaled Q-criterion to determine vortex regions and then perform vortex core profiling in these regions to calculate vortex core radii. This method provides an efficient way of visualizing and quantifying the blade tip vortices. Moreover, the vortices radii are displayed graphically in a plane.

Flow Structure on a Flapping Wing: Quasi-Steady Limit

NASA Astrophysics Data System (ADS)

Ozen, Cem; Rockwell, Donald

2011-11-01

The flapping motion of an insect wing typically involves quasi-steady motion between extremes of unsteady motion. This investigation characterizes the flow structure for the quasi-steady limit via a rotating wing in the form of a thin rectangular plate having a low aspect ratio (AR =1). Particle Image Velocimetry (PIV) is employed, in order to gain insight into the effects of centripetal and Coriolis forces. Vorticity, velocity and streamline patterns are used to describe the overall flow structure with an emphasis on the leading-edge vortex. A stable leading-edge vortex is maintained over effective angles of attack from 30° to 75° and it is observed that at each angle of attack the flow structure remains relatively same over the Reynolds number range from 3,600 to 14,500. The dimensionless circulation of the leading edge vortex is found to be proportional to the effective angle of attack. Quasi-three-dimensional construction of the flow structure is used to identify the different regimes along the span of the wing which is then complemented by patterns on cross flow planes to demonstrate the influence of root and tip swirls on the spanwise flow. The rotating wing results are also compared with the equivalent of translating wing to further illustrate the effects of the rotation.

Analytical investigation of aerodynamic characteristics of highly swept wings with separated flow

NASA Technical Reports Server (NTRS)

Reddy, C. S.

1980-01-01

Many modern aircraft designed for supersonic speeds employ highly swept-back and low-aspect-ratio wings with sharp or thin edges. Flow separation occurs near the leading and tip edges of such wings at moderate to high angles of attack. Attempts have been made over the years to develop analytical methods for predicting the aerodynamic characteristics of such aircraft. Before any method can really be useful, it must be tested against a standard set of data to determine its capabilities and limitations. The present work undertakes such an investigation. Three methods are considered: the free-vortex-sheet method (Weber et al., 1975), the vortex-lattice method with suction analogy (Lamar and Gloss, 1975), and the quasi-vortex lattice method of Mehrotra (1977). Both flat and cambered wings of different configurations, for which experimental data are available, are studied and comparisons made.

Tracking Blade Tip Vortices for Numerical Flow Simulations of Hovering Rotorcraft

NASA Technical Reports Server (NTRS)

Kao, David L.

2016-01-01

Blade tip vortices generated by a helicopter rotor blade are a major source of rotor noise and airframe vibration. This occurs when a vortex passes closely by, and interacts with, a rotor blade. The accurate prediction of Blade Vortex Interaction (BVI) continues to be a challenge for Computational Fluid Dynamics (CFD). Though considerable research has been devoted to BVI noise reduction and experimental techniques for measuring the blade tip vortices in a wind tunnel, there are only a handful of post-processing tools available for extracting vortex core lines from CFD simulation data. In order to calculate the vortex core radius, most of these tools require the user to manually select a vortex core to perform the calculation. Furthermore, none of them provide the capability to track the growth of a vortex core, which is a measure of how quickly the vortex diffuses over time. This paper introduces an automated approach for tracking the core growth of a blade tip vortex from CFD simulations of rotorcraft in hover. The proposed approach offers an effective method for the quantification and visualization of blade tip vortices in helicopter rotor wakes. Keywords: vortex core, feature extraction, CFD, numerical flow visualization

Vortex coupling in trailing vortex-wing interactions

NASA Astrophysics Data System (ADS)

Chen, C.; Wang, Z.; Gursul, I.

2018-03-01

The interaction of trailing vortices of an upstream wing with rigid and flexible downstream wings has been investigated experimentally in a wind tunnel, using particle image velocimetry, hot-wire, force, and deformation measurements. Counter-rotating upstream vortices exhibit increased meandering when they are close to the tip of the downstream wing. The upstream vortex forms a pair with the vortex shed from the downstream wing and then exhibits large displacements around the wing tip. This coupled motion of the pair has been found to cause large lift fluctuations on the downstream wing. The meandering of the vortex pair occurs at the natural meandering frequency of the isolated vortex, with a low Strouhal number, and is not affected by the frequency of the large-amplitude wing oscillations if the downstream wing is flexible. The displacement of the leading vortex is larger than that of the trailing vortex; however, it causes highly correlated variations of the core radius, core vorticity, and circulation of the trailing vortex with the coupled meandering motion. In contrast, co-rotating vortices do not exhibit any increased meandering.

The effect of asymmetric vortex wake characteristics on a slender delta wing undergoing wing rock motion

NASA Technical Reports Server (NTRS)

Arena, A. S., Jr.; Nelson, R. C.

1989-01-01

An experimental investigation into the fluid mechanisms responsible for wing rock on a slender delta wing with 80 deg leading edge sweep has been conducted. Time history and flow visualization data are presented for a wide angle-of-attack range. The use of an air bearing spindle has allowed the motion of the wing to be free from bearing friction or mechanical hysteresis. A bistable static condition has been found in vortex breakdown at an angle of attack of 40 deg which causes an overshoot of the steady state rocking amplitude. Flow visualization experiments also reveal a difference in static and dynamic breakdown locations on the wing. A hysteresis loop in dynamic breakdown location similar to that seen on pitching delta wings was observed as the wing was undergoing the limit cycle oscillation.

Vortex formation and saturation for low-aspect-ratio rotating flat-plate fins

NASA Astrophysics Data System (ADS)

Devoria, Adam C.; Ringuette, Matthew J.

2012-02-01

We investigate experimentally the unsteady, three-dimensional vortex formation of low-aspect-ratio, trapezoidal flat-plate fins undergoing rotation from rest at a 90° angle of attack and Reynolds numbers of O(103). The objectives are to characterize the unsteady three-dimensional vortex structure, examine vortex saturation, and understand the effects of the root-to-tip flow for different velocity programs. The experiments are conducted in a water tank facility, and the diagnostic tools are dye flow visualization and digital particle image velocimetry. The dye visualizations show that the low-aspect-ratio plate produces symmetric ring-like vortices comprised mainly of tip-edge vorticity. They also indicate the presence of the root-to-tip velocity. For large rotational amplitudes, the primary ring-like vortex sheds and a secondary ring-like vortex is generated while the plate is still in motion, indicating saturation of the leading vortex. The time-varying vortex circulation in the flow symmetry plane provides quantitative evidence of vortex saturation. The phenomenon of saturation is observed for several plate velocity programs. The temporal development of the vortex circulation is often complex, which prevents an objective determination of an exact saturation time. This is the result of an interaction between the developing vortex and the root-to-tip flow, which breaks apart the vortex. However, it is possible to define a range of time during which the vortex reaches saturation. A formation-parameter definition is investigated and is found to reasonably predict the state corresponding to the pinch-off of the initial tip vortex across the velocity programs tested. This event is the lower bound on the saturation time range.

Measurement of vortex velocities over a wide range of vortex age, downstream distance and free stream velocity

NASA Technical Reports Server (NTRS)

Rorke, J. B.; Moffett, R. C.

1977-01-01

A wind tunnel test was conducted to obtain vortex velocity signatures over a wide parameter range encompassing the data conditions of several previous researchers while maintaining a common instrumentation and test facility. The generating wing panel was configured with both a revolved airfoil tip shape and a square tip shape and had a semispan aspect of 4.05/1.0 with a 121.9 cm span. Free stream velocity was varied from 6.1 m/sec to 76.2 m/sec and the vortex core velocities were measured at locations 3, 6, 12, 24 and 48 chordlengths downstream of the wing trailing edge, yielding vortex ages up to 2.0 seconds. Wing pitch angles of 6, 8, 9 and 12 deg were investigated. Detailed surface pressure distributions and wing force measurements were obtained for each wing tip configuration. Correlation with vortex velocity data taken in previous experiments is good. During the rollup process, vortex core parameters appear to be dependent primarily on vortex age. Trending in the plateau and decay regions is more complex and the machanisms appear to be more unstable.

Computational investigation of cicada aerodynamics in forward flight.

PubMed

Wan, Hui; Dong, Haibo; Gai, Kuo

2015-01-06

Free forward flight of cicadas is investigated through high-speed photogrammetry, three-dimensional surface reconstruction and computational fluid dynamics simulations. We report two new vortices generated by the cicada's wide body. One is the thorax-generated vortex, which helps the downwash flow, indicating a new phenomenon of lift enhancement. Another is the cicada posterior body vortex, which entangles with the vortex ring composed of wing tip, trailing edge and wing root vortices. Some other vortex features include: independently developed left- and right-hand side leading edge vortex (LEV), dual-core LEV structure at the mid-wing region and near-wake two-vortex-ring structure. In the cicada forward flight, approximately 79% of the total lift is generated during the downstroke. Cicada wings experience drag in the downstroke, and generate thrust during the upstroke. Energetics study shows that the cicada in free forward flight consumes much more power in the downstroke than in the upstroke, to provide enough lift to support the weight and to overcome drag to move forward.

Computational investigation of cicada aerodynamics in forward flight

PubMed Central

Wan, Hui; Dong, Haibo; Gai, Kuo

2015-01-01

Free forward flight of cicadas is investigated through high-speed photogrammetry, three-dimensional surface reconstruction and computational fluid dynamics simulations. We report two new vortices generated by the cicada's wide body. One is the thorax-generated vortex, which helps the downwash flow, indicating a new phenomenon of lift enhancement. Another is the cicada posterior body vortex, which entangles with the vortex ring composed of wing tip, trailing edge and wing root vortices. Some other vortex features include: independently developed left- and right-hand side leading edge vortex (LEV), dual-core LEV structure at the mid-wing region and near-wake two-vortex-ring structure. In the cicada forward flight, approximately 79% of the total lift is generated during the downstroke. Cicada wings experience drag in the downstroke, and generate thrust during the upstroke. Energetics study shows that the cicada in free forward flight consumes much more power in the downstroke than in the upstroke, to provide enough lift to support the weight and to overcome drag to move forward. PMID:25551136

Experimental Study of the Structure of a Wingtip Vortex

NASA Technical Reports Server (NTRS)

Anderson, Elgin A.; Wright, Christopher T.

2000-01-01

A complete look at the near-field development and subsequent role-up of a wingtip vortex from a NACA 0015 wing section is investigated. Two separate but equally important surveys of the vortex structure in the region adjacent to the wingtip and approximately one chord length downstream of the trailing edge are performed. The two surveys provide qualitative flow-visualization an quantitative velocity measurement data. The near-field development and subsequent role-up of the vortex structures is strongly influenced by the angle-of-attack and the end-cap treatment of the wing section. The velocity field near the wingtip of the NACA 0015 wing section was measured with a triple-sensor hot wire probe and compared to flow visualization images produced with titanium tetrachloride smoke injection and laser illumination. The flat end-cap results indicate the formation of multiple, relatively strong vortex structures as opposed to the formation of a single vortex produced with the round end-cap. The multiple vortices generated by the flat end-cap are seen to rotate around a common ce te in a helical pattern until they eventually merge into a single vortex. Compared to a non-dimensional loading parameter, the results of the velocity and flow visualization data shows a "jetlike" axial velocity profile for loading parameter values on the order of 0.1 and a "wakelike" profile for much lower loading parameter values.

Effects of Passive Porosity on Interacting Vortex Flows At Supersonic Speeds

NASA Technical Reports Server (NTRS)

Erickson, Gary E.

2000-01-01

A wind tunnel experiment was conducted in the NASA Langley Research Center (LaRC) Unitary Plan Wind Tunnel (UPWT) to determine the effects of passive surface porosity on vortex flow interaction about a general research fighter configuration at supersonic speeds. Optical flow measurement and flow visualization techniques were used and included pressure-sensitive paint (PSP), schlieren, and laser vapor screen (LVS) These techniques were combined with force and moment and conventional electronically-scanned pressure (ESP) measurements to quantify and to visualize the effects of flow-through porosity applied to a wing leading-edge extension (LEX) mounted to a 65 deg cropped delta wing model.

Control of Interacting Vortex Flows at Subsonic and Transonic Speeds Using Passive Porosity

NASA Technical Reports Server (NTRS)

Erickson, Gary E.

2003-01-01

A wind tunnel experiment was conducted in the NASA Langley Research Center (LaRC) 8-Foot Transonic Pressure Tunnel (TPT) to determine the effects of passive surface porosity on vortex flow interactions about a general research fighter configuration at subsonic and transonic speeds. Flow-through porosity was applied to a wing leading-edge extension (LEX) mounted to a 65 deg cropped delta wing model to promote large nose-down pitching moment increments at high angles of attack. Porosity decreased the vorticity shed from the LEX, which weakened the LEX vortex and altered the global interactions of the LEX and wing vortices at high angles of attack. Six-component forces and moments and wing upper surface static pressure distributions were obtained at free-stream Mach numbers of 0.50, 0.85, and 1.20, Reynolds number of 2.5(10(exp 6)) per foot, angles of attack up to 30 deg, and angles of sideslip to +/- 8 deg. The off-surface flow field was visualized in selected cross-planes using a laser vapor screen flow visualization technique. Test data were obtained with a centerline vertical tail and with alternate twin, wing-mounted vertical fins having 0 deg and 30 deg cant angles. In addition, the porosity of the LEX was compartmentalized to determine the sensitivity of the vortex-dominated aerodynamics to the location and level of porosity applied to the LEX.

The screw propeller

NASA Astrophysics Data System (ADS)

Larrabee, E. E.

1980-07-01

Marine and air screw propellers are considered in terms of theoretical hydrodynamics as developed by Joukowsky, Prandtl, and Betz. Attention is given to the flow around wings of finite span where spanwise flow exists and where lift and the bound vorticity must all go smoothly to zero at the wing tips. The concept of a trailing vortex sheet made up of infinitesimal line vortexes roughly aligned with the direction of flight is discussed in this regard. Also considered is induced velocity, which tends to convect the sheet downward at every stage in the roll-up process, the vortex theory of propellers and the Betz-Prandtl circulation distribution. The performance of the Gossamer Albatross and of a pedal-driven biplane called the Chrysalis are also discussed.

Water Tunnel Flow Visualization Study Through Poststall of 12 Novel Planform Shapes

NASA Technical Reports Server (NTRS)

Gatlin, Gregory M.; Neuhart, Dan H.

1996-01-01

To determine the flow field characteristics of 12 planform geometries, a flow visualization investigation was conducted in the Langley 16- by 24-Inch Water Tunnel. Concepts studied included flat plate representations of diamond wings, twin bodies, double wings, cutout wing configurations, and serrated forebodies. The off-surface flow patterns were identified by injecting colored dyes from the model surface into the free-stream flow. These dyes generally were injected so that the localized vortical flow patterns were visualized. Photographs were obtained for angles of attack ranging from 10' to 50', and all investigations were conducted at a test section speed of 0.25 ft per sec. Results from the investigation indicate that the formation of strong vortices on highly swept forebodies can improve poststall lift characteristics; however, the asymmetric bursting of these vortices could produce substantial control problems. A wing cutout was found to significantly alter the position of the forebody vortex on the wing by shifting the vortex inboard. Serrated forebodies were found to effectively generate multiple vortices over the configuration. Vortices from 65' swept forebody serrations tended to roll together, while vortices from 40' swept serrations were more effective in generating additional lift caused by their more independent nature.

Spanwise loading distribution and wake velocity surveys of a semi-span wing

NASA Technical Reports Server (NTRS)

Felker, F. F., III; Piziali, R. A.; Gall, J. K.

1982-01-01

The spanwise distribution of bound circulation on a semi-span wing and the flow velocities in its wake were measured in a wind tunnel. Particular attention was given to documenting the flow velocities in and around the development tip vortex. A two-component laser velocimeter was used to make the velocity measurements. The spanwise distribution of bound circulation, three components of the time-averaged velocities throughout the near wake their standard deviations, and the integrated forces and moments on a metric tip as measured by an internal strain gage balance are presented without discussion.

Definition of the unsteady vortex flow over a wing/body configuration

NASA Technical Reports Server (NTRS)

Liou, S. G.; Debry, B.; Lenakos, J.; Caplin, J.; Komerath, N. M.

1991-01-01

A problem of current interest in computational aerodynamics is the prediction of unsteady vortex flows over aircraft at high angles of attack. A six-month experimental effort was conducted at the John H. Harper Wind Tunnel to acquire qualitative and quantitative information on the unsteady vortex flow over a generic wing-body configuration at high angles of attack. A double-delta flat-plate wing with beveled edges was combined with a slender sharp-nosed body-of-revolution fuselage to form the generic configuration. This configuration produces a strong attached leading edge vortex on the wing, as well as sharply-peaked flow velocity spectra above the wing. While it thus produces flows with several well-defined features of current interest, the model was designed for efficiency of representation in computational codes. A moderate number of surface pressure ports and two unsteady pressure sensors were used to study the pressure distribution over the wing and body surface at high angles of attack; the unsteady pressure sensing did not succeed because of inadequate signal-to-noise ratio. A pulsed copper vapor laser sheet was used to visualize the vortex flow over the model, and vortex trajectories, burst locations, mutual induction of vortex systems from the forebody, strake, and wing, were quantified. Laser Doppler velocimetry was used to quantify all 3 components of the time-average velocity in 3 data planes perpendicular to the freestream direction. Statistics of the instantaneous velocity were used to study intermittency and fluctuation intensity. Hot-film anemometry was used to study the fluctuation energy content in the velocity field, and the spectra of these fluctuations. In addition, a successful attempt was made to measure velocity spectra, component by component, using laser velocimetry, and these were compared with spectra measured by hot-film anemometry at several locations.

Subsonic flow investigations on a cranked wing designed for high maneuverability

NASA Technical Reports Server (NTRS)

Rao, D. M.

1986-01-01

The characteristic pitching moment nonlinearity of cranked wings limits their usable lift coefficient well below C sub L max. The potential of several aerodynamic devices, viz., fences, pylon vortex generators (PVG), mid-span strakes and cavity flaps, in delaying the pitch up onset on a 70/50 deg cranked wing was explored in low speed tunnel tests. Upper surface pressure measurements and low visualizations were conducted on a semi-span wing model to observe the vortex flow development with increasing angle of attack, and then to assess the effectiveness of the devices in controlling the collapse of vortex lift over the wing panel outboard of the crank. Force tests on a full span wing and body model were also conducted to assess the fence and PVG in improving the usable C sub L.

An exploratory study of apex fence flaps on a 74 deg delta wing

NASA Technical Reports Server (NTRS)

Wahls, R. A.; Vess, R. J.

1985-01-01

An exploratory wind tunnel investigation was performed to observe the flow field effects produced by vertically deployed apex fences on a planar 74 degree delta wing. The delta shaped fences, each comprising approximately 3.375 percent of the wing area, were affixed along the first 25 percent of the wing leading edge in symmetric as well as asymmetric (i.e., fence on one side only) arrangements. The vortex flow field was visualized at angles of attack from 0 to 20 degrees using helium bubble and oil flow techniques; upper surface pressures were also measured along spanwise rows. The results were used to construct a preliminary description of the vortex patterns and induced pressures associated with vertical apex fence deployment. The objective was to obtain an initial evaluation of the potential of apex fences as vortex devices for subsonic lift modulation as well as lateral directional control of delta wing aircraft.

Flow Visualization Techniques in Wind Tunnel Tests of a Full-Scale F/A-18 Aircraft

NASA Technical Reports Server (NTRS)

Lanser, Wendy R.; Botha, Gavin J.; James, Kevin D.; Bennett, Mark; Crowder, James P.; Cooper, Don; Olson, Lawrence (Technical Monitor)

1994-01-01

The proposed paper presents flow visualization performed during experiments conducted on a full-scale F/A-18 aircraft in the 80- by 120-Foot Wind-Tunnel at NASA Ames Research Center. The purpose of the flow-visualization experiments was to document the forebody and leading edge extension (LEX) vortex interaction along with the wing flow patterns at high angles of attack and low speed high Reynolds number conditions. This investigation used surface pressures in addition to both surface and off-surface flow visualization techniques to examine the flow field on the forebody, canopy, LEXS, and wings. The various techniques used to visualize the flow field were fluorescent tufts, flow cones treated with reflective material, smoke in combination with a laser light sheet, and a video imaging system for three-dimension vortex tracking. The flow visualization experiments were conducted over an angle of attack range from 20 deg to 45 deg and over a sideslip range from -10 deg to 10 deg. The various visualization techniques as well as the pressure distributions were used to understand the flow field structure. The results show regions of attached and separated flow on the forebody, canopy, and wings as well as the vortical flow over the leading-edge extensions. This paper will also present flow visualization comparisons with the F-18 HARV flight vehicle and small-scale oil flows on the F-18.

Wind Tunnel Investigation of Passive Vortex Control and Vortex-Tail Interactions on a Slender Wing at Subsonic and Transonic Speeds

NASA Technical Reports Server (NTRS)

Erickson, Gary E.

2013-01-01

A wind tunnel experiment was conducted in the NASA Langley 8-Foot Transonic Pressure Tunnel to determine the effects of passive porosity on vortex flow interactions about a slender wing configuration at subsonic and transonic speeds. Flow-through porosity was applied in several arrangements to a leading-edge extension, or LEX, mounted to a 65-degree cropped delta wing as a longitudinal instability mitigation technique. Test data were obtained with LEX on and off in the presence of a centerline vertical tail and twin, wing-mounted vertical fins to quantify the sensitivity of the aerodynamics to tail placement and orientation. A close-coupled canard was tested as an alternative to the LEX as a passive flow control device. Wing upper surface static pressure distributions and six-component forces and moments were obtained at Mach numbers of 0.50, 0.85, and 1.20, unit Reynolds number of 2.5 million, angles of attack up to approximately 30 degrees, and angles of sideslip to +/-8 degrees. The off-surface flow field was visualized in cross planes on selected configurations using a laser vapor screen flow visualization technique. Tunnel-to-tunnel data comparisons and a Reynolds number sensitivity assessment were also performed. 15.

Experimental and Computational Study of the Flow past a Simplified Geometry of an Engine/Pylon/Wing Installation at low velocity/moderate incidence flight conditions

NASA Astrophysics Data System (ADS)

Bury, Yannick; Lucas, Matthieu; Bonnaud, Cyril; Joly, Laurent; ISAE Team; Airbus Team

2014-11-01

We study numerically and experimentally the vortices that develop past a model geometry of a wing equipped with pylon-mounted engine at low speed/moderate incidence flight conditions. For such configuration, the presence of the powerplant installation under the wing initiates a complex, unsteady vortical flow field at the nacelle/pylon/wing junctions. Its interaction with the upper wing boundary layer causes a drop of aircraft performances. In order to decipher the underlying physics, this study is initially conducted on a simplified geometry at a Reynolds number of 200000, based on the chord wing and on the freestream velocity. Two configurations of angle of attack and side-slip angle are investigated. This work relies on unsteady Reynolds Averaged Navier Stokes computations, oil flow visualizations and stereoscopic Particle Image Velocimetry measurements. The vortex dynamics thus produced is described in terms of vortex core position, intensity, size and turbulent intensity thanks to a vortex tracking approach. In addition, the analysis of the velocity flow fields obtained from PIV highlights the influence of the longitudinal vortex initiated at the pylon/wing junction on the separation process of the boundary layer near the upper wing leading-edge.

Skin friction fields on delta wings

NASA Astrophysics Data System (ADS)

Woodiga, S. A.; Liu, Tianshu

2009-12-01

The normalized skin friction fields on a 65° delta wing and a 76°/40° double-delta wing are measured by using a global luminescent oil-film skin friction meter. The detailed topological structures of skin friction fields on the wings are revealed for different angles of attack and the important features are detected such as reattachment lines, secondary separation lines, vortex bursting and vortex interaction. The comparisons with the existing flow visualization results are discussed.

An experimental investigation of vortex breakdown on a delta wing

NASA Technical Reports Server (NTRS)

Payne, F. M.; Nelson, R. C.

1986-01-01

An experimental investigation of vortex breakdown on delta wings at high angles is presented. Thin delta wings having sweep angles of 70, 75, 80 and 85 degrees are being studied. Smoke flow visualization and the laser light sheet technique are being used to obtain cross-sectional views of the leading edge vortices as they break down. At low tunnel speeds (as low as 3 m/s) details of the flow, which are usually imperceptible or blurred at higher speeds, can be clearly seen. A combination of lateral and longitudinal cross-sectional views provides information on the three dimensional nature of the vortex structure before, during and after breakdown. Whereas details of the flow are identified in still photographs, the dynamic characteristics of the breakdown process were recorded using high speed movies. Velocity measurements were obtained using a laser Doppler anemometer with the 70 degree delta wing at 30 degrees angle of attack. The measurements show that when breakdown occurs the core flow transforms from a jet-like flow to a wake-like flow.

A water tunnel flow visualization study of the F-15

NASA Technical Reports Server (NTRS)

Lorincz, D. J.

1978-01-01

Water tunnel studies were performed to qualitatively define the flow field of the F-15 aircraft. Two lengthened forebodies, one with a modified cross-sectional shape, were tested in addition to the basic forebody. Particular emphasis was placed on defining vortex flows generated at high angles of attack. The flow visualization tests were conducted in the Northrop diagnostic water tunnel using a 1/48-scale model of the F-15. Flow visualization pictures were obtained over an angle-of-attack range to 55 deg and sideslip angles up to 10 deg. The basic aircraft configuration was investigated in detail to determine the vortex flow field development, vortex path, and vortex breakdown characteristics as a function of angle of attack and sideslip. Additional tests showed that the wing upper surface vortex flow fields were sensitive to variations in inlet mass flow ratio and inlet cowl deflection angle. Asymmetries in the vortex systems generated by each of the three forebodies were observed in the water tunnel at zero sideslip and high angles of attack.

Farfield structure of an aircraft trailing vortex, including effects of mass injection

NASA Technical Reports Server (NTRS)

Mason, W. H.; Marchman, J. F., III

1972-01-01

Wind tunnel tests to predict the aircraft wake turbulence due to the tip trailing vortex are discussed. A yawhead pressure probe was used in a subsonic wind tunnel to obtain detailed mean flow measurements at stations up to 30 chordlengths downstream in an aircraft trailing vortex. Mass injection at the wingtip was shown to hasten the decay of the trailing vortex. A theoretical method is presented to show the effect which the circulation distribution on the wing has on the structure of the outer portion of the vortex.

On the aerodynamic forces of flapping finite-wings in forward flight: a numerical study

NASA Astrophysics Data System (ADS)

Gonzalo, Alejandro; Uhlmann, Markus; Garcia-Villalba, Manuel; Flores, Oscar

2017-11-01

We study the flow around two flapping wings in forward flight at a low Reynolds number, Re = 500 , with 3D direct numerical simulations. The flow solver used is TUCAN, an in-house code which solves the Navier-Stokes equations for incompressible flow using an immersed boundary method to model the presence of the wings. The wings are rectangular with a NACA0012 airfoil of chord c as a cross-section. They are located side by side at a distance 0.5 c between their inboard tips. The wings flap with respect to an axis parallel to the streamwise velocity, without pitching. The angle of rotation is defined using a sinusoidal function with a reduced frequency k = 1 and an amplitude such that the maximum height of the outboard tips is c in all cases. We perform several simulations varying the aspect ratio of the wings (AR = 2 and 4) and the distance between the inboard tip of the wings and the axis of rotation (R = 0 , 2 and ∞), the latter case corresponding to wings in heaving motion. In this way we can study the variation of the fictitious forces on the wings and the induced spanwise flows, and their relation to the vortical structures on the wing (i.e. leading edge vortex, trailing edge votex, tip vortices) and the resulting aerodynamic forces. This work was funded by project TRA2013-41103-P (Mineco/Feder UE). The simulations were partially performed at the Steinbuch Centre for Computing, Karlsruhe, whose support is thankfully acknowledged.

ARC-1964-AC-33500-2

NASA Image and Video Library

1964-10-01

DURING APPROACH. OGEE Wing Planform on modified F5D-1 SkylancerAirplane Flight Tests. 'Flow Visualization Photographs'. In landing approach trials at Moffett Field, vapor trails are generated by low pressure in votex flow near wing leading edge on upper wing surface. Studies were undertaken in efforts to determine if there were adverse effects of vortex flow on the dynamic stability of the aircraft.

Lift enhancement by trapped vortex

NASA Technical Reports Server (NTRS)

Rossow, Vernon J.

1992-01-01

The viewgraphs and discussion of lift enhancement by trapped vortex are provided. Efforts are continuously being made to find simple ways to convert wings of aircraft from an efficient cruise configuration to one that develops the high lift needed during landing and takeoff. The high-lift configurations studied here consist of conventional airfoils with a trapped vortex over the upper surface. The vortex is trapped by one or two vertical fences that serve as barriers to the oncoming stream and as reflection planes for the vortex and the sink that form a separation bubble on top of the airfoil. Since the full three-dimensional unsteady flow problem over the wing of an aircraft is so complicated that it is hard to get an understanding of the principles that govern the vortex trapping process, the analysis is restricted here to the flow field illustrated in the first slide. It is assumed that the flow field between the two end plates approximates a streamwise strip of the flow over a wing. The flow between the endplates and about the airfoil consists of a spanwise vortex located between the suction orifices in the endplates. The spanwise fence or spoiler located near the nose of the airfoil serves to form a separated flow region and a shear layer. The vorticity in the shear layer is concentrated into the vortex by withdrawal of fluid at the suction orifices. As the strength of the vortex increases with time, it eventually dominates the flow in the separated region so that a shear or vertical layer is no longer shed from the tip of the fence. At that point, the vortex strength is fixed and its location is such that all of the velocity contributions at its center sum to zero thereby making it an equilibrium point for the vortex. The results of a theoretical analysis of such an idealized flow field are described.

On the Use pf Active Flow Control to Trim and Control a Tailles Aircraft Model

NASA Astrophysics Data System (ADS)

Jentzsch, Marvin

The Stability And Control CONfiguration (SACCON) model represents an emerging trend in airplane design where the classical tube, wing and empennage are replaced by a single tailless configuration. The challenge is to assure that these designs are stable and controllable. Nonlinear aerodynamic behavior is observed on the SACCON at higher incidence angles due to leading edge vortex structures. Active Flow Control (AFC) used in preliminary design represents a promising solution to the longitudinal stability problems and this was demonstrated experimentally on a semi span model. AFC can be used to trim the SACCON in pitch and it alters forces and moments comparable to common control surface deflections. A combination of AFC and control surface deflection may increase the overall efficiency and opens up a variety of maneuvering possibilities. This implies that AFC should be treated concomitantly with other design parameters and should be considered in the preliminary design process already and not as an add-on tool. Integral force and moment data was supplemented by observations using Pressure Sensitive Paint (PSP) and flow visualization. Two arrays of individually controlled sweeping jets, one located along the leading edge and the other along the flap hinge provided the AFC input needed to alter the flow. The array positioned over the flap-hinge of the model was most effective in stabilizing the wing by decreasing the pitching moment at lower and intermediate angles of incidence. This effect was achieved by reducing the spanwise flow on the swept back portion of the wing through jet-entrainment that also affected the leading edge vortex. Leading edge actuation showed some beneficial effects by inhibiting the formation of the leading edge vortex near the wing tip. A preliminary study using suction was carried out. The tests were carried out at Mach numbers smaller than 0.2 and Reynolds numbers based on the root chord of the model that approached 106.

NACA 0015 wing pressure and trailing vortex measurements

NASA Technical Reports Server (NTRS)

Mcalister, K. W.; Takahashi, R. K.

1991-01-01

A NACA 0015 semispan wing was placed in a low-speed wind tunnel, and measurements were made of the pressure on the upper and lower surface of the wing and of velocity across the vortex trailing downstream from the tip of the wing. Pressure data were obtained for both 2-D and 3-D configurations. These data feature a detailed comparison between wing tips with square and round lateral edges. A two-component laser velocimeter was used to measure velocity profiles across the vortex at numerous stations behind the wing and for various combinations of conditions. These conditions include three aspect ratios, three chord lengths, a square- and a round lateral-tip, presence or absence of a boundary-layer trip, and three image plane positions located opposite the wing tip. Both pressure and velocity measurements were made for the angles of attack 4 deg less than or equal to alpha less than or equal to 12 deg and for Reynolds numbers 1 x 10(exp 6) less than or equal to Re less than or equal to 3 x 10(exp 6).

Control of submersible vortex flows

NASA Technical Reports Server (NTRS)

Bushnell, D. M.; Donaldson, C. D.

1990-01-01

Vortex flows produced by submersibles typically unfavorably influence key figures of merit such as acoustic and nonacoustic stealth, control effectiveness/maneuverability, and propulsor efficiency/body drag. Sources of such organized, primarily longitudinal, vorticity include the basic body (nose and sides) and appendages (both base/intersection and tip regions) such as the fairwater, dive planes, rear control surfaces, and propulsor stators/tips. Two fundamentally different vortex control approaches are available: (1) deintensification of the amplitude and/or organization of the vortex during its initiation process; and (2) downstream vortex disablement. Vortex control techniques applicable to the initiation region (deintensification approach) include transverse pressure gradient minimization via altered body cross section, appendage dillets, fillets, and sweep, and various appendage tip and spanload treatment along with the use of active controls to minimize control surface size and motions. Vortex disablement can be accomplished either via use of control vortices (which can also be used to steer the vortices off-board), direct unwinding, inducement of vortex bursting, or segmentation/tailoring for enhanced dissipation. Submersible-applicable vortex control technology is also included derived from various aeronautical applications such as mitigation of the wing wake vortex hazard and flight aircraft maneuverability at high angle of attack as well as the status of vortex effects upon, and mitigation of, nonlinear control forces on submersibles. Specific suggestions for submersible-applicable vortex control techniques are presented.

Investigation of aerodynamic characteristics of subsonic wings

NASA Technical Reports Server (NTRS)

Dejarnette, F. R.; Frink, N. T.

1979-01-01

An analytical strake design procedure is investigated. A numerical solution to the governing strake design equation is used to generate a series of strakes which are tested in a water tunnel to study their vortex breakdown characteristics. The strakes are scaled for use on a half-scale model of the NASA-LaRC general research fuselage with a 44 degrees trapezoidal wing. An analytical solution to the governing design equation is obtained. The strake design procedure relates the potential-flow leading-edge suction and pressure distributions to vortex stability. Several suction distributions are studied and those which are more triangular and peak near the tip generate strakes that reach higher angles of attack before vortex breakdown occurs at the wing trailing edge. For the same suction distribution, a conical rather than three dimensional pressure specification results in a better strake shape as judged from its vortex breakdown characteristics.

An experimental study of the vortex wake at Mach number of 3

NASA Astrophysics Data System (ADS)

Shmakov, A. S.; Shevchenko, A. M.

2017-10-01

The results of experimental study of the flow in the wing wake at Mach number of 3 are presented. These experiments extends the data obtained in the same experimental setup at Mach numbers of 2.5 and 4 [1]. Experiments were carried out in supersonic wind tunnel T-325 of ITAM SB RAS. Rectangular half-wing with sharp edges with a chord length of 30 mm and semispan of 95 mm was used to generate vortex wake. Experimental data were obtained in two cross sections located 1.5 and 6 chord length downstream of the trailing edge at wing angle of attack of 10 degrees. Constant temperature hot-wire anemometer was used to measure disturbances in supersonic flow. Hot-wire aemometer was made of a tungsten wire with a diameter of 10 µm and length of 1.5 mm. Shlieren flow visualization were performed. As a result, the position and size of the vortex core in the wake of a rectangular wing were determined. For the first time mass flow distribution and its pulsations in the supersonic longitudinal vortex was measured at Mach number of 3.

Forebody vortex control for suppressing wing rock on a highly-swept wing configuration

NASA Technical Reports Server (NTRS)

Suarez, Carlos J.; Kramer, Brian R.; Ayers, Bert; Malcolm, Gerald N.

1992-01-01

Free-to-roll tests were conducted in a wind tunnel with a configuration that consisted of a highly-slender forebody and a 78 deg swept delta wing. A limit cycle oscillation was observed for angles of attack between 22 and 30 deg. In general, the main flow phenomena responsible for the wing-body-tail wing rock are the interactions between the forebody and the wing vortices. Various blowing techniques were evaluated as means of wing rock suppression. Blowing tangentially aft from leeward side nozzles near the forebody tip can damp the roll motion at low blowing rates and stop it completely at higher blowing rates. At the high rates, significant vortex asymmetries are created, causing the model to stop at a non-zero roll angle. Forward blowing and alternating right/left pulsed blowing appear to be more efficient techniques for suppressing wing rock. The oscillations can be damped almost completely at lower blowing coefficients, and, apparently, no major vortex asymmetries are induced. Good agreement is observed between this study and previous water tunnel tests on the same configuration.

Three-dimensional vortex wake structure of flapping wings in hovering flight.

PubMed

Cheng, Bo; Roll, Jesse; Liu, Yun; Troolin, Daniel R; Deng, Xinyan

2014-02-06

Flapping wings continuously create and send vortices into their wake, while imparting downward momentum into the surrounding fluid. However, experimental studies concerning the details of the three-dimensional vorticity distribution and evolution in the far wake are limited. In this study, the three-dimensional vortex wake structure in both the near and far field of a dynamically scaled flapping wing was investigated experimentally, using volumetric three-component velocimetry. A single wing, with shape and kinematics similar to those of a fruitfly, was examined. The overall result of the wing action is to create an integrated vortex structure consisting of a tip vortex (TV), trailing-edge shear layer (TESL) and leading-edge vortex. The TESL rolls up into a root vortex (RV) as it is shed from the wing, and together with the TV, contracts radially and stretches tangentially in the downstream wake. The downwash is distributed in an arc-shaped region enclosed by the stretched tangential vorticity of the TVs and the RVs. A closed vortex ring structure is not observed in the current study owing to the lack of well-established starting and stopping vortex structures that smoothly connect the TV and RV. An evaluation of the vorticity transport equation shows that both the TV and the RV undergo vortex stretching while convecting downwards: a three-dimensional phenomenon in rotating flows. It also confirms that convection and secondary tilting and stretching effects dominate the evolution of vorticity.

Delta wing vortex manipulation using pulsed and steady blowing during ramp pitching

NASA Technical Reports Server (NTRS)

Moreira, J.; Johari, H.

1995-01-01

The effectiveness of steady and pulsed blowing as a method of controlling delta wing vortices during ramp pitching has been investigated in flow visualization experiments conducted in a water tunnel. The recessed angled spanwise blowing technique was utilized for vortex manipulation. This technique was implemented on a beveled 60 delta wing using a pair of blowing ports located beneath the vortex core at 40% chord. The flow was injected primarily in the spanwise direction but was also composed of a component normal to the wing surface. The location of vortex burst was measured as a function of blowing intensity and pulsing frequency under static conditions, and the optimum blowing case was applied at three different wing pitching rates. Experimental results have shown that, when the burst location is upstream of the blowing port, pulsed blowing delays vortex breakdown in static and dynamic cases. Dynamic tests verified the existence of a hysteresis effect and demonstrated the improvements offered by pulsed blowing over both steady blowing and no-blowing scenarios. The application of blowing, at the optimum pulsing frequency, made the vortex breakdown location comparable in static and ramp pitch-up conditions.

Fundamental Characterization of Spanwise Loading and Trailed Wake Vortices

DTIC Science & Technology

2016-07-01

the close interaction of the tip vortex with a following blade . Such vortex interactions are fundamental determinants of rotor performance, loads, and...wing loading distribution differs from a typical loading on a hovering rotor blade in that the maximum bound circulation occurs at the blade root...and not close to the tip; this is similar to a very highly twisted rotor blade , like a tilt-rotor, in hover. The wing-vortex interaction alters the

Experimental Investigation of the Influence of a Reverse Delta Type Add-on Device on the Flap-tip Vortex of a Wing

NASA Astrophysics Data System (ADS)

Altaf, A.; Thong, T. B.; Omar, A. A.; Asrar, W.

2017-03-01

Particle Image Velocimetry was used in a low speed wind tunnel to investigate the effect of interactions of vortices produced by an outboard flap-tip of a half wing (NACA 23012 in landing configuration) and a slender reverse delta type add-on device, placed in the proximity of the outboard flap-tip, on the upper surface of the half wing. This work investigates the characteristics of the vortex interactions generated downstream in planes perpendicular to the free stream direction at a chord-based Reynolds number of Rec=2.74×105 . It was found that the add-on device significantly reduces the tangential velocity magnitude and enlarges the vortex core of the resultant vortex by up to 36.1% and 36.8%, respectively.

Sensitivity of F-106B Leading-Edge-Vortex Images to Flight and Vapor-Screen Parameters

NASA Technical Reports Server (NTRS)

Lamar, John E.; Johnson, Thomas D., Jr.

1988-01-01

A flight test was undertaken at NASA Langley Research Center with vapor-screen and image-enhancement techniques to obtain qualitative and quantitative information about near-field vortex flows above the wings of fighter aircraft. In particular, the effects of Reynolds and Mach numbers on the vortex system over an angle-of-attack range were sought. The relevance of these flows stems from their present and future use at many points in the flight envelope, especially during transonic maneuvers. The aircraft used in this flight program was the F-106B because it was available and had sufficient wing sweep (60 deg) to generate a significant leading-edge vortex system. The sensitivity of the visual results to vapor screen hardware and to onset flow changes is discussed.

Pressure-Sensitive Paint Investigation of Double-Delta Wing Vortex Flow Manipulation

NASA Technical Reports Server (NTRS)

Erickson, Gary E.; Gonzalez, Hugo A.

2004-01-01

A pressure-sensitive paint (PSP) technique was applied in a wind tunnel experiment in the NASA Langley Research Center 8-Foot Transonic Pressure Tunnel to quantify the effect of wing fillets on the global vortex-induced surface static pressure field about a sharp leading-edge 76o/40o double delta wing, or strake-wing, model at subsonic and transonic speeds. Global calibrations of the PSP were obtained at M = 0.50, 0.70, 0.85, 0.95, and 1.20, a Reynolds number per unit length of 2.0 million, and angles of attack from 10 degrees to 20 degrees using an in-situ method featuring the simultaneous acquisition of electronically-scanned pressures (ESP) at discrete locations on the model. The mean error in the PSP measurements relative to the ESP data was approximately 2 percent or less at M = 0.50 to 0.85 but increased to several percent at M =0.95 and 1.20. The PSP pressure distributions and pseudo-colored planform view pressure maps clearly revealed the vortex-induced pressure signatures at all Mach numbers and angles of attack. Small fillets having a parabolic or diamond planform situated at the strake-wing intersection were designed to manipulate the vortical flows by, respectively, removing the leading-edge discontinuity or introducing additional discontinuities. The fillets caused global changes in the vortex-dominated surface pressure field that were effectively captured in the PSP measurements. The vortex surface pressure signatures were compared to available off-surface vortex cross-flow structures obtained using a laser vapor screen (LVS) flow visualization technique. The fillet effects on the PSP pressure distributions and the observed leading-edge vortex flow characteristics were consistent with the trends in the measured lift, drag, and pitching moment coefficients.

Pressure-Sensitive Paint Investigation of Double-Delta Wing Vortex Flow Manipulation

NASA Technical Reports Server (NTRS)

Erickson, Gary E.; Gonzalez, Hugo A.

2005-01-01

A pressure-sensitive paint (PSP) technique was applied in a wind tunnel experiment in the NASA Langley Research Center 8-Foot Transonic Pressure Tunnel to quantify the effect of wing fillets on the global vortex-induced surface static pressure field about a sharp leading-edge 76 deg/40 deg double delta wing, or strake-wing, model at subsonic and transonic speeds. Global calibrations of the PSP were obtained at M = 0.50, 0.70, 0.85, 0.95, and 1.20, a Reynolds number per unit length of 2.0 million, and angles of attack from 10 degrees to 30 degrees using an in-situ method featuring the simultaneous acquisition of electronically-scanned pressures (ESP) at discrete locations on the model. The mean error in the PSP measurements relative to the ESP data was approximately 2 percent or less at M = 0.50 to 0.85 but increased to several percent at M = 0.95 and 1.20. The PSP pressure distributions and pseudo-colored planform view pressure maps clearly revealed the vortex-induced pressure signatures at all Mach numbers and angles of attack. Small fillets having a parabolic or diamond planform situated at the strake-wing intersection were designed to manipulate the vortical flows by, respectively, removing the leading-edge discontinuity or introducing additional discontinuities. The fillets caused global changes in the vortex-dominated surface pressure field that were effectively captured in the PSP measurements. The vortex surface pressure signatures were compared to available off-surface vortex cross-flow structures obtained using a laser vapor screen (LVS) flow visualization technique. The fillet effects on the PSP pressure distributions and the observed leading-edge vortex flow characteristics were consistent with the trends in the measured lift, drag, and pitching moment coefficients.

Control of Interacting Vortex Flows at Subsonic and Transonic Speeds Using Passive Porosity

NASA Technical Reports Server (NTRS)

Erickson, Gary E.

2003-01-01

A wind tunnel experiment was conducted in the NASA Langley Research Center (LaRC) 8-foot Transonic Pressure Tunnel (TPT) to determine the effects of passive surface porosity on vortex flow interactions about a general research fighter configuration at subsonic and transonic speeds. Flow- through porosity was applied to a wind leading-edge extension (LEX) mounted to a 65 deg cropped delta wind model to promote large nose-down pitching moment increments at high angles of attack. Porosity decreased the vorticity shed from the LEX, which weakened the LEX vortex and altered the global interactions of the LEX and wing vortices at high angles of attack. Six-component forces and moments and wing upper surface static pressure distributions were obtained at free- stream Mach numbers of 0.50, 0.85, and 1.20, Reynolds number of 2.5(10(exp-6) per foot, angles of attack up to 30 deg and angles of sideslip to plus or minus 8 deg. The off-surface flow field was visualized in selected cross-planes using a laser vapor screen flow visualization technique. Test data were obtained with a centerline vertical tail and with alternate twin, wing-mounted vertical fins having 0 deg and 30 deg cant angles. In addition, the porosity of the LEX was compartmentalized to determine the sensitivity of the vortex- dominated aerodynamics to the location and level of porosity applied to the LEX.

Calculation of the rotor induced download on airfoils

NASA Technical Reports Server (NTRS)

Lee, C. S.

1989-01-01

Interactions between the rotors and wing of a rotary wing aircraft in hover have a significant detrimental effect on its payload performance. The reduction of payload results from the wake of lifting rotors impinging on the wing, which is at 90 deg angle of attack in hover. This vertical drag, often referred as download, can be as large as 15 percent of the total rotor thrust in hover. The rotor wake is a three-dimensional, unsteady flow with concentrated tip vortices. With the rotor tip vortices impinging on the upper surface of the wing, the flow over the wing is not only three-dimensional and unsteady, but also separated from the leading and trailing edges. A simplified two-dimensional model was developed to demonstrate the stability of the methodology. The flow model combines a panel method to represent the rotor and the wing, and a vortex method to track the wing wake. A parametric study of the download on a 20 percent thick elliptical airfoil below a rotor disk of uniform inflow was performed. Comparisons with experimental data are made where the data are available. This approach is now being extended to three-dimensional flows. Preliminary results on a wing at 90 deg angle of attack in free stream is presented.

Coherent dynamics in the rotor tip shear layer of utility-scale wind turbines

DOE Office of Scientific and Technical Information (OSTI.GOV)

Yang, Xiaolei; Hong, Jiarong; Barone, Matthew

Here, recent field experiments conducted in the near wake (up to 0.5 rotor diameters downwind of the rotor) of a Clipper Liberty C96 2.5 MW wind turbine using snow-based super-large-scale particle image velocimetry (SLPIV) were successful in visualizing tip vortex cores as areas devoid of snowflakes. The so-visualized snow voids, however, suggested tip vortex cores of complex shape consisting of circular cores with distinct elongated comet-like tails. We employ large-eddy simulation (LES) to elucidate the structure and dynamics of the complex tip vortices identified experimentally. We show that the LES, with inflow conditions representing as closely as possible the statemore » of the flow approaching the turbine when the SLPIV experiments were carried out, reproduce vortex cores in good qualitative agreement with the SLPIV results, essentially capturing all vortex core patterns observed in the field in the tip shear layer. The computed results show that the visualized vortex patterns are formed by the tip vortices and a second set of counter-rotating spiral vortices intertwined with the tip vortices. To probe the dependence of these newly uncovered coherent flow structures on turbine design, size and approach flow conditions, we carry out LES for three additional turbines: (i) the Scaled Wind Farm Technology (SWiFT) turbine developed by Sandia National Laboratories in Lubbock, TX, USA; (ii) the wind turbine developed for the European collaborative MEXICO (Model Experiments in Controlled Conditions) project; and (iii) the model turbine, and the Clipper turbine under varying inflow turbulence conditions. We show that similar counter-rotating vortex structures as those observed for the Clipper turbine are also observed for the SWiFT, MEXICO and model wind turbines. However, the strength of the counter-rotating vortices relative to that of the tip vortices from the model turbine is significantly weaker. We also show that incoming flows with low level turbulence attenuate the elongation of the tip and counter-rotating vortices. Sufficiently high turbulence levels in the incoming flow, on the other hand, tend to break up the coherence of spiral vortices in the near wake. To elucidate the physical mechanism that gives rise to such rich coherent dynamics we examine the stability of the turbine tip shear layer using the theory. We show that for all simulated cases the theory consistently indicates the flow to be unstable exactly in the region where counter-rotating spirals emerge. We thus postulate that centrifugal instability of the rotating turbine tip shear layer is a possible mechanism for explaining the phenomena we have uncovered herein.« less

Coherent dynamics in the rotor tip shear layer of utility-scale wind turbines

DOE PAGES

Yang, Xiaolei; Hong, Jiarong; Barone, Matthew; ...

2016-09-08

Here, recent field experiments conducted in the near wake (up to 0.5 rotor diameters downwind of the rotor) of a Clipper Liberty C96 2.5 MW wind turbine using snow-based super-large-scale particle image velocimetry (SLPIV) were successful in visualizing tip vortex cores as areas devoid of snowflakes. The so-visualized snow voids, however, suggested tip vortex cores of complex shape consisting of circular cores with distinct elongated comet-like tails. We employ large-eddy simulation (LES) to elucidate the structure and dynamics of the complex tip vortices identified experimentally. We show that the LES, with inflow conditions representing as closely as possible the statemore » of the flow approaching the turbine when the SLPIV experiments were carried out, reproduce vortex cores in good qualitative agreement with the SLPIV results, essentially capturing all vortex core patterns observed in the field in the tip shear layer. The computed results show that the visualized vortex patterns are formed by the tip vortices and a second set of counter-rotating spiral vortices intertwined with the tip vortices. To probe the dependence of these newly uncovered coherent flow structures on turbine design, size and approach flow conditions, we carry out LES for three additional turbines: (i) the Scaled Wind Farm Technology (SWiFT) turbine developed by Sandia National Laboratories in Lubbock, TX, USA; (ii) the wind turbine developed for the European collaborative MEXICO (Model Experiments in Controlled Conditions) project; and (iii) the model turbine, and the Clipper turbine under varying inflow turbulence conditions. We show that similar counter-rotating vortex structures as those observed for the Clipper turbine are also observed for the SWiFT, MEXICO and model wind turbines. However, the strength of the counter-rotating vortices relative to that of the tip vortices from the model turbine is significantly weaker. We also show that incoming flows with low level turbulence attenuate the elongation of the tip and counter-rotating vortices. Sufficiently high turbulence levels in the incoming flow, on the other hand, tend to break up the coherence of spiral vortices in the near wake. To elucidate the physical mechanism that gives rise to such rich coherent dynamics we examine the stability of the turbine tip shear layer using the theory. We show that for all simulated cases the theory consistently indicates the flow to be unstable exactly in the region where counter-rotating spirals emerge. We thus postulate that centrifugal instability of the rotating turbine tip shear layer is a possible mechanism for explaining the phenomena we have uncovered herein.« less

A Computational and Experimental Investigation of a Delta Wing with Vertical Tails

NASA Technical Reports Server (NTRS)

Krist. Sherrie L.; Washburn, Anthony E.; Visser, Kenneth D.

2004-01-01

The flow over an aspect ratio 1 delta wing with twin vertical tails is studied in a combined computational and experimental investigation. This research is conducted in an effort to understand the vortex and fin interaction process. The computational algorithm used solves both the thin-layer Navier-Stokes and the inviscid Euler equations and utilizes a chimera grid-overlapping technique. The results are compared with data obtained from a detailed experimental investigation. The laminar case presented is for an angle of attack of 20 and a Reynolds number of 500; 000. Good agreement is observed for the physics of the flow field, as evidenced by comparisons of computational pressure contours with experimental flow-visualization images, as well as by comparisons of vortex-core trajectories. While comparisons of the vorticity magnitudes indicate that the computations underpredict the magnitude in the wing primary-vortex-core region, grid embedding improves the computational prediction.

A computational and experimental investigation of a delta wing with vertical tails

NASA Technical Reports Server (NTRS)

Krist, Sherrie L.; Washburn, Anthony E.; Visser, Kenneth D.

1993-01-01

The flow over an aspect ratio 1 delta wing with twin vertical tails is studied in a combined computational and experimental investigation. This research is conducted in an effort to understand the vortex and fin interaction process. The computational algorithm used solves both the thin-layer Navier-Stokes and the inviscid Euler equations and utilizes a chimera grid-overlapping technique. The results are compared with data obtained from a detailed experimental investigation. The laminar case presented is for an angle of attack of 20 deg and a Reynolds number of 500,000. Good agreement is observed for the physics of the flow field, as evidenced by comparisons of computational pressure contours with experimental flow-visualization images, as well as by comparisons of vortex-core trajectories. While comparisons of the vorticity magnitudes indicate that the computations underpredict the magnitude in the wing primary-vortex-core region, grid embedding improves the computational prediction.

Effects of boundary layer forcing on wing-tip vortices

NASA Astrophysics Data System (ADS)

Shaw-Ward, Samantha

The nature of turbulence within wing-tip vortices has been a topic of research for decades, yet accurate measurements of Reynolds stresses within the core are inherently difficult due to the bulk motion wandering caused by initial and boundary conditions in wind tunnels. As a result, characterization of a vortex as laminar or turbulent is inconclusive and highly contradicting. This research uses several experimental techniques to study the effects of broadband turbulence, introduced within the wing boundary layer, on the development of wing-tip vortices. Two rectangular wings with a NACA 0012 profile were fabricated for the use of this research. One wing had a smooth finish and the other rough, introduced by P80 grade sandpaper. Force balance measurements showed a small reduction in wing performance due to surface roughness for both 2D and 3D configurations, although stall characteristics remained relatively unchanged. Seven-hole probes were purpose-built and used to assess the mean velocity profiles of the vortices five chord lengths downstream of the wing at multiple angles of attack. Above an incidence of 4 degrees, the vortices were nearly axisymmetric, and the wing roughness reduced both velocity gradients and peak velocity magnitudes within the vortex. Laser Doppler velocimetry was used to further assess the time-resolved vortex at an incidence of 5 degrees. Evidence of wake shedding frequencies and wing shear layer instabilities at higher frequencies were seen in power spectra within the vortex. Unlike the introduction of freestream turbulence, wing surface roughness did not appear to increase wandering amplitude. A new method for removing the effects of vortex wandering is proposed with the use of carefully selected high-pass filters. The filtered data revealed that the Reynolds stress profiles of the vortex produced by the smooth and rough wing were similar in shape, with a peak occurring away from the vortex centre but inside of the core. Single hot-wire measurements in the 2D wing wake revealed the potential origin of dominant length-scales observed in the vortex power spectra. At angles above 5 degrees, the 2D wing wake had both higher velocity deficits and higher levels of total wake kinetic energy for the rough wing as compared to the smooth wing.

The application of laser Doppler velocimetry to trailing vortex definition and alleviation

NASA Technical Reports Server (NTRS)

Orloff, K. L.; Grant, G. R.

1973-01-01

A laser Doppler velocimeter whose focal volume can be rapidly traversed through a flowfield has been used to overcome the problem introduced by excursions of the central vortex filament within a wind tunnel test section. The basic concepts of operation of the instrument are reviewed and data are presented which accurately define the trailing vortex from a square-tipped rectangular wing. Measured axial and tangential velocity distributions are given, both with and without a vortex dissipator panel installed at the wing tip. From the experimental data, circulation and vorticity distributions are obtained and the effect of turbulence injection into the vortex structure is discussed.

A PIV Study of Baseline and Controlled Flow over the Highly Deflected Flap of a Generic Low Aspect Ratio Trapezoidal Wing

NASA Astrophysics Data System (ADS)

Tewes, Philipp; Genschow, Konstantin; Little, Jesse; Wygnanski, Israel

2017-11-01

A detailed flow survey using PIV was conducted over a highly-deflected flap (55°) of a low-aspect ratio trapezoidal wing. The wing section is a NACA 0012 with 45° sweep at both the leading and trailing edges, an aspect ratio of 1.5 and a taper ratio of 0.27. The main element is equipped with 7 equally spaced fluidic oscillators, covering the inner 60 % of the span, located near the flap hinge. Experiments were carried out at 0° and 8° incidence at a Reynolds number of 1.7 .106 for both baseline and active flow control (AFC) cases. Velocity ISO-surfaces, x-vorticity and streamlines are analyzed / discussed. A flap leading edge vortex governs the baseline flow field for 0°. This vortical structure interacts with the jets emitted by the actuators (Cμ = 1 %). Its development is hampered and the vortex is redirected toward the trailing edge resulting in a CL increase. At 8°, the dominant flap leading edge vortex could not be detected and is believed to have already merged with the tip vortex. AFC attached the flow over the flap and enhanced the lift by up to 20 % while maintaining longitudinal stability. The dominant flow features in the AFC cases are actuator-generated streamwise vortices which appear stronger at 8°. This work was supported by the Office of Naval Research under ONR Grant No. N00014-14-1-0387.

Effects of forebody strakes and Mach number on overall aerodynamic characteristics of configuration with 55 deg cropped delta wing

NASA Technical Reports Server (NTRS)

Erickson, Gary E.; Rogers, Lawrence W.

1992-01-01

A wind tunnel data base was established for the effects of chine-like forebody strakes and Mach number on the longitudinal and lateral-directional characteristics of a generalized 55 degree cropped delta wing-fuselage-centerline vertical tail configuration. The testing was conducted in the 7- by 10-Foot Transonic Tunnel at the David Taylor Research Center at free-stream Mach numbers of 0.40 to 1.10 and Reynolds numbers based on the wing mean aerodynamic chord of 1.60 x 10(exp 6) to 2.59 x 10(exp 6). The best matrix included angles of attack from 0 degree to a maximum of 28 degree, angles of sidesip of 0, +5, and -5 degrees, and wing leading-edge flat deflection angles of 0 and 30 degrees. Key flow phenomena at subsonic and transonic conditions were identified by measuring off-body flow visualization with a laser screen technique. These phenomena included coexisting and interacting vortex flows and shock waves, vortex breakdown, vortex flow interactions with the vertical tail, and vortices induced by flow separation from the hinge line of the deflected wing flap. The flow mechanisms were correlated with the longitudinal and lateral-directional aerodynamic data trends.

Local measurement and numerical modeling of mass/heat transfer from a turbine blade in a linear cascade with tip clearance

NASA Astrophysics Data System (ADS)

Jin, Peitong

2000-11-01

Local mass/heat transfer measurements from the turbine blade near-tip and the tip surfaces are performed using the naphthalene sublimation technique. The experiments are conducted in a linear cascade consisting of five high-pressure blades with a central test-blade configuration. The incoming flow conditions are close to those of the gas turbine engine environment (boundary layer displacement thickness is about 0.01 of chord) with an exit Reynolds number of 6.2 x 105. The effects of tip clearance level (0.86%--6.90% of chord), mainstream Reynolds number and turbulence intensity (0.2 and 12.0%) are investigated. Two methods of flow visualization---oil and lampblack, laser light sheet smoke wire---as well as static pressure measurement on the blade surface are used to study the tip leakage flow and vortex in the cascade. In addition, numerical modeling of the flow and heat transfer processes in the linear cascade with different tip clearances is conducted using commercial software incorporating advanced turbulence models. The present study confirms many important results on the tip leakage flow and vortex from the literature, contributes to the current understanding in the effects of tip leakage flow and vortex on local heat transfer from the blade near-tip and the tip surfaces, and provides detailed local and average heat/mass transfer data applicable to turbine blade tip cooling design.

The Aerodynamics of Deforming Wings at Low Reynolds Number

NASA Astrophysics Data System (ADS)

Medina, Albert

Flapping flight has gained much attention in the past decade driven by the desire to understand capabilities observed in nature and the desire to develop agile small-scale aerial vehicles. Advancing our current understanding of unsteady aerodynamics is an essential component in the development of micro-air vehicles (MAV) intended to utilize flight mechanics akin to insect flight. Thus the efforts undertaken that of bio-mimicry. The complexities of insect wing motion are dissected and simplified to more tractable problems to elucidate the fundamentals of unsteady aerodynamics in biologically inspired kinematics. The MAV's fruition would satisfy long established needs in both the military and civilian sectors. Although recent studies have provided great insight into the lift generating mechanisms of flapping wings the deflection response of such wings remains poorly understood. This dissertation numerically and experimentally investigates the aerodynamic performance of passively and actively deflected wings in hover and rotary kinematics. Flexibility is distilled to discrete lines of flexion which acknowledging major flexion lines in insect wings to be the primary avenue for deformation. Of primary concern is the development of the leading-edge vortex (LEV), a high circulation region of low pressure above the wing to which much of the wing's lift generation is attributed. Two-dimensional simulations of wings with chord-wise flexibility in a freestream reveal a lift generating mechanism unavailable to rigid wings with origins in vortical symmetry breaking. The inclusion of flexibility in translating wings accelerated from rest revealed the formation time of the initial LEV was very weakly dependent on the flexible stiffness of the wing, maintaining a universal time scale of four to five chords of travel before shedding. The frequency of oscillatory shedding of the leading and trailing-edge vortices that develops after the initial vortex shedding was shown to be responsive to flexibility satisfying an inverse proportionality to stiffness. In hover, an effective pitch angle can be defined in a flexible wing that accounts for deflection which shifts results toward trend lines of rigid wings. Three-dimensional simulations examining the effects of two distinct deformation modes undergoing prescribed deformation associated with root and tip deflection demonstrated a greater aerodynamic response to tip deflection in hover. Efficiency gains in flexion wings over rigid wing counterpart were shown to be dependent on Reynolds number with efficiency in both modes increasing with increased Reynolds number. Additionally, while the leading-edge vortex axis proved insensitive to deformation, the shape and orientation of the LEV core is modified. Experiments on three-dimensional dynamically-scaled fruit fly wings with passive deformation operating in the bursting limit Reynolds number regime revealed enhanced leading-edge vortex bursting with tip deflection promoting greater LEV core flow deceleration in stroke. Experimental studies on rotary wings highlights a universal formation time of the leading-edge vortex independent of Reynolds number, acceleration profile and aspect ratio. Efforts to replicate LEV bursting phenomena of higher aspect ratio wings in a unity aspect ratio wing such that LEV growth is no limited by span but by the LEV traversing the chord revealed a flow regime of oscillatory lift generation reminiscent of behavior exhibited in translating wings that also maintains magnitude peak to peak.

Effects of Double-Leakage Tip Clearance Flow on the Performance of a Compressor Stage with a Large Rotor Tip Gap

NASA Technical Reports Server (NTRS)

Hah, Chunill

2016-01-01

Effects of a large rotor tip gap on the performance of a one and half stage axial compressor are investigated in detail with a numerical simulation based on LES and available PIV data. The current paper studies the main flow physics, including why and how the loss generation is increased with the large rotor tip gap. The present study reveals that when the tip gap becomes large, tip clearance fluid goes over the tip clearance core vortex and enters into the next blade's tip gap, which is called double-leakage tip clearance flow. As the tip clearance flow enters into the adjacent blade's tip gap, a vortex rope with a lower pressure core is generated. This vortex rope breaks up the tip clearance core vortex of the adjacent blade, resulting in a large additional mixing. This double-leakage tip clearance flow occurs at all operating conditions, from design flow to near stall condition, with the large tip gap for the current compressor stage. The double-leakage tip clearance flow, its interaction with the tip clearance core vortex of the adjacent blade, and the resulting large mixing loss are the main flow mechanism of the large rotor tip gap in the compressor. When the tip clearance is smaller, flow near the end wall follows more closely with the main passage flow and this double-leakage tip clearance flow does not happen near the design flow condition for the current compressor stage. When the compressor with a large tip gap operates at near stall operation, a strong vortex rope is generated near the leading edge due to the double-leakage flow. Part of this vortex separates from the path of the tip clearance core vortex and travels from the suction side of the blade toward the pressure side of the blade. This vortex is generated periodically at near stall operation with a large tip gap. As the vortex travels from the suction side to the pressure side of the blade, a large fluctuation of local pressure forces blade vibration. Nonsynchronous blade vibration occurs due to this vortex as the frequency of this vortex generation is not the same as the rotor. The present investigation confirms that this vortex is a part of separated tip clearance vortex, which is caused by the double-leakage tip clearance flow.

A Rotor Tip Vortex Tracing Algorithm for Image Post-Processing

NASA Technical Reports Server (NTRS)

Overmeyer, Austin D.

2015-01-01

A neurite tracing algorithm, originally developed for medical image processing, was used to trace the location of the rotor tip vortex in density gradient flow visualization images. The tracing algorithm was applied to several representative test images to form case studies. The accuracy of the tracing algorithm was compared to two current methods including a manual point and click method and a cross-correlation template method. It is shown that the neurite tracing algorithm can reduce the post-processing time to trace the vortex by a factor of 10 to 15 without compromising the accuracy of the tip vortex location compared to other methods presented in literature.

Numerical study of the trailing vortex of a wing with wing-tip blowing

NASA Technical Reports Server (NTRS)

Lim, Hock-Bin

1994-01-01

Trailing vortices generated by lifting surfaces such as helicopter rotor blades, ship propellers, fixed wings, and canard control surfaces are known to be the source of noise, vibration, cavitation, degradation of performance, and other hazardous problems. Controlling these vortices is, therefore, of practical interest. The formation and behavior of the trailing vortices are studied in the present research. In addition, wing-tip blowing concepts employing axial blowing and spanwise blowing are studied to determine their effectiveness in controlling these vortices and their effects on the performance of the wing. The 3D, unsteady, thin-layer compressible Navier-Stokes equations are solved using a time-accurate, implicit, finite difference scheme that employs LU-ADI factorization. The wing-tip blowing is simulated using the actuator plane concept, thereby, not requiring resolution of the jet slot geometry. Furthermore, the solution blanking feature of the chimera scheme is used to simplify the parametric study procedure for the wing-tip blowing. Computed results are shown to compare favorably with experimental measurements. It is found that axial wing-tip blowing, although delaying the rolling-up of the trailing vortices and the near-field behavior of the flowfield, does not dissipate the circulation strength of the trailing vortex farther downstream. Spanwise wing-tip blowing has the effect of displacing the trailing vortices outboard and upward. The increased 'wing-span' due to the spanwise wing-tip blowing has the effect of lift augmentation on the wing and the strengthening of the trailing vortices. Secondary trailing vortices are created at high spanwise wing-tip blowing intensities.

Petiolate wings: effects on the leading-edge vortex in flapping flight.

PubMed

Phillips, Nathan; Knowles, Kevin; Bomphrey, Richard J

2017-02-06

The wings of many insect species including crane flies and damselflies are petiolate (on stalks), with the wing planform beginning some distance away from the wing hinge, rather than at the hinge. The aerodynamic impact of flapping petiolate wings is relatively unknown, particularly on the formation of the lift-augmenting leading-edge vortex (LEV): a key flow structure exploited by many insects, birds and bats to enhance their lift coefficient. We investigated the aerodynamic implications of petiolation P using particle image velocimetry flow field measurements on an array of rectangular wings of aspect ratio 3 and petiolation values of P = 1-3. The wings were driven using a mechanical device, the 'Flapperatus', to produce highly repeatable insect-like kinematics. The wings maintained a constant Reynolds number of 1400 and dimensionless stroke amplitude Λ * (number of chords traversed by the wingtip) of 6.5 across all test cases. Our results showed that for more petiolate wings the LEV is generally larger, stronger in circulation, and covers a greater area of the wing surface, particularly at the mid-span and inboard locations early in the wing stroke cycle. In each case, the LEV was initially arch-like in form with its outboard end terminating in a focus-sink on the wing surface, before transitioning to become continuous with the tip vortex thereafter. In the second half of the wing stroke, more petiolate wings exhibit a more detached LEV, with detachment initiating at approximately 70% and 50% span for P = 1 and 3, respectively. As a consequence, lift coefficients based on the LEV are higher in the first half of the wing stroke for petiolate wings, but more comparable in the second half. Time-averaged LEV lift coefficients show a general rise with petiolation over the range tested.

Petiolate wings: effects on the leading-edge vortex in flapping flight

PubMed Central

2017-01-01

The wings of many insect species including crane flies and damselflies are petiolate (on stalks), with the wing planform beginning some distance away from the wing hinge, rather than at the hinge. The aerodynamic impact of flapping petiolate wings is relatively unknown, particularly on the formation of the lift-augmenting leading-edge vortex (LEV): a key flow structure exploited by many insects, birds and bats to enhance their lift coefficient. We investigated the aerodynamic implications of petiolation P using particle image velocimetry flow field measurements on an array of rectangular wings of aspect ratio 3 and petiolation values of P = 1–3. The wings were driven using a mechanical device, the ‘Flapperatus’, to produce highly repeatable insect-like kinematics. The wings maintained a constant Reynolds number of 1400 and dimensionless stroke amplitude Λ* (number of chords traversed by the wingtip) of 6.5 across all test cases. Our results showed that for more petiolate wings the LEV is generally larger, stronger in circulation, and covers a greater area of the wing surface, particularly at the mid-span and inboard locations early in the wing stroke cycle. In each case, the LEV was initially arch-like in form with its outboard end terminating in a focus-sink on the wing surface, before transitioning to become continuous with the tip vortex thereafter. In the second half of the wing stroke, more petiolate wings exhibit a more detached LEV, with detachment initiating at approximately 70% and 50% span for P = 1 and 3, respectively. As a consequence, lift coefficients based on the LEV are higher in the first half of the wing stroke for petiolate wings, but more comparable in the second half. Time-averaged LEV lift coefficients show a general rise with petiolation over the range tested. PMID:28163876

A water tunnel flow visualization study of the vortex flow structures on the F/A-18 aircraft

NASA Technical Reports Server (NTRS)

Sandlin, Doral R.; Ramirez, Edgar J.

1991-01-01

The vortex flow structures occurring on the F/A-18 aircraft at high angles of attack were studied. A water tunnel was used to gather flow visualization data on the forebody vortex and the wing leading edge extension vortex. The longitudinal location of breakdown of the leading edge vortex was found to be consistently dependent on the angle of attack. Other parameters such as Reynolds number, model scale, and model fidelity had little influence on the overall behavior of the flow structures studied. The lateral location of the forebody vortex system was greatly influenced by changes in the angle of sideslip. Strong interactions can occur between the leading edge extension vortex and the forebody vortex. Close attention was paid to vortex induced flows on various airframe components of the F/A-18. Reynolds number and angle of attack greatly affected the swirling intensity, and therefore the strength of the studied vortices. Water tunnel results on the F/A-18 correlated well with those obtained in similar studies at both full and sub scale levels. The water tunnel can provide, under certain conditions, good simulations of realistic flows in full scale configurations.

Experimental Study of Tip Vortex Flow from a Periodically Pitched Airfoil Section

NASA Technical Reports Server (NTRS)

Zaman, Khairul; Fagan, Amy; Mankbadi, Mina

2016-01-01

An experimental investigation of tip vortex flow from a NACA0012 airfoil, pitched periodically at various frequencies, is conducted in a low-speed wind tunnel. Initially, data for stationary airfoil held fixed at various angles-of-attack are gathered. Flow visualization pictures as well as detailed cross-sectional properties areobtained at various streamwise locations using hot-wire anemometry. Data include mean velocity, streamwise vorticity as well as various turbulent stresses. Preliminary data are also acquired for periodically pitched airfoil. These results are briefly presented in this extended abstract.

Aerodynamic performance of a wing with a deflected tip-mounted reverse half-delta wing

NASA Astrophysics Data System (ADS)

Lee, T.; Su, Y. Y.

2012-11-01

The impact of a tip-mounted 65°-sweep reverse half-delta wing (RHDW), set at different deflections, on the aerodynamic performance of a rectangular NACA 0012 wing was investigated experimentally at Re = 2.45 × 105. This study is a continuation of the work of Lee and Su (Exp Fluids 52(6):1593-1609, 2012) on the passive control of wing tip vortex by the use of a reverse half-delta wing. The present results show that for RHDW deflection with -5° ≤ δ ≤ +15°, the lift was found to increase nonlinearly with increasing δ compared to the baseline wing. The lift increment was accompanied by an increased total drag. For negative RHDW deflection with δ < -5°, the RHDW-induced lift decrement was, however, accompanied by an improved drag. The deflected RHDW also significantly modified and weakened the tip vortex, leading to a persistently lowered lift-induced drag, regardless of its deflection, compared to the baseline wing. Physical mechanisms responsible for the observed RHDW-induced phenomenon were also discussed.

Tip Vortices of Isolated Wings and Helicopter Rotor Blades.

DTIC Science & Technology

1987-12-01

root to tip, as expected due to the induced downwash of the tip vor- tex and wake vortex sheet. Although the three different tip-caps produce very...the inherent limitation of not being able to model the vortex wake with these equations, although the Euler formulation has in it the necessary...physics to model vorticity transport correctly. These equations basically lack the physical mecha- nism needed to generate the vortex wake . However, in

Experimental examination of vorticity stripping from a wing-tip vortex in free-stream turbulence

NASA Astrophysics Data System (ADS)

Ghimire, Hari C.; Bailey, Sean C. C.

2018-03-01

Time-resolved stereoscopic particle image velocimetry measurements were conducted of a wing-tip vortex decaying in free-stream turbulence. The objective of the research was to experimentally investigate the mechanism causing the increased rate of decay of the vortex in the presence of turbulence. It was observed that the circulation of the vortex core experienced periods of rapid loss and recovery when immersed in free-stream turbulence. These events were not observed when the vortex was in a laminar free stream. A connection was made between these events and distortion of the vortex, coinciding with stripping of core fluid from the vortex core. Specifically, vortex stripping events were connected to asymmetry in the vortex core, and this asymmetry was associated with instances of rapid circulation loss. The increased rate of decay of the vortex in turbulence coincided with the formation of secondary vortical structures which wrapped azimuthally around the primary vortex.

On the nonlinear aerodynamic and stability characteristics of a generic chine-forebody slender-wing fighter configuration

NASA Technical Reports Server (NTRS)

Erickson, Gary E.; Brandon, Jay M.

1987-01-01

An exploratory investigation was conducted of the nonlinear aerodynamic and stability characteristics of a tailless generic fighter configuration featuring a chine-shaped forebody coupled to a slender cropped delta wing in the NASA Langley Research Center's 12-Foot Low-Speed Wind Tunnel. Forebody and wing vortex flow mechanisms were identified through off-body flow visualizations to explain the trends in the longitudinal and lateral-directional characteristics at extreme attitudes (angles of attack and sideslip). The interactions of the vortical motions with centerline and wing-mounted vertical tail surfaces were studied and the flow phenomena were correlated with the configuration forces and moments. Single degree of freedom, free-to-roll tests were used to study the wing rock susceptibility of the generic fighter model. Modifications to the nose region of the chine forebody were examined and fluid mechanisms were established to account for their ineffectiveness in modulating the highly interactive forebody and wing vortex systems.

Parametric analysis of swept-wing geometry with sheared wing tips

NASA Technical Reports Server (NTRS)

Fremaux, C. M.; Vijgen, P. M. H. W.; Van Dam, C. P.

1990-01-01

A computational parameter study is presented of potential reductions in induced drag and increases in lateral-directional stability due to sheared wing tips attached to an untwisted wing of moderate sweep and aspect ratio. Sheared tips are swept and tapered wing-tip devices mounted in the plane of the wing. The induced-drag results are obtained using an inviscid, incompressible surface-panel method that models the nonlinear effects due to the deflected and rolled-up wake behind the lifting surface. The induced-drag results with planar sheared tips are compared to straight-tapered tip extensions and nonplanar winglet geometries. The lateral-directional static-stability characteristics of the wing with sheared tips are estimated using a quasi-vortex-lattice method. For certain combinations of sheared-tip sweep and taper, both the induced efficiency of the wing and the relevant static-stability derivatives are predicted to increase compared to the wing with a straight-tapered tip modification.

The effect of aspect ratio on the leading-edge vortex over an insect-like flapping wing.

PubMed

Phillips, Nathan; Knowles, Kevin; Bomphrey, Richard J

2015-10-09

Insect wing shapes are diverse and a renowned source of inspiration for the new generation of autonomous flapping vehicles, yet the aerodynamic consequences of varying geometry is not well understood. One of the most defining and aerodynamically significant measures of wing shape is the aspect ratio, defined as the ratio of wing length (R) to mean wing chord (c). We investigated the impact of aspect ratio, AR, on the induced flow field around a flapping wing using a robotic device. Rigid rectangular wings ranging from AR = 1.5 to 7.5 were flapped with insect-like kinematics in air with a constant Reynolds number (Re) of 1400, and a dimensionless stroke amplitude of 6.5c (number of chords traversed by the wingtip). Pseudo-volumetric, ensemble-averaged, flow fields around the wings were captured using particle image velocimetry at 11 instances throughout simulated downstrokes. Results confirmed the presence of a high-lift, separated flow field with a leading-edge vortex (LEV), and revealed that the conical, primary LEV grows in size and strength with increasing AR. In each case, the LEV had an arch-shaped axis with its outboard end originating from a focus-sink singularity on the wing surface near the tip. LEV detachment was observed for AR > 1.5 around mid-stroke at ~70% span, and initiated sooner over higher aspect ratio wings. At AR > 3 the larger, stronger vortex persisted under the wing surface well into the next half-stroke leading to a reduction in lift. Circulatory lift attributable to the LEV increased with AR up to AR = 6. Higher aspect ratios generated proportionally less lift distally because of LEV breakdown, and also less lift closer to the wing root due to the previous LEV's continuing presence under the wing. In nature, insect wings go no higher than AR ~ 5, likely in part due to architectural and physiological constraints but also because of the reducing aerodynamic benefits of high AR wings.

Wind tunnel investigation of vortex flows on F/A-18 configuration at subsonic through transonic speed

NASA Technical Reports Server (NTRS)

Erickson, Gary E.

1991-01-01

A wind tunnel experiment was conducted in the David Taylor Research Center 7- by 10-Foot Transonic Tunnel of the wing leading-edge extension (LEX) and forebody vortex flows at subsonic and transonic speeds about a 0.06-scale model of the F/A-18. The primary goal was to improve the understanding and control of the vortical flows, including the phenomena of vortex breakdown and vortex interactions with the vertical tails. Laser vapor screen flow visualizations, LEX, and forebody surface static pressures, and six-component forces and moments were obtained at angles of attack of 10 to 50 degrees, free-stream Mach numbers of 0.20 to 0.90, and Reynolds numbers based on the wing mean aerodynamic chord of 0.96 x 10(exp 6) to 1.75 x 10(exp 6). The wind tunnel results were correlated with in-flight flow visualizations and handling qualities trends obtained by NASA using an F-18 High-Alpha Research Vehicle (HARV) and by the Navy and McDonnell Douglas on F-18 aircraft with LEX fences added to improve the vertical tail buffet environment. Key issues that were addressed include the sensitivity of the vortical flows to the Reynolds number and Mach number; the reduced vertical tail excitation, and the corresponding flow mechanism, in the presence of the LEX fence; the repeatability of data obtained during high angle-of-attack wind tunnel testing of F-18 models; the effects of particle seeding for flow visualization on the quantitative model measurements; and the interpretation of off-body flow visualizations obtained using different illumination and particle seeding techniques.

Cessna-172R Airplane in Cruise and Landing Configurations: A Numerical Study of the Wing Loads and Wake

NASA Astrophysics Data System (ADS)

Jha, Pankaj

2013-11-01

The present work deals with the analysis of flight test data on a Cessna 172R airplane near University Park airport in Pennsylvania. Several tests pertaining to rate-of-climb, cruise, stall and landing were performed. Those of aerodynamic nature will be discussed. The wing loads for the cruise as well as landing configurations with various flap angles were computed using a vortex method considering horse-shoe and bound vortices. The stall speed and maximum lift coefficient of the airplane for these flap settings at a particular altitude were determined. The comparison against the processed flight data was generally very good. A detailed study will be presented. A CFD approach inspired by the author's work (Jha et al., 2013) to model wind turbine blades and wakes and classical aerodynamics problems was taken to model the airplane wings. The simulation results were also compared against the flight data. In addition, these simulations facilitated visualization and analysis of flow features of interest, like wing tip trailing vortices and their turbulence characterization. Graduate Research Assistant, Aerospace Engineering.

A Discrete-Vortex Method for Studying the Wing Rock of Delta Wings

NASA Technical Reports Server (NTRS)

Gainer, Thomas G.

2002-01-01

A discrete-vortex method is developed to investigate the wing rock problem associated with highly swept wings. The method uses two logarithmic vortices placed above the wing to represent the vortex flow field and uses boundary conditions based on conical flow, vortex rate of change of momentum, and other considerations to position the vortices and determine their strengths. A relationship based on the time analogy and conical-flow assumptions is used to determine the hysteretic positions of the vortices during roll oscillations. Static and dynamic vortex positions and wing rock amplitudes and frequencies calculated by using the method are generally in good agreement with available experimental data. The results verify that wing rock is caused by hysteretic deflections of the vortices and indicate that the stabilizing moments that limit wing rock amplitudes are the result of the one primary vortex moving outboard of the wing where it has little influence on the wing.

B-747 in Flight during Vortex Study with Learjet and T-37 Fly Through the Wake

NASA Technical Reports Server (NTRS)

1974-01-01

In this 1974 NASA Flight Research Center (FRC) photograph, the two chase aircraft, a Learjet and a Cessna T-37, are shown in formation off the right wing tip of the Boeing B-747 jetliner. The two chase aircraft were used to probe the trailing wake vortices generated by the airflow around the wings of the B-747 aircraft. The vortex trail behind the right wing tip was made visible by a smoke generator mounted under the wing of the B-747 aircraft. In 1974 the NASA Flight Research Center (later Dryden Flight Research Center, Edwards, California) used a Boeing 747 as part of the overall NASA study of trailing vortices. Trailing vortices are the invisible flow of spiraling air that trails from the wings of large aircraft and can 'upset' smaller aircraft flying behind them. The 747 that NASA used was on loan from the Johnson Space Center where it was part of the Space Shuttle Program. The data gathered in the 747 studies complemented data from the previous (1973-74) joint NASA Flight Research Center and Federal Aviation Administration (FAA) Boeing727 wake vortices study. Six smoke generators were installed under the wings of the 747 to provide a visual image of the trailing vortices. The object of the experiments was to test different configurations and mechanical devices on the747 that could be used to break up or lessen the strength of the vortices. The results of the tests could lead to shorter spacing between landings and takeoffs, which, in turn, could alleviate air-traffic congestion. For approximately 30 flights the 747 was flown using various combinations of wing air spoilers in an attempt to reduce wake vortices. To evaluate the effectiveness of the different configurations, chase aircraft were flown into the vortex sheets to probe their strengths and patterns at different times. Two of the chase planes used were the Flight Research Center's Cessna T-37 and the NASA Ames Research Center's Learjet. These aircraft represented the types of smaller business jets and other small aircraft that might encounter large passenger aircraft on approach or landings around major airports or in flight. Tests without the 747's wing spoilers deployed produced violent 'upset' problems for the T-37 aircraft at a distance of approximately 3 miles. From the magnitude of the problems found, distances of as much as ten miles might be required if spoilers were not used. With two spoilers on the outer wing panels, the T-37 could fly at a distance of three miles and not experience the 'upset' problem. The wake vortex study continued even after the 747 was returned to its primary mission of carrying the Space Shuttle.

Power reduction and the radial limit of stall delay in revolving wings of different aspect ratio

PubMed Central

Kruyt, Jan W.; van Heijst, GertJan F.; Altshuler, Douglas L.; Lentink, David

2015-01-01

Airplanes and helicopters use high aspect ratio wings to reduce the power required to fly, but must operate at low angle of attack to prevent flow separation and stall. Animals capable of slow sustained flight, such as hummingbirds, have low aspect ratio wings and flap their wings at high angle of attack without stalling. Instead, they generate an attached vortex along the leading edge of the wing that elevates lift. Previous studies have demonstrated that this vortex and high lift can be reproduced by revolving the animal wing at the same angle of attack. How do flapping and revolving animal wings delay stall and reduce power? It has been hypothesized that stall delay derives from having a short radial distance between the shoulder joint and wing tip, measured in chord lengths. This non-dimensional measure of wing length represents the relative magnitude of inertial forces versus rotational accelerations operating in the boundary layer of revolving and flapping wings. Here we show for a suite of aspect ratios, which represent both animal and aircraft wings, that the attachment of the leading edge vortex on a revolving wing is determined by wing aspect ratio, defined with respect to the centre of revolution. At high angle of attack, the vortex remains attached when the local radius is shorter than four chord lengths and separates outboard on higher aspect ratio wings. This radial stall limit explains why revolving high aspect ratio wings (of helicopters) require less power compared with low aspect ratio wings (of hummingbirds) at low angle of attack and vice versa at high angle of attack. PMID:25788539

Ear-body lift and a novel thrust generating mechanism revealed by the complex wake of brown long-eared bats (Plecotus auritus)

NASA Astrophysics Data System (ADS)

Johansson, L. Christoffer; Håkansson, Jonas; Jakobsen, Lasse; Hedenström, Anders

2016-04-01

Large ears enhance perception of echolocation and prey generated sounds in bats. However, external ears likely impair aerodynamic performance of bats compared to birds. But large ears may generate lift on their own, mitigating the negative effects. We studied flying brown long-eared bats, using high resolution, time resolved particle image velocimetry, to determine the aerodynamics of flying with large ears. We show that the ears and body generate lift at medium to cruising speeds (3-5 m/s), but at the cost of an interaction with the wing root vortices, likely reducing inner wing performance. We also propose that the bats use a novel wing pitch mechanism at the end of the upstroke generating thrust at low speeds, which should provide effective pitch and yaw control. In addition, the wing tip vortices show a distinct spiraling pattern. The tip vortex of the previous wingbeat remains into the next wingbeat and rotates together with a newly formed tip vortex. Several smaller vortices, related to changes in circulation around the wing also spiral the tip vortex. Our results thus show a new level of complexity in bat wakes and suggest large eared bats are less aerodynamically limited than previous wake studies have suggested.

Unified Application of Vapor Screen Flow Visualization and Pressure Sensitive Paint Measurement Techniques to Vortex- and Shock Wave-Dominated Flow Fields

NASA Technical Reports Server (NTRS)

Erickson, Gary E.

2010-01-01

Laser vapor screen (LVS) flow visualization and pressure sensitive paint (PSP) techniques were applied in a unified approach to wind tunnel testing of slender wing and missile configurations dominated by vortex flows and shock waves at subsonic, transonic, and supersonic speeds. The off-surface cross-flow patterns using the LVS technique were combined with global PSP surface static pressure mappings to characterize the leading-edge vortices and shock waves that coexist and interact at high angles of attack. The synthesis of LVS and PSP techniques was also effective in identifying the significant effects of passive surface porosity and the presence of vertical tail surfaces on the flow topologies. An overview is given of LVS and PSP applications in selected experiments on small-scale models of generic slender wing and missile configurations in the NASA Langley Research Center (NASA LaRC) Unitary Plan Wind Tunnel (UPWT) and 8-Foot Transonic Pressure Tunnel (8-Foot TPT).

Unified Application Vapor Screen Flow Visualization and Pressure Sensitive Paint Measurement Techniques to Vortex- and Shock Wave-Dominated Flow Fields

NASA Technical Reports Server (NTRS)

Erickson, Gary E.

2008-01-01

Laser vapor screen (LVS) flow visualization and pressure sensitive paint (PSP) techniques were applied in a unified approach to wind tunnel testing of slender wing and missile configurations dominated by vortex flows and shock waves at subsonic, transonic, and supersonic speeds. The off-surface cross-flow patterns using the LVS technique were combined with global PSP surface static pressure mappings to characterize the leading-edge vortices and shock waves that coexist and interact at high angles of attack (alpha). The synthesis of LVS and PSP techniques was also effective in identifying the significant effects of passive surface porosity and the presence of vertical tail surfaces on the flow topologies. An overview is given of LVS and PSP applications in selected experiments on small-scale models of generic slender wing and missile configurations in the NASA Langley Research Center (NASA LaRC) Unitary Plan Wind Tunnel (UPWT) and 8-Foot Transonic Pressure Tunnel (8-Foot TPT).

Influence of Finite Span and Sweep on Active Flow Control Efficacy

NASA Technical Reports Server (NTRS)

Greenblatt, David; Washburn, Anthony E.

2008-01-01

Active flow control efficacy was investigated by means of leading-edge and flap-shoulder zero mass-flux blowing slots on a semispan wing model that was tested in unswept (standard) and swept configurations. On the standard configuration, stall commenced inboard, but with sweep the wing stalled initially near the tip. On both configurations, leading-edge perturbations increased CL,max and post stall lift, both with and without deflected flaps. Without sweep, the effect of control was approximately uniform across the wing span but remained effective to high angles of attack near the tip; when sweep was introduced a significant effect was noted inboard, but this effect degraded along the span and produced virtually no meaningful lift enhancement near the tip, irrespective of the tip configuration. In the former case, control strengthened the wingtip vortex; in the latter case, a simple semi-empirical model, based on the trajectory or "streamline" of the evolving perturbation, served to explain the observations. In the absence of sweep, control on finite-span flaps did not differ significantly from their nominally twodimensional counterpart. Control from the flap produced expected lift enhancement and CL,max improvements in the absence of sweep, but these improvements degraded with the introduction of sweep.

Numerical Study of Tip Vortex Flows

NASA Technical Reports Server (NTRS)

Dacles-Mariani, Jennifer; Hafez, Mohamed

1998-01-01

This paper presents an overview and summary of the many different research work related to tip vortex flows and wake/trailing vortices as applied to practical engineering problems. As a literature survey paper, it outlines relevant analytical, theoretical, experimental and computational study found in literature. It also discusses in brief some of the fundamental aspects of the physics and its complexities. An appendix is also included. The topics included in this paper are: 1) Analytical Vortices; 2) Experimental Studies; 3) Computational Studies; 4) Wake Vortex Control and Management; 5) Wake Modeling; 6) High-Lift Systems; 7) Issues in Numerical Studies; 8) Instabilities; 9) Related Topics; 10) Visualization Tools for Vertical Flows; 11) Further Work Needed; 12) Acknowledgements; 13) References; and 14) Appendix.

Fundamental study of flow field generated by rotorcraft blades using wide-field shadowgraph

NASA Technical Reports Server (NTRS)

Parthasarathy, S. P.; Cho, Y. I.; Back, L. H.

1985-01-01

The vortex trajectory and vortex wake generated by helicopter rotors are visualized using a wide-field shadowgraph technique. Use of a retro-reflective Scotchlite screen makes it possible to investigate the flow field generated by full-scale rotors. Tip vortex trajectories are visible in shadowgraphs for a range of tip Mach number of 0.38 to 0.60. The effect of the angle of attack is substantial. At an angle of attack greater than 8 degrees, the visibility of the vortex core is significant even at relatively low tip Mach numbers. The theoretical analysis of the sensitivity is carried out for a rotating blade. This analysis demonstrates that the sensitivity decreases with increasing dimensionless core radius and increases with increasing tip Mach number. The threshold value of the sensitivity is found to be 0.0015, below which the vortex core is not visible and above which it is visible. The effect of the optical path length is also discussed. Based on this investigation, it is concluded that the application of this wide-field shadowgraph technique to a large wind tunnel test should be feasible. In addition, two simultaneous shadowgraph views would allow three-dimensional reconstruction of vortex trajectories.

Effect of planform and body on supersonic aerodynamics of multibody configurations

NASA Technical Reports Server (NTRS)

Mcmillin, S. Naomi; Bauer, Steven X. S.; Howell, Dorothy T.

1992-01-01

An experimental and theoretical investigation of the effect of the wing planform and bodies on the supersonic aerodynamics of a low-fineness-ratio, multibody configuration has been conducted in the Langley Unitary Plan Wind Tunnel at Mach numbers of 1.60, 1.80, 2.00, and 2.16. Force and moment data, flow-visualization data, and surface-pressure data were obtained on eight low-fineness-ratio, twin-body configurations. These configurations varied in inboard wing planform shape, outboard wing planform shape, outboard wing planform size, and presence of the bodies. The force and moment data showed that increasing the ratio of outboard wing area to total wing area or increasing the leading-edge sweep of the inboard wing influenced the aerodynamic characteristics. The flow-visualization data showed a complex flow-field system of shocks, shock-induced separation, and body vortex systems occurring between the side bodies. This flow field was substantially affected by the inboard wing planform shape but minimally affected by the outboard wing planform shape. The flow-visualization and surface-pressure data showed that flow over the outboard wing developed as expected with changes in angle of attack and Mach number and was affected by the leading-edge sweep of the inboard wing and the presence of the bodies. Evaluation of the linear-theory prediction methods revealed their general inability to consistently predict the characteristics of these multibody configurations.

In-flight total forces, moments and static aeroelastic characteristics of an oblique-wing research airplane

NASA Technical Reports Server (NTRS)

Curry, R. E.; Sim, A. G.

1984-01-01

A low-speed flight investigation has provided total force and moment coefficients and aeroelastic effects for the AD-1 oblique-wing research airplane. The results were interpreted and compared with predictions that were based on wind tunnel data. An assessment has been made of the aeroelastic wing bending design criteria. Lateral-directional trim requirements caused by asymmetry were determined. At angles of attack near stall, flow visualization indicated viscous flow separation and spanwise vortex flow. These effects were also apparent in the force and moment data.

A wind tunnel investigation of the effects of micro-vortex generators and Gurney flaps on the high-lift characteristics of a business jet wing. M.S. Thesis

NASA Technical Reports Server (NTRS)

Martuccio, Michelle Therese

1994-01-01

A study of a full-scale, semi-span business jet wing has been conducted to investigate the potential of two types of high-lift devices for improving aircraft high-lift performance. The research effort involved low-speed wind-tunnel tests of micro-vortex generators and Gurney flaps applied to the flap system of the business jet wing and included force and moment measurements, surface pressure surveys and flow visualization on the wing and flap. Results showed that the micro-vortex generators tested had no beneficial effects on the longitudinal force characteristics in this particular application, while the Gurney flaps were an effective means of increasing lift. However, the Gurney flaps also caused an increase in drag in most circumstances.

Study of lee-side flows over conically cambered Delta wings at supersonic speeds, part 2

NASA Technical Reports Server (NTRS)

Wood, Richard M.; Watson, Carolyn B.

1987-01-01

An experimental investigation was performed in which surface pressure data, flow visualization data, and force and moment data were obtained on four conical delta wing models which differed in leading edge camber only. Wing leading edge camber was achieved through a deflection of the outboard 30% of the local wing semispan of a reference 75 deg swept flat delta wing. The four wing models have leading edge deflection angles delta sub F of 0, 5, 10, and 15 deg measured streamwise. Data for the wings with delta sub F = 10 and 15 deg showed that hinge line separation dominated the lee-side wing loading and prohibited the development of leading edge separation on the deflected portion of wing leading edge. However, data for the wing with delta sub F = 5 deg showed that at an angle of attack of 5 deg, a vortex was positioned on the deflected leading edge with reattachment at the hinge line. Flow visualization results were presented which detail the influence of Mach number, angle of attack, and camber on the lee-side flow characteristics of conically cambered delta wings. Analysis of photographic data identified the existence of 12 distinctive lee-side flow types.

The Effect of Pitching Phase on the Vortex Circulation for a Flapping Wing During Stroke Reversal

NASA Astrophysics Data System (ADS)

Burge, Matthew; Ringuette, Matthew

2017-11-01

We study the effect of pitching-phase on the circulation behavior for the 3D flow structures produced during stroke reversal for a 2-degree-of-freedom flapping wing executing hovering kinematics. Previous research has related the choice in pitching-phase with respect to the wing rotation during stroke reversal (advanced vs. symmetric pitch-timing) to a lift peak preceding stroke reversal. However, results from experiments on the time-varying circulation contributions from the 3D vortex structures across the span produced by both rotation and pitching are lacking. The objective of this research is to quantitatively examine how the spanwise circulation of these structures is affected by the pitching-phase for several reduced pitching frequencies. We employ a scaled wing model in a glycerin-water mixture and measure the time-varying velocity using multiple planes of stereo digital particle image velocimetry. Data-plane positions along the wing span are informed by the unsteady behavior of the 3D vortex structures found in our prior flow visualization movies. Individual vortices are identified to calculate their circulation. This work is aimed at understanding how the behavior of the vortex structures created during stroke reversal vary with key motion parameters. This work is supported by the National Science Foundation, Award Number 1336548, supervised by Dr. Ronald Joslin.

Experimental study of wing leading-edge devices for improved maneuver performance of a supercritical maneuvering fighter configuration. [Langley 7- by 10-ft high speed tunnel tests

NASA Technical Reports Server (NTRS)

Mann, M. J.; Huffman, J. K.; Fox, C. H., Jr.; Campbell, R. L.

1983-01-01

Wind tunnel tests were conducted to examine the use of wing leading-edge devices for improved subsonic and transonic maneuver performance. These devices were tested on a fighter configuration which utilized supercritical-wing technology. The configuration had a leading-edge sweep of 45 deg and an aspect ratio of 3.28. The tests were conducted at Mach numbers of 0.60 and 0.85 with angles of attack from -0.5 deg to 22 deg. At both Mach numbers, sharp leading-edge flaps produced vortices which greatly altered the flow pattern on the wing and resulted in substantial reductions in drag at high lift. Underwing or pylon-type vortex generators also reduced drag at high lift. The vortex generators worked better at a Mach number of 0.60. The vortex generators gave the best overall results with zero toe-in angle and when mounted on either the outboard part of the wing or at both an outboard location and halfway out the semispan. Both the flaps and the vortex generators had a minor effect on the pitching moment. Fluorescent minitufts were found to be useful for flow visualization at transonic maneuver conditions.

Wing rock suppression using forebody vortex control

NASA Technical Reports Server (NTRS)

Ng, T. T.; Ong, L. Y.; Suarez, C. J.; Malcolm, G. N.

1991-01-01

Static and free-to-roll tests were conducted in a water tunnel with a configuration that consisted of a highly-slender forebody and 78-deg sweep delta wings. Flow visualization was performed and the roll angle histories were obtained. The fluid mechanisms governing the wing rock of this configuration were identified. Different means of suppressing wing rock by controlling the forebody vortices using small blowing jets were also explored. Steady blowing was found to be capable of suppressing wing rock, but significant vortex asymmetries had to be induced at the same time. On the other hand, alternating pulsed blowing on the left and right sides of the forebody was demonstrated to be potentially an effective means of suppressing wing rock and eliminating large asymmetric moments at high angles of attack.

Reactive Flow Control of Delta Wing Vortex (Postprint)

DTIC Science & Technology

2006-08-01

wing aircraft. A substantial amount of research has been dedicated to the control of aerodynamic flows using both passive and active control mechanisms...Passive vortex control devices such as vortex generators and winglets attach to the wing and require no energy input. Passive vortex control...leading edges is also effective for changing the aerodynamic characteristics of delta wings [2] [3]. Gutmark and Guillot [5] proposed controlling

Can Wing Tip Vortices Be Accurately Simulated?

DTIC Science & Technology

2011-07-01

additional tail buffeting.2 In commercial applications, winglets have been installed on passenger aircraft to minimize vortex formation and reduce lift...air. In military applications, wing tip In commercial applications, winglets have been installed on passenger aircraft to minimize increases with downstream distances.

Studies in Forecasting Upper-Level Turbulence

DTIC Science & Technology

2006-09-01

path, where they begin 9 to dissipate. Vortex size is reduced by the use of winglets , smaller “wings” that curve upward from aircraft wing tips. b...the flight path, where they begin to dissipate. Vortex size is reduced by the use of winglets , smaller “wings” that curve upward from aircraft wing

Features of flow around the flying wing model at various attack and slip angle

NASA Astrophysics Data System (ADS)

Pavlenko, A. M.; Zanin, B. Yu.; Katasonov, M. M.

2017-10-01

Experimental study of flow features around aircraft model having "flying wing" form and belonging to the category of small-unmanned aerial vehicleswas carried out. Hot-wire anemometry and flow visualization techniques were used in the investigation to get quantitative data and streamlines pictures ofthe flow near the model surface. Evolution of vortex structures depending on the attack and slip angle was demonstrated. The possibility of flow control and reduction of flow separation zones on the wing surface by means of ledges in the form of cones was also investigated. It was shown, that the laminar-turbulent transition scenario on the flying wing model is identical to the one on a straight wing and occurs through the development of a package of unstable oscillations in the boundary layer separation.

Flight test operations using an F-106B research airplane modified with a wing leading-edge vortex flap

NASA Technical Reports Server (NTRS)

Dicarlo, Daniel J.; Brown, Philip W.; Hallissy, James B.

1992-01-01

Flight tests of an F-106B aircraft equipped with a leading-edge vortex flap, which represented the culmination of a research effort to examine the effectiveness of the flap, were conducted at the NASA Langley Research Center. The purpose of the flight tests was to establish a data base on the use of a wing leading-edge vortex flap as a means to validate the design and analysis methods associated with the development of such a vortical flow-control concept. The overall experiment included: refinements of the design codes for vortex flaps; numerous wind tunnel entries to aid in verifying design codes and determining basic aerodynamic characteristics; design and fabrication of the flaps, structural modifications to the wing tip and leading edges of the test aircraft; development and installation of an aircraft research instrumentation system, including wing and flap surface pressure measurements and selected structural loads measurements; ground-based simulation to assess flying qualities; and finally, flight testing. This paper reviews the operational aspects associated with the flight experiment, which includes a description of modifications to the research airplane, the overall flight test procedures, and problems encountered. Selected research results are also presented to illustrate the accomplishments of the research effort.

Development of a High-Performance Fin-and-Tube Heat Exchanger with Vortex Generators for a Vending Machine

NASA Astrophysics Data System (ADS)

Iwasaki, Masamichi; Saito, Hiroshi; Mochizuki, Sadanari; Murata, Akira

The effect of delta-wing-vortex generators (combination of a delta wing and a delta winglet pair) on the heat transfer performance of fin-and-tube heat exchangers for vending machines has been investegated. Flow visualizations, numerical simulations and heat transfer experiments were conducted to find an optimum geometrical shape and arrangement of the vortex generators. Maximum heat transfer enhancement was achieved by the combination of (a) the delta wing with the apex angle of 86 degrees and (b) the delta winglet pair with the inline angle of 45 degrees. In relatively low Reynolds number range, about 40 % increase in heat transfer coefficient was attained with the above mentioned combination of the vortex generators compared to the ordinary heat exchangers with plain fins. It was revealed that the heat transfer enhancement was attributed to (1) the longitudinal vortexes generated by the delta wing and (2) the reduction of wake area behind the tube. It was also found that an increase in the apex angle of the delta wing brought about heat transfer enhancement, and the scale as well as the streggth of the induced longitudinal vortices played an important role in the heat transfer performance.

Flow Structure along the 1303 UCAV

NASA Astrophysics Data System (ADS)

Kosoglu, Mehmet A.; Rockwell, Donald

2007-11-01

The 1303 Unmanned Combat Air Vehicle is representative of a variety of UCAVs with blended wing-body configurations. Flow structure along a scale model of this configuration was investigated using dye visualization and particle image velocimetry for variations of Reynolds number and angle-of-attack. Both of these parameters substantially influence onset and structure of the leading-edge vortex (LEV) and a separation bubble/stall region along the tip. The onset of formation of the LEV initially occurs at a location well downstream of the apex and moves upstream for increasing values of either Reynolds number or angle-of-attack. In cases where a separation bubble or stall region exists, quantitative information on its structure was obtained via PIV imaging on a plane nearly parallel to the surface of the wing. By acquiring images on planes at successively larger elevations from the surface, it was possible to gain insight into the space-time features of the three-dimensional and highly time-dependent structure of the bubble or stall region. Time-averaged images indicate that maximum velocity defect decreases in magnitude and moves downstream with increasing elevation from the surface.

Forward rotor vortex effects on counter rotating propeller noise

NASA Technical Reports Server (NTRS)

Laur, Michele; Squires, Becky; Nagel, Robert T.

1992-01-01

Three configurations of a model counter rotating propeller manipulate the blade tip flow by: placing the CRP at angle of attack, installing shrouds, and turning the upstream blades to provide forward sweep. Flow visualization and flow measurements with thermal anemometry show no evidence of a tip vortex; however, a leading edge vortex was detected on aft swept blades. The modifications served to alter the strength and/or path of the leading edge vortex. The vortical flow is eliminated by forward sweep on the upstream propeller blades. Far field acoustic data from each test indicate only small influences on the level and directivity of the BPFs. The interaction tone at the sum of the two BPF's was significantly altered in a consistent manner. As the vortex system varied, the interaction tone was affected: far field noise levels in the forward quandrant increased and the characteristic noise minimum near the plane of rotation became less pronounced and in some cases were eliminated. If the forward propeller leading edge vortex system does not impact the rear propeller in the standard manner, a net increase in the primary interaction tone occurs for the model tested. If the leading edge vortex is removed, the interaction tone increases.

Near- and far-field aerodynamics in insect hovering flight: an integrated computational study.

PubMed

Aono, Hikaru; Liang, Fuyou; Liu, Hao

2008-01-01

We present the first integrative computational fluid dynamics (CFD) study of near- and far-field aerodynamics in insect hovering flight using a biology-inspired, dynamic flight simulator. This simulator, which has been built to encompass multiple mechanisms and principles related to insect flight, is capable of 'flying' an insect on the basis of realistic wing-body morphologies and kinematics. Our CFD study integrates near- and far-field wake dynamics and shows the detailed three-dimensional (3D) near- and far-field vortex flows: a horseshoe-shaped vortex is generated and wraps around the wing in the early down- and upstroke; subsequently, the horseshoe-shaped vortex grows into a doughnut-shaped vortex ring, with an intense jet-stream present in its core, forming the downwash; and eventually, the doughnut-shaped vortex rings of the wing pair break up into two circular vortex rings in the wake. The computed aerodynamic forces show reasonable agreement with experimental results in terms of both the mean force (vertical, horizontal and sideslip forces) and the time course over one stroke cycle (lift and drag forces). A large amount of lift force (approximately 62% of total lift force generated over a full wingbeat cycle) is generated during the upstroke, most likely due to the presence of intensive and stable, leading-edge vortices (LEVs) and wing tip vortices (TVs); and correspondingly, a much stronger downwash is observed compared to the downstroke. We also estimated hovering energetics based on the computed aerodynamic and inertial torques, and powers.

Wake Geometry Measurements and Analytical Calculations on a Small-Scale Rotor Model

NASA Technical Reports Server (NTRS)

Ghee, Terence A.; Berry, John D.; Zori, Laith A. J.; Elliott, Joe W.

1996-01-01

An experimental investigation was conducted in the Langley 14- by 22-Foot Subsonic Tunnel to quantify the rotor wake behind a scale model helicopter rotor in forward level flight at one thrust level. The rotor system in this test consisted of a four-bladed fully articulated hub with blades of rectangular planform and an NACA 0012 airfoil section. A laser light sheet, seeded with propylene glycol smoke, was used to visualize the vortex geometry in the flow in planes parallel and perpendicular to the free-stream flow. Quantitative measurements of wake geometric proper- ties, such as vortex location, vertical skew angle, and vortex particle void radius, were obtained as well as convective velocities for blade tip vortices. Comparisons were made between experimental data and four computational method predictions of experimental tip vortex locations, vortex vertical skew angles, and wake geometries. The results of these comparisons highlight difficulties of accurate wake geometry predictions.

Wind Tunnel Application of a Pressure-Sensitive Paint Technique to a Double Delta Wing Model at Subsonic and Transonic Speeds

NASA Technical Reports Server (NTRS)

Erickson, Gary E.; Gonzalez, Hugo A.

2006-01-01

A pressure-sensitive paint (PSP) technique was applied in a wind tunnel experiment in the NASA Langley Research Center 8-Foot Transonic Pressure Tunnel to study the effect of wing fillets on the global vortex induced surface static pressure field about a sharp leading-edge 76 deg./40 deg. double delta wing, or strake-wing, model at subsonic and transonic speeds. Global calibrations of the PSP were obtained at M(sub infinity) = 0.50, 0.70, 0.85, 0.95, and 1.20, a Reynolds number per unit length of 2.0 million, and angles of attack from 10 degrees to 20 degrees using an insitu method featuring the simultaneous acquisition of electronically scanned pressures (ESP) at discrete locations on the model. The mean error in the PSP measurements relative to the ESP data was approximately 2 percent or less at M(sub infinity) = 0.50 to 0.85 but increased to several percent at M(sub infinity) =0.95 and 1.20. The PSP pressure distributions and pseudo-colored, planform-view pressure maps clearly revealed the vortex-induced pressure signatures at all Mach numbers and angles of attack. Small fillets having parabolic or diamond planforms situated at the strake-wing intersection were respectively designed to manipulate the vortical flows by removing the leading-edge discontinuity or introducing additional discontinuities. The fillets caused global changes in the vortex-dominated surface pressure field that were effectively captured in the PSP measurements. The vortex surface pressure signatures were compared to available off-surface vortex cross-flow structures obtained using a laser vapor screen (LVS) flow visualization technique. The fillet effects on the PSP pressure distributions and the observed leading-edge vortex flow characteristics were consistent with the trends in the measured lift, drag, and pitching moment coefficients.

Experimental investigation of forebody and wing leading-edge vortex interactions at high angles of attack

NASA Technical Reports Server (NTRS)

Erickson, G. E.; Gilbert, W. P.

1983-01-01

An experimental investigation was conducted to assess the vortex flow-field interactions on an advanced, twin-jet fighter aircraft configuration at high angles of attack. Flow-field surveys were conducted on a small-scale model in the Northrop 0.41 - by 0.60-meter water tunnel and, where appropriate, the qualitative observations were correlated with low-speed wind tunnel data trends obtained on a large-scale model of the advanced fighter in the NASA Langley Research Center 30- by 60-foot (9.1- by 18.3-meter) facility. Emphasis was placed on understanding the interactions of the forebody and LEX-wing vortical flows, defining the effects on rolling moment variation with sideslip, and identifying modifications to control or regulate the vortex interactions at high angles of attack. The water tunnel flow visualization results and wind tunnel data trend analysis revealed the potential for strong interactions between the forebody and LEX vortices at high angles of attack. In particular, the forebody flow development near the nose could be controlled by means of carefully-positioned radome strakes. The resultant strake-induced flow-field changes were amplified downstream by the more powerful LEX vortical motions with subsequent large effects on wing flow separation characteristics.

Volumetric three-component velocimetry measurements of the turbulent flow around a Rushton turbine

NASA Astrophysics Data System (ADS)

Sharp, Kendra V.; Hill, David; Troolin, Daniel; Walters, Geoffrey; Lai, Wing

2010-01-01

Volumetric three-component velocimetry measurements have been taken of the flow field near a Rushton turbine in a stirred tank reactor. This particular flow field is highly unsteady and three-dimensional, and is characterized by a strong radial jet, large tank-scale ring vortices, and small-scale blade tip vortices. The experimental technique uses a single camera head with three apertures to obtain approximately 15,000 three-dimensional vectors in a cubic volume. These velocity data offer the most comprehensive view to date of this flow field, especially since they are acquired at three Reynolds numbers (15,000, 107,000, and 137,000). Mean velocity fields and turbulent kinetic energy quantities are calculated. The volumetric nature of the data enables tip vortex identification, vortex trajectory analysis, and calculation of vortex strength. Three identification methods for the vortices are compared based on: the calculation of circumferential vorticity; the calculation of local pressure minima via an eigenvalue approach; and the calculation of swirling strength again via an eigenvalue approach. The use of two-dimensional data and three-dimensional data is compared for vortex identification; a `swirl strength' criterion is less sensitive to completeness of the velocity gradient tensor and overall provides clearer identification of the tip vortices. The principal components of the strain rate tensor are also calculated for one Reynolds number case as these measures of stretching and compression have recently been associated with tip vortex characterization. Vortex trajectories and strength compare favorably with those in the literature. No clear dependence of trajectory on Reynolds number is deduced. The visualization of tip vortices up to 140° past blade passage in the highest Reynolds number case is notable and has not previously been shown.

Navier-Stokes, dynamics and aeroelastic computations for vortical flows, buffet and flutter applications

NASA Technical Reports Server (NTRS)

Kandil, Osama A.

1993-01-01

Research on Navier-Stokes, dynamics, and aeroelastic computations for vortical flows, buffet, and flutter applications was performed. Progress during the period from 1 Oct. 1992 to 30 Sep. 1993 is included. Papers on the following topics are included: vertical tail buffet in vortex breakdown flows; simulation of tail buffet using delta wing-vertical tail configuration; shock-vortex interaction over a 65-degree delta wing in transonic flow; supersonic vortex breakdown over a delta wing in transonic flow; and prediction and control of slender wing rock.

In-flight flow visualization characteristics of the NASA F-18 high alpha research vehicle at high angles of attack

NASA Technical Reports Server (NTRS)

Fisher, David F.; Delfrate, John H.; Richwine, David M.

1991-01-01

Surface and off-surface flow visualization techniques were used to visualize the 3-D separated flows on the NASA F-18 high alpha research vehicle at high angles of attack. Results near the alpha = 25 to 26 deg and alpha = 45 to 49 deg are presented. Both the forebody and leading edge extension (LEX) vortex cores and breakdown locations were visualized using smoke. Forebody and LEX vortex separation lines on the surface were defined using an emitted fluid technique. A laminar separation bubble was also detected on the nose cone using the emitted fluid technique and was similar to that observed in the wind tunnel test, but not as extensive. Regions of attached, separated, and vortical flow were noted on the wing and the leading edge flap using tufts and flow cones, and compared well with limited wind tunnel results.

A Study of the Flow Structure of Tip Vortices on a Hydrofoil

DTIC Science & Technology

1986-11-28

as measured from the flow visualization imager. . . 0 . . . 61 III.10 The vertical location of the tip vortex center as measured from the flow...pressure gra- dients of opposite sign exist on both sides of an airfoil . These gradients induce an inward lateral flow on the suc- tion side and an...And most recently, Cebeci et al. (1986) developed a viscous/inviscid interaction method to calculate the flow around airfoils , emphasizing the

Flow field over the wing of a delta-wing fighter model with vortex control devices at Mach 0.6 to 1.2

NASA Technical Reports Server (NTRS)

Bare, E. Ann; Reubush, David E.; Haddad, Raymond C.

1992-01-01

As part of a cooperative research program between NASA, McDonnell Douglas Corporation, and Wright Research and Development Center, a flow field investigation was conducted on a 7.52 percent scale windtunnel model of an advanced fighter aircraft design. The investigation was conducted in the Langley 16 ft Transonic Tunnel at Mach numbers of 0.6, 0.9, and 1.2. Angle of attack was varied from -4 degrees to 30 degrees and the model was tested at angles of sideslip of 0, 5, and -5 degrees. Data for the over the wing flow field were obtained at four axial survey stations by the use of six 5 hole conical probes mounted on a survey mechanism. The wing leading edge primary vortex exerted the greatest influence in terms of total pressure loss on the over the wing flow field in the area surveyed. A number of vortex control devices were also investigated. They included two different apex flaps, wing leading edge vortex flaps, and small large wing fences. The vortex flap and both apex flaps were beneficial in controlling the wing leading edge primary vortex.

Wind tunnel investigation of the interaction and breakdown characteristics of slender wing vortices at subsonic, transonic, and supersonic speeds

NASA Technical Reports Server (NTRS)

Erickson, Gary E.

1991-01-01

The vortex dominated aerodynamic characteristics of a generic 65 degree cropped delta wing model were studied in a wind tunnel at subsonic through supersonic speeds. The lee-side flow fields over the wing-alone configuration and the wing with leading edge extension (LEX) added were observed at M (infinity) equals 0.40 to 1.60 using a laser vapor screen technique. These results were correlated with surface streamline patterns, upper surface static pressure distributions, and six-component forces and moments. The wing-alone exhibited vortex breakdown and asymmetry of the breakdown location at the subsonic and transonic speeds. An earlier onset of vortex breakdown over the wing occurred at transonic speeds due to the interaction of the leading edge vortex with the normal shock wave. The development of a shock wave between the vortex and wing surface caused an early separation of the secondary boundary layer. With the LEX installed, wing vortex breakdown asymmetry did not occur up to the maximum angle of attack in the present test of 24 degrees. The favorable interaction of the LEX vortex with the wing flow field reduced the effects of shock waves on the wing primary and secondary vortical flows. The direct interaction of the wing and LEX vortex cores diminished with increasing Mach number. The maximum attainable vortex-induced pressure signatures were constrained by the vacuum pressure limit at the transonic and supersonic speeds.

Study of potential aerodynamic benefits from spanwise blowing at wingtip. Ph.D. Thesis - George Washington Univ., 1992

NASA Technical Reports Server (NTRS)

Mineck, Raymond E.

1995-01-01

Comprehensive experimental and analytical studies have been conducted to assess the potential aerodynamic benefits from spanwise blowing at the tip of a moderate-aspect-ratio swept wing. Previous studies on low-aspect-ratio wings indicated that blowing from the wingtip can diffuse the tip vortex and displace it outward. The diffused and displaced vortex will induce a smaller downwash at the wing, and consequently the wing will have increased lift and decreased induced drag at a given angle of attack. Results from the present investigation indicated that blowing from jets with a short chord had little effect on lift or drag, but blowing from jets with a longer chord increased lift near the tip and reduced drag at low Mach numbers. A Navier-Stokes solver with modified boundary conditions at the tip was used to extrapolate the results to a Mach number of 0.72. Calculations indicated that lift and drag increase with increasing jet momentum coefficient. Because the momentum of the jet is typically greater than the reduction in the wing drag and the increase in the wing lift due to spanwise blowing is small, spanwise blowing at the wingtip does not appear to be a practical means of improving the aerodynamic efficiency of moderate-aspectratio swept wings at high subsonic Mach numbers.

Mind the gap - tip leakage vortex in axial turbines

NASA Astrophysics Data System (ADS)

Dreyer, M.; Decaix, J.; Münch-Alligné, C.; Farhat, M.

2014-03-01

The tendency of designing large Kaplan turbines with a continuous increase of output power is bringing to the front the cavitation erosion issue. Due to the flow in the gap between the runner and the discharge ring, axial turbine blades may develop the so called tip leakage vortex (TLV) cavitation with negative consequences. Such vortices may interact strongly with the wake of guide vanes leading to their multiple collapses and rebounds. If the vortex trajectory remains close to the blade tip, these collapses may lead to severe erosion. One is still unable today to predict its occurrence and development in axial turbines with acceptable accuracy. Numerical flow simulations as well as the actual scale-up rules from small to large scales are unreliable. The present work addresses this problematic in a simplified case study representing TLV cavitation to better understand its sensitivity to the gap width. A Naca0009 hydrofoil is used as a generic blade in the test section of EPFL cavitation tunnel. A sliding mounting support allowing an adjustable gap between the blade tip and wall was manufactured. The vortex trajectory is visualized with a high speed camera and appropriate lighting. The three dimensional velocity field induced by the TLV is investigated using stereo particle image velocimetry. We have taken into account the vortex wandering in the image processing to obtain accurate measurements of the vortex properties. The measurements were performed in three planes located downstream of the hydrofoil for different values of the flow velocity, the incidence angle and the gap width. The results clearly reveal a strong influence of the gap width on both trajectory and intensity of the tip leakage vortex.

Vortex boundary-layer interactions

NASA Technical Reports Server (NTRS)

Bradshaw, P.

1986-01-01

Parametric studies to identify a vortex generator were completed. Data acquisition in the first chosen configuration, in which a longitudinal vortex pair generated by an isolated delta wing starts to merge with a turbulent boundary layer on a flat plate fairly close to the leading edge is nearly completed. Work on a delta-wing/flat-plate combination, consisting of a flow visualization and hot wire measurements taken with a computer controlled traverse gear and data logging system were completed. Data taking and analysis have continued, and sample results for another cross stream plane are presented. Available data include all mean velocity components, second order mean products of turbulent fluctuations, and third order mean products. Implementation of a faster data logging system was accomplished.

Numerical investigation of influence of tip leakage flow on secondary flow in transonic centrifugal compressor at design condition

NASA Astrophysics Data System (ADS)

Kaneko, Masanao; Tsujita, Hoshio

2015-04-01

In a centrifugal compressor, the leakage flow through the tip clearance generates the tip leakage vortex by the interaction with the main flow, and consequently makes the flow in the impeller passage more complex by the interaction with the passage vortex. In addition, the tip leakage vortex interacts with the shock wave on the suction surface near the blade tip in the transonic centrifugal compressor impeller. Therefore, the detailed examination for the influence of the tip leakage vortex becomes seriously important to improve the aerodynamic performance especially for the transonic centrifugal compressor. In this study, the flows in the transonic centrifugal compressor with and without the tip clearance at the design condition were analyzed numerically by using the commercial CFD code. The computed results revealed that the tip leakage vortex induced by the high loading at the blade tip around the leading edge affected the loss generation by the reduction or the suppression of the shock wave on the suction surface of the blade.

Vorticity dynamics of revolving wings: The role of planetary vortex tilting on the stability of leading-edge vortex

NASA Astrophysics Data System (ADS)

Werner, Nathaniel; Chung, Hojae; Wang, Junshi; Liu, Geng; Cimbala, John; Dong, Haibo; Cheng, Bo

2017-11-01

This work investigates the radial vorticity dynamics and the stability of leading-edge vortices (LEVs) in revolving wings. Previous studies have shown that Coriolis acceleration plays a key role in stabilizing the LEV; however, the exact mechanism remains unclear. This study tests a new hypothesis based on the curl of the Coriolis acceleration in the vorticity equation, which corresponds to the radial tilting of the planetary vortex (PVTr). The PVTr could reorient planetary vorticity into radial vorticity that reduces the strength of the LEV, preventing the LEV from growing and becoming unstable. To test this, an in-house immersed-boundary-method-based flow solver was used to generate velocity and vorticity fields of revolving wings of different aspect ratio (AR = 3, 5, 7) and Reynolds number (Re = 110, 1400). It is found that the PVTr consistently negates the LEV vorticity for all the AR and Re investigated, although its effect is outweighed by other 3D effects at Re =1400. It is also found that the strength of the PVTr increases along the wing span until approximately a chord length from the wing tip. The averaged magnitude of PVTr within the LEV and the dependency of its relative strength on the aspect ratio and Reynolds number are also investigated.

Cross-flow vortex structure and transition measurements using multi-element hot films

NASA Technical Reports Server (NTRS)

Agarwal, Naval K.; Mangalam, Siva M.; Maddalon, Dal V.; Collier, Fayette S., Jr.

1991-01-01

An experiment on a 45-degree swept wing was conducted to study three-dimensional boundary-layer characteristics using surface-mounted, micro-thin, multi-element hot-film sensors. Cross-flow vortex structure and boundary-layer transition were measured from the simultaneously acquired signals of the hot films. Spanwise variation of the root-mean-square (RMS) hot-film signal show a local minima and maxima. The distance between two minima corresponds to the stationary cross-flow vortex wavelength and agrees with naphthalene flow-visualization results. The chordwise and spanwise variation of amplified traveling (nonstationary) cross-flow disturbance characteristics were measured as Reynolds number was varied. The frequency of the most amplified cross-flow disturbances agrees with linear stability theory.

Theoretical study of aerodynamic characteristics of wings having vortex flow

NASA Technical Reports Server (NTRS)

Reddy, C. S.

1979-01-01

The aerodynamic characteristics of slender wings having separation induced vortex flows are investigated by employing three different computer codes--free vortex sheet, quasi vortex lattice, and suction analogy methods. Their capabilities and limitations are examined, and modifications are discussed. Flat wings of different configurations: arrow, delta, and diamond shapes, as well as cambered delta wings, are studied. The effect of notch ratio on the load distributions and the longitudinal characteristics of a family of arrow and diamond wings is explored. The sectional lift coefficients and the accumulated span loadings are determined for an arrow wing and are seen to be unusual in comparison with the attached flow results. The theoretically predicted results are compared with the existing experimental values.

The tubercles on humpback whales' flippers: application of bio-inspired technology.

PubMed

Fish, Frank E; Weber, Paul W; Murray, Mark M; Howle, Laurens E

2011-07-01

The humpback whale (Megaptera novaeangliae) is exceptional among the large baleen whales in its ability to undertake aquabatic maneuvers to catch prey. Humpback whales utilize extremely mobile, wing-like flippers for banking and turning. Large rounded tubercles along the leading edge of the flipper are morphological structures that are unique in nature. The tubercles on the leading edge act as passive-flow control devices that improve performance and maneuverability of the flipper. Experimental analysis of finite wing models has demonstrated that the presence of tubercles produces a delay in the angle of attack until stall, thereby increasing maximum lift and decreasing drag. Possible fluid-dynamic mechanisms for improved performance include delay of stall through generation of a vortex and modification of the boundary layer, and increase in effective span by reduction of both spanwise flow and strength of the tip vortex. The tubercles provide a bio-inspired design that has commercial viability for wing-like structures. Control of passive flow has the advantages of eliminating complex, costly, high-maintenance, and heavy control mechanisms, while improving performance for lifting bodies in air and water. The tubercles on the leading edge can be applied to the design of watercraft, aircraft, ventilation fans, and windmills.

Theory, Computation and Experiment on Criticality and Stability of Vortices Separating from Edges

DTIC Science & Technology

2016-08-15

aerospace engineering research. These include dynamic stall in wind turbines and helicopter rotors, and flapping-wing vehicle (micro-air vehicle) design...and Robinson, M., “Blade Three-Dimensional Dynamic Stall Response to Wind Turbine Operating Condition,” Journal of Solar Energy Engineering , Vol...Snapshots of TEV shedding in vortex ring representation. . . . . . . . . . . . . . . . 57 7.3 Schematic description of separated tip flow model

A low-speed wind tunnel study of vortex interaction control techniques on a chine-forebody/delta-wing configuration

NASA Technical Reports Server (NTRS)

Rao, Dhanvada M.; Bhat, M. K.

1992-01-01

A low speed wind tunnel evaluation was conducted of passive and active techniques proposed as a means to impede the interaction of forebody chine and delta wing vortices, when such interaction leads to undesirable aerodynamic characteristics particularly in the post stall regime. The passive method was based on physically disconnecting the chine/wing junction; the active technique employed deflection of inboard leading edge flaps. In either case, the intent was to forcibly shed the chine vortices before they encountered the downwash of wing vortices. Flow visualizations, wing pressures, and six component force/moment measurements confirmed the benefits of forced vortex de-coupling at post stall angles of attack and in sideslip, viz., alleviation of post stall zero beta asymmetry, lateral instability and twin tail buffet, with insignificant loss of maximum lift.

B-747 in Flight during Vortex Study

NASA Technical Reports Server (NTRS)

1974-01-01

In this 1974 NASA Flight Research Center photograph, a Boeing B-747 jetliner is shown taking part in the trailing wake vortex study. In the photograph, the two wing tip vortex trails, being the strongest, stay in tight cylindrical rolls. The 'strength' of the vortices decreases toward the midspan of each wing, and the trails become less defined. In 1974 the NASA Flight Research Center (later Dryden Flight Research Center, Edwards, California) used a Boeing 747 as part of the overall NASA study of trailing vortices. Trailing vortices are the invisible flow of spiraling air that trails from the wings of large aircraft and can 'upset' smaller aircraft flying behind them. The 747 that NASA used was on loan from the Johnson Space Center where it was part of the Space Shuttle Program. The data gathered in the 747 studies complemented data from the previous (1973-74) joint NASA Flight Research Center and Federal Aviation Administration (FAA) Boeing727 wake vortices study. Six smoke generators were installed under the wings of the 747 to provide a visual image of the trailing vortices. The object of the experiments was to test different configurations and mechanical devices on the747 that could be used to break up or lessen the strength of the vortices. The results of the tests could lead to shorter spacing between landings and takeoffs, which, in turn, could alleviate air-traffic congestion. For approximately 30 flights the 747 was flown using various combinations of wing air spoilers in an attempt to reduce wake vortices. To evaluate the effectiveness of the different configurations, chase aircraft were flown into the vortex sheets to probe their strengths and patterns at different times. Two of the chase planes used were the Flight Research Center's Cessna T-37 and the NASA Ames Research Center's Learjet. These aircraft represented the types of smaller business jets and other small aircraft that might encounter large passenger aircraft on approach or landings around major airports or in flight. Tests without the 747's wing spoilers deployed produced violent 'upset' problems for the T-37 aircraft at a distance of approximately 3 miles. From the magnitude of the problems found, distances of as much as ten miles might be required if spoilers were not used. With two spoilers on the outer wing panels, the T-37 could fly at a distance of three miles and not experience the 'upset' problem. The wake vortex study continued even after the 747 was returned to its primary mission of carrying the Space Shuttle.

NASA Trapezoidal Wing Computations Including Transition and Advanced Turbulence Modeling

NASA Technical Reports Server (NTRS)

Rumsey, C. L.; Lee-Rausch, E. M.

2012-01-01

Flow about the NASA Trapezoidal Wing is computed with several turbulence models by using grids from the first High Lift Prediction Workshop in an effort to advance understanding of computational fluid dynamics modeling for this type of flowfield. Transition is accounted for in many of the computations. In particular, a recently-developed 4-equation transition model is utilized and works well overall. Accounting for transition tends to increase lift and decrease moment, which improves the agreement with experiment. Upper surface flap separation is reduced, and agreement with experimental surface pressures and velocity profiles is improved. The predicted shape of wakes from upstream elements is strongly influenced by grid resolution in regions above the main and flap elements. Turbulence model enhancements to account for rotation and curvature have the general effect of increasing lift and improving the resolution of the wing tip vortex as it convects downstream. However, none of the models improve the prediction of surface pressures near the wing tip, where more grid resolution is needed.

Lift production through asymmetric flapping

NASA Astrophysics Data System (ADS)

Jalikop, Shreyas; Sreenivas, K. R.

2009-11-01

At present, there is a strong interest in developing Micro Air Vehicles (MAV) for applications like disaster management and aerial surveys. At these small length scales, the flight of insects and small birds suggests that unsteady aerodynamics of flapping wings can offer many advantages over fixed wing flight, such as hovering-flight, high maneuverability and high lift at large angles of attack. Various lift generating mechanims such as delayed stall, wake capture and wing rotation contribute towards our understanding of insect flight. We address the effect of asymmetric flapping of wings on lift production. By visualising the flow around a pair of rectangular wings flapping in a water tank and numerically computing the flow using a discrete vortex method, we demonstrate that net lift can be produced by introducing an asymmetry in the upstroke-to-downstroke velocity profile of the flapping wings. The competition between generation of upstroke and downstroke tip vortices appears to hold the key to understanding this lift generation mechanism.

Investigation of Unsteady Tip Clearance Flow in a Low-Speed One and Half Stage Axial Compressor with LES And PIV

NASA Technical Reports Server (NTRS)

Hah, Chunill; Hathaway, Michael; Katz, Joseph; Tan, David

2015-01-01

The primary focus of this paper is to investigate how a rotor's unsteady tip clearance flow structure changes in a low speed one and half stage axial compressor when the rotor tip gap size is increased from 0.5 mm (0.49% of rotor tip blade chord, 2% of blade span) to 2.4 mm (2.34% chord, 4% span) at the design condition are investigated. The changes in unsteady tip clearance flow with the 0.62 % tip gap as the flow rate is reduced to near stall condition are also investigated. A Large Eddy Simulation (LES) is applied to calculate the unsteady flow field at these three flow conditions. Detailed Stereoscopic PIV (SPIV) measurements of the current flow fields were also performed at the Johns Hopkins University in a refractive index-matched test facility which renders the compressor blades and casing optically transparent. With this setup, the unsteady velocity field in the entire flow domain, including the flow inside the tip gap, can be measured. Unsteady tip clearance flow fields from LES are compared with the PIV measurements and both LES and PIV results are used to study changes in tip clearance flow structures. The current study shows that the tip clearance vortex is not a single structure as traditionally perceived. The tip clearance vortex is formed by multiple interlaced vorticities. Therefore, the tip clearance vortex is inherently unsteady. The multiple interlaced vortices never roll up to form a single structure. When phased-averaged, the tip clearance vortex appears as a single structure. When flow rate is reduced with the same tip gap, the tip clearance vortex rolls further upstream and the tip clearance vortex moves further radially inward and away from the suction side of the blade. When the tip gap size is increased at the design flow condition, the overall tip clearance vortex becomes stronger and it stays closer to the blade suction side and the vortex core extends all the way to the exit of the blade passage. Measured and calculated unsteady flow fields inside the tip gap agree fairly well. Instantaneous velocity vectors inside the tip gap from both the PIV and LES do show flow separation and reattachment at the entrance of tip gap as some earlier studies suggested. This area at the entrance of tip gap flow (the pressure side of the blade) is confined very close to the rotor tip section. With a small tip gap (0.5mm), the gap flow looks like a simple two-dimensional channel flow with larger velocity near the casing for both flow rates. A small area with a sharp velocity gradient is observed just above the rotor tip. This strong shear layer is turned radially inward when it collides with the incoming flow and forms the core structure of the tip clearance vortex. When tip gap size is increased to 2.4 mm at the design operation, the radial profile of the tip gap flow changes drastically. With the large tip gap, the gap flow looks like a two-dimensional channel flow only near the casing. Near the rotor top section, a bigger region with very large shear and reversed flow is observed.

Contrail Formation in Aircraft Wakes Using Large-Eddy Simulations

NASA Technical Reports Server (NTRS)

Paoli, R.; Helie, J.; Poinsot, T. J.; Ghosal, S.

2002-01-01

In this work we analyze the issue of the formation of condensation trails ("contrails") in the near-field of an aircraft wake. The basic configuration consists in an exhaust engine jet interacting with a wing-tip training vortex. The procedure adopted relies on a mixed Eulerian/Lagrangian two-phase flow approach; a simple micro-physics model for ice growth has been used to couple ice and vapor phases. Large eddy simulations have carried out at a realistic flight Reynolds number to evaluate the effects of turbulent mixing and wake vortex dynamics on ice-growth characteristics and vapor thermodynamic properties.

The effect of wing dihedral and section suction distribution on vortex bursting

NASA Technical Reports Server (NTRS)

Washburn, K. E.; Gloss, B. B.

1975-01-01

Eleven semi-span wing models were tested in the 1/8-scale model of the Langley V/STOL tunnel to qualitatively study vortex bursting. Flow visualization was achieved by using helium filled soap bubbles introduced upstream of the model. The angle of attack range was from 0 deg to 45 deg. The results show that the vortex is unstable, that is, the bursting point location is not fixed at a given angle of attack but moves within certain bounds. Upstream of the trailing edge, the bursting point location has a range of two inches; downstream, the range is about six inches. Anhedral and dihedral appear to have an insignificant effect on the vortex and its bursting point location. Altering the section suction distribution by improving the triangularity generally increases the angle of attack at which vortex bursting occurs at the trailing edge.

Study on tip leakage vortex cavitating flows using a visualization method

NASA Astrophysics Data System (ADS)

Zhao, Yu; Jiang, Yutong; Cao, Xiaolong; Wang, Guoyu

2018-01-01

Experimental investigations of unsteady cavitating flows in a hydrofoil tip leakage region with different gap sizes are conducted to highlight the development of gap cavitation. The experiments were taken in a closed cavitation tunnel, during which high-speed camera had been used to capture the cavitation patterns. A new visualization method based on image processing was developed to capture time-dependent cavitation patterns. The results show that the visualization method can effectively capture the cavitation patterns in the tip region, including both the attached cavity in the gap and the tip leakage vortex (TLV) cavity near the trailing edge. Moreover, with the decrease of cavitation number, the TLV cavity develops from a rapid onset-growth-collapse process to a continuous process, and extends both upstream and downstream. The attached cavity in the gap develops gradually stretching beyond the gap and combines with the vortex cavity to form the triangle cavitating region. Furthermore, the influences of gap size on the cavitation are also discussed. The gap size has a great influence on the loss across the gap, and hence the locations of the inception attached cavity. Besides, inception locations and extending direction of the TLV cavity with different gap sizes also differ. The TLV in the case with τ = 0.061 is more likely to be jet-like compared with that in the case with τ = 0.024, and the gap size has a great influence on the TLV strength.

CFD Analysis of a T-38 Wing Fence

DTIC Science & Technology

2007-06-01

or making major adjustments to the existing airframe. The answer lies in flow control. Flow control devices like vortex generators, winglets , and wing...proposed by the Air Force Test Pilot School. The driving force for considering a wing fence as opposed to vane vortex generators or winglets 3 was a row of...devices are vortex generators, fences, high lift flaps, and winglets . Active flow control injects the boundary layer with energy from small jets of

Vortex Core Size in the Rotor Near-Wake

NASA Technical Reports Server (NTRS)

Young, Larry A.

2003-01-01

Using a kinetic energy conservation approach, a number of simple analytic expressions are derived for estimating the core size of tip vortices in the near-wake of rotors in hover and axial-flow flight. The influence of thrust, induced power losses, advance ratio, and vortex structure on rotor vortex core size is assessed. Experimental data from the literature is compared to the analytical results derived in this paper. In general, three conclusions can be drawn from the work in this paper. First, the greater the rotor thrust, t h e larger the vortex core size in the rotor near-wake. Second, the more efficient a rotor is with respect to induced power losses, the smaller the resulting vortex core size. Third, and lastly, vortex core size initially decreases for low axial-flow advance ratios, but for large advance ratios core size asymptotically increases to a nominal upper limit. Insights gained from this work should enable improved modeling of rotary-wing aerodynamics, as well as provide a framework for improved experimental investigations of rotor a n d advanced propeller wakes.

Membrane wing aerodynamics for micro air vehicles

NASA Astrophysics Data System (ADS)

Lian, Yongsheng; Shyy, Wei; Viieru, Dragos; Zhang, Baoning

2003-10-01

The aerodynamic performance of a wing deteriorates considerably as the Reynolds number decreases from 10 6 to 10 4. In particular, flow separation can result in substantial change in effective airfoil shape and cause reduced aerodynamic performance. Lately, there has been growing interest in developing suitable techniques for sustained and robust flight of micro air vehicles (MAVs) with a wingspan of 15 cm or smaller, flight speed around 10 m/ s, and a corresponding Reynolds number of 10 4-10 5. This paper reviews the aerodynamics of membrane and corresponding rigid wings under the MAV flight conditions. The membrane wing is observed to yield desirable characteristics in delaying stall as well as adapting to the unsteady flight environment, which is intrinsic to the designated flight speed. Flow structures associated with the low Reynolds number and low aspect ratio wing, such as pressure distribution, separation bubble and tip vortex are reviewed. Structural dynamics in response to the surrounding flow field is presented to highlight the multiple time-scale phenomena. Based on the computational capabilities for treating moving boundary problems, wing shape optimization can be conducted in automated manners. To enhance the lift, the effect of endplates is evaluated. The proper orthogonal decomposition method is also discussed as an economic tool to describe the flow structure around a wing and to facilitate flow and vehicle control.

Endplate effect on aerodynamic characteristics of threedimensional wings in close free surface proximity

NASA Astrophysics Data System (ADS)

Jung, Jae Hwan; Kim, Mi Jeong; Yoon, Hyun Sik; Hung, Pham Anh; Chun, Ho Hwan; Park, Dong Woo

2012-12-01

We investigated the aerodynamic characteristics of a three-dimensional (3D) wing with an endplate in the vicinity of the free surface by solving incompressible Navier-Stokes equations with the turbulence closure model. The endplate causes a blockage effect on the flow, and an additional viscous effect especially near the endplate. These combined effects of the endplate significantly reduce the magnitudes of the velocities under the lower surface of the wing, thereby enhancing aerodynamic performance in terms of the force coefficients. The maximum lift-to-drag ratio of a wing with an endplate is increased 46% compared to that of wing without an endplate at the lowest clearance. The tip vortex of a wing-with-endplate (WWE) moved laterally to a greater extent than that of a wing-without-endplate (WOE). This causes a decrease in the induced drag, resulting in a reduction in the total drag.

Large eddy simulation of tip-leakage flow in an axial flow fan

NASA Astrophysics Data System (ADS)

Park, Keuntae; Choi, Haecheon; Choi, Seokho; Sa, Yongcheol; Kwon, Oh-Kyoung

2016-11-01

An axial flow fan with a shroud generates a complicated tip-leakage flow by the interaction of the axial flow with the fan blades and shroud near the blade tips. In this study, large eddy simulation is performed for tip-leakage flow in a forward-swept axial flow fan inside an outdoor unit of an air-conditioner, operating at the design condition of the Reynolds number of 547,000 based on the radius of blade tip and the tip velocity. A dynamic global model is used for a subgrid-scale model, and an immersed boundary method in a non-inertial reference frame is adopted. The present simulation clearly reveals the generation and evolution of tip-leakage vortex near the blade tip by the leakage flow. At the inception of the leakage vortex near the leading edge of the suction-side of the blade tip, the leakage vortex is composed of unsteady multiple vortices containing high-frequency fluctuations. As the leakage vortex develops downstream along a slant line toward the following blade, large and meandering movements of the leakage vortex are observed. Thus low-frequency broad peaks of velocity and pressure occur near the pressure surface. Supported by the KISTI Supercomputing Center (KSC-2016-C3-0027).

Aeroacoustic flowfield and acoustics of a model helicopter tail rotor at high advance ratio

NASA Technical Reports Server (NTRS)

Shenoy, Rajarama K.

1989-01-01

Some results, relevant to rotorcraft noise generation process at high advance ratio, are presented in this paper from schlieren flow visualization and acoustic tests of a model tail rotor. The measured in-plane noise trends are consistent with the growth of the tip supersonic region seen in the schlieren visuals. Schlieren flow visuals reveal a propagating pressure wave in the second quadrant. Simultaneously measured acoustic data and the results of two-dimensional transonic Blade-Vortex Interaction analysis code ATRAN-2 indicate that this pressure wave is attributable to BVI activity in the first quadrant. This paper establishes that the transonic Blade-Vortex Interactions contribute to noise at high advance ratio level flight conditions.

High Speed Vortex Flows

NASA Technical Reports Server (NTRS)

Wood, Richard M.; Wilcox, Floyd J., Jr.; Bauer, Steven X. S.; Allen, Jerry M.

2000-01-01

A review of the research conducted at the National Aeronautics and Space Administration (NASA), Langley Research Center (LaRC) into high-speed vortex flows during the 1970s, 1980s, and 1990s is presented. The data reviewed is for flat plates, cavities, bodies, missiles, wings, and aircraft. These data are presented and discussed relative to the design of future vehicles. Also presented is a brief historical review of the extensive body of high-speed vortex flow research from the 1940s to the present in order to provide perspective of the NASA LaRC's high-speed research results. Data are presented which show the types of vortex structures which occur at supersonic speeds and the impact of these flow structures to vehicle performance and control is discussed. The data presented shows the presence of both small- and large scale vortex structures for a variety of vehicles, from missiles to transports. For cavities, the data show very complex multiple vortex structures exist at all combinations of cavity depth to length ratios and Mach number. The data for missiles show the existence of very strong interference effects between body and/or fin vortices and the downstream fins. It was shown that these vortex flow interference effects could be both positive and negative. Data are shown which highlights the effect that leading-edge sweep, leading-edge bluntness, wing thickness, location of maximum thickness, and camber has on the aerodynamics of and flow over delta wings. The observed flow fields for delta wings (i.e. separation bubble, classical vortex, vortex with shock, etc.) are discussed in the context of' aircraft design. And data have been shown that indicate that aerodynamic performance improvements are available by considering vortex flows as a primary design feature. Finally a discussing of a design approach for wings which utilize vortex flows for improved aerodynamic performance at supersonic speed is presented.

Prediction and control of vortex-dominated and vortex-wake flows

NASA Technical Reports Server (NTRS)

Kandil, Osama

1993-01-01

This progress report documents the accomplishments achieved in the period from December 1, 1992 until November 30, 1993. These accomplishments include publications, national and international presentations, NASA presentations, and the research group supported under this grant. Topics covered by documents incorporated into this progress report include: active control of asymmetric conical flow using spinning and rotary oscillation; supersonic vortex breakdown over a delta wing in transonic flow; shock-vortex interaction over a 65-degree delta wing in transonic flow; three dimensional supersonic vortex breakdown; numerical simulation and physical aspects of supersonic vortex breakdown; and prediction of asymmetric vortical flows around slender bodies using Navier-Stokes equations.

The migration and growth of nuclei in an ideal vortex flow

NASA Astrophysics Data System (ADS)

Zhang, Lingxin; Chen, Linya; Shao, Xueming

2016-12-01

Tip vortex cavitation occurs on ship propellers which can cause significant noise compared to the wet flow. In order to predict the inception of tip vortex cavitation, numerous researches have been investigated about the detailed flow field around the tip. According to informed studies, the inception of tip vortex cavitation is affected by many factors. To understand the effect of water quality on cavitation inception, the motion of nuclei in an ideal vortex flow, i.e., the Rankine vortex flow, was investigated. The one-way coupling point-particle tracking model was employed to simulate the trajectory of nuclei. Meanwhile, Rayleigh-Plesset equation was introduced to describe the growth of nuclei. The results show that the nucleus size has a significant effect on nucleus' trajectory. The capture time of a nucleus is approximately inversely proportional to its radius. The growth of nucleus accelerates its migration in the vortex flow and shortens its capture time, especially for the case of explosive growth.

Tip Vortex Cavitation

NASA Astrophysics Data System (ADS)

Maines, Brant H.; Arndt, Roger E. A.

2000-11-01

Cavitation in vortical flows is a problem of practical importance, that is relatively unexplored. Vortical structures of importance range from the eddies occurring randomly in space and time in turbulent flows to the developed vortices that occur at the tips of lifting surfaces and at the hubs of propellers and hydraulic turbines. A variety of secondary flow phenomena such as the horse shoe vortices that form around bridge piers, chute blocks and struts, and the secondary vortices found in the clearance passages of turbomachinery are also important cavitation sites. Tip vortex cavitation can be viewed as a canonical problem that captures many of the essential physics associated with vortex cavitation in general. This paper describes the inception process and focuses on the high levels of tension that can be sustained in the flow, which appears to scale with the blade loading. High speed video visualization indicates that the details of how free stream nuclei are ingested plays a major role in the nucleation and inception process. A new photographic technique was used to obtain high quality images of the bubble growth process at framing rates as high as 40,000 fps. Sponsored by the Office of Naval Research

Investigation of Unsteady Flow Field in a Low-Speed One and a Half Stage Axial Compressor. Part 2; Effects of Tip Gap Size On the Tip Clearance Flow Structure at Near Stall Operation

NASA Technical Reports Server (NTRS)

Hah, Chunill; Hathaway, Michael; Katz, Joseph

2014-01-01

The primary focus of this paper is to investigate the effect of rotor tip gap size on how the rotor unsteady tip clearance flow structure changes in a low speed one and half stage axial compressor at near stall operation (for example, where maximum pressure rise is obtained). A Large Eddy Simulation (LES) is applied to calculate the unsteady flow field at this flow condition with both a small and a large tip gaps. The numerically obtained flow fields at the small clearance matches fairly well with the available initial measurements obtained at the Johns Hopkins University with 3-D unsteady PIV in an index-matched test facility which renders the compressor blades and casing optically transparent. With this setup, the unsteady velocity field in the entire flow domain, including the flow inside the tip gap, can be measured. The numerical results are also compared with previously published measurements in a low speed single stage compressor (Maerz et al. [2002]). The current study shows that, with the smaller rotor tip gap, the tip clearance vortex moves to the leading edge plane at near stall operating condition, creating a nearly circumferentially aligned vortex that persists around the entire rotor. On the other hand, with a large tip gap, the clearance vortex stays inside the blade passage at near stall operation. With the large tip gap, flow instability and related large pressure fluctuation at the leading edge are observed in this one and a half stage compressor. Detailed examination of the unsteady flow structure in this compressor stage reveals that the flow instability is due to shed vortices near the leading edge, and not due to a three-dimensional separation vortex originating from the suction side of the blade, which is commonly referred to during a spike-type stall inception. The entire tip clearance flow is highly unsteady. Many vortex structures in the tip clearance flow, including the sheet vortex system near the casing, interact with each other. The core tip clearance vortex, which is formed with the rotor tip gap flows near the leading edge, is also highly unsteady or intermittent due to pressure oscillations near the leading edge and varies from passage to passage. For the current compressor stage, the evidence does not seem to support that a classical vortex breakup occurs in any organized way, even with the large tip gap. Although wakes from the IGV influence the tip clearance flow in the rotor, the major characteristics of rotor tip clearance flows in isolated or single stage rotors are observed in this one and a half stage axial compressor.

Crossflow Stability and Transition Experiments in Swept-Wing Flow

NASA Technical Reports Server (NTRS)

Dagenhart, J. Ray; Saric, William S.

1999-01-01

An experimental examination of crossflow instability and transition on a 45deg swept wing was conducted in the Arizona State University Unsteady Wind Tunnel. The stationary-vortex pattern and transition location are visualized by using both sublimating chemical and liquid-crystal coatings. Extensive hot-wire measurements were obtained at several measurement stations across a single vortex track. The mean and travelling wave disturbances were measured simultaneously. Stationary crossflow disturbance profiles were determined by subtracting either a reference or a span-averaged velocity profile from the mean velocity data. Mean, stationary crossflow, and traveling wave velocity data were presented as local boundary layer profiles and contour plots across a single stationary crossflow vortex track. Disturbance mode profiles and growth rates were determined. The experimental data are compared with predictions from linear stability theory.

Leading-edge vortex research: Some nonplanar concepts and current challenges

NASA Technical Reports Server (NTRS)

Campbell, J. F.; Osborn, R. F.

1986-01-01

Some background information is provided for the Vortex Flow Aerodynamics Conference and that current slender wing airplanes do not use variable leading edge geometry to improve transonic drag polar is shown. Highlights of some of the initial studies combining wing camber, or flaps, with vortex flow are presented. Current vortex flap studies were reviewed to show that there is a large subsonic data base and that transonic and supersonic generic studies have begun. There is a need for validated flow field solvers to calculate vortex/shock interactions at transonic and supersonic speeds. Many important research opportunities exist for fundamental vortex flow investigations and for designing advanced fighter concepts.

Aeroelastic loads prediction for an arrow wing. Task 1: Evaluation of R. P. White's method

NASA Technical Reports Server (NTRS)

Borland, C. J.; Manro, M. E.

1983-01-01

The separated flow method is evaluated. This method was developed for moderately swept wings with multiple, constant strength vortex systems. The flow on the highly swept wing used in this evaluation is characterized by a single vortex system of continuously varying strength.

Multi-cored vortices support function of slotted wing tips of birds in gliding and flapping flight

PubMed Central

2017-01-01

Slotted wing tips of birds are commonly considered an adaptation to improve soaring performance, despite their presence in species that neither soar nor glide. We used particle image velocimetry to measure the airflow around the slotted wing tip of a jackdaw (Corvus monedula) as well as in its wake during unrestrained flight in a wind tunnel. The separated primary feathers produce individual wakes, confirming a multi-slotted function, in both gliding and flapping flight. The resulting multi-cored wingtip vortex represents a spreading of vorticity, which has previously been suggested as indicative of increased aerodynamic efficiency. Considering benefits of the slotted wing tips that are specific to flapping flight combined with the wide phylogenetic occurrence of this configuration, we propose the hypothesis that slotted wings evolved initially to improve performance in powered flight. PMID:28539482

An Evaluation of the Measurement Requirements for an In-Situ Wake Vortex Detection System

NASA Technical Reports Server (NTRS)

Fuhrmann, Henri D.; Stewart, Eric C.

1996-01-01

Results of a numerical simulation are presented to determine the feasibility of estimating the location and strength of a wake vortex from imperfect in-situ measurements. These estimates could be used to provide information to a pilot on how to avoid a hazardous wake vortex encounter. An iterative algorithm based on the method of secants was used to solve the four simultaneous equations describing the two-dimensional flow field around a pair of parallel counter-rotating vortices of equal and constant strength. The flow field information used by the algorithm could be derived from measurements from flow angle sensors mounted on the wing-tip of the detecting aircraft and an inertial navigation system. The study determined the propagated errors in the estimated location and strength of the vortex which resulted from random errors added to theoretically perfect measurements. The results are summarized in a series of charts and a table which make it possible to estimate these propagated errors for many practical situations. The situations include several generator-detector airplane combinations, different distances between the vortex and the detector airplane, as well as different levels of total measurement error.

Effect of radius of gyration on a wing rotating at low Reynolds number: A computational study

NASA Astrophysics Data System (ADS)

Tudball Smith, Daniel; Rockwell, Donald; Sheridan, John; Thompson, Mark

2017-06-01

This computational study analyzes the effect of variation of the radius of gyration (rg), expressed as the Rossby number Ro=rg/C , with C the chord, on the aerodynamics of a rotating wing at a Reynolds number of 1400. The wing is represented as an aspect-ratio-unity rectangular flat plate aligned at 45 ∘ . This plate is accelerated near impulsively to a constant rotational velocity and the flow is allowed to develop. Flow structures are analyzed and force coefficients evaluated. Trends in velocity field degradation with increasing Ro are consistent with previous experimental studies. At low Ro the flow structure generated initially is mostly retained with a strong laminar leading-edge vortex (LEV) and tip vortex (TV). As both Ro and travel distance increase, the flow structure degrades such that at high Ro it begins to resemble that of a translating wing. Additionally, the present study has shown the following. (i) At low Ro the LEV and TV structure is laminar and steady; as Ro increases this structure breaks down, and the location at which it breaks down shifts closer to the wing root. (ii) For moderate Ro of 1.4 and higher, the LEV is no longer steady but enters a shedding regime fed by the leading-edge shear layer. (iii) At the lowest Ro of 0.7 the lift force rises during start-up and then stabilizes, consistent with the flow structure being retained, while for higher Ro a force peak occurs after the initial acceleration is complete, followed by a reduction in lift which appears to correspond to shedding of excess leading-edge vorticity generated during start-up. (iv) All rotating wings produced greater lift than a translating wing, this increase varied from ˜65 % at the lowest Ro=0.7 down to ˜5 % for the highest Ro examined of 9.1.

Wake Development behind Paired Wings with Tip and Root Trailing Vortices: Consequences for Animal Flight Force Estimates

PubMed Central

Horstmann, Jan T.; Henningsson, Per; Thomas, Adrian L. R.; Bomphrey, Richard J.

2014-01-01

Recent experiments on flapping flight in animals have shown that a variety of unrelated species shed a wake behind left and right wings consisting of both tip and root vortices. Here we present an investigation using Particle Image Velocimetry (PIV) of the behaviour and interaction of trailing vortices shed by paired, fixed wings that simplify and mimic the wake of a flying animal with a non-lifting body. We measured flow velocities at five positions downstream of two adjacent NACA 0012 aerofoils and systematically varied aspect ratio, the gap between the wings (corresponding to the width of a non-lifting body), angle of attack, and the Reynolds number. The range of aspect ratios and Reynolds number where chosen to be relevant to natural fliers and swimmers, and insect flight in particular. We show that the wake behind the paired wings deformed as a consequence of the induced flow distribution such that the wingtip vortices convected downwards while the root vortices twist around each other. Vortex interaction and wake deformation became more pronounced further downstream of the wing, so the positioning of PIV measurement planes in experiments on flying animals has an important effect on subsequent force estimates due to rotating induced flow vectors. Wake deformation was most severe behind wings with lower aspect ratios and when the distance between the wings was small, suggesting that animals that match this description constitute high-risk groups in terms of measurement error. Our results, therefore, have significant implications for experimental design where wake measurements are used to estimate forces generated in animal flight. In particular, the downstream distance of the measurement plane should be minimised, notwithstanding the animal welfare constraints when measuring the wake behind flying animals. PMID:24632825

Influence of blade tip rounding on tip leakage vortex cavitation of axial flow pump

NASA Astrophysics Data System (ADS)

Wu, S. Q.; Shi, W. D.; Zhang, D. S.; Yao, J.; Cheng, C.

2013-12-01

Tip leakage flow in axial flow pumps is mainly caused by the tip clearance, which is the main cause of tip leakage vortex cavitation and blade tip cavitation erosion. In order to improve tip clearance flow and reduce TLV cavitation, four schemes were adopted to the round blade tip. These are: no tip rounding, one time tip clearance tip rounding, two times tip clearance tip rounding, four times tip clearance tip rounding. Using SST k-ω turbulence model and Zwart cavitation model in CFX software, this simulation obtained four kinds of inner flow field results. The numerical results indicated that with the increase of r*, NPSHc gradually increased and the cavitation performance reduced. However, corner vortex was eliminated so that cavitation in gap was restrained. But TLV vorticity increased and cavitation's range here had a little expansion. Combined with the research of this paper and the different analyses of four schemes, we recommend adopting the two times of the tip clearance rounding.

Dynamics of the vortex wakes of flying and swimming vertebrates.

PubMed

Rayner, J M

1995-01-01

The vortex wakes of flying and swimming animals provide evidence of the history of aero- and hydrodynamic force generation during the locomotor cycle. Vortex-induced momentum flux in the wake is the reaction of forces the animal imposes on its environment, which must be in equilibrium with inertial and external forces. In flying birds and bats, the flapping wings generate lift both to provide thrust and to support the weight. Distinct wingbeat and wake movement patterns can be identified as gaits. In flow visualization experiments, only two wake patterns have been identified: a vortex ring gait with inactive upstroke, and a continuous vortex gait with active upstroke. These gaits may be modelled theoretically by free vortex and lifting line theory to predict mechanical energy consumption, aerodynamic forces and muscle activity. Longer-winged birds undergo a distinct gait change with speed, but shorter-winged species use the vortex ring gait at all speeds. In swimming fish, the situation is more complex: the wake vortices form a reversed von Kármán vortex street, but little is known about the mechanism of generation of the wake, or about how it varies with speed and acceleration or with body form and swimming mode. An unresolved complicating factor is the interaction between the drag wake of the flapping fish body and the thrusting wake from the tail.

A comparison with theory of peak to peak sound level for a model helicopter rotor generating blade slap at low tip speeds

NASA Technical Reports Server (NTRS)

Fontana, R. R.; Hubbard, J. E., Jr.

1983-01-01

Mini-tuft and smoke flow visualization techniques have been developed for the investigation of model helicopter rotor blade vortex interaction noise at low tip speeds. These techniques allow the parameters required for calculation of the blade vortex interaction noise using the Widnall/Wolf model to be determined. The measured acoustics are compared with the predicted acoustics for each test condition. Under the conditions tested it is determined that the dominating acoustic pulse results from the interaction of the blade with a vortex 1-1/4 revolutions old at an interaction angle of less than 8 deg. The Widnall/Wolf model predicts the peak sound pressure level within 3 dB for blade vortex separation distances greater than 1 semichord, but it generally over predicts the peak S.P.L. by over 10 dB for blade vortex separation distances of less than 1/4 semichord.

Influence of airfoil geometry on delta wing leading-edge vortices and vortex-induced aerodynamics at supersonic speeds

NASA Technical Reports Server (NTRS)

Wood, Richard M.; Byrd, James E.; Wesselmann, Gary F.

1992-01-01

An assessment of the influence of airfoil geometry on delta wing leading edge vortex flow and vortex induced aerodynamics at supersonic speeds is discussed. A series of delta wing wind tunnel models were tested over a Mach number range from 1.7 to 2.0. The model geometric variables included leading edge sweep and airfoil shape. Surface pressure data, vapor screen, and oil flow photograph data were taken to evaluate the complex structure of the vortices and shocks on the family of wings tested. The data show that airfoil shape has a significant impact on the wing upper surface flow structure and pressure distribution, but has a minimal impact on the integrated upper surface pressure increments.

Vortex Flows at Supersonic Speeds

NASA Technical Reports Server (NTRS)

Wood, Richard M.; Wilcox, Floyd J., Jr.; Bauer, Steven X. S.; Allen, Jerry M.

2003-01-01

A review of research conducted at the National Aeronautics and Space Administration (NASA) Langley Research Center (LaRC) into high-speed vortex flows during the 1970s, 1980s, and 1990s is presented. The data are for flat plates, cavities, bodies, missiles, wings, and aircraft with Mach numbers of 1.5 to 4.6. Data are presented to show the types of vortex structures that occur at supersonic speeds and the impact of these flow structures on vehicle performance and control. The data show the presence of both small- and large-scale vortex structures for a variety of vehicles, from missiles to transports. For cavities, the data show very complex multiple vortex structures exist at all combinations of cavity depth to length ratios and Mach number. The data for missiles show the existence of very strong interference effects between body and/or fin vortices. Data are shown that highlight the effect of leading-edge sweep, leading-edge bluntness, wing thickness, location of maximum thickness, and camber on the aerodynamics of and flow over delta wings. Finally, a discussion of a design approach for wings that use vortex flows for improved aerodynamic performance at supersonic speeds is presented.

An experimental study of an airfoil with a bio-inspired leading edge device at high angles of attack

NASA Astrophysics Data System (ADS)

Mandadzhiev, Boris A.; Lynch, Michael K.; Chamorro, Leonardo P.; Wissa, Aimy A.

2017-09-01

Robust and predictable aerodynamic performance of unmanned aerial vehicles at the limits of their design envelope is critical for safety and mission adaptability. Deployable aerodynamic surfaces from the wing leading or trailing edges are often used to extend the aerodynamic envelope (e.g. slats and flaps). Birds have also evolved feathers at the leading edge (LE) of their wings, known as the alula, which enables them to perform high angles of attack maneuvers. In this study, a series of wind tunnel experiments are performed to quantify the effect of various deployment parameters of an alula-like LE device on the aerodynamic performance of a cambered airfoil (S1223) at stall and post stall conditions. The alula relative angle of attack, measured from the mean chord of the airfoil, is varied to modulate tip-vortex strength, while the alula deflection angle is varied to modulate the distance between the tip vortex and the wing surface. Integrated lift force measurements were collected at various alula-inspired device configurations. The effect of the alula-inspired device on the boundary layer velocity profile and turbulence intensity were investigated through hot-wire anemometer measurements. Results show that as alula deflection angle increases, the lift coefficient also increase especially at lower alula relative angles of attack. Moreover, at post stall wing angles of attack, the wake velocity deficit is reduced in the presence of alula device, confirming the mitigation of the wing adverse pressure gradient. The results are in strong agreement with measurements taken on bird wings showing delayed flow reversal and extended range of operational angles of attack. An engineered alula-inspired device has the potential to improve mission adaptability in small unmanned air vehicles during low Reynolds number flight.

Fundamental aerodynamic characteristics of delta wings with leading-edge vortex flows

NASA Technical Reports Server (NTRS)

Wood, R. M.; Miller, D. S.

1985-01-01

An investigation of the aerodynamics of sharp leading-edge delta wings at supersonic speeds has been conducted. The supporting experimental data for this investigation were taken from published force, pressure, and flow-visualization data in which the Mach number normal to the wing leading edge is always less than 1.0. The individual upper- and lower-surface nonlinear characteristics for uncambered delta wings are determined and presented in three charts. The upper-surface data show that both the normal-force coefficient and minimum pressure coefficient increase nonlinearly with a decreasing slope with increasing angle of attack. The lower-surface normal-force coefficient was shown to be independent of Mach number and to increase nonlinearly, with an increasing slope, with increasing angle of attack. These charts are then used to define a wing-design space for sharp leading-edge delta wings.

Calculation of Wing Bending Moments and Tail Loads Resulting from the Jettison of Wing Tips During a Symmetrical Pull-Up

NASA Technical Reports Server (NTRS)

Boshar, John

1947-01-01

A preliminary analytical investigation was made to determine the feasibility of the basic idea of controlled failure points as safety valves for the primary airplane structure. The present analysis considers the possibilities of the breakable wing tip which, in failing as a weak link, would relieve the bending moments on the wing structure. The analysis was carried out by computing the time histories of the wing and stabilizer angle of attack in a 10g pull-up for an XF8F airplane with tips fixed and comparing the results with those for the same maneuver, that is, elevator motion but with tips jettisoned at 8g. The calculations indicate that the increased stability accompanying the loss of the wing tips reduces the bending moment an additional amount above that which would be expected from the initial loss in lift and the inboard shift in load. The vortex shed when the tips are lost may induce a transient load requiring that the tail be made stronger than otherwise.

Vortex modeling for rotor aerodynamics - The 1991 Alexander A. Nikolsky Lecture

NASA Technical Reports Server (NTRS)

Gray, Robin B.

1992-01-01

The efforts toward realistic vortex modeling for rotary wings which began under the guidance of professor A. A. Nikolsky of Princeton University in 1955-1956 are discussed. Attention is given to Nikolsky's flow-visualization studies and major theoretical considerations for vortex modeling. More recent efforts by other researchers have led to models of increasing complexity. The neglect of compressibility and viscous effects in the classical approach is noted to be a major limiting factor in full-scale rotor applications of the classical vortex theory; it has nevertheless been valuable for the delineation of problem areas and the guiding of both experimental and theoretical investigations.

High-Speed Boundary-Layer Transition: Study of Stationary Crossflow Using Spectral Analysis

NASA Astrophysics Data System (ADS)

McGuire, Patrick Joseph

Crossflow instability is primary cause of boundary-layer transition on swept wings used in high-speed applications. Delaying the downstream location of transition would drastically reduce the viscous drag over the wing surface, and subsequently improves the overall aircraft efficiency. By studying the development of instability growth rates and how they interact with the surroundings, researchers can control the crossflow transition location. Experiments on the 35° swept-wing model were performed in the NASA Langley 20-Inch Supersonic Wind Tunnel with Mach 2.0 flow conditions and 20 μm tall discrete roughness elements (DRE) with varying spacing placed along the leading edge. Fluorene was used as the sublimating chemical in the surface flow visualization technique to observe the transition front and stationary crossflow vortex patterns in the laminar flow region. Spatial spectral decomposition was completed on high-resolution images of sublimating chemical runs using a newly developed image processing technique. Streamwise evolution of the vortex track wavelengths within the laminar boundary-layer region was observed. The spectral information was averaged to produce dominant modes present throughout the laminar region.

Multi-cored vortices support function of slotted wing tips of birds in gliding and flapping flight.

PubMed

KleinHeerenbrink, Marco; Johansson, L Christoffer; Hedenström, Anders

2017-05-01

Slotted wing tips of birds are commonly considered an adaptation to improve soaring performance, despite their presence in species that neither soar nor glide. We used particle image velocimetry to measure the airflow around the slotted wing tip of a jackdaw ( Corvus monedula ) as well as in its wake during unrestrained flight in a wind tunnel. The separated primary feathers produce individual wakes, confirming a multi-slotted function, in both gliding and flapping flight. The resulting multi-cored wingtip vortex represents a spreading of vorticity, which has previously been suggested as indicative of increased aerodynamic efficiency. Considering benefits of the slotted wing tips that are specific to flapping flight combined with the wide phylogenetic occurrence of this configuration, we propose the hypothesis that slotted wings evolved initially to improve performance in powered flight. © 2017 The Author(s).

Crossflow Stability and Transition Experiments in a Swept-Wing Flow. Ph.D. Thesis

NASA Technical Reports Server (NTRS)

Dagenhart, John Ray

1992-01-01

An experimental examination of crossflow instability and transition on a 45 degree swept wing is conducted in the Arizona State University Unsteady Wind Tunnel. The stationary-vortex pattern and transition location are visualized using both sublimating-chemical and liquid-crystal coatings. Extensive hot-wire measurements are conducted at several measurement stations across a single vortex track. The mean and travelling-wave disturbances are measured simultaneously. Stationary-crossflow disturbance profiles are determined by subtracting either a reference or a span-averaged velocity profile from the mean-velocity data. Mean, stationary-crossflow, and travelling-wave velocity data are presented as local boundary-layer profiles and as contour plots across a single stationary-crossflow vortex track. Disturbance-mode profiles and growth rates are determined. The experimental data are compared to predictions from linear stability theory.

Further Studies of the Response of Single Rotor Helicopters to Vortex Encounters

DOT National Transportation Integrated Search

1985-09-01

This report is a continuation of the studies described in Reference where a simplified approach to the problem of predicting the uncontrolled response of a single rotor helicopter to an encounter with the wing tip vortex of a large transport aircraft...

A Three-Dimensional Solution of Flows over Wings with Leading-Edge Vortex Separation. Part 1: Engineering Document

NASA Technical Reports Server (NTRS)

Brune, G. W.; Weber, J. A.; Johnson, F. T.; Lu, P.; Rubbert, P. E.

1975-01-01

A method of predicting forces, moments, and detailed surface pressures on thin, sharp-edged wings with leading-edge vortex separation in incompressible flow is presented. The method employs an inviscid flow model in which the wing and the rolled-up vortex sheets are represented by piecewise, continuous quadratic doublet sheet distributions. The Kutta condition is imposed on all wing edges. Computed results are compared with experimental data and with the predictions of the leading-edge suction analogy for a selected number of wing planforms over a wide range of angle of attack. These comparisons show the method to be very promising, capable of producing not only force predictions, but also accurate predictions of detailed surface pressure distributions, loads, and moments.

Flow-field in a vortex with breakdown above sharp edged delta wings

NASA Technical Reports Server (NTRS)

Hayashi, Y.; Nakaya, T.

1978-01-01

The behavior of vortex-flow, accompanied with breakdown, formed above sharp-edged delta wings, was studied experimentally as well as theoretically. Emphasis is placed particularly on the criterion for the breakdown at sufficiently large Reynolds numbers

Slender wing theory including regions of embedded total pressure loss

NASA Technical Reports Server (NTRS)

Mccune, James E.; Tavares, T. Sean; Lee, Norman K. W.; Weissbein, David

1988-01-01

An aerodynamic theory of the flow about slender delta wings is described. The theory includes a treatment of the self-consistent development of the vortex wake patterns above the wing necessary to maintain smooth flow at the wing edges. The paper focuses especially on the formation within the wake of vortex 'cores' as embedded regions of total pressure loss, fed and maintained by umbilical vortex sheets emanating from the wing edges. Criteria are developed for determining the growing size and location of these cores, as well as the distribution and strength of the vorticity within them. In this paper, however, the possibility of vortex breakup is omitted. The aerodynamic consequences of the presence and evolution of the cores and the associated wake structure are illustrated and discussed. It is noted that wake history effects can have substantial influence on the distribution of normal force on the wing as well as on its magnitude.

Subsonic investigations of vortex interaction control for enhanced high-alpha aerodynamics of a chine forebody/Delta wing configuration

NASA Technical Reports Server (NTRS)

Rao, Dhanvada M.; Bhat, M. K.

1992-01-01

A proposed concept to alleviate high alpha asymmetry and lateral/directional instability by decoupling of forebody and wing vortices was studied on a generic chine forebody/ 60 deg. delta configuration in the NASA Langley 7 by 10 foot High Speed Tunnel. The decoupling technique involved inboard leading edge flaps of varying span and deflection angle. Six component force/moment characteristics, surface pressure distributions and vapor-screen flow visualizations were acquired, on the basic wing-body configuration and with both single and twin vertical tails at M sub infinity = 0.1 and 0.4, and in the range alpha = 0 to 50 deg and beta = -10 to +10 degs. Results are presented which highlight the potential of vortex decoupling via leading edge flaps for enhanced high alpha lateral/directional characteristics.

Aerodynamic control of NASP-type vehicles through vortex manipulation. Volume 3: Wing rock experiments

NASA Technical Reports Server (NTRS)

Suarez, Carlos J.; Smith, Brooke C.; Kramer, Brian R.; Ng, T. Terry; Ong, Lih-Yenn; Malcolm, Gerald N.

1993-01-01

Free-to-roll tests were conducted in water and wind tunnels in an effort to investigate the mechanisms of wing rock on a NASP-type vehicle. The configuration tested consisted of a highly-slender forebody and a 78 deg swept delta wing. In the water tunnel test, extensive flow visualization was performed and roll angle histories were obtained. In the wind tunnel test, the roll angle, forces and moments, and limited forebody and wing surface pressures were measured during the wing rock motion. A limit cycle oscillation was observed for angles of attack between 22 deg and 30 deg. In general, the experiments confirmed that the main flow phenomena responsible for the wing-body-tail wing rock are the interactions between the forebody and the wing vortices. The variation of roll acceleration (determined from the second derivative of the roll angle time history) with roll angle clearly slowed the energy balance necessary to sustain the limit cycle oscillation. Different means of suppressing wing rock by controlling the forebody vortices using small blowing jets were also explored. Steady blowing was found to be capable of suppressing wing rock, but significant vortex asymmetrices are created, causing the model to stop at a non-zero roll angle. On the other hand, alternating pulsed blowing on the left and right sides of the fore body was demonstrated to be a potentially effective means of suppressing wing rock and eliminating large asymmetric moments at high angles of attack.

An experimental study of the aerodynamic characteristics of planar and non-planar outboard wing planforms

NASA Technical Reports Server (NTRS)

Naik, D. A.; Ostowari, C.

1987-01-01

A series of wind tunnel experiments have been conducted to investigate the aerodynamic characteristics of several planar and nonplanar wingtip planforms. Seven different configurations: base-line rectangular, elliptical, swept and tapered, swept and tapered with dihedral, swept and tapered with anhedral, rising arc, and drooping arc, were investigated for two different spans. The data are available in terms of coefficient plots of force data, flow visualization photographs, and velocity and pressure flowfield surveys. All planforms, particularly the nonplanar, have some advantages over the baseline rectangular planform. Span efficiencies up to 20-percent greater than baseline are a possibility. However, it is suggested that the span efficiency concept might need refinement for nonplanar wings. Flow survey data show the change in effective span with vortex roll-up. The flow visualization shows the occurrence of mushroom-cell-separation flow patterns at angles of attack corresponding to stall. These grow with an increase in post-stall angle of attack. For the larger aspect ratios, the cells are observed to split into sub-cells at the higher angles of attack. For all angles of attack, some amount of secondary vortex flow is observed for the planar and nonplanar out-board planforms with sweep and taper.

A Preliminary Study of the Response of Single Rotor Helicopters to Vortex Encounters

DOT National Transportation Integrated Search

1985-04-01

This report examines some aspects of the uncontrolled dynamic response of a single rotor helicopter to an encounter with the wing tip vortex of a large transport aircraft. The primary emphasis in the study was to investigate the importance of various...

The leading-edge vortex of yacht sails

NASA Astrophysics Data System (ADS)

Arredondo-Galeana, Abel; Viola, Ignazio Maria

2017-11-01

We experimentally show, for the first time, that a stable Leading-Edge Vortex (LEV) can be formed on an asymmetric spinnaker, which is a high-lift sail used by yachts to sail downwind. We tested a 3D printed rigid sail in a water flume at a chord-based Reynolds number of ca. 104. We found that on the leeward side of the sail (the suction side), the flow separates at the leading edge reattaching further downstream and forming a stable LEV. The LEV grows in diameter from the root to the tip of the sail, where it merges with the tip vortex. We detected the LEV using the γ criterion, and we verified its stability over time. The lift contribution provided by the LEV was computed solving a complex potential model of each sail section. This analysis indicated that the LEV provides a substantial contribution to the total sail's lift. These findings suggest that the maximum lift of low-aspect-ratio wings with a sharp leading edge, such as spinnakers, can be enhanced by promoting a stable LEV. This work was funded by the Consejo Nacional de Ciencia y Tecnologia (CONACYT).

Flow Modulation and Force Control of Flapping Wings

DTIC Science & Technology

2014-10-29

evolution of which reflect the wing morphology and kinematics. While the near-wake vortex system directly reflects the action of the wing on the...at 8 different stroke positions, which demonstrate the evolution of the vortex wake structure. The contour plot of Z vorticity at X-Y plane (Z...20 Figure 14. Smoke patterns showing the evolution of the flow structure in an

Flow structures around a beetle in a tethered flight

NASA Astrophysics Data System (ADS)

Lee, Boogeon; Oh, Sehyeong; Park, Hyungmin; Choi, Haecheon

2017-11-01

In the present study, through a wind-tunnel experiment, we visualize the flow in a tethered flight of a rhinoceros beetle using a smoke-wire visualization technique. Measurements are done at five side planes along the wind span while varying the body angle (angle between the horizontal and the body axis) to investigate the influence of the stroke plane angle that was observed to change depending on the flight mode such as hovering, forward and takeoff flights so on. Observing that a large attached leading-edge vortex is only found on the hindwing, it is inferred that most of the aerodynamic forces would be generated by hindwings (flexible inner wings) compared to the elytra (hard outer wings). In addition, it is observed to use unsteady lift-generating mechanisms such as clap-and-fling, wing-wing interaction and wake capture. Finally, we discuss the relation between the advance ratio and Strouhal number by adjusting free-stream velocity and the body angle (i.e., angle of wake-induced flow). Supported by a Grant to Bio-Mimetic Robot Research Center Funded by Defense Acquisition Program Administration, and by ADD, Korea (UD130070ID).

Some observations of tip-vortex cavitation

NASA Astrophysics Data System (ADS)

Arndt, R. E. A.; Arakeri, V. H.; Higuchi, H.

1991-08-01

Cavitation has been observed in the trailing vortex system of an elliptic platform hydrofoil. A complex dependence on Reynolds number and gas content is noted at inception. Some of the observations can be related to tension effects associated with the lack of sufficiently large-sized nuclei. Inception measurements are compared with estimates of pressure in the vortex obtained from LDV measurements of velocity within the vortex. It is concluded that a complete correlation is not possible without knowledge of the fluctuating levels of pressure in tip-vortex flows. When cavitation is fully developed, the observed tip-vortex trajectory flows. When cavitation is fully developed, the observed tip-vortex trajectory shows a surprising lack of dependence on any of the physical parameters varied, such as angle of attack, Reynolds number, cavitation number, and dissolved gas content.

Design and wind tunnel tests of winglets on a DC-10 wing

NASA Technical Reports Server (NTRS)

Gilkey, R. D.

1979-01-01

Results are presented of a wind tunnel test utilizing a 4.7 percent scale semi-span model in the Langley Research Center 8-foot transonic pressure wind tunnel to establish the cruise drag improvement potential of winglets as applied to the DC-10 wide body transport aircraft. Winglets were investigated on both the DC-10 Series 10 (domestic) and 30/40 (intercontinental) configurations and compared with the Series 30/40 configuration. The results of the investigation confirm that for the DC-10 winglets provide approximately twice the cruise drag reduction of wing-tip extensions for about the same increase in bending moment at the wing fuselage juncture. Furthermore, the winglet configurations achieved drag improvements which were in close agreement to analytical estimates. It was observed that relatively small changes in wing-winglet tailoring effected large improvements in drag and visual flow characteristics. All final winglet configurations exhibited visual flow characteristics on the wing and winglets

Calculation of vortex lift effect for cambered wings by the suction analogy

NASA Technical Reports Server (NTRS)

Lan, C. E.; Chang, J. F.

1981-01-01

An improved version of Woodward's chord plane aerodynamic panel method for subsonic and supersonic flow is developed for cambered wings exhibiting edge separated vortex flow, including those with leading edge vortex flaps. The exact relation between leading edge thrust and suction force in potential flow is derived. Instead of assuming the rotated suction force to be normal to wing surface at the leading edge, new orientation for the rotated suction force is determined through consideration of the momentum principle. The supersonic suction analogy method is improved by using an effective angle of attack defined through a semi-empirical method. Comparisons of predicted results with available data in subsonic and supersonic flow are presented.

Tip leakage vortex dynamics and inception

NASA Astrophysics Data System (ADS)

Oweis, Ghanem; Ceccio, Steven; Jessup, Stuart; Chesnakas, Christopher; Fry, David

2002-11-01

The McCormick rule for tip vortex cavitation scaling predicts that cavitation should take place in the vortex where the average core pressure deficit from the free stream is the largest along the vortex tube. The average core pressure deficit can be calculated from the vortex core size and circulation and these can be measured by LDV or hot wire, among other methods. The same rule applies to the tip vortex from a wall-bounded hydrofoil. Recent cavitation inception experiments on a ducted propeller in the NSWCCD 36 inch water tunnel combined with PIV and LDV measurements of the tip vortex flow are described. These tests reveal a disagreement between the actual inception location and that predicted by the McCormick rule. It is hypothesized that in this case the inception mechanism is related to local flow phenomena associated with local vortex unsteadiness, as opposed to the average vortex parameters (core size and circulation) used in the viscous scaling rule of McCormick. Discussion of the flow field measurements, bubble population, and the noise production from the inception events is given.

Effect of tip flange on tip leakage flow of small axial flow fans

NASA Astrophysics Data System (ADS)

Zhang, Li; Jin, Yingzi; Jin, Yuzhen

2014-02-01

Aerodynamic performance of an axial flow fan is closely related to its tip clearance leakage flow. In this paper, the hot-wire anemometer is used to measure the three dimensional mean velocity near the blade tips. Moreover, the filtered N-S equations with finite volume method and RNG k-ɛ turbulence model are adopted to carry out the steady simulation calculation of several fans that differ only in tip flange shape and number. The large eddy simulation and the FW-H noise models are adopted to carry out the unsteady numerical calculation and aerodynamic noise prediction. The results of simulation calculation agree roughly with that of tests, which proves the numerical calculation method is feasible.The effects of tip flange shapes and numbers on the blade tip vortex structure and the characteristics are analyzed. The results show that tip flange of the fan has a certain influence on the characteristics of the fan. The maximum efficiencies for the fans with tip flanges are shifted towards partial flow with respect to the design point of the datum fan. Furthermore, the noise characteristics for the fans with tip flanges have become more deteriorated than that for the datum fan. Tip flange contributes to forming tip vortex shedding and the effect of the half-cylinder tip flange on tip vortex shedding is obvious. There is a distinct relationship between the characteristics of the fan and tip vortex shedding. The research results provide the profitable reference for the internal flow mechanism of the performance optimization of small axial flow fans.

Water tunnel flow visualization study of a 4.4 percent scale X-31 forebody

NASA Technical Reports Server (NTRS)

Cobleigh, Brent R.; Delfrate, John

1994-01-01

A water-tunnel test of a 4.4 percent-scale, forebody-only model of the X-31 aircraft with different forebody strakes and nosebooms has been performed in the Flow Visualization Facility at the NASA Dryden Flight Research Center. The focus of the study was to determine the relative effects of the different configurations on the stability and symmetry of the high-angle-of-attack forebody vortex flow field. The clean, noseboom-off configuration resisted the development of asymmetries in the primary vortices through 70 deg angle of attack. The wake of the X-31 flight test noseboom configuration significantly degraded the steadiness of the primary vortex cores and promoted asymmetries. An alternate L-shaped noseboom mounted underneath the forebody had results similar to those seen with the configuration, enabling stable, symmetrical vortices up to 70 deg angle of attack. The addition of strakes near the radome tip along the waterline increased the primary vortex strength while it simultaneously caused the vortex breakdown location co move forward. Forebody strakes did not appear to significantly reduce the asymmetries in the forebody vortex field in the presence of the flight test noseboom.

Trailing Vortex Measurements in the Wake of a Hovering Rotor Blade with Various Tip Shapes

NASA Technical Reports Server (NTRS)

Martin, Preston B.; Leishman, J. Gordon

2003-01-01

This work examined the wake aerodynamics of a single helicopter rotor blade with several tip shapes operating on a hover test stand. Velocity field measurements were conducted using three-component laser Doppler velocimetry (LDV). The objective of these measurements was to document the vortex velocity profiles and then extract the core properties, such as the core radius, peak swirl velocity, and axial velocity. The measured test cases covered a wide range of wake-ages and several tip shapes, including rectangular, tapered, swept, and a subwing tip. One of the primary differences shown by the change in tip shape was the wake geometry. The effect of blade taper reduced the initial peak swirl velocity by a significant fraction. It appears that this is accomplished by decreasing the vortex strength for a given blade loading. The subwing measurements showed that the interaction and merging of the subwing and primary vortices created a less coherent vortical structure. A source of vortex core instability is shown to be the ratio of the peak swirl velocity to the axial velocity deficit. The results show that if there is a turbulence producing region of the vortex structure, it will be outside of the core boundary. The LDV measurements were supported by laser light-sheet flow visualization. The results provide several benchmark test cases for future validation of theoretical vortex models, numerical free-wake models, and computational fluid dynamics results.

Experimental Investigation of the Unsteady Flow Structures of Two Interacting Pitching Wings

NASA Astrophysics Data System (ADS)

Kurt, Melike; Moored, Keith

2015-11-01

Birds, insects and fish propel themselves with unsteady motions of their wings and fins. Many of these animals are also found to fly or swim in three-dimensional flocks and schools. Numerous studies have explored the three-dimensional steady flow interactions and the two-dimensional unsteady flow interactions in collectives. Yet, the characterization of the three-dimensional unsteady interactions remains relatively unexplored. This study aims to characterize the flow structures and interactions between two sinusoidally pitching finite-span wings. The arrangement of the wings varies from a tandem to a bi-plane configuration. The vortex structures for these various arrangements are quantified by using particle image velocimetry. The vortex-wing interactions are also characterized as the synchrony between the wings is modified.

Flight validation of a pulsed smoke flow visualization system

NASA Technical Reports Server (NTRS)

Ward, Donald T.; Dorsett, Kenneth M.

1993-01-01

A flow visualization scheme, designed to measure vortex fluid dynamics on research aircraft, was validated in flight. Strake vortex trajectories and axial core velocities were determined using pulsed smoke, high-speed video images, and semiautomated image edge detection hardware and software. Smoke was pulsed by using a fast-acting three-way valve. After being redesigned because of repeatedly jamming in flight, the valve shuttle operated flawlessly during the last two tests. A 25-percent scale, Gothic strake was used to generate vortex over the wing of a GA-7 Cougar and was operated at a local angle of attack of 22 degrees and Reynolds number of approximately 7.8 x 10(exp 5)/ft. Maximum axial velocities measured in the vortex core were between 1.75 and 1.95 times the freestream velocity. Analysis of the pulsed smoke system's affect on forebody vortices indicates that the system may reorient the forebody vortex system; however, blowing momentum coefficients normally used will have no appreciable affect on the leading-edge extension vortex system. It is recommended that a similar pulsed smoke system be installed on the F/A-18 High Angle Research Vehicle and that this approach be used to analyze vortex core dynamics during the remainder of its high-angle-of-attack research flights.

Study of Flapping Flight Using Discrete Vortex Method Based Simulations

NASA Astrophysics Data System (ADS)

Devranjan, S.; Jalikop, Shreyas V.; Sreenivas, K. R.

2013-12-01

In recent times, research in the area of flapping flight has attracted renewed interest with an endeavor to use this mechanism in Micro Air vehicles (MAVs). For a sustained and high-endurance flight, having larger payload carrying capacity we need to identify a simple and efficient flapping-kinematics. In this paper, we have used flow visualizations and Discrete Vortex Method (DVM) based simulations for the study of flapping flight. Our results highlight that simple flapping kinematics with down-stroke period (tD) shorter than the upstroke period (tU) would produce a sustained lift. We have identified optimal asymmetry ratio (Ar = tD/tU), for which flapping-wings will produce maximum lift and find that introducing optimal wing flexibility will further enhances the lift.

A qualitative study of vortex trapping capability for lift enhancement on unconventional wing

NASA Astrophysics Data System (ADS)

Salleh, M. B.; Kamaruddin, N. M.; Mohamed-Kassim, Z.

2018-05-01

Lift enhancement by using passive vortex trapping technique offers great advantage in small aircraft design as it can improve aerodynamics performance and reduce weight of the wing. To achieve this aim, a qualitative study on the flow structures across wing models with cavities has been performed using smoke wire visualisation technique. An experiment has been conducted at low Reynolds number of 26,000 with angle of attack (α) = 0°, 5°, 10° and 15° to investigate the vortex trapping capability of semi-circular leading edge (SCLE) flat-plate wing model and elliptical leading edge (ELE) flat-plate wing model with cavities, respectively. Results from the qualitative study indicated unique characteristics in the flow structures between the tested wing models. The SCLE wing models were able to trap stable rotating vortices for α ≤ 10° whereas the ability of ELE wing models to suppress flow separation allowed stable clockwise vortices to be trapped inside the cavities even at α > 10°. The trapped vortices found to have the potential to increase lift on the unconventional wing models.

Finite-span rotating wings: three-dimensional vortex formation and variations with aspect ratio

NASA Astrophysics Data System (ADS)

Carr, Z. R.; Chen, C.; Ringuette, M. J.

2013-02-01

We investigate experimentally the effect of aspect ratio ( [InlineMediaObject not available: see fulltext.] ) on the time-varying, three-dimensional flow structure of flat-plate wings rotating from rest at 45° angle of attack. Plates of [InlineMediaObject not available: see fulltext.] = 2 and 4 are tested in a 50 % by mass glycerin-water mixture, with a total rotation of ϕ = 120° and a matched tip Reynolds number of 5,000. The time-varying, three-component volumetric velocity field is reconstructed using phase-locked, phase-averaged stereoscopic digital particle image velocimetry in multiple, closely-spaced chordwise planes. The vortex structure is analyzed using the {Q}-criterion, helicity density, and spanwise quantities. For both [InlineMediaObject not available: see fulltext.] s, the flow initially consists of a connected and coherent leading-edge vortex (LEV), tip vortex (TV), and trailing-edge vortex (TEV) loop; the LEV increases in size with span and tilts aft. Smaller, discrete vortices are present in the separated shear layers at the trailing and tip edges, which wrap around the primary TEV and TV. After about ϕ = 20°, the outboard-span LEV lifts off the plate and becomes arch-like. A second, smaller LEV and the formation of corner vortex structures follow. For [InlineMediaObject not available: see fulltext.] = 4, the outboard LEV moves farther aft, multiple LEVs form ahead of it, and after about ϕ = 50° a breakdown of the lifted-off LEV and the TV occurs. However, for [InlineMediaObject not available: see fulltext.] = 2, the outboard LEV lift-off is not progressive, and the overall LEV-TV flow remains more coherent and closer to the plate, with evidence of breakdown late in the motion. Inboard of about 50 % span, the [InlineMediaObject not available: see fulltext.] = 4 LEV is stable for the motion duration. Up to approximately 60 % span, the [InlineMediaObject not available: see fulltext.] = 2 LEV is distinct from the TV and is similarly stable. The [InlineMediaObject not available: see fulltext.] = 2 LEV exhibits substantially higher spanwise vorticity and velocity. The latter possesses a "four-lobed" distribution at the periphery of the LEV core having adjacent positive (outboard) and negative (inboard) components, corresponding to a helical streamline structure. Both [InlineMediaObject not available: see fulltext.] s show substantial root-to-tip velocity aft of the stable LEV, which drives outboard spanwise vorticity flux; flux toward the root is also present in the front portion of the LEV. For [InlineMediaObject not available: see fulltext.] = 2, there is a strong flux of spanwise vorticity from the outboard LEV to the tip, which may mitigate LEV lift-off and is not found for [InlineMediaObject not available: see fulltext.] = 4. The TV circulation for each [InlineMediaObject not available: see fulltext.] is similar in magnitude and growth when plotted versus the chord lengths travelled by the tip, prior to breakdown. Streamwise vorticity due to the TV induces high spanwise velocity, and for [InlineMediaObject not available: see fulltext.] = 2, the tilted LEV creates further streamwise vorticity which corresponds well to spanwise-elongated regions of spanwise velocity. For [InlineMediaObject not available: see fulltext.] = 2, the TV influences a relatively greater portion of the span and is more coherent at later times, which coupled with the tilted LEV strongly contributes to the higher overall spanwise velocity and vorticity flux.

Effect of Thickness-to-Chord Ratio on Flow Structure of Low Swept Delta Wing

NASA Astrophysics Data System (ADS)

Gulsacan, Burak; Sencan, Gizem; Yavuz, Mehmet Metin

2017-11-01

The effect of thickness-to-chord (t/C) ratio on flow structure of a delta wing with sweep angle of 35 degree is characterized in a low speed wind tunnel using laser illuminated smoke visualization, particle image velocimetry, and surface pressure measurements. Four different t/C ratio varying from 4.75% to 19% are tested at angles of attack 4, 6, 8, and 10 degrees for Reynolds numbers Re =10,000 and 35,000. The results indicate that the effect of thickness-to-chord ratio on flow structure is quite substantial, such that, as the wing thickness increases, the flow structure transforms from leading edge vortex to three-dimensional separated flow regime. The wing with low t/C ratio of 4.75% experiences pronounced surface separation at significantly higher angle of attack compared to the wing with high t/C ratio. The results might explain some of the discrepancies reported in previously conducted studies related to delta wings. In addition, it is observed that the thickness of the shear layer separated from windward side of the wing is directly correlated with the thickness of the wing. To conclude, the flow structure on low swept delta wing is highly affected by t/C ratio, which in turn might indicate the potential usage of wing thickness as an effective flow control parameter.

Experimental and computational investigation of the tip clearance flow in a transonic axial compressor rotor

NASA Astrophysics Data System (ADS)

Suder, Kenneth L.; Celestina, Mark L.

1995-06-01

Experimental and computational techniques are used to investigate tip clearance flows in a transonic axial compressor rotor at design and part speed conditions. Laser anemometer data acquired in the endwall region are presented for operating conditions near peak efficiency and near stall at 100% design speed and at near peak efficiency at 60% design speed. The role of the passage shock/leakage vortex interaction in generating endwall blockage is discussed. As a result of the shock/vortex interaction at design speed, the radial influence of the tip clearance flow extends to 20 times the physical tip clearance height. At part speed, in the absence of the shock, the radial extent is only 5 times the tip clearance height. Both measurements and analysis indicate that under part-speed operating conditions a second vortex, which does not originate from the tip leakage flow, forms in the endwall region within the blade passage and exits the passage near midpitch. Mixing of the leakage vortex with primary flow downstream of the rotor at both design and part speed conditions is also discussed.

Experimental and Computational Investigation of the Tip Clearance Flow in a Transonic Axial Compressor Rotor

NASA Technical Reports Server (NTRS)

Suder, Kenneth L.; Celestina, Mark L.

1995-01-01

Experimental and computational techniques are used to investigate tip clearance flows in a transonic axial compressor rotor at design and part speed conditions. Laser anemometer data acquired in the endwall region are presented for operating conditions near peak efficiency and near stall at 100% design speed and at near peak efficiency at 60% design speed. The role of the passage shock/leakage vortex interaction in generating endwall blockage is discussed. As a result of the shock/vortex interaction at design speed, the radial influence of the tip clearance flow extends to 20 times the physical tip clearance height. At part speed, in the absence of the shock, the radial extent is only 5 times the tip clearance height. Both measurements and analysis indicate that under part-speed operating conditions a second vortex, which does not originate from the tip leakage flow, forms in the endwall region within the blade passage and exits the passage near midpitch. Mixing of the leakage vortex with primary flow downstream of the rotor at both design and part speed conditions is also discussed.

Flow Behavior in Side-View Plane of Pitching Delta Wing

NASA Astrophysics Data System (ADS)

Pektas, Mehmet Can; Tasci, Mehmet Oguz; Karasu, Ilyas; Sahin, Besir; Akilli, Huseyin

2018-06-01

In the present investigation, a delta wing which has 70° sweep angle, Λ was oscillated on its midcord according to the equation of α(t)=αm+α0sin(ωet). This study focused on understanding the effect of pitching and characterizing the interaction of vortex breakdown with oscillating leading edges under different yaw angles, β over a slender delta wing. The value of mean angle of attack, αm was taken as 25°. The yaw angle, β was varied with an interval of 4° over the range of 0°≤β≤ 16°. The delta wing was sinusoidally pitched within the range of period of time 5s≤Te≤60s and reduced frequency was set as K=0.16, 0.25, 0.49, 1.96 and lastly amplitude of pitching motion was arranged as α0=±5°.Formations and locations of vortex breakdown were investigated by using the dye visualization technique in side view plane.

Stereo-Video Data Reduction of Wake Vortices and Trailing Aircrafts

NASA Technical Reports Server (NTRS)

Alter-Gartenberg, Rachel

1998-01-01

This report presents stereo image theory and the corresponding image processing software developed to analyze stereo imaging data acquired for the wake-vortex hazard flight experiment conducted at NASA Langley Research Center. In this experiment, a leading Lockheed C-130 was equipped with wing-tip smokers to visualize its wing vortices, while a trailing Boeing 737 flew into the wake vortices of the leading airplane. A Rockwell OV-10A airplane, fitted with video cameras under its wings, flew at 400 to 1000 feet above and parallel to the wakes, and photographed the wake interception process for the purpose of determining the three-dimensional location of the trailing aircraft relative to the wake. The report establishes the image-processing tools developed to analyze the video flight-test data, identifies sources of potential inaccuracies, and assesses the quality of the resultant set of stereo data reduction.

Vortex maneuver lift for super-cruise configurations

NASA Technical Reports Server (NTRS)

Campbell, J. F.; Gloss, B. B.; Lamar, J. E.

1976-01-01

Some of the theoretical and experimental research conducted at the NASA Langley Research Center is presented to investigate the subsonic vortex-lift producing capabilities for two classes of Super-Cruise designs: a close-coupled wing-canard arrangement and a slender wing configuration. In addition, several analytical methods are discussed for estimating critical structural design loads for thin, highly swept wings having separated leading-edge vortex flows.

The Vortex Lattice Method for the Rotor-Vortex Interaction Problem

NASA Technical Reports Server (NTRS)

Padakannaya, R.

1974-01-01

The rotor blade-vortex interaction problem and the resulting impulsive airloads which generate undesirable noise levels are discussed. A numerical lifting surface method to predict unsteady aerodynamic forces induced on a finite aspect ratio rectangular wing by a straight, free vortex placed at an arbitrary angle in a subsonic incompressible free stream is developed first. Using a rigid wake assumption, the wake vortices are assumed to move downsteam with the free steam velocity. Unsteady load distributions are obtained which compare favorably with the results of planar lifting surface theory. The vortex lattice method has been extended to a single bladed rotor operating at high advance ratios and encountering a free vortex from a fixed wing upstream of the rotor. The predicted unsteady load distributions on the model rotor blade are generally in agreement with the experimental results. This method has also been extended to full scale rotor flight cases in which vortex induced loads near the tip of a rotor blade were indicated. In both the model and the full scale rotor blade airload calculations a flat planar wake was assumed which is a good approximation at large advance ratios because the downwash is small in comparison to the free stream at large advance ratios. The large fluctuations in the measured airloads near the tip of the rotor blade on the advance side is predicted closely by the vortex lattice method.

Numerical investigation of tip clearance cavitation in Kaplan runners

NASA Astrophysics Data System (ADS)

Nikiforova, K.; Semenov, G.; Kuznetsov, I.; Spiridonov, E.

2016-11-01

There is a gap between the Kaplan runner blade and the shroud that makes for a special kind of cavitation: cavitation in the tip leakage flow. Two types of cavitation caused by the presence of clearance gap are known: tip vortex cavitation that appears at the core of the rolled up vortex on the blade suction side and tip clearance cavitation that appears precisely in the gap between the blade tip edge and the shroud. In the context of this work numerical investigation of the model Kaplan runner has been performed taking into account variable tip clearance for several cavitation regimes. The focus is put on investigation of structure and origination of mechanism of cavitation in the tip leakage flow. Calculations have been performed with the help of 3-D unsteady numerical model for two-phase medium. Modeling of turbulent flow in this work has been carried out using full equations of Navier-Stokes averaged by Reynolds with correction for streamline curvature and system rotation. For description of this medium (liquid-vapor) simplification of Euler approach is used; it is based on the model of interpenetrating continuums, within the bounds of this two- phase medium considered as a quasi-homogeneous mixture with the common velocity field and continuous distribution of density for both phases. As a result, engineering techniques for calculation of cavitation conditioned by existence of tip clearance in model turbine runner have been developed. The detailed visualization of the flow was carried out and vortex structure on the suction side of the blade was reproduced. The range of frequency with maximum value of pulsation was assigned and maximum energy frequency was defined; it is based on spectral analysis of the obtained data. Comparison between numerical computation results and experimental data has been also performed. The location of cavitation zone has a good agreement with experiment for all analyzed regimes.

An experimental study of pressures on 60 deg Delta wings with leading edge vortex flaps

NASA Technical Reports Server (NTRS)

Marchman, J. F., III; Terry, J. E.; Donatelli, D. A.

1983-01-01

An experimental study was conducted in the Virginia Tech Stability Wind Tunnel to determine surface pressures over a 60 deg sweep delta wing with three vortex flap designs. Extensive pressure data was collected to provide a base data set for comparison with computational design codes and to allow a better understanding of the flow over vortex flaps. The results indicated that vortex flaps can be designed which will contain the leading edge vortex with no spillage onto the wing upper surface. However, the tests also showed that flaps designed without accounting for flap thickness will not be optimum and the result can be oversized flaps, early flap vortex reattachment and a second separation and vortex at the wing/flap hinge line.

Load distribution on a close-coupled wing canard at transonic speeds

NASA Technical Reports Server (NTRS)

Gloss, B. B.; Washburn, K. E.

1977-01-01

This paper reports on a wind-tunnel test where load distributions were obtained at transonic speeds on both the canard and wing surfaces of a closely-coupled wing-canard configuration. The investigation included detailed component and configuration arrangement studies to provide insight into the various aerodynamic interference effects for the leading-edge vortex flow conditions encountered. Data indicate that increasing the Mach number from 0.70 to 0.95 caused the wing leading-edge vortex to burst over the wing when the wing was in the presence of the high canard. For some of the outboard span locations, the leading-edge vortex reattachment streamline intersects the wing trailing edge inboard of these span locations, thus, the Kutta condition was not satisfied. In general, the effect of adding a canard was to reduce the lift inboard and somewhat increase the lift outboard similar to the trends that would have been expected had the flow been attached.

Aerodynamics and vortical structures in hovering fruitflies

NASA Astrophysics Data System (ADS)

Meng, Xue Guang; Sun, Mao

2015-03-01

We measure the wing kinematics and morphological parameters of seven freely hovering fruitflies and numerically compute the flows of the flapping wings. The computed mean lift approximately equals to the measured weight and the mean horizontal force is approximately zero, validating the computational model. Because of the very small relative velocity of the wing, the mean lift coefficient required to support the weight is rather large, around 1.8, and the Reynolds number of the wing is low, around 100. How such a large lift is produced at such a low Reynolds number is explained by combining the wing motion data, the computed vortical structures, and the theory of vorticity dynamics. It has been shown that two unsteady mechanisms are responsible for the high lift. One is referred as to "fast pitching-up rotation": at the start of an up- or downstroke when the wing has very small speed, it fast pitches down to a small angle of attack, and then, when its speed is higher, it fast pitches up to the angle it normally uses. When the wing pitches up while moving forward, large vorticity is produced and sheds at the trailing edge, and vorticity of opposite sign is produced near the leading edge and on the upper surface, resulting in a large time rate of change of the first moment of vorticity (or fluid impulse), hence a large aerodynamic force. The other is the well known "delayed stall" mechanism: in the mid-portion of the up- or downstroke the wing moves at large angle of attack (about 45 deg) and the leading-edge-vortex (LEV) moves with the wing; thus, the vortex ring, formed by the LEV, the tip vortices, and the starting vortex, expands in size continuously, producing a large time rate of change of fluid impulse or a large aerodynamic force.

Problem of Vortex Turbulence behind Wings (II),

DTIC Science & Technology

1980-09-23

these winglets would give a resultant aerodynamic force directed towards the front which would decrease the wing drag. Such winglets will affect the...Fig. 30 Whitcomb winglets Pig. 31 Set of winglets for wake dissipation Surfaces on wing tips, winglets (Fig. 30), proposed by Whitcomb to diminish...anyway - to decrease the induced drag of the wing by putting some winglets at a certain angle in different planes, as shown in Fig. 31. The total

Aeroelastic loads prediction for an arrow wing. Task 3: Evaluation of the Boeing three-dimensional leading-edge vortex code

NASA Technical Reports Server (NTRS)

Manro, M. E.

1983-01-01

Two separated flow computer programs and a semiempirical method for incorporating the experimentally measured separated flow effects into a linear aeroelastic analysis were evaluated. The three dimensional leading edge vortex (LEV) code is evaluated. This code is an improved panel method for three dimensional inviscid flow over a wing with leading edge vortex separation. The governing equations are the linear flow differential equation with nonlinear boundary conditions. The solution is iterative; the position as well as the strength of the vortex is determined. Cases for both full and partial span vortices were executed. The predicted pressures are good and adequately reflect changes in configuration.

Turbulence Measurements in the Near Field of a Wingtip Vortex

NASA Technical Reports Server (NTRS)

Chow, Jim; Zilliac, Greg; Bradshaw, Peter

1997-01-01

The roll-up of a wingtip vortex, at Reynolds number based on chord of 4.6 million was studied with an emphasis on suction side and near wake measurements. The research was conducted in a 32 in. x 48 in. low-speed wind tunnel. The half-wing model had a semi-span of 36 in. a chord of 48 in. and a rounded tip. Seven-hole pressure probe measurements of the velocity field surrounding the wingtip showed that a large axial velocity of up to 1.77 U(sub infinity) developed in the vortex core. This level of axial velocity has not been previously measured. Triple-wire probes have been used to measure all components of the Reynolds stress tensor. It was determined from correlation measurements that meandering of the vortex was small and did not appreciably contribute to the turbulence measurements. The flow was found to be turbulent in the near-field (as high as 24 percent RMS w - velocity on the edge of the core) and the turbulence decayed quickly with streamwise distance because of the nearly solid body rotation of the vortex core mean flow. A streamwise variation of the location of peak levels of turbulence, relative to the core centerline, was also found. Close to the trailing edge of the wing, the peak shear stress levels were found at the edge of the vortex core, whereas in the most downstream wake planes they occurred at a radius roughly equal to one-third of the vortex core radius. The Reynolds shear stresses were not aligned with the mean strain rate, indicating that an isotropic-eddy-viscosity based prediction method cannot accurately model the turbulence in the cortex. In cylindrical coordinates, with the origin at the vortex centerline, the radial normal stress was found to be larger than the circumferential.

Wingtip vortex turbine

NASA Technical Reports Server (NTRS)

Patterson, James C., Jr. (Inventor)

1990-01-01

A means for extracting rotational energy from the vortex created at aircraft wing tips which consists of a turbine with blades located in the crossflow of the vortex and attached downstream of the wingtip. The turbine has blades attached to a core. When the aircraft is in motion, rotation of a core transmits energy to a centrally attached shaft. The rotational energy thus generated may be put to use within the airfoil or aircraft fuselage.

Velocity field measurements in the wake of a propeller model

NASA Astrophysics Data System (ADS)

Mukund, R.; Kumar, A. Chandan

2016-10-01

Turboprop configurations are being revisited for the modern-day regional transport aircrafts for their fuel efficiency. The use of laminar flow wings is an effort in this direction. One way to further improve their efficiency is by optimizing the flow over the wing in the propeller wake. Previous studies have focused on improving the gross aerodynamic characteristics of the wing. It is known that the propeller slipstream causes early transition of the boundary layer on the wing. However, an optimized design of the propeller and wing combination could delay this transition and decrease the skin friction drag. Such a wing design would require the detailed knowledge of the development of the slipstream in isolated conditions. There are very few studies in the literature addressing the requirements of transport aircraft having six-bladed propeller and cruising at a high propeller advance ratio. Low-speed wind tunnel experiments have been conducted on a powered propeller model in isolated conditions, measuring the velocity field in the vertical plane behind the propeller using two-component hot-wire anemometry. The data obtained clearly resolved the mean velocity, the turbulence, the ensemble phase averages and the structure and development of the tip vortex. The turbulence in the slipstream showed that transition could be close to the leading edge of the wing, making it a fine case for optimization. The development of the wake with distance shows some interesting flow features, and the data are valuable for flow computation and optimization.

A study of the rotor wake of a small-scale rotor model in forward flight using laser light sheet flow visualization with comparisons to analytical models

NASA Technical Reports Server (NTRS)

Ghee, Terence A.; Elliott, Joe W.

1992-01-01

An experimental investigation was conducted in the 14 by 22 ft subsonic tunnel at NASA Langley Research Center to quantify the rotor wake behind a scale model helicopter rotor in forward flight (mu = 0.15 and 0.23) at one thrust level (C sub T = 0.0064). The rotor system used in the present test consisted of a four-bladed, fully articulated hub and utilized blades of rectangular planform with a NACA-0012 airfoil section. A laser light sheet, seeded with propylene glycol smoke, was used to visualize the flow in planes parallel and perpendicular to the freestream flow. Quantitative measurements of vortex location, vertical skew angle, and vortex particle void radius were obtained for vortices in the flow; convective velocities were obtained for blade tip vortices. Comparisons were made between the experimental results and the wake geometry generated by computational predictions. The results of these comparisons show that the interaction between wake vortex structures is an important consideration for correctly predicting the wake geometry.

Influence of Initial Vorticity Distribution on Axisymmetric Vortex Breakdown and Reconnection

NASA Technical Reports Server (NTRS)

Young, Larry A.

2007-01-01

An analytical treatment has been developed to study some of the axisymmetric vortex breakdown and reconnection fluid dynamic processes underlying body-vortex interactions that are frequently manifested in rotorcraft and propeller-driven fixed-wing aircraft wakes. In particular, the presence of negative vorticity in the inner core of a vortex filament (one example of which is examined in this paper) subsequent to "cutting" by a solid body has a profound influence on the vortex reconnection, leading to analog flow behavior similar to vortex breakdown phenomena described in the literature. Initial vorticity distributions (three specific examples which are examined) without an inner core of negative vorticity do not exhibit vortex breakdown and instead manifest diffusion-like properties while undergoing vortex reconnection. Though this work focuses on laminar vortical flow, this work is anticipated to provide valuable insight into rotary-wing aerodynamics as well as other types of vortical flow phenomena.

Numerical simulation of incidence and sweep effects on delta wing vortex breakdown

NASA Technical Reports Server (NTRS)

Ekaterinaris, J. A.; Schiff, Lewis B.

1994-01-01

The structure of the vortical flowfield over delta wings at high angles of attack was investigated. Three-dimensional Navier-Stokes numerical simulations were carried out to predict the complex leeward-side flowfield characteristics, including leading-edge separation, secondary separation, and vortex breakdown. Flows over a 75- and a 63-deg sweep delta wing with sharp leading edges were investigated and compared with available experimental data. The effect of variation of circumferential grid resolution grid resolution in the vicinity of the wing leading edge on the accuracy of the solutions was addressed. Furthermore, the effect of turbulence modeling on the solutions was investigated. The effects of variation of angle of attack on the computed vortical flow structure for the 75-deg sweep delta wing were examined. At moderate angles of attack no vortex breakdown was observed. When a critical angle of attack was reached, bubble-type vortex breakdown was found. With further increase in angle of attack, a change from bubble-type breakdown to spiral-type vortex breakdown was predicted by the numerical solution. The effects of variation of sweep angle and freestream Mach number were addressed with the solutions on a 63-deg sweep delta wing.

Non-linear unsteady wing theory, part 1. Quasi two-dimensional behavior: Airfoils and slender wings

NASA Technical Reports Server (NTRS)

Mccune, J. E.

1987-01-01

The initial phases of a study of the large-amplitude unsteady aerodynamics of wings in severe maneuver are reported. The research centers on vortex flows, their initiation at wing surfaces, their subsequent convection, and interaction dynamically with wings and control surfaces. The focus is on 2D and quasi-2D aspects of the problem and features the development of an exact nonlinear unsteady airfoil theory as well as an approach to the crossflow problem for slender wing applications including leading-edge separation. The effective use of interactive on-line computing in quantifying and visualizing the nonsteady effects of severe maneuver is demonstrated. Interactive computational work is now possible, in which a maneuver can be initiated and its effects observed and analyzed immediately.

An Investigation of the Effects of Discrete Wing Tip Jets on Wake Vortex Roll Up.

DTIC Science & Technology

1983-08-01

failure of these devices does not mean that the vortex structure cannot be altered such as to reduce rolling moment. On the contrary, Yuan and Bloom (43...which has demonstrated a capabilitv, to e:ra induced rolling moment - the downward blowing jet of .𔄁, ,and Bloom (43)- was also the only jet...eliminated the large vortex excursions associated with close approaches. Bloom and Jen (83) used the method of Kuwahara and Takami to calculate vortex roll up

Aerodynamics of dynamic wing flexion in translating wings

NASA Astrophysics Data System (ADS)

Liu, Yun; Cheng, Bo; Sane, Sanjay P.; Deng, Xinyan

2015-06-01

We conducted a systematic experimental study to investigate the aerodynamic effects of active trailing-edge flexion on a high-aspect-ratio wing translating from rest at a high angle of attack. We varied the timing and speed of the trailing-edge flexion and measured the resulting aerodynamic effects using a combination of direct force measurements and two-dimensional PIV flow measurements. The results indicated that the force and flow characteristics depend strongly on the timing of flexion, but relatively weakly on its speed. This is because the force and vortical flow structure are more sensitive to the timing of flexion relative to the shedding of starting vortex and leading-edge vortex. When the trailing-edge flexion occurred slightly before the starting vortex was shed, the lift production was greatly improved with the instantaneous peak lift increased by 54 % and averaged lift increased by 21 % compared with the pre-flexed case where the trailing-edge flexed before wing translation. However, when the trailing-edge flexed during or slightly after the leading-edge vortex shedding, the lift was significantly reduced by the disturbed development of leading-edge vortex. The force measurement results also imply that the trailing-edge flexion prior to wing translation does not augment lift but increases drag, thus resulting in a lower lift-drag ratio as compared to the case of flat wing.

Tip vortex computer code SRATIP. User's guide

NASA Technical Reports Server (NTRS)

Levy, R.; Lin, S. J.

1985-01-01

This User's Guide applies to the three dimensional viscous flow forward marching analysis, PEPSIG, as used for the calculation of the helicopter tip vortex flow field. The guide presents a discussion of the program flow and subroutines, as well as a list of sample input and output.

Coupled Vortex-Lattice Flight Dynamic Model with Aeroelastic Finite-Element Model of Flexible Wing Transport Aircraft with Variable Camber Continuous Trailing Edge Flap for Drag Reduction

NASA Technical Reports Server (NTRS)

Nguyen, Nhan; Ting, Eric; Nguyen, Daniel; Dao, Tung; Trinh, Khanh

2013-01-01

This paper presents a coupled vortex-lattice flight dynamic model with an aeroelastic finite-element model to predict dynamic characteristics of a flexible wing transport aircraft. The aircraft model is based on NASA Generic Transport Model (GTM) with representative mass and stiffness properties to achieve a wing tip deflection about twice that of a conventional transport aircraft (10% versus 5%). This flexible wing transport aircraft is referred to as an Elastically Shaped Aircraft Concept (ESAC) which is equipped with a Variable Camber Continuous Trailing Edge Flap (VCCTEF) system for active wing shaping control for drag reduction. A vortex-lattice aerodynamic model of the ESAC is developed and is coupled with an aeroelastic finite-element model via an automated geometry modeler. This coupled model is used to compute static and dynamic aeroelastic solutions. The deflection information from the finite-element model and the vortex-lattice model is used to compute unsteady contributions to the aerodynamic force and moment coefficients. A coupled aeroelastic-longitudinal flight dynamic model is developed by coupling the finite-element model with the rigid-body flight dynamic model of the GTM.

Unsteady flow challenges tracking performance at vortex shedding frequencies without disrupting lift mechanisms

NASA Astrophysics Data System (ADS)

Matthews, Megan; Sponberg, Simon

2017-11-01

Birds, insects, and many animals use unsteady aerodynamic mechanisms to achieve stable hovering flight. Natural environments are often characterized by unsteady flows causing animals to dynamically respond to perturbations while performing complex tasks, such as foraging. Little is known about how unsteady flow around an animal interacts with already unsteady flow in the environment or how this impacts performance. We study how the environment impacts maneuverability to reveal any coupling between body dynamics and aerodynamics for hawkmoths, Manduca sexta,tracking a 3D-printed robotic flower in a wind tunnel. We also observe the leading-edge vortex (LEV), a known lift-generating mechanism for insect flight with smoke visualization. Moths in still and unsteady air exhibit near perfect tracking at low frequencies, but tracking in the flower wake results in larger overshoot at mid-range. Smoke visualization of the flower wake shows that the dominant vortex shedding corresponds to the same frequency band as the increased overshoot. Despite the large effect on flight dynamics, the LEV remains bound to the wing and thorax throughout the wingstroke. In general, unsteady wind seems to decrease maneuverability, but LEV stability seems decoupled from changes in flight dynamics.

Effect of wing-tip dihedral on the longitudinal and lateral aerodynamic characteristics of a supersonic cruise configuration at subsonic speeds

NASA Technical Reports Server (NTRS)

Washburn, K. E.; Gloss, B. B.

1976-01-01

Force and moment data studies were conducted to determine the effect of wing-tip dihedral on the longitudinal and lateral aerodynamic characteristics of a supersonic cruise fighter configuration. Oil flow studies were also performed to investigate the model surface flow. Three models were tested: a flat (0 deg dihedral) wing tip, a dihedral, and an anhedral wing tip. The tests were conducted at the NASA Langley high-speed 7- by 10-foot wind tunnel.

Experiments with a wing from which the boundary layer is removed by pressure or suction

NASA Technical Reports Server (NTRS)

Wieland, K

1928-01-01

With an unsymmetrical wing and a rotating Magnus cylinder, the lift is produced by the superposition of parallel and circulatory flows. An explanation of the circulatory flow is furnished by the boundary-layer theory of Prandtl and the consequent vortex formation. According to this explanation, it must evidently be possible to increase the circulation either by increasing the size of the stronger (lower) vortex or by decreasing the size of the weaker (upper) vortex. In this sense, according to Professor H. Zickendraht, we have a new type of wing from which the boundary layer is removed by forcing air out or sucking it in through openings in the upper surface of the wing near its trailing edge.

An integrated Navier-Stokes - full potential - free wake method for rotor flows

NASA Astrophysics Data System (ADS)

Berkman, Mert Enis

1998-12-01

The strong wake shed from rotary wings interacts with almost all components of the aircraft, and alters the flow field thus causing performance and noise problems. Understanding and modeling the behavior of this wake, and its effect on the aerodynamics and acoustics of helicopters have remained as challenges. This vortex wake and its effect should be accurately accounted for in any technique that aims to predict rotor flow field and performance. In this study, an advanced and efficient computational technique for predicting three-dimensional unsteady viscous flows over isolated helicopter rotors in hover and in forward flight is developed. In this hybrid technique, the advantages of various existing methods have been combined to accurately and efficiently study rotor flows with a single numerical method. The flow field is viewed in three parts: (i) an inner zone surrounding each blade where the wake and viscous effects are numerically captured, (ii) an outer zone away from the blades where wake is modeled, and (iii) a Lagrangean wake which induces wake effects in the outer zone. This technique was coded in a flow solver and compared with experimental data for hovering and advancing rotors including a two-bladed rotor, the UH-60A rotor and a tapered tip rotor. Detailed surface pressure, integrated thrust and torque, sectional thrust, and tip vortex position predictions compared favorably against experimental data. Results indicated that the hybrid solver provided accurate flow details and performance information typically in one-half to one-eighth cost of complete Navier-Stokes methods.

The gust-mitigating potential of flapping wings.

PubMed

Fisher, Alex; Ravi, Sridhar; Watkins, Simon; Watmuff, Jon; Wang, Chun; Liu, Hao; Petersen, Phred

2016-08-02

Nature's flapping-wing flyers are adept at negotiating highly turbulent flows across a wide range of scales. This is in part due to their ability to quickly detect and counterract disturbances to their flight path, but may also be assisted by an inherent aerodynamic property of flapping wings. In this study, we subject a mechanical flapping wing to replicated atmospheric turbulence across a range of flapping frequencies and turbulence intensities. By means of flow visualization and surface pressure measurements, we determine the salient effects of large-scale freestream turbulence on the flow field, and on the phase-average and fluctuating components of pressure and lift. It is shown that at lower flapping frequencies, turbulence dominates the instantaneous flow field, and the random fluctuating component of lift contributes significantly to the total lift. At higher flapping frequencies, kinematic forcing begins to dominate and the flow field becomes more consistent from cycle to cycle. Turbulence still modulates the flapping-induced flow field, as evidenced in particular by a variation in the timing and extent of leading edge vortex formation during the early downstroke. The random fluctuating component of lift contributes less to the total lift at these frequencies, providing evidence that flapping wings do indeed provide some inherent gust mitigation.

The effects of free stream turbulence on the flow field through a compressor cascade

NASA Astrophysics Data System (ADS)

Muthanna Kolera, Chittiappa

The flow through a compressor cascade with tip leakage has been studied experimentally. The cascade of GE rotor B section blades had an inlet angle of 65.1°, a stagger angle of 56.9°, and a solidity of 1.08. The final turning angle of the cascade was 11.8°. This compressor configuration was representative of the core compressor of an aircraft engine. The cascade was operated with a tip gap of 1.65%, and operated at a Reynolds number based on the chord length (0.254 m) of 388,000. Measurements were made at 8 axial locations to reveal the structure of the flow as it evolved through the cascade. Measurements were also made to reveal the effects of grid generated turbulence on this flow. The data set is unique in that not only does it give a comparison of elevated free stream turbulence effects, but also documents the developing flow through the blade row of a compressor cascade with tip leakage. Measurements were made at a total of 8 locations 0.8, 0.23 axial chords upstream and 0, 0.27, 0.48, 0.77, 0.98, and 1.26 axial chords downstream of the leading edge of the blade row for both inflow turbulence cases. The measurements revealed the formation and development of the tip leakage vortex within the passage. The tip leakage vortex becomes apparent at approximately X/ca = 0.27 and dominated much of the endwall flow. The tip leakage vortex is characterized by high streamwise velocity deficits, high vorticity and high turbulence kinetic energy levels. The result showed that between 0.77 and 0.98 axial chords downstream of the leading edge, the vortex structure and behavior changes. The effects of grid generated turbulence were also documented. The results revealed significant effects on the flow field. The results showed a 4% decrease in the blade loading and a 20% reduction in the vorticity levels within tip leakage vortex. There was also a shift in the vortex path, showing a shift close to the suction side with grid generated turbulence, indicating the strength of the vortex was decreased. Circulation calculations showed this reduction, and also indicated that the tip leakage vortex increased in size by about 30%. The results revealed that overall, the turbulence kinetic energy levels in the tip leakage vortex were increased, with the most drastic change occurring at X/ca = 0.77.

Aerodynamics of wings at low Reynolds numbers: Boundary layer separation and reattachment

NASA Astrophysics Data System (ADS)

McArthur, John

Due to advances in electronics technology, it is now possible to build small scale flying and swimming vehicles. These vehicles will have size and velocity scales similar to small birds and fish, and their characteristic Reynolds number will be between 104 and 105. Currently, these flying and swimming vehicles do not perform well, and very little research has been done to characterize them, or to explain why they perform so poorly. This dissertation documents three basic investigations into the performance of small scale lifting surfaces, with Reynolds numbers near 104. Part I. Low Reynolds number aerodynamics. Three airfoil shapes were studied at Reynolds numbers of 1 and 2x104: a flat plate airfoil, a circular arc cambered airfoil, and the Eppler 387 airfoil. Lift and drag force measurements were made on both 2D and 3D conditions, with the 3D wings having an aspect ratio of 6, and the 2D condition being approximated by placing end plates at the wing tips. Comparisons to the limited number of previous measurements show adequate agreement. Previous studies have been inconclusive on whether lifting line theory can be applied to this range of Re, but this study shows that lifting line theory can be applied when there are no sudden changes in the slope of the force curves. This is highly dependent on the airfoil shape of the wing, and explains why previous studies have been inconclusive. Part II. The laminar separation bubble. The Eppler 387 airfoil was studied at two higher Reynolds numbers: 3 and 6x10 4. Previous studies at a Reynolds number of 6x104 had shown this airfoil experiences a drag increase at moderate lift, and a subsequent drag decrease at high lift. Previous studies suggested that the drag increase is caused by a laminar separation bubble, but the experiments used to show this were conducted at higher Reynolds numbers and extrapolated down. Force measurements were combined with flow field measurements at Reynolds numbers 3 and 6x104 to determine whether the drag increase is really caused by the formation of a laminar separation bubble. The results clearly indicate that the reverse is true, and that the subsequent drag decrease is caused by the laminar separation bubble. Part III. The leading edge vortex. Four wings with different sweep angles were studied at Reynolds number 5x104: sweep angles of 0, 20, 40, and 60 degrees. The wings had a simple cambered plate airfoil similar to the cambered airfoil of part I above. Each wing was built to have the same aspect ratio, wing area, and streamwise airfoil shape. Previous studies on bird wings speculate that simply sweeping the wings can cause a leading edge vortex to form, which could cause substantial improvements in performance. However, these studies were not well controlled, and were conducted from a biological perspective. Qualitative and quantitative flow field measurements were combined with force measurements to conduct a well controlled engineering experiment on the formation and effect of a leading edge vortex on simple swept wings. A stable vortex was found to form over the 60 degree swept wing at one particular angle of attack, but it was not similar to the traditional notion of a leading edge vortex. The vortex has a small radius, and extends over little of the span. Force measurements indicate that the vortex has no significant impact on the forces measured. Thus, simply sweeping a wing is not sufficient to form a significant leading edge vortex, and other effects must be considered.

A Novel Approach for Reducing Rotor Tip-Clearance Induced Noise in Turbofan Engines

NASA Technical Reports Server (NTRS)

Khorrami, Mehdi R.; Li, Fei; Choudhari, Meelan

2001-01-01

Rotor tip-clearance induced noise, both in the form of rotor self noise and rotor-stator interaction noise , constitutes a significant component of total fan noise. Innovative yet cost effective techniques to suppress rotor-generated noise are, therefore, of foremost importance for improving the noise signature of turbofan engines. To that end, the feasibility of a passive porous treatment strategy to positively modify the tip-clearance flow field is addressed. The present study is focused on accurate viscous flow calculations of the baseline and the treated rotor flow fields. Detailed comparison between the computed baseline solution and experimental measurements shows excellent agreement. Tip-vortex structure, trajectory, strength, and other relevant aerodynamic quantities are extracted from the computed database. Extensive comparison between the untreated and treated tip-clearance flow fields is performed. The effectiveness of the porous treatment for altering the rotor-tip vortex flow field in general and reducing the intensity of the tip vortex, in particular, is demonstrated. In addition, the simulated flow field for the treated tip clearly shows that substantial reduction in the intensity of both the shear layer roll-up and boundary layer separation on the wall is achieved.

Analytical observations on the aerodynamics of a delta wing with leading edge flaps

NASA Technical Reports Server (NTRS)

Oh, S.; Tavella, D.

1986-01-01

The effect of a leading edge flap on the aerodynamics of a low aspect ratio delta wing is studied analytically. The separated flow field about the wing is represented by a simple vortex model composed of a conical straight vortex sheet and a concentrated vortex. The analysis is carried out in the cross flow plane by mapping the wing trace, by means of the Schwarz-Christoffel transformation into the real axis of the transformed plane. Particular attention is given to the influence of the angle of attack and flap deflection angle on lift and drag forces. Both lift and drag decrease with flap deflection, while the lift-to-drag ratioe increases. A simple coordinate transformation is used to obtain a closed form expression for the lift-to-drag ratio as a function of flap deflection. The main effect of leading edge flap deflection is a partial suppression of the separated flow on the leeside of the wing. Qualitative comparison with experiments is presented, showing agreement in the general trends.

On the three-dimensional interaction of a rotor-tip vortex with a cylindrical surface

NASA Astrophysics Data System (ADS)

Radcliff, Thomas D.; Burggraf, Odus R.; Conlisk, A. T.

2000-12-01

The collision of a strong vortex with a surface is an important problem because significant impulsive loads may be generated. Prediction of helicopter fatigue lifetime may be limited by an inability to predict these loads accurately. Experimental results for the impingement of a helicopter rotor-tip vortex on a cylindrical airframe show a suction peak on the top of the airframe that strengthens and then weakens within milliseconds. A simple line-vortex model can predict the experimental results if the vortex is at least two vortex-core radii away from the airframe. After this, the model predicts continually deepening rather than lessening suction as the vortex stretches. Experimental results suggest that axial flow within the core of a tip vortex has an impact on the airframe pressure distribution upon close approach. The mechanism for this is hypothesized to be the inviscid redistribution of the vorticity field within the vortex as the axial velocity stagnates. Two models of a tip vortex with axial flow are considered. First, a classical axisymmetric line vortex with a cutoff parameter is superimposed with vortex ringlets suitably placed to represent the helically wound vortex shed by the rotor tip. Thus, inclusion of axial flow is found to advect vortex core thinning away from the point of closest interaction as the vortex stretches around the cylindrical surface during the collision process. With less local thinning, vorticity in the cutoff parameter model significantly overlaps the solid cylinder in an unphysical manner, highlighting the fact that the vortex core must deform from its original cylindrical shape. A second model is then developed in which axial and azimuthal vorticity are confined within a rectangular-section vortex. Area and aspect ratio of this vortex can be varied independently to simulate deformation of the vortex core. Both axial velocity and core deformation are shown to be important to calculate the local induced pressure loads properly. The computational results are compared with experiments conducted at the Georgia Institute of Technology.

Interaction of a Vortex with Axial Flow and a Cylindrical Surface

NASA Astrophysics Data System (ADS)

Radcliff, T. D.; Burgraff, O. R.; Conlisk, A. T.

1998-11-01

The direct collision of a vortex with a surface is an important problem because significant impulsive loads may be generated leading to premature fatigue. Experimental results for the impingement of a tip-vortex on a cylindrical airframe indicate that a suction peak forms on the top of the airframe which is subsequently reduced within milliseconds of vortex-surface contact. A simple line-vortex model can predict the experimental results until the vortex is within a vortex-core radius of the airframe. After this the model predicts continually deepening rather than lessening suction. Study of the experimental results suggests that axial flow within the core of a tip-vortex has an impact on the airframe pressure distribution upon close approach. The mechanism for this is hypothesized to be the inviscid redistribution of the vorticity field within the vortex coupled with deformation of the vortex core. Two models of a tip-vortex with axial flow are considered. First a classical line vortex with a cut-off parameter is superimposed with suitably placed vortex rings. This model simulates the helically wound vortex shed by the rotor tip. Inclusion of axial flow is found to prevent thinning of the vortex core as the vortex stretches around the cylindrical surface during the collision process. With less thinning, vorticity is observed to overlap the solid cylinder, highlighting the fact that the vortex core must deform from its original cylindrical shape. A second model is developed in which axial and azimuthal vorticity are uniformly distributed throughout a rectangular-section vortex. Area and aspect ratio of this vortex can be varied independently to simulate deformation of the vortex core. Both vorticity redistribution and core deformation are shown to be important to properly calculate the local induced pressure loads. The computational results are compared with the results of experiments conducted at the Georgia Institute of Technology.

In-flight leading-edge extension vortex flow-field survey measurements on a F-18 aircraft at high angle of attack

NASA Technical Reports Server (NTRS)

Richwine, David M.; Fisher, David F.

1992-01-01

Flow-field measurements on the leading-edge extension (LEX) of the F-18 High Alpha Research Vehicle (HARV) were obtained using a rotating rake with 16 hemispherical-tipped five-hole probes. Detailed pressure, velocity, and flow direction data were obtained through the LEX vortex core. Data were gathered during 1-g quasi-stabilized flight conditions at angles of attack alpha from 10 degrees to 52 degrees and at Reynolds numbers based on mean aerodynamic cord up to 16 x 10(exp 6). Normalized dynamic pressures and crossflow velocities clearly showed the primary vortex above the LEX and formation of a secondary vortex at higher angles of attack. The vortex was characterized by a ring of high dynamic pressure surrounding a region of low dynamic pressure at the vortex core center. The vortex core, subcore diameter, and vertical location of the core above the LEX increased with angle of attack. Minimum values for static pressure were obtained in the vortex subcore and decreased nearly linearly with increasing angle of attack until vortex breakdown. Rake-measured static pressures were consistent with previously documented surface pressures and showed good agreement with flow visualization flight test results. Comparison of the LEX vortex flight test data to computational solutions at alpha approximately equals 19 degrees and 30 degrees showed fair correlation.

F/A-18 Performance Benefits Measured During the Autonomous Formation Flight Project

NASA Technical Reports Server (NTRS)

Vachon, M. Jake; Ray, Ronald J.; Walsh, Kevin R.; Ennix, Kimberly

2003-01-01

The Autonomous Formation Flight (AFF) project at the NASA Dryden Flight Research Center (Edwards, California) investigated performance benefits resulting from formation flight, such as reduced aerodynamic drag and fuel consumption. To obtain data on performance benefits, a trailing F/A-18 airplane flew within the wing tip-shed vortex of a leading F/A-18 airplane. The pilot of the trail airplane used advanced station-keeping technology to aid in positioning the trail airplane at precise locations behind the lead airplane. The specially instrumented trail airplane was able to obtain accurate fuel flow measurements and to calculate engine thrust and vehicle drag. A maneuver technique developed for this test provided a direct comparison of performance values while flying in and out of the vortex. Based on performance within the vortex as a function of changes in vertical, lateral, and longitudinal positioning, these tests explored design-drivers for autonomous stationkeeping control systems. Observations showed significant performance improvements over a large range of trail positions tested. Calculations revealed maximum drag reductions of over 20 percent, and demonstrated maximum reductions in fuel flow of just over 18 percent.

Numerical studies of incompressible flow around delta and double-delta wings

NASA Technical Reports Server (NTRS)

Krause, E.; Liu, C. H.

1989-01-01

The subject has been jointly investigated at NASA Langley Research Center and the Aerodynamisches Institut of the RWTH Aachen over a substantial period. The aim of this investigation has been to develop numerical integration procedures for the Navier-Stokes equations - particularly for incompressible three-dimensional viscous flows about simple and double delta wings - and to study the low speed flow behavior, with its complex vortex structures on the leeward side of the wing. The low speed flight regime poses unusual problems because high incidence flight conditions may, for example, encounter symmetric and asymmetric vortex breakdown. Because of the many difficulties to be expected in solving the problem, it was divided into two - analysis of the flow without vortex breakdown and analysis of the breakdown of isolated vortices. The major results obtained so far on the two topics are briefly described.

X-29 vortex flow control tests

NASA Technical Reports Server (NTRS)

Hancock, Regis; Fullerton, Gordon

1992-01-01

A joint Air Force/NASA X-29 aircraft program to improve yaw control at high angle of attack using vortex flow control (VFC) is described. Directional VFC blowing proved to a be a powerful yaw moment generator and was very effective in overriding natural asymmetries, but was essentially ineffective in suppressing wing rock. Symmetric aft blowing also had little effect on suppressing wing rock.

Measurements of the tip leakage vortex structures and turbulence in the meridional plane of an axial water-jet pump

NASA Astrophysics Data System (ADS)

Wu, Huixuan; Miorini, Rinaldo L.; Katz, Joseph

2011-04-01

Particle image velocimetry (PIV) measurements at varying resolutions focus on the flow structures in the tip region of a water-jet pump rotor, including the tip-clearance flow and the rollup process of a tip leakage vortex (TLV). Unobstructed views of these regions are facilitated by matching the optical refractive index of the transparent pump with that of the fluid. High-magnification data reveal the flow non-uniformities and associated turbulence within the tip gap. Instantaneous data and statistics of spatial distributions and strength of vortices in the rotor passage reveal that the leakage flow emerges as a wall jet with a shear layer containing a train of vortex filaments extending from the tip of the blade. These vortices are entrained into the TLV, but do not have time to merge. TLV breakdown in the aft part of the blade passage further fragments these structures, increasing their number and reducing their size. Analogy is made between the circumferential development of the TLV in the blade passage and that of the starting jet vortex ring rollup. Subject to several assumptions, these flows display similar trends, including conditions for TLV separation from the shear layer feeding vorticity into it.

Study of tip clearance flow in a turbomachinery cascade using large eddy simulation

NASA Astrophysics Data System (ADS)

You, Donghyun

In liquid handling systems like pumps and ducted propulsors, low pressure events in the vicinity and downstream of the rotor tip gap can induce tip-leakage cavitation which leads to noise, vibration, performance loss, and erosions of blade and casing wall. In order to analyze the dynamics of the tip-clearance flow and determine the underlying mechanism for the low pressure events, a newly developed large-eddy simulation (LES) solver which combines an immersed-boundary method with a generalized curvilinear structured grid has been employed. An analysis of the LES results has been performed to understand the mean flow field, turbulence characteristics, vortex dynamics, and pressure fluctuations in the turbomachinery cascade with tip gap. In the cascade passage, the tip-leakage jet, which is generated by the pressure difference between the pressure and suction sides of the blade tip, is found to produce highly enhanced vorticity magnitude and significant levels of turbulent kinetic energy. Based on the understanding of the flow field, a guideline for reducing viscous loss in the cascade is provided. Analyses of the energy spectra and space-time correlations of the velocity fluctuations suggest that the tip-leakage vortex is subject to pitchwise wandering motion. The largest pressure drop and most intense pressure fluctuations due to the formation of the tip-leakage vortex are found at the location where the strongest portion of the tip-leakage vortex is found. Present study suggests that the tip-leakage vortex needs to be controlled in its origin to reduce cavitation in the present configuration. The effects of tip-gap size on the end-wall vortical structures and on the velocity and pressure fields have been investigated. The present analysis indicates that the mechanism for the generation of the vorticity and turbulent kinetic energy is mostly unchanged by the tip-gap size variation. However, larger tip-gap sizes are found to be more inductive to tip-leakage cavitation judged by the levels of negative mean pressure and pressure fluctuations.

Propeller tip and hub vortex dynamics in the interaction with a rudder

NASA Astrophysics Data System (ADS)

Felli, Mario; Falchi, Massimo

2011-11-01

In the present paper, the interaction mechanisms of the vortices shed by a single-screw propeller with a rudder installed in its wake are addressed; in particular, following the works by Felli et al. (Exp Fluids 6(1):1-11, 2006a, Exp Fluids 46(1):147-1641, 2009a, Proceedings of the 8th international symposium on particle image velocimetry: Piv09, Melbourne, 2009b), the attention is focused on the analysis of the evolution, instability, breakdown and recovering mechanisms of the propeller tip and hub vortices during the interaction with the rudder. To investigate these mechanisms in detail, a wide experimental activity consisting in time-resolved visualizations, velocity measurements by particle image velocimetry (PIV) and laser Doppler velocimetry (LDV) along horizontal chordwise, vertical chordwise and transversal sections of the wake have been performed in the Cavitation Tunnel of the Italian Navy. Collected data allows to investigate the major flow features that distinguish the flow field around a rudder operating in the wake of a propeller, as, for example, the spiral breakdown of the vortex filaments, the rejoining mechanism of the tip vortices behind the rudder and the mechanisms governing the different spanwise misalignment of the vortex filaments in the pressure and suction sides of the appendage.

Wind-tunnel studies of advanced cargo aircraft concepts. [leading edge vortex flaps for drag reduction

NASA Technical Reports Server (NTRS)

Rao, D. M.; Goglia, G. L.

1981-01-01

Accomplishments in vortex flap research are summarized. A singular feature of the vortex flap is that, throughout the range of angle of attack range, the flow type remains qualitatively unchanged. Accordingly, no large or sudden change in the aerodynamic characteristics, as happens when forcibly maintained attached flow suddenly reverts to separation, will occur with the vortex flap. Typical wind tunnel test data are presented which show the drag reduction potential of the vortex flap concept applied to a supersonic cruise airplane configuration. The new technology offers a means of aerodynamically augmenting roll-control effectiveness on slender wings at higher angles of attack by manipulating the vortex flow generated from leading edge separation. The proposed manipulator takes the form of a flap hinged at or close to the leading edge, normally retracted flush with the wing upper surface to conform to the airfoil shape.

Cut-cell method based large-eddy simulation of tip-leakage flow

NASA Astrophysics Data System (ADS)

Pogorelov, Alexej; Meinke, Matthias; Schröder, Wolfgang

2015-07-01

The turbulent low Mach number flow through an axial fan at a Reynolds number of 9.36 × 105 based on the outer casing diameter is investigated by large-eddy simulation. A finite-volume flow solver in an unstructured hierarchical Cartesian setup for the compressible Navier-Stokes equations is used. To account for sharp edges, a fully conservative cut-cell approach is applied. A newly developed rotational periodic boundary condition for Cartesian meshes is introduced such that the simulations are performed just for a 72° segment, i.e., the flow field over one out of five axial blades is resolved. The focus of this numerical analysis is on the development of the vortical flow structures in the tip-gap region. A detailed grid convergence study is performed on four computational grids with 50 × 106, 250 × 106, 1 × 109, and 1.6 × 109 cells. Results of the instantaneous and the mean fan flow field are thoroughly analyzed based on the solution with 1 × 109 cells. High levels of turbulent kinetic energy and pressure fluctuations are generated by a tip-gap vortex upstream of the blade, the separating vortices inside the tip gap, and a counter-rotating vortex on the outer casing wall. An intermittent interaction of the turbulent wake, generated by the tip-gap vortex, with the downstream blade, leads to a cyclic transition with high pressure fluctuations on the suction side of the blade and a decay of the tip-gap vortex. The disturbance of the tip-gap vortex results in an unsteady behavior of the turbulent wake causing the intermittent interaction. For this interaction and the cyclic transition, two dominant frequencies are identified which perfectly match with the characteristic frequencies in the experimental sound power level and therefore explain their physical origin.

Implementation of Parallel Computing Technology to Vortex Flow

NASA Technical Reports Server (NTRS)

Dacles-Mariani, Jennifer

1999-01-01

Mainframe supercomputers such as the Cray C90 was invaluable in obtaining large scale computations using several millions of grid points to resolve salient features of a tip vortex flow over a lifting wing. However, real flight configurations require tracking not only of the flow over several lifting wings but its growth and decay in the near- and intermediate- wake regions, not to mention the interaction of these vortices with each other. Resolving and tracking the evolution and interaction of these vortices shed from complex bodies is computationally intensive. Parallel computing technology is an attractive option in solving these flows. In planetary science vortical flows are also important in studying how planets and protoplanets form when cosmic dust and gases become gravitationally unstable and eventually form planets or protoplanets. The current paradigm for the formation of planetary systems maintains that the planets accreted from the nebula of gas and dust left over from the formation of the Sun. Traditional theory also indicate that such a preplanetary nebula took the form of flattened disk. The coagulation of dust led to the settling of aggregates toward the midplane of the disk, where they grew further into asteroid-like planetesimals. Some of the issues still remaining in this process are the onset of gravitational instability, the role of turbulence in the damping of particles and radial effects. In this study the focus will be with the role of turbulence and the radial effects.

Wake Characteristics of a Flapping Wing Optimized for both Aerial and Aquatic Flight

NASA Astrophysics Data System (ADS)

Izraelevitz, Jacob; Kotidis, Miranda; Triantafyllou, Michael

2017-11-01

Multiple aquatic bird species (including murres, puffins, and other auks) employ a single actuator to propel themselves in two different fluid media: both flying and swimming using primarily their flapping wings. This impressive design compromise could be adopted by engineered implementations of dual aerial/aquatic robotic platforms, as it offers an existence proof for favorable flow physics. We discuss one realization of a 3D flapping wing actuation system for use in both air and water. The wing oscillates by the root and employs an active in-line motion degree-of-freedom. An experiment-coupled optimization routine generates the wing trajectories, controlling the unsteady forces throughout each flapping cycle. We elucidate the wakes of these wing trajectories using dye visualization, correlating the wake vortex structures with simultaneous force measurements. After optimization, the wing generates the large force envelope necessary for propulsion in both fluid media, and furthermore, demonstrate improved control over the unsteady wake.

Vortical flow management for improved configuration aerodynamics: Recent experiences

NASA Technical Reports Server (NTRS)

Rao, D. M.

1983-01-01

Recent progress in vortex-control applications for alleviating the adverse consequences of three dimensional separation and vortical interactions on slender body/swept wing configurations is reported. Examples include helical separation trip to alleviate the side force due to forebody vortex asymmetry; hinged strakes to avoid vortex breakdown effects; compartmentation of swept leading edge separation to delay the pitch-up instability; under wing vortex trip and vortex trip and vortex flaps for drag reduction at high lift; and an apex-flap trimmer to fully utilize the lift capability of trailing-edge flaps for take off and landing of delta wings. Experimental results on generic wind-tunnel models are presented to illustrate the vortex-management concepts involved and to indicate their potential for enhancing the subsonic aerodynamics of supersonic-cruise type vehicles.

Vortex sensing tests at NAFEC.

DOT National Transportation Integrated Search

1972-01-01

The report describes the results of a series of tests to determine and evaluate three experimental techniques for remote sensing of the wing-tip vortices generated by heavy commercial and military aircraft. These techniques involved a pulsed, bistati...

Control of Flow Structure on Non-Slender Delta Wing: Bio-inspired Edge Modifications, Passive Bleeding, and Pulsed Blowing

NASA Astrophysics Data System (ADS)

Yavuz, Mehmet Metin; Celik, Alper; Cetin, Cenk

2016-11-01

In the present study, different flow control approaches including bio-inspired edge modifications, passive bleeding, and pulsed blowing are introduced and applied for the flow over non-slender delta wing. Experiments are conducted in a low speed wind tunnel for a 45 degree swept delta wing using qualitative and quantitative measurement techniques including laser illuminated smoke visualization, particle image velocimety (PIV), and surface pressure measurements. For the bio-inspired edge modifications, the edges of the wing are modified to dolphin fluke geometry. In addition, the concept of flexion ratio, a ratio depending on the flexible length of animal propulsors such as wings, is introduced. For passive bleeding, directing the free stream air from the pressure side of the planform to the suction side of the wing is applied. For pulsed blowing, periodic air injection through the leading edge of the wing is performed in a square waveform with 25% duty cycle at different excitation frequencies and compared with the steady and no blowing cases. The results indicate that each control approach is quite effective in terms of altering the overall flow structure on the planform. However, the success level, considering the elimination of stall or delaying the vortex breakdown, depends on the parameters in each method.

Effects of Various Fillet Shapes on a 76/40 Double Delta Wing from Mach 0.18 to 0.7

NASA Technical Reports Server (NTRS)

Erickson, Gary E.; Bell, James H.; Gonzalez, Hugo A.; McLachlan, Blair G.

2003-01-01

The effects of linear, diamond, and parabolic fillets on a double delta wing were investigated in the NASA Langley 7 x 10 ft High Speed Tunnel from Mach 0.18 to 0.7 and angles of attack from 4 deg. to 42 deg. Force and moment, pneumatic pressures, pressure sensitive paint, and vapor screen flow visualization measurements were used to characterize the flow field and to determine longitudinal forces and moments. The fillets increased lift coefficient and reduced induced drag without significantly affecting pitching moment. Pressure sensitive paint showed the increase in lift is caused by an increase in suction and broadening of the vortex suction footprint. Vapor screen results showed the mixing and coalescing of the strake fillet and wing vortices causes the footprint to broaden.

Load distribution on a closed-coupled wing canard at transonic speeds

NASA Technical Reports Server (NTRS)

Gloss, B. B.; Washburn, K. E.

1977-01-01

A wind tunnel test where load distributions were obtained at transonic speeds on both the canard and wing surfaces of a closely coupled wing canard configuration is reported. Detailed component and configuration arrangement studies to provide insight into the various aerodynamic interference effects for the leading edge vortex flow conditions encountered are included. Data indicate that increasing the Mach number from 0.70 to 0.95 caused the wing leading edge vortex to burst over the wing when the wing was in the presence of the high canard.

Clap and Fling Interaction of Bristled Wings: Effects of Varying Reynolds Number and Bristle Spacing on Force Generation and Flow Structures

NASA Astrophysics Data System (ADS)

Kasoju, Vishwa Teja

The smallest flying insects with body lengths under 1 mm, such as thrips and fairyflies, typically show the presence of long bristles on their wings. Thrips have been observed to use wing-wing interaction via 'clap and fling' for flapping flight at low Reynolds number (Re) on the order of 10, where a wing pair comes into close contact at the end of upstroke and fling apart at the beginning of downstroke. We examined the effects of varying the following parameters on force generation and flow structures formed during clap and fling: (1) Re ranging from 5 to 15 for a bristled wing pair (G/D = 17) and a geometrically equivalent solid wing pair; and (2) ratio of spacing between bristles to bristle diameter (G/D) for Re = 10. The G/D ratio in 70 thrips species were quantified from published forewing images. Scaled-up physical models of three bristled wing pairs of varying G/D (5, 11, 17) and a solid wing pair (G/D = 0) were fabricated. A robotic model was used for this study, in which a wing pair was immersed in an aquarium tank filled with glycerin and driven by stepper motors to execute clap and fling kinematics. Dimensionless lift and drag coefficients were determined from strain gauge measurements. Phase-locked particle image velocimetry (PIV) measurements were used to examine flow through the bristles. Chordwise PIV was used to visualize the leading edge vortex (LEV) and trailing edge vortex (TEV) formed over the wings during clap and fling. With increasing G/D, larger reduction was observed in peak drag coefficients as compared to reduction in peak lift coefficients. Net circulation, defined as the difference in circulation (strength) of LEV and TEV, diminished with increasing G/D. Reduction in net circulation resulted in reducing lift generated by bristled wings as compared to solid wings. Leaky, recirculating flow through the bristles provided large drag reduction during fling of a bristled wing pair. If flight efficiency is defined as the ratio of lift to drag, largest peak lift to peak drag ratios were obtained in bristled wings as compared to the solid wings across the entire range of Re and G/D tested.

Aerodynamics Investigation of Faceted Airfoils at Low Reynolds Number

NASA Astrophysics Data System (ADS)

Napolillo, Zachary G.

The desire and demand to fly farther and faster has progressively integrated the concept of optimization with airfoil design, resulting in increasingly complex numerical tools pursuing efficiency often at diminishing returns; while the costs and difficulty associated with fabrication increases with design complexity. Such efficiencies may often be necessary due to the power density limitations of certain aircraft such as small unmanned aerial vehicles (UAVs) and micro air vehicles (MAVs). This research, however, focuses on reducing the complexity of airfoils for applications where aerodynamic performance is less important than the efficiency of manufacturing; in this case a Hybrid Projectile. By employing faceted sections to approximate traditional contoured wing sections it may be possible to expedite manufacturing and reduce costs. We applied this method to the development of a low Reynolds number, disposable Hybrid Projectile requiring a 4.5:1 glide ratio, resulting in a series of airfoils which are geometric approximations to highly contoured cross-sections called ShopFoils. This series of airfoils both numerically and experimentally perform within a 10% margin of the SD6060 airfoil at low Re. Additionally, flow visualization has been conducted to qualitatively determine what mechanisms, if any, are responsible for the similarity in performance between the faceted ShopFoil sections and the SD6060. The data obtained by these experiments did not conclusively reveal how the faceted surfaces may influence low Re flow but did indicate that the ShopFoil s did not maintain flow attachment at higher angles of attack than the SD6060. Two reasons are provided for the unexpected performance of the ShopFoil: one is related to downwash effects, which are suspected of placing the outer portion of the span at an effective angle of attack where the ShopFoils outperform the SD6060; the other is the influence of the tip vortex on separation near the wing tips, which possibly provides a 'comparative advantage' to the ShopFoil because it has more to gain from a reduction in its pressure drag component.

Advance Ratio Effects on the Dynamic-stall Vortex of a Rotating Blade in Steady Forward Flight

DTIC Science & Technology

2014-08-06

dependence on advance ratio is used to relate the stability of the dynamic-stall vortex to Coriolis effects . Advance ratio effects on the dynamic-stall vortex...relate the stability of the dynamic-stall vortex to Coriolis effects . Keywords: Leading-edge vortex, Dynamic stall vortex, Vortex flows, Rotating wing...Reynolds number are not decoupled. 3. Radial flow field In the rotating environment the coupled effect of centripetal and Coriolis accelerations is ex

Investigation of rotor blade tip-vortex aerodynamics

NASA Technical Reports Server (NTRS)

Lewellen, W. S.

1971-01-01

Several aspects of the aerodynamics of rotor blade tip vortices are examined. Two particular categories are dealt with; (1) dynamic loads on a blade passing close to or intersecting a trailing vortex, and (2) the response of the trailing vortex core to changes in the flow. Results for both categories are in reasonable agreement with existing data, although lower pressure gradients were obtained than anticipated for category one. A correlation between trailing edge sweep angle at the tip and vortex core size was noted for category two.

The effect of leading edge tubercles on dynamic stall

NASA Astrophysics Data System (ADS)

Hrynuk, John

The effect of the leading edge tubercles of humpback whales has been heavily studied for their static benefits. These studies have shown that tubercles inhibit flow separation, limit spanwise flow, and extend the operating angle of a wing beyond the static stall point while maintaining lift, all while having a comparatively low negative impact on drag. The current study extends the prior work to investigating the effect of tubercles on dynamic stall, a fundamental flow phenomenon that occurs when wings undergo dynamic pitching motions. Flow fields around the wing models tested were studied using Laser Induced Fluorescence (LIF) and Molecular Tagging Velocimetry (MTV).Resulting velocity fields show that the dynamics of the formation and separation of the leading edge vortex were fundamentally different between the straight wing and the tubercled wing. Tracking of the Dynamic Stall Vortex (DSV) and Shear Layer Vortices (SLVs), which may have a significant impact on the overall flow behavior, was done along with calculations of vortex circulation. Proximity to the wing surface and total circulation were used to evaluate potential dynamic lift increases provided by the tubercles. The effects of pitch rate on the formation process and benefits of the tubercles were also studied and were generally consistent with prior dynamic stall studies. However, tubercles were shown to affect the SLV formation and the circulation differently at higher pitch rates.

Three-dimensional flow structures and evolution of the leading-edge vortices on a flapping wing.

PubMed

Lu, Yuan; Shen, Gong Xin

2008-04-01

Following the identification and confirmation of the substructures of the leading-edge vortex (LEV) system on flapping wings, it is apparent that the actual LEV structures could be more complex than had been estimated in previous investigations. In this experimental study, we reveal for the first time the detailed three-dimensional (3-D) flow structures and evolution of the LEVs on a flapping wing in the hovering condition at high Reynolds number (Re=1624). This was accomplished by utilizing an electromechanical model dragonfly wing flapping in a water tank (mid-stroke angle of attack=60 degrees) and applying phase-lock based multi-slice digital stereoscopic particle image velocimetry (DSPIV) to measure the target flow fields at three typical stroke phases: at 0.125 T (T=stroke period), when the wing was accelerating; at 0.25 T, when the wing had maximum speed; and at 0.375 T, when the wing was decelerating. The result shows that the LEV system is a collection of four vortical elements: one primary vortex and three minor vortices, instead of a single conical or tube-like vortex as reported or hypothesized in previous studies. These vortical elements are highly time-dependent in structure and show distinct ;stay properties' at different spanwise sections. The spanwise flows are also time-dependent, not only in the velocity magnitude but also in direction.

Wind Tunnel Test of a Risk-Reduction Wing/Fuselage Model to Examine Juncture-Flow Phenomena

NASA Technical Reports Server (NTRS)

Kegerise, Michael A.; Neuhart, Dan H.

2016-01-01

A wing/fuselage wind-tunnel model was tested in the Langley 14- by 22-foot Subsonic Wind Tunnel in preparation for a highly-instrumented Juncture Flow Experiment to be conducted in the same facility. This test, which was sponsored by the NASA Transformational Tool and Technologies Project, is part of a comprehensive set of experimental and computational research activities to develop revolutionary, physics-based aeronautics analysis and design capability. The objectives of this particular test were to examine the surface and off-body flow on a generic wing/body combination to: 1) choose a final wing for a future, highly instrumented model, 2) use the results to facilitate unsteady pressure sensor placement on the model, 3) determine the area to be surveyed with an embedded laser-doppler velocimetry (LDV) system, 4) investigate the primary juncture corner- flow separation region using particle image velocimetry (PIV) to see if the particle seeding is adequately entrained and to examine the structure in the separated region, and 5) to determine the similarity of observed flow features with those predicted by computational fluid dynamics (CFD). This report documents the results of the above experiment that specifically address the first three goals. Multiple wing configurations were tested at a chord Reynolds number of 2.4 million. Flow patterns on the surface of the wings and in the region of the wing/fuselage juncture were examined using oil- flow visualization and infrared thermography. A limited number of unsteady pressure sensors on the fuselage around the wing leading and trailing edges were used to identify any dynamic effects of the horseshoe vortex on the flow field. The area of separated flow in the wing/fuselage juncture near the wing trailing edge was observed for all wing configurations at various angles of attack. All of the test objectives were met. The staff of the 14- by 22-foot Subsonic Wind Tunnel provided outstanding support and delivered exceptional value to the experiment, which exceeded expectations. The results of this test will directly inform the planning for the first of a series of instrumented-model tests at the same Reynolds number. These tests will be performed on a slightly larger-scale model with the selected wing, and will include off-body measurements with LDV and PIV, steady and unsteady pressure measurements, and the flow-visualization techniques that are discussed in this report.

Development of a nonlinear vortex method

NASA Technical Reports Server (NTRS)

Kandil, O. A.

1982-01-01

Steady and unsteady Nonliner Hybrid Vortex (NHV) method, for low aspect ratio wings at large angles of attack, is developed. The method uses vortex panels with first-order vorticity distribution (equivalent to second-order doublet distribution) to calculate the induced velocity in the near field using closed form expressions. In the far field, the distributed vorticity is reduced to concentrated vortex lines and the simpler Biot-Savart's law is employed. The method is applied to rectangular wings in steady and unsteady flows without any restriction on the order of magnitude of the disturbances in the flow field. The numerical results show that the method accurately predicts the distributed aerodynamic loads and that it is of acceptable computational efficiency.

Flowfield survey over a 75 deg swept delta wing at an angle of attack of 20.5 deg

NASA Technical Reports Server (NTRS)

Kjelgaard, S. O.; Sellers, W. L., III; Weston, R. P.

1986-01-01

An experimental investigation of the flowfield over a 75 deg swept delta wing at an angle of attack of 20.5 deg has been conducted. The data include pitot pressure surveys and two types of flow visualization. Surface and flowfield visualization data were obtained at Reynolds number, Rn, ranging from 0.5 to 2.0 million in increments of 0.25 million. Detailed pitot pressure surveys were made at five longitudinal stations at Rn = 0.5, 1.0, and 1.5 million in both the primary and secondary vortices. The results indicate that Reynolds number has only a minor effect on the global structure of the flowfield in the Reynolds number range that was investigated. The boundary layer transitions from laminar to turbulent at the trailing edge of the wing at Rn = 1.0 x 10 to the 6th, and the transition moves forward to x/L = 0.4 at Rn = 2.0 x 10 to the 6th. The positions of the primary vortex cores are insensitive to Reynolds number in this range; however, the lateral position of the secondary vortex core moves outboard aft of the region where the boundary layer transitions from laminar to turbulent.

Measurements of Supersonic Wing Tip Vortices

NASA Technical Reports Server (NTRS)

Smart, Michael K.; Kalkhoran, Iraj M.; Benston, James

1994-01-01

An experimental survey of supersonic wing tip vortices has been conducted at Mach 2.5 using small performed 2.25 chords down-stream of a semi-span rectangular wing at angle of attack of 5 and 10 degrees. The main objective of the experiments was to determine the Mach number, flow angularity and total pressure distribution in the core region of supersonic wing tip vortices. A secondary aim was to demonstrate the feasibility of using cone probes calibrated with a numerical flow solver to measure flow characteristics at supersonic speeds. Results showed that the numerically generated calibration curves can be used for 4-hole cone probes, but were not sufficiently accurate for conventional 5-hole probes due to nose bluntness effects. Combination of 4-hole cone probe measurements with independent pitot pressure measurements indicated a significant Mach number and total pressure deficit in the core regions of supersonic wing tip vortices, combined with an asymmetric 'Burger like' swirl distribution.

Wing flapping with minimum energy. [minimize the drag for a bending moment at the wing root

NASA Technical Reports Server (NTRS)

Jones, R. T.

1980-01-01

For slow flapping motions it is found that the minimum energy loss occurs when the vortex wake moves as a rigid surface that rotates about the wing root - a condition analogous to that determined for a slow-turning propeller. The optimum circulation distribution determined by this condition differs from the elliptic distribution, showing a greater concentration of lift toward the tips. It appears that very high propulsive efficiencies are obtained by flapping.

Application of slender wing benefits to military aircraft

NASA Technical Reports Server (NTRS)

Polhamus, E. C.

1983-01-01

A review is provided of aerodynamic research conducted at the Langley Research Center with respect to the application of slender wing benefits in the design of high-speed military aircraft, taking into account the supersonic performance and leading-edge vortex flow associated with very highly sweptback wings. The beginning of the development of modern classical swept wing jet aircraft is related to the German Me 262 project during World War II. In the U.S., a theoretical study conducted by Jones (1945) pointed out the advantages of the sweptback wing concept. Developments with respect to variable sweep wings are discussed, taking into account early research in 1946, a joint program of the U.S. with the United Kingdom, the tactical aircraft concept, and the important part which the Langley variable-sweep research program played in the development of the F-111, F-14, and B-1. Attention is also given to hybrid wings, vortex flow theory development, and examples of flow design technology.

A Numerical Model of Unsteady, Subsonic Aeroelastic Behavior. Ph.D Thesis

NASA Technical Reports Server (NTRS)

Strganac, Thomas W.

1987-01-01

A method for predicting unsteady, subsonic aeroelastic responses was developed. The technique accounts for aerodynamic nonlinearities associated with angles of attack, vortex-dominated flow, static deformations, and unsteady behavior. The fluid and the wing together are treated as a single dynamical system, and the equations of motion for the structure and flow field are integrated simultaneously and interactively in the time domain. The method employs an iterative scheme based on a predictor-corrector technique. The aerodynamic loads are computed by the general unsteady vortex-lattice method and are determined simultaneously with the motion of the wing. Because the unsteady vortex-lattice method predicts the wake as part of the solution, the history of the motion is taken into account; hysteresis is predicted. Two models are used to demonstrate the technique: a rigid wing on an elastic support experiencing plunge and pitch about the elastic axis, and an elastic wing rigidly supported at the root chord experiencing spanwise bending and twisting. The method can be readily extended to account for structural nonlinearities and/or substitute aerodynamic load models. The time domain solution coupled with the unsteady vortex-lattice method provides the capability of graphically depicting wing and wake motion.

Leading edge vortex control on a delta wing with dielectric barrier discharge plasma actuators

NASA Astrophysics Data System (ADS)

Shen, Lu; Wen, Chih-yung

2017-06-01

This paper presents an experimental investigation of the application of dielectric barrier discharge (DBD) plasma actuators on a slender delta wing to control the leading edge vortices (LEVs). The experiments are conducted in a wind tunnel with a Reynolds number of 50 000 based on the chord length. The smoke flow visualization reveals that the DBD plasma actuators at the leading edges significantly modify the vortical flow structure over the delta wing. It is noted that symmetric control at both semi-spans and asymmetric control at a single semi-span leads to opposite effects on the local LEVs. Particle image velocimetry (PIV) indicates that the shear layer is deformed by the actuators. Therefore, both the strength and the shape of the LEV cores are deeply affected. The six-component force measurement shows that the DBD plasma actuators have a limited effect on lift and drag while inducing relatively large moments. This suggests that the DBD plasma actuator is a promising technique for delta wing maneuvering.

Tables for the Rapid Estimation of Downwash and Sidewash Behind Wings Performing Various Motions at Supersonic Speeds

NASA Technical Reports Server (NTRS)

Bobbitt, Percy J.

1959-01-01

Equations for the downwash and sidewash due to supersonic yawed and unswept horseshoe vortices have been utilized in formulating tables and charts to permit a rapid estimation of the flow velocities behind wings performing various steady motions. Tabulations are presented of the downwash and sidewash in the wing vertical plane of symmetry due to a unit-strength yawed horseshoe vortex located at 20 equally spaced spanwise positions along lifting lines of various sweeps. (The bound portion of the yawed vortex is coincident with the lifting line.) Charts are presented for the purpose of estimating the spanwise variations of the flow-field velocities and give longitudinal variations of the downwash and sidewash at a nuMber of vertical and spanwise locations due to a unit-strength unswept horseshoe vortex. Use of the tables and charts to calculate wing downwash or sidewash requires a knowledge of the wing spanwise distribution of circulation. Sample computations for the rolling sidewash and angle-of-attack downwash behind a typical swept wing are presented to demonstrate the use of the tables and charts.

Parametric Dependence of Initial LEV Behavior on Maneuvering Wings

NASA Astrophysics Data System (ADS)

Berdon, Randall; Wabick, Kevin; Buchholz, James; Johnson, Kyle; Thurow, Brian; University of Iowa Team; Auburn University Team

2017-11-01

A maneuvering rectangular wing of aspect ratio 2 is examined experimentally using dye visualization and PIV to characterize the initial development of the leading-edge vortex (LEV) during a rolling maneuver in a uniform free stream. Understanding the underlying physics during the early evolution of the vortex is important for developing strategies to manipulate vortex evolution. Varying the dimensionless radius of gyration of the wing (Rg/c, where Rg is the radius of gyration and c is the chord) and the advance ratio (J=U/ ΩRg, where U is the free-stream velocity and Ω is the roll rate) affects the structure of the vortex and its propensity to remain attached. The influence of these parameters will be discussed, toward identification of similarity parameters governing vortex development. This work is supported by the Air Force Office of Scientific Research (Grant Number FA9550-16-1-0107, Dr. Douglas Smith, program manager).

Can Wing Tip Vortices Be Accurately Simulated?

DTIC Science & Technology

2011-07-01

additional tail buffeting.2 In commercial applications, winglets have been installed on passenger aircraft to minimize vortex formation and reduce lift...towed vehicles and cause additional tail buffeting (Ref 2). In commercial applications, winglets have been installed on passenger aircraft to

Numerical Simulation of the Interaction of a Vortex with Stationary Airfoil in Transonic Flow,

DTIC Science & Technology

1984-01-12

Goorjian, P. M., "Implicit Vortex Wakes ," AIAA Journal, Vol. 15, No. 4, April Finite- Difference Computations of Unsteady Transonic 1977, pp. 581-590... Difference Simulations of Three- tion of Wing- Vortex Interaction in Transonic Flow Dimensional Flow," AIAA Journal, Vol. 18, No. 2, Using Implicit...assumptions are made in p = density modeling the nonlinear vortex wake structure. Numerical algorithms based on the Euler equations p_ = free stream density

Nonlinear Aerodynamics and the Design of Wing Tips

NASA Technical Reports Server (NTRS)

Kroo, Ilan

1991-01-01

The analysis and design of wing tips for fixed wing and rotary wing aircraft still remains part art, part science. Although the design of airfoil sections and basic planform geometry is well developed, the tip regions require more detailed consideration. This is important because of the strong impact of wing tip flow on wing drag; although the tip region constitutes a small portion of the wing, its effect on the drag can be significant. The induced drag of a wing is, for a given lift and speed, inversely proportional to the square of the wing span. Concepts are proposed as a means of reducing drag. Modern computational methods provide a tool for studying these issues in greater detail. The purpose of the current research program is to improve the understanding of the fundamental issues involved in the design of wing tips and to develop the range of computational and experimental tools needed for further study of these ideas.

FUN3D and CFL3D Computations for the First High Lift Prediction Workshop

NASA Technical Reports Server (NTRS)

Park, Michael A.; Lee-Rausch, Elizabeth M.; Rumsey, Christopher L.

2011-01-01

Two Reynolds-averaged Navier-Stokes codes were used to compute flow over the NASA Trapezoidal Wing at high lift conditions for the 1st AIAA CFD High Lift Prediction Workshop, held in Chicago in June 2010. The unstructured-grid code FUN3D and the structured-grid code CFL3D were applied to several different grid systems. The effects of code, grid system, turbulence model, viscous term treatment, and brackets were studied. The SST model on this configuration predicted lower lift than the Spalart-Allmaras model at high angles of attack; the Spalart-Allmaras model agreed better with experiment. Neglecting viscous cross-derivative terms caused poorer prediction in the wing tip vortex region. Output-based grid adaptation was applied to the unstructured-grid solutions. The adapted grids better resolved wake structures and reduced flap flow separation, which was also observed in uniform grid refinement studies. Limitations of the adaptation method as well as areas for future improvement were identified.

Enhancing the hydrodynamic performance of a tapered swept-back wing through leading-edge tubercles

NASA Astrophysics Data System (ADS)

Wei, Zhaoyu; Lian, Lian; Zhong, Yisen

2018-06-01

The hydrodynamic benefit of implementing leading-edge (LE) tubercles on wings at very low Reynolds numbers ( Res) has not been thoroughly elucidated to date, though their benefits at relatively higher Res are well-studied. Through wind tunnel testing at Re = 5.5 × 104, we found that the LE tubercles increase the lift at all pitch angles tested and slightly reduce the drag at a pitch angle of 4° < α < 10°, which finally results in a significant hydrodynamic performance enhancement at lower pitch angles. Flow visualization reveals that the hydrodynamic performance enhancement is due to the favourable attached flows downstream of the tubercle peaks. The attached flows are believed to be closely related to the downwash and momentum exchange within the boundary layers, which originate from surface and streamwise-aligned counter-rotating vortex pairs (CVPs).

An experimental analysis of critical factors involved in the breakdown process of leading edge vortex flows. Ph.D. Thesis

NASA Technical Reports Server (NTRS)

Visser, Kenneth D.

1991-01-01

Experimental crosswire measurements of the flowfield above a 70 and 75 degree flat plate delta wing were performed at a Reynolds number of 250,000. Survey grids were taken normal to the platform at a series of chordwise locations for angles of attack of 20 and 30 degrees. Axial and azimuthal vorticity distributions were derived from the velocity fields. The dependence of circulation on distance from the vortex core as well as on chordwise location was examined. The effects of nondimensionalization in comparison with other experimental data was made. The circulation distribution scales with the local semispan and grows approximately linearly in the chordwise direction. For regions of the flow outside of the vortex subcore, the circulation at any chordwise station was observed to vary logarithmically with distance from the vortex axis. The circulation was also found to increase linearly with angle of incidence at a given chordwise station. A reduction in the local circulation about the vortex axis occurred at breakdown. The spanwise distribution of axial vorticity was severely altered through the breakdown region and the spanwise distribution of axial vorticity present appeared to reach a maximum immediately preceding breakdown. The local concentration of axial vorticity about the vortex axis was reduced while the magnitude of the azimuthal vorticity decreased throughout the breakdown zone. The axial vorticity components with a negative sense, found in the secondary vortex, remained unaffected by changes in wing sweep or angle of attack, in direct contrast to the positive components. The inclusion of the local wing geometry into a previously derived correlation parameter indicated that the circulation of growing leading edge vortex flows were similar at corresponding radii from the vortex axis. It was concluded that the flow over a delta wing, upstream of the breakdown regions and away from the apex and trailing edge regions, is conical. In addition, the dominating factors leading to the onset of breakdown are felt to be the local circulation of the vortex and the accompanying pressure field.

A computational study on the influence of insect wing geometry on bee flight mechanics

PubMed Central

Feaster, Jeffrey; Bayandor, Javid

2017-01-01

ABSTRACT Two-dimensional computational fluid dynamics (CFD) is applied to better understand the effects of wing cross-sectional morphology on flow field and force production. This study investigates the influence of wing cross-section on insect scale flapping flight performance, for the first time, using a morphologically representative model of a bee (Bombus pensylvanicus) wing. The bee wing cross-section was determined using a micro-computed tomography scanner. The results of the bee wing are compared with flat and elliptical cross-sections, representative of those used in modern literature, to determine the impact of profile variation on aerodynamic performance. The flow field surrounding each cross-section and the resulting forces are resolved using CFD for a flight speed range of 1 to 5 m/s. A significant variation in vortex formation is found when comparing the ellipse and flat plate with the true bee wing. During the upstroke, the bee and approximate wing cross-sections have a much shorter wake structure than the flat plate or ellipse. During the downstroke, the flat plate and elliptical cross-sections generate a single leading edge vortex, while the approximate and bee wings generate numerous, smaller structures that are shed throughout the stroke. Comparing the instantaneous aerodynamic forces on the wing, the ellipse and flat plate sections deviate progressively with velocity from the true bee wing. Based on the present findings, a simplified cross-section of an insect wing can misrepresent the flow field and force production. We present the first aerodynamic study using a true insect wing cross-section and show that the wing corrugation increases the leading edge vortex formation frequency for a given set of kinematics. PMID:29061734

A computational study on the influence of insect wing geometry on bee flight mechanics.

PubMed

Feaster, Jeffrey; Battaglia, Francine; Bayandor, Javid

2017-12-15

Two-dimensional computational fluid dynamics (CFD) is applied to better understand the effects of wing cross-sectional morphology on flow field and force production. This study investigates the influence of wing cross-section on insect scale flapping flight performance, for the first time, using a morphologically representative model of a bee ( Bombus pensylvanicus ) wing. The bee wing cross-section was determined using a micro-computed tomography scanner. The results of the bee wing are compared with flat and elliptical cross-sections, representative of those used in modern literature, to determine the impact of profile variation on aerodynamic performance. The flow field surrounding each cross-section and the resulting forces are resolved using CFD for a flight speed range of 1 to 5 m/s. A significant variation in vortex formation is found when comparing the ellipse and flat plate with the true bee wing. During the upstroke, the bee and approximate wing cross-sections have a much shorter wake structure than the flat plate or ellipse. During the downstroke, the flat plate and elliptical cross-sections generate a single leading edge vortex, while the approximate and bee wings generate numerous, smaller structures that are shed throughout the stroke. Comparing the instantaneous aerodynamic forces on the wing, the ellipse and flat plate sections deviate progressively with velocity from the true bee wing. Based on the present findings, a simplified cross-section of an insect wing can misrepresent the flow field and force production. We present the first aerodynamic study using a true insect wing cross-section and show that the wing corrugation increases the leading edge vortex formation frequency for a given set of kinematics. © 2017. Published by The Company of Biologists Ltd.

URANS simulations of the tip-leakage cavitating flow with verification and validation procedures

NASA Astrophysics Data System (ADS)

Cheng, Huai-yu; Long, Xin-ping; Liang, Yun-zhi; Long, Yun; Ji, Bin

2018-04-01

In the present paper, the Vortex Identified Zwart-Gerber-Belamri (VIZGB) cavitation model coupled with the SST-CC turbulence model is used to investigate the unsteady tip-leakage cavitating flow induced by a NACA0009 hydrofoil. A qualitative comparison between the numerical and experimental results is made. In order to quantitatively evaluate the reliability of the numerical data, the verification and validation (V&V) procedures are used in the present paper. Errors of numerical results are estimated with seven error estimators based on the Richardson extrapolation method. It is shown that though a strict validation cannot be achieved, a reasonable prediction of the gross characteristics of the tip-leakage cavitating flow can be obtained. Based on the numerical results, the influence of the cavitation on the tip-leakage vortex (TLV) is discussed, which indicates that the cavitation accelerates the fusion of the TLV and the tip-separation vortex (TSV). Moreover, the trajectory of the TLV, when the cavitation occurs, is close to the side wall.

Instantaneous and Time Averaged Flow Fields of Multiple Vortices in the Tip Region of a Ducted Propulsor

NASA Astrophysics Data System (ADS)

Oweis, Ghanem; Steven, Ceccio

2003-11-01

PIV data of the flow field in the immediate vicinity of the trailing edge of a ducted propeller at the tip revealed the existence of multiple vorticity concentrations. The multiple vortices in each instantaneous PIV field were identified and individually characterized. The measurements of the multiple vortices were combined with a Gaussian vortex model to reconstruct the vorticity and velocity fields. The major features of the original experimental field were recovered, and the correlation between the two fields was good. The time averaged field and velocity fluctuations were also measured. We will discuss why the "typical" instantaneous tip vortex and the tip vortex from the time averaged field are substantially different. We attempt to explain the cause of these differences. Knowledge of the instantaneous flow field variability is used to understand the causes of the measured velocity fluctuations. The results from this study have an impact on the understanding of the roll-up of tip vortices, and the dynamics of multiple vortices.

Aerodynamic Inner Workings of Circumferential Grooves in a Transonic Axial Compressor

NASA Technical Reports Server (NTRS)

Hah, Chunill; Mueller, Martin; Schiffer, Heinz-Peter

2007-01-01

The current paper reports on investigations of the fundamental flow mechanisms of circumferential grooves applied to a transonic axial compressor. Experimental results show that the compressor stall margin is significantly improved with the current set of circumferential grooves. The primary focus of the current investigation is to advance understanding of basic flow mechanics behind the observed improvement of stall margin. Experimental data and numerical simulations of a circumferential groove were analyzed in detail to unlock the inner workings of the circumferential grooves in the current transonic compressor rotor. A short length scale stall inception occurs when a large flow blockage is built on the pressure side of the blade near the leading edge and incoming flow spills over to the adjacent blade passage due to this blockage. The current study reveals that a large portion of this blockage is created by the tip clearance flow originating from 20% to 50% chord of the blade from the leading edge. Tip clearance flows originating from the leading edge up to 20% chord form a tip clearance core vortex and this tip clearance core vortex travels radially inward. The tip clearance flows originating from 20% to 50% chord travels over this tip clearance core vortex and reaches to the pressure side. This part of tip clearance flow is of low momentum as it is coming from the casing boundary layer and the blade suction surface boundary layer. The circumferential grooves disturb this part of the tip clearance flow close to the casing. Consequently the buildup of the induced vortex and the blockage near the pressure side of the passage is reduced. This is the main mechanism of the circumferential grooves that delays the formation of blockage near the pressure side of the passage and delays the onset of short length scale stall inception. The primary effect of the circumferential grooves is preventing local blockage near the pressure side of the blade leading edge that directly determines flow spillage around the leading edge. The circumferential grooves do not necessarily reduce the over all blockage built up at the rotor tip section.

Investigation of Unsteady Flow Behavior in Transonic Compressor Rotors with LES and PIV Measurements

NASA Technical Reports Server (NTRS)

Hah, Chunill; Voges, Melanie; Mueller, Martin; Schiffer, Heinz-Peter

2009-01-01

In the present study, unsteady flow behavior in a modern transonic axial compressor rotor is studied in detail with large eddy simulation (LES) and particle image velocimetry (PIV). The main purpose of the study is to advance the current understanding of the flow field near the blade tip in an axial transonic compressor rotor near the stall and peak-efficiency conditions. Flow interaction between the tip leakage vortex and the passage shock is inherently unsteady in a transonic compressor. Casing-mounted unsteady pressure transducers have been widely applied to investigate steady and unsteady flow behavior near the casing. Although many aspects of flow have been revealed, flow structures below the casing cannot be studied with casing-mounted pressure transducers. In the present study, unsteady velocity fields are measured with a PIV system and the measured unsteady flow fields are compared with LES simulations. The currently applied PIV measurements indicate that the flow near the tip region is not steady even at the design condition. This self-induced unsteadiness increases significantly as the compressor rotor operates near the stall condition. Measured data from PIV show that the tip clearance vortex oscillates substantially near stall. The calculated unsteady characteristics of the flow from LES agree well with the PIV measurements. Calculated unsteady flow fields show that the formation of the tip clearance vortex is intermittent and the concept of vortex breakdown from steady flow analysis does not seem to apply in the current flow field. Fluid with low momentum near the pressure side of the blade close to the leading edge periodically spills over into the adjacent blade passage. The present study indicates that stall inception is heavily dependent on unsteady behavior of the flow field near the leading edge of the blade tip section for the present transonic compressor rotor.

An experimental study of the effects of aft blowing on a 3.0 caliber tangent ogive body at high angles of attack

NASA Technical Reports Server (NTRS)

Gittner, Nathan M.; Chokani, Ndaona

1991-01-01

An experimental study of the effects of aft blowing on the forebody vortex asymmetry over a 3.0 caliber tangent ogive body at high angles of attack was conducted. The tip of the ogive body was equipped with a single blowing nozzle whose position could be adjusted. The tests were conducted in a subsonic wind tunnel at laminar flow conditions. The effects of model roll, angle of attack, blowing coefficient, and blowing nozzle axial position were independently studied. Surface pressure measurements and flow visualization results were obtained. Aft blowing was observed to alleviate the degree of vortex asymmetry at all angles of attack. The blowing was found to be more effective at the higher angles of attack. However, proportional control of the degree of vortex asymmetry was not observed, because the initial flowfield was highly asymmetric.

Wing Wake Vortices and Temporal Vortex Pair Instabilities

NASA Astrophysics Data System (ADS)

Williamson, C. H. K.; Leweke, T.; Miller, G. D.

In this presentation we include selected results which have originated from vortex dynamics studies conducted at Cornell, in collaboration with IRPHE, Marseille. These studies concern, in particular, the spatial development of delta wing trailing vortices, and the temporal development of counter-rotating vortex pairs. There are, as might be expected, similarities in the instabilities of both of these basic flows, as shown in our laboratory-scale studies. In the case of the spatial development of vortex pairs in the wake of a delta wing, either in free flight or towed from an XY carriage system in a towing tank, we have found three distinct instability length scales as the trailing vortex pair travels downstream. The first (smallest-scale) instability is found immediately behind the delta wing, and this scales on the thickness of the two shear layers separating from the wing trailing edge. The second (short-wave) instability, at an intermediate distance downstream, scales on the primary vortex core dimensions. The third (long-wave) instability far downstream represents the classical "Crow" instability (Crow, 1970), scaling on the distance between the two primary vortices. By imposing disturbances on the delta wing incident velocity, we find that the long-wave instability is receptive to a range of wavelengths. Our experimental measurements of instability growth rates are compared with theoretical predictions, which are based on the theory of Widnall et al. (1971), and which require, as input, DPIV measurements of axial and circumferential velocity profiles. This represents the first time that theoretical and experimental growth rates have been compared, without the imposition of ad-hoc assumptions regarding the vorticity distribution. The agreement with theory appears to be good. The ease with which a Delta wing may be flown in free flight was demonstrated at the Symposium, using a giant polystyrene triangular wing, launched from the back of the auditorium, and ably caught by Professor Sid Leibovich, in whose honour the Symposium was held. In the case of the temporal growth of vortex pairs, formed by the closing of a pair of long flaps underwater, we find two principal instabilities; namely, a longwavelength Crow instability, and a short-wavelength "elliptic" instability. Comparisons between experiment and theory for the growth rates of the long-wave instability, over a range of perturbed wavelengths, appears to be very good. The vortex pair "pinches off", or reconnects, to form vortex rings in the manner assumed to occur in contrails behind jet aircraft. We discover a symmetry-breaking phase relationship for the short wave disturbances growing in the two vortices, which we 380 C.H.K. Williamson et al. show to be consistent with a kinematic matching condition between the two disturbances. Further results demonstrate that this instability is a manifestation of an elliptic instability, which is here identified for the first time in a real open flow. We therefore refer to this flow as a "cooperative elliptic" instability. The long-term evolution of the flow involves the inception of secondary miniscule vortex pairs, which are perpendicular to the primary vortex pair.

Determination of corrections to flow direction measurements obtained with a wing-tip mounted sensor. M.S. Thesis

NASA Technical Reports Server (NTRS)

Moul, T. M.

1983-01-01

The nature of corrections for flow direction measurements obtained with a wing-tip mounted sensor was investigated. Corrections for the angle of attack and sideslip, measured by sensors mounted in front of each wing tip of a general aviation airplane, were determined. These flow corrections were obtained from both wind-tunnel and flight tests over a large angle-of-attack range. Both the angle-of-attack and angle-of-sideslip flow corrections were found to be substantial. The corrections were a function of the angle of attack and angle of sideslip. The effects of wing configuration changes, small changes in Reynolds number, and spinning rotation on the angle-of-attack flow correction were found to be small. The angle-of-attack flow correction determined from the static wind-tunnel tests agreed reasonably well with the correction determined from flight tests.

A Numerical Study of Cavitation Inception in Complex Flow Fields

DTIC Science & Technology

2007-12-01

field in a tip vortex flow of an open propeller to better describe the interaction between the blade wake and the tip vortex (i.e. the roll-up... WAKE INTERACTION ON CAVITATION INCEPTION IN AN OPEN PROPELLER ................15 2.5 NON-SPHERICAL BUBBLE EFFECTS ON CAVITATION INCEPTION [14,15...18 2.6 STUDY OF CAVITATION INCEPTION NOISE [16,17,18

New Insights on Insect's Silent Flight. Part I: Vortex Dynamics and Wing Morphing

NASA Astrophysics Data System (ADS)

Ren, Yan; Liu, Geng; Dong, Haibo; Geng, Biao; Zheng, Xudong; Xue, Qian

2016-11-01

Insects are capable of conducting silent flights. This is attributed to its specially designed wing material properties for the control of vibration and surface morphing during the flapping flight. In current work, we focus on the roles of dynamic wing morphing on the unsteady vortex dynamics of a cicada in steady flight. A 3D image-based surface reconstruction method is used to obtain kinematical and morphological data of cicada wings from high-quality high-speed videos. The observed morphing wing kinematics is highly complex and a singular value decomposition method is used to decompose the wing motion to several dominant modes with distinct motion features. A high-fidelity immersed-boundary-based flow solver is then used to study the vortex dynamics in details. The results show that vortical structures closely relate to the morphing mode, which plays key role in the development and attachment of leading-edge vortex (LEV), thus helps the silent flapping of the cicada wings. This work is supported by AFOSR FA9550-12-1-0071 and NSF CBET-1313217.

Pressurized air injection in an axial hydro-turbine model for the mitigation of tip leakage cavitation

NASA Astrophysics Data System (ADS)

Rivetti, A.; Angulo, M.; Lucino, C.; Liscia, S.

2015-12-01

Tip leakage vortex cavitation in axial hydro-turbines may cause erosion, noise and vibration. Damage due to cavitation can be found at the tip of the runner blades on the low pressure side and the discharge ring. In some cases, the erosion follows an oscillatory pattern that is related to the number of guide vanes. That might suggest that a relationship exists between the flow through the guide vanes and the tip vortex cavitating core that induces this kind of erosion. On the other hand, it is known that air injection has a beneficial effect on reducing the damage by cavitation. In this paper, a methodology to identify the interaction between guide vanes and tip vortex cavitation is presented and the effect of air injection in reducing this particular kind of erosion was studied over a range of operating conditions on a Kaplan scale model. It was found that air injection, at the expense of slightly reducing the efficiency of the turbine, mitigates the erosive potential of tip leakage cavitation, attenuates the interaction between the flow through the guide vanes and the tip vortex and decreases the level of vibration of the structural components.

Vortical Flow Structures in the Near-Wake of a Heaving Airfoil with Passively Actuated Leading and Trailing Flaps.

NASA Astrophysics Data System (ADS)

Siala, Firas; Totpal, Alexander; Liburdy, James

2015-11-01

The flow physics of flying animals has recently received significant attention, mostly in the context of developing bio-inspired micro air vehicles and oscillating flow energy harvesters. Of particular interest is the understanding of the impact of airfoil flexibility on the flow physics. Research efforts showed that some degree of surface flexibility enhanced the strength and size of the leading edge vortex. In this study, the influence of flexibility on the near-wake dynamics and flow structures is investigated using 2D PIV measurements. The experiments are conducted in a wind tunnel at a Reynolds number of 30,000 and a range of reduced frequencies from 0.09 to 0.2. The flexibility is attained using a torsion rod forming a hinge between the flap and the main wing. Vortex flow structures are visualized using large eddy scale decomposition technique and quantified using swirling strength analysis. It is found that trailing edge flexibility increases the vortex swirling strength compared to a rigid airfoil, whereas leading edge flexibility decreases the swirling strength. Furthermore, the integral length scale determined from the autocorrelation of the velocity fluctuations is found to be approximately equal to the actual vortex size. The vortex convective velocity is shown to be independent of flexibility and oscillation frequency, and it is represented by a trimodal distribution, with peak values at 0.8, 0.95 and 1 times the free stream velocity. Oregon State University.

Experimental and analytical studies of a model helicopter rotor in hover

NASA Technical Reports Server (NTRS)

Caradonna, F. X.; Tung, C.

1981-01-01

A benchmark test to aid the development of various rotor performance codes was conducted. Simultaneous blade pressure measurements and tip vortex surveys were made for a wide range of tip Mach numbers including the transonic flow regime. The measured tip vortex strength and geometry permit effective blade loading predictions when used as input to a prescribed wake lifting surface code. It is also shown that with proper inflow and boundary layer modeling, the supercritical flow regime can be accurately predicted.

Experimental Verification of the Streamline Curvature Numerical Analysis Method Applied to the Flow through an Axial Flow Fan.

DTIC Science & Technology

1980-05-28

Total Deviation Angles and Measured Inlet Axial Velocity . . . . 55 ix LIST OF FIGURES (Continued) Figure Page 19 Points Defining Blade Sections of...distance from leading edge to point of maximum camber along chord line ar tip vortex core radius AVR axial velocity ratio (Vx /V x c chord length CL tip...yaw ceoefficient d longitudinal distance from leading edge to tip vortex calculation point G distance from chord line to maximum camber point K cascade

Rotating coherent flow structures as a source for narrowband tip clearance noise from axial fans

NASA Astrophysics Data System (ADS)

Zhu, Tao; Lallier-Daniels, Dominic; Sanjosé, Marlène; Moreau, Stéphane; Carolus, Thomas

2018-03-01

Noise from axial fans typically increases significantly as the tip clearance is increased. In addition to the broadband tip clearance noise at the design flow rate, narrowband humps also associated with the tip flow are observed in the far-field acoustic spectra at lower flow rate. In this study, both experimental and numerical methods are used to shed more light on the noise generation mechanism of this narrowband tip clearance noise and provide a unified description of this source. Unsteady aeroacoustic predictions with the Lattice-Boltzmann Method (LBM) are successfully compared with experiment. Such a validation allows using LBM data to conduct a detailed modal analysis of the pressure field for detecting rotating coherent flow structures which might be considered as noise sources. As previously found in ring fans the narrowband humps in the far-field noise spectra are found to be related to the tip clearance noise that is generated by an interaction of coherent flow structures present in the tip region with the leading edge of the impeller blades. The visualization of the coherent structures shows that they are indeed part of the unsteady tip clearance vortex structures. They are hidden in a complex, spatially and temporally inhomogeneous flow field, but can be recovered by means of appropriate filtering techniques. Their pressure trace corresponds to the so-called rotational instability identified in previous turbomachinery studies, which brings a unified picture of this tip-noise phenomenon for the first time.

Detailed flow-field measurements over a 75 deg swept delta wing

NASA Technical Reports Server (NTRS)

Kjelgaard, Scott O.; Sellers, William L., III

1990-01-01

Results from an experimental investigation documenting the flowfield over a 75 deg swept delta wing at an angle-of-attack of 20.5 deg are presented. Results obtained include surface flow visualization, off-body flow visualization, and detailed flowfield surveys for various Reynolds numbers. Flowfield surveys at Reynolds numbers of 0.5, 1.0, and 1.5 million based on the root chord were conducted with both a Pitot pressure probe and a 5-hole pressure probe; and 3-component laser velocimeter surveys were conducted at a Reynolds number of 1.0 million. The Pitot pressure surveys were obtained at 5 chordwise stations, the 5-hole probe surveys were obtained at 3 chordwise stations and the laser velocimeter surveys were obtained at one station. The results confirm the classical roll up of the flow into a pair of primary vortices over the delta wing. The velocity measurements indicate that Reynolds number has little effect on the global structure of the flowfield for the Reynolds number range investigated. Measurements of the non-dimensional axial velocity in the core of the vortex indicate a jet like flow with values greater than twice freestream. Comparisons between velocity measurements from the 5-hole pressure probe and the laser velocimeter indicate that the pressure probe does a reasonable job of measuring the flowfield quantities where the velocity gradients in the flowfield are low.

A Computational and Experimental Study of Nonlinear Aspects of Induced Drag

NASA Technical Reports Server (NTRS)

Smith, Stephen C.

1996-01-01

Despite the 80-year history of classical wing theory, considerable research has recently been directed toward planform and wake effects on induced drag. Nonlinear interactions between the trailing wake and the wing offer the possibility of reducing drag. The nonlinear effect of compressibility on induced drag characteristics may also influence wing design. This thesis deals with the prediction of these nonlinear aspects of induced drag and ways to exploit them. A potential benefit of only a few percent of the drag represents a large fuel savings for the world's commercial transport fleet. Computational methods must be applied carefully to obtain accurate induced drag predictions. Trefftz-plane drag integration is far more reliable than surface pressure integration, but is very sensitive to the accuracy of the force-free wake model. The practical use of Trefftz plane drag integration was extended to transonic flow with the Tranair full-potential code. The induced drag characteristics of a typical transport wing were studied with Tranair, a full-potential method, and A502, a high-order linear panel method to investigate changes in lift distribution and span efficiency due to compressibility. Modeling the force-free wake is a nonlinear problem, even when the flow governing equation is linear. A novel method was developed for computing the force-free wake shape. This hybrid wake-relaxation scheme couples the well-behaved nature of the discrete vortex wake with viscous-core modeling and the high-accuracy velocity prediction of the high-order panel method. The hybrid scheme produced converged wake shapes that allowed accurate Trefftz-plane integration. An unusual split-tip wing concept was studied for exploiting nonlinear wake interaction to reduced induced drag. This design exhibits significant nonlinear interactions between the wing and wake that produced a 12% reduction in induced drag compared to an equivalent elliptical wing at a lift coefficient of 0.7. The performance of the split-tip wing was also investigated by wing tunnel experiments. Induced drag was determined from force measurements by subtracting the estimated viscous drag, and from an analytical drag-decomposition method using a wake survey. The experimental results confirm the computational prediction.

Helicopter theory

NASA Technical Reports Server (NTRS)

Johnson, W.

1980-01-01

A comprehensive presentation is made of the engineering analysis methods used in the design, development and evaluation of helicopters. After an introduction covering the fundamentals of helicopter rotors, configuration and operation, rotary wing history, and the analytical notation used in the text, the following topics are discussed: (1) vertical flight, including momentum, blade element and vortex theories, induced power, vertical drag and ground effect; (2) forward flight, including in addition to momentum and vortex theory for this mode such phenomena as rotor flapping and its higher harmonics, tip loss and root cutout, compressibility and pitch-flap coupling; (3) hover and forward flight performance assessment; (4) helicopter rotor design; (5) rotary wing aerodynamics; (6) rotary wing structural dynamics, including flutter, flap-lag dynamics ground resonance and vibration and loads; (7) helicopter aeroelasticity; (8) stability and control (flying qualities); (9) stall; and (10) noise.

VORCAM: A computer program for calculating vortex lift effect of cambered wings by the suction analogy

NASA Technical Reports Server (NTRS)

Lan, C. E.; Chang, J. F.

1981-01-01

A user's guide to an improved version of Woodward's chord plane aerodynamic panel computer code is presumed. The guide can be applied to cambered wings exhibiting edge separated flow, including those with leading edge vortex flow at subsonic and supersonic speeds. New orientations for the rotated suction force are employed based on the momentum principal. The supersonic suction analogy method is improved by using an effective angle of attack defined through a semiempirical method.

Experimental Study of Tip Vortex Flow from a Periodically Pitched Airfoil Section

NASA Technical Reports Server (NTRS)

Zaman, KBMQ; Fagan, A. F.; Mankbadi, M. R.

2016-01-01

An experimental investigation of a tip vortex from a NACA0012 airfoil is conducted in a low-speed wind tunnel at a chord Reynolds number of 4x10(exp 4). Initially, data for a stationary airfoil held at various angles-of-attack (alpha) are gathered. Detailed surveys are done for two cases: alpha=10 deg with attached flow and alpha=25 deg with massive flow separation on the upper surface. Distributions of various properties are obtained using hot-wire anemometry. Data include mean velocity, streamwise vorticity and turbulent stresses at various streamwise locations. For all cases, the vortex core is seen to involve a mean velocity deficit. The deficit apparently traces to the airfoil wake, part of which gets wrapped by the tip vortex. At small alpha, the vortex is laminar within the measurement domain. The strength of the vortex increases with increasing alpha but undergoes a sudden drop around alpha (is) greater than 16 deg. The drop in peak vorticity level is accompanied by transition and a sharp rise in turbulence within the core. Data are also acquired with the airfoil pitched sinusoidally. All oscillation cases pertain to a mean alpha=15 deg while the amplitude and frequency are varied. An example of phase-averaged data for an amplitude of +/-10 deg and a reduced frequency of k=0.2 is discussed. All results are compared with available data from the literature shedding further light on the complex dynamics of the tip vortex.

Effects of Sweep Angle on the Boundary-Layer Stability Characteristics of an Untapered Wing at Low Speeds

NASA Technical Reports Server (NTRS)

Boltz, Frederick W.; Kenyon, George C.; Allen, Clyde Q.

1960-01-01

An investigation was conducted in the Ames 12-Foot Low-Turbulence Pressure Tunnel to determine the effects of sweep on the boundary-layer stability characteristics of an untapered variable-sweep wing having an NACA 64(2)A015 section normal to the leading edge. Pressure distribution and transition were measured on the wing at low speeds at sweep angles of 0, 10, 20, 30, 40, and 50 deg. and at angles of attack from -3 to 3 deg. The investigation also included flow-visualization studies on the surface at sweep angles from 0 to 50 deg. and total pressure surveys in the boundary layer at a sweep angle of 30 deg. for angles of attack from -12 to 0 deg. It was found that sweep caused premature transition on the wing under certain conditions. This effect resulted from the formation of vortices in the boundary layer when a critical combination of sweep angle, pressure gradient, and stream Reynolds number was attained. A useful parameter in indicating the combined effect of these flow variables on vortex formation and on beginning transition is the crossflow Reynolds number. The critical values of crossflow Reynolds number for vortex formation found in this investigation range from about 135 to 190 and are in good agreement with those reported in previous investigations. The values of crossflow Reynolds number for beginning transitions were found to be between 190 and 260. For each condition (i.e., development of vortices and initiation of transition at a given location) the lower values in the specified ranges were obtained with a light coating of flow-visualization material on the surface. A method is presented for the rapid computation of crossflow Reynolds number on any swept surface for which the pressure distribution is known. From calculations based on this method, it was found that the maximum values of crossflow Reynolds number are attained under conditions of a strong pressure gradient and at a sweep angle of about 50 deg. Due to the primary dependence on pressure gradient, effects of sweep in causing premature transition are generally first encountered on the lower surfaces of wings operating at positive angles of attack.

A vortex model for forces and moments on low-aspect-ratio wings in side-slip with experimental validation

PubMed Central

DeVoria, Adam C.

2017-01-01

This paper studies low-aspect-ratio () rectangular wings at high incidence and in side-slip. The main objective is to incorporate the effects of high angle of attack and side-slip into a simplified vortex model for the forces and moments. Experiments are also performed and are used to validate assumptions made in the model. The model asymptotes to the potential flow result of classical aerodynamics for an infinite aspect ratio. The → 0 limit of a rectangular wing is considered with slender body theory, where the side-edge vortices merge into a vortex doublet. Hence, the velocity fields transition from being dominated by a spanwise vorticity monopole ( ≫ 1) to a streamwise vorticity dipole ( ∼ 1). We theoretically derive a spanwise loading distribution that is parabolic instead of elliptic, and this physically represents the additional circulation around the wing that is associated with reattached flow. This is a fundamental feature of wings with a broad-facing leading edge. The experimental measurements of the spanwise circulation closely approximate a parabolic distribution. The vortex model yields very agreeable comparison with direct measurement of the lift and drag, and the roll moment prediction is acceptable for ≤ 1 prior to the roll stall angle and up to side-slip angles of 20°. PMID:28293139

Numerical and experimental investigations on unsteady aerodynamics of flapping wings

NASA Astrophysics Data System (ADS)

Yu, Meilin

The development of a dynamic unstructured grid high-order accurate spectral difference (SD) method for the three dimensional compressible Navier-Stokes (N-S) equations and its applications in flapping-wing aerodynamics are carried out in this work. Grid deformation is achieved via an algebraic blending strategy to save computational cost. The Geometric Conservation Law (GCL) is imposed to ensure that grid deformation will not contaminate the flow physics. A low Mach number preconditioning procedure is conducted in the developed solver to handle the bio-inspired flow. The capability of the low Mach number preconditioned SD solver is demonstrated by a series of two dimensional (2D) and three dimensional (3D) simulations of the unsteady vortex dominated flow. Several topics in the flapping wing aerodynamics are numerically and experimentally investigated in this work. These topics cover some of the cutting-edge issues in flapping wing aerodynamics, including the wake structure analysis, airfoil thickness and kinematics effects on the aerodynamic performances, vortex structure analysis around 3D flapping wings and the kinematics optimization. Wake structures behind a sinusoidally pitching NACA0012 airfoil are studied with both experimental and numerical approaches. The experiments are carried out with Particle Image Velocimetry (PIV) and two types of wake transition processes, namely the transition from a drag-indicative wake to a thrust-indicative wake and that from the symmetric wake to the asymmetric wake are distinguished. The numerical results from the developed SD solver agree well with the experimental results. It is numerically found that the deflective direction of the asymmetric wake is determined by the initial conditions, e.g. initial phase angle. As most insects use thin wings (i. e., wing thickness is only a few percent of the chord length) in flapping flight, the effects of airfoil thickness on thrust generation are numerically investigated by simulating the flow fields around a series of plunging NACA symmetric airfoils with thickness ratio ranging from 4.0% to 20.0% of the airfoil chord length. The contribution of viscous force to flapping propulsion is accessed and it is found that viscous force becomes thrust producing, instead of drag producing, and plays a non-negligible role in thrust generation for thin airfoils. This is closely related to the variations of the dynamics of the unsteady vortex structures around the plunging airfoils. As nature flyers use complex wing kinematics in flapping flight, kinematics effects on the aerodynamic performance with different airfoil thicknesses are numerically studied by using a series of NACA symmetric airfoils. It is found that the combined plunging and pitching motion can outperform the pure plunging or pitching motion by sophisticatedly adjusting the airfoil gestures during the oscillation stroke. The thin airfoil better manipulates leading edge vortices (LEVs) than the thick airfoil (NACA0030) does in studied cases, and there exists an optimal thickness for large thrust generation with reasonable propulsive efficiency. With the present kinematics and dynamic parameters, relatively low reduced frequency is conducive for thrust production and propulsive efficiency for all tested airfoil thicknesses. In order to obtain the optimal kinematics parameters of flapping flight, a kinematics optimization is then performed. A gradient-based optimization algorithm is coupled with a second-order SD Navier-Stokes solver to search for the optimal kinematics of a certain airfoil undergoing a combined plunging and pitching motion. Then a high-order SD scheme is used to verify the optimization results and reveal the detailed vortex structures associated with the optimal kinematics of the flapping flight. It is found that for the case with maximum propulsive efficiency, there exists no leading edge separation during most of the oscillation cycle. In order to provide constructive suggestions to the design of micro-air-vehicles (MAVs), 3D simulations of the flapping wings are carried out in this work. Both the rectangular and bio-inspired wings with different kinematics are investigated. The formation process of two-jet-like wake patterns behind the finite-span flapping wing is found to be closely related to the interaction between trailing edge vortices and tip vortices. Then the effects of the wing planforms on the aerodynamics performance of the finite-span flapping wings are elucidated in terms of the evolution and dynamic interaction of unsteady vortex structures.

Pitching effect on transonic wing stall of a blended flying wing with low aspect ratio

NASA Astrophysics Data System (ADS)

Tao, Yang; Zhao, Zhongliang; Wu, Junqiang; Fan, Zhaolin; Zhang, Yi

2018-05-01

Numerical simulation of the pitching effect on transonic wing stall of a blended flying wing with low aspect ratio was performed using improved delayed detached eddy simulation (IDDES). To capture the discontinuity caused by shock wave, a second-order upwind scheme with Roe’s flux-difference splitting is introduced into the inviscid flux. The artificial dissipation is also turned off in the region where the upwind scheme is applied. To reveal the pitching effect, the implicit approximate-factorization method with sub-iterations and second-order temporal accuracy is employed to avoid the time integration of the unsteady Navier-Stokes equations solved by finite volume method at Arbitrary Lagrange-Euler (ALE) form. The leading edge vortex (LEV) development and LEV circulation of pitch-up wings at a free-stream Mach number M = 0.9 and a Reynolds number Re = 9.6 × 106 is studied. The Q-criterion is used to capture the LEV structure from shear layer. The result shows that a shock wave/vortex interaction is responsible for the vortex breakdown which eventually causes the wing stall. The balance of the vortex strength and axial flow, and the shock strength, is examined to provide an explanation of the sensitivity of the breakdown location. Pitching motion has great influence on shock wave and shock wave/vortex interactions, which can significantly affect the vortex breakdown behavior and wing stall onset of low aspect ratio blended flying wing.

DoD High Performance Computing Modernization Program FY16 Annual Report

DTIC Science & Technology

2018-05-02

vortex shedding from rotor blade tips using adaptive mesh refinement gives Helios the unique capability to assess the interaction of these vortices...with the fuselage and nearby rotor blades . Helios provides all the benefits for rotary-winged aircraft that Kestrel does for fixed-wing aircraft...rotor blade upgrade of the CH-47F Chinook helicopter to achieve up to an estimated 2,000 pounds increase in hover thrust (~10%) with limited

Vortex leading edge flap assembly for supersonic airplanes

NASA Technical Reports Server (NTRS)

Rudolph, Peter K. C. (Inventor)

1997-01-01

A leading edge flap (16) for supersonic transport airplanes is disclosed. In its stowed position, the leading edge flap forms the lower surface of the wing leading edge up to the horizontal center of the leading edge radius. For low speed operation, the vortex leading edge flap moves forward and rotates down. The upward curve of the flap leading edge triggers flow separation on the flap and rotational flow on the upper surface of the flap (vortex). The rounded shape of the upper fixed leading edge provides the conditions for a controlled reattachment of the flow on the upper wing surface and therefore a stable vortex. The vortex generates lift and a nose-up pitching moment. This improves maximum lift at low speed, reduces attitude for a given lift coefficient and improves lift to drag ratio. The mechanism (27) to move the vortex flap consists of two spanwise supports (24) with two diverging straight tracks (64 and 68) each and a screw drive mechanism (62) in the center of the flap panel (29). The flap motion is essentially normal to the airloads and therefore requires only low actuation forces.

New techniques for experimental generation of two-dimensional blade-vortex interaction at low Reynolds numbers

NASA Technical Reports Server (NTRS)

Booth, E., Jr.; Yu, J. C.

1986-01-01

An experimental investigation of two dimensional blade vortex interaction was held at NASA Langley Research Center. The first phase was a flow visualization study to document the approach process of a two dimensional vortex as it encountered a loaded blade model. To accomplish the flow visualization study, a method for generating two dimensional vortex filaments was required. The numerical study used to define a new vortex generation process and the use of this process in the flow visualization study were documented. Additionally, photographic techniques and data analysis methods used in the flow visualization study are examined.

Vorticity Transfer in Shock Wave Interactions with Turbulence and Vortices

NASA Astrophysics Data System (ADS)

Agui, J. H.; Andreopoulos, J.

1998-11-01

Time-dependent, three-dimensional vorticity measurements of shock waves interacting with grid generated turbulence and concentrated tip vortices were conducted in a large diameter shock tube facility. Two different mesh size grids and a NACA-0012 semi-span wing acting as a tip vortex generator were used to carry out different relative Mach number interactions. The turbulence interactions produced a clear amplification of the lateral and spanwise vorticity rms, while the longitudinal component remained mostly unaffected. By comparison, the tip vortex/shock wave interactions produced a two fold increase in the rms of longitudinal vorticity. Considerable attention was given to the vorticity source terms. The mean and rms of the vorticity stretching terms dominated by 5 to 7 orders of magnitude over the dilitational compression terms in all the interactions. All three signals of the stretching terms manifested very intermittent, large amplitude peak events which indicated the bursting character of the stretching process. Distributions of these signals were characterized by extremely large levels of flatness with varying degrees of skewness. These distribution patterns were found to change only slightly through the turbulence interactions. However, the tip vortex/shock wave interactions brought about significant changes in these distributions which were associated with the abrupt structural changes of the vortex after the interaction.

A computer program for calculating aerodynamic characteristics of low aspect-ratio wings with partial leading-edge separation

NASA Technical Reports Server (NTRS)

Mehrotra, S. C.; Lan, C. E.

1978-01-01

The necessary information for using a computer program to predict distributed and total aerodynamic characteristics for low aspect ratio wings with partial leading-edge separation is presented. The flow is assumed to be steady and inviscid. The wing boundary condition is formulated by the Quasi-Vortex-Lattice method. The leading edge separated vortices are represented by discrete free vortex elements which are aligned with the local velocity vector at midpoints to satisfy the force free condition. The wake behind the trailing edge is also force free. The flow tangency boundary condition is satisfied on the wing, including the leading and trailing edges. The program is restricted to delta wings with zero thickness and no camber. It is written in FORTRAN language and runs on CDC 6600 computer.

System Identification of a Vortex Lattice Aerodynamic Model

NASA Technical Reports Server (NTRS)

Juang, Jer-Nan; Kholodar, Denis; Dowell, Earl H.

2001-01-01

The state-space presentation of an aerodynamic vortex model is considered from a classical and system identification perspective. Using an aerodynamic vortex model as a numerical simulator of a wing tunnel experiment, both full state and limited state data or measurements are considered. Two possible approaches for system identification are presented and modal controllability and observability are also considered. The theory then is applied to the system identification of a flow over an aerodynamic delta wing and typical results are presented.

Applicability of linearized-theory attached-flow methods to design and analysis of flap systems at low speeds for thin swept wings with sharp leading edges

NASA Technical Reports Server (NTRS)

Carlson, Harry W.; Darden, Christine M.

1987-01-01

Low-speed experimental force and data on a series of thin swept wings with sharp leading edges and leading and trailing-edge flaps are compared with predictions made using a linearized-theory method which includes estimates of vortex forces. These comparisons were made to assess the effectiveness of linearized-theory methods for use in the design and analysis of flap systems in subsonic flow. Results demonstrate that linearized-theory, attached-flow methods (with approximate representation of vortex forces) can form the basis of a rational system for flap design and analysis. Even attached-flow methods that do not take vortex forces into account can be used for the selection of optimized flap-system geometry, but design-point performance levels tend to be underestimated unless vortex forces are included. Illustrative examples of the use of these methods in the design of efficient low-speed flap systems are included.

Numerical Analysis of Incipient Separation on 53 Deg Swept Diamond Wing

NASA Technical Reports Server (NTRS)

Frink, Neal T.

2015-01-01

A systematic analysis of incipient separation and subsequent vortex formation from moderately swept blunt leading edges is presented for a 53 deg swept diamond wing. This work contributes to a collective body of knowledge generated within the NATO/STO AVT-183 Task Group titled 'Reliable Prediction of Separated Flow Onset and Progression for Air and Sea Vehicles'. The objective is to extract insights from the experimentally measured and numerically computed flow fields that might enable turbulence experts to further improve their models for predicting swept blunt leading-edge flow separation. Details of vortex formation are inferred from numerical solutions after establishing a good correlation of the global flow field and surface pressure distributions between wind tunnel measurements and computed flow solutions. From this, significant and sometimes surprising insights into the nature of incipient separation and part-span vortex formation are derived from the wealth of information available in the computational solutions.

Wind Tunnel Testing of a 120th Scale Large Civil Tilt-Rotor Model in Airplane and Helicopter Modes

NASA Technical Reports Server (NTRS)

Theodore, Colin R.; Willink, Gina C.; Russell, Carl R.; Amy, Alexander R.; Pete, Ashley E.

2014-01-01

In April 2012 and October 2013, NASA and the U.S. Army jointly conducted a wind tunnel test program examining two notional large tilt rotor designs: NASA's Large Civil Tilt Rotor and the Army's High Efficiency Tilt Rotor. The approximately 6%-scale airframe models (unpowered) were tested without rotors in the U.S. Army 7- by 10-foot wind tunnel at NASA Ames Research Center. Measurements of all six forces and moments acting on the airframe were taken using the wind tunnel scale system. In addition to force and moment measurements, flow visualization using tufts, infrared thermography and oil flow were used to identify flow trajectories, boundary layer transition and areas of flow separation. The purpose of this test was to collect data for the validation of computational fluid dynamics tools, for the development of flight dynamics simulation models, and to validate performance predictions made during conceptual design. This paper focuses on the results for the Large Civil Tilt Rotor model in an airplane mode configuration up to 200 knots of wind tunnel speed. Results are presented with the full airframe model with various wing tip and nacelle configurations, and for a wing-only case also with various wing tip and nacelle configurations. Key results show that the addition of a wing extension outboard of the nacelles produces a significant increase in the lift-to-drag ratio, and interestingly decreases the drag compared to the case where the wing extension is not present. The drag decrease is likely due to complex aerodynamic interactions between the nacelle and wing extension that results in a significant drag benefit.

Visualizing Time-Varying Phenomena In Numerical Simulations Of Unsteady Flows

NASA Technical Reports Server (NTRS)

Lane, David A.

1996-01-01

Streamlines, contour lines, vector plots, and volume slices (cutting planes) are commonly used for flow visualization. These techniques are sometimes referred to as instantaneous flow visualization techniques because calculations are based on an instant of the flowfield in time. Although instantaneous flow visualization techniques are effective for depicting phenomena in steady flows,they sometimes do not adequately depict time-varying phenomena in unsteady flows. Streaklines and timelines are effective visualization techniques for depicting vortex shedding, vortex breakdown, and shock waves in unsteady flows. These techniques are examples of time-dependent flow visualization techniques, which are based on many instants of the flowfields in time. This paper describes the algorithms for computing streaklines and timelines. Using numerically simulated unsteady flows, streaklines and timelines are compared with streamlines, contour lines, and vector plots. It is shown that streaklines and timelines reveal vortex shedding and vortex breakdown more clearly than instantaneous flow visualization techniques.

A computer program to calculate the longitudinal aerodynamic characteristics of upper-surface-blown wing-flap configurations

NASA Technical Reports Server (NTRS)

Mendenhall, M. R.

1978-01-01

A user's manual is presented for a computer program in which a vortex-lattice lifting-surface method is used to model the wing and multiple flaps. The engine wake model consists of a series of closely spaced vortex rings with rectangular cross sections. The jet wake is positioned such that the lower boundary of the jet is tangent to the wing and flap upper surfaces. The two potential flow models are used to calculate the wing-flap loading distribution including the influence of the wakes from up to two engines on the semispan. The method is limited to the condition where the flow and geometry of the configurations are symmetric about the vertical plane containing the wing root chord. The results include total configuration forces and moments, individual lifting-surface load distributions, pressure distributions, flap hinge moments, and flow field calculation at arbitrary field points. The use of the program, preparation of input, the output, program listing, and sample cases are described.

Oil-flow study of a Space Shuttle orbiter tip-fin controller

NASA Technical Reports Server (NTRS)

Helms, V. T., III

1983-01-01

Possible use of tip-fin controllers instead of a vertical tail on advanced winged entry vehicles was examined. Elimination of the vertical tail and using tip-fins offers the advantages of positive yaw control at high angles of attack and a potential weight savings. Oil-flow technique was used to obtain surface flow patterns on a tip-fin installed on a 0.01-scale Space Shuttle orbiter model for the purpose of assessing the extent of flow interference effects on the wing and tip-fin which might lead to serious heating problems. Tests were conducted in air at Mach 10 for a free-stream Reynolds numbers of .000113 at 20, 30, and 40 degree angle of attack and sideslip angles of 0 and 2 degree. Elevon deflections of -10, 0, and 10 degree and tip-fin control-surface deflections of 0, 20, and 40 degree were employed. Test results were also used to aid in the interpretation of heating data obtained on a Shuttle orbiter tip-fin on another model in a different facility. A limited comparison of oil-flow patterns and heat-transfer data is included. It was determined that elevon deflection angles from -10 to 10 degree and sideslip angles up to 2 degree have very little effect on tip-fin surface flow patterns. Also, there is a minimum of interference between the tip-fin and the wing. The most significant flow interactions occur on the tip-fin onboard surface as a result of its control-surface deflections.

Tactical missile aerodynamics - General topics. Progress in Astronautics and Aeronautics. Vol. 141

DOE Office of Scientific and Technical Information (OSTI.GOV)

Hemsch, M.J.

1992-01-01

The present volume discusses the development history of tactical missile airframes, aerodynamic considerations for autopilot design, a systematic method for tactical missile design, the character and reduction of missile observability by radar, the visualization of high angle-of-attack flow phenomena, and the behavior of low aspect ratio wings at high angles of attack. Also discussed are airbreathing missile inlets, 'waverider' missile configurations, bodies with noncircular cross-sections and bank-to-turn missiles, asymmetric flow separation and vortex shedding on bodies-of-revolution, unsteady missile flows, swept shock-wave/boundary-layer interactions, pylon carriage and separation of stores, and internal stores carriage and separation.

Application of Circulation Control Technology to Airframe Noise Reduction

NASA Technical Reports Server (NTRS)

Ahuja, K. K.; Sankar, L. N.; Englar, R. J.; Munro, Scott E.; Li, Yi; Gaeta, R. J.

2003-01-01

This report is a summary of the work performed by Georgia Tech Research Institute (GTRI) under NASA Langley Grant NAG-1-2146, which was awarded as a part of NASA's Breakthrough Innovative Technologies (BIT) initiative. This was a three-year program, with a one-year no-cost extension. Each year's study has been an integrated effort consisting of computational fluid dynamics, experimental aerodynamics, and detailed noise and flow measurements. Year I effort examined the feasibility of reducing airframe noise by replacing the conventional wing systems with a Circulation Control Wing (CCW), where steady blowing was used through the trailing edge of the wing over a Coanda surface. It was shown that the wing lift increases with CCW blowing and indeed for the same lift, a CCW wing was shown to produce less noise. Year 2 effort dealt with a similar study on the role of pulsed blowing on airframe noise. The main objective of this portion of the study was to assess whether pulse blowing from the trailing edge of a CCW resulted in more, less, or the same amount of radiated noise to the farfield. Results show that a reduction in farfield noise of up to 5 dB is measured when pulse flow is compared with steady flow for an equivalent lift configuration. This reduction is in the spectral region associated with the trailing edge jet noise. This result is due to the unique advantage that pulsed flow has over steady flow. For a range of frequencies, more lift is experienced with the same mass flow as the steady case. Thus, for an equivalent lift and slot height, the pulsed system can operate at lower jet velocities, and hence lower jet noise. The computational analysis showed that for a given time-averaged mass flow rate, pulsed jets give a higher value of C(sub l) and a higher L/D than equivalent steady jets. This benefit is attributable to higher instantaneous jet velocities, and higher instantaneous C(sub mu) values for the pulsed jet. Pulsed jet benefits increase at higher frequencies. However, these advantages are somewhat offset by the unsteadiness in the loads, which will cause structural vibrations and fatigue. Additional studies must be done, perhaps with multiple jets on the upper and lower surfaces, to smooth out the fluctuations in lift while retaining the benefits. The rest of the effort was devoted to examining ways of reducing flap edge noise by blowing air through a Coanda nozzle over a rounded tip of the flap. In this case, we were successful in moving the tip vortex away from the tip, but the device producing the blowing was noisy and we were unable to examine the noise benefits, although we believe that the movement of the tip vortex far from the tip should provide noise benefits. It should be noted that in an effort to understand the fluid dynamics and the aeroacoustics of a jet blowing over a Coanda surface, we also carried out a very extensive study of the high aspect ratio slot jets. A first-ever set of far-field noise spectra were measured for jets exhausting from slots with aspect ratios in the range 100 to 3000. Parallel measurements of velocity profiles, length scales and convection velocities were measured to understand the noise generation of high aspect ratio jets. Attempts were also made to develop jet noise prediction schemes for such jets. Much of the work done under this effort has been described in five conference papers and two doctoral theses. The first year s work on the use of steady blowing was described in two AIAA papers presented at the 2001 AIAA Aerospace Sciences Meeting in Reno. Subsequent work was presented at the 9th AIMCEAS Aeroacoustics Conference and Exhibit held at Hilton Head May 12-13. Another paper is to be presented at the 2004 AIAA Aerospace Sciences Meeting in Reno in January 2004. All six papers are included with this report as Appendices. The bulk of the experimental work done in an effort to produce a pulsed flow that is free of upstream noise is also attached as an Appendix.

Experimental study of a generic high-speed civil transport: Tabulated data

NASA Technical Reports Server (NTRS)

Belton, Pamela S.; Campbell, Richard L.

1992-01-01

An experimental study of a generic high-speed civil transport was conducted in LaRC's 8-Foot Transonic Pressure Tunnel. The data base was obtained for the purpose of assessing the accuracy of various levels of computational analysis. Two models differing only in wing tip geometry were tested with and without flow-through nacelles. The baseline model has a curved or crescent wing tip shape while the second model has a more conventional straight wing tip shape. The study was conducted at Mach numbers from 0.30-1.19. Force data were obtained on both the straight and curved wing tip models. Only the curved wing tip model was instrumented for measuring pressures. Longitudinal and lateral-directional aerodynamic data are presented without analysis in tabulated form. Pressure coefficients for the curved wing tip model are also presented in tabulated form.

Physical mechanisms of longitudinal vortexes formation, appearance of zones with high heat fluxes and early transition in hypersonic flow over delta wing with blunted leading edges

NASA Astrophysics Data System (ADS)

Alexandrov, S. V.; Vaganov, A. V.; Shalaev, V. I.

2016-10-01

Processes of vortex structures formation and they interactions with the boundary layer in the hypersonic flow over delta wing with blunted leading edges are analyzed on the base of experimental investigations and numerical solutions of Navier-Stokes equations. Physical mechanisms of longitudinal vortexes formation, appearance of abnormal zones with high heat fluxes and early laminar turbulent transition are studied. These phenomena were observed in many high-speed wind tunnel experiments; however they were understood only using the detailed analysis of numerical modeling results with the high resolution. Presented results allowed explaining experimental phenomena. ANSYS CFX code (the DAFE MIPT license) on the grid with 50 million nodes was used for the numerical modeling. The numerical method was verified by comparison calculated heat flux distributions on the wing surface with experimental data.

Experimental framework to study tip vortex interactions in multirotor wakes

NASA Astrophysics Data System (ADS)

Yao, Rongnan; Araya, Daniel

2017-11-01

We present an experimental study to compare the dynamic characteristics of tip vortices shed from a propeller in a crossflow to similar characteristics of an isolated vortex column generated in a closed system. Our aim is to evaluate the feasibility of using this simple isolated system to study the more complicated three-dimensional vortex interactions inherent to multirotor wakes, where the local unsteadiness generated by one rotor can strongly impact the performance of nearby rotors. Time-resolved particle image velocimetry is used to measure the velocity field of the propeller wake flow in a wind tunnel and the vortex column in a water tank. Specific attention is placed on analyzing the observed vortex core precession in the isolated system and comparing this to characteristic tip-vortex wandering phenomenon.

B-47A on ramp

NASA Technical Reports Server (NTRS)

1953-01-01

Boeing B-47A (NACA 150) shown on the ramp near NACA High-Speed Flight Research Station at South Base of Edwards Air Force Base, California, in 1953. The B-47A Stratojet's wing is mounted high on the fuselage with a sweep back of 36 degrees and a span of 116 feet, with wing vortex generators installed. A two engine pod under each wing, and an additional engine pod at each wing tip using General Electric J-47-GE-23 turbojets. The airplane is fitted with a nose boom for measuring airspeed, altitude, angle-of-attack and angle-of-sideslip, and an optigraph for measuring the movements of target lights on the wing and tail.

Near wakes of advanced turbopropellers

NASA Technical Reports Server (NTRS)

Hanson, D. B.; Patrick, W. P.

1989-01-01

The flow in the wake of a model single rotation Prop-Fan rotor operating in a wind tunnel was traversed with a hot-wire anemometer system designed to determine the 3 periodic velocity components. Special data acquisition and data reduction methods were required to deal with the high data frequency, narrow wakes, and large fluctuating air angles in the tip vortex region. The model tip helical Mach number was 1.17, simulating the cruise condition. Although the flow field is complex, flow features such as viscous velocity defects, vortex sheets, tip vortices, and propagating acoustic pulses are clearly identified with the aid of a simple analytical wake theory.

Dynamics of Isolated Tip Vortex Cavitation

NASA Astrophysics Data System (ADS)

Pennings, Pepijn; Bosschers, Johan; van Terwisga, Tom

2014-11-01

Performance of ship propellers and comfort levels in the surroundings are limited by various forms of cavitation. Amongst these forms tip vortex cavitation is one of the first appearing forms and is expected to be mainly responsible for the emission of broadband pressure fluctuations typically occurring between the 4th to the 7th blade passing frequency (approx. 40--70 Hz). These radiated pressure pulses are likely to excite parts of the hull structure resulting in a design compromise between efficiency and comfort. Insight is needed in the mechanism of acoustic emission from the oscillations by a tip vortex cavity. In the current experimental study the tip vortex cavity from a blade with an elliptic planform and sections based on NACA 662 - 415 with meanline a = 0 . 8 is observed using high speed shadowgraphy in combination with blade force and acoustic measurements. An analytic model describing three main cavity deformation modes is verified and used to explain the origin of a cavity eigenfrequency or ``vortex singing'' phenomenon observed by Maines and Arndt (1997) on the tip vortex cavity originating from the same blade. As no hydrodynamic sound originating from the tip vortex cavity was observed it is posed that a tip flow instability is essential for ``vortex singing.'' This research was funded by the Lloyd's Register Foundation as part of the International Institute for Cavitation Research.

Effect of cavitation on flow structure of a tip vortex

NASA Astrophysics Data System (ADS)

Matthieu, Dreyer; Reclari, Martino; Farhat, Mohamed

2013-11-01

Tip vortices, which may develop in axial turbines and marine propellers, are often associated with the occurrence of cavitation because of the low pressure in their core. Although this issue has received a great deal of attention, it is still unclear how the phase transition affects the flow structure of such a vortex. In the present work, we investigate the change of the vortex structure due to cavitation incipience. The measurement of the velocity field is performed in the case of a tip vortex generated by an elliptical hydrofoil placed in the test section of EPFL high speed cavitation tunnel. To this end, a 3D stereo PIV is used with fluorescent seeding particles. A cost effective method is developed to produce in-house fluorescent seeding material, based on polyamide particles and Rhodamine-B dye. The amount of cavitation in the vortex core is controlled by the inlet pressure in the test section, starting with the non-cavitating case. We present an extensive analysis of the vorticity distribution, the vortex intensity and core size for various cavitation developments. This research is supported by CCEM and swisselectric research.

F/A-18 1/9th scale model tail buffet measurements

NASA Technical Reports Server (NTRS)

Martin, C. A.; Glaister, M. K.; Maclaren, L. D.; Meyn, L. A.; Ross, J.

1991-01-01

Wind tunnel tests were carried out on a 1/9th scale model of the F/A-18 at high angles of attack to investigate the characteristics of tail buffet due to bursting of the wing leading edge extension (LEX) vortices. The tests were carried out at the Aeronautical Research Laboratory low-speed wind tunnel facility and form part of a collaborative activity with NASA Ames Research Center, organized by The Technical Cooperative Program (TTCP). Information from the program will be used in the planning of similar collaborative tests, to be carried out at NASA Ames, on a full-scale aircraft. The program covered the measurement of unsteady pressures and fin vibration for cases with and without the wing LEX fences fitted. Fourier transform methods were used to analyze the unsteady data, and information on the spatial and temporal content of the vortex burst pressure field was obtained. Flow visualization of the vortex behavior was carried out using smoke and a laser light sheet technique.

Rotor blade system with reduced blade-vortex interaction noise

NASA Technical Reports Server (NTRS)

Leishman, John G. (Inventor); Han, Yong Oun (Inventor)

2005-01-01

A rotor blade system with reduced blade-vortex interaction noise includes a plurality of tube members embedded in proximity to a tip of each rotor blade. The inlets of the tube members are arrayed at the leading edge of the blade slightly above the chord plane, while the outlets are arrayed at the blade tip face. Such a design rapidly diffuses the vorticity contained within the concentrated tip vortex because of enhanced flow mixing in the inner core, which prevents the development of a laminar core region.

VORSTAB: A computer program for calculating lateral-directional stability derivatives with vortex flow effect

NASA Technical Reports Server (NTRS)

Lan, C. Edward

1985-01-01

A computer program based on the Quasi-Vortex-Lattice Method of Lan is presented for calculating longitudinal and lateral-directional aerodynamic characteristics of nonplanar wing-body combination. The method is based on the assumption of inviscid subsonic flow. Both attached and vortex-separated flows are treated. For the vortex-separated flow, the calculation is based on the method of suction analogy. The effect of vortex breakdown is accounted for by an empirical method. A summary of the theoretical method, program capabilities, input format, output variables and program job control set-up are described. Three test cases are presented as guides for potential users of the code.

Multiple leading edge vortices of unexpected strength in freely flying hawkmoth

PubMed Central

Johansson, L. Christoffer; Engel, Sophia; Kelber, Almut; Heerenbrink, Marco Klein; Hedenström, Anders

2013-01-01

The Leading Edge Vortex (LEV) is a universal mechanism enhancing lift in flying organisms. LEVs, generally illustrated as a single vortex attached to the wing throughout the downstroke, have not been studied quantitatively in freely flying insects. Previous findings are either qualitative or from flappers and tethered insects. We measure the flow above the wing of freely flying hawkmoths and find multiple simultaneous LEVs of varying strength and structure along the wingspan. At the inner wing there is a single, attached LEV, while at mid wing there are multiple LEVs, and towards the wingtip flow separates. At mid wing the LEV circulation is ~40% higher than in the wake, implying that the circulation unrelated to the LEV may reduce lift. The strong and complex LEV suggests relatively high flight power in hawmoths. The variable LEV structure may result in variable force production, influencing flight control in the animals. PMID:24253180

Flow-around modes for a rhomboid wing with a stall vortex in the shock layer

NASA Astrophysics Data System (ADS)

Zubin, M. A.; Maximov, F. A.; Ostapenko, N. A.

2017-12-01

The results of theoretical and experimental investigation of an asymmetrical hypersonic flow around a V-shaped wing with the opening angle larger than π on the modes with attached shockwaves on forward edges, when the stall flow is implemented on the leeward wing cantilever behind the kink point of the cross contour. In this case, a vortex of nonviscous nature is formed in which the velocities on the sphere exceeding the speed of sound and resulting in the occurrence of pressure shocks with an intensity sufficient for the separation of the turbulent boundary layer take place in the reverse flow according to the calculations within the framework of the ideal gas. It is experimentally established that a separation boundary layer can exist in the reverse flow, and its structure is subject to the laws inherent to the reverse flow in the separation region of the turbulent boundary layer arising in the supersonic conic flow under the action of a shockwave incident to the boundary layer.

Investigation of Vortex Flaps and Other Flow Control Devices on Generic High-Speed Civil Transport Planforms

NASA Technical Reports Server (NTRS)

Kjerstad, Kevin J.; Campbell, Bryan A.; Gile, Brenda E.; Kemmerly, Guy T.

1999-01-01

A parametric cranked delta planform study has been conducted in the Langley 14- by 22-Foot Subsonic Tunnel with the following objectives: (1) to evaluate the vortex flap design methodology for cranked delta wings, (2) to determine the influence of leading-edge sweep and the outboard wing on vortex flap effectiveness, (3) to evaluate novel flow control concepts, and (4) to validate unstructured grid Euler computer code predictions with modeled vortex and trailing-edge flaps. Two families of cranked delta planforms were investigated. One family had constant aspect ratio, while the other had a constant nondimensional semispan location of the leading-edge break. The inboard leading-edge sweep of the planforms was varied between 68 deg., 71 deg., and 74 deg., while outboard leading-edge sweep was varied between 48 deg. and 61 deg. Vortex flaps for the different planforms were designed by an analytical vortex flap design method. The results indicate that the effectiveness of the vortex flaps was only slightly influenced by the variations in the parametric planforms. The unstructured grid Euler computer code was successfully used to model the configurations with vortex flaps. The vortex trap concept was successfully demonstrated.

Investigation of tip clearance flow physics in axial flow turbine rotors

NASA Astrophysics Data System (ADS)

Xiao, Xinwen

In axial turbines, the tip clearance between casing wall and rotating blades results in a tip leakage flow, which significantly affects loss production, heat protection, vibration and noise. It is important to minimize these effects for a better turbine engine performance and higher reliability. Most of previous efforts were concentrated on turbine cascades that however may not completely and correctly simulate the flow physics in practical turbine rotors. An investigation has to be performed in turbine rotors to reveal the real tip leakage flow physics in order to provide a scientific basis for minimizing its effects. This is the objective of this thesis research. The three dimensional flow field near the end wall/tip clearance region in a turbine rotor has been investigated experimentally, complemented by a numerical simulation to study the influences of inlet turbulence intensities on the development of the tip leakage flow. The experimental investigation is carried out in a modern unshrouded high pressure turbine stage. The survey region covers 20% span near the end wall, and extends axially from 10% chord upstream of the leading edge, through the rotor passage, and to 20% chord downstream of the trailing edge. It has been found that the tip leakage effects extend only to the surveyed region. The three dimensional LDV technique is used to measure the velocity and turbulence field upstream of the rotor, inside the rotor passage, and near the trailing edge. The static pressure on blade surfaces is surveyed from the rotating frame. The transient pressure on the casing wall is measured using a dynamic pressure sensor with a shaft encoder. A rotating Five Hole Probe is employed to measure the losses as well as the pressure and the three dimensional velocity field at 20% chord downstream of the rotor. The unsteady flow field is also investigated at this location by using a slanted single-element Hot Wire technique. The physics of the tip leakage flow and vortex in turbine rotors, including its inception location, development, interaction with the main stream and the passage vortex, and decay, are revealed. The rotation effects on the boundary layer flow and the turbulence structure are discussed. The effects of the relative motion between the blade and the casing wall on the flow field near the tip clearance region are also investigated. The structure of the rotor wake, the nozzle wake, and their interaction are interpreted based on the instantaneous Hot Wire data. The numerical simulation on the influence of the inlet turbulence intensity on the development of the tip leakage flow is based on previous efforts. The results indicate that the tip leakage vortex diffuses very quickly under a high inlet turbulence intensity, resulting in a very weak tip leakage vortex and less losses.

An improved panel method for the solution of three-dimensional leading-edge vortex flows. Volume 1: Theory document

NASA Technical Reports Server (NTRS)

Johnson, F. T.; Lu, P.; Tinoco, E. N.

1980-01-01

An improved panel method for the solution of three dimensional flow and wing and wing-body combinations with leading edge vortex separation is presented. The method employs a three dimensional inviscid flow model in which the configuration, the rolled-up vortex sheets, and the wake are represented by quadratic doublet distributions. The strength of the singularity distribution as well as shape and position of the vortex spirals are computed in an iterative fashion starting with an assumed initial sheet geometry. The method calculates forces and moments as well as detail surface pressure distributions. Improvements include the implementation of improved panel numerics for the purpose of elimination the highly nonlinear effects of ring vortices around double panel edges, and the development of a least squares procedure for damping vortex sheet geometry update instabilities. A complete description of the method is included. A variety of cases generated by the computer program implementing the method are presented which verify the mathematical assumptions of the method and which compare computed results with experimental data to verify the underlying physical assumptions made by the method.

Experimental study of effects of forebody geometry on high angle of attack static and dynamic stability and control

NASA Technical Reports Server (NTRS)

Brandon, J. M.; Murri, D. G.; Nguyen, L. T.

1986-01-01

A series of low-speed wind tunnel tests on a generic airplane model with a cylindrical fuselage were made to investigate the effects of forebody shape and fitness ratio, and fuselage/wing proximity on static and dynamic lateral/directional stability. In addition, some preliminary testing to determine the effectiveness of deflectable forebody strakes for high angle of attack yaw control was conducted. During the stability investigation, 11 forebodies were tested including three different cross-sectional shapes with fineness ratios of 2, 3, and 4. In addition, the wing was tested at two longitudinal positions to provide a substantial variation in forebody/wing proximity. Conventional force tests were conducted to determine static stability characteristics, and single-degree-of-freedom free-to-roll tests were conducted to study the wing rock characteristics of the model with the various forebodies. Flow visualization data were obtained to aid in the analysis of the complex flow phenomena involved. The results show that the forebody cross-sectional shape and fineness ratio and forebody/wing proximity can strongly affect both static and dynamic (roll) stability at high angles of attack. These characteristics result from the impact of these factors on forebody vortex development, the behavior of the vortices in sideslip, and their interaction with the wing flow field. Preliminary results from the deflectable strake investigation indicated that forebody flow control using this concept can provide very large yaw control moments at stall and post-stall angles of attack.

Navier-Stokes Computations of a Wing-Flap Model With Blowing Normal to the Flap Surface

NASA Technical Reports Server (NTRS)

Boyd, D. Douglas, Jr.

2005-01-01

A computational study of a generic wing with a half span flap shows the mean flow effects of several blown flap configurations. The effort compares and contrasts the thin-layer, Reynolds averaged, Navier-Stokes solutions of a baseline wing-flap configuration with configurations that have blowing normal to the flap surface through small slits near the flap side edge. Vorticity contours reveal a dual vortex structure at the flap side edge for all cases. The dual vortex merges into a single vortex at approximately the mid-flap chord location. Upper surface blowing reduces the strength of the merged vortex and moves the vortex away from the upper edge. Lower surface blowing thickens the lower shear layer and weakens the merged vortex, but not as much as upper surface blowing. Side surface blowing forces the lower surface vortex farther outboard of the flap edge by effectively increasing the aerodynamic span of the flap. It is seen that there is no global aerodynamic penalty or benefit from the particular blowing configurations examined.

Effects of leading-edge devices on the low-speed aerodynamic characteristics of a highly-swept arrow-wing

NASA Technical Reports Server (NTRS)

Scott, S. J.; Nicks, O. W.; Imbrie, P. K.

1985-01-01

An investigation was conducted in the Texas A&M University 7 by 10 foot Low Speed Wind Tunnel to provide a direct comparison of the effect of several leading edge devices on the aerodynamic performance of a highly swept wing configuration. Analysis of the data indicates that for the configuration with undeflected leading edges, vortex separation first occurs on the outboard wing panel for angles of attack of approximately 2, and wing apex vorticies become apparent for alpha or = 4 deg. However, the occurrence of the leading edge vortex flow may be postponed with leading edge devices. Of the devices considered, the most promising were a simple leading edge deflection of 30 deg and a leading edge slat system. The trailing edge flap effectiveness was found to be essentially the same for the configuration employing either of these more promising leading edge devices. Analysis of the lateral directional data showed that for all of the concepts considered, deflecting leading edge downward in an attempt to postpone leading edge vortex flows, has the favorable effect of reducing the effective dihedral.

Infrared Imaging Of Flows Seeded With SF6

NASA Technical Reports Server (NTRS)

Manuel, Gregory S.; Daryabeigi, Kamran; Alderfer, David W.; Obara, Clifford J.

1993-01-01

Novel technique enables repeated measurements of flow patterns during flight. Wing-tip vorticity studied in flight by observing infrared emissions from SF6 gas entrained in wing-tip flow. System makes vortical flows visible throughout all altitude and speed ranges of all subsonic aircraft. Also useful for transonic and supersonic speeds. Primary application is testing of aircraft in flight, also proves useful in testing fast land vehicles and structures or devices subject to strong winds.

Flow-visualization study of the X-29A aircraft at high angles of attack using a 1/48-scale model

NASA Technical Reports Server (NTRS)

Cotton, Stacey J.; Bjarke, Lisa J.

1994-01-01

A water-tunnel study on a 1/48-scale model of the X-29A aircraft was performed at the NASA Dryden Flow Visualization Facility. The water-tunnel test enhanced the results of the X-29A flight tests by providing flow-visualization data for comparison and insights into the aerodynamic characteristics of the aircraft. The model was placed in the water tunnel at angles of attack of 20 to 55 deg. and with angles of sideslip from 0 to 5 deg. In general, flow-visualization techniques provided useful information on vortex formation, separation, and breakdown and their role in yaw asymmetries and tail buffeting. Asymmetric forebody vortices were observed at angles of attack greater than 30 deg. with 0 deg. sideslip and greater than 20 deg. with 5 deg. sideslip. While the asymmetric flows observed in the water tunnel did not agree fully with the flight data, they did show some of the same trends. In addition, the flow visualization indicated that the interaction of forebody vortices and the wing wake at angles of attack between 20 and 35 deg. may cause vertical-tail buffeting observed in flight.

A theoretical investigation of the aerodynamics of low-aspect-ratio wings with partial leading-edge separation

NASA Technical Reports Server (NTRS)

Mehrotra, S. C.; Lan, C. E.

1978-01-01

A numerical method is developed to predict distributed and total aerodynamic characteristics for low aspect-ratio wings with partial leading-edge separation. The flow is assumed to be steady and inviscid. The wing boundary condition is formulated by the quasi-vortex-lattice method. The leading-edge separated vortices are represented by discrete free vortex elements which are aligned with the local velocity vector at mid-points to satisfy the force free condition. The wake behind the trailing-edge is also force free. The flow tangency boundary condition is satisfied on the wing, including the leading- and trailing-edges. Comparison of the predicted results with complete leading-edge separation has shown reasonably good agreement. For cases with partial leading-edge separation, the lift is found to be highly nonlinear with angle of attack.

Numerical investigation of the unsteady tip leakage flow and rotating stall inception in a transonic compressor

NASA Astrophysics Data System (ADS)

Zhang, Yanfeng; Lu, Xingen; Chu, Wuli; Zhu, Junqiang

2010-08-01

It is well known that tip leakage flow has a strong effect on the compressor performance and stability. This paper reports on a numerical investigation of detailed flow structures in an isolated transonic compressor rotor-NASA Rotor 37 at near stall and stalled conditions aimed at improving understanding of changes in 3D tip leakage flow structures with rotating stall inception. Steady and unsteady 3D Navier-Stokes analyses were conducted to investigate flow structures in the same rotor. For steady analysis, the predicted results agree well with the experimental data for the estimation of compressor rotor global performance. For unsteady flow analysis, the unsteady flow nature caused by the breakdown of the tip leakage vortex in blade tip region in the transonic compressor rotor at near stall condition has been captured with a single blade passage. On the other hand, the time-accurate unsteady computations of multi-blade passage at near stall condition indicate that the unsteady breakdown of the tip leakage vortex triggered the short length-scale — spike type rotating stall inception at blade tip region. It was the forward spillage of the tip leakage flow at blade leading edge resulting in the spike stall inception. As the mass flow ratio is decreased, the rotating stall cell was further developed in the blade passage.

Tactical missile aerodynamics

NASA Technical Reports Server (NTRS)

Hemsch, Michael J. (Editor); Nielsen, Jack N. (Editor)

1986-01-01

The present conference on tactical missile aerodynamics discusses autopilot-related aerodynamic design considerations, flow visualization methods' role in the study of high angle-of-attack aerodynamics, low aspect ratio wing behavior at high angle-of-attack, supersonic airbreathing propulsion system inlet design, missile bodies with noncircular cross section and bank-to-turn maneuvering capabilities, 'waverider' supersonic cruise missile concepts and design methods, asymmetric vortex sheding phenomena from bodies-of-revolution, and swept shock wave/boundary layer interaction phenomena. Also discussed are the assessment of aerodynamic drag in tactical missiles, the analysis of supersonic missile aerodynamic heating, the 'equivalent angle-of-attack' concept for engineering analysis, the vortex cloud model for body vortex shedding and tracking, paneling methods with vorticity effects and corrections for nonlinear compressibility, the application of supersonic full potential method to missile bodies, Euler space marching methods for missiles, three-dimensional missile boundary layers, and an analysis of exhaust plumes and their interaction with missile airframes.

Separated Flow Control with Actuated Membrane Wings

NASA Astrophysics Data System (ADS)

Bohnker, Jillian; Breuer, Kenneth

2017-11-01

By perturbing shear layer instabilities, some level of control over highly separated flows can be established, as has been demonstrated on rigid wings using synthetic jet actuators or acoustic excitation. Here, we demonstrate similar phenomena using sinusoidal actuation of a dielectric membrane wing. The effect of actuation on lift is examined as a function of freestream velocity (5-25 m/s), angle of attack (10°-40°), and actuation frequency (0.1

Influence of optimized leading-edge deflection and geometric anhedral on the low-speed aerodynamic characteristics of a low-aspect-ratio highly swept arrow-wing configuration. [langley 7 by 10 foot tunnel

NASA Technical Reports Server (NTRS)

Coe, P. L., Jr.; Huffman, J. K.

1979-01-01

An investigation conducted in the Langley 7 by 10 foot tunnel to determine the influence of an optimized leading-edge deflection on the low speed aerodynamic performance of a configuration with a low aspect ratio, highly swept wing. The sensitivity of the lateral stability derivative to geometric anhedral was also studied. The optimized leading edge deflection was developed by aligning the leading edge with the incoming flow along the entire span. Owing to spanwise variation of unwash, the resulting optimized leading edge was a smooth, continuously warped surface for which the deflection varied from 16 deg at the side of body to 50 deg at the wing tip. For the particular configuration studied, levels of leading-edge suction on the order of 90 percent were achieved. The results of tests conducted to determine the sensitivity of the lateral stability derivative to geometric anhedral indicate values which are in reasonable agreement with estimates provided by simple vortex-lattice theories.

Video photographic considerations for measuring the proximity of a probe aircraft with a smoke seeded trailing vortex

NASA Technical Reports Server (NTRS)

Childers, Brooks A.; Snow, Walter L.

1990-01-01

Considerations for acquiring and analyzing 30 Hz video frames from charge coupled device (CCD) cameras mounted in the wing tips of a Beech T-34 aircraft are described. Particular attention is given to the characterization and correction of optical distortions inherent in the data.

An Investigation into the Potential Benefits of Distributed Electric Propulsion on Small UAVs at Low Reynolds Numbers

NASA Astrophysics Data System (ADS)

Baris, Engin

Distributed electric propulsion systems benefit from the inherent scale independence of electric propulsion. This property allows the designer to place multiple small electric motors along the wing of an aircraft instead of using a single or several internal combustion motors with gear boxes or other power train components. Aircraft operating at low Reynolds numbers are ideal candidates for benefiting from increased local flow velocities as provided by distributed propulsion systems. In this study, a distributed electric propulsion system made up of eight motor/propellers was integrated into the leading edge of a small fixed wing-body model to investigate the expected improvements on the aerodynamics available to small UAVs operating at low Reynolds numbers. Wind tunnel tests featuring a Design of Experiments (DOE) methodology were used for aerodynamic characterization. Experiments were performed in four modes: all-propellers-on, wing-tip-propellers-alone-on, wing-alone mode, and two-inboard-propellers-on-alone mode. In addition, the all-propeller-on, wing-alone, and a single-tractor configuration were analyzed using VSPAERO, a vortex lattice code, to make comparisons between these different configurations. Results show that the distributed propulsion system has higher normal force, endurance, and range features, despite a potential weight penalty.

Flow Field Characteristics and Lift Changing Mechanism for Half-Rotating Wing in Hovering Flight

NASA Astrophysics Data System (ADS)

Li, Q.; Wang, X. Y.; Qiu, H.; Li, C. M.; Qiu, Z. Z.

2017-12-01

Half-rotating wing (HRW) is a new similar-flapping wing system based on half-rotating mechanism which could perform rotating-type flapping instead of oscillating-type flapping. The characteristics of flow field and lift changing mechanism for HRW in hovering flight are important theoretical basis to improve the flight capability of HRW aircraft. The driving mechanism and work process of HRW were firstly introduced in this paper. Aerodynamic simulation model of HRW in hovering flight was established and solved using XFlow software, by which lift changing rule of HRW was drawn from the simulation solution. On the other hand, the development and shedding of the distal vortex throughout one stroke would lead to the changes of the lift force. Based on analyzing distribution characteristics of vorticity, velocity and pressure around wing blade, the main features of the flow field for HRW were further given. The distal attached vortex led to the increase of the lift force, which would gradually shed into the wake with a decline of lift in the later downstroke. The wake ring directed by the distal end of the blade would generate the downward accelerating airflow which produced the upward anti-impulse to HRW. The research results mentioned above illustrated that the behavior characteristics of vortex formed in flow field were main cause of lift changing for HRW.

Vortex Flap Technology: a Stability and Control Assessment

NASA Technical Reports Server (NTRS)

Carey, K. M.; Erickson, G. E.

1984-01-01

A comprehensive low-speed wind tunnel investigation was performed of leading edge vortex flaps applied to representative aircraft configurations. A determination was made of the effects of analytically- and empirically-designed vortex flaps on the static longitudinal and lateral-directional aerodynamics, stability, and control characteristics of fighter wings having leading-edge sweep angles of 45 to 76.5 degrees. The sensitivity to several configuration modifications was assessed, which included the effects of flap planform, leading- and trailing-edge flap deflection angles, wing location on the fuselage, forebody strakes, canards, and centerline and outboard vertical tails. Six-component forces and moments, wing surface static pressure distributions, and surface flow patterns were obtained using the Northrop 21- by 30-inch low-speed wind tunnel.

Unsteady aerodynamics of membrane wings with adaptive compliance

NASA Astrophysics Data System (ADS)

Kiser, Jillian; Breuer, Kenneth

2016-11-01

Membrane wings are known to provide superior aerodynamic performance at low Reynolds numbers (Re =104 -105), primarily due to passive shape adaptation to flow conditions. In addition to this passive deformation, active control of the fluid-structure interaction and resultant aerodynamic properties can be achieved through the use of dielectric elastomer actuators as the wing membrane material. When actuated, membrane pretension is decreased and wing camber increases. Additionally, actuation at resonance frequencies allows additional control over wing camber. We present results using synchronized (i) time-resolved particle image velocimetry (PIV) to resolve the flow field, (ii) 3D direct linear transformation (DLT) to recover membrane shape, (iii) lift/drag/torque measurements and (iv) near-wake hot wire anemometry measurements to characterize the fluid-structure interactions. Particular attention is paid to cases in which the vortex shedding frequency, the membrane resonance, and the actuation frequency coincide. In quantitatively examining both flow field and membrane shape at a range of actuation frequencies and vortex shedding frequencies, this work seeks to find actuation parameters that allow for active control of boundary layer separation over a range of flow conditions. Also at Naval Undersea Warfare Center, Division Newport.

The application of experimental data to blade wake interaction noise prediction

NASA Technical Reports Server (NTRS)

Glegg, Stewart A. L.; Devenport, William J.

1991-01-01

Blade wake interaction noise (BWI) has been defined as the broadband noise generated by the ingestion of turbulent trailing tip vortices by helicopter rotors. This has been shown to be the dominant contributor to the subjectively important part of the acoustic spectrum for the approach stage of a helicopter flyover. A prediction method for BWI noise based on the calculated trailing vortex trajectories has been developed and estimates of the vortex turbulence have been made. These measurements were made on a trailing vortex from a split wing arrangement and did not give the spectrum of the velocity fluctuations. A recent experiment carried out to measure the turbulence associated with a trailing vortex and the application of the results to BWI noise prediction is described.

Numerical analyses of evolution of unsteady flow structures in the wake of flapping starling wing model

NASA Astrophysics Data System (ADS)

Krishnan, Krishnamoorthy; Naqavi, Iftekhar Z.; Gurka, Roi

2017-11-01

Understanding the physics of flapping wings at moderate Reynolds number flows takes on greater importance in the context of avian aerodynamics as well as in the design of miniature-aerial-vehicles. Analyzing the characteristics of wake vortices generated downstream of flapping wings can help to explain the unsteady contribution to the aerodynamics loads. In this study, numerical simulations of flow over a bio-inspired pseudo-2D flapping wing model was conducted to characterize the evolution of unsteady flow structures in the downstream wake of flapping wing. The wing model was based on a European starling's wing and wingbeat kinematics were incorporated to simulate a free-forward flight. The starling's wingbeat kinematics were extracted from experiments conducted in a wind tunnel where freely flying starling was measured using high-speed PIV as well as high-speed imaging yielding a series of kinematic images sampled at 500 Hz. The average chord of the wing section was 6 cm and simulations were carried out at a Reynolds number of 54,000, reduced frequency of 0.17, and Strouhal number of 0.16. Large eddy simulation was performed using a second order, finite difference code ParLES. Characteristics of wake vortex structures during the different phases of the wing strokes were examined. The role of wingbeat kinematics in the configuration of downstream vortex patterns is discussed. Evaluated wake topology and lift-drag characteristics are compared with the starling's wind tunnel results.

Experimental investigation into wing span and angle-of-attack effects on sub-scale race car wing/wheel interaction aerodynamics

NASA Astrophysics Data System (ADS)

Diasinos, S.; Gatto, A.

2008-09-01

This paper details a quantitative 3D investigation using LDA into the interaction aerodynamics on a sub-scale open wheel race car inverted front wing and wheel. Of primary importance to this study was the influence of changing wing angle of attack and span on the resulting near-field and far-field flow characteristics. Results obtained showed that both variables do have a significant influence on the resultant flow-field, particularly on wing vortex and wheel wake development and propagation.

Comparison of Experimental Surface and Flow Field Measurements to Computational Results of the Juncture Flow Model

NASA Technical Reports Server (NTRS)

Roozeboom, Nettie H.; Lee, Henry C.; Simurda, Laura J.; Zilliac, Gregory G.; Pulliam, Thomas H.

2016-01-01

Wing-body juncture flow fields on commercial aircraft configurations are challenging to compute accurately. The NASA Advanced Air Vehicle Program's juncture flow committee is designing an experiment to provide data to improve Computational Fluid Dynamics (CFD) modeling in the juncture flow region. Preliminary design of the model was done using CFD, yet CFD tends to over-predict the separation in the juncture flow region. Risk reduction wind tunnel tests were requisitioned by the committee to obtain a better understanding of the flow characteristics of the designed models. NASA Ames Research Center's Fluid Mechanics Lab performed one of the risk reduction tests. The results of one case, accompanied by CFD simulations, are presented in this paper. Experimental results suggest the wall mounted wind tunnel model produces a thicker boundary layer on the fuselage than the CFD predictions, resulting in a larger wing horseshoe vortex suppressing the side of body separation in the juncture flow region. Compared to experimental results, CFD predicts a thinner boundary layer on the fuselage generates a weaker wing horseshoe vortex resulting in a larger side of body separation.

3-D flow and scour near a submerged wing dike: ADCP measurements on the Missouri River

USGS Publications Warehouse

Jamieson, E.C.; Rennie, C.D.; Jacobson, R.B.; Townsend, R.D.

2011-01-01

Detailed mapping of bathymetry and three-dimensional water velocities using a boat-mounted single-beam sonar and acoustic Doppler current profiler (ADCP) was carried out in the vicinity of two submerged wing dikes located in the Lower Missouri River near Columbia, Missouri. During high spring flows the wing dikes become submerged, creating a unique combination of vertical flow separation and overtopping (plunging) flow conditions, causing large-scale three-dimensional turbulent flow structures to form. On three different days and for a range of discharges, sampling transects at 5 and 20 m spacing were completed, covering the area adjacent to and upstream and downstream from two different wing dikes. The objectives of this research are to evaluate whether an ADCP can identify and measure large-scale flow features such as recirculating flow and vortex shedding that develop in the vicinity of a submerged wing dike; and whether or not moving-boat (single-transect) data are sufficient for resolving complex three-dimensional flow fields. Results indicate that spatial averaging from multiple nearby single transects may be more representative of an inherently complex (temporally and spatially variable) three-dimensional flow field than repeated single transects. Results also indicate a correspondence between the location of calculated vortex cores (resolved from the interpolated three-dimensional flow field) and the nearby scour holes, providing new insight into the connections between vertically oriented coherent structures and local scour, with the unique perspective of flow and morphology in a large river.

Theoretical characteristics in supersonic flow of two types of control surfaces on triangular wings

NASA Technical Reports Server (NTRS)

Tucker, Warren A; Nelson, Robert L

1949-01-01

Methods based on the linearized theory for supersonic flow were used to find the characteristics of two types of control surfaces on thin triangular wings. The first type, the constant-chord partial-span flap, was considered to extend either outboard from the center of the wing or inboard from the wing tip. The second type, the full-triangular-tip flap, was treated only for the case in which the Mach number component normal to the leading edge is supersonic. For each type, expressions were found for the lift, rolling-moment, pitching-moment, and hinge-moment characteristics.

Turbulent flow around a wing-fuselage type juncture

NASA Technical Reports Server (NTRS)

Kubendran, L. R.; Mcmahon, H. M.; Hubbartt, J. E.

1985-01-01

The flow over a 58-mm-thick uniform-thickness winglike body having a 1.5:1 elliptical leading edge and joined to a large flat plate (representing an aircraft fuselage) is characterized experimentally at freestream velocity 15 m/s, corresponding to Reynolds number 940,000/m, using hot-wire anemometry. The results are presented graphically, and it is found that the horseshoe vortex formed by the separation of the fuselage boundary layer ahead of the wing leading edge is effective in transporting turbulence and modifying the mean-flow characteristics and the turbulent-stress distribution. It is suggested that the slenderness ratio of the leading edge is the dominant factor affecting the strength and location of the vortex.

The leading-edge vortex of swift-wing shaped delta wings

NASA Astrophysics Data System (ADS)

Muir, Rowan; Arredondo-Galeana, Abel; Viola, Ignazio Maria

2017-11-01

Recent investigations on the aerodynamics of natural fliers have illuminated the significance of the Leading-Edge Vortex (LEV) for lift generation in a variety of flight conditions. In this investigation, a model non-slender delta shaped wing with a sharp leading-edge is tested at low Reynolds Number, along with a delta wing of the same design, but with a modified trailing edge inspired by the wing of a common swift Apus apus. The effect of the tapering swift wing on LEV development and stability is compared with the flow structure over the un-modified delta wing model through particle image velocimetry. For the first time, a leading-edge vortex system consisting of a dual or triple LEV is recorded on a swift-wing shaped delta wing, where such a system is found across all tested conditions. It is shown that the spanwise location of LEV breakdown is governed by the local chord rather than Reynolds Number or angle of attack. These findings suggest that the trailing-edge geometry of the swift wing alone does not prevent the common swift from generating an LEV system comparable with that of a delta shaped wing. This work received funding from the Engineering and Physical Sciences Research Council [EP/M506515/1] and the Consejo Nacional de Ciencia y Tecnología (CONACYT).

Sailplane Glide Performance and Control Using Fixed and Articulating Winglets. M.S. Thesis

NASA Technical Reports Server (NTRS)

Colling, James David

1995-01-01

An experimental study was conducted to investigate the effects of controllable articulating winglets on glide performance and yawing moments of high performance sailplanes. Testing was conducted in the Texas A&M University 7 x 10 foot Low Speed Wind Tunnel using a full-scale model of the outboard 5.6 feet of a 15 meter class high performance sailplane wing. Different wing tip configurations could be easily mounted to the wing model. A winglet was designed in which the cant and toe angles as well as a rudder on the winglet could be adjusted to a range of positions. Cant angles used in the investigation consisted of 5, 25, and 40 degrees measured from the vertical axis. Toe-out angles ranged from 0 to 22.5 degrees. A rudder on the winglet was used to study the effects of changing the camber of the winglet airfoil on wing performance and wing yawing moments. Rudder deflections consisted of-10, 0, and 10 degrees. Test results for a fixed geometry winglet and a standard wing tip are presented to show the general behavior of winglets on sailplane wings, and the effects of boundary-layer turbulators on the winglets are also presented. By tripping the laminar boundary-layer to turbulent before laminar separation occurs, the wing performance was increased at low Reynolds numbers. The effects on the lift and drag, yawing moment, pitching moment, and wing root bending moment of the model are presented. Oil flows were used on the wing model with the fixed geometry winglet and the standard wing tip to visualize flow directions and areas of boundary layer transition. A cant angle of 25 degrees and a toe-out angle of 2.5 degrees provided an optimal increase in wing performance for the cant and toe angles tested. Maximum performance was obtained when the winglet rudder remained in the neutral position of zero degrees. By varying the cant, toe, and rudder angles from their optimized positions, wing performance decreases. Although the winglet rudder proved to be more effective in increasing the yawing moment compared to varying the cant and toe angles, the amount of increased yawing moment was insignificant when compared to that produced by the vertical tail. A rudder on the winglet was determined to be ineffective for providing additional yaw control.

Propeller noise caused by blade tip radial forces

NASA Technical Reports Server (NTRS)

Hanson, D. B.

1986-01-01

New experimental evidence which indicates the presence of leading edge and tip edge vortex flow on Prop-Fans is examined, and performance and noise consequences are addressed. It was shown that the tip edge vortex is a significant noise source, particularly for unswept Prop-Fan blades. Preliminary calculations revealed that the addition of the tip side edge source to single rotation Prop-Fans during take off conditions improved the agreement between experiment and theory at blade passing frequency. At high-speed conditions such as the Prop-Fan cruise point, the tip loading effect tends to cancel thickness noise.

Theoretical prediction of thick wing and pylon-fuselage-fanpod-nacelle aerodynamic characteristics at subcritical speeds. Part 1: Theory and results

NASA Technical Reports Server (NTRS)

Tulinius, J. R.

1974-01-01

The theoretical development and the comparison of results with data of a thick wing and pylon-fuselage-fanpod-nacelle analysis are presented. The analysis utilizes potential flow theory to compute the surface velocities and pressures, section lift and center of pressure, and the total configuration lift, moment, and vortex drag. The skin friction drag is also estimated in the analysis. The perturbation velocities induced by the wing and pylon, fuselage and fanpod, and nacelle are represented by source and vortex lattices, quadrilateral vortices, and source frustums, respectively. The strengths of these singularities are solved for simultaneously including all interference effects. The wing and pylon planforms, twists, cambers, and thickness distributions, and the fuselage and fanpod geometries can be arbitrary in shape, provided the surface gradients are smooth. The flow through nacelle is assumed to be axisymmetric. An axisymmetric center engine hub can also be included. The pylon and nacelle can be attached to the wing, fuselage, or fanpod.

Large-Amplitude, High-Rate Roll Oscillations of a 65 deg Delta Wing at High Incidence

NASA Technical Reports Server (NTRS)

Chaderjian, Neal M.; Schiff, Lewis B.

2000-01-01

The IAR/WL 65 deg delta wing experimental results provide both detail pressure measurements and a wide range of flow conditions covering from simple attached flow, through fully developed vortex and vortex burst flow, up to fully-stalled flow at very high incidence. Thus, the Computational Unsteady Aerodynamics researchers can use it at different level of validating the corresponding code. In this section a range of CFD results are provided for the 65 deg delta wing at selected flow conditions. The time-dependent, three-dimensional, Reynolds-averaged, Navier-Stokes (RANS) equations are used to numerically simulate the unsteady vertical flow. Two sting angles and two large- amplitude, high-rate, forced-roll motions and a damped free-to-roll motion are presented. The free-to-roll motion is computed by coupling the time-dependent RANS equations to the flight dynamic equation of motion. The computed results are compared with experimental pressures, forces, moments and roll angle time history. In addition, surface and off-surface flow particle streaks are also presented.

Unsteady aerodynamic flow field analysis of the space shuttle configuration. Part 1: Orbiter aerodynamics

NASA Technical Reports Server (NTRS)

Ericsson, L. E.; Reding, J. P.

1976-01-01

An analysis of the steady and unsteady aerodynamics of the space shuttle orbiter has been performed. It is shown that slender wing theory can be modified to account for the effect of Mach number and leading edge roundness on both attached and separated flow loads. The orbiter unsteady aerodynamics can be computed by defining two equivalent slender wings, one for attached flow loads and another for the vortex-induced loads. It is found that the orbiter is in the transonic speed region subject to vortex-shock-boundary layer interactions that cause highly nonlinear or discontinuous load changes which can endanger the structural integrity of the orbiter wing and possibly cause snap roll problems. It is presently impossible to simulate these interactions in a wind tunnel test even in the static case. Thus, a well planned combined analytic and experimental approach is needed to solve the problem.

NASA aircraft trailing vortex research

NASA Technical Reports Server (NTRS)

Mcgowan, W. A.

1971-01-01

A brief description is given of NASA's comprehensive program to study the aircraft trailing vortex problem. Wind tunnel experiments are used to develop the detailed processes of wing tip vortex formation and explore different means to either prevent trailing vortices from forming or induce early break-up. Flight tests provide information on trailing vortex system behavior behind large transport aircraft, both near the ground, as in the vicinity of the airport, and at cruise/holding pattern altitudes. Results from some flight tests are used to show how pilots might avoid the dangerous areas when flying in the vicinity of large transport aircraft. Other flight tests will be made to verify and evaluate trailing vortex elimination schemes developed in the model tests. Laser Doppler velocimeters being developed for use in the research program and to locate and measure vortex winds in the airport area are discussed. Field tests have shown that the laser Doppler velocimeter measurements compare well with those from cup anemometers.

Aperiodicity Correction for Rotor Tip Vortex Measurements

NASA Technical Reports Server (NTRS)

Ramasamy, Manikandan; Paetzel, Ryan; Bhagwat, Mahendra J.

2011-01-01

The initial roll-up of a tip vortex trailing from a model-scale, hovering rotor was measured using particle image velocimetry. The unique feature of the measurements was that a microscope was attached to the camera to allow much higher spatial resolution than hitherto possible. This also posed some unique challenges. In particular, the existing methodologies to correct for aperiodicity in the tip vortex locations could not be easily extended to the present measurements. The difficulty stemmed from the inability to accurately determine the vortex center, which is a prerequisite for the correction procedure. A new method is proposed for determining the vortex center, as well as the vortex core properties, using a least-squares fit approach. This approach has the obvious advantage that the properties are derived from not just a few points near the vortex core, but from a much larger area of flow measurements. Results clearly demonstrate the advantage in the form of reduced variation in the estimated core properties, and also the self-consistent results obtained using three different aperiodicity correction methods.

Flow Field Characteristics of Finite-span Hydrofoils with Leading Edge Protuberances

NASA Astrophysics Data System (ADS)

Custodio, Derrick; Henoch, Charles; Johari, Hamid; Office of Naval Research Collaboration

2011-11-01

Past work has shown that humpback whale-like leading edge protuberances can significantly alter the load characteristics of both 2D and finite-span hydrofoils. To understand the mechanisms responsible for observed performance changes, the flow field characteristics of a baseline hydrofoil and models with leading edge protuberances were examined using the Stereo Particle Image Velocimetry (SPIV) technique. The near surface flow field on the hydrofoils was measured along with the tip vortex flow field on finite-span hydrofoils. Angles of attack ranging from 6 to 24 degrees were examined at freestream velocities of 1.8 m/s and 4.5 m/s, corresponding to Reynolds numbers of 180 and 450 thousand, respectively. While Reynolds number does not play a major role in establishing the flow field trends, both the protuberance geometry and spatial proximity to protuberances affect the velocity and vorticity characteristics near the foil surface, and in the wake and tip vortex. Near surface measurements reveal counter-rotating vortices on protuberance shoulders, while tip vortex measurements show that streamwise vorticity can be strongly affected by the presence of protuberances. The observed flow field characteristics will be presented. Sponsored by the ONR-ULI program.

On vortex-airfoil interaction noise including span-end effects, with application to open-rotor aeroacoustics

NASA Astrophysics Data System (ADS)

Roger, Michel; Schram, Christophe; Moreau, Stéphane

2014-01-01

A linear analytical model is developed for the chopping of a cylindrical vortex by a flat-plate airfoil, with or without a span-end effect. The major interest is the contribution of the tip-vortex produced by an upstream rotating blade in the rotor-rotor interaction noise mechanism of counter-rotating open rotors. Therefore the interaction is primarily addressed in an annular strip of limited spanwise extent bounding the impinged blade segment, and the unwrapped strip is described in Cartesian coordinates. The study also addresses the interaction of a propeller wake with a downstream wing or empennage. Cylindrical vortices are considered, for which the velocity field is expanded in two-dimensional gusts in the reference frame of the airfoil. For each gust the response of the airfoil is derived, first ignoring the effect of the span end, assimilating the airfoil to a rigid flat plate, with or without sweep. The corresponding unsteady lift acts as a distribution of acoustic dipoles, and the radiated sound is obtained from a radiation integral over the actual extent of the airfoil. In the case of tip-vortex interaction noise in CRORs the acoustic signature is determined for vortex trajectories passing beyond, exactly at and below the tip radius of the impinged blade segment, in a reference frame attached to the segment. In a second step the same problem is readdressed accounting for the effect of span end on the aerodynamic response of a blade tip. This is achieved through a composite two-directional Schwarzschild's technique. The modifications of the distributed unsteady lift and of the radiated sound are discussed. The chained source and radiation models provide physical insight into the mechanism of vortex chopping by a blade tip in free field. They allow assessing the acoustic benefit of clipping the rear rotor in a counter-rotating open-rotor architecture.

Flow interactions lead to orderly formations of flapping wings in forward flight

NASA Astrophysics Data System (ADS)

Ramananarivo, Sophie; Fang, Fang; Oza, Anand; Zhang, Jun; Ristroph, Leif

2016-11-01

Classic models of fish schools and flying formations of birds are built on the hypothesis that the preferred locations of an individual are determined by the flow left by its upstream neighbor. Lighthill posited that arrangements may in fact emerge passively from hydro- or aerodynamic interactions, drawing an analogy to the formation of crystals by intermolecular forces. Here, we carry out physical experiments aimed at testing the Lighthill conjecture and find that self-propelled flapping wings spontaneously assume one of multiple arrangements due to flow interactions. Wings in a tandem pair select the same forward speed, which tends to be faster than a single wing, while maintaining a separation distance that is an integer multiple of the wavelength traced out by each body. When perturbed, these locomotors robustly return to the same arrangement, and direct hydrodynamic force measurements reveal springlike restoring forces that maintain group cohesion. We also use these data to construct an interaction potential, showing how the observed positions of the follower correspond to stable wells in an energy landscape. Flow visualization and vortex-based theoretical models reveal coherent interactions in which the follower surfs on the periodic wake left by the leader. These results indicate that, for the high-Reynolds-number flows characteristic of schools and flocks, collective locomotion at enhanced speed and in orderly formations can emerge from flow interactions alone. If true for larger groups, then the view of collectives as ordered states of matter may prove to be a useful analogy.

Study of vortex generator influence on the flow in the wake of high-lift system wing

NASA Astrophysics Data System (ADS)

Bragin, N. N.; Ryabov, D. I.; Skomorokhov, S. I.; Slitinskaya, A. Yu.

2016-10-01

Passive vortex generators (VG) are known as one of the ways to improve the flow of the wings and other surfaces in the presence of flow separation. In particular, the VG are installed on the wings and nacelles of many foreign airplanes, including the most recent ones (for example, Boeing 787, Airbus A-350). The principle of the passive VG effects on flow is to transfer the kinetic energy of the external flow separation region by the vortices system arising from the flow VG themselves. For example, by increasing the angle of attack of the wing separation it is highly three-dimensional picture of the flow and sufficiently sensitive to external influences. Thus separated flow can be controlled when using the VG destroy large separation vortices. The VG effectiveness depends on many parameters. This is primarily the relative position of the second harmonic and the separation region on the wing and their size and position relative to each other, the orientation of the second harmonic relative to the local flow direction of the external flow, etc. Obviously, the VG effect will depend essentially on the intensity ratio of the second harmonic vortexes and nature of flow separation in the separation area. In the presence of intense flow separation the effect of conventional VG may be reduced or not occur at all. Until recently, investigations and selection of position of conventional VG were made only experimentally. Currently, the possibilities of calculation methods allow estimating the VG effect on the flow in the separation area. However, due to the phenomenon complexity the accuracy of these calculations is low. The experimental data are required to validate the computational methods, including information not only about the total impact, but also about the flow structure in the separation area. To obtain such information is the subject of this paper. In the test model of high-lift devices swept wing with modern supercritical profile the parametric studies were performed on the VG effects on the flow in the intensive separation zone on flaps. A number of VG types is considered that differ by configuration, size, location in relation to the area of flow separation on the flap, as well as the orientation relative to the incoming flow. The major part of standard of VG positions is investigated. The VG influence on head velocity loss and the characteristics of the amplitude-frequency spectra of pressure fluctuations in the wake of the wing are obtained, as well as the flow spectra are obtained by means of fluorescent mini-tufts.

The DELTA MONSTER: An RPV designed to investigate the aerodynamics of a delta wing platform

NASA Technical Reports Server (NTRS)

Connolly, Kristen; Flynn, Mike; Gallagher, Randy; Greek, Chris; Kozlowski, Marc; Mcdonald, Brian; Mckenna, Matt; Sellar, Rich; Shearon, Andy

1989-01-01

The mission requirements for the performance of aerodynamic tests on a delta wind planform posed some problems, these include aerodynamic interference; structural support; data acquisition and transmission instrumentation; aircraft stability and control; and propulsion implementation. To eliminate the problems of wall interference, free stream turbulence, and the difficulty of achieving dynamic similarity between the test and actual flight aircraft that are associated with aerodynamic testing in wind tunnels, the concept of the remotely piloted vehicle which can perform a basic aerodynamic study on a delta wing was the main objective for the Green Mission - the Delta Monster. The basic aerodynamic studies were performed on a delta wing with a sweep angle greater than 45 degrees. These tests were performed at various angles of attack and Reynolds numbers. The delta wing was instrumented to determine the primary leading edge vortex formation and location, using pressure measurements and/or flow visualization. A data acquisition system was provided to collect all necessary data.

A point vortex model for the formation of ocean eddies by flow separation

NASA Astrophysics Data System (ADS)

Southwick, O. R.; Johnson, E. R.; McDonald, N. R.

2015-01-01

A simple model for the formation of ocean eddies by flow separation from sharply curved horizontal boundary topography is developed. This is based on the Brown-Michael model for two-dimensional vortex shedding, which is adapted to more realistically model mesoscale oceanic flow by including a deforming free surface. With a free surface, the streamfunction for the flow is not harmonic so the conformal mapping methods used in the standard Brown-Michael approach cannot be used and the problem must be solved numerically. A numerical scheme is developed based on a Chebyshev spectral method for the streamfunction partial differential equation and a second order implicit timestepping scheme for the vortex position ordinary differntial equations. This method is used to compute shed vortex trajectories for three background flows: (A) a steady flow around a semi-infinite plate, (B) a free vortex moving around a semi-infinite plate, and (C) a free vortex moving around a right-angled wedge. In (A), the inclusion of surface deformation dramatically slows the vortex and changes its trajectory from a straight path to a curved one. In (B) and (C), without the inclusion of flow separation, free vortices traverse fully around the tip along symmetrical trajectories. With the effects of flow separation included, very different trajectories are found: for all values of the model parameter—the Rossby radius—the free and shed vortices pair up and move off to infinity without passing around the tip. Their final propagation angle depends strongly and monotonically on the Rossby radius.

Reynolds-averaged Navier-Stokes computation on tip clearance flow in a compressor cascade using an unstructured grid

NASA Astrophysics Data System (ADS)

Shin, Sangmook

2001-07-01

A three-dimensional unstructured incompressible RANS code has been developed using artificial compressibility and Spalart-Allmaras eddy viscosity model. A node-based finite volume method is used in which all flow variables are defined at the vertices of tetrahedrons in an unstructured grid. The inviscid fluxes are computed by using the Roe's flux difference splitting method, and higher order accuracy is attained by data reconstruction based on Taylor series expansion. Gauss theorem is used to formulate necessary gradients. For time integration, an implicit scheme based on linearized Euler backward method is used. A tetrahedral unstructured grid generation code has been also developed and applied to the tip clearance flow in a highly staggered cascade. Surface grids are first generated in the flow passage and blade tip by using several triangulation methods including Delaunay triangulation, advancing front method and advancing layer method. Then the whole computational domain including tip gap region is filled with prisms using the surface grids. The code has been validated by comparisons with available computational and experimental results for several test cases: inviscid flow around NACA section, laminar and turbulent flow over a flat plate, turbulent flow through double-circular arc cascade and laminar flow through a square duct with 90° bend. Finally the code is applied to a linear cascade that has GE rotor B section with tip clearance and a high stagger angle of 56.9°. The overall structure of the tip clearance flow is well predicted. Loss of loading due to tip leakage flow and reloading due to tip leakage vortex are presented. On the end wall, separation line of the tip leakage vortex and reattachment line of passage vortex are identified. Prediction of such an interaction presents a challenge to RANS computations. The effects of blade span on the flow structure have been also investigated. Two cascades with blades of aspect ratios of 0.5 and 1.0 are considered. By comparing pressure distributions on the blade, it is shown that the aspect ratio has strong effects on loading distribution on the blade although the tip gap height is very small (0.016 chord). Grid convergence study has been carried out with three different grids for pressure distributions and limiting streamlines on the end wall. (Abstract shortened by UMI.)

The complex aerodynamic footprint of desert locusts revealed by large-volume tomographic particle image velocimetry

PubMed Central

Henningsson, Per; Michaelis, Dirk; Nakata, Toshiyuki; Schanz, Daniel; Geisler, Reinhard; Schröder, Andreas; Bomphrey, Richard J.

2015-01-01

Particle image velocimetry has been the preferred experimental technique with which to study the aerodynamics of animal flight for over a decade. In that time, hardware has become more accessible and the software has progressed from the acquisition of planes through the flow field to the reconstruction of small volumetric measurements. Until now, it has not been possible to capture large volumes that incorporate the full wavelength of the aerodynamic track left behind during a complete wingbeat cycle. Here, we use a unique apparatus to acquire the first instantaneous wake volume of a flying animal's entire wingbeat. We confirm the presence of wake deformation behind desert locusts and quantify the effect of that deformation on estimates of aerodynamic force and the efficiency of lift generation. We present previously undescribed vortex wake phenomena, including entrainment around the wing-tip vortices of a set of secondary vortices borne of Kelvin–Helmholtz instability in the shear layer behind the flapping wings. PMID:26040598

Calculation of wake vortex structures in the near-field wake behind cruising aircraft

NASA Astrophysics Data System (ADS)

Ehret, T.; Oertel, H.

Wake flows behind cruising aircraft influence the distribution of the exhaust gases. A three-dimensional vortex filament method was developed to calculate the vortex structures and the velocity field of the vorticity dominated wake flows as an integration of the Biot-Savart law. For three-dimensional vortex filament calculations, self-induction singularities were prevented using a finite vortex core for each vortex filament. Numerical simulations show the vortex structures and the velocity field in the wake behind a cruising Boeing 747 as a result of the integration of the Biot-Savart law. It is further shown how the structures of the fully rolled-up trailing vortices depend on the wing span loading, i.e. the circulation distribution.

Aerodynamic Analysis of Variable Geometry Raked Wingtips for Mid-Range Transonic Transport Aircraft

NASA Astrophysics Data System (ADS)

Jingeleski, David J.

Previous applications have shown that a wingtip treatment on a commercial airliner will reduce drag and increase fuel efficiency and the most common types of treatment are blended winglets and raked wingtips. With Boeing currently investigating novel designs for its next generation of airliners, a variable geometry raked wingtip novel control effector (VGRWT/NCE) was studied to determine the aerodynamic performance benefits over an untreated wingtip. The Boeing SUGAR design employing a truss-braced wing was selected as the baseline. Vortex lattice method (VLM) and computational fluid dynamics (CFD) software was implemented to analyze the aerodynamic performance of such a configuration applied to a next-generation, transonic, mid-range transport aircraft. Several models were created to simulate various sweep positions for the VGRWT/NCE tip, as well as a baseline model with an untreated wingtip. The majority of investigation was conducted using the VLM software, with CFD used largely as a validation of the VLM analysis. The VGRWT/NCE tip was shown to increase the lift of the wing while also decreasing the drag. As expected, the unswept VGRWT/NCE tip increases the amount of lift available over the untreated wingtip, which will be very beneficial for take-off and landing. Similarly, the swept VGRWT/NCE tip reduced the drag of the wing during cruise compared to the unmodified tip, which will favorably impact the fuel efficiency of the aircraft. Also, the swept VGRWT/NCE tip showed an increase in moment compared to the unmodified wingtip, implying an increase in stability, as well providing an avenue for roll control and gust alleviation for flexible wings. CFD analysis validated VLM as a useful low fidelity tool that yielded quite accurate results. The main results of this study are tabulated "deltas" in the forces and moments on the VGRWT/NCE tip as a function of sweep angle and aileron deflection compared to the baseline wing. A side study of the effects of the joint between the main wing and the movable tip showed that the drag impact can be kept small by careful design.

Vortex-flow aerodynamics - An emerging design capability

NASA Technical Reports Server (NTRS)

Campbell, J. F.

1981-01-01

Promising current theoretical and simulational developments in the field of leading edge vortex-generating delta, arrow ogival wings are reported, along with the history of theory and experiment leading to them. The effects of wing slenderness, leading edge nose radius, Mach number and incidence variations, and planform on the onset of vortex generation and redistribution of aerodynamic loads are considered. The range of design possibilities in this field are consequential for the future development of strategic aircraft, supersonic transports and commercial cargo aircraft which will possess low-speed, high-lift capability by virtue of leading edge vortex generation and control without recourse to heavy and expensive leading edge high-lift devices and compound airfoils. Attention is given to interactive graphics simulation devices recently developed.

Flowfield And Download Measurements And Computation of a Tiltrotor Aircraft In Hover

NASA Technical Reports Server (NTRS)

Brand, Albert G.; Peryea, Martin A.; Wood, Tom L.; Meakin, Robert L.

2001-01-01

A multipart study of the V-22 hover flowfield was conducted. Testing involved a 0.15-scale semispan model with multiple independent force balance systems. The velocity flowfield surrounding the airframe was measured using a robotic positioning system and anemometer. Both time averaged and cycle-averaged results are reported. It is shown that the fuselage download in hover can be significantly reduced using a small download reduction device. Measurements indicate that the success of the device is attributed to the substantial elimination of tiltrotor fountain flow. As part of.the study, an unsteady CFD prediction is time-averaged, and shown to have excellent agreement in predicting the baseline configuration fountain flow. Some discrepancies at the outboard edge of the rotor are discussed. An &&sessment of an advanced tip shape rotor comp"'Ietes the study. Derived from a nonrotating study, the advanced tip shape rotor was developed and tested on the Bell 0.15 scale semi-span V-22 model. The tip shape was intended to diffuse the tip vortex and reduce BVI noise. Rotor wake vorticity is extracted from the measured velocity dam to show that the advanced tip shape produces a tip vortex that is only slightly more diffuse than the baseline tip blade. The results indicate that nonrotating tests may overpredict the amount of tip vortex diffusion achieved by tip shape design in a rotating environment.

Validation of a pair of computer codes for estimation and optimization of subsonic aerodynamic performance of simple hinged-flap systems for thin swept wings

NASA Technical Reports Server (NTRS)

Carlson, Harry W.; Darden, Christine M.

1988-01-01

Extensive correlations of computer code results with experimental data are employed to illustrate the use of linearized theory attached flow methods for the estimation and optimization of the aerodynamic performance of simple hinged flap systems. Use of attached flow methods is based on the premise that high levels of aerodynamic efficiency require a flow that is as nearly attached as circumstances permit. A variety of swept wing configurations are considered ranging from fighters to supersonic transports, all with leading- and trailing-edge flaps for enhancement of subsonic aerodynamic efficiency. The results indicate that linearized theory attached flow computer code methods provide a rational basis for the estimation and optimization of flap system aerodynamic performance at subsonic speeds. The analysis also indicates that vortex flap design is not an opposing approach but is closely related to attached flow design concepts. The successful vortex flap design actually suppresses the formation of detached vortices to produce a small vortex which is restricted almost entirely to the leading edge flap itself.

A computer program to calculate the longitudinal aerodynamic characteristics of wing-flap configurations with externally blown flaps

NASA Technical Reports Server (NTRS)

Mendenhall, M. R.; Goodwin, F. K.; Spangler, S. B.

1976-01-01

A vortex lattice lifting-surface method is used to model the wing and multiple flaps. Each lifting surface may be of arbitrary planform having camber and twist, and the multiple-slotted trailing-edge flap system may consist of up to ten flaps with different spans and deflection angles. The engine wakes model consists of a series of closely spaced vortex rings with circular or elliptic cross sections. The rings are normal to a wake centerline which is free to move vertically and laterally to accommodate the local flow field beneath the wing and flaps. The two potential flow models are used in an iterative fashion to calculate the wing-flap loading distribution including the influence of the waves from up to two turbofan engines on the semispan. The method is limited to the condition where the flow and geometry of the configurations are symmetric about the vertical plane containing the wing root chord. The calculation procedure starts with arbitrarily positioned wake centerlines and the iterative calculation continues until the total configuration loading converges within a prescribed tolerance. Program results include total configuration forces and moments, individual lifting-surface load distributions, including pressure distributions, individual flap hinge moments, and flow field calculation at arbitrary field points.

Development of an aerodyanmic theory capable of predicting surface loads on slender wings with vortex flow

NASA Technical Reports Server (NTRS)

Gloss, B. B.; Johnson, F. T.

1976-01-01

The Boeing Commercial Airplane Company developed an inviscid three-dimensional lifting surface method that shows promise in being able to accurately predict loads, subsonic and supersonic, on wings with leading-edge separation and reattachment.

Control of vortex on a non-slender delta wing by a nanosecond pulse surface dielectric barrier discharge

NASA Astrophysics Data System (ADS)

Zhao, Guang-yin; Li, Ying-hong; Liang, Hua; Han, Meng-hu; Hua, Wei-zhuo

2015-01-01

Wind tunnel experiments are conducted for improving the aerodynamic performance of delta wing using a leading-edge pulsed nanosecond dielectric barrier discharge (NS-DBD). The whole effects of pulsed NS-DBD on the aerodynamic performance of the delta wing are studied by balanced force measurements. Pressure measurements and particle image velocimetry (PIV) measurements are conducted to investigate the formation of leading-edge vortices affected by the pulsed NS-DBD, compared to completely stalled flow without actuation. Various pulsed actuation frequencies of the plasma actuator are examined with the freestream velocity up to 50 m/s. Stall has been delayed substantially and significant shifts in the aerodynamic forces can be achieved at the post-stall regions when the actuator works at the optimum reduced frequency of F + = 2. The upper surface pressure measurements show that the largest change of static pressure occurs at the forward part of the wing at the stall region. The time-averaged flow pattern obtained from the PIV measurement shows that flow reattachment is promoted with excitation, and a vortex flow pattern develops. The time-averaged locations of the secondary separation line and the center of the vortical region both move outboard with excitation.

A Method for Computing the Core Flow in Three-Dimensional Leading-Edge Vortices. Ph.D. Thesis - North Carolina State Univ.

NASA Technical Reports Server (NTRS)

Luckring, J. M.

1985-01-01

A theory is presented for calculating the flow in the core of a separation-induced leading-edge vortex. The method is based on matching inner and outer representations of the vortex. The inner model of the vortex is based on the quasicylindrical Navier-Stokes equations; the flow is assumed to be steady, axially symmetric, and incompressible and in addition, gradients in the radial direction are assumed to be much larger then gradients in the axial direction. The outer model is based on the three-dimensional free-vortex-sheet theory, a higher-order panel method which solves the Prandtl-Glauert equation including nonlinear boundary conditions pertinent to the concentrated vorticity representation of the leading edge vortex. The resultant flow is evaluated a posteriori for evidence of incipient vortex breakdown and the critical helix angle concept, in conjunction with an adverse longitudinal pressure gradient, is found to correlate well with the occurrence of vortex breakdown at the trailing edge of delta, arrow, and diamond wings.

VORCOR: A computer program for calculating characteristics of wings with edge vortex separation by using a vortex-filament and-core model

NASA Technical Reports Server (NTRS)

Pao, J. L.; Mehrotra, S. C.; Lan, C. E.

1982-01-01

A computer code base on an improved vortex filament/vortex core method for predicting aerodynamic characteristics of slender wings with edge vortex separations is developed. The code is applicable to camber wings, straked wings or wings with leading edge vortex flaps at subsonic speeds. The prediction of lifting pressure distribution and the computer time are improved by using a pair of concentrated vortex cores above the wing surface. The main features of this computer program are: (1) arbitrary camber shape may be defined and an option for exactly defining leading edge flap geometry is also provided; (2) the side edge vortex system is incorporated.

The leading-edge vortex of swift wing-shaped delta wings

NASA Astrophysics Data System (ADS)

Muir, Rowan Eveline; Arredondo-Galeana, Abel; Viola, Ignazio Maria

2017-08-01

Recent investigations on the aerodynamics of natural fliers have illuminated the significance of the leading-edge vortex (LEV) for lift generation in a variety of flight conditions. A well-documented example of an LEV is that generated by aircraft with highly swept, delta-shaped wings. While the wing aerodynamics of a manoeuvring aircraft, a bird gliding and a bird in flapping flight vary significantly, it is believed that this existing knowledge can serve to add understanding to the complex aerodynamics of natural fliers. In this investigation, a model non-slender delta-shaped wing with a sharp leading edge is tested at low Reynolds number, along with a delta wing of the same design, but with a modified trailing edge inspired by the wing of a common swift Apus apus. The effect of the tapering swift wing on LEV development and stability is compared with the flow structure over the unmodified delta wing model through particle image velocimetry. For the first time, a leading-edge vortex system consisting of a dual or triple LEV is recorded on a swift wing-shaped delta wing, where such a system is found across all tested conditions. It is shown that the spanwise location of LEV breakdown is governed by the local chord rather than Reynolds number or angle of attack. These findings suggest that the trailing-edge geometry of the swift wing alone does not prevent the common swift from generating an LEV system comparable with that of a delta-shaped wing.

The leading-edge vortex of swift wing-shaped delta wings

PubMed Central

Muir, Rowan Eveline; Arredondo-Galeana, Abel

2017-01-01

Recent investigations on the aerodynamics of natural fliers have illuminated the significance of the leading-edge vortex (LEV) for lift generation in a variety of flight conditions. A well-documented example of an LEV is that generated by aircraft with highly swept, delta-shaped wings. While the wing aerodynamics of a manoeuvring aircraft, a bird gliding and a bird in flapping flight vary significantly, it is believed that this existing knowledge can serve to add understanding to the complex aerodynamics of natural fliers. In this investigation, a model non-slender delta-shaped wing with a sharp leading edge is tested at low Reynolds number, along with a delta wing of the same design, but with a modified trailing edge inspired by the wing of a common swift Apus apus. The effect of the tapering swift wing on LEV development and stability is compared with the flow structure over the unmodified delta wing model through particle image velocimetry. For the first time, a leading-edge vortex system consisting of a dual or triple LEV is recorded on a swift wing-shaped delta wing, where such a system is found across all tested conditions. It is shown that the spanwise location of LEV breakdown is governed by the local chord rather than Reynolds number or angle of attack. These findings suggest that the trailing-edge geometry of the swift wing alone does not prevent the common swift from generating an LEV system comparable with that of a delta-shaped wing. PMID:28878968

The leading-edge vortex of swift wing-shaped delta wings.

PubMed

Muir, Rowan Eveline; Arredondo-Galeana, Abel; Viola, Ignazio Maria

2017-08-01

Recent investigations on the aerodynamics of natural fliers have illuminated the significance of the leading-edge vortex (LEV) for lift generation in a variety of flight conditions. A well-documented example of an LEV is that generated by aircraft with highly swept, delta-shaped wings. While the wing aerodynamics of a manoeuvring aircraft, a bird gliding and a bird in flapping flight vary significantly, it is believed that this existing knowledge can serve to add understanding to the complex aerodynamics of natural fliers. In this investigation, a model non-slender delta-shaped wing with a sharp leading edge is tested at low Reynolds number, along with a delta wing of the same design, but with a modified trailing edge inspired by the wing of a common swift Apus apus . The effect of the tapering swift wing on LEV development and stability is compared with the flow structure over the unmodified delta wing model through particle image velocimetry. For the first time, a leading-edge vortex system consisting of a dual or triple LEV is recorded on a swift wing-shaped delta wing, where such a system is found across all tested conditions. It is shown that the spanwise location of LEV breakdown is governed by the local chord rather than Reynolds number or angle of attack. These findings suggest that the trailing-edge geometry of the swift wing alone does not prevent the common swift from generating an LEV system comparable with that of a delta-shaped wing.

Recent theoretical developments and experimental studies pertinent to vortex flow aerodynamics - With a view towards design

NASA Technical Reports Server (NTRS)

Lamar, J. E.; Luckring, J. M.

1978-01-01

A review is presented of recent progress in a research program directed towards the development of an improved vortex-flow technology base. It is pointed out that separation induced vortex-flows from the leading and side edges play an important role in the high angle-of-attack aerodynamic characteristics of a wide range of modern aircraft. In the analysis and design of high-speed aircraft, a detailed knowledge of this type of separation is required, particularly with regard to critical wind loads and the stability and performance at various off-design conditions. A description of analytical methods is presented. The theoretical methods employed are divided into two classes which are dependent upon the underlying aerodynamic assumptions. One conical flow method is considered along with three different nonconical flow methods. Comparisons are conducted between the described methods and available aerodynamic data. Attention is also given to a vortex flow drag study and a vortex flow wing design using suction analogy.

On the stability and control of a trailing vortex

NASA Astrophysics Data System (ADS)

Edstrand, Adam M.

Trailing vortices are both a fundamental and practical problem of fluid mechanics. Fundamentally, they provide a canonical vortex flow that is pervasive in finite aspect ratio lifting bodies, practically producing many adverse effects across aeronautical and maritime applications. These adverse effects coupled with the broad range of applicability make their active control desirable; however, they remain robust to control efforts. Experimental baseline results provided an explanation of vortex wandering, the side-to-side motion often attributed to wind-tunnel unsteadiness or a vortex instability. We extracted the wandering motion and found striking similarities with the eigenmodes, growth rates, and frequencies from a stability analysis of the Batchelor vortex. After concluding that wandering is a result of a vortex instability, we applied control to the trailing vortex flow field through blowing from a slot at the wingtip. We experimentally obtained modest reductions in the metrics, but found the parameter space for optimization unwieldy. With the ultimate goal of designing control, we performed a physics-based stability analysis in the wake of a NACA0012 wing with an aspect ratio of 1.25 positioned at a geometric angle of attack of 5 degrees. Numerically computing the base flow at a chord Reynolds number of 1000, we perform a parallel temporal and spatial stability analysis three chords downstream of the trailing edge finding seven instabilities: three temporal, four spatial. The three temporal contain a wake instability, a vortex instability, and a mixed instability, which is a higher-order wake instability. The primary instability localized to the wake results from the two-dimensional wake, while the secondary instability is the mixed instability, containing higher-order spanwise structures in the wake. These instabilities imply that although it may be intuitive to place control at the wingtip, these results show that control may be more effective at the trailing edge, which would excite these instabilities that result with the eventual break up of the vortex. Further, by performing a wave-packet analysis, we found the wave packets contained directivity, coming inward toward the vortex above and below the wing, and traveling outward in the spanwise directions. We conjecture that this directivity can be translated to receptivity, with free-stream disturbances above and below the wing being more receptive than spanwise disturbances. With this, we provide two methods for instability excitation: utilizing control devices on the wing to excite near-field instabilities directly and utilizing free-stream disturbances to such as a speaker to excite near-field instabilities through receptivity.

Assessment of computational prediction of tail buffeting

NASA Technical Reports Server (NTRS)

Edwards, John W.

1990-01-01

Assessments of the viability of computational methods and the computer resource requirements for the prediction of tail buffeting are made. Issues involved in the use of Euler and Navier-Stokes equations in modeling vortex-dominated and buffet flows are discussed and the requirement for sufficient grid density to allow accurate, converged calculations is stressed. Areas in need of basic fluid dynamics research are highlighted: vorticity convection, vortex breakdown, dynamic turbulence modeling for free shear layers, unsteady flow separation for moderately swept, rounded leading-edge wings, vortex flows about wings at high subsonic speeds. An estimate of the computer run time for a buffeting response calculation for a full span F-15 aircraft indicates that an improvement in computer and/or algorithm efficiency of three orders of magnitude is needed to enable routine use of such methods. Attention is also drawn to significant uncertainties in the estimates, in particular with regard to nonlinearities contained within the modeling and the question of the repeatability or randomness of buffeting response.

Aerodynamic sound generation of flapping wing.

PubMed

Bae, Youngmin; Moon, Young J

2008-07-01

The unsteady flow and acoustic characteristics of the flapping wing are numerically investigated for a two-dimensional model of Bombus terrestris bumblebee at hovering and forward flight conditions. The Reynolds number Re, based on the maximum translational velocity of the wing and the chord length, is 8800 and the Mach number M is 0.0485. The computational results show that the flapping wing sound is generated by two different sound generation mechanisms. A primary dipole tone is generated at wing beat frequency by the transverse motion of the wing, while other higher frequency dipole tones are produced via vortex edge scattering during a tangential motion. It is also found that the primary tone is directional because of the torsional angle in wing motion. These features are only distinct for hovering, while in forward flight condition, the wing-vortex interaction becomes more prominent due to the free stream effect. Thereby, the sound pressure level spectrum is more broadband at higher frequencies and the frequency compositions become similar in all directions.

An analytical design procedure for the determination of effective leading edge extensions on thick delta wings

NASA Technical Reports Server (NTRS)

Ghaffari, F.; Chaturvedi, S. K.

1984-01-01

An analytical design procedure for leading edge extensions (LEE) was developed for thick delta wings. This LEE device is designed to be mounted to a wing along the pseudo-stagnation stream surface associated with the attached flow design lift coefficient of greater than zero. The intended purpose of this device is to improve the aerodynamic performance of high subsonic and low supersonic aircraft at incidences above that of attached flow design lift coefficient, by using a vortex system emanating along the leading edges of the device. The low pressure associated with these vortices would act on the LEE upper surface and the forward facing area at the wing leading edges, providing an additional lift and effective leading edge thrust recovery. The first application of this technique was to a thick, round edged, twisted and cambered wing of approximately triangular planform having a sweep of 58 deg and aspect ratio of 2.30. The panel aerodynamics and vortex lattice method with suction analogy computer codes were employed to determine the pseudo-stagnation stream surface and an optimized LEE planform shape.

Influence of numerical dissipation in computing supersonic vortex-dominated flows

NASA Technical Reports Server (NTRS)

Kandil, O. A.; Chuang, A.

1986-01-01

Steady supersonic vortex-dominated flows are solved using the unsteady Euler equations for conical and three-dimensional flows around sharp- and round-edged delta wings. The computational method is a finite-volume scheme which uses a four-stage Runge-Kutta time stepping with explicit second- and fourth-order dissipation terms. The grid is generated by a modified Joukowski transformation. The steady flow solution is obtained through time-stepping with initial conditions corresponding to the freestream conditions, and the bow shock is captured as a part of the solution. The scheme is applied to flat-plate and elliptic-section wings with a leading edge sweep of 70 deg at an angle of attack of 10 deg and a freestream Mach number of 2.0. Three grid sizes of 29 x 39, 65 x 65 and 100 x 100 have been used. The results for sharp-edged wings show that they are consistent with all grid sizes and variation of the artificial viscosity coefficients. The results for round-edged wings show that separated and attached flow solutions can be obtained by varying the artificial viscosity coefficients. They also show that the solutions are independent of the way time stepping is done. Local time-stepping and global minimum time-steeping produce same solutions.

Developments and Validations of Fully Coupled CFD and Practical Vortex Transport Method for High-Fidelity Wake Modeling in Fixed and Rotary Wing Applications

NASA Technical Reports Server (NTRS)

Anusonti-Inthra, Phuriwat

2010-01-01

A novel Computational Fluid Dynamics (CFD) coupling framework using a conventional Reynolds-Averaged Navier-Stokes (BANS) solver to resolve the near-body flow field and a Particle-based Vorticity Transport Method (PVTM) to predict the evolution of the far field wake is developed, refined, and evaluated for fixed and rotary wing cases. For the rotary wing case, the RANS/PVTM modules are loosely coupled to a Computational Structural Dynamics (CSD) module that provides blade motion and vehicle trim information. The PVTM module is refined by the addition of vortex diffusion, stretching, and reorientation models as well as an efficient memory model. Results from the coupled framework are compared with several experimental data sets (a fixed-wing wind tunnel test and a rotary-wing hover test).

Mitigation of tip vortex cavitation by means of air injection on a Kaplan turbine scale model

NASA Astrophysics Data System (ADS)

Rivetti, A.; Angulo, M.; Lucino, C.; Liscia, S.

2014-03-01

Kaplan turbines operating at full-load conditions may undergo excessive vibration, noise and cavitation. In such cases, damage by erosion associated to tip vortex cavitation can be observed at the discharge ring. This phenomenon involves design features such as (1) overhang of guide vanes; (2) blade profile; (3) gap increasing size with blade opening; (4) suction head; (5) operation point; and (6) discharge ring stiffness, among others. Tip vortex cavitation may cause erosion at the discharge ring and draft tube inlet following a wavy pattern, in which the number of vanes can be clearly identified. Injection of pressurized air above the runner blade centerline was tested as a mean to mitigate discharge ring cavitation damage on a scale model. Air entrance was observed by means of a high-speed camera in order to track the air trajectory toward its mergence with the tip vortex cavitation core. Post-processing of acceleration signals shows that the level of vibration and the RSI frequency amplitude decrease proportionally with air flow rate injected. These findings reveal the potential mitigating effect of air injection in preventing cavitation damage and will be useful in further tests to be performed on prototype, aiming at determining the optimum air flow rate, size and distribution of the injectors.

Wind Tunnel Investigation of Passive Porosity Applied to the Leading-Edge Extension and Leading-Edge Flaps on a Slender Wing at Subsonic Speed

NASA Technical Reports Server (NTRS)

Erickson, Gary E.

2017-01-01

A wind tunnel experiment was conducted in the NASA Langley Research Center 7- by 10-Foot High Speed Tunnel to determine the effects of passive surface porosity on the subsonic vortex flow interactions about a general research fighter configuration. Flow-through porosity was applied to the leading-edge extension, or LEX, and leading-edge flaps mounted to a 65deg cropped delta wing model as a potential vortex flow control technique at high angles of attack. All combinations of porous and nonporous LEX and flaps were investigated. Wing upper surface static pressure distributions and six-component forces and moments were obtained at a free-stream Mach number of 0.20 corresponding to a Reynolds number of 1.35(106) per foot, angles of attack up to 45deg, angles of sideslip of 0deg and +/-5deg, and leading-edge flap deflections of 0deg and 30deg.

CFD simulations of a wind turbine for analysis of tip vortex breakdown

NASA Astrophysics Data System (ADS)

Kimura, K.; Tanabe, Y.; Aoyama, T.; Matsuo, Y.; Arakawa, C.; Iida, M.

2016-09-01

This paper discusses about the wake structure of wind turbine via the use of URANS and Quasi-DNS, focussing on the tip vortex breakdown. The moving overlapped structured grids CFD Solver based on a fourth-order reconstruction and an all-speed scheme, rFlow3D is used for capturing the characteristics of tip vortices. The results from the Model Experiments in Controlled Conditions project (MEXICO) was accordingly selected for executing wake simulations through the variation of tip speed ratio (TSR); in an operational wind turbine, TSR often changes in value. Therefore, it is important to assess the potential effects of TSR on wake characteristics. The results obtained by changing TSR show the variations of the position of wake breakdown and wake expansion. The correspondence between vortices and radial/rotational flow is also confirmed.

Recent Advancements in the Infrared Flow Visualization System for the NASA Ames Unitary Plan Wind Tunnels

NASA Technical Reports Server (NTRS)

Garbeff, Theodore J., II; Baerny, Jennifer K.

2017-01-01

The following details recent efforts undertaken at the NASA Ames Unitary Plan wind tunnels to design and deploy an advanced, production-level infrared (IR) flow visualization data system. Highly sensitive IR cameras, coupled with in-line image processing, have enabled the visualization of wind tunnel model surface flow features as they develop in real-time. Boundary layer transition, shock impingement, junction flow, vortex dynamics, and buffet are routinely observed in both transonic and supersonic flow regimes all without the need of dedicated ramps in test section total temperature. Successful measurements have been performed on wing-body sting mounted test articles, semi-span floor mounted aircraft models, and sting mounted launch vehicle configurations. The unique requirements of imaging in production wind tunnel testing has led to advancements in the deployment of advanced IR cameras in a harsh test environment, robust data acquisition storage and workflow, real-time image processing algorithms, and evaluation of optimal surface treatments. The addition of a multi-camera IR flow visualization data system to the Ames UPWT has demonstrated itself to be a valuable analyses tool in the study of new and old aircraft/launch vehicle aerodynamics and has provided new insight for the evaluation of computational techniques.

Large-Vortex Capture by a Wing at Very High Angles of Attack

NASA Technical Reports Server (NTRS)

Wu, J. M.; Wu, J. Z.; Denny, G. A.; Lu, X. Y.

1996-01-01

In generating the lift on a wing, the static stall is a severe barrier. As the angle of attack, alpha, increases to the stall angle, alpha(sub stall) the flow separation point on the upper surface of the wing moves to the leading edge, so that on a two-dimensional airfoil or a large-aspect-ratio wing, the lift abruptly drops to a very low level. Therefore, the first generation of aeronautical flow type, i.e., the attached steady flow, has been limited to alpha less than alpha(sub stall). Owing to the obvious importance in applications, therefore, a great effort has been made in the past two decades to enlarge the range of usable angles of attack by various flow controls for a large-aspect-ratio wing. Basically, relevant works fall into two categories. The first category is usually refereed to as separation control, which concentrates on partially separated flow at alpha less than alpha(sub stall). Since the first experimental study of Collins and Zelenevitz, there has been ample literature showing that a partially separated flow can be turned to almost fully attached by flow controls, so that the lift is recovered and the stall is delayed (for a recent work see Seifert et al.). It has been well established that, in this category, unsteady controls are much more effective than steady ones and can be realized at a very low power-input level (Wu et al.; Seifert et al.). The second and more ambitious category of relevant efforts is the post-stall lift enhancement. Its possibility roots at the existence of a second lift peak at a very high angle of attack. In fact, As alpha further increases from alpha(sub stall), the completely separated flow develops and gradually becomes a bluff-body flow. This flow gives a normal force to the airfoil with a lift component, which reaches a peak at a maximum utilizable angle of attack, alpha(sub m) approx.= 40 deg. This second peak is of the same level as the first lift peak at alpha(sub stall). Meanwhile, the drag is also quickly increased (e.g., Fage and Johansen ; Critzos et al.). Figure 1 shows a typical experimental lift and drag coefficients of NACA-0012 airfoil in this whole range of angle of attack. Obviously, without overcoming the lift crisis at alpha(sub stall) the second lift peak is completely useless. Thus, the ultimate goal of post-stall lift enhancement is to fill the lift valley after stall by flow controls, so that a wing and/or flap can work at the whole range of 0 deg less than alpha less than alpha(sub m). Relevant early experimental studies have been extensively reviewed by Wu et al., who concluded that, first, similar to the leading-edge vortex on a slender wing, the lift enhancement on a large-aspect-ratio wing should be the result of capturing a vortex on the upper surface of the wing; and, second, using steady controls cannot reach the goal, and one must rely on unsteady controls with low-level power input as well. Wu et al. also conjectured that the underlying physics of post-stall lift enhancement by unsteady controls consists of a chain of mechanisms: vortex layer instability - receptivity resonance - nonlinear streaming.

Breaking down the delta wing vortex: The role of vorticity in the breakdown process. Ph.D. Thesis Final Report

NASA Technical Reports Server (NTRS)

Nelson, Robert C.; Visser, Kenneth D.

1990-01-01

Experimental x-wire measurements of the flowfield above a 70 and 75 deg flat plate delta wing were performed at a Reynolds number of 250,000. Grids were taken normal to the wing at various chordwise locations for angles of attack of 20 and 30 deg. Axial and azimuthal vorticity distributions were derived from the velocity fields. The dependence of circulation on distance from the vortex core and on chordwise location was also examined. The effects of nondimensionalization in comparison with other experimental data is made. The results indicate that the circulation distribution scales with the local semispan and grows in a nearly linear fashion in the chordwise direction. The spanwise distribution of axial vorticity is severely altered through the breakdown. The axial vorticity components with a negative sense, such as that found in the secondary vortex, seem to remain unaffected by changes in wind sweep or angle of attack, in direct contrast to the positive components. In addition, the inclusion of the local wing geometry into a previously derived correlation parameter allows the circulation of growing leading edge vortex flows to be reduced into a single curve.

Wingtip Vortices and Free Shear Layer Interaction in the Vicinity of Maximum Lift to Drag Ratio Lift Condition

NASA Astrophysics Data System (ADS)

Memon, Muhammad Omar

Cost-effective air-travel is something everyone wishes for when it comes to booking flights. The continued and projected increase in commercial air travel advocates for energy efficient airplanes, reduced carbon footprint, and a strong need to accommodate more airplanes into airports. All of these needs are directly affected by the magnitudes of drag these aircraft experience and the nature of their wingtip vortex. A large portion of the aerodynamic drag results from the airflow rolling from the higher pressure side of the wing to the lower pressure side, causing the wingtip vortices. The generation of this particular drag is inevitable however, a more fundamental understanding of the phenomenon could result in applications whose benefits extend much beyond the relatively minuscule benefits of commonly-used winglets. Maximizing airport efficiency calls for shorter intervals between takeoffs and landings. Wingtip vortices can be hazardous for following aircraft that may fly directly through the high-velocity swirls causing upsets at vulnerably low speeds and altitudes. The vortex system in the near wake is typically more complex since strong vortices tend to continue developing throughout the near wake region. Several chord lengths distance downstream of a wing, the so-called fully rolled up wing wake evolves into a combination of a discrete wingtip vortex pair and a free shear layer. Lift induced drag is generated as a byproduct of downwash induced by the wingtip vortices. The parasite drag results from a combination of form/pressure drag and the upper and lower surface boundary layers. These parasite effects amalgamate to create the free shear layer in the wake. While the wingtip vortices embody a large portion of the total drag at lifting angles, flow properties in the free shear layer also reveal their contribution to the aerodynamic efficiency of the aircraft. Since aircraft rarely cruise at maximum aerodynamic efficiency, a better understanding of the balance between the lift induced drag (wingtip vortices) and parasite drag (free shear layer) can have a significant impact. Particle Image Velocimetry (PIV) experiments were performed at a) a water tunnel at ILR Aachen, Germany, and b) at the University of Dayton Low Speed Wind Tunnel in the near wake of an AR 6 wing with a Clark-Y airfoil to investigate the characteristics of the wingtip vortex and free shear layer at angles of attack in the vicinity of maximum aerodynamic efficiency for the wing. The data was taken 1.5 and 3 chord lengths downstream of the wing at varying free-stream velocities. A unique exergy-based technique was introduced to quantify distinct changes in the wingtip vortex axial core flow. The existence of wingtip vortex axial core flow transformation from wake-like (velocity less-than the freestream) to jet-like (velocity greater-than the freestream) behavior in the vicinity of the maximum (L/D) angles was observed. The exergy-based technique was able to identify the change in the out of plane profile and corresponding changes in the L/D performance. The resulting velocity components in and around the free shear layer in the wing wake showed counter flow in the cross-flow plane presumably corresponding to behavior associated with the flow over the upper and lower surfaces of the wing. Even though the velocity magnitudes in the free shear layer in cross-flow plane are a small fraction of the freestream velocity ( 10%), significant directional flow was observed. An indication of the possibility of the transfer of momentum (from inboard to outboard of the wing) was identified through spanwise flow corresponding to the upper and lower surfaces through the free shear layer in the wake. A transition from minimal cross flow in the free shear layer to a well-established shear flow in the spanwise direction occurs in the vicinity of maximum lift-to-drag ratio (max L/D) angle of attack. A distinctive balance between the lift induced drag and parasite drag was identified. Improved understanding of this relationship could be extended not only to improve aircraft performance through the reduction of lift induced drag, but also to air vehicle performance in off-design cruise conditions.

Technical Evaluation Report, Part A - Vortex Flow and High Angle of Attack

NASA Technical Reports Server (NTRS)

Luckring, James M.

2003-01-01

A symposium entitled Vortex Flow and High Angle of Attack was held in Loen, Norway, from May 7 through May 11, 2001. The Applied Vehicle Technology (AVT) panel, under the auspices of the Research and Technology Organization (RTO), sponsored this symposium. Forty-eight papers, organized into nine sessions, addressed computational and experimental studies of vortex flows pertinent to both aircraft and maritime applications. The studies also ranged from fundamental fluids investigations to flight test results, and significant results were contributed from a broad range of countries. The principal emphasis of this symposium was on "the understanding and prediction of separation-induced vortex flows and their effects on military vehicle performance, stability, control, and structural design loads." It was further observed by the program committee that "separation- induced vortex flows are an important part of the design and off-design performance of conventional fighter aircraft and new conventional or unconventional manned or unmanned advanced vehicle designs (UAVs, manned aircraft, missiles, space planes, ground-based vehicles, and ships)." The nine sessions addressed the following topics: vortical flows on wings and bodies, experimental techniques for vortical flows, numerical simulations of vortical flows, vortex stability and breakdown, vortex flows in maritime applications, vortex interactions and control, vortex dynamics, flight testing, and vehicle design. The purpose of this paper is to provide brief reviews of these papers along with some synthesizing perspectives toward future vortex flow research opportunities. The paper includes the symposium program. (15 refs.)

An analytical parametric study of the broadband noise from axial-flow fans

NASA Technical Reports Server (NTRS)

Chou, Shau-Tak; George, Albert R.

1987-01-01

The rotating dipole analysis of Ffowcs Williams and Hawkings (1969) is used to predict the far field noise radiation due to various rotor broadband noise mechanisms. Consideration is given to inflow turbulence noise, attached boundary layer/trailing-edge interaction noise, tip-vortex formation noise, and trailing-edge thickness noise. The parametric dependence of broadband noise from unducted axial-flow fans on several critical variables is studied theoretically. The angle of attack of the rotor blades, which is related to the rotor performance, is shown to be important to the trailing-edge noise and to the tip-vortex formation noise.

Theoretical analysis of an augmentor wing for a VTOL fighter

NASA Technical Reports Server (NTRS)

Dillenius, M. F. E.; Mendenhall, M. R.

1979-01-01

A method based on potential flow theory was developed for predicting forces and moments acting on augmentor wings for prescribed ejector jet characteristics. A three dimensional nonplanar vortex lattice is laid out on the chordal planes of the augmentor wing components. Jet induced effects are included in the boundary condition from which the horseshoe vortex strengths are obtained. The jet within the diffusor is made to expand from the primary nozzles to the diffusor exit and is represented by a distribution of vorticity on the jet boundary to provide proper entrainment. The jet downstream of the diffusor exit is modeled by a vorticity distribution and blockage panels and its centerline location and spreading rate are taken from experimental data. The vortex lattice and jet models are used in an iterative manner until the predicted diffusor exit velocity matches the specified one. Some comparisons with available data show good agreement at lower power settings.

Interaction of a Vortex Pair with a Free Surface.

DTIC Science & Technology

1987-09-01

larger than normal in a wing in a wind tunnel), or the unsteady nature of the flow (for vortices generated by a wing in a tow basin) makes the mean...tension can be assumed to be negligible. This observation has been verified experimentally by Gray [Ref. 17]. Thus, if the flow is considered inviscid...analytical solutions in the exterior regions could be obtained without too much difficulty. In particular, treatment of steady flows in a uniform

Secondary Vortex Structures in Vortex Generator Induced Flow

NASA Astrophysics Data System (ADS)

Velte, Clara; Okulov, Valery; Hansen, Martin

2010-11-01

Passive rectangular vane actuators can induce a longitudinal vortex that redistributes the momentum in the boundary layer to control the flow. Recent experiments [1] as well as previous studies [2] have shown that a secondary vortex of opposite sign is generated along with the primary one, supposedly from local separation of the boundary layer due to the primary vortex. 2D flow visualizations of a vortex in the vicinity of a boundary support this hypothesis [3]. These secondary vortices are studied for various configurations -- single generator, counter- and co-rotating cascades. The objective is to study their removal through cancelation in cascades using Stereoscopic Particle Image Velocimetry and flow visualization.[4pt] [1] Velte, Hansen and Okulov, J. Fluid Mech. 619, 2009.[0pt] [2] Zhang, Int. J. Heat Fluid Flow 21 2000.[0pt] [3] Harris, Miller and Williamson, APS abstract 2009.

Boeing 747 with Smoke Generator Installed for Vortex Study

NASA Technical Reports Server (NTRS)

1979-01-01

The Boeing 747 used for wingtip vortex research flights sits on the ramp at NASA's Flight Research Center in Edwards, California. Note the smoke generator mounted underneath the jet's wing. Smoke from underwing smoke generators made it possible for researchers to actually see the vortices created by the 747's wings in flight. In 1974 the NASA Flight Research Center (later Dryden Flight Research Center, Edwards, California) used a Boeing 747 as part of the overall NASA study of trailing vortices. Trailing vortices are the invisible flow of spiraling air that trails from the wings of large aircraft and can 'upset' smaller aircraft flying behind them. The 747 that NASA used was on loan from the Johnson Space Center where it was part of the Space Shuttle Program. The data gathered in the 747 studies complemented data from the previous (1973-74) joint NASA Flight Research Center and Federal Aviation Administration (FAA) Boeing727 wake vortices study. Six smoke generators were installed under the wings of the 747 to provide a visual image of the trailing vortices. The object of the experiments was to test different configurations and mechanical devices on the747 that could be used to break up or lessen the strength of the vortices. The results of the tests could lead to shorter spacing between landings and takeoffs, which, in turn, could alleviate air-traffic congestion. For approximately 30 flights the 747 was flown using various combinations of wing air spoilers in an attempt to reduce wake vortices. To evaluate the effectiveness of the different configurations, chase aircraft were flown into the vortex sheets to probe their strengths and patterns at different times. Two of the chase planes used were the Flight Research Center's Cessna T-37 and the NASA Ames Research Center's Learjet. These aircraft represented the types of smaller business jets and other small aircraft that might encounter large passenger aircraft on approach or landings around major airports or in flight. Tests without the 747's wing spoilers deployed produced violent 'upset' problems for the T-37 aircraft at a distance of approximately 3 miles. From the magnitude of the problems found, distances of as much as ten miles might be required if spoilers were not used. With two spoilers on the outer wing panels, the T-37 could fly at a distance of three miles and not experience the 'upset' problem. The wake vortex study continued even after the 747 was returned to its primary mission of carrying the Space Shuttle.

Lift Augmentation on a Delta Wing via Leading Edge Fences and the Gurney Flap

NASA Technical Reports Server (NTRS)

Buchholz, Mark D.; Tso, Jin

1993-01-01

Wind tunnel tests have been conducted on two devices for the purpose of lift augmentation on a 60 deg delta wing at low speed. Lift, drag, pitching moment, and surface pressures were measured. Detailed flow visualization was also obtained. Both the leading edge fence and the Gurney flap are shown to increase lift. The fences and flap shift the lift curve by as much as 5 deg and 10 deg, respectively. The fences aid in trapping vortices on the upper surface, thereby increasing suction. The Gurney flap improves circulation at the trailing edge. The individual influences of both devices are roughly additive, creating high lift gain. However, the lower lift to drag ratio and the precipitation of vortex burst caused by the fences, and the nose down pitching moment created by the flap are also significant factors.

Visualization of flow separation and control by vortex generators on an single flap in landing configuration

NASA Astrophysics Data System (ADS)

Součková, Natálie; Kuklová, Jana; Popelka, Lukáš; Matějka, Milan

2012-04-01

This paper focuses on a suppression of the flow separation, which occurs on a deflected flap, by means of vortex generators (VG's). An airfoil NACA 63A421 with a simple flap and vane-type vortex generators were used. The investigation was carried out by using experimental and numerical methods. The data from the numerical simulation of the flapped airfoil without VG's control were used for the vortex generator design. Two sizes, two different shapes and various spacing of the vortex generators were tested. The flow past the airfoil was visualized through three methods, namely tuft filaments technique, oil and thermo camera visualization. The experiments were performed in closed circuit wind tunnels with closed and open test sections. The lift curves for both cases without and with vortex generators were acquired for a lift coefficient improvement determination. The improvement was achieved for several cases by means all of the applied methods.

Buffet characteristics of the F-8 supercritical wing airplane

NASA Technical Reports Server (NTRS)

Deangelis, V. M.; Monaghan, R. C.

1977-01-01

The buffet characteristics of the F-8 supercritical wing airplane were investigated. Wing structural response was used to determine the buffet characteristics of the wing and these characteristics are compared with wind tunnel model data and the wing flow characteristics at transonic speeds. The wingtip accelerometer was used to determine the buffet onset boundary and to measure the buffet intensity characteristics of the airplane. The effects of moderate trailing edge flap deflections on the buffet onset boundary are presented. The supercritical wing flow characteristics were determined from wind tunnel and flight static pressure measurements and from a dynamic pressure sensor mounted on the flight test airplane in the vicinity of the shock wave that formed on the upper surface of the wing at transonic speeds. The comparison of the airplane's structural response data to the supercritical flow characteristics includes the effects of a leading edge vortex generator.

Aeroacoustic Study of a 26%-Scale Semispan Model of a Boeing 777 Wing in the NASA Ames 40- by 80-Foot Wind Tunnel

NASA Technical Reports Server (NTRS)

Horne, W. Clifton; Burnside, Nathan J.; Soderman, Paul T.; Jaeger, Stephen M.; Reinero, Bryan R.; James, Kevin D.; Arledge, Thomas K.

2004-01-01

An acoustic and aerodynamic study was made of a 26%-scale unpowered Boeing 777 aircraft semispan model in the NASA Ames 40- by 80-Foot Wind Tunnel for the purpose of identifying and attenuating airframe noise sources. Simulated approach and landing configurations were evaluated at Mach numbers between 0.12 and 0.24. Cruise configurations were evaluated at Mach numbers between 0.24 and 0.33. The research team used two Ames phased-microphone arrays, a large fixed array and a small traversing array, mounted under the wing to locate and compare various noise sources in the wing high-lift system and landing gear. Numerous model modifications and noise alleviation devices were evaluated. Simultaneous with acoustic measurements, aerodynamic forces were recorded to document aircraft conditions and any performance changes caused by the geometric modifications. Numerous airframe noise sources were identified that might be important factors in the approach and landing noise of the full-scale aircraft. Several noise-control devices were applied to each noise source. The devices were chosen to manipulate and control, if possible, the flow around the various tips and through the various gaps of the high-lift system so as to minimize the noise generation. Fences, fairings, tip extensions, cove fillers, vortex generators, hole coverings, and boundary-layer trips were tested. In many cases, the noise-control devices eliminated noise from some sources at specific frequencies. When scaled to full-scale third-octave bands, typical noise reductions ranged from 1 to 10 dB without significant aerodynamic performance loss.

Vortex topology of rolling and pitching wings

NASA Astrophysics Data System (ADS)

Johnson, Kyle; Thurow, Brian; Wabick, Kevin; Buchholz, James; Berdon, Randall

2017-11-01

A flat, rectangular plate with an aspect ratio of 2 was articulated in roll and pitch, individually and simultaneously, to isolate the effects of each motion. The plate was immersed into a Re = 10,000 flow (based on chord length) to simulate forward, flapping flight. Measurements were made using a 3D-3C plenoptic PIV system to allow for the study of vortex topology in the instantaneous flow, in addition to phase-averaged results. The prominent focus is leading-edge vortex (LEV) stability and the lifespan of shed LEVs. The parameter space involves multiple values of advance coefficient J and reduced frequency k for roll and pitch, respectively. This space aims to determine the influence of each parameter on LEVs, which has been identified as an important factor for the lift enhancement seen in flapping wing flight. A variety of results are to be presented characterizing the variations in vortex topology across this parameter space. This work is supported by the Air Force Office of Scientific Research (Grant Number FA9550-16-1-0107, Dr. Douglas Smith, program manager).

Aerodynamic performance and particle image velocimetery of piezo actuated biomimetic manduca sexta engineered wings towards the design and application of a flapping wing flight vehicle

NASA Astrophysics Data System (ADS)

DeLuca, Anthony M.

Considerable research and investigation has been conducted on the aerodynamic performance, and the predominate flow physics of the Manduca Sexta size of biomimetically designed and fabricated wings as part of the AFIT FWMAV design project. Despite a burgeoning interest and research into the diverse field of flapping wing flight and biomimicry, the aerodynamics of flapping wing flight remains a nebulous field of science with considerable variance into the theoretical abstractions surrounding aerodynamic mechanisms responsible for aerial performance. Traditional FWMAV flight models assume a form of a quasi-steady approximation of wing aerodynamics based on an infinite wing blade element model (BEM). An accurate estimation of the lift, drag, and side force coefficients is a critical component of autonomous stability and control models. This research focused on two separate experimental avenues into the aerodynamics of AFIT's engineered hawkmoth wings|forces and flow visualization. 1. Six degree of freedom force balance testing, and high speed video analysis was conducted on 30°, 45°, and 60° angle stop wings. A novel, non-intrusive optical tracking algorithm was developed utilizing a combination of a Gaussian Mixture Model (GMM) and ComputerVision (OpenCV) tools to track the wing in motion from multiple cameras. A complete mapping of the wing's kinematic angles as a function of driving amplitude was performed. The stroke angle, elevation angle, and angle of attack were tabulated for all three wings at driving amplitudes ranging from A=0.3 to A=0.6. The wing kinematics together with the force balance data was used to develop several aerodynamic force coefficient models. A combined translational and rotational aerodynamic model predicted lift forces within 10%, and vertical forces within 6%. The total power consumption was calculated for each of the three wings, and a Figure of Merit was calculated for each wing as a general expression of the overall efficiency of the wing. Th 60° angle stop wing achieved the largest total stroke angle and generated the most lift for the lowest power consumption of the wings tested. 2. Phase averaged stereo Particle Image Velocimetry (PIV) data was collected at eight phases through the flap cycle on the 30°, 45°, and 60° angle stop wings. Wings were mounted transverse and parallel to the interrogating laser sheet, and planar velocity intersections at the wing mid-span, one chord below the wing, were compared to one another to verify data fidelity. A Rankine-Froude actuator disk model was adapted to calculate the approximate vertical thrust generated from the total momentum flux through the flapping semi-disk using the velocity field measurements. Three component stereo u, v, and w-velocity contour measurements confirmed the presence of extensive vortical structures in the vicinity of the wing. The leading edge vortex was successfully tracked through the stroke cycle appearing at approximately 25% span, increasing in circulatory strength and translational velocity down the span toward the tip, and dissipating just after 75% span. Thrust calculations showed the vertically mounted wing more accurately represented the vertical forces when compared to its corresponding force balance measurement than the horizontally mounted wing. The mid-span showed the highest vertical velocity profile below the wing; and hence, was the location responsible for the majority of lift production along the span.

An improved panel method for the solution of three-dimensional leading edge vortex flows Volume 2: User's guide and programmer's document

NASA Technical Reports Server (NTRS)

Tinoco, E. N.; Lu, P.; Johnson, F. T.

1980-01-01

A computer program developed for solving the subsonic, three dimensional flow over wing-body configurations with leading edge vortex separation is presented. Instructions are given for the proper set up and input of a problem into the computer code. Program input formats and output are described, as well as the overlay structure of the program. The program is written in FORTRAN.

The role of the leading edge vortex in lift augmentation of steadily revolving wings: a change in perspective.

PubMed

Nabawy, Mostafa R A; Crowther, William J

2017-07-01

The presence of a stable leading edge vortex (LEV) on steadily revolving wings increases the maximum lift coefficient that can be generated from the wing and its role is important to understanding natural flyers and flapping wing vehicles. In this paper, the role of LEV in lift augmentation is discussed under two hypotheses referred to as 'additional lift' and 'absence of stall'. The 'additional lift' hypothesis represents the traditional view. It presumes that an additional suction/circulation from the LEV increases the lift above that of a potential flow solution. This behaviour may be represented through either the 'Polhamus leading edge suction' model or the so-called 'trapped vortex' model. The 'absence of stall' hypothesis is a more recent contender that presumes that the LEV prevents stall at high angles of attack where flow separation would normally occur. This behaviour is represented through the so-called 'normal force' model. We show that all three models can be written in the form of the same potential flow kernel with modifiers to account for the presence of a LEV. The modelling is built on previous work on quasi-steady models for hovering wings such that model parameters are determined from first principles, which allows a fair comparison between the models themselves, and the models and experimental data. We show that the two models which directly include the LEV as a lift generating component are built on a physical picture that does not represent the available experimental data. The simpler 'normal force' model, which does not explicitly model the LEV, performs best against data in the literature. We conclude that under steady conditions the LEV as an 'absence of stall' model/mechanism is the most satisfying explanation for observed aerodynamic behaviour. © 2017 The Author(s).

The Hydrodynamics of Plesiosaurs

NASA Astrophysics Data System (ADS)

Muscutt, Luke; Ganapathisubramani, Bharathram; Dyke, Gareth; Weymouth, Gabriel

2015-11-01

Plesiosaurs are extinct marine reptiles that existed at the same time as the dinosaurs, and are the only known animals to swim by actively flapping their four wing-like flippers. This can be viewed as a tandem flapping wing problem, where the hind wing is operating in the wake of the fore wing. Experiments using full-scale robotic plesiosaur flippers in a large flume tank have been used to investigate the kinematics and interaction of the flippers. The flippers are actuated in heave and pitch, and a combination of force measurements and flow visualization are used to analyze the characteristics of the vortex interaction between the flippers. Previous two-dimensional numerical simulations have shown that certain kinematics give an increase in thrust of the hind flipper of up to 50%. The current experiments determine if such thrust augmentation is present for a three-dimensional flowfield and determine the kinematics that give the highest possible thrust. This will help to answer paleo-biological questions about the function and evolution of the plesiosaur flippers, along with helping to determine if tandem flapping wings could be a viable propulsion system for autonomous underwater vehicles. Support from EPSRC and Ginko Investments Ltd.

Rotor/Wing Interactions in Hover

NASA Technical Reports Server (NTRS)

Young, Larry A.; Derby, Michael R.

2002-01-01

Hover predictions of tiltrotor aircraft are hampered by the lack of accurate and computationally efficient models for rotor/wing interactional aerodynamics. This paper summarizes the development of an approximate, potential flow solution for the rotor-on-rotor and wing-on-rotor interactions. This analysis is based on actuator disk and vortex theory and the method of images. The analysis is applicable for out-of-ground-effect predictions. The analysis is particularly suited for aircraft preliminary design studies. Flow field predictions from this simple analytical model are validated against experimental data from previous studies. The paper concludes with an analytical assessment of the influence of rotor-on-rotor and wing-on-rotor interactions. This assessment examines the effect of rotor-to-wing offset distance, wing sweep, wing span, and flaperon incidence angle on tiltrotor inflow and performance.

Vortex information display system program description manual. [data acquisition from laser Doppler velocimeters and real time operation

NASA Technical Reports Server (NTRS)

Conway, R.; Matuck, G. N.; Roe, J. M.; Taylor, J.; Turner, A.

1975-01-01

A vortex information display system is described which provides flexible control through system-user interaction for collecting wing-tip-trailing vortex data, processing this data in real time, displaying the processed data, storing raw data on magnetic tape, and post processing raw data. The data is received from two asynchronous laser Doppler velocimeters (LDV's) and includes position, velocity, and intensity information. The raw data is written onto magnetic tape for permanent storage and is also processed in real time to locate vortices and plot their positions as a function of time. The interactive capability enables the user to make real time adjustments in processing data and provides a better definition of vortex behavior. Displaying the vortex information in real time produces a feedback capability to the LDV system operator allowing adjustments to be made in the collection of raw data. Both raw data and processing can be continually upgraded during flyby testing to improve vortex behavior studies. The post-analysis capability permits the analyst to perform in-depth studies of test data and to modify vortex behavior models to improve transport predictions.

Recent developments in rotary-wing aerodynamic theory

NASA Technical Reports Server (NTRS)

Johnson, W.

1986-01-01

Current progress in the computational analysis of rotary-wing flowfields is surveyed, and some typical results are presented in graphs. Topics examined include potential theory, rotating coordinate systems, lifting-surface theory (moving singularity, fixed wing, and rotary wing), panel methods (surface singularity representations, integral equations, and compressible flows), transonic theory (the small-disturbance equation), wake analysis (hovering rotor-wake models and transonic blade-vortex interaction), limitations on computational aerodynamics, and viscous-flow methods (dynamic-stall theories and lifting-line theory). It is suggested that the present algorithms and advanced computers make it possible to begin working toward the ultimate goal of turbulent Navier-Stokes calculations for an entire rotorcraft.

Study of Near-Stall Flow Behavior in a Modern Transonic Fan with Composite Sweep

NASA Technical Reports Server (NTRS)

Hah, Chunill; Shin, Hyoun-Woo

2011-01-01

Detailed flow behavior in a modern transonic fan with a composite sweep is investigated in this paper. Both unsteady Reynolds-averaged Navier-Stokes (URANS) and Large Eddy Simulation (LES) methods are applied to investigate the flow field over a wide operating range. The calculated flow fields are compared with the data from an array of high-frequency response pressure transducers embedded in the fan casing. The current study shows that a relatively fine computational grid is required to resolve the flow field adequately and to calculate the pressure rise across the fan correctly. The calculated flow field shows detailed flow structure near the fan rotor tip region. Due to the introduction of composite sweep toward the rotor tip, the flow structure at the rotor tip is much more stable compared to that of the conventional blade design. The passage shock stays very close to the leading edge at the rotor tip even at the throttle limit. On the other hand, the passage shock becomes stronger and detaches earlier from the blade passage at the radius where the blade sweep is in the opposite direction. The interaction between the tip clearance vortex and the passage shock becomes intense as the fan operates toward the stall limit, and tip clearance vortex breakdown occurs at near-stall operation. URANS calculates the time-averaged flow field fairly well. Details of measured RMS static pressure are not calculated with sufficient accuracy with URANS. On the other hand, LES calculates details of the measured unsteady flow features in the current transonic fan with composite sweep fairly well and reveals the flow mechanism behind the measured unsteady flow field.

In-flight flow visualization with pressure measurements at low speeds on the NASA F-18 high alpha research vehicle

NASA Technical Reports Server (NTRS)

Delfrate, John H.; Fisher, David F.; Zuniga, Fanny A.

1990-01-01

In-flight results from surface and off-surface flow visualizations and from extensive pressure distributions document the vortical flow on the leading edge extensions (LEX) and forebody of the NASA F-18 high alpha research vehicle for low speeds and angles of attack up to 50 degs. Surface flow visualization data, obtained using the emitted fluid technique, were used to define separation lines and laminar separation bubbles. Off-surface flow visualization data, obtained by smoke injection, were used to document both the path of the vortex cores and the location of vortex core breakdown. The location of vortex core breakdown correlated well with the loss of suction pressure on the LEX and with the flow visualization results from ground facilities. Surface flow separation lines on the LEX and forebody corresponded well with the end of pressure recovery under the vortical flows. Correlation of the pressures with wind tunnel results show fair to good correlation.

Transonic Aerodynamic Characteristics of a Wing-Body Combination having a 52.5 deg Sweptback Wing of Aspect Ratio 3 with Conical Camber and Designed for a Mach Number of the Square Root of 2

NASA Technical Reports Server (NTRS)

Igoe, William B.; Re, Richard J.; Cassetti, Marlowe

1961-01-01

An investigation has been made of the effects of conical wing camber and supersonic body indentation on the aerodynamic characteristics of a wing-body configuration at transonic speeds. Wing aspect ratio was 3.0, taper ratio was 0.1, and quarter-chord line sweepback was 52.5 deg with airfoil sections of 0.03 thickness ratio. The tests were conducted in the Langley 16-foot transonic tunnel at various Mach numbers from 0.80 to 1.05 at angles of attack from -4 deg to 14 deg. The cambered-wing configuration achieved higher lift-drag ratios than a similar plane-wing configuration. The camber also reduced the effects of wing-tip flow separation on the aerodynamic characteristics. In general, no stability or trim changes below wing-tip flow separation resulted from the use of camber. The use of supersonic body indentation improved the lift-drag ratios at Mach numbers from 0.96 to 1.05.

Prediction of aerodynamic noise in a ring fan based on wake characteristics

NASA Astrophysics Data System (ADS)

Sasaki, Soichi; Fukuda, Masaharu; Tsujino, Masao; Tsubota, Haruhiro

2011-06-01

A ring fan is a propeller fan that applies an axial-flow impeller with a ring-shaped shroud on the blade tip side. In this study, the entire flow field of the ring fan is simulated using computational fluid dynamics (CFD); the accuracy of the CFD is verified through a comparison with the aerodynamic characteristics of a propeller fan of current model. Moreover, the aerodynamic noise generated by the fan is predicted on the basis of the wake characteristics. The aerodynamic characteristic of the ring fan based on CFD can represent qualitatively the variation in the measured value. The main flow domain of the ring fan is formed at the tip side of the blade because blade tip vortex is not formed at that location. Therefore, the relative velocity of the ring fan is increased by the circumferential velocity. The sound pressure levels of the ring fan within the frequency band of less than 200 Hz are larger than that of the propeller fan. In the analysis of the wake characteristics, it revealed that Karman vortex shedding occurred in the main flow domain in the frequency domain lower than 200 Hz; the aerodynamic noise of the ring fan in the vortex shedding frequency enlarges due to increase in the relative velocity and the velocity fluctuation.

Development of Hairpin Vortices in Turbulent Spots and End-Wall Transition

NASA Technical Reports Server (NTRS)

Smith, Charles R.

2007-01-01

The end-stage phase of boundary layer transition is characterized by the development of hairpin-like vortices which evolve rapidly into patches of turbulent behavior. In general, the characteristics of the evolution form this hairpin stage to the turbulent stage is poorly understood, which has prompted the present experimental examination of hairpin vortex development and growth processes. Two topics of particular relevance to the workshop focus will be covered: 1) the growth of turbulent spots through the generatio and amalgamation of hairpin-like vortices, and 2) the development of hairpin vortices during transition in an end-wall junction flow. Brief summaries of these studies are described below. Using controlled generation of hairpin vortices by surface injection in a critical laminar boundary layer, detailed flow visualization studies have been done of the phases of growth of single hairpin vortices, from the initial hairgin generation, through the systematic generation of secondary hairpin-like flow structures, culminating in the evolution to a turbulent spot. The key to the growth process is strong vortex-surface interactions, which give rise to strong eruptive events adjacent to the surface, which results in the generation of subsequent hairpin vortex structures due to inviscid-viscuous interactions between the eruptive events and the free steam fluid. The general process of vortex-surface fluid interaction, coupled with subsequent interactions and amalgamation of the generated multiple hairpin-type vortices, is demonstrated as a physical mechanism for the growth and development of turbulent spots. When a boundary layer flow along a surface encounters a bluff body obstruction extending from the surface (such as cylinder or wing), the strong adverse pressure gradients generated by these types of flows result in the concentration of the impinging vorticity into a system of discrete vortices near the end-wall juncture of the obstruction, with the extensions of the vortices engirdling the obstruction to form "necklace" or "horseshoe" vortices. Recent hydrogen bubble and particle image visualization have shown that as Reynolds number is increased for a laminar approach flow, the flow will become critical, and a destabilization of the necklace vortices results in the development of an azimuthal waviness, or "kinks", in the vortices. These vortex kinks are accentuated by Biot-Savart effects, causing portions of a distorted necklace vortex to make a rapid approach to the surface, precipitating processes of localized, three-dimensional surface interactions. These interactions result in the rapid generation, focussing, and ejection of thin tongues of surface fluid, which rapidly roll-over and appear as hairpin vortices in the junction region. Subsequent amalgamation of these hairpin vortices with the necklace vortices produces a complex transitional-type flow. A presentation of key results from both these studies will be done, emphasizing both the ubiquity of such hairpin-type flow structures in manifold transitional-type flows, and the importance of vortex-surface interactions n the development of hairpin vortices.

Flow past a Flat Plate with a Vortex/sink Combination

NASA Technical Reports Server (NTRS)

Mourtos, N. J.

1984-01-01

An attempt was made to model the so called leading edge vortex which forms over the leading edge of delta wings at high angles of attack. A simplified model was considered, namely that of a two-dimensional, inviscid, incompressible steady flow around a flat plate at an angle of attack with a stationary vortex detached on top, as well as a sink to simulate the strong spanwise flow. The results appear to agree qualitatively with experiments. A comparison was also made between the lift and the drag of this model and the corresponding results for two classical solutions: (1) that of totally attached flow over the plate with the Kutta condition satisfied at the trailing edge only: and (2) the Helmholtz solution of totally separated flow over the plate.

Flow past a flat plat with a vortex/sink combination

NASA Technical Reports Server (NTRS)

Mourtos, N. J.

1985-01-01

An attempt was made to model the so called leading edge vortex which forms over the leading edge of delta wings at high angles of attack. A simplified model was considered, namely that of a two-dimensional, inviscid, incompressible steady flow around a flat plate at an angle of attack with a stationary vortex detached on top, as well as a sink to simulate the strong spanwise flow. The results appear to agree qualitatively with experiments. A comparison was also made between the lift and the drag of this model and the corresponding results for two classical solutions: (1) that of totally attached flow over the plate with the Kutta condition satisfied at the trailing edge only; and (2) the Helmholtz solution of totally separated flow over the plate.

Three-dimensional boundary layer stability and transition

NASA Technical Reports Server (NTRS)

Malik, M. R.; Li, F.

1992-01-01

Nonparallel and nonlinear stability of a three-dimensional boundary layer, subject to crossflow instability, is investigated using parabolized stability equations (PSEs). Both traveling and stationary disturbances are considered and nonparallel effect on crossflow instability is found to be destabilizing. Our linear PSE results for stationary disturbances agree well with the results from direct solution of Navier-Stokes equations obtained by Spalart (1989). Nonlinear calculations have been carried out for stationary vortices and the computed wall vorticity pattern results in streamwise streaks which resemble remarkably well with the surface oil-flow visualizations in swept-wing experiments. Other features of the stationary vortex development (half-mushroom structure, inflected velocity profiles, vortex doubling, etc.) are also captured in our nonlinear calculations. Nonlinear interaction of the stationary amplitude of the stationary vortex is large as compared to the traveling mode, and the stationary vortex dominates most of the downstream development. When the two modes have the same initial amplitude, the traveling mode dominates the downstream development owing to its higher growth rate, and there is a tendency for the stationary mode to be suppressed. The effect of nonlinear wave development on the skin-friction coefficient is also computed.

Aerodynamics of a finite wing with simulated ice

NASA Technical Reports Server (NTRS)

Bragg, M. B.; Khodadoust, A.; Kerho, M.

1992-01-01

The effect of a simulated glaze ice accretion on the aerodynamic performance of a three-dimensional wing is studied experimentally. Results are reviewed from earlier two-dimensional tests which show the character of the large leading-edge separation bubbles caused by the simulated ice accretion. The 2-D bubbles are found to closely resemble well known airfoil laminar separation bubbles. For the 3-D experiments a semispan wing of effective aspect ratio five was mounted from the sidewall of the UIUC subsonic wind tunnel. The model uses a NACA 0012 airfoil section on a rectangular planform with interchangeable tip and root sections to allow for 0- and 30-degree sweep. A three-component sidewall balance was used to measure lift, drag and pitching moment on the clean and iced model. Fluorescent oil flow visualization has been performed on the iced model and reveals extensive spanwise and vortical flow in the separation bubble aft of the upper surface horn. Sidewall interaction and spanwise nonuniformity are also seen on the unswept model. Comparisons to the computed flow fields are shown. Results are also shown for roughness effects on the straight wing. Sand grain roughness on the ice shape is seen to have a different effect than isolated 3-D roughness elements.

On the combination of kinematics with flow visualization to compute total circulation - Application to vortex rings in a tube

NASA Technical Reports Server (NTRS)

Brasseur, J. G.; Chang, I.-D.

1980-01-01

To date the computation of the total circulation, or strength of a vortex has required detailed measurements of the velocity field within the vortex. In this paper a method is described in which the kinematics of the vortical flow field is exploited to calculate the strength of a vortex from relatively simple flow visualization measurements. There are several advantages in the technique, the most important being the newly acquired ability to calculate the transient changes in strength of a single vortex as it evolves. The method is applied to the study of vortex rings, although the development can be carried over directly to study vortex pairs, and it is expected that it can be generalized to other flows which contain regions of concentrated vorticity. The accuracy of the method as applied to vortex rings, assessed in part by comparing with the laser Doppler velocimeter (LDV) measurements of Sullivan et al., is shown to be excellent.

A coupled CFD and wake model simulation of helicopter rotor in hover

NASA Astrophysics Data System (ADS)

Zhao, Qinghe; Li, Xiaodong

2018-03-01

The helicopter rotor wake plays a dominant role since it affects the flow field structure. It is very difficult to predict accurately of the flow-field. The numerical dissipation is so excessive that it eliminates the vortex structure. A hybrid method of CFD and prescribed wake model was constructed by applying the prescribed wake model as much as possible. The wake vortices were described as a single blade tip vortex in this study. The coupling model is used to simulate the flow field. Both non-lifting and lifting cases have been calculated with subcritical and supercritical tip Mach numbers. Surface pressure distributions are presented and compared with experimental data. The calculated results agree well with the experimental data.

Vortical Flow Prediction Using an Adaptive Unstructured Grid Method

NASA Technical Reports Server (NTRS)

Pirzadeh, Shahyar Z.

2001-01-01

A computational fluid dynamics (CFD) method has been employed to compute vortical flows around slender wing/body configurations. The emphasis of the paper is on the effectiveness of an adaptive grid procedure in "capturing" concentrated vortices generated at sharp edges or flow separation lines of lifting surfaces flying at high angles of attack. The method is based on a tetrahedral unstructured grid technology developed at the NASA Langley Research Center. Two steady-state, subsonic, inviscid and Navier-Stokes flow test cases are presented to demonstrate the applicability of the method for solving practical vortical flow problems. The first test case concerns vortex flow over a simple 65deg delta wing with different values of leading-edge bluntness, and the second case is that of a more complex fighter configuration. The superiority of the adapted solutions in capturing the vortex flow structure over the conventional unadapted results is demonstrated by comparisons with the windtunnel experimental data. The study shows that numerical prediction of vortical flows is highly sensitive to the local grid resolution and that the implementation of grid adaptation is essential when applying CFD methods to such complicated flow problems.

Flight, Wind-Tunnel, and Computational Fluid Dynamics Comparison for Cranked Arrow Wing (F-16XL-1) at Subsonic and Transonic Speeds

NASA Technical Reports Server (NTRS)

Lamar, John E.; Obara, Clifford J.; Fisher, Bruce D.; Fisher, David F.

2001-01-01

Geometrical, flight, computational fluid dynamics (CFD), and wind-tunnel studies for the F-16XL-1 airplane are summarized over a wide range of test conditions. Details are as follows: (1) For geometry, the upper surface of the airplane and the numerical surface description compare reasonably well. (2) For flight, CFD, and wind-tunnel surface pressures, the comparisons are generally good at low angles of attack at both subsonic and transonic speeds, however, local differences are present. In addition, the shock location at transonic speeds from wind-tunnel pressure contours is near the aileron hinge line and generally is in correlative agreement with flight results. (3) For boundary layers, flight profiles were predicted reasonably well for attached flow and underneath the primary vortex but not for the secondary vortex. Flight data indicate the presence of an interaction of the secondary vortex system and the boundary layer and the boundary-layer measurements show the secondary vortex located more outboard than predicted. (4) Predicted and measured skin friction distributions showed qualitative agreement for a two vortex system. (5) Web-based data-extraction and computational-graphical tools have proven useful in expediting the preceding comparisons. (6) Data fusion has produced insightful results for a variety of visualization-based data sets.

An application of the suction analog for the analysis of asymmetric flow situations

NASA Technical Reports Server (NTRS)

Luckring, J. M.

1976-01-01

A recent extension of the suction analogy for estimation of vortex loads on asymmetric configurations is reviewed. This extension includes asymmetric augmented vortex lift and the forward sweep effect on side edge suction. Application of this extension to a series of skewed wings has resulted in an improved estimating capability for a wide range of asymmetric flow situations. Hence, the suction analogy concept now has more general applicability for subsonic lifting surface analysis.

An Experimental Study of a Turbulent Wing-Body Junction and Wake Flow

DTIC Science & Technology

1991-09-01

10 ................... 323 Figure 5.37. Comparison of Contours of (a(W)/aY) 5fU ,’I, Plane 11 ......... 324 Figure 5.38. Comparison of Contours of h...the mean velocity in the boundary layer and the vortex position. The vortex depresses the contours of Urc by drawing in high momentum fluid from the

Computer prediction of three-dimensional potential flow fields in which aircraft propellers operate. M.S. Thesis

NASA Technical Reports Server (NTRS)

Jumper, S. J.

1982-01-01

A computer program was developed to calculate the three dimensional, steady, incompressible, inviscid, irrotational flow field at the propeller plane (propeller removed) located upstream of an arbitrary airframe geometry. The program uses a horseshoe vortex of known strength to model the wing. All other airframe surfaces are modeled by a network source panels of unknown strength which is exposed to a uniform free stream and the wing-induced velocity field. By satisfying boundary conditions on each panel (the Neumann problem), relaxed boundary conditions being used on certain panels to simulate inlet inflow, the source strengths are determined. From the known source and wing vortex strengths, the resulting velocity fields on the airframe surface and at the propeller plane are obtained. All program equations are derived in detail, and a brief description of the program structure is presented. A user's manual which fully documents the program is cited. Computer predictions of the flow on the surface of a sphere and at a propeller plane upstream of the sphere are compared with the exact mathematical solutions. Agreement is good, and correct program operation is verified.

Aerodynamics of high frequency flapping wings

NASA Astrophysics Data System (ADS)

Hu, Zheng; Roll, Jesse; Cheng, Bo; Deng, Xinyan

2010-11-01

We investigated the aerodynamic performance of high frequency flapping wings using a 2.5 gram robotic insect mechanism developed in our lab. The mechanism flaps up to 65Hz with a pair of man-made wing mounted with 10cm wingtip-to-wingtip span. The mean aerodynamic lift force was measured by a lever platform, and the flow velocity and vorticity were measured using a stereo DPIV system in the frontal, parasagittal, and horizontal planes. Both near field (leading edge vortex) and far field flow (induced flow) were measured with instantaneous and phase-averaged results. Systematic experiments were performed on the man-made wings, cicada and hawk moth wings due to their similar size, frequency and Reynolds number. For insect wings, we used both dry and freshly-cut wings. The aerodynamic force increase with flapping frequency and the man-made wing generates more than 4 grams of lift at 35Hz with 3 volt input. Here we present the experimental results and the major differences in their aerodynamic performances.

Exploring bird aerodynamics using radio-controlled models.

PubMed

Hoey, Robert G

2010-12-01

A series of radio-controlled glider models was constructed by duplicating the aerodynamic shape of soaring birds (raven, turkey vulture, seagull and pelican). Controlled tests were conducted to determine the level of longitudinal and lateral-directional static stability, and to identify the characteristics that allowed flight without a vertical tail. The use of tail-tilt for controlling small bank-angle changes, as observed in soaring birds, was verified. Subsequent tests, using wing-tip ailerons, inferred that birds use a three-dimensional flow pattern around the wing tip (wing tip vortices) to control adverse yaw and to create a small amount of forward thrust in gliding flight.

F-8 supercritical wing flight pressure, Boundary layer, and wake measurements and comparisons with wind tunnel data

NASA Technical Reports Server (NTRS)

Montoya, L. C.; Banner, R. D.

1977-01-01

Data for speeds from Mach 0.50 to Mach 0.99 are presented for configurations with and without fuselage area-rule additions, with and without leading-edge vortex generators, and with and without boundary-layer trips on the wing. The wing pressure coefficients are tabulated. Comparisons between the airplane and model data show that higher second velocity peaks occurred on the airplane wing than on the model wing. The differences were attributed to wind tunnel wall interference effects that caused too much rear camber to be designed into the wing. Optimum flow conditions on the outboard wing section occurred at Mach 0.98 at an angle of attack near 4 deg. The measured differences in section drag with and without boundary-layer trips on the wing suggested that a region of laminar flow existed on the outboard wing without trips.

Observations of the Effect of Wing Appendages and Flaps on the Spread of Separation of Flow over the Wing

NASA Technical Reports Server (NTRS)

Hartwig, G

1941-01-01

The spread of the separation of flow on three tapered wings insymmetrical and unsymmetrical flow was observed with silk tufts. By equal thickness and chord distribution the wings manifested a different form of lifting line. The principal result of the study was that the wings alone first disclosed complete breakdown of the flow at the tips, even the one with twist, but that after adding fuselage and engine nacelles, the twisted wing broke down completely first in the wing center. The observed boundary layer motions transverse to the main flow direction were briefly explored as to their possible influence on the spread of the separation. On top of that certain disclosures were afforded in which the transverse motions observed in the boundary layer became perceptible even above the boundary layer.

Numerical Study of Steady and Unsteady Canard-Wing-Body Aerodynamics

NASA Technical Reports Server (NTRS)

Eugene, L. Tu

1996-01-01

The use of canards in advanced aircraft for control and improved aerodynamic performance is a topic of continued interest and research. In addition to providing maneuver control and trim, the influence of canards on wing aerodynamics can often result in increased maximum lift and decreased trim drag. In many canard-configured aircraft, the main benefits of canards are realized during maneuver or other dynamic conditions. Therefore, the detailed study and understanding of canards requires the accurate prediction of the non-linear unsteady aerodynamics of such configurations. For close-coupled canards, the unsteady aerodynamic performance associated with the canard-wing interaction is of particular interest. The presence of a canard in close proximity to the wing results in a highly coupled canard-wing aerodynamic flowfield which can include downwash/upwash effects, vortex-vortex interactions and vortex-surface interactions. For unsteady conditions, these complexities of the canard-wing flowfield are further increased. The development and integration of advanced computational technologies provide for the time-accurate Navier-Stokes simulations of the steady and unsteady canard-wing-body flox,fields. Simulation, are performed for non-linear flight regimes at transonic Mach numbers and for a wide range of angles of attack. For the static configurations, the effects of canard positioning and fixed deflection angles on aerodynamic performance and canard-wing vortex interaction are considered. For non-static configurations, the analyses of the canard-wing body flowfield includes the unsteady aerodynamics associated with pitch-up ramp and pitch oscillatory motions of the entire geometry. The unsteady flowfield associated with moving canards which are typically used as primary control surfaces are considered as well. The steady and unsteady effects of the canard on surface pressure integrated forces and moments, and canard-wing vortex interaction are presented in detail including the effects of the canard on the static and dynamic stability characteristics. The current study provides an understanding of the steady and unsteady canard-wing-body flowfield. Emphasis is placed on the effects of the canard on aerodynamic performance as well as the detailed flow physics of the canard-wing flowfield interactions. The computational tools developed to accurately predict the time-accurate flowfield of moving canards provides for the capability of coupled fluids-controls simulations desired in the detailed design and analysis of advanced aircraft.

Measurements of Tip Vortices from a Full-Scale UH-60A Rotor by Retro- Reflective Background Oriented Schlieren and Stereo Photogrammetry

NASA Technical Reports Server (NTRS)

Schairer, Edward; Kushner, Laura K.; Heineck, James T.

2013-01-01

Positions of vortices shed by a full-scale UH-60A rotor in forward flight were measured during a test in the National Full- Scale Aerodynamics Complex at NASA Ames Research Center. Vortices in a region near the tip of the advancing blade were visualized from two directions by Retro-Reflective Background-Oriented Schlieren (RBOS). Correspondence of points on the vortex in the RBOS images from both cameras was established using epipolar geometry. The object-space coordinates of the vortices were then calculated from the image-plane coordinates using stereo photogrammetry. One vortex from the tip of the blade that had most recently passed was visible in most of the data. The visibility of the vortices was greatest at high thrust and low advance ratios. At these favorable conditions, vortices from the most recent passages of all four blades were detected. The vortex positions were in good agreement with PIV data for a case where PIV measurements were also made. RBOS and photogrammetry provided measurements of the angle at which each vortex passed through the PIV plane.

Unsteady, Transonic Flow Around Delta Wings Undergoing Coupled and Natural Modes Response: A Multidisciplinary Problem

NASA Technical Reports Server (NTRS)

Menzies, Margaret Anne

1996-01-01

The unsteady, three-dimensional Navier-Stokes equations coupled with the Euler equations of rigid-body dynamics are sequentially solved to simulate and analyze the aerodynamic response of a high angle of attack delta wing undergoing oscillatory motion. The governing equations of fluid flow and dynamics of the multidisciplinary problem are solved using a time-accurate solution of the laminar, unsteady, compressible, full Navier- Stokes equations with the implicit, upwind, Roe flux-difference splitting, finite-volume scheme and a four-stage Runge-Kutta scheme, respectively. The primary model under consideration consists of a 65 deg swept, sharp-edged, cropped delta wing of zero thickness at 20 deg angle of attack. In a freestream of Mach 0.85 and Reynolds number of 3.23 x 10(exp 6), the flow over the upper surface of the wing develops a complex shock system which interacts with the leading-edge primary vortices producing vortex breakdown. The effect of the oscillatory motion of the wing on the vortex breakdown and overall aerodynamic response is detailed to provide insight to the complicated physics associated with unsteady flows and the phenomenon of wing rock. Forced sinusoidal single and coupled mode rolling and pitching motion is presented for the wing in a transonic freestream. The Reynolds number, frequency of oscillation, and the phase angle are varied. Comparison between the single and coupled mode forced rolling and pitching oscillation cases illustrate the effects of coupling the motion. This investigation shows that even when coupled, forced rolling oscillation at a reduced frequency of 2(pi) eliminates the vortex breakdown which results in an increase in lift. The coupling effect for in phase forced oscillations show that the lift coefficient of the pitching-alone case and the rolling-moment coefficient of the rolling-alone case dominate the resulting response. However, with a phase lead in the pitching motion, the coupled motion results in a non-periodic response of the rolling moment. The second class of problems involve releasing the wing in roll to respond to the flowfield. Two models of sharp-edged delta wings, the previous 65 deg swept model and an 80 deg swept, sharp-edged delta wing, are used to observe the aerodynamic response of a wing free to roll in a transonic and subsonic freestream, respectively. These cases demonstrate damped oscillations, self-sustained limit cycle oscillations, and divergent rolling oscillations. Ultimately, an active control model using a mass injection system was applied on the surface of the wing to suppress the self-sustained limit cycle oscillation known as wing rock. Comparisons with experimental investigations complete this study, validating the analysis and illustrating the complex details afforded by computational investigations.

An experimental study of tip shape effects on the flutter of aft-swept, flat-plate wings

NASA Technical Reports Server (NTRS)

Dansberry, Bryan E.; Rivera, Jose A., Jr.; Farmer, Moses G.

1990-01-01

The effects of tip chord orientation on wing flutter are investigated experimentally using six cantilever-mounted, flat-plate wing models. Experimentally determined flutter characteristics of the six models are presented covering both the subsonic and transonic Mach number ranges. While all models have a 60 degree leading edge sweep, a 40.97 degree trailing edge sweep, and a root chord of 34.75 inches, they are subdivided into two series characterized by a higher aspect ratio and a lower aspect ratio. Each series is made up of three models with tip chord orientations which are parallel to the free-stream flow, perpendicular to the model mid-chord line, and perpendicular to the free-stream flow. Although planform characteristics within each series of models are held constant, structural characteristics such as mode shapes and natural frequencies are allowed to vary.

Schlieren photography on freely flying hawkmoth.

PubMed

Liu, Yun; Roll, Jesse; Van Kooten, Stephen; Deng, Xinyan

2018-05-01

The aerodynamic force on flying insects results from the vortical flow structures that vary both spatially and temporally throughout flight. Due to these complexities and the inherent difficulties in studying flying insects in a natural setting, a complete picture of the vortical flow has been difficult to obtain experimentally. In this paper, Schlieren , a widely used technique for highspeed flow visualization, was adapted to capture the vortex structures around freely flying hawkmoth ( Manduca ). Flow features such as leading-edge vortex, trailing-edge vortex, as well as the full vortex system in the wake were visualized directly. Quantification of the flow from the Schlieren images was then obtained by applying a physics-based optical flow method, extending the potential applications of the method to further studies of flying insects. © 2018 The Author(s).

An experimental study of the unsteady vortex structures in the wake of a root-fixed flapping wing

NASA Astrophysics Data System (ADS)

Hu, Hui; Clemons, Lucas; Igarashi, Hirofumi

2011-08-01

An experimental study was conducted to characterize the evolution of the unsteady vortex structures in the wake of a root-fixed flapping wing with the wing size, stroke amplitude, and flapping frequency within the range of insect characteristics for the development of novel insect-sized nano-air-vehicles (NAVs). The experiments were conducted in a low-speed wing tunnel with a miniaturized piezoelectric wing (i.e., chord length, C = 12.7 mm) flapping at a frequency of 60 Hz (i.e., f = 60 Hz). The non-dimensional parameters of the flapping wing are chord Reynolds number of Re = 1,200, reduced frequency of k = 3.5, and non-dimensional flapping amplitude at wingtip h = A/C = 1.35. The corresponding Strouhal number (Str) is 0.33 , which is well within the optimal range of 0.2 < Str < 0.4 used by flying insects and birds and swimming fishes for locomotion. A digital particle image velocimetry (PIV) system was used to achieve phased-locked and time-averaged flow field measurements to quantify the transient behavior of the wake vortices in relation to the positions of the flapping wing during the upstroke and down stroke flapping cycles. The characteristics of the wake vortex structures in the chordwise cross planes at different wingspan locations were compared quantitatively to elucidate underlying physics for a better understanding of the unsteady aerodynamics of flapping flight and to explore/optimize design paradigms for the development of novel insect-sized, flapping-wing-based NAVs.

A two dimensional study of rotor/airfoil interaction in hover

NASA Technical Reports Server (NTRS)

Lee, Chyang S.

1988-01-01

A two dimensional model for the chordwise flow near the wing tip of the tilt rotor in hover is presented. The airfoil is represented by vortex panels and the rotor is modeled by doublet panels. The rotor slipstream and the airfoil wake are simulated by free point vortices. Calculations on a 20 percent thick elliptical airfoil under a uniform rotor inflow are performed. Variations on rotor size, spacing between the rotor and the airfoil, ground effect, and the influence upper surface blowing in download reduction are analyzed. Rotor size has only a minor influence on download when it is small. Increase of the rotor/airfoil spacing causes a gradual decrease on download. Proximity to the ground effectively reduces the download and makes the wake unsteady. The surface blowing changes the whole flow structure and significantly reduces the download within the assumption of a potential solution. Improvement on the present model is recommended to estimate the wall jets induced suction on the airfoil lower surface.

Lift augmentation on a delta wing via leading edge fences and the Gurney flap. M.S. Thesis

NASA Technical Reports Server (NTRS)

Buchholz, Mark D.

1992-01-01

Wind tunnel tests were conducted on two devices for the purpose of lift augmentation on a 60 deg delta wing at low speed. Lift, drag, pitching moment, and surface pressures were measured. Detailed flow visualization was also obtained. Both the leading edge fence and the Gurney flap are shown to increase lift. The fences and flap shift the lift curve as much as 5 deg and 10 deg, respectively. The fences aid in trapping vortices on the upper surface, thereby increasing suction. The Gurney flap improves circulation at the trailing edge. The individual influences of both devices are roughly additive, creating high lift gain. However, the lower lift to drag ratio and the precipitation of vortex burst caused by the fences, and the nose down pitching moment created by the flap are also significant factors.

LDV Surveys Over a Fighter Model at Moderate to High Angles of Attack

NASA Technical Reports Server (NTRS)

Sellers, William L., III; Meyers, James F.; Hepner, Timothy E.

2004-01-01

The vortex flowfield over an advanced twin-tailed fighter configuration was measured in a low-speed wind tunnel at two angles of attack. The primary test data consisted of 3-component velocity surveys obtained using a Laser Doppler Velocimeter. Laser light sheet and surface flow visualization were also obtained to provide insight into the flowfield structure. Time-averaged velocities and the root mean square of the velocity fluctuations were obtained at two cross-sections above the model. At 15 degrees angle of attack, the vortices generated by the wing leading edge extension (LEX) were unburst over the model and passed outboard of the vertical tail. At 25 degrees angle of attack, the vortices burst in the vicinity of the wing-LEX intersection and impact directly on the vertical tails. The RMS levels of the velocity fluctuations reach values of approximately 30% in the region of the vertical tails.

Mathematical Fluid Dynamics of Plasma Flow Control Over High Speed Wings

DTIC Science & Technology

2009-02-01

decreased voltage; e= 8, d= 1 mm. electrode u fe ^mmm^^n/* Fyd electrode Fig. 23 Schematics of momentum and heat source distributions for...For a>25°, the influence of DBD on the vortex breakdown is not so clear, because the breakdown point is very close to the wing apex in all three

Analytical and Experimental Investigations of Delta Wings in Incompressible Flow

DTIC Science & Technology

1976-08-01

posi- tion unless so designated by other official documents. Rep roduct ion Reproduction in whole or in part is permitted for any purpose of the...Trailing Edge of Free-Wake Model 56 19 Polar Coordinate System 5i 20 Free-Wake Geometry Prediction 5 21 Finite-Size Core 58 22 Vortex Core Position of Smith...k=0.4 70 34 Details of a Helical Type Burst 71 35 Vortex Burst-Steady Flow 72 36 Location of Pressure Ports 73 37 Pressure Destribution on a Delta

Visualization of a vortex flow in a rotating tank

NASA Astrophysics Data System (ADS)

Kawano, Yosuke

Flow structures of a vortex in a rotating tank were studied employing tracer method. The velocity measurements were made by photographing the motions of small polystyrene particles and analyzing strobo flash light pictures. The vortex flow is confined to a cylindrical region which is composed of a spiral upward flow in the center surrounded by an annular downward flow.

A lift formula applied to low-Reynolds-number unsteady flows

NASA Astrophysics Data System (ADS)

Wang, Shizhao; Zhang, Xing; He, Guowei; Liu, Tianshu

2013-09-01

A lift formula for a wing in a rectangular control volume is given in a very simple and physically lucid form, providing a rational foundation for calculation of the lift of a flapping wing in highly unsteady and separated flows at low Reynolds numbers. Direct numerical simulations on the stationary and flapping two-dimensional flat plate and rectangular flat-plate wing are conducted to assess the accuracy of the lift formula along with the classical Kutta-Joukowski theorem. In particular, the Lamb vector integral for the vortex force and the acceleration term of fluid for the unsteady inertial effect are evaluated as the main contributions to the unsteady lift generation of a flapping wing.

Surface-Streamline Flow Visualization

NASA Technical Reports Server (NTRS)

Langston, L.; Boyle, M.

1985-01-01

Matrix of ink dots covers matte surface of polyester drafting film. Film placed against wind-tunnel wall. Layer of methyl salicylate (oil of wintergreen) sprayed over dotted area. Ink dot streaklines show several characteristics of flow, including primary saddle point of separations, primary horseshoe vortex and smaller vortex at cylinder/ endwall junction. Surface streamline flow visualization technique suitable for use in low-speed windtunnels or other low-speed gas flows.

NAS (Numerical Aerodynamic Simulation Program) technical summaries, March 1989 - February 1990

NASA Technical Reports Server (NTRS)

1990-01-01

Given here are selected scientific results from the Numerical Aerodynamic Simulation (NAS) Program's third year of operation. During this year, the scientific community was given access to a Cray-2 and a Cray Y-MP supercomputer. Topics covered include flow field analysis of fighter wing configurations, large-scale ocean modeling, the Space Shuttle flow field, advanced computational fluid dynamics (CFD) codes for rotary-wing airloads and performance prediction, turbulence modeling of separated flows, airloads and acoustics of rotorcraft, vortex-induced nonlinearities on submarines, and standing oblique detonation waves.

Trimmed noncoplanar planforms with minimum vortex drag

NASA Technical Reports Server (NTRS)

Lamar, J. E.

1977-01-01

Vortex-lattice subsonic method determines mean camber surface for trimmed noncoplanar planforms with minimum vortex drag. Multiple surfaces can be designed together to yield trimmed configuration with minimum induced drag at some specified lift coefficient. Program is applicable to isolated wings, wing-canard configuration, tandem wing, and wing-winglet configuration.

Similitude relations for buffet and wing rock on delta wings

NASA Astrophysics Data System (ADS)

Mabey, D. G.

1997-08-01

Vortex flow phenomena at high angles of incidence are of great interest to the designers of advanced combat aircraft. The steady phenomena (such as steady lift and pitching moments) are understood fairly well, whereas the unsteady phenomena are still uncertain. This paper addresses two important unsteady phenomena on delta wings. With regard to the frequency parameter of the quasi-periodic excitation caused by vortex bursting, a new correlation is established covering a range of sweep back from 60 to 75°. With regard to the much lower frequency parameter of limit-cycle rigid-body wing-rock, a new experiment shows conclusively that although the motion is non-linear, the frequency parameter can be predicted by quasi-steady theory. As a consequence, for a given sweep angle, the frequency parameter is inversely proportional to the square root of the inertia in roll. This is an important observation when attempting to extrapolate from model tests in wind tunnels to predict the wing-rock characteristics of aircraft.

Visualization and analysis of flow structures in an open cavity

NASA Astrophysics Data System (ADS)

Liu, Jun; Cai, Jinsheng; Yang, Dangguo; Wu, Junqiang; Wang, Xiansheng

2018-05-01

A numerical study is performed on the supersonic flow over an open cavity at Mach number of 1.5. A newly developed visualization method is employed to visualize the complicated flow structures, which provide an insight into major flow physics. Four types of shock/compressive waves which existed in experimental schlieren are observed in numerical visualization results. Furthermore, other flow structures such as multi-scale vortices are also obtained in the numerical results. And a new type of shocklet which is beneath large vortices is found. The shocklet beneath the vortex originates from leading edge, then, is strengthened by successive interactions between feedback compressive waves and its attached vortex. Finally, it collides against the trailing surface and generates a large number of feedback compressive waves and intensive pressure fluctuations. It is suggested that the shocklets beneath vortex play an important role of cavity self-sustained oscillation.

Vortical structures of supersonic flow over a delta-wing on a flat plate

NASA Astrophysics Data System (ADS)

Wang, D. P.; Xia, Z. X.; Zhao, Y. X.; Wang, Q. H.; Liu, B.

2013-02-01

Employing the nanoparticle-based planar laser scattering (NPLS), supersonic flow over a delta-winged vortex generator on a flat plate was experimentally investigated in a supersonic quiet wind tunnel at Ma = 2.68. The fine structures of the flow field, shock waves, separation vortices, wake, and boundary layer transition were observed in the NPLS images. According to the time-correlation of the NPLS images and the measurement results of particle image velocimetry, the structural model of the flow field was improved further, and coherent wake structures were observed, which is of significance theoretically and in engineering application.

Infrared Images of Boundary Layer Transition on the D8 Transport Configuration in the LaRC 14- by 22-Foot Subsonic Tunnel

NASA Technical Reports Server (NTRS)

Mason, Michelle L.; Gatlin, Gregory M.

2015-01-01

Grit, trip tape, or trip dots are routinely applied on the leading-edge regions of the fuselage, wings, tails or nacelles of wind tunnel models to trip the flow from laminar to turbulent. The thickness of the model's boundary layer is calculated for nominal conditions in the wind tunnel test to determine the effective size of the trip dots, but the flow over the model may not transition as intended for runs with different flow conditions. Temperature gradients measured with an infrared camera can be used to detect laminar to turbulent boundary layer transition on a wind tunnel model. This non-intrusive technique was used in the NASA Langley 14- by 22-Foot Subsonic Tunnel to visualize the behavior of the flow over a D8 transport configuration model. As the flow through the wind tunnel either increased to or decreased from the run conditions, a sufficient temperature difference existed between the air and the model to visualize the transition location (due to different heat transfer rates through the laminar and the turbulent boundary layers) for several runs in this test. Transition phenomena were visible without active temperature control in the atmospheric wind tunnel, whether the air was cooler than the model or vice-versa. However, when the temperature of the model relative to the air was purposely changed, the ability to detect transition in the infrared images was enhanced. Flow characteristics such as a wing root horseshoe vortex or the presence of fore-body vortical flows also were observed in the infrared images. The images of flow features obtained for this study demonstrate the usefulness of current infrared technology in subsonic wind tunnel tests.

B-747 in Flight during Vortex Study

NASA Technical Reports Server (NTRS)

1974-01-01

Two chase aircraft, a Learjet and a Cessna T-37, are shown in formation with a Boeing B-747 jetliner in this 1974 NASA Flight Research Center (FRC) photograph. The two chase aircraft were used to probe the trailing wake vortices generated by the airflow around the wings of the B-747 aircraft. The vortex trails were made visible by smoke generators mounted under the wings of the B-747 aircraft. In 1974 the NASA Flight Research Center (later Dryden Flight Research Center, Edwards, California) used a Boeing 747 as part of the overall NASA study of trailing vortices. Trailing vortices are the invisible flow of spiraling air that trails from the wings of large aircraft and can 'upset' smaller aircraft flying behind them. The 747 that NASA used was on loan from the Johnson Space Center where it was part of the Space Shuttle Program. The data gathered in the 747 studies complemented data from the previous (1973-74) joint NASA Flight Research Center and Federal Aviation Administration (FAA) Boeing727 wake vortices study. Six smoke generators were installed under the wings of the 747 to provide a visual image of the trailing vortices. The object of the experiments was to test different configurations and mechanical devices on the747 that could be used to break up or lessen the strength of the vortices. The results of the tests could lead to shorter spacing between landings and takeoffs, which, in turn, could alleviate air-traffic congestion. For approximately 30 flights the 747 was flown using various combinations of wing air spoilers in an attempt to reduce wake vortices. To evaluate the effectiveness of the different configurations, chase aircraft were flown into the vortex sheets to probe their strengths and patterns at different times. Two of the chase planes used were the Flight Research Center's Cessna T-37 and the NASA Ames Research Center's Learjet. These aircraft represented the types of smaller business jets and other small aircraft that might encounter large passenger aircraft on approach or landings around major airports or in flight. Tests without the 747's wing spoilers deployed produced violent 'upset' problems for the T-37 aircraft at a distance of approximately 3 miles. From the magnitude of the problems found, distances of as much as ten miles might be required if spoilers were not used. With two spoilers on the outer wing panels, the T-37 could fly at a distance of three miles and not experience the 'upset' problem. The wake vortex study continued even after the 747 was returned to its primary mission of carrying the Space Shuttle.

Crossflow-Vortex Breakdown on Swept Wings: Correlation of Nonlinear Physics

NASA Technical Reports Server (NTRS)

Joslin, R. D.; Streett, C. L.

1994-01-01

The spatial evolution of cross flow-vortex packets in a laminar boundary layer on a swept wing are computed by the direct numerical simulation of the incompressible Navier- Stokes equations. A wall-normal velocity distribution of steady suction and blowing at the wing surface is used to generate a strip of equally spaced and periodic disturbances along the span. Three simulations are conducted to study the effect of initial amplitude on the disturbance evolution, to determine the role of traveling cross ow modes in transition, and to devise a correlation function to guide theories of transition prediction. In each simulation, the vortex packets first enter a chordwise region of linear independent growth, then, the individual packets coalesce downstream and interact with adjacent packets, and, finally, the vortex packets nonlinearly interact to generate inflectional velocity profiles. As the initial amplitude of the disturbance is increased, the length of the evolution to breakdown decreases. For this pressure gradient, stationary modes dominate the disturbance evolution. A two-coeffcient function was devised to correlate the simulation results. The coefficients, combined with a single simulation result, provide sufficient information to generate the evolution pattern for disturbances of any initial amplitude.

Review of the physics of enhancing vortex lift by unsteady excitation

NASA Technical Reports Server (NTRS)

Wu, J. Z.; Vakili, A. D.; Wu, J. M.

1991-01-01

A review aimed at providing a physical understanding of the crucial mechanisms for obtaining super lift by means of unsteady excitations is presented. Particular attention is given to physical problems, including rolled-up vortex layer instability and receptivity, wave-vortex interaction and resonance, nonlinear streaming, instability of vortices behind bluff bodies and their shedding, and vortex breakdown. A general theoretical framework suitable for handling the unsteady vortex flows is introduced. It is suggested that wings with swept and sharp leading edges, equipped with devices for unsteady excitations, could yield the first breakthrough of the unsteady separation barrier and provide super lift at post-stall angle of attack.

Wind Tunnel Measurements of the Wake of a Full-Scale UH-60A Rotor in Forward Flight

NASA Technical Reports Server (NTRS)

Wadcock, Alan J.; Yamauchi, Gloria K.; Schairer, Edward T.

2013-01-01

A full-scale UH-60A rotor was tested in the National Full-Scale Aerodynamics Complex (NFAC) 40- by 80-Foot Wind Tunnel in May 2010. The test was designed to acquire a suite of measurements to validate state-of-the-art modeling tools. Measurements include blade airloads (from a single pressure-instrumented blade), blade structural loads (strain gages), rotor performance (rotor balance and torque measurements), blade deformation (stereo-photogrammetry), and rotor wake measurements (Particle Image Velocimetry (PIV) and Retro-reflective Backward Oriented Schlieren (RBOS)). During the test, PIV measurements of flow field velocities were acquired in a stationary cross-flow plane located on the advancing side of the rotor disk at approximately 90 deg rotor azimuth. At each test condition, blade position relative to the measurement plane was varied. The region of interest (ROI) was 4-ft high by 14-ft wide and covered the outer half of the blade radius. Although PIV measurements were acquired in only one plane, much information can be gleaned by studying the rotor wake trajectory in this plane, especially when such measurements are augmented by blade airloads and RBOS data. This paper will provide a comparison between PIV and RBOS measurements of tip vortex position and vortex filament orientation for multiple rotor test conditions. Blade displacement measurements over the complete rotor disk will also be presented documenting blade-to-blade differences in tip-path-plane and providing additional information for correlation with PIV and RBOS measurements of tip vortex location. In addition, PIV measurements of tip vortex core diameter and strength will be presented. Vortex strength will be compared with measurements of maximum bound circulation on the rotor blade determined from pressure distributions obtained from 235 pressure sensors distributed over 9 radial stations.

Evolution of a Collection of Bubbles with Application to Wakes, Bubble Screens, and Cloud Noise

DTIC Science & Technology

1994-08-01

Hydrodynamics", Santa Barbara, CA, August 1994.. 2. G.L. CHAIIINE, "Bubble Interactions with Vortices," in "Vortex Flows," S. GREEN , ed., to be published by...2. G.L. CHAHINE, "Bubble Interactions with Vortices," in "Vortex Flows," S. GREEN , ed., to be published by Klttwer Academic, (1993). 3. G.L. CHAHINE...Tip Vortei, ASME Cavitation and Multiphase Flow Forum, Washington D.C., FED-VoL 153, pp. 93-99. (24] Green , S.I., 1991, "Correlatiag Single Phase Flow

An airborne system for vortex flow visualization on the F-18 high-alpha research vehicle

NASA Technical Reports Server (NTRS)

Curry, Robert E.; Richwine, David M.

1988-01-01

A flow visualization system for the F-18 high-alpha research vehicle is described which allows direct observation of the separated vortex flows over a wide range of flight conditions. The system consists of a smoke generator system, on-board photographic and video systems, and instrumentation. In the present concept, smoke is entrained into the low-pressure vortex core, and vortice breakdown is indicated by a rapid diffusion of the smoke. The resulting pattern is observed using photographic and video images and is correlated with measured flight conditions.

Unsteady aerodynamic analysis of space shuttle vehicles. Part 2: Steady and unsteady aerodynamics of sharp-edged delta wings

NASA Technical Reports Server (NTRS)

Ericsson, L. E.; Reding, J. P.

1973-01-01

An analysis of the steady and unsteady aerodynamics of sharp-edged slender wings has been performed. The results show that slender wing theory can be modified to give the potential flow static and dynamic characteristics in incompressible flow. A semiempirical approximation is developed for the vortex-induced loads, and it is shown that the analytic approximation for sharp-edged slender wings gives good prediction of experimentally determined steady and unsteady aerodynamics at M = 0 and M = 1. The predictions are good not only for delta wings but also for so-called arrow and diamond wings. The results indicate that the effects of delta planform lifting surfaces can be included in a simple manner when determining elastic launch vehicle dynamic characteristics. For Part 1 see (N73-32763).

A three dimensional unsteady iterative panel method with vortex particle wakes and boundary layer model for bio-inspired multi-body wings

NASA Astrophysics Data System (ADS)

Dhruv, Akash; Blower, Christopher; Wickenheiser, Adam M.

2015-03-01

The ability of UAVs to operate in complex and hostile environments makes them useful in military and civil operations concerning surveillance and reconnaissance. However, limitations in size of UAVs and communication delays prohibit their operation close to the ground and in cluttered environments, which increase risks associated with turbulence and wind gusts that cause trajectory deviations and potential loss of the vehicle. In the last decade, scientists and engineers have turned towards bio-inspiration to solve these issues by developing innovative flow control methods that offer better stability, controllability, and maneuverability. This paper presents an aerodynamic load solver for bio-inspired wings that consist of an array of feather-like flaps installed across the upper and lower surfaces in both the chord- and span-wise directions, mimicking the feathers of an avian wing. Each flap has the ability to rotate into both the wing body and the inbound airflow, generating complex flap configurations unobtainable by traditional wings that offer improved aerodynamic stability against gusting flows and turbulence. The solver discussed is an unsteady three-dimensional iterative doublet panel method with vortex particle wakes. This panel method models the wake-body interactions between multiple flaps effectively without the need to define specific wake geometries, thereby eliminating the need to manually model the wake for each configuration. To incorporate viscous flow characteristics, an iterative boundary layer theory is employed, modeling laminar, transitional and turbulent regions over the wing's surfaces, in addition to flow separation and reattachment locations. This technique enables the boundary layer to influence the wake strength and geometry both within the wing and aft of the trailing edge. The results obtained from this solver are validated using experimental data from a low-speed suction wind tunnel operating at Reynolds Number 300,000. This method enables fast and accurate assessment of aerodynamic loads for initial design of complex wing configurations compared to other methods available.

Characteristics of the flow around tandem flapping wings

NASA Astrophysics Data System (ADS)

Muscutt, Luke; Ganapathisubramani, Bharathram; Weymouth, Gabriel; The University of Southampton Team

2014-11-01

Vortex recapture is a fundamental fluid mechanics phenomenon which is important to many fields. Any large scale vorticity contained within a freestream flow may affect the aerodynamic properties of a downstream body. In the case of tandem flapping wings, the front wing generates strong large scale vorticity which impinges on the hind wing. The characteristics of this interaction are greatly affected by the spacing, and the phase of flapping between the front and rear wings. The interaction of the vorticity of the rear wing with the shed vorticity of the front wing may be constructive or destructive, increasing thrust or efficiency of the hind wing when compared to a wing operating in isolation. Knowledge of the parameter space where the maximum increases in these are obtained is important for the development of tandem wing unmanned air and underwater vehicles, commercial aerospace and renewable energy applications. This question is addressed with a combined computational and experimental approach, and a discussion of these is presented.

Development of the triplet singularity for the analysis of wings and bodies in supersonic flow

NASA Technical Reports Server (NTRS)

Woodward, F. A.

1981-01-01

A supersonic triplet singularity was developed which eliminates internal waves generated by panels having supersonic edges. The triplet is a linear combination of source and vortex distributions which gives directional properties to the perturbation flow field surrounding the panel. The theoretical development of the triplet singularity is described together with its application to the calculation of surface pressures on wings and bodies. Examples are presented comparing the results of the new method with other supersonic methods and with experimental data.

The Spectral and Statistical Properties of Turbulence Generated by a Vortex/Blade-Tip Interaction

NASA Technical Reports Server (NTRS)

Devenport, William J.; Wittmer, Kenneth S.; Wenger, Christian W.

1997-01-01

The perpendicular interaction of a streamwise vortex with the tip of a lifting blade was studied in incompressible flow to provide information useful to the accurate prediction of helicopter rotor noise and the understanding of vortex dominated turbulent flows. The vortex passed 0.3 chord lengths to the suction side of the blade tip, providing a weak interaction. Single and two-point turbulence measurements were made using sub-miniature four sensor hot-wire probes 15 chord lengths downstream of the blade trailing edge; revealing the mean velocity and Reynolds stress tensor distributions of the turbulence, as well as its spanwise length scales as a function of frequency. The single point measurements show the flow downstream of the blade to be dominated by the interaction of the original tip vortex and the vortex shed by the blade. These vortices rotate about each other under their mutual induction, winding up the turbulent wakes of the blades. This interaction between the vortices appears to be the source of new turbulence in their cores and in the region between them. This turbulence appears to be responsible for some decay in the core of the original vortex, not seen when the blade is removed. The region between the vortices is not only a region of comparatively large stresses, but also one of intense turbulence production. Velocity autospectra measured near its center suggests the presence quasi-periodic large eddies with axes roughly parallel to a line joining the vortex cores. Detailed two-point measurements were made on a series of spanwise cuts through the flow so as to reveal the turbulence scales as they would be seen along the span of an intersecting airfoil. The measurements were made over a range of probe separations that enabled them to be analyzed not only in terms of coherence and phase spectra but also in terms of wave-number frequency (kappa-omega) spectra, computed by transforming the measured cross-spectra with respect to the spanwise separation of the probes. These data clearly show the influence of the coherent eddies in the spiral wake and the turbulent region between the cores. These eddies produce distinct peaks in the upwash velocity kappa-omega spectra, and strong anisotropy manifested both in the decay of the kappa-omega spectrum at larger wave-numbers and in differences between the kappa-omega spectra of different components. None of these features are represented in the von Karman spectrum for isotropic turbulence that is often used in broadband noise computations. Wave-number frequency spectra measured in the cores appear to show some evidence that the turbulence outside sets tip core waves, as has previously been hypothesized. These spectra also provide for the first time a truly objective method for distinguishing velocity fluctuations produced by core wandering from other motions.

On the leading edge vortex of thin wings

NASA Astrophysics Data System (ADS)

Arredondo, Abel; Viola, Ignazio Maria

2016-11-01

On thin wings, the sharp leading edge triggers laminar separation followed by reattachment, forming a Leading Edge Vortex (LEV). This flow feature is of paramount importance because, if periodically shed, it leads to large amplitude load fluctuations, while if stably attached to the wing, it can provide lift augmentation. We found that on asymmetric-spinnaker-type yacht sails, the LEV can be stable despite the relatively low sweep (30°). This finding, which was recently predicted numerically by Viola et al., has been confirmed through current flume tests on a 1:115th model scale sail. Forces were measured and Particle Image Velocimetry was performed on four horizontal sail sections at a Reynolds number of 1.7x104. Vortex detection revealed that the LEV becomes progressively larger and more stable towards the highest sections, where its axis has a smaller angle with respect to the freestream velocity. Mapping the sail section on a rotating cylinder through a Joukowski transformation, we quantified the lift augmentation provided by the LEV on each sail section. These results open up new sail design strategies based on the manipulation of the LEV and can be applicable to the wings of unmanned aerial vehicles and underwater vehicles. Project funded by Conacyt.

Leading-edge vortex lifts swifts.

PubMed

Videler, J J; Stamhuis, E J; Povel, G D E

2004-12-10

The current understanding of how birds fly must be revised, because birds use their hand-wings in an unconventional way to generate lift and drag. Physical models of a common swift wing in gliding posture with a 60 degrees sweep of the sharp hand-wing leading edge were tested in a water tunnel. Interactions with the flow were measured quantitatively with digital particle image velocimetry at Reynolds numbers realistic for the gliding flight of a swift between 3750 and 37,500. The results show that gliding swifts can generate stable leading-edge vortices at small (5 degrees to 10 degrees) angles of attack. We suggest that the flow around the arm-wings of most birds can remain conventionally attached, whereas the swept-back hand-wings generate lift with leading-edge vortices.

A generalized vortex lattice method for subsonic and supersonic flow applications

NASA Technical Reports Server (NTRS)

Miranda, L. R.; Elliot, R. D.; Baker, W. M.

1977-01-01

If the discrete vortex lattice is considered as an approximation to the surface-distributed vorticity, then the concept of the generalized principal part of an integral yields a residual term to the vorticity-induced velocity field. The proper incorporation of this term to the velocity field generated by the discrete vortex lines renders the present vortex lattice method valid for supersonic flow. Special techniques for simulating nonzero thickness lifting surfaces and fusiform bodies with vortex lattice elements are included. Thickness effects of wing-like components are simulated by a double (biplanar) vortex lattice layer, and fusiform bodies are represented by a vortex grid arranged on a series of concentrical cylindrical surfaces. The analysis of sideslip effects by the subject method is described. Numerical considerations peculiar to the application of these techniques are also discussed. The method has been implemented in a digital computer code. A users manual is included along with a complete FORTRAN compilation, an executed case, and conversion programs for transforming input for the NASA wave drag program.

Wind Tunnel Visualization of the Flow Over a Full-Scale F/A-18 Aircraft

NASA Technical Reports Server (NTRS)

Lanser, Wendy R.; Botha, Gavin J.; James, Kevin D.; Crowder, James P.; Schmitz, Fredric H. (Technical Monitor)

1994-01-01

The proposed paper presents flow visualization performed during experiments conducted on a full-scale F/A-18 aircraft in the 80- by 120-Foot Wind-Tunnel at NASA Ames Research Center. This investigation used both surface and off-surface flow visualization techniques to examine the flow field on the forebody, canopy, leading edge extensions (LEXs), and wings. The various techniques used to visualize the flow field were fluorescent tufts, flow cones treated with reflective material, smoke in combination with a laser light sheet, and a video imaging system. The flow visualization experiments were conducted over an angle of attack range from 20deg to 45deg and over a sideslip range from -10deg to 10deg. The results show regions of attached and separated flow on the forebody, canopy, and wings. Additionally, the vortical flow is clearly visible over the leading-edge extensions, canopy, and wings.