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Showing papers on "Airfoil published in 1996"


Journal ArticleDOI
TL;DR: In this article, it was demonstrated that oscillatory blowing can delay separation from a symmetrical airfoil much more effectively than the steady blowing used traditionally for this purpose than the traditional slow blowing.
Abstract: It was recently demonstrated that oscillatory blowing can delay separation from a symmetrical airfoil much more effectively than the steady blowing used traditionally for this purpose. Experiments carried out on different airfoils revealed that this flow depends on many parameters such as, the location of the blowing slot, the steady and oscillatory momentum coefficients of the jet, the frequency of imposed oscillations, and the shape and incidence of the particular airfoil. In airfoils equipped with slotted flaps, the flow is also dependent on the geometry of the slot and on the Reynolds number in addition to the flap deflection that is considered as a part of the airfoil shape. The incremental improvements in single element airfoil characteristics are generally insensitive to a change in Reynolds number, provided the latter is sufficiently large. The imposed oscillations do not generate large oscillatory lift nor do they cause a periodic meander of the c.p. C* C D = dp Ct =

669 citations


Journal ArticleDOI
TL;DR: An implicit code for computing inviscid and viscous incompressible flows on unstructured grids is described and results are compared with an exact solution for theInviscid flow over a four-element airfoil.

304 citations


Journal ArticleDOI
TL;DR: In this paper, the effects of Reynolds number and angle of attack on boundary-layer separation from an Eppler 387 airfoil at low Reynolds number were investigated numerically.
Abstract: Unsteady boundary-layer separation from an Eppler 387 airfoil at low Reynolds number is studied numerically. Through a series of computations, the effects of Reynolds number and angle of attack are investigated. For all cases, vortex shedding is observed from the separated shear layer. From linear stability analysis, a KelvinHelmholtz instability is identified as causing shear layer unsteadiness. The low-turbulence wind-tunnel tests of the Eppler 387 airfoil are used to compare with the time-averaged results of the present unsteady computations. The favorable comparison between computational and experimental results strongly suggests that the unsteady largescale structure controls the low-Reynolds-number separation bubble reattachment with small-scale turbulence playing a secondary role. Nomenclature C = chord length CD - drag coefficient CL = lift coefficient Cp = pressure coefficient / = shedding frequency Re = chord Reynolds number R P - reattachment point S P = separation point Sr = Strouhal number U = velocity 9 = momentum thickness Subscripts sep = conditions at separation oo = freestream conditions

233 citations


Journal ArticleDOI
TL;DR: In this article, the effect of compressibility on dynamic stall behavior has been comprehensively studied, including a review of work performed on both aircraft and helicopters, and offers insight into the impact of compressible airfoils on the complex aerodynamic phenomenon known as dynamic stall.

160 citations


Journal ArticleDOI
TL;DR: In this article, a multiblock Navier-Stokes solver is employed to compute unsteady flow fields around a single flapping airfoil and the unstaired potential flow code is also computed.
Abstract: Thrust generation on a single flapping airfoil and a flapping/stationary airfoil combination in tandem is studied parametrically. A multiblock Navier-Stokes solver is employed to compute unsteady flowfields. The unsteady flowfield around a single flapping airfoil is also computed by an unsteady potential flow code. The numerical solutions predict thrust generation in flapping airfoils and a significant augmentation of thrust in flapping/stationary airfoil combinations in tandem. The propulsive efficiency is found to be a strong function of reduced frequency and the amplitude of the flapping motion. At a flapping amplitude of 0.40 chord lengths and a reduced frequency of 0.10, the propulsive efficiency of a single NACA 0012 airfoil was computed to be more than 70 %. For the airfoil combination in tandem, the propulsive efficiency was augmented more than 40% at a reduced frequency of 0.75 and a flapping amplitude of 0.20 chord lengths when the airfoils are separated by about two chord lengths.

154 citations



Journal ArticleDOI
TL;DR: In this paper, a two-dimensional, low-Mach-number laminar flow past a NACA 0012 airfoil at the chord Reynolds number of 10 4 was analyzed.
Abstract: Acoustic analogy computations of vortex shedding noise were carried out in the context of a two-dimensional, low-Mach-number laminar flow past a NACA 0012 airfoil at chord Reynolds number of 10 4 . The incompressible Navier-Stokes equations were solved numerically to give an approximate description of the near-field flow dynamics and the acoustic source functions. The radiated far-field noise was computed based on Curle's extension to the Lighthill analogy. This study emphasizes an accurate evaluation of the Reynolds stress quadrupoles in the presence of an extensive wake. An effective method for separating the physical noise source from spurious boundary contributions caused by eddies crossing a permeable computational boundary is presented. The effect of retarded-time variations across the source region is also examined. Computational solutions confirm that the quadrupole noise is weak compared with the noise due to lift and drag dipoles when the freestream Mach number is small. The techniques developed in this study are equally applicable to flows in which the volume quadrupoles act as a prominent noise source.

133 citations


Journal ArticleDOI
TL;DR: In this paper, a genetic algorithm was applied to optimize target pressure distributions for inverse design methods, where pressure distributions around airfoils were parameterized by B-spline polygons, and the airfoil drag was minimized under constraints on lift, air foil thickness, and other design principles.
Abstract: A genetic algorithm has been applied to optimize target pressure distributions for inverse design methods. Pressure distributions around airfoils are parameterized by B-spline polygons, and the airfoil drag is minimized under constraints on lift, airfoil thickness, and other design principles. Once target pressure distributions are obtained, corresponding airfoil/wing geometries can be computed by an inverse design code coupled with a Navier-Stokes solver. Successful design results were obtained for transonic cases with and without a shock wave.

132 citations


Journal ArticleDOI
TL;DR: In this paper, the dynamic stall process on a pitching NACA 0012 airfoil was investigated by two experimental techniques, particle image velocimetry (PIV) and laser-sheet visualizations.
Abstract: The dynamic stall process on a pitching NACA 0012 airfoil was investigated by two experimental techniques-particle image velocimetry (PIV) and laser-sheet visualizations-and a numerical code based on the Navier-Stokes equations. The freestream velocity was 28 m/s, leading to a Reynolds number (based on airfoil chord) of 3.73 X 10 5 . The airfoil motion was a sinusoidal function between 5 and 25 deg of incidence, with a frequency of 6.67 Hz corresponding to a reduced frequency (based on airfoil half-chord) of 0.15. The out-of-plane component of the vorticity could be derived from the PIV velocity fields. The comparison between experimental and numerical results was conducted for the four main phases of the dynamic stall process, i.e., attached flow, development of the dynamic stall vortex, poststall vortex shedding, and reattachment. In general, the computational results agreed very well with the experimental results. However, some discrepancies were observed and discussed. The cycle-to-cycle nonreproducibility of the flowfield during the phase of massive separation is also mentioned.

129 citations


Journal ArticleDOI
TL;DR: An aerodynamic shape optimization procedure based on discrete sensitivity analysis is extended to treat three-dimensional geometries and results in a significant factor of 50 decrease in computational time and a factor of eight reduction in memory over the most efficient design strategies in current use.
Abstract: An aerodynamic shape optimization procedure based on discrete sensitivity analysis is extended to treat three-dimensional geometries. The function of sensitivity analysis is to directly couple computational fluid dynamics (CFD) with numerical optimization techniques, which facilitates the construction of efficient direct-design methods. The development of a practical three-dimensional design procedures entails many challenges, such as: (1) the demand for significant efficiency improvements over current design methods; (2) a general and flexible three-dimensional surface representation; and (3) the efficient solution of very large systems of linear algebraic equations. It is demonstrated that each of these challenges is overcome by: (1) employing fully implicit (Newton) methods for the CFD analyses; (2) adopting a Bezier-Bernstein polynomial parameterization of two- and three-dimensional surfaces; and (3) using preconditioned conjugate gradient-like linear system solvers. Whereas each of these extensions independently yields an improvement in computational efficiency, the combined effect of implementing all the extensions simultaneously results in a significant factor of 50 decrease in computational time and a factor of eight reduction in memory over the most efficient design strategies in current use. The new aerodynamic shape optimization procedure is demonstrated in the design of both two- and three-dimensional inviscid aerodynamic problems including a two-dimensional supersonic internal/external nozzle, two-dimensional transonic airfoils (resulting in supercritical shapes), three-dimensional transport wings, and three-dimensional supersonic delta wings. Each design application results in realistic and useful optimized shapes.

117 citations


Journal ArticleDOI
TL;DR: In this article, an approach based on indicial concepts is described to model the unsteady airloads on a thin airfoil in subsonic compressible flow caused by the arbitrary motion of a trailing-edge flap.
Abstract: An approach based on indicial concepts is described to model the unsteady airloads on a thin airfoil in subsonic compressible flow caused by the arbitrary motion of a trailing-edge flap. Exact indicial aerodynamic responses at small values of time as a result of flap deflection and angular deflection rate about the flap hinge are obtained from linear unsteady subsonic theory in conjunction with the aerodynamic reverse flow theorems. Using the known exact initial (piston theory) and asymptotic values of the airloads, along with an assumed analytic form for the indicial functions, these exact results are used to help obtain complete approximations for the respective indicial responses. The airloads from arbitrary flap motion in subsonic flow are subsequently obtained in state - space form. Validation of the method is conducted with experimental data for time-dependent flap motions.

Journal ArticleDOI
Earl H. Dowell1
TL;DR: In this article, the Lanczos algorithm is used to solve the sparse eigenvalue problem of the unsteady flow model and a reduced-order model of the flow is constructed.
Abstract: A conceptually novel and computationally efficient technique for computing unsteady flow about isolated airfoils, wings, and turbomachinery cascades is presented. Starting with either a time-domain or frequency-domain computational fluid dynamics analysis of unsteady aerodynamic or aeroacoustic flows, a large, sparse eigenvalue problem is solved using the Lanczos algorithm. Then, using just a few of the resulting eigenmodes, a reduced-order model of the unsteady flow is constructed. With this model, one can rapidly and accurately predict the unsteady aerodynamic response of the system over a wide range of reduced frequencies. Moreover, the eigenmode information provides important insights into the physics of unsteady flows.

Journal ArticleDOI
TL;DR: In this article, the effect of Reynolds number on the aerodynamic characteristics of an airfoil with ground effect in viscous flow is investigated by numerical method, based on the standard k-e turbulence model, generalized body-fixed coordinates and the finite volume method.
Abstract: The effect of Reynolds number on the aerodynamic characteristics of an airfoil with ground effect in viscous flow is investigated by numerical method. A numerical scheme, based on the standard k-e turbulence model, generalized body-fixed coordinates and the finite volume method, is developed to solve the two-dimensional wing-in-ground problem hi viscous flow. The steady, incompressible Navier-Stokes equations are solved using a grid generation program developed by the authors, and the PHOENICS code. Some numerical results are presented to show the effects of Reynolds number, ground clearance, and angles of attack on the aerodynamic characteristics of a NACA 4412 airfoil.

Journal ArticleDOI
TL;DR: In this article, an aeroelastic model of flexible membrane wing aerodynamics which incorporates the Reynolds-averaged Navier-Stokes equations is presented, where the Reynolds stresses are prescribed by the k-ω shear-stress transport eddy-viscosity model recently proposed by Menter.
Abstract: In the present paper an aeroelastic model of flexible membrane wing aerodynamics which incorporates the Reynolds‐averaged Navier–Stokes equations is presented. The Reynolds stresses are prescribed by the k–ω shear‐stress transport eddy‐viscosity model recently proposed by Menter. The computed coefficients are compared with classical inviscid membrane airfoil theory and with a portion of the available experimental data for membrane wings. The results indicate that classical potential‐based membrane airfoil theory can provide a meaningful description of membrane wing aerodynamics only for a small range of incidence angles near ideal and then only for membrane airfoils with small excess length ratios. For larger excess lengths and incidence angles viscous effects dominate the aerodynamics. The agreement of the computed results with the experimental data is mixed. The current status of the available experimental data for membrane airfoils is also reviewed.

Patent
05 Feb 1996
TL;DR: In this article, a method and an apparatus are provided for optimizing the aerodynamic effect of the airfoil of an aircraft by defined changes in camber, which includes the following steps: a. determining the flow for the flight condition caused by the change in Camber, b. comparing the ascertained characteristic values with stored nominal reference values for an optimal flow, c. deriving actuator signals from the differential values, and e.
Abstract: A method and an apparatus are provided for optimizing the aerodynamic effect of the airfoil of an aircraft by defined changes in camber. The method includes the following steps: a. determining the flow for the flight condition caused by the change in camber, b. comparing the ascertained characteristic values with stored nominal reference values for an optimal flow, c. forming differential values between the characteristic values and the stored nominal reference values, d. deriving actuator signals from the differential values, and e. changing the camber by motor, based on the actuator signals, for minimizing the differential values. The optimum wing flow is thereby maintained more exactly. For transonic wings, the position and strength of compression shocks is also effectively controlled, which leads to a reduction of the direct shock induced separation.

Proceedings ArticleDOI
01 Jan 1996
TL;DR: Selig and Guglielmo as discussed by the authors measured the lift and drag characteristics of 34 airfoils at low Reynolds numbers in an attempt to develop a consistent database for use in design studies that require accurate low Reynolds number airfoil data.
Abstract: Lift and drag measurements were taken on 34 airfoils at low Reynolds numbers in an attempt to develop a consistent database for use in design studies that require accurate low Reynolds number airfoil data. Prom these data emerged several interesting results related to the behavior of the laminar separation bubbles and their effect on the lift characteristics. A plateau hi the lift curve of symmetrical airfoils hi the vicinity of an angle of attack of 0 deg was found to be common in the Reynolds number range of 40,000 to 100,000. Through the use of zig-zag type boundary-layer trips, this nonlinearity can be reduced owing to a reduction in the size of the laminar separation bubbles. The influence of laminar separation bubbles was also found to dominate the performance of several high-lift airfoils in the Reynolds number range of 80,000 to 150,000. In particular, hysteresis loops in the lift curve were present, and these are related to the size of the laminar separation bubble as deduced from drag data. The data reveals that some airfoils exhibit both counterclockwise aoid clockwise hysteresis loops for a given Reynolds number. Moreover, depending on the airfoil, either type of loop can occur first. Introduction A wide variety of small unmanned aerial vehicles (UAVs) operate in the low Reynolds number regime in which the airfoil aerodynamics play a key role in aircraft performance and handling. In this regime, natural boundary-layer transition takes place through a laminar separation bubble that forms as the laminar boundary layer first separates, then becomes unstable, makes a transition to turbulent flow and reattaches to the airfoil to form a laminar separation bubble.' This bubble often results in a notable airfoil-performance degradation that is characterized by undesirable high drag, nonlinearCopyright © 1996 by Michael S. Selig, James J. Guglielmo, Andy P. Broeren and Philippe Giguere. Published by the American Institute of Aeronautics and Astronautics, Inc. with permission. 'Assistant Professor. Member AIAA. fGraduate Research Assistant. Student Member AIAA. ^Graduate Research Assistant. ities in the lift-curve characteristics and sometimes static hysteresis in the section lift, drag and moment data.' In an effort to understand the flow phenomena at low Reynolds numbers, there have been numerous theoretical and experimental investigations that have resulted in three conferences" and a special AGARD publication. Because of the difficulty in modeling the laminar separation bubble, computational efforts, while vital for use in design, have not been entirely reliable hi predicting these complex flows. Nonetheless, considerable progress has been made in recent years." Despite the high level of interest in this area, few systematic experiments have been performed to document the performance of a wide selection of airfoils for use in conceptual and detailed design studies. This has been particularly problematic because comparisons of data between different wind tunnels facilities regularly show discrepencies owing to the documented difficulties in measuring low Reynolds number airfoil performance. Thus, the mixture of different data sets is not ideally suited for the purposes of examining performance trade-offs involving different airfoils. A key objective of the present research was to produce a large and consistent low Reynolds number airfoil performance database for use in the design of small UAVs. Specifically, the experiments involved measuring the lift and drag characteristics of many airfoils over the nominal Reynolds number range of 60,000 to 300,000. The collection of airfoils tested is depicted in Fig. 1. The more limited objective of this paper is to highlight and explain two interesting features observed in the lift characteristics of some of the airfoil tested. In particular, this paper presents and discusses (1) the nonlinear lift characteristics of two symmetrical airfoils (NACA 0009 and SD8020) as well as improvements made through the use of boundary-layer trips, and (2) the hysteresis in the lift characteristics of six high-lift airfoils (FX 63-137, S1210, modified FX 74CL5-140, CH 10-48-13, M06-13-128 and S1223). As will be described, both features, the nonlinearity and the hysteresis of the lift curves, can be linked to the behavior of the laminar separation bubbles. Anti-Turbulence Screens Diffuser Silencer N. / Frequency Controller / Fan \

Journal ArticleDOI
TL;DR: In this paper, the authors further document the characteristics of the low-frequency flow oscillation on airfoils near stall and gain additional insight into the flow mechanisms, including the mechanism of selection of the long time scale involved in the oscillation.
Abstract: The low-frequency flow oscillation on airfoils near stall has been observed on several airfoils over a range of Reynolds numbers. However, the mechanism of selection of the long time scale involved in the oscillation remained far from clearly understood. This provided motivation to continue to pursue the topic. Here, efforts are made to further document the characteristics of the phenomenon and gain additional insight into the flow mechanisms.


Journal ArticleDOI
TL;DR: In this article, a bifurcation analysis of a two-dimensional airfoil with a structural nonlinearity in the pitch direction and subject to incompressible flow is presented.
Abstract: A bifurcation analysis of a two-dimensional airfoil with a structural nonlinearity in the pitch direction and subject to incompressible flow is presented. The nonlinearity is an analytical third-order rational curve fitted to a structural freeplay. The aeroelastic equations-of-motion are reformulated into a system of eight first-order ordinary differential equations. An eigenvalue analysis of the linearized equations is used to give the linear flutter speed. The nonlinear equations of motion are either integrated numerically using a fourth-order Runge-Kutta method or analyzed using the AUTO software package. Fixed points of the system are found analytically and regions of limit cycle oscillations are detected for velocities well below the divergent flutter boundary. Bifurcation diagrams showing both stable and unstable periodic solutions are calculated, and the types of bifurcations are assessed by evaluating the Floquet multipliers. In cases where the structural preload is small, regions of chaotic motion are obtained, as demonstrated by bifurcation diagrams, power spectral densities, phase-plane plots and Poincare sections of the airfoil motion; the existence of chaos is also confirmed via calculation of the Lyapunov exponents. The general behaviour of the system is explained by the effectiveness of the freeplay part of the nonlinearity in a complete cycle of oscillation. Results obtained using this reformulated set of equations and the analytical nonlinearity are in good agreement with previously obtained finite difference results for a freeplay nonlinearity.

Proceedings ArticleDOI
17 Jun 1996
TL;DR: It is shown that, due to its neat-elimination of CPU cost dependence on the number of design variables, the adjoint method is preferred over the direct and finite difference methods for practical single-discipline aerodynamic optimization.
Abstract: A methodology for performing optimization on 2D and 3D unstructured grids based on the Euler equations is presented. The same, low-memory-cost explicit relaxation algorithm is used to resolve the discrete equations which govern the flow, linearized direct and adjoint problems. The analysis schemes, for both 2D and 3D, are high resolution Local-Extremum-Diminishing (LED) schemes and use Roe decomposition for the dissipative fluxes. The local timestepping relaxation scheme is based on a multidimensional equivalent of a TVD CFL-like condition guaranteeing convergence of flow and sensitivity computations to machine accuracy. Mesh movement is performed in such a way that optimization of arbitrary geometries is allowed. Sensitivities based on direct and adjoint methods are validated and sample optimizations are performed: the inverse pressure design of a multielement airfoil in high-lift mode, an infinite-span straight transonic wing and a transonic wing/body configuration. It is shown that, due to its neat-elimination of CPU cost dependence on the number of design variables, the adjoint method is preferred over the direct and finite difference methods for practical single-discipline aerodynamic optimization.

Journal ArticleDOI
Sang-Hyun Kim1, In Lee1
TL;DR: In this paper, a two-dimensional flexible airfoil with a free-play non-linearity in pitch has been analyzed in the subsonic flow range, and nonlinear aeroelastic analyses for both the frequency domain and time domain are performed for rigid and flexible air-foil models to investigate the flexibility effect.

Journal ArticleDOI
TL;DR: In this article, an experimental investigation was made of the gust field generated by a rotating slotted cylinder installed in the Duke University low-speed, closed-circuit wind tunnel.
Abstract: An experimental investigation was made of the gust field generated by a rotating slotted cylinder installed in the Duke University low-speed, closed-circuit wind tunnel. The system has a very simple configuration with low cost and can produce a controllable single or multiple harmonic gust wave in the lateral and longitudinal directions. It requires minimal power and torque input. A simplified theoretical aerodynamic model and a design estimation of the lateral and longitudinal gust flowfield is also proposed in this article. The design estimate is based on a two-dimensional dynamic lift coefficient that is given by the theoretical and experimental results. An interfering wake vortex effect is the major disadvantage of this system. Nomenclature C(k) = Theodorsen's function ICleql = magnitude of equivalent lift coefficient for rotating slotted cylinder/airfoil c = airfoil chord d = cylinder diameter dLa = airfoil lift force per span length dLrsc = rotating slotted cylinder lift force per span length e, e = gap between the o.d. of the rotating slotted cylinder and trailing edge of the airfoil, elc H, H = vertical position from tunnel bottom, HIHW Hw = height of the tunnel test section

Patent
29 Jan 1996
TL;DR: In this article, a composite airfoil was proposed for a wide chord fan blade having a high degree of twist in a large high bypass ratio turbofan engine. But it was only suitable for a single-passenger aircraft.
Abstract: The present invention provides a composite airfoil, particularly useful as a wide chord fan blade having a high degree of twist in a large high bypass ratio turbofan engine. The composite airfoil of the present invention has a reinforced region of its airfoil that extends a portion of its span from its tip and a portion of its chord from its trailing edge. The region is covered by thin metallic sheathing bonded to trailing edge surfaces of the blade in a manner to reinforce that portion of the composite blade.

Proceedings ArticleDOI
15 Jan 1996
TL;DR: In this paper, the effect of large-droplet ice accretion on aircraft control and in particular lateral control is examined and it is shown that a ridge of ice aft of the boot can lead to large losses in lift, increases in drag and changes in the pitching moment.
Abstract: The effect of large-droplet ice accretion on aircraft control and in particular lateral control is examined. Supercooled large droplet icing conditions can result in the formation of a ridge of ice aft of the upper surface boot. By comparing this ice shape to data acquired with a spanwise protuberance on a different airfoil, it is clear that a ridge of ice aft of the boot can lead to large losses in lift, increases in drag and changes in the pitching moment. This effect is most likely due to the formation of a large separation bubble aft of the ice accretion which grows with angle of attack and eventually fails to reattach, leading to premature airfoil stall. The bubble alters the pressure distribution about the airfoil resulting in a more trailing edge up (negative) hinge moment on the aileron and the resulting change in aileron stick force. This can lead to aileron hinge moment reversal and aileron snatch. In aileron snatch the hinge moments are altered to the extent that the aileron is pulled up by the low pressure over the upper surface of the aileron with sufficient force to induce a rapid roll if a large stick force is not immediately exerted to oppose it. There is evidence in the literature which shows that similar lateral control problems are possible with other types of ice accretions and airfoil types.

Patent
15 Nov 1996
TL;DR: A swept turbomachinery blade for use in a cascade of such blades is disclosed in this article, where the blade has an airfoil uniquely swept so that an endwall shock of limited radial extent and a passage shock are coincident and a working medium flowing through interblade passages is subjected to a single coincident shock rather than the individual shocks.
Abstract: A swept turbomachinery blade for use in a cascade of such blades is disclosed. The blade ( 12 ) has an airfoil ( 22 ) uniquely swept so that an endwall shock ( 64 ) of limited radial extent and a passage shock ( 66 ) are coincident and a working medium ( 48 ) flowing through interblade passages ( 50 ) is subjected to a single coincident shock rather than the individual shocks. In one embodiment of the invention the forwardmost extremity of the airfoil defines an inner transition point ( 40 ) located at an inner transition radius r t -inner. The sweep angle of the airfoil is nondecreasing with increasing radius from the inner transition radius to an outer transition radius r t-outer , radially inward of the airfoil tip ( 26 ), and is nonincreasing with increasing radius between the outer transition radius and the airfoil tip.

Proceedings ArticleDOI
17 Jun 1996
TL;DR: In this article, the effectiveness of the actuators in improving airfoil performance degraded with increases in Mach number; however, even at the highest Mach number, modest lift enhancements were observed.
Abstract: Separation control experiments were performed on airfoil models with freestream Mach numbers ranging from 0.1 to 0.5. This range of Mach numbers reaches into the transonic flow regime for the models tested. Pulsed jet control actuators were located near the airfoil leading edge. Experiments were performed to determine the effectiveness of the control actuators in reducing flow separation and augmenting lift. At low Mach numbers, the pulsed jet actuators produced up to a 50 percent increase in lift with only a small change in drag. The effectiveness of the actuators in improving airfoil performance degraded with increases in Mach number; however, even at the highest Mach number, modest lift enhancements were observed. In general, maximum performance benefits were observed for angles of attack equal to and above those corresponding to Cl(max) where severe flow separation exists in the uncontrolled flow. (Author)

Patent
25 Apr 1996
TL;DR: In this paper, a hollow airfoil for a gas turbine engine having a leading edge (12), a trailing edge (14), a pressure side (20), and a suction side (22) includes a plurality of internal spanwise stiffening ribs (31-35) that are arranged in a logarithmic pattern.
Abstract: A hollow airfoil (10) for a gas turbine engine having a leading edge (12), a trailing edge (14), a pressure side (20), and a suction side (22) includes a plurality of internal spanwise stiffening ribs (31-35) that are arranged in a logarithmic pattern. The particular arrangement of internal ribs (31-35) optimizes stiffness of the airfoil (10) without significantly increasing the weight thereof.

Journal ArticleDOI
TL;DR: Indicial approximations for the lift on an airfoil penetrating a stationary sharp-edge gust in two-dimensional al subsonic flow have been derived in this article.
Abstract: Indicial approximations are derived for the lift on an airfoil penetrating a stationary sharp-edge gust in two-dimension al subsonic flow. Using an assumed exponential form, the approximations have been generalized in terms of Mach number alone by means of an optimization algorithm where certain coefficients of the approximations are free parameters. The optimization is subject to prescribed constraints in terms of the initial and asymptotic behavior of the gust response, and by requiring the response closely match the known exact solutions given by subsonic linear theory at earlier values of time. An alternative approximation is obtained by using results from a direct numerical simulation of the gust problem using computational fluid dynamcs (CFD). For an airfoil-vortex interaction problem, comparisons were made with experimental data and CFD results. Finally, the indicial method was integrated into a three-dimensional rotor simulation, and the near- and far-field acoustics were computed using the Ffowcs WilliamsHawkins equation. Good agreement was found with simultaneously measured airloads and acoustics data.

Journal ArticleDOI
TL;DR: In this paper, a new technique for measuring skin friction was employed to help document the flow on an airfoil at angles of attack from -0.5 to 11.5 deg.
Abstract: A new technique for measuring skin friction was employed to help document the flow on an airfoil at angles of attack from -0.5 to 11.5 deg. Surface pressures were also measured on both the wing and wind-tunnel walls. The experiment was conducted at a freestream Mach number of 0.2 and Reynolds numbers of 0.6, 2, and 6 x 10 6 . The objective of the study was to provide data and boundary condition information sufficient for the validation of numerical simulations. Such a simulation of the experiment was conducted using the INS2D Navier-Stokes code with the shear-stress-transport turbulence model. The computations provide a good description of both laminar and turbulent shear levels, except for turbulent flow on the top surface of the wing at the higher angles of attack.

Journal ArticleDOI
TL;DR: In this paper, the results of viscous drag reduction using riblets from 3M on a NACA 0012 airfoil model up to moderate angles of attack are presented.
Abstract: Results of viscous drag reduction using riblets from 3M on a NACA 0012 airfoil model up to moderate angles of attack are presented. Measurements made consisted of model surface pressure distributions, mean velocity and streamwise turbulence intensity profiles in the boundary layer (just ahead of the trailing edge), and total airfoil drag for two riblet heights of 0.152 and 0.076 mm. Results show significantly higher skin friction drag reduction with incidence compared to plat plate flows; the reduction was as high as 16% at ? = 6 deg. Results of mean velocity profiles show that a larger contribution to drag reduction results from the suction side of the airfoil, indicating increased effectiveness of riblets in adverse pressure gradients. Examination of turbulence intensity profiles in the wall region indicates an appreciable reduction in the presence of riblets; correspondingly, the spectra show reduced energy levels at low frequencies.