Showing papers on "Airfoil published in 1999"

Detached-eddy simulation of an airfoil at high angle of attack

Abstract: We present the first true applications of Detached-Eddy Simulation (DES), in the sense of being three-dimensional. DES was defined in 1997 with hopes of combining the strengths of Reynolds-averaged methods and of Large-Eddy Simulations, in a non-zonal manner, to treat separated flows at high Reynolds numbers. We first simulate isotropic turbulence, to check the concept in LES mode and set its adjustable constant. Smooth inertial ranges are obtained up to the cutoff in the spectra. We then treat an airfoil in the challenging regime of massive separation and do so very successfully, in that lift and drag are within 10% of the experimental results at all angles of attack, to 90°. Such an accuracy is not achieved with traditional modelling, even unsteady, which gives up to 40% error. Cost puts a pure LES of the same flow (at Reynolds number 105 and beyond) out of reach on any computer, yet we use personal computers for the DES, and about 200,000 grid points. On the other hand, grid refinement, domain-size and Reynolds-number studies have not been completed yet. Hysteresis in the 15 - 25° range has not been addressed.

Parametric Airfoils and Wings

Abstract: Explicit mathematical functions are used for 2D curve definition for airfoil design. Flowphe-nomena-oriented parameters control geometrical and aerodynamic properties. Airfoil shapes are blended with known analytical section formulae. Generic variable camber wing sections and multicomponent airfoils are generated. For 3D wing definition all parameters are made functions of a third spanwise coordinate. High lift systems are defined kinematically by modelled track gear geometries, translation and rotation in 3D space. Examples for parameter variation in numerical optimization, mechanical adaptation and for unsteady coupling of flow and configuration are presented.

Jet characteristics of a plunging airfoil

Abstract: Water-tunnel tests of a NACA 0012 airfoil that was oscillated sinusoidally in plunge are described. The flowered downstream of the airfoil was explored by dye flow visualization and single-component laser Doppler velocimetry (LDV) measurements for a range of freestream speeds, frequencies, and amplitudes of oscillation. The dye visualizations show that the vortex patterns generated by the plunging airfoil change from drag-producing wake flows to thrust-producing jet flows as soon as the ratio of maximum plunge velocity to freestream speed, i.e., the nondimensional plunge velocity, exceeds approximately 0.4. The LDV measurements show that the nondimensional plunge velocity is the appropriate parameter to collapse the maximum streamwise velocity data covering a nondimensional plunge velocity range from 0.18 to 9.3

Oscillatory Control of Separation at High Reynolds Numbers

Abstract: An experiment conducted in a pressurized, cryogenic wind tunnel demonstrates that unsteady flow control using oscillatory blowing (with essentially zero mass flux) can effectively delay flow separation and reattach separated flow on an airfoil at chord Reynolds numbers as high as 38 × 10 6 . Oscillatory blowing at frequencies that generate one to three vortices over the controlled region at all times are effective over the entire Reynolds number range, in accordance with previous low-Reynolds-number tests. Stall is delayed and poststall characteristics are improved when oscillatory blowing is applied from the leading-edge region of the airfoil, whereas flap effectiveness is increased when control is applied at the flap shoulder. Similar gains in airfoil performance require steady blowing with a momentum coefficient that is two orders of magnitude greater. A detailed experimental and theoretical investigation was undertaken to characterize the oscillatory blowing disturbance, in the absence of external flow, and to estimate the oscillatory blowing momentum coefficient used in the cryogenic wind tunnel experiment. Possible approaches toward closed-loop active separation control are also presented

Airfoil Design on Unstructured Grids for Turbulent Flows

Abstract: An aerodynamic design algorithm for turbulent e ows using unstructured grids is described. The current approachusesadjoint (costate)variablestoobtainderivativesofthecostfunction.Thesolutionoftheadjointequations is obtained by using an implicit formulation in which the turbulence model is fully coupled with the e ow equations when solving for the costate variables. The accuracy of the derivatives is demonstrated by comparison with e nite difference gradients, and a few sample computations are shown. Recommendations on directions of further research into the Navier ‐Stokes design process are made. Nomenclature A = area of control volume a = speed of sound C ¤ = constant used in Sutherland’ s law for viscosity cb1;cb2;cv1; = constants used in Spalart ‐Allmaras cw1;cw2;cw3 turbulence model cd = drag cl = lift c1

Effects of Dynamic Stall on Propulsive Efficiency and Thrust of Flapping Airfoil

Abstract: Numerical simulations of dynamic stall phenomena around an airfoil oscillating in a coupled mode, in which the pitching and heaving oscillations have some phase difference, have been performed with a Navier ‐Stokes code. The propulsive efe ciency and the thrust have been calculated for various combinations of the phase difference and the reduced frequency for two different amplitude ratios. The effects of the dynamic stall phenomena on the behaviors of the propulsive efe ciency and thrust are discussed in detail by examination of each e ow pattern obtained. Highest efe ciency has been observed for the case in which the pitching oscillation advances 90 deg ahead of the heaving oscillation and the reduced frequency is at some optimum value, for which there appears no appreciable e ow separation in spite of large-amplitude oscillations. For phase angles and reduced frequency other than thisbestcondition, efe ciency israpidly degraded by theoccurrenceof thelarge-scaleleading-edgeseparation.

Heat Transfer and Flow on the First Stage Blade Tip of a Power Generation Gas Turbine: Part 1 — Experimental Results

Abstract: A combined experimental and computational study has been performed to investigate the detailed distribution of convective heat transfer coefficients on the first-stage blade tip surface for a geometry typical of large power generation turbines (> 100 MW). This paper is concerned with the design and execution of the experimental portion of the study, which represents the first reported investigation to obtain nearly full surface information on heat transfer coefficients within an environment that develops an appropriate pressure distribution about an airfoil blade tip and shroud model. A stationary blade cascade experiment has been run consisting of three airfoils, the center airfoil having a variable tip gap clearance. The airfoil models the aerodynamic tip section of a high-pressure turbine blade with inlet Mach number of 0.30, exit Mach number of 0.75, pressure ratio of 1.45, exit Reynolds number based on axial chord of 2.57 x 10{sup 6}, and total turning of about 110 degrees. A hue detection based liquid crystal method is used to obtain the detailed heat transfer coefficient distribution on the blade tip surface for flat, smooth tip surfaces with both sharp and rounded edges. The cascade inlet turbulence intensity level took on values of either 5 or 9%.more » The cascade also models the casing recess in the shroud surface ahead of the blade. Experimental results are shown for the pressure distribution measurements on the airfoil near the tip gap, on the blade tip surface, and on the opposite shroud surface. Tip surface heat transfer coefficient distributions are shown for sharp edge and rounded edge tip geometries at each of the inlet turbulence intensity levels.« less

A Finite Element Method Study of Eulerian Droplets Impingement Models

Abstract: To compute droplet impingement on airfoils, an Eulerian model for air flows containing water droplets is proposed as an alternative to the traditional Lagrangian particle tracking approach. Appropriate boundary conditions are presented for the droplets equations, with a stability analysis of the solution near the airfoil surface. Several finite element formulations are proposed to solve the droplets equations, based on conservative and non-conservative forms and using different stabilization terms. Numerical results on single and multi-elements airfoils for droplets of mean volume diameter, as well as for a Langmuir distribution of diameters, are presented and validated against measured values

Airfoil and Wing Design Through Hybrid Optimization Strategies

Abstract: Real-world design problems need robust and effective system-level optimization tools inasmuch as they are ruled by several criteria, most often in multidisciplinary environments. In this work a hybrid optimization algorithm has been obtained by adding a gradient-based technique to the set of operators of a multiobjective genetic algorithm. This makes it possible to increase the computational efficiency of the genetic algorithm while preserving its favorable features of robustness, problem independence, and multiobjective optimization capabilities. Aerodynamic shape design problems, including both airfoil and wing designs, are considered

Oscillatory Excitation of Unsteady Compressible Flows over Airfoils at Flight Reynolds Numbers

Abstract: An experimental investigation, aimed at delaying flow separation due to the occurrence of a shock-wave-boundary-layer interaction, is reported. The experiment was performed using a NACA 0012 airfoil and a NACA 0015 airfoil at high Reynolds number incompressible and compressible flow conditions. The effects of Mach and Reynolds numbers were identified, using the capabilities of the cryogenic-pressurized facility to maintain one parameter fixed and change the other. Significant Reynolds number effects were identified in the baseline compressible flow conditions even at Reynolds number of 10 and 20 million. The main objectives of the experiment were to study the effects of periodic excitation on airfoil drag-divergence and to alleviate the severe unsteadiness associated with shock-induced separation (known as "buffeting"). Zero-mass-flux oscillatory blowing was introduced through a downstream directed slot located at 10% chord on the upper surface of the NACA 0015 airfoil. The effective frequencies generated 2-4 vortices over the separated region, regardless of the Mach number. Even though the excitation was introduced upstream of the shock-wave, due to experimental limitations, it had pronounced effects downstream of it. Wake deficit (associated with drag) and unsteadiness (associated with buffeting) were significantly reduced. The spectral content of the wake pressure fluctuations indicates of steadier flow throughout the frequency range when excitation was applied. This is especially important at low frequencies which are more likely to interact with the airframe.

Reduced-order modelling of unsteady small-disturbance flows using a frequency-domain proper orthogonal decomposition technique

Abstract: A new method for constructing reduced-order models (ROM) of unsteady small-disturbance flows is presented. The reduced-order models are constructed using basis vectors determined from the proper orthogonal decomposition (POD) of an ensemble of small-disturbance frequencydomain solutions. Each of the individual frequency-domain solutions is computed using an efficient time-linearized flow solver. We show that reduced-order models can be constructed using just a handful of POD basis vectors, producing low-order but highly accurate models of the unsteady flow over a wide range of frequencies. In this paper, we apply the POD/ROM technique to compute the unsteady aerodynamic and aeroelastic behavior of an isolated nansonic airfoil, and to a two-dimensional cascade of airfoils. Nomenclature A = matrix defining homogeneous part of discretized aerodynamic operator A = reduced-order form of A. b = airfoil semi-chord b = vector defining inhomogeneous part of discretized aerodynamic operator Ba, B1 = matrices relating airfoil motion h and h to b c c

Coolable airfoil structure

Abstract: A coolable airfoil structure having internal heat transfer features through which cooling air is flowed under operative conditions is disclosed. Various construction details and features are developed which affect the castability and core strength during manufacture and strength and cooling effectiveness of the airfoil after manufacture. In one particular embodiment, the airfoil has a plurality of heat transfer members disposed in the rearmost section of the trailing edge region which comprises a single impingement rib, a single row of pedestals and a single row of chordwisely extending flow dividers.

Twin rib turbine blade

Abstract: A turbine blade includes an airfoil (24) and an integral dovetail. The airfoil includes first and second sidewalls (28,30) joined together at leasing and trailing edges (32,34), and extending from a root to a tip plate 48. Twin tip ribs (50,52) extend outwardly from the tip plate between the leading and trailing edges, and are spaced laterally apart to define an open-top tip channel (54) therebetween. Each of the tip ribs (50,52) has an airfoil profile for extracting energy from combustion gases flowable around the turbine blade.

Limit Cycle Oscillations of a Cantilevered Wing in Low Subsonic Flow

Abstract: A nonlinear, aeroelastic analysis of a low-aspect, rectangular wing modeled as a plate of constant thickness demonstratesthatlimitcycleoscillationsoftheorderoftheplatethicknessarepossible.Thestructuralnonlinearity arisesfrom doublebendinginboth thechordwiseand spanwisedirections.Thepresentresultsusing a vortex lattice aerodynamic model for low-Mach-numbere owscomplementearlierstudiesforhigh supersonicspeed thatshowed similar qualitative results. Also, the theoretical results are consistent with experimental data reported by other investigators for low-aspect-ratio delta wings.

Rigid and Flexible Low Reynolds Number Airfoils

Abstract: Issues related to the design of low Reynolds number airfoils, such as the thickness, camber, and surface proe les, are investigated. To contrast the issues involved, NACA 0012 and CLARK-Y, two well-known airfoils, a recently proposed low Reynolds number airfoil S1223, and a modie ed airfoil UF, are compared under varied Reynolds numbers and angles of attack. These airfoilsrange from 0% (NACA 0012)to 8.89%(S1223)camber, and from 6% (UF)to12.9%(CLARK-Y)thickness,and allowusto makea broadcomparison of thelift-and-drag characteristics with varying Reynolds numbers, from 7.5 £ 104 to 2.0 £ 106. Furthermore, the concept of a e exible airfoil is assessed in an unsteady, low Reynolds number environment. To facilitate the present study, we have employed techniques treating either inviscid or coupled inviscid/boundary-layer e ows around rigid airfoils, as well as a moving boundary technique to handle an elastic, massless membrane in a portion of the upper airfoil surface. The results show that within the range of Reynolds numbers and airfoil shapes, increased camber and reduced thickness provide more favorable lift-and-drag characteristics when the Reynolds number becomes lower. The results also indicate that a e exible proe le yields better overall performance than a similar rigid proe le in an oscillating freestream.

Reduced-Order Models Based on Linear and Nonlinear Aerodynamic Impulse Responses

Abstract: This paper discusses a method for the identification and application of reduced-order models based on linear and nonlinear aerodynamic impulse responses. The Volterra theory of nonlinear systems and an appropriate kernel identification technique are described. Insight into the nature of kernels is provided by applying the method to the nonlinear Riccati equation in a non-aerodynamic application. The method is then applied to a nonlinear aerodynamic model of an RAE 2822 supercritical airfoil undergoing plunge motions using the CFL3D Navier-Stokes flow solver with the Spalart-Allmaras turbulence model. Results demonstrate the computational efficiency of the technique.

Hopf-Bifurcation Analysis of Airfoil Flutter at Transonic Speeds

Abstract: Hopf-bifurcation analysis is used to determine flutter onset for a pitch-and-plunge airfoil at transonic Mach number conditions. The pitch-and-plunge model is a coupling of the Euler equations and a twodegree-of-freedom structural model composed of linear and torsional springs. The Euler equations are discretized using the upwind total variation diminishing scheme of Harten and Yee. Equilibrium solutions of the aeroelastic model are computed using Newton's method, and dynamic solutions are explicitly integrated in time with first-order accuracy. The Hopf-bifurcation point, which models the flutter condition, is computed directly using a modified form of the Griewank and Reddien algorithm. A path of Hopf points is computed as a function of Mach number to produce a Mach flutter boundary. The flutter boundary is validated by time integration. Flutter boundaries are also obtained through variation of static pretwist and pitch and plunge damping. The direct, Hopf-point method is found to be precise and efficient for grids typical of inviscid, transonic airfoil calculations.

Method using laser shock peening to process airfoil weld repairs pertaining to blade cut and weld techniques

Abstract: A method is disclosed for repairing damage to an airfoil. The method provides for the removal of a section of the airfoil that substantially encompasses the damaged area, which consequently leaves a void and a cut-away surface in the airfoil. A replacement piece larger than the residual void is provided for use in replacing the section removed from the airfoil. A joining operation welds or otherwise joins the replacement piece to the airfoil at the cut-away surface to form a joined airfoil. The joined airfoil has a seam between the airfoil and the replacement piece. At least a portion of the seam is processed by laser shock peening to induce compressive residual stresses therein.

Development of Advanced Compressor Airfoils for Heavy-Duty Gas Turbines: Part I — Design and Optimization

Abstract: A new family of subsonic compressor airfoils, which are characterized by low losses and wide operating ranges, has been designed for use in heavy-duty gas turbines. In particular the influence of the higher airfoil Reynolds numbers compared to aeroengine compressors and the impact of these differences on the location of transition are taken into account. The design process itself is carried out by the combination of a geometric code for the airfoil description, with a blade-to-blade solver and a numerical optimization algorithm. The optimization process includes the design-point losses for a specified Q3D flow problem and the off-design performance for the entire operating range. The family covers a wide range of inlet flow angle, Mach number, flow turning, blade thickness, solidity and AVDR in order to consider the entire range of flow conditions that occur in practical compressor design. The superior performance of the new airfoil family is demonstrated by a comparison with conventional controlled diffusion airfoils (CDA). The advantage in performance has been confirmed by detailed experimental investigations, which will be presented in Part II of the paper. This leads to the conclusion that CDA airfoils that have been primarily developed for aeroengine applications are not the optimum solution, if directly transferred to heavy-duty gas turbines. A significant improvement in compressor efficiency is possible, if the new profiles are used instead of conventional airfoils. @S0889-504X~00!02102-4#

Airborne vehicle having deployable wing and control surface

Abstract: An airborne vehicle having a deployable airfoil with an elevon wherein the deployment of the airfoil and the control of the elevon are both powered by a single servo mechanism. A shear pin prevents relative movement between the elevon and the airfoil until the airfoil is in the deployed position. A stop mechanism locks the airfoil in the deployed position, whereafter operation of the drive mechanism fractures the shear pin, thereby allowing the elevon to be controlled by the drive mechanism.

Experimental Investigation of Simulated Large-Droplet Ice Shapes on Airfoil Aerodynamics

Abstract: An experimental investigation was conducted to study the aerodynamic effect of simulated supercooled largedroplet ice accretion on a modie ed NACA 23012 airfoil. Forward-facing quarter-round simulations with heightto-chord ratios of 0.0083 and 0.0139 were used at a Reynolds number of 1 :8 £ 10 6 . When the simulated ice was placed at critical chordwise locations, a long separation bubbleformed downstream of thesimulated ice shape and effectively eliminated the formation of a large leading-edge suction peak that was observed on the clean NACA 23012 airfoil. This resulted in a dramatic reduction in the maximum lift coefe cient, as low as 0.27, when the larger simulated ice shape was located at 12% chord. Because the airfoil loading distribution was severely altered, large changes in airfoil pitching moments and e ap-hinge moments were also observed.

Feasibility Study of e Transition Prediction in Navier-Stokes Methods for Airfoils

Abstract: When Navier-Stokes methods are applied in the design process for laminar airfoils, the prediction of the transition location still represents an unresolved problem. By the acceptence of the E N method as representing a convenient transition prediction tool, the requirements for coupling E N to Navier-Stokes methods will be demonstrated. In particular, the possibility to determine the laminar and turbulent viscous length scales is outlined. Based on the knowledge of the viscous layer thickness, a mesh adaption procedure can be applied. The Navier-Stokes results produced using adapted meshes are shown to be Identical to boundary-layer results, when the computed wall pressure from the Navier-Stokes solution is used as input for the boundary-layer calculation. For the validation of the necessary steps in the coupling procedure, the laminar airfoil DoAL3 was selected. This airfoil was measured in the Transonic Wind Tunnel Braunschweig facility at the DLR, German Aerospace Research Establishment. The limiting N factor for that wind tunnel was determined beforehand