Showing papers on "Airfoil published in 1992"

Multipoint inverse airfoil design method based on conformal mapping

Abstract: An exact method of multipoint inverse airfoil design for incompressible flow is presented. Multipoint design is handled by dividing the airfoil into a number of desired segments. For each segment the velocity distribution is prescribed together with an angle of attack at which the prescribed velocity distribution is to be achieved. In this manner, multipoint design objectives can be taken into account in the initial specification of the velocity distribution. In order for the multipoint inverse airfoil design problem to be well posed, three integral constraints and several conditions arise which must be satisfied. Further restrictions are imposed if the airfoil is to have a specified pitching moment, thickness ratio, or other constraints. The system of equations is solved partly as a linear system and partly through multidimensional Newton iteration. Since the velocity distribution is prescribed about the circle in the angular coordinate, specification of the velocity in terms of arc length is handled through the multidimensional Newton iteration as well. The current formulation sets the stage for a more general multipoint inverse airfoil design method in which it will be possible to specify the velocity distribution, some boundary-layer development, or the surface geometry along a segment.

Forcing level effects of internal acoustic excitation on the improvement of airfoil performance

Abstract: The effects of internal acoustic excitation on the leading-edge, separated boundary layers and the aerodynamic performance of NACA 633-018 cross section airfoil are examined as a function of forcing level and forcing frequency of the introduced acoustics. Tests are separately conducted in two suction, open-typed wind tunnels at the Reynolds number of 3.0 x 10 s for the measurements and 1.0 x 10 4 for the visualization. Results indicate that the flow separation is delayed at the angles of attack higher than the stalled angle of small level excitation with the forcing frequency fe near the shear layer instability frequency ft. As the forcing level is increased to some extent, the velocity fluctuations around the slot exit are demonstrated to be the primary governing parameter for modifying the separated boundary layers. Data also show that the effective forcing frequency (and the Strouhal number, 50 extends over wider range as compared to the lower level excitation. Meanwhile, the pressure distributions on the airfoil surface exhibit recovery behaviors with different forcing frequencies. The corresponding boundary layers are visualized to be reattached to the surface to form a recirculation region when the airfoil is around at the stalled angles.

Unsteady flow past an airfoil pitching at a constant rate

Abstract: The unsteady flow past a NACA 0012 airfoil that is undertaking a constant-rate pitching up motion is investigated experimentally by the PIDV technique in a water towing tank. The Reynolds number is 5000, based upon the airfoil's chord and the free-stream velocity. The airfoil is pitching impulsively from 0 to 30 deg. with a dimensionless pitch rate alpha of 0.131. Instantaneous velocity and associated vorticity data have been acquired over the entire flow field. The primary vortex dominates the flow behavior after it separates from the leading edge of the airfoil. Complete stall emerges after this vortex detaches from the airfoil and triggers the shedding of a counter-rotating vortex near the trailing edge. A parallel computational study using the discrete vortex, random walk approximation has also been conducted. In general, the computational results agree very well with the experiment.

An Experimental Study of Heat Transfer in a Large-Scale Turbine Rotor Passage

Abstract: An experimental study of the heat transfer distribution in a turbine rotor passage was conducted in a large-scale, ambient temperature, rotating turbine model. Heat transfer was measured for both the full-span suction and pressure surfaces of the airfoil as well as for the hub endwall surface. The objective of this program was to document the effects of flow three-dimensionality on the heat transfer in a rotating blade row (vs a stationary cascade). Of particular interest were the effects of the hub and tip secondary flows, tip leakage and the leading-edge horseshoe vortex system. The effect of surface roughness on the passage heat transfer was also investigated. Midspan results are compared with both smooth-wall and rough-wall finite-difference two-dimensional heat transfer predictions. Contour maps of Stanton number for both the rotor airfoil and endwall surfaces revealed numerous regions of high heat transfer produced by the three-dimensional flows within the rotor passage. Of particular importance are regions of local enhancement (as much as 100 percent over midspan values) produced on the airfoil suction surface by the secondary flows and tip-leakage vortices and on the hub endwall by the leading-edge horseshoe vortex system.

Calculation of Three-Dimensional Unsteady Flows in Turbomachinery Using the Linearized Harmonic Euler Equations

Abstract: An efficient three-dimensional Euler analysis of unsteady flows in turbomachinery is presented. The unsteady flow is modeled as the sun of a steady or mean flow field plus a harmonically varying small perturbation flow. The linearized Euler equations, which describe the small perturbation unsteady flow, are found to be linear, variable coefficient differential equations whose coefficients depend on the mean flow. A pseudo-time time-marching finite-volume Lax-Wendroff scheme is used to discretize and solve the linearized equations for the unknown perturbation flow quantities. Local time stepping and multiple-grid acceleration techniques are used to speed convergence. For unsteady flow problems involving blade motion, a harmonically deforming computational grid, which conforms to the motion of the vibrating blades, is used to eliminate large error-producing extrapolation terms that would otherwise appear in the airfoil surface boundary conditions and in the evaluation of the unsteady surface pressure. Results are presented for both linear and annular cascade geometries, and for the latter, both rotating and nonrotating blade row.

Flutter and Stall Response of a Helicopter Blade with Structural Nonlinearity

Abstract: The purpose of the present paper is to study the flutter instability and forced response of a nonrotating helicopter blade model with a NACA-0012 airfoil and a pitch freeplay structural nonlinearity. In this paper, three typical combinations of linear and nonlinear structure with a linear and nonlinear (ONERA) aerodynamic model are considered. Characteristic results are used to display the limit cycle oscillation and chaotic behavior of both the flutter instability and forced response for all three cases. The effects of various initial disturbance amplitudes on the forced response behavior are discussed. Comparisons of the results for the three cases are helpful in understanding physically the nonlinear aeroelasticity phenomena and chaotic oscillations.

NACA 0012 benchmark model experimental flutter results with unsteady pressure distributions

Abstract: The Structural Dynamics Division at NASA Langley Research Center has started a wind tunnel activity referred to as the Benchmark Models Program. The primary objective of this program is to acquire measured dynamic instability and corresponding pressure data that will be useful for developing and evaluating aeroelastic type computational fluid dynamics codes currently in use or under development. The program is a multi-year activity that will involve testing of several different models to investigate various aeroelastic phenomena. This paper describes results obtained from a second wind tunnel test of the first model in the Benchmark Models Program. This first model consisted of a rigid semispan wing having a rectangular planform and a NACA 0012 airfoil shape which was mounted on a flexible two degree of freedom mount system. Experimental flutter boundaries and corresponding unsteady pressure distribution data acquired over two model chords located at the 60 and 95 percent span stations are presented.

Tip cooled blade

Abstract: A gas turbine engine blade includes an airfoil having first and second sides and a dovetail extending from the airfoil root. The airfoil includes a tip having a tip floor with first and second tip walls extending from the floor and spaced apart to define therebetween a tip pienum. The first tip wall is recessed at least in part from the airfoil first side to define an outwardly facing tip shelf, with the tip shelf and the first tip wall defining therebetween a trough. A plurality of cooling holes extend through the tip floor at the tip shelf for channeling cooling air from a flow channel inside the airfoil into the trough for cooling the blade tip.

Unsteady Aerodynamics and Gust Response in Compressors and Turbines

Abstract: A comprehensive series of experiments and analyses was performed on compressor and turbine blading to evaluate the ability of current, practical, engineering/analysis models to predict unsteady aerodynamic loading of modern gas turbine blading. This is part of an ongoing effort to improve methods for preventing blading failure. The experiments were conducted in low-speed research facilities capable of simulating the relevant aerodynamic features of turbomachinery. Unsteady loading on compressor and turbine blading was generated by upstream wakes and, additionally for compressors, by a rotating inlet distortion. Fast-response hot-wire anemometry and pressure transducers embedded in the airfoil surfaces were used to determine the aerodynamic gusts and resulting unsteady pressure responses acting on the airfoils. This is the first time that gust response measurements for turbines have been reported in the literature. Several different analyses were used to predict the unsteady component of the blade loading: (1) a classical flat-plate analysis, (2) a two-dimensional linearized flow analysis with a frozen gust model, (3) a two-dimensional linearized flow analysis with a distorted gust model, (4) a two-dimensional linearized Euler analysis, and (5) a two-dimensional nonlinear Euler analysis. Also for the first time, a detailed comparison of these analyses methods is made and the importancemore » of properly accounting for both vortical and potential disturbances is demonstrated. The predictions are compared with experiment and their abilities assessed to help guide designers in using these prediction schemes.« less

Unsteady pressure field and vorticity production over a pitching airfoil

Abstract: The unsteady pressure field and the accompanying variations in the flux of spanwise vorticity from the surface were measured over a range of dimensionless pitch rates for a two-dimensional NACA 0012 airfoil model undergoing a single pitch-up motion. The results were examined to identify the mechanisms that play key roles in the initiation, development, growth, and movement of the dynamic-stall vortex. The unsteady pressure distribution over the airfoil was dominated by three features, whose emergence and evolution were used to distinguish between two classes of behavior, corresponding to low and high pitch rates. Further, it was found that the flux of vorticity from the surface originated primarily from five concentrated regions or sources, the majority of which were located over the forward portion of the airfoil surface. The behavior of vorticity flux from these sources was related to the interacting mechanisms responsible for the development of the flowfield. The change of these features in the pressure and surface vorticity flux variations with the pitch rate is described.

Generalized multipoint inverse airfoil design

Abstract: In a rather general sense, inverse airfoil design can be taken to mean the problem of specifying a desired set of airfoil characteristics, such as the airfoil maximum thickness ratio, pitching moment, part of the velocity distribution, or boundary-layer development. From thie information, the corresponding airfoil shape is determined. We present a method that approaches the design problem from this perspective. In particular, the airfoil is divided into segments along which, together with the design conditions, either the velocity distribution or boundary-layer development may be prescribed

Effect of acoustic excitation on stalled flows over an airfoil

Abstract: The effect of acoustic excitation on poststalled flows over an airfoil, ie, flows that are fully separated from near the leading edge, is investigated The excitation results in a tendency toward reattachment, which is accompanied by an increased lift and reduced drag, although the flow may still remain fully separated It is found that with increasing excitation amplitude, the effect becomes more pronounced but shifts to a Strouhal number which is much lower than that expected from linear, inviscid instability of the separated shear layer

Shroud geometry for unmanned aerial-vehicles

Abstract: An unmanned aerial vehicle (10) having a toroidal fuselage (20) that surrounds a pair of coaxial, multibladed, counter-ro-tating rotors (40). The toroidal fuselage (20) has an airfoil profile that is optimized to provide high hover efficiency and produce a pressure distribution that provide high lift forces. The airfoil profile is further optimized to counteract the undesirable nose-up pitching moments experienced by ducted rotary-type UAVs in forward translational flight.

Computational evaluation of an airfoil with a Gurney flap

Abstract: A 2D numerical investigation was performed to determine the effect of a Gurney flap on a NACA 4412 airfoil. A Gurney flap is a flat plate on the order of 1 to 3 percent of the airfoil chord length, oriented perpendicular to the airfoil chord line and located at the trailing edge of the airfoil. An incompressible Navier Stokes code, INS2D, was used to calculate the flow field about the airfoil. The fully turbulent results were obtained using the Baldwin-Barth one-equation turbulence model. Gurney flap sizes of 0.5 , 1, 1.25, 1.5, 2, and 3 percent of the airfoil chord were studied. Computational results were compared with experimental results where possible. The numerical solutions show that the Gurney flap increases airfoil lift coefficient with only a slight increase in drag coefficient. Use of a 1.5 percent chord Gurney flap increases the maximum lift coefficient by approximately 0.3 and decreases the angle of attack for a given lift coefficient by more than 3 deg. The numerical solutions exhibit detailed flow structures at the trialing edge and provide a possible explanation for the increased aerodynamic performance.

Results of an icing test on a NACA 0012 airfoil in the NASA Lewis Icing Research Tunnel

Abstract: Tests were conducted in the Icing Research Tunnel (IRT) at the NASA Lewis Research Center to document the current capability of the IRT, focused mainly on the repeatability of the ice shape over a range of icing conditions. Measurements of drag increase due to the ice accretion were also made to document the repeatability of drag. Surface temperatures of the model were obtained to show the effects of latent-heat release by the freezing droplets and heat transfer through the ice layer. The repeatability of the ice shape was very good at low temperatures, but only fair at near freezing temperatures. In general, drag data shows good repeatability.

Design and processing method for manufacturing hollow airfoils

Abstract: A method for forming a hollow, internally-ribbed airfoil having a complex geometry. The process forms airfoil halves from alloy members in essentially final machined form. The airfoil halves optionally include a gasket around their periphery and optionally, a gasket around the internal sequence of ribs and cavities. Because the airfoil halves are in final form before joining, important interior airfoil dimensions such as skin thickness, cavity size, rib location and gasket location as well as exterior dimensions such as camber and twist may be verified prior to joining. The airfoil halves are then joined by welding, thereby sealing their interiors and diffusion bonding at relatively low pressures. The gaskets eliminate the use of high pressures and associated skin buckling and permit a concentration of the loads at key locations despite the low pressures. After diffusion bonding, any voids are removed by hot isostatically pressing the airfoil in the superplastic temperature range of the alloy.

NREL Combined Experimental Final Report--Phase II

Abstract: Predicting peak power and loads on a fixed-pitch wind turbine. How does the performance of the airfoil in the wind tunnel differ from the performance of an operating horizontal-axis wind turbine (HAWT)?

Time domain flutter analysis of cascades using a full-potential solver

Abstract: A time domain approach is used to determine the aeroelastic stability of a cascade of blades. The structural model for each blade is a typical section with two degrees of freedom. The aerodynamic model is unsteady, two-dimensional, full-potential flow through the cascade of airfoils. The unsteady equations of motion for the structure and the fluid are integrated simultaneously in time starting with a steady flowfield and a small initial disturbance applied to the airfoils. Each blade is allowed to move independently, and the motion of all blades is analyzed to determine the aeroelastic stability of the cascade

Experimental Study of a Low Reynolds Number Tandem Airfoil Configuration

Abstract: Experiments were concluded with two identical Wortmann FX63-137 airfoils in closely coupled tandem configurations at a Reynolds number of 8.5 x 10 4. For the data presented here, the values of the stagger and gap were 1.5 and 0, respectively. The decalage angles were 0 and ±10 deg. Direct measurement of lift, drag, and 1/4-chord pitching moment, as well as static pressure distributions, were acquired for each airfoil. Flow visualization using kerosene smoke was performed to complement the experimental data. The total drag reduction and lift increase resulted in a significant increase in the lift-to-drag ratio for a number of configurations. Nomenclature Aw = wing area Cd = section drag coefficient, D/(Awqx) Ci = section lift coefficient, L/(AwqJ) Cm = section 1/4-chord moment coefficient, Cp = pressure coefficient, l-q/qx c = chord length D = drag force G = gap, \y\lc L = lift force M = 1/4-chord pitching moment q = dynamic pressure, \pU2 Rc = Reynolds number based on chord, Uxc/v St = stagger, x/c (positive when upstream airfoil is above the downstream airfoil) U = flow velocity x = distance in streamwise direction y - distance normal to streamwise directions a = angle of attack 8 = decalage, aua - ada

Experimental and computational ice shapes and resulting drag increase for a NACA 0012 airfoil

Abstract: Tests were conducted in the Icing Research Tunnel (IRT) at LeRC to document the repeatability of the ice shape over the range of temperatures varying from -15 to 28 F. Measurements of drag increase due to the ice accretion were also made. The ice shape and drag coefficient data, with varying total temperatures at two different airspeeds, were compared with the computational predictions. The calculations were made with the 2D LEWICE/IBL code which is a combined code of LEWICE and the interactive boundary layer method developed for iced airfoils. Comparisons show good agreement with the experimental data in ice shapes. The calculations show the ability of the code to predict drag increases as the ice shape changes from a rime shape to a glaze shape.

Ultra-soft, ultra-elastic airfoils

Abstract: A novel rotating aerodynamic toy comprising a ultra-elastic gel airfoil which is suitable for launch in light or heavy wind conditions and capable of performing various aerodynamic effects including climb, stall, return, straight-line flight and other aerobatics. The ultra-elastic properties of the airfoil allow the airfoil to transform its aerodynamic profile at launch and while in flight.

Design and development of a compressible dynamic stall facility

Abstract: A dynamic stall facility offering a unique new capability for studies of compressibility effects on dynamic stall is described. This facility features complete visual access by mounting the test airfoil between optical-quality glass windows which are rotated in unison to produce the oscillating airfoil motion associated with helicopter rotor dynamic stall. By using the density gradients associated with the rapidly changing dynamic stall flow field, this facility permits simultaneous detailed investigation of the flow on the surface as well as in the flow field surrounding airfoils experiencing dynamic stall.

Turbine blade trailing edge cooling construction

Abstract: Internal cooling of the trailing edge of the airfoil of the turbine blade for a gas turbine engine includes a cascade formed from juxtaposed rows of longitudinally extending spaced airfoil shaped vanes or ribs leading cooling air from a supply source through the space between adjacent vanes and discharging out of the blade.

Skin for a deicer

Abstract: An aircraft deicing assembly 101 for attachment to an airfoil 99 includes a deflection shell 101 disposed over a deflection means 102. The deflection shell is comprised of a primary layer 116 comprised of a high strength fabric reinforced with a phenolic resin and a backing layer 115 comprised of a high strength fabric reinforced with an epoxy resin. The thickness of the backing layer is varied as a function of curvature of the airfoil and spanwise length of the airfoil.

Simplified Linear Stability Transition Prediction Method for Separated Boundary Layers

Abstract: An existing transition prediction method for attached, two-dimensional, incompressible boundary layers based on linear stability analysis is extended to separated, two-dimensional, incompressible boundary layers such as those found in laminar (transitional) separation bubbles. It is shown why the present method, which tracks the growth of disturbances at many different frequencies, is more accurate than the so-called envelope methods for nonsimilar boundary-layer developments. Reliance on a database of precalculated stability characteristics of known velocity profiles makes this method much faster than traditional stability calculations of similar accuracy. The Falkner-Skan self-similar profiles are used for attached flow, and a new, very general family of profiles is used for separated flow. Comparisons with measured transition locations inside the bubble show good agreement over the range of chord Reynolds numbers and airfoil angles of attack of interest.

A fast implicit upwind solution algorithm for three-dimensional unstructured dynamic meshes

Abstract: A fast implicit upwind algorithm for the solution of the time-dependent Euler equations is presented for aerodynamic analysis involving unstructured dynamic meshes. The spatial discretization of the scheme is based on the upwind approach of Roe, referred to as flux-difference splitting (FDS). The FDS approach is naturally dissipative and captures shock waves and contact discontinuities sharply. The temporal discretization of the scheme involves an implicit time-integration using a two-sweep Gauss-Seidel relaxation procedure. The procedure is computationally efficient for either steady or unsteady flow problems. A detailed description is given of the implicit upwind solution algorithm along with results which assess the capability. The results are presented for the NACA 0012 airfoil and for the Boeing 747 aircraft. The 747 geometry includes the fuselage, wing, horizontal and vertical tails, under-wing pylons, and flow-through engine nacelles. Euler solutions for the 747 aircraft on an unstructured tetrahedral mesh containing approximately 100,000 cells were obtained to engineering accuracy in less than one hour CPU time on a Cray-2 computer.