Showing papers on "Airfoil published in 1987"

Viscous-inviscid analysis of transonic and low Reynolds number airfoils

Abstract: A method of accurately calculating transonic and low Reynolds number airfoil flows, implemented in the viscous-inviscid design/analysis code ISES, is presented. The Euler equations are discretized on a conservative streamline grid and are strongly coupled to a two-equation integral boundary-layer formulation, using the displacement thickness concept. A transition prediction formulation of the e type is derived and incorporated into the viscous formulation. The entire discrete equation set, including the viscous and transition formulations, is solved as a fully coupled nonlinear system by a global Newton method. This is a rapid and reliable method for dealing with strong viscous-inviscid interactions, which invariably occur in transonic and low Reynolds number airfoil flows. The results presented demonstrate the ability of the ISES code to predict transitioning separation bubbles and their associated losses. The rapid airfoil performance degradation with decreasing Reynolds number is thus accurately predicted. Also presented is a transonic airfoil calculation involving shock-induced separation, showing the robustness of the global Newton solution procedure. Good agreement with experiment is obtained, further demonstrating the performance of the present integral boundary-layer formulation.

Vortical patterns in the wake of an oscillating airfoil

Abstract: The vortical flow patterns in the wake of a NACA 0012 airfoil pitching at small amplitudes are studied in a low speed water channel. it is shown that a great deal of control can be exercised on the structure of the wake by the control of the frequency, amplitude and also the shape of the oscillation waveform. An important observation in this study has been the existence of an axial flow along the cores of the wake vortices. Estimates of the magnitude of the axial flow suggest a linear dependence on the oscillation frequency and amplitude.

Navier-Stokes Computations of Vortical Flows Over Low-Aspect-Ratio Wings

Abstract: An upwind-biased finite-volume algorithm is applied to the low-speed flow over a low aspect ratio delta wing from zero to forty degrees angle of attack. The differencing is second-order accurate spatially, and a multigrid algorithm is used to promote convergence to the steady state. The results compare well with the detailed experiments of Hummel (1983) and others for a Re(L) of 0.95 x 10 to the 6th. The predicted maximum lift coefficient of 1.10 at thirty-five degrees angle of attack agrees closely with the measured maximum lift of 1.06 at thirty-three degrees. At forty degrees angle of attack, a bubble type of vortex breakdown is evident in the computations, extending from 0.6 of the root chord to just downstream of the trailing edge.

Laminar separation bubble characteristics on an airfoil at low Reynolds numbers

Abstract: An experimental investigation was conducted in order to document the structure and behavior of laminar separation bubbles at low Reynolds numbers. Data of this type are necessary if the currently insufficient analytical and numerical models are to be improved. The laminar separation bubble that forms on a NACA 663-018 airfoil model was surveyed at chord Reynolds numbers of 50,000-200,000 at angles of attack of 8-12 deg. The effects of the various testing conditions on the separation bubble were isolated and the data were analyzed in relation to existing separation bubble correlations in order to test their low Reynolds number applicability. This analysis indicated that the chord Reynolds number and the disturbance environment strongly influence the experimental pressure distributions. These effects must be included in any analytic prediction technique applied to the low Reynolds number flight regime.

Navier-Stokes Simulations of Rotor/Stator Interaction Using Patched and Overlaid Grids

Abstract: An accurate numerical analysis of the flows associated with rotor/stator configurations in turbomachinery can be very helpful in optimizing performance. In this study, the unsteady, thin-layer, Navier-Stokes equations are solved using a system of patched and overlaid grids for a rotor/stator configuration of an axial turbine. The equations necessary for an accurate transfer of information between the several grids are briefly described within the framework of an iterative, implicit algorithm. Results in the form of Mach number contours, time-averaged pressures, unsteady pressures, amplitudes, and phase are presented. The numerical results are also compared with experimental data and the agreement is found to be good. HE aerodynamic processes associated with the flow of fluid through turbomachines pose one of the toughest challenges to the computational fluid dynamicist. The un- steady nature of the flow, the complex geometries involved, the motion of some parts of the system relative to others, and the periodic transition of the flow from laminar to tur- bulent are some of the factors that contribute to the com- plexity of the problem. A clear understanding of these types of flows is essential for the optimization of the performance of turbomachiner y. In this study, an attempt is made to analyze two-dimensional flow past the rotor/stator con- figuration of an axial turbine using state-of-the-art computa- tional tools and computing facilities. The two-dimensional analysis of stator airfoils in isolation or rotor airfoils in isolation is a relatively straightforward task. Such an analysis is valid when the two rows of blades are set far enough apart so that the interaction effects are minimal. However, the desire to minimize engine length re- quires the stator and rotor airfoils to be closely spaced. Clearly, the interaction effects will become more important as the axial gap between the rows is reduced. In fact, the flow becomes periodically unsteady for small values of the axial gap. The experimental results of Ref. 1 show that the temporal pressure fluctuation (the difference between the minimum and maximum pressure values) near the leading edge of the rotor can be as much as 12% of the exit dynamic pressure when the axial gap is reduced to 15% of the chord length (for the operating conditions and geometry chosen). Hence, it is important that the rotor and stator airfoils be treated as a single system when the interaction effects become predominant. A computational tool that provides the design engineer with the necessary aerodynamic data can be used to great advantage in redesigning rotor and stator airfoils to enhance performance. Such a tool has to accurately simulate the unsteady flow associated with rotor/stator con- figurations exhibiting a strong interaction. A finite-differe nce solution to the Navier-Stokes equations requires the generation of a computational grid for the

Two-Dimensional Transonic Aerodynamic Design Method

Abstract: This paper demonstrates the capabilities of a new transonic, two-dimensional design method, based on the simultaneous solution of multiple streamtubes, coupled through the position of, and pressure at, the streamline interfaces. This allows the specification of either the airfoil shape (direct, analysis mode) or the airfoil surface pressure distribution (inverse, design mode). The nonlinear system of equations is formulated in a conservative manner, which guarantees the correct treatment of shocks, and is solved by a rapid Newton solution method. Viscous effects can also be included through a coupled integral boundary-layer analysis. The first set of results shows the effect of different far-field treatments, demonstrating the improvement in accuracy obtained by including the second-order doublet terms in addition to the usual first-order vortex term. The results are also compared to those obtained by specifying straight far-field streamlines (corresponding to solid-wall wind-tunnel experiments) or constant far-field pressure (corresponding to freejet experiments) to show the sensitivity to the farfield distance. In the second set of results, the design method is used to design a transonic airfoil with C/ = 1.000 at A/oo = 0.70. It is shown that the off-design performance is improved by specifying a surface pressure distribution with a very weak shock.

A critical assessment of wind tunnel results for the NACA 0012 airfoil

Abstract: A large body of experimental results, obtained in more than 40 wind tunnels on a single, well-known two-dimensional configuration, has been critically examined and correlated. An assessment of some of the possible sources of error has been made for each facility, and data which are suspect have been identified. It was found that no single experiment provided a complete set of reliable data, although one investigation stands out as superior in many respects. However, from the aggregate of data the representative properties of the NACA 0012 airfoil can be identified with reasonable confidence over wide ranges of Mach number, Reynolds number, and angles of attack. This synthesized information can now be used to assess and validate existing and future wind tunnel results and to evaluate advanced Computational Fluid Dynamics codes.

Effect of acoustic excitation on the flow over a low-Re airfoil

Abstract: Wind-tunnel measurements of lift, drag, and wake velocity spectra were carried out under (tonal) acoustic excitation for a smooth airfoil in the chord-Reynolds-number Re(c) range of 40,000-140,000. The data were supported by smoke-wire flow-visualization pictures. Small-amplitude excitation in a wide, low-frequency range is found to eliminate laminar separation that otherwise degrades the airfoil performance at low Re(c) near the design angle of attack. Excitation at high frequencies eliminates a prestall, periodic shedding of large-scale vortices. Significant improvement in lift is also achieved during poststall, but with large-amplitude excitation. Wind-tunnel resonances strongly influence the results, especially in cases requiring large amplitudes.

Experimental study of airfoil performance with vortex generators.

Abstract: The canard airfoil from the Voyager aircraft was tested in Ohio State University's subsonic wind tunnel. This highly optimized laminar flow section had good clean airfoil performance, but suffered severe lift and drag penalties with early boundary-layer transition. These performance penalties resulted from a midchord boundary-layer separation. An experimental program was conducted to document this problem and then to design and test vortex generators in order to improve the tripped airfoil performance while having the least effect on the clean airfoil. A set of properly designed vortex generators were found to increase the lift and reduce the drag of the contaminated airfoil. A brief study documented a significant drag rise due to a rough surface in the turbulent boundary-layer region.

Lift-curve characteristics for an airfoil pitching at constant rate

Abstract: This paper reports the results of wind-tunnel studies of dynamic stall for an NACA 0015 airfoil pitching about the midchord at a constant rate. Time-varying pressure readings from 16 locations on the airfoil were collected and used to determine the lift, pressure-drag, and moment coefficients as functions of angle of attack for 100 test cases, covering 20 dynamic airspeed/pitch rate combinations. The dynamic-stall effects of the change (from steady flow) in the angle of attack at which separation occurs at the quarter chord and the change in the angle of attack at which stall occurs were extracted from these data and found to collapse onto a nondimensional pitch rate given by the chord times the pitch rate divided by two times the freestream velocity. The results showed that relatively slow pitch rates had dramatic effects on both the delay of stall and the magnitude of the maximum lift coefficient. The nondimensional rate is a measure of the speed of the leading edge divided by the speed of the freestream; it was found that nondimensional rates of less than 0.03 more than doubled the maximum coefficient of lift. The reduced data also clearly indicate that quarter-chord separation is systematically linked to dynamic stall.

Successes and challenges in computational aerodynamics

Abstract: The purpose of this paper is to survey some of the highlights of computational fluid dynamics as an emerging branch of aeronautical science, and to identify several remaining unsurmounted challenges. Prior to the advent of the computer there was already in place a rather comprehensive mathematical formulation of fluid mechanics. This had been developed by elegant mathetical analysis, frequently guided by brilliant insights. Well known examples include the airfoil theory of Kutta and Joukowski, Prandtl's wing and boundary layer theories, von Karman's analysis of the vortex street, and more recently Jones' slender wing theory [I], and Hayes' theory of linearized supersonic flow [ Z ] . These methods required simplifying assumptions of various kinds, and could not be used to make quantitative predictions of complex flows dominated by nonlinear effects. The computer opens up new possibilities for attacking these problems by direct calculation of solutions to more complete mathematical models.

Computational methods for unsteady transonic flows

Abstract: Computational methods for unsteady transonic flows are surveyed with emphasis on prediction. Computational difficulty is discussed with respect to type of unsteady flow; attached, mixed (attached/separated) and separated. Significant early computations of shock motions, aileron buzz and periodic oscillations are discussed. The maturation of computational methods towards the capability of treating complete vehicles with reasonable computational resources is noted and a survey of recent comparisons with experimental results is compiled. The importance of mixed attached and separated flow modeling for aeroelastic analysis is discussed, and recent calculations of periodic aerodynamic oscillations for an 18 percent thick circular arc airfoil are given.

A hybrid algorithm for transonic airfoil and wing design

Abstract: The present method for the design of transonic airfoils and wings employs a predictor/corrector approach in which an analysis code calculates the flowfield for an initial geometry, then modifies it on the basis of the difference between calculated and target pressures. This allows the design method to be straightforwardly coupled with any existing analysis code, as presently undertaken with several two- and three-dimensional potential flow codes. The results obtained indicate that the method is robust and accurate, even in the cases of airfoils with strongly supercritical flow and shocks. The design codes are noted to require computational resources typical of current pure-inverse methods.

Separation control over an airfoil at high angles of attack by sound emanating from the surface

Abstract: Active control by sound emanating from a narrow gap in the vicinity of the leading edge of a symmetrical airfoil is used to study the influence of sound on the pressure distribution and the wake at high angles of attack The results from experiments conducted at a Reynolds number based on the chord of 35,000 show that, with injection of sound at twice the shedding frequency of the shear layer, the region of separation becomes drastically reduced The shear layer is found to be very sensitive to sound excitation in the vicinity of the separation point The excitation sufficiently alters the global circulation to cause an increase in lift and reduction in drag Furthermore, experimental results describing stall and post-stall conditions compare well with the limited data available and indicate that stall is delayed by sound injection into the separated region

Numerical simulation of viscous transonic airfoil flows

Abstract: Numerical simulations of transonic airfoil flows using the Reynolds-averaged Navier-Stokes equations and various turbulence models are presented and compared with experimental data. Three different airfoils were investigated under varying flow conditions ranging from subcritical unseparated flows to supercritical separated flows. The turbulence models investigated consisted of three zero-equation models and one two-equation model. For unseparated flows involving weak viscous-inviscid interactions, the four models were comparable in their agreement with experiment. For separated flows involving strong viscous-inviscid interactions, the nonequilibrium zero-equation model of Johnson and King gave the best overall agreement with experiment.

Viscous Transonic Airfoil Workshop compendium of results

Abstract: Results from the Viscous Transonic Airfoil Workshop held at the AIAA 25th Aerospace Sciences Meeting at Reno, NV in January 1987, are compared with each other and with experimental data. Test cases used in this workshop include attached and separated transonic flows for three different airfoils: the NACA 0012 airfoil, the RAE 2822 airfoil, and the Jones airfoil. A total of 23 sets of numerical results from 15 different author groups are included. The numerical methods used vary widely and include: 16 Navier-Stokes methods, 2 Euler/boundary-layer methods, and 5 full-potential/boundary-layer methods. The results indicate a high degree of sophistication among the numerical methods with generally good agreement between the various computed and experimental results for attached or moderately-separated cases. The agreement for cases with larger separation is only fair and suggests additional work is required in this area.

The effects of inlet turbulence and rotor/stator interactions on the aerodynamics and heat transfer of a large-scale rotating turbine model, volume 1

Abstract: A combined experimental and analytical program was conducted to examine the effects of inlet turbulence on airfoil heat transfer. Heat transfer measurements were obtained using low conductivity airfoils with miniature thermocouples welded to a thin, electrically heated surface skin. Heat transfer data were acquired for various combinations of low or high inlet turbulence intensity, flow coefficient (incidence), first-stator/rotor axial spacing, Reynolds number, and relative circumferential position of the first and second stators. Aerodynamic measurements include distributions of the mean and fluctuating velocities at the turbine inlet and, for each airfoil row, midspan airfoil surface pressures and circumferential distributions of the downstream steady state pressures and fluctuating velocities. Analytical results include airfoil heat transfer predictions and a examination of solutions of the unstead boundary layer equipment.

Experimental investigations of the circular wall jet on a circulation control airfoil

Abstract: The circular wall jet on a circulation control airfoil model's flowfield has been examined in detail using modern Laser Velocimetry techniques in a specially designed test program. Prime motivation for these studies was to aid in the develoment and correlation of Computational Fluid 'Jm A'Jm 9 Dynamics (CFD) models. The tests &re designed to u, provide mean flow information, turbulence intensities, and Reynolds stress information over the aft region of circulation control airfoil. The test condition used were: zero angle of attack, slot to chard ratio, h/c=0.002, and blowing momentum coefficients of Cll=O.lOO, and C,,=0.030. uv The free stream velocity used in these studies was 42.5 meterdsec which gave a chord Reynolds number y of 1 million. Y1/2 Ym In addition to the velocity data, surface pressure y ' data was obtained such that the performance aspects of the model airfoil tested were determined. Wind a tunnel wall-surface pressure data was also acquired so that methods to determine the equivalent free e air angle of attack could be employed for use in :?,v, experimental/numerical, comparisons. PP pa V

Investigation of low-speed turbulent separated flow around airfoils

Abstract: Described is a low-speed wind tunnel experiment to measure the flowfield around a two-dimensional airfoil operating close to maximum lift Boundary layer separation occurs on the upper surface at x/c=085 A three-component laser velocimeter, coupled with a computer-controlled data acquisition system, was used to obtain three orthogonal mean velocity components and three components of the Reynolds stress tensor in both the boundary layer and wake of the airfoil Pressure distributions on the airfoil, skin friction distribution on the upper surface of the airfoil, and integral properties of the airfoil boudary layer are also documented In addition to these near-field flow properties, static pressure distributions, both upstream and downstream from the airfoil and on the walls of the wind tunnel, are also presented

Low Reynolds Number Tests of NACA 64-210, NACA 0012, and Wortmann FX67-K170 Airfoils in Rain

Abstract: Wind-tunnel experiments were conducted on Wortmann FX67-K170, NACA 0012, and NACA 64-210 airfoils at a simulated rain rate of 1000 mm/h and Reynolds number of 3.1 X105 to compare the aerodynamic performance degradation of the airfoils in heavy rain conditions and to identify the various mechanisms that affect airfoil performance in rain conditions. Lift and drag were measured in both dry and wet conditions, and a variety of flow-visualiz ation techniques were employed. At low angles of attack, the lift degradation in wet conditions varied significantly between the airfoils. The Wortmann section had the greatest lift degradation (-25%) and the NACA 64-210 airfoil had the least (-5%). At high angles of attack, the NACA 64-210 and NACA 0012 airfoils were observed to have improved aerodynamic performance in rain conditions due to a reduction of boundary-layer separation. Performance degradation in heavy rain for all three airfoils at low angles of attack could be emulated by forced boundary-layer transition near the leading edge. Time-resolved measurements indicate two primary mechanisms are responsible for the observed performance degradation. The initial effect of rain is to cause premature boundary-layer transition at the leading edge. The second effect occurs at time scales consistent with top surface water runback (1-10 s). The runback layer is thought to alter the airfoil geometry effectively, but this effect is most likely exaggerated in these tests due to the small scale. The severity of the performance degradation for the airfoils varied. The relative differences appeared to be related to the susceptibility of each airfoil to premature boundary-layer transition.

Buffeting Prediction for Transport Aircraft Applications Based on Unsteady Pressure Measurements

Abstract: At ONERA a buffet onset prediction method based on computation of aircraft dynamic responses from the measurement of the unsteady pressure field on a model with scaled static deformation and geometry is under development. To build up such a method several experimental tests were performed on both RA16SC1 airfoil and rectangular wing to analyze in details unsteady separated flows. The existence of organized separated flow fluctuations at transonic speeds was observed. But this type of flow does not seem to be relevant at buffet onset for transport aircraft. Dynamic buffet responses have also been calculated on a 1/38 scale wing model from more complete unsteady pressure measurements. Reasonable agreement was obtained with experimental results at least for transonic speeds.

Studies of the flow field near a NACA 4412 aerofoil at nearly maximum lift

Abstract: : Measurements made at a Mach number of 0.18 and a chord-based Reynolds number of 4.2 x million on a constant-chord model having a NACA 4412 aerofoil section are described and compared with the results of flow field calculations. Both the experimental arrangement and the difficulties initially experienced in achieving an adequate approximation to two-dimensional flow above the wing are briefly outlined. The measurements include static pressure distributions on the wing surface and on the wind tunnel walls above and below the mid-span section of the wing. The main emphasis in the experiment was, however, on defining the development of the upper surface boundary layer through separation (at about 20% chord ahead of the trailing-edge) and on into the wake, making extensive use of laser anemometry. The flow field calculations are the semi-inverse kind in which an inverse momentum-integral treatment of the shear flow, used to avoid difficulties at separation, is coupled to a direct solution of the inviscid flow problem. The main features of the method are outlined. Keywords: Turbulent boundary layers; Wakes; Flow separation; Aerofoil flow. (Great Britain).

Inverse aerodynamic design method for aircraft components

Abstract: N existing, semi-inverse, aerodynamic design algorithm is ified to permit the unrestricted geometric design of aircraft components with prescribed aerodynamic surface pressures. A brief description of the present design procedure is given and computed results are presented for both a twodimensional airfoil and a three-dimensional nacelle configuration.

Unsteady Stall Penetration Experiments at High Reynolds Number

Abstract: : An experiment was performed to examine the unsteady aerodynamics of stall penetration at constant pitch rate and high Reynolds number, in an attempt to more accurately model conditions during aircraft post-stall maneuvers and during helicopter high speed forward flight The model spanned the 8 ft wind tunnel and consisted of a 173 in chord wing with a Sikorsky SSC-AOQ airfoil section Two forms of pitching motion were used: constant pitch rate ramps and sinusoidal oscillations Ramp data were obtained for 36 test points at pitch rates between 0001 and 0020, Mach numbers between 02 and 04, and Reynolds numbers between 2 and 4 million Sinusoidal data were obtained for an additional 9 conditions The results demonstrate the influence of the leading edge stall vortex on the unsteady aerodynamic response during and after stall The vortex- related unsteady increments to the lift, drag, and pitching moment increase with pitch rate; the maximum delta C sub L is 12 at A =002 Angular delays in stall events also increase with pitch rate Vortex strength and propagation velocity were determined from pressures induced on the airfoil surface The vortex is strengthened by increasing the pitch rate, and is weakened both by increasing the Mach number and by starting the motion close to the steady-state stall angle Propagation velocity increases linearly with pitch rate

Status of the special-purpose airfoil families

Abstract: This work is directed at developing thin and thick airfoil families, for rotors with diameters of 10 to 30 m, that enhance energy output at low to medium wind speeds and provide more consistent operating characteristics with lower fatigue loads at high wind speeds. Performance is enhanced through the use of laminar flow, while more consistent rotor operating characteristics at high wind speeds are achieved by tailoring the airfoil such that the maximum lift coefficient C/sub 1,max/ is largely independent of roughness effects. Using the Eppler airfoil design code, two thin and one thick airfoil family were designed; each family has a root, outboard, and tip airfoil. Two-dimensional wind-tunnel tests were conducted to verify the predicted performance characteristics for both a thin and thick outboard airfoil from these families. Atmospheric tests on full-scale wind turbines will complete the verification process. 3 refs., 7 figs., 3 tabs.

Pressure distributions from high Reynolds number transonic tests of an NACA 0012 airfoil in the Langley 0.3-meter transonic cryogenic tunnel

Abstract: Tests were conducted in the 2-D test section of the Langley 0.3-meter Transonic Cryogenic Tunnel on a NACA 0012 airfoil to obtain aerodynamic data as a part of the Advanced Technology Airfoil Test (ATAT) program. The test program covered a Mach number range of 0.30 to 0.82 and a Reynolds number range of 3.0 to 45.0 x 10 to the 6th power. The stagnation pressure was varied between 1.2 and 6.0 atmospheres and the stagnation temperature was varied between 300 K and 90 K to obtain these test conditions. Tabulated pressure distributions and integrated force and moment coefficients are presented as well as plots of the surface pressure distributions. The data are presented uncorrected for wall interference effects and without analysis.

A unique measurement technique to study laminar-separation bubble characteristics on an airfoil

Abstract: A 'nonintrusive', multielement heat-transfer sensor was designed to study laminar-separation bubble characteristics on a NASA LRN (1)-1010 low-Reynolds number airfoil. The sensor consists of 30 individual nickel films, vacuum-deposited on a thin substrate (0.05 mm) that was bonded to the airfoil model with the sensor array placed streamwise on the airfoil upper surface. Experiments were conducted on a 15-cm chord model in the 50,000-300,000 chord Reynolds number range. Time history as well as spectral analysis of signals from surface film gauges were simultaneously obtained to determine the location of laminar separation and the subsequent behavior of the separated shear layer. In addition to the successful determination of laminar separation, a new phenomenon involving a large phase shift in dynamic shear stresses across the separation and reattachment points was observed.

Airfoils with leading edge pockets for reduced heat transfer

Abstract: A gas turbine engine airfoil leading edge includes depressions therein longitudinally spaced apart and centered on stagnation points. The depressions stay filled with relatively stationary air during engine operation and reduce the heat load at the leading edge, which is primarily cooled by an internal supply of cooling fluid.