The acoustics and unsteady wall pressure of a circulation control airfoil
NASA Astrophysics Data System (ADS)
Silver, Jonathan C.
A Circulation Control (CC) airfoil uses a wall jet exiting onto a rounded trailing edge to generate lift via the Coanda effect. The aerodynamics of the CC airfoil have been studied extensively. The acoustics of the airfoil are, however, much less understood. The primary goal of the present work was to study the radiated sound and unsteady surface pressures of a CC airfoil. The focus of this work can be divided up into three main categories: characterizing the unsteady surface pressures, characterizing the radiated sound, and understanding the acoustics from surface pressures. The present work is the first to present the unsteady surface pressures from the trailing edge cylinder of a circulation control airfoil. The auto-spectral density of the unsteady surface pressures at various locations around the trailing edge are presented over a wide range of the jets momentum coefficient. Coherence of pressure and length scales were computed and presented. Single microphone measurements were made at a range of angles for a fixed observer distance in the far field. Spectra are presented for select angles to show the directivity of the airfoil's radiated sound. Predictions of the acoustics were made from unsteady surface pressures via Howe's curvature noise model and a modified Curle's analogy. A summary of the current understanding of the acoustics from a CC airfoil is given along with suggestions for future work.
Global Pressure Measurement of Unsteady-State Flow and Motion on Fluttering Airfoil
NASA Astrophysics Data System (ADS)
Okabe, Taika; Miyazaki, Takeshi; Saitoh, Kenichi; Sakaue, Hirotaka
2012-11-01
Pressure-sensitive paint (PSP) measurement has been applied to a fluid dynamic measurement. It can be applied to a steady-state flow in the transonic and supersonic wind tunnels. To extend the PSP measurement, an unsteady-state measurement is paid attention. It can be categorized by the unsteady-state flow and unsteady motion of the PSP-coated model. The former can be captured by using a porous PSP. The fastest PSP gives the response time on the order of ten microseconds. For the latter, the motion-capturing PSP method is studied to capture the unsteady motion. It consists of a two-color PSP and color camera. One color corresponds to the pressure-independent luminescence, and the other to the pressure-dependent luminescence. The former is used to cancel the pressure-independent distribution of a pressure-dependent image. The two-color images are simultaneously captured by the color camera. A combination of the unsteady-state flow and the motion can be occurred for an unsteady-state measurement. We present one of the cases, which is a flutter on an airfoil. By combining a porous PSP and the motion-capturing method, a pressure distribution on a fluttering airfoil (2-D YXX 30% span) is captured.
Unsteady pressure measurements on a supercritical airfoil at high Reynolds numbers
NASA Technical Reports Server (NTRS)
Hess, R. W.
1989-01-01
Steady and unsteady pressures were measured on a 14 percent supercritical airfoil at transonic Mach numbers at Reynolds numbers from 6,000,000 to 35,000,000. Instrumentation techniques were developed to measure unsteady pressures in a cryogenic tunnel at flight Reynolds numbers. Experimental steady data, corrected for wall effects show very good agreement with calculations from a full potential code with an interacted boundary layer. The steady and unsteady pressures both show a shock position that is dependent on Reynolds number. For a supercritical pressure distribution at a chord Reynolds number of 35,000,000 laminar flow was observed between the leading edge and the shock wave at 45 percent chord.
Unsteady pressure measurements on a biconvex airfoil in a transonic oscillating cascade
NASA Technical Reports Server (NTRS)
Shaw, L. M.; Boldman, D. R.; Buggele, A. E.; Buffum, D. H.
1985-01-01
Flush-mounted dynamic pressure transducers were installed on the center airfoil of a transonic oscillating cascade to measure the unsteady aerodynamic response as nine airfroils were simultaneously driven to provide 1.2 deg of pitching motion about the midchord. Initial tests were performed at an incidence and angle of 0 deg and A Mach number of 0.65 in order to obtain results in a shock-free compressible flowfield. Subsequent tests were performed at an incidence angle of 7 deg and Mach number of 0.8 in order to observe the surface pressures with an oscillating shock near the leading edge of the airfoil. Results are presented for interblade phase angles of 90 and -90 deg and at blade oscillatory frequencies of 200 and 500 Hz (semi-chord reduced frequencies up to about 0.5 at a Mach number of 0.8). Results from the zero-incidence cascade are compared with a classical unsteady flat-plate analysis. Flow visualization results depicting the shock motion on the airfoils in the high-incidence cascade are discussed. The airfoil pressure data are tabulated.
An Experimental Investigation of Unsteady Surface Pressure on an Airfoil in Turbulence
NASA Technical Reports Server (NTRS)
Mish, Patrick F.; Devenport, William J.
2003-01-01
Measurements of fluctuating surface pressure were made on a NACA 0015 airfoil immersed in grid generated turbulence. The airfoil model has a 2 ft chord and spans the 6 ft Virginia Tech Stability Wind Tunnel test section. Two grids were used to investigate the effects of turbulence length scale on the surface pressure response. A large grid which produced turbulence with an integral scale 13% of the chord and a smaller grid which produced turbulence with an integral scale 1.3% of the chord. Measurements were performed at angles of attack, alpha from 0 to 20 . An array of microphones mounted subsurface was used to measure the unsteady surface pressure. The goal of this measurement was to characterize the effects of angle of attack on the inviscid response. Lift spectra calculated from pressure measurements at each angle of attack revealed two distinct interaction regions; for omega(sub r) = omega b / U(sub infinity) is less than 10 a reduction in unsteady lift of up to 7 decibels (dB) occurs while an increase occurs for omega(sub r) is greater than 10 as the angle of attack is increased. The reduction in unsteady lift at low omega(sub r) with increasing angle of attack is a result that has never before been shown either experimentally or theoretically. The source of the reduction in lift spectral level appears to be closely related to the distortion of inflow turbulence based on analysis of surface pressure spanwise correlation length scales. Furthermore, while the distortion of the inflow appears to be critical in this experiment, this effect does not seem to be significant in larger integral scale (relative to the chord) flows based on the previous experimental work of McKeough suggesting the airfoils size relative to the inflow integral scale is critical in defining how the airfoil will respond under variation of angle of attack. A prediction scheme is developed that correctly accounts for the effects of distortion when the inflow integral scale is small relative
Unsteady Airloads on Airfoils in Reverse Flow
NASA Astrophysics Data System (ADS)
Lind, Andrew; Jones, Anya
2014-11-01
This work gives insight into the influence of airfoil characteristics on unsteady airloads for rotor applications where local airfoil sections may operate at high and/or reverse flow angles of attack. Two-dimensional wind tunnel experiments have been performed on four airfoil sections to investigate the effects of thickness, camber, and trailing edge shape on unsteady airloads (lift, pressure drag, and pitching moment). These model rotor blades were tested through 360 deg of incidence for 104 <=Re <=106 . Unsteady pressure transducers were mounted on the airfoil surface to measure the high frequency, dynamic pressure variations. The temporal evolution of chordwise pressure distributions and resulting airloads is quantified for each airfoil in each of the three unsteady wake regimes present in reverse flow. Specifically, the influence of the formation, growth, and shedding of vortices on the surface pressure distribution is quantified and compared between airfoils with a sharp geometric trailing edge and those with a blunt geometric trailing edge. These findings are integral to mitigation of rotor blade vibrations for applications where airfoil sections are subjected to reverse flow, such as high-speed helicopters and tidal turbines.
Unsteady Newton-Busemann flow theory. I - Airfoils
NASA Technical Reports Server (NTRS)
Hui, W. H.; Tobak, M.
1981-01-01
Newtonian flow theory for unsteady flow at very high Mach numbers is completed by the addition of a centrifugal force correction to the impact pressures. The correction term is the unsteady counterpart of Busemann's centrifugal force correction to impact pressures in steady flow. For airfoils of arbitary shape, exact formulas for the unsteady pressure and stiffness and damping-in-pitch derivatives are obtained in closed form, which require only numerical quadratures of terms involving the airfoil shape. They are applicable to airfoils of arbitrary thickness having sharp or blunt leading edges. For wedges and thin airfoils these formulas are greatly simplified, and it is proved that the pitching motions of thin airfoils of convex shape and of wedges of arbitrary thickness are always dynamically stable according to Newton-Busemann theory. Leading-edge bluntness is shown to have a favorable effect on the dynamic stability; on the other hand, airfoils of concave shape tend toward dynamic instability over a range of axis positions if the surface curvature exceeds a certain limit. As a byproduct, it is also shown that a pressure formula recently given by Barron and Mandl for unsteady Newtonian flow over a pitching power-law shaped airfoil is erroneous and that their conclusion regarding the effect of pivot position on the dynamic stability is misleading.
Numerical studies of unsteady transonic flow over oscillating airfoil
NASA Technical Reports Server (NTRS)
Chyu, W. J.; Davis, S. S.
1985-01-01
A finite difference solution to the Navier-Stokes equations combined with a time varying grid generation technique was used to compute unsteady transonic flow over an oscillating airfoil. These computations were compared with experimental data (obtained at Ames Research Center) which form part of the AGARD standard configuration for aeroelastic analysis. A variety of approximations to the full Navier-Stokes equations was used to determine the effect of frequency, shock wave motion, flow separation, and airfoil geometry on unsteady pressures and overall air loads. Good agreement is shown between experiment and theory with the limiting factor being the lack of a reliable turbulence model for high Reynolds number, unsteady transonic flows.
Unsteady transonic flow control around an airfoil in a channel
NASA Astrophysics Data System (ADS)
Hamid, Md. Abdul; Hasan, A. B. M. Toufique; Ali, Mohammad; Mitsutake, Yuichi; Setoguchi, Toshiaki; Yu, Shen
2016-04-01
Transonic internal flow around an airfoil is associated with self-excited unsteady shock wave oscillation. This unsteady phenomenon generates buffet, high speed impulsive noise, non-synchronous vibration, high cycle fatigue failure and so on. Present study investigates the effectiveness of perforated cavity to control this unsteady flow field. The cavity has been incorporated on the airfoil surface. The degree of perforation of the cavity is kept constant as 30%. However, the number of openings (perforation) at the cavity upper wall has been varied. Results showed that this passive control reduces the strength of shock wave compared to that of baseline airfoil. As a result, the intensity of shock wave/boundary layer interaction and the root mean square (RMS) of pressure oscillation around the airfoil have been reduced with the control method.
Synthesized airfoil data method for prediction of dynamic stall and unsteady airloads
NASA Technical Reports Server (NTRS)
Gangwani, S. T.
1983-01-01
A detailed analysis of dynamic stall experiments has led to a set of relatively compact analytical expressions, called synthesized unsteady airfoil data, which accurately describe in the time-domain the unsteady aerodynamic characteristics of stalled airfoils. An analytical research program was conducted to expand and improve this synthesized unsteady airfoil data method using additional available sets of unsteady airfoil data. The primary objectives were to reduce these data to synthesized form for use in rotor airload prediction analyses and to generalize the results. Unsteady drag data were synthesized which provided the basis for successful expansion of the formulation to include computation of the unsteady pressure drag of airfoils and rotor blades. Also, an improved prediction model for airfoil flow reattachment was incorporated in the method. Application of this improved unsteady aerodynamics model has resulted in an improved correlation between analytic predictions and measured full scale helicopter blade loads and stress data.
NASA Technical Reports Server (NTRS)
Batina, J. T.
1985-01-01
The effects of airfoil shape, thickness, camber, and mean angle of attack on transonic unsteady airloads were investigated as calculated by the transonic small-disturbance computer code XTRAN2L. Shape effects were investigated by examining the pressure distributions, shock locations, and unsteady airloads for three 10 percent thick airfoils. NACA 0010, NACA 64A010, and parabolic arc. Thickness effects were determined by studying a single airfoil shape with three different thicknesses: NACA 0008, NACA 0010, and NACA 0012. Angle of attack and camber effects were studied by including mean angle of attack or by adding a simple parabolic camber distribution to the originally symmetric airfoils. Comparisons of unsteady airloads for different airfoil configurations show similar results caused by variations in airfoil shape, thickness, camber, or mean angle of attack. Computer costs can be reduced by limiting the number of transonic unsteady aerodynamic calculations for small changes in airfoil geometry or angle of attack.
Unsteady Thick Airfoil Aerodynamics: Experiments, Computation, and Theory
NASA Technical Reports Server (NTRS)
Strangfeld, C.; Rumsey, C. L.; Mueller-Vahl, H.; Greenblatt, D.; Nayeri, C. N.; Paschereit, C. O.
2015-01-01
An experimental, computational and theoretical investigation was carried out to study the aerodynamic loads acting on a relatively thick NACA 0018 airfoil when subjected to pitching and surging, individually and synchronously. Both pre-stall and post-stall angles of attack were considered. Experiments were carried out in a dedicated unsteady wind tunnel, with large surge amplitudes, and airfoil loads were estimated by means of unsteady surface mounted pressure measurements. Theoretical predictions were based on Theodorsen's and Isaacs' results as well as on the relatively recent generalizations of van der Wall. Both two- and three-dimensional computations were performed on structured grids employing unsteady Reynolds-averaged Navier-Stokes (URANS). For pure surging at pre-stall angles of attack, the correspondence between experiments and theory was satisfactory; this served as a validation of Isaacs theory. Discrepancies were traced to dynamic trailing-edge separation, even at low angles of attack. Excellent correspondence was found between experiments and theory for airfoil pitching as well as combined pitching and surging; the latter appears to be the first clear validation of van der Wall's theoretical results. Although qualitatively similar to experiment at low angles of attack, two-dimensional URANS computations yielded notable errors in the unsteady load effects of pitching, surging and their synchronous combination. The main reason is believed to be that the URANS equations do not resolve wake vorticity (explicitly modeled in the theory) or the resulting rolled-up un- steady flow structures because high values of eddy viscosity tend to \\smear" the wake. At post-stall angles, three-dimensional computations illustrated the importance of modeling the tunnel side walls.
NASA Technical Reports Server (NTRS)
Vinci, Samuel, J.
2012-01-01
This report is the third part of a three-part final report of research performed under an NRA cooperative Agreement contract. The first part was published as NASA/CR-2012-217415. The second part was published as NASA/CR-2012-217416. The study of the very high lift low-pressure turbine airfoil L1A in the presence of unsteady wakes was performed computationally and compared against experimental results. The experiments were conducted in a low speed wind tunnel under high (4.9%) and then low (0.6%) freestream turbulence intensity for Reynolds number equal to 25,000 and 50,000. The experimental and computational data have shown that in cases without wakes, the boundary layer separated without reattachment. The CFD was done with LES and URANS utilizing the finite-volume code ANSYS Fluent (ANSYS, Inc.) under the same freestream turbulence and Reynolds number conditions as the experiment but only at a rod to blade spacing of 1. With wakes, separation was largely suppressed, particularly if the wake passing frequency was sufficiently high. This was validated in the 3D CFD efforts by comparing the experimental results for the pressure coefficients and velocity profiles, which were reasonable for all cases examined. The 2D CFD efforts failed to capture the three dimensionality effects of the wake and thus were less consistent with the experimental data. The effect of the freestream turbulence intensity levels also showed a little more consistency with the experimental data at higher intensities when compared with the low intensity cases. Additional cases with higher wake passing frequencies which were not run experimentally were simulated. The results showed that an initial 25% increase from the experimental wake passing greatly reduced the size of the separation bubble, nearly completely suppressing it.
Viscous effect on airfoils for unsteady transonic flows
NASA Technical Reports Server (NTRS)
Lee, S. C.
1982-01-01
The viscous effect on aerodynamic performance of an arbitrary airfoil executing low frequency maneuvers during transonic flight was investigated. The small disturbance code, LTRAN2, was modified by using a conventional integral method, BLAYER, for the boundary layer and an empirical relation, viscous wedge, for simulating the suddenly thickened boundary layer behind the shock. Before the shock, only the boundary layer displacement thickness was evaluated. After the shock, the empirical wedge thickness was superimposed on the boundary layer thickness along the surface as well as in the wake region. The pressure coefficients were calculated for both steady and unsteady states. The viscous solution takes fewer iterations to obtain the converged steady state solution. Comparisons made with experimental data and the inviscid solution show that the viscous solution agrees better with the experimental data with about the same (or slightly less) amount of computational time.
Unsteady flow past an airfoil pitched at constant rate
NASA Technical Reports Server (NTRS)
Lourenco, L.; Vandommelen, L.; Shib, C.; Krothapalli, A.
1992-01-01
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.
On the Theory of the Unsteady Motion of an Airfoil
NASA Technical Reports Server (NTRS)
Sedov, L. I.
1947-01-01
The paper presents a systematical analysis of the problem of the determination of the unsteady motion about an airfoil moving in an infinite fluid that contains a system of vortices and the determination of the hydrodynamical forces acting on the airfoil. The hydrodynamical problem is reduced to the determination of the function f (xi) which transforms conformally the external region of the airfoil into the interior of a circle. The proposed methods of determining the irrotational motion of a fluid that is produced by any motion of the airfoil are especially simple and effective if the function f (xi) is rational. As an example the flow is determined for the case of an arbitrary motion of an airfoil of the Joukowsky type. The formulas obtained for the determination of the hydrodynamical forces by means of contour integration are similar to those given by S. Chaplygin. These formulas are used to determine the force acting on the airfoil in the cases where the unsteady motion is potential throughout and the circulation about the airfoil is constant and also when the fluid contains a system of vortices. A full discussion is given of the concept of virtual masses together with practical formulas for computing the virtual mass coefficients.
Transonic unsteady aerodynamic and aeroelastic calculations about airfoils and wings
NASA Technical Reports Server (NTRS)
Goorjian, Peter M.; Guruswamy, Guru P.
1988-01-01
Recent advances in the numerical simulation of unsteady transonic flow around airfoils and wings are surveyed, with an emphasis on the treatment of aeroelastic effects. The fundamental physical principles involved are discussed, and the numerical implementation of the methods is considered. Typical results are presented in extensive graphs and diagrams and briefly characterized, with reference to experimental data.
Vortex scale of unsteady separation on a pitching airfoil.
Fuchiwaki, Masaki; Tanaka, Kazuhiro
2002-10-01
The streaklines of unsteady separation on two kinds of pitching airfoils, the NACA65-0910 and a blunt trailing edge airfoil, were studied by dye flow visualization and by the Schlieren method. The latter visualized the discrete vortices shed from the leading edge. The results of these visualization studies allow a comparison between the dynamic behavior of the streakline of unsteady separation and that of the discrete vortices shed from the leading edge. The influence of the airfoil configuration on the flow characteristics was also examined. Furthermore, the scale of a discrete vortex forming the recirculation region was investigated. The non-dimensional pitching rate was k = 0.377, the angle of attack alpha(m) = 16 degrees and the pitching amplitude was fixed to A = +/-6 degrees for Re = 4.0 x 10(3) in this experiment.
Theory and Low-Order Modeling of Unsteady Airfoil Flows
NASA Astrophysics Data System (ADS)
Ramesh, Kiran
Unsteady flow phenomena are prevalent in a wide range of problems in nature and engineering. These include, but are not limited to, aerodynamics of insect flight, dynamic stall in rotorcraft and wind turbines, leading-edge vortices in delta wings, micro-air vehicle (MAV) design, gust handling and flow control. The most significant characteristics of unsteady flows are rapid changes in the circulation of the airfoil, apparent-mass effects, flow separation and the leading-edge vortex (LEV) phenomenon. Although experimental techniques and computational fluid dynamics (CFD) methods have enabled the detailed study of unsteady flows and their underlying features, a reliable and inexpensive loworder method for fast prediction and for use in control and design is still required. In this research, a low-order methodology based on physical principles rather than empirical fitting is proposed. The objective of such an approach is to enable insights into unsteady phenomena while developing approaches to model them. The basis of the low-order model developed here is unsteady thin-airfoil theory. A time-stepping approach is used to solve for the vorticity on an airfoil camberline, allowing for large amplitudes and nonplanar wakes. On comparing lift coefficients from this method against data from CFD and experiments for some unsteady test cases, it is seen that the method predicts well so long as LEV formation does not occur and flow over the airfoil is attached. The formation of leading-edge vortices (LEVs) in unsteady flows is initiated by flow separation and the formation of a shear layer at the airfoil's leading edge. This phenomenon has been observed to have both detrimental (dynamic stall in helicopters) and beneficial (high-lift flight in insects) effects. To predict the formation of LEVs in unsteady flows, a Leading Edge Suction Parameter (LESP) is proposed. This parameter is calculated from inviscid theory and is a measure of the suction at the airfoil's leading edge. It
Steady and Unsteady Aerodynamics of Thin Airfoils with Porosity Gradients
NASA Astrophysics Data System (ADS)
Hajian, Rozhin; Jaworski, Justin W.
2015-11-01
Porous treatments have been shown in previous studies to reduce turbulence noise generation from the edges of wings and blades. However, this acoustical benefit can come at the cost of aerodynamic performance that is degraded by seepage flow through the wing. To better understand the trade-off between acoustic stealth and the desired airfoil performance, the aerodynamic loads of a thin airfoil in uniform flow with a prescribed porosity distribution are determined analytically in closed form, provided that the distribution is Hölder-continuous. The theoretical model is extended to include unsteady heaving and pitching motions of the airfoil section, which has applications to the performance estimation of biologically-inspired swimmers and fliers and to the future assessment of vortex noise production from porous airfoils.
NASA Technical Reports Server (NTRS)
Ehlers, F. E.; Weatherill, W. H.
1982-01-01
A finite difference method for solving the unsteady transonic flow about harmonically oscillating wings is investigated. The procedure is based on separating the velocity potential into steady and unsteady parts and linearizing the resulting unsteady differential equation for small disturbances. The differential equation for the unsteady velocity potential is linear with spatially varying coefficients and with the time variable eliminated by assuming harmonic motion. A study is presented of the shock motion associated with an oscillating airfoil and its representation by the harmonic procedure. The effects of the shock motion and the resulting pressure pulse are shown to be included in the harmonic pressure distributions and the corresponding generalized forces. Analytical and experimental pressure distributions for the NACA 64A010 airfoil are compared for Mach numbers of 0.75, 0.80 and 0.842. A typical section, two-degree-of-freedom flutter analysis of a NACA 64A010 airfoil is performed. The results show a sharp transonic bucket in one case and abrupt changes in instability modes.
Some examples of unsteady transonic flows over airfoils
NASA Technical Reports Server (NTRS)
Ballhaus, W. F.; Magnus, R.; Yoshihara, H.
1975-01-01
A finite difference flutter analysis is presented for the NACA 64A-410 airfoil at M equals 0.72, where the incidence is abruptly changed from 2 to 4 degrees. The effect of gust loads is studied, and the unsteady flow adjusting process is displayed. The semi-implicit procedure of Ballhaus and Lomax (1974) is used to solve the small disturbance transonic potential equation. The physical aspects of the results, rather than the numerical details, are emphasized.
NASA Technical Reports Server (NTRS)
Fromme, J.; Golberg, M.
1978-01-01
The numerical calculation of unsteady two dimensional airloads which act upon thin airfoils in subsonic ventilated wind tunnels was studied. Neglecting certain quadrature errors, Bland's collocation method is rigorously proved to converge to the mathematically exact solution of Bland's integral equation, and a three way equivalence was established between collocation, Galerkin's method and least squares whenever the collocation points are chosen to be the nodes of the quadrature rule used for Galerkin's method. A computer program displayed convergence with respect to the number of pressure basis functions employed, and agreement with known special cases was demonstrated. Results are obtained for the combined effects of wind tunnel wall ventilation and wind tunnel depth to airfoil chord ratio, and for acoustic resonance between the airfoil and wind tunnel walls. A boundary condition is proposed for permeable walls through which mass flow rate is proportional to pressure jump.
Oscillatory Excitation of Unsteady Compressible Flows over Airfoils at Flight Reynolds Numbers
NASA Technical Reports Server (NTRS)
Seifert, Avi; Pack, LaTunia G.
1999-01-01
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.
Unsteady transonic aerodynamic and aeroelastic calculations about airfoils and wings
NASA Technical Reports Server (NTRS)
Goorjian, P. M.; Guruswamy, G. P.
1985-01-01
The development and application of transonic small disturbance codes for computing two dimensional flows, using the code ATRAN2, and for computing three dimensional flows, using the code ATRAN3S, are described. Calculated and experimental results are compared for unsteady flows about airfoils and wings, including several of the cases from the AGARD Standard Aeroelastic Configurations. In two dimensions, the results include AGARD priority cases for the NACA 54A006, NACA 64A010, NACA 0012, and MBB-A3 airfoils. In three dimensions, the results include flow about the F-5 wing, a typical wing, and the AGARD rectangular wings. Viscous corrections are included in some calculations, including those for the AGARD rectangular wing. For several cases, the aerodynamic and aeroelastic calculations are compared with experimental results.
NASA Astrophysics Data System (ADS)
Sun, Min; Yang, Bo; Peng, Tianxiang; Lei, Mingkai
2016-06-01
Unsteady dielectric barrier discharge (DBD) plasma aerodynamic actuation technology is employed to suppress airfoil stall separation and the technical parameters are explored with wind tunnel experiments on an NACA0015 airfoil by measuring the surface pressure distribution of the airfoil. The performance of the DBD aerodynamic actuation for airfoil stall separation suppression is evaluated under DBD voltages from 2000 V to 4000 V and the duty cycles varied in the range of 0.1 to 1.0. It is found that higher lift coefficients and lower threshold voltages are achieved under the unsteady DBD aerodynamic actuation with the duty cycles less than 0.5 as compared to that of the steady plasma actuation at the same free-stream speeds and attack angles, indicating a better flow control performance. By comparing the lift coefficients and the threshold voltages, an optimum duty cycle is determined as 0.25 by which the maximum lift coefficient and the minimum threshold voltage are obtained at the same free-stream speed and attack angle. The non-uniform DBD discharge with stronger discharge in the positive half cycle due to electrons deposition on the dielectric slabs and the suppression of opposite momentum transfer due to the intermittent discharge with cutoff of the negative half cycle are responsible for the observed optimum duty cycle. supported by National Natural Science Foundation of China (No. 21276036), Liaoning Provincial Natural Science Foundation of China (No. 2015020123) and the Fundamental Research Funds for the Central Universities of China (No. 3132015154)
Nonlinear effects of flow unsteadiness on the acoustic radiation of a heaving airfoil
NASA Astrophysics Data System (ADS)
Manela, Avshalom
2013-12-01
The study considers the combined effects of boundary animation (small-amplitude heaving) and incoming flow unsteadiness (incident vorticity) on the vibroacoustic signature of a thin rigid airfoil in low-Mach number flow. The potential-flow problem is analysed using the Brown and Michael equation, yielding the incident vortex trajectory and time evolution of trailing edge wake. The dynamical description serves as an effective source term to evaluate the far-field sound using Powell-Howe analogy. The results identify the fluid-airfoil system as a dipole-type source, and demonstrate the significance of nonlinear eddy-airfoil interactions on the acoustic radiation. Based on the value of scaled heaving frequency ωa/U (with ω the dimensional heaving frequency, a the airfoil half-chord, and U the mean flow speed), the system behaviour can be divided into two characteristic regimes: (i) for ωa/U≪1, the effect of heaving is minor, and the acoustic response is well approximated by considering the interaction of a line vortex with a stationary airfoil; (ii) for ωa/U≫1, the impact of heaving is dominant, radiating sound through an “airfoil motion” dipole oriented along the direction of heaving. In between (for ωa/U~O(1)), an intermediate regime takes place. The results indicate that trailing edge vorticity has a two-fold impact on the acoustic far field: while reducing pressure fluctuations generated by incident vortex interaction with the airfoil, trailing edge vortices transmit sound along the mean-flow direction, characterized by airfoil heaving frequency. The “silencing” effect of trailing edge vorticity is particularly efficient when the incident vortex passes close to the airfoil trailing edge: at that time, application of the Kutta condition implies the release of a trailing edge vortex in the opposite direction to the incident vortex; the released vortex then detaches from the airfoil and follows the incident vortex, forming a “silent” vortex pair
Control of unsteady separated flow associated with the dynamic stall of airfoils
NASA Technical Reports Server (NTRS)
Wilder, M. C.
1995-01-01
An effort to understand and control the unsteady separated flow associated with the dynamic stall of airfoils was funded for three years through the NASA cooperative agreement program. As part of this effort a substantial data base was compiled detailing the effects various parameters have on the development of the dynamic stall flow field. Parameters studied include Mach number, pitch rate, and pitch history, as well as Reynolds number (through two different model chord lengths) and the condition of the boundary layer at the leading edge of the airfoil (through application of surface roughness). It was found for free stream Mach numbers as low as 0.4 that a region of supersonic flow forms on the leading edge of the suction surface of the airfoil at moderate angles of attack. The shocks which form in this supersonic region induce boundary-layer separation and advance the dynamic stall process. Under such conditions a supercritical airfoil profile is called for to produce a flow field having a weaker leading-edge pressure gradient and no leading-edge shocks. An airfoil having an adaptive-geometry, or dynamically deformable leading edge (DDLE), is under development as a unique active flow-control device. The DDLE, formed of carbon-fiber composite and fiberglass, can be flexed between a NACA 0012 profile and a supercritical profile in a controllable fashion while the airfoil is executing an angle-of-attack pitch-up maneuver. The dynamic stall data were recorded using point diffraction interferometry (PDI), a noninvasive measurement technique. A new high-speed cinematography system was developed for recording interferometric images. The system is capable of phase-locking with the pitching airfoil motion for real-time documentation of the development of the dynamic stall flow field. Computer-aided image analysis algorithms were developed for fast and accurate reduction of the images, improving interpretation of the results.
Unsteady Adjoint Approach for Design Optimization of Flapping Airfoils
NASA Technical Reports Server (NTRS)
Lee, Byung Joon; Liou, Meng-Sing
2012-01-01
This paper describes the work for optimizing the propulsive efficiency of flapping airfoils, i.e., improving the thrust under constraining aerodynamic work during the flapping flights by changing their shape and trajectory of motion with the unsteady discrete adjoint approach. For unsteady problems, it is essential to properly resolving time scales of motion under consideration and it must be compatible with the objective sought after. We include both the instantaneous and time-averaged (periodic) formulations in this study. For the design optimization with shape parameters or motion parameters, the time-averaged objective function is found to be more useful, while the instantaneous one is more suitable for flow control. The instantaneous objective function is operationally straightforward. On the other hand, the time-averaged objective function requires additional steps in the adjoint approach; the unsteady discrete adjoint equations for a periodic flow must be reformulated and the corresponding system of equations solved iteratively. We compare the design results from shape and trajectory optimizations and investigate the physical relevance of design variables to the flapping motion at on- and off-design conditions.
Recent transonic unsteady pressure measurements at the NASA Langley Research Center
NASA Technical Reports Server (NTRS)
Sandford, M. C.; Ricketts, R. H.; Hess, R. W.
1985-01-01
Four semispan wing model configurations were studied in the Transonic Dynamics Tunnel (TDT). The first model had a clipped delta planform with a circular arc airfoil, the second model had a high aspect ratio planform with a supercritical airfoil, the third model has a rectangular planform with a supercritical airfoil and the fourth model had a high aspect ratio planform with a supercritical airfoil. To generate unsteady flow, the first and third models were equipped with pitch oscillation mechanisms and the first, second and fourth models were equipped with control surface oscillation mechanisms. The fourth model was similar in planform and airfoil shape to the second model, but it is the only one of the four models that has an elastic wing structure. The unsteady pressure studies of the four models are described and some typical results for each model are presented. Comparison of selected experimental data with analytical results also are included.
NASA Technical Reports Server (NTRS)
Capece, Vincent R.; Platzer, Max F.
2003-01-01
A major challenge in the design and development of turbomachine airfoils for gas turbine engines is high cycle fatigue failures due to flutter and aerodynamically induced forced vibrations. In order to predict the aeroelastic response of gas turbine airfoils early in the design phase, accurate unsteady aerodynamic models are required. However, accurate predictions of flutter and forced vibration stress at all operating conditions have remained elusive. The overall objectives of this research program are to develop a transition model suitable for unsteady separated flow and quantify the effects of transition on airfoil steady and unsteady aerodynamics for attached and separated flow using this model. Furthermore, the capability of current state-of-the-art unsteady aerodynamic models to predict the oscillating airfoil response of compressor airfoils over a range of realistic reduced frequencies, Mach numbers, and loading levels will be evaluated through correlation with benchmark data. This comprehensive evaluation will assess the assumptions used in unsteady aerodynamic models. The results of this evaluation can be used to direct improvement of current models and the development of future models. The transition modeling effort will also make strides in improving predictions of steady flow performance of fan and compressor blades at off-design conditions. This report summarizes the progress and results obtained in the first year of this program. These include: installation and verification of the operation of the parallel version of TURBO; the grid generation and initiation of steady flow simulations of the NASA/Pratt&Whitney airfoil at a Mach number of 0.5 and chordal incidence angles of 0 and 10 deg.; and the investigation of the prediction of laminar separation bubbles on a NACA 0012 airfoil.
Numerical calculations of velocity and pressure distribution around oscillating airfoils
NASA Technical Reports Server (NTRS)
Bratanow, T.; Ecer, A.; Kobiske, M.
1974-01-01
An analytical procedure based on the Navier-Stokes equations was developed for analyzing and representing properties of unsteady viscous flow around oscillating obstacles. A variational formulation of the vorticity transport equation was discretized in finite element form and integrated numerically. At each time step of the numerical integration, the velocity field around the obstacle was determined for the instantaneous vorticity distribution from the finite element solution of Poisson's equation. The time-dependent boundary conditions around the oscillating obstacle were introduced as external constraints, using the Lagrangian Multiplier Technique, at each time step of the numerical integration. The procedure was then applied for determining pressures around obstacles oscillating in unsteady flow. The obtained results for a cylinder and an airfoil were illustrated in the form of streamlines and vorticity and pressure distributions.
Acoustic radiation and surface pressure characteristics of an airfoil due to incident turbulence
NASA Technical Reports Server (NTRS)
Paterson, R. W.
1976-01-01
A theoretical and experimental investigation of the noise and unsteady surface pressure characteristics of an isolated airfoil in a uniform mean velocity, homogeneous, nearly-isotropic turbulence field was conducted. Wind tunnel experiments were performed with a 23 cm chord, two dimensional NACA 0012 airfoil over a free stream Mach number range of 0.1 to 0.5. Far-field noise spectra and directivity were measured in an anechoic chamber that surrounded the tunnel open jet test section. Spanwise and chordwise distribution of unsteady airfoil surface pressure spectra and surface pressure cross-spectra were obtained. Incident turbulence intensities, length scales, spectra, and spanwise cross-spectra, required in the calculation of far-field noise and surface pressure characteristics were also measured.
Control of unsteady separated flow associated with the dynamic stall of airfoils
NASA Technical Reports Server (NTRS)
Wilder, M. C.
1994-01-01
A unique active flow-control device is proposed for the control of unsteady separated flow associated with the dynamic stall of airfoils. The device is an adaptive-geometry leading-edge which will allow controlled, dynamic modification of the leading-edge profile of an airfoil while the airfoil is executing an angle-of-attack pitch-up maneuver. A carbon-fiber composite skin has been bench tested, and a wind tunnel model is under construction. A baseline parameter study of compressible dynamic stall was performed for flow over an NACA 0012 airfoil. Parameters included Mach number, pitch rate, pitch history, and boundary layer tripping. Dynamic stall data were recorded via point-diffraction interferometry and the interferograms were analyzed with in-house developed image processing software. A new high-speed phase-locked photographic image recording system was developed for real-time documentation of dynamic stall.
Unsteady-Pressure and Dynamic-Deflection Measurements on an Aeroelastic Supercritical Wing
NASA Technical Reports Server (NTRS)
Seidel, David A.; Sandford, Maynard C.; Eckstrom, Clinton V.
1991-01-01
Transonic steady and unsteady pressure tests were conducted on a large elastic wing. The wing has a supercritical airfoil, a full span aspect ratio of 10.3, a leading edge sweepback angle of 28.8 degrees, and two inboard and one outboard trailing edge control surfaces. Only the outboard control surface was deflected statically and dynamically to generate steady and unsteady flow over the wing. The unsteady surface pressure and dynamic deflection measurements of this elastic wing are presented to permit correlations of the experimental data with theoretical predictions.
Numerical Simulations of Unsteady Low-Reynolds-Number Flows Over the APEX Airfoil
NASA Technical Reports Server (NTRS)
Tatineni, Mahidhar; Zhong, Xiao-Lin
1998-01-01
Laminar and transitional separation bubbles are an important feature of low-Reynolds-number flows over airfoils. The separation bubbles are unsteady and have a significant impact on the aerodynamic properties of the airfoils. In this paper unsteady low-Reynolds-number separated flows over the APEX airfoil are calculated using a Navier-Stokes solver. The numerical results show the presence of unsteady separation bubbles in the flowfield. An analysis of the numerical results shows that flowfield disturbances are amplified significantly in the separation bubble, leading to periodic vortex shedding. A linear stability analysis of the separated boundary layer is performed and the results show that the dominant wavenumber and frequency in the numerical simulations agree with the most unstable wavenumber and frequency from the linear stability analysis. The numerical results also show the growth and interaction of disturbance waves in the separation bubble. For transonic flows over the APEX airfoil, the calculations show that the presence of shocks causes significant changes in the separation location and consequently, the overall flowfield.
Tables of properties of airfoil polynomials
NASA Technical Reports Server (NTRS)
Desmarais, Robert N.; Bland, Samuel R.
1995-01-01
This monograph provides an extensive list of formulas for airfoil polynomials. These polynomials provide convenient expansion functions for the description of the downwash and pressure distributions of linear theory for airfoils in both steady and unsteady subsonic flow.
Measurement of unsteady pressures in rotating systems
NASA Technical Reports Server (NTRS)
Kienappel, K.
1978-01-01
The principles of the experimental determination of unsteady periodic pressure distributions in rotating systems are reported. An indirect method is discussed, and the effects of the centrifugal force and the transmission behavior of the pressure measurement circuit were outlined. The required correction procedures are described and experimentally implemented in a test bench. Results show that the indirect method is suited to the measurement of unsteady nonharmonic pressure distributions in rotating systems.
Simulation of self-induced unsteady motion in the near wake of a Joukowski airfoil
NASA Technical Reports Server (NTRS)
Ghia, K. N.; Osswald, G. A.; Ghia, U.
1986-01-01
The unsteady Navier-Stokes analysis is shown to be capable of analyzing the massively separated, persistently unsteady flow in the post-stall regime of a Joukowski airfoil for an angle of attack as high as 53 degrees. The analysis has provided the detailed flow structure, showing the complex vortex interaction for this configuration. The aerodynamic coefficients for lift, drag, and moment were calculated. So far only the spatial structure of the vortex interaction was computed. It is now important to potentially use the large-scale vortex interactions, an additional energy source, to improve the aerodynamic performance.
Unsteady transonic flow past airfoils in rigid-body motion. [UFLO5
Chang, I C
1981-03-01
With the aim of developing a fast and accurate computer code for predicting the aerodynamic forces needed for a flutter analysis, some basic concepts in computational transonics are reviewed. The unsteady transonic flow past airfoils in rigid body motion is adequately described by the potential flow equation as long as the boundary layer remains attached. The two dimensional unsteady transonic potential flow equation in quasilinear form with first order radiation boundary conditions is solved by an alternating direction implicit scheme in an airfoil attached sheared parabolic coordinate system. Numerical experiments show that the scheme is very stable and is able to resolve the higher nonlinear transonic effects for filter analysis within the context of an inviscid theory.
Validation of DYSTOOL for unsteady aerodynamic modeling of 2D airfoils
NASA Astrophysics Data System (ADS)
González, A.; Gomez-Iradi, S.; Munduate, X.
2014-06-01
From the point of view of wind turbine modeling, an important group of tools is based on blade element momentum (BEM) theory using 2D aerodynamic calculations on the blade elements. Due to the importance of this sectional computation of the blades, the National Renewable Wind Energy Center of Spain (CENER) developed DYSTOOL, an aerodynamic code for 2D airfoil modeling based on the Beddoes-Leishman model. The main focus here is related to the model parameters, whose values depend on the airfoil or the operating conditions. In this work, the values of the parameters are adjusted using available experimental or CFD data. The present document is mainly related to the validation of the results of DYSTOOL for 2D airfoils. The results of the computations have been compared with unsteady experimental data of the S809 and NACA0015 profiles. Some of the cases have also been modeled using the CFD code WMB (Wind Multi Block), within the framework of a collaboration with ACCIONA Windpower. The validation has been performed using pitch oscillations with different reduced frequencies, Reynolds numbers, amplitudes and mean angles of attack. The results have shown a good agreement using the methodology of adjustment for the value of the parameters. DYSTOOL have demonstrated to be a promising tool for 2D airfoil unsteady aerodynamic modeling.
Time domain numerical calculations of unsteady vortical flows about a flat plate airfoil
NASA Technical Reports Server (NTRS)
Hariharan, S. I.; Yu, Ping; Scott, J. R.
1989-01-01
A time domain numerical scheme is developed to solve for the unsteady flow about a flat plate airfoil due to imposed upstream, small amplitude, transverse velocity perturbations. The governing equation for the resulting unsteady potential is a homogeneous, constant coefficient, convective wave equation. Accurate solution of the problem requires the development of approximate boundary conditions which correctly model the physics of the unsteady flow in the far field. A uniformly valid far field boundary condition is developed, and numerical results are presented using this condition. The stability of the scheme is discussed, and the stability restriction for the scheme is established as a function of the Mach number. Finally, comparisons are made with the frequency domain calculation by Scott and Atassi, and the relative strengths and weaknesses of each approach are assessed.
Unsteady Aerodynamic Response of a Linear Cascade of Airfoils in Separated Flow
NASA Technical Reports Server (NTRS)
Capece, Vincent R.; Ford, Christopher; Bone, Christopher; Li, Rui
2004-01-01
The overall objective of this research program was to investigate methods to modify the leading edge separation region, which could lead to an improvement in aeroelastic stability of advanced airfoil designs. The airfoil section used is representative of current low aspect ratio fan blade tip sections. The experimental potion of this study investigated separated zone boundary layer from removal through suction slots. Suction applied to a cavity in the vicinity of the separation onset point was found to be the most effective location. The computational study looked into the influence of front camber on flutter stability. To assess the influence of the change in airfoil shape on stability the work-per-cycle was evaluated for torsion mode oscillations. It was shown that the front camberline shape can be an important factor for stabilizing the predicted work-per-cycle and reducing the predicted extent of the separation zone. In addition, data analysis procedures are discussed for reducing data acquired in experiments that involve periodic unsteady data. This work was conducted in support of experiments being conducted in the NASA Glenn Research Center Transonic Flutter Cascade. The spectral block averaging method is presented. This method is shown to be able to account for variations in airfoil oscillation frequency that can occur in experiments that force oscillate the airfoils to simulate flutter.
NASA Astrophysics Data System (ADS)
Liu, Yan; Li, Kailun; Zhang, Jiazhong; Wang, Hang; Liu, Liguang
2012-08-01
By the finite element method combined with Arbitrary-Lagrangian-Eulerian (ALE) frame and explicit Characteristic Based Split Scheme (CBS), the complex flows around stationary and sinusoidal pitching airfoil are studied numerically. In particular, the static and dynamic stalls are analyzed in detail, and the natures of the static stall of NACA0012 airfoil are given from viewpoint of bifurcations. Following the bifurcation in Map, the static stall is proved to be the result from saddle-node bifurcation which involves both the hysteresis and jumping phenomena, by introducing a Map and its Floquet multiplier, which is constructed in the numerical simulation of flow field and related to the lift of the airfoil. Further, because the saddle-node bifurcation is sensitive to imperfection or perturbation, the airfoil is then subjected to a perturbation which is a kind of sinusoidal pitching oscillation, and the flow structure and aerodynamic performance are studied numerically. The results show that the large-scale flow separation at the static stall on the airfoil surface can be removed or delayed feasibly, and the ensuing lift could be enhanced significantly and also the stalling incidence could be delayed effectively. As a conclusion, it can be drawn that the proper external excitation can be considered as a powerful control strategy for the stall. As an unsteady aerodynamic behavior of high angle of attack, the dynamic stall can be investigated from viewpoint of nonlinear dynamics, and there exists a rich variety of nonlinear phenomena, which are related to the lift enhancement and drag reduction.
Pressure Distribution Over Airfoils at High Speeds
NASA Technical Reports Server (NTRS)
Briggs, L J; Dryden, H L
1927-01-01
This report deals with the pressure distribution over airfoils at high speeds, and describes an extension of an investigation of the aerodynamic characteristics of certain airfoils which was presented in NACA Technical Report no. 207. The results presented in report no. 207 have been confirmed and extended to higher speeds through a more extensive and systematic series of tests. Observations were also made of the air flow near the surface of the airfoils, and the large changes in lift coefficients were shown to be associated with a sudden breaking away of the flow from the upper surface. The tests were made on models of 1-inch chord and comparison with the earlier measurements on models of 3-inch chord shows that the sudden change in the lift coefficient is due to compressibility and not to a change in the Reynolds number. The Reynolds number still has a large effect, however, on the drag coefficient. The pressure distribution observations furnish the propeller designer with data on the load distribution at high speeds, and also give a better picture of the air-flow changes.
Three-dimensional unsteady viscous flow analysis over airfoil sections
NASA Technical Reports Server (NTRS)
Weinberg, B. C.; Shamroth, S. J.
1984-01-01
A three-dimensional solution procedure for the approximate form of the Navier-Stokes equation was exercised in the two- and three-dimensional modes to compute the unsteady turbulent boundary layer on a flat plate corresponding to the data of Karlsson. The procedure is based on the use of a consistently split Linearized Block Implicit technique in conjunction with a QR operator scheme. New time-dependent upstream boundary conditions were developed that yielded realistic solutions for the interior in the vicinity of the upstream boundary. Comparisons of the computation employing these boundary conditions with the data indicate that both qualitative and quantitative agreement was obtained for the mean velocity and the in phase and out of phase components of the first harmonic of the velocity. In addition, the calculation gave results for the skin friction phase angle that had expected physical behavior for large distances downstream of the inflow boundary. For the three-dimensional case, the two-dimensional data of Karlsson was considered, but in a coordinate system skewed at 45 deg to the free stream direction. The results of the calculations were in excellent agreement with the data and the two-dimensional computations.
Effect of pre-strain and excess length on unsteady fluid-structure interactions of membrane airfoils
NASA Astrophysics Data System (ADS)
Rojratsirikul, P.; Wang, Z.; Gursul, I.
2010-04-01
Aerodynamic characteristics of two-dimensional membrane airfoils were experimentally investigated in a wind tunnel. The effects of the membrane pre-strain and excess length on the unsteady aspects of the fluid-structure interaction were studied. The deformation of the membrane as a function of angle of attack and free-stream velocity was measured using a high-speed camera. These measurements were complemented by the measurements of unsteady velocity field with a high frame-rate Particle Image Velocimetry (PIV) system as well as smoke visualization. Membrane airfoils with excess length exhibit higher vibration modes, earlier roll-up of vortices, and smaller separated flow regions, whereas the membranes with pre-strain generally behave more similarly to a rigid airfoil. Measured frequencies of the membrane vibrations suggest a possible coupling with the wake instabilities at high incidences for all airfoils.
Transonic Aerodynamic Characteristics of Two Wedge Airfoil Sections Including Unsteady Flow Studies
NASA Technical Reports Server (NTRS)
Johnston, Patrick J.
1959-01-01
A two-dimensional wind-tunnel investigation has been conducted on a 20-percent-thick single-wedge airfoil section. Steady-state forces and moments were determined from pressure measurements at Mach numbers from 0.70 to about 1.25. Additional information on the flows about the single wedge is provided by means of instantaneous pressure measurements at Mach numbers up to unity. Pressure distributions were also obtained on a symmetrical double-wedge or diamond-shaped profile which had the same leading-edge included angle as the single-wedge airfoil. A comparison of the data on the two profiles to provide information on the effects of the afterbody showed that with the exception of drag, the single-wedge profile proved to be aerodynamically superior to the diamond profile in all respects. The lift effectiveness of the single-wedge airfoil section far exceeded that of conventional thin airfoil sections over the speed range of the investigation. Pitching-moment irregularities, caused by negative loadings near the trailing edge, generally associated with conventional airfoils of equivalent thicknesses were not exhibited by the single-wedge profile. Moderately high pulsating pressures existing over the base of the single-wedge airfoil section were significantly reduced as the Mach number was increased beyond 0.92 and the boundaries of the dead airspace at the base of the model converged to eliminate the vortex street in the wake. Increasing the leading-edge radius from 0 to 1 percent of the chord had a minor effect on the steady-state forces and generally raised the level of pressure pulsations over the forward part of the single-wedge profile.
Dynamic stall simulation of a pitching airfoil under unsteady freestream velocity
NASA Astrophysics Data System (ADS)
Gharali, Kobra; Johnson, David A.
2013-10-01
Effects of horizontal oscillations of the freestream velocity superimposed on a pitch oscillating NACA0012 airfoil were investigated using Computational Fluid Dynamics (CFD). The SST k-ω model coupled with a low-Reynolds number correction was applied for Re ≈105, when the airfoil was undergoing dynamic stall. The main parameter Φ, the phase difference between the freestream oscillation and the airfoil oscillation, was varied from 0 to π. The Φ variation resulted in several times amplitude dynamic loads when Φ≤π/2 to several times damping dynamic loads for Φ>π/2 where some dynamic stall loads were damped even below static stall load values. It was found that Φ variation was divided into two main ranges based on the values of the unsteady freestream velocity at dynamic stall. The load variation also appeared with some differences including the shape of the trailing edge vortex sheet before stall, the circulation of the dynamic stall vortex pairs, the critical angles, vortex growth time and the secondary lift peak location that are discussed in detail.
Unsteady Pressures on a Generic Capsule Shape
NASA Technical Reports Server (NTRS)
Burnside, Nathan; Ross, James C.
2015-01-01
While developing the aerodynamic database for the Orion spacecraft, the low-speed flight regime (transonic and below) proved to be the most difficult to predict and measure accurately. The flow over the capsule heat shield in descent flight was particularly troublesome for both computational and experimental efforts due to its unsteady nature and uncertainty about the boundary layer state. The data described here were acquired as part of a study to improve the understanding of the overall flow around a generic capsule. The unsteady pressure measurements acquired on a generic capsule shape are presented along with a discussion about the effects of various flight conditions and heat-shield surface roughness on the resulting pressure fluctuations.
Measuring unsteady pressure on rotating compressor blades
NASA Technical Reports Server (NTRS)
Englund, D. R.; Grant, H. P.; Lanati, G. A.
1979-01-01
Miniature semiconductor strain gage pressure transducers mounted in several arrangements were studied. Both surface mountings and recessed flush mountings were tested. Test parameters included mounting arrangement, blade material, temperature, local strain in the acceleration normal to the transducer diaphragm, centripetal acceleration, and pressure. Test results show no failures of transducers or mountings and indicate an uncertainty of unsteady pressure measurement of approximately + or - 6 percent + 0.1 kPa for a typical application. Two configurations were used on a rotating fan flutter program. Examples of transducer data and correction factors are presented.
Unsteady pressures under impinging jets in crossflows
NASA Astrophysics Data System (ADS)
Knowles, K.; Bray, D.; Wilson, M. J.
The flowfield surrounding single and twin jets impinging in a cross-flow has been investigated experimentally. Two regions of the flowfield are seen to be particularly unsteady: the ground vortex formed by separation of the wall jet in the cross-flow; and the fountain formed between two adjacent impinging jets. The ground vortex formed by twin (side-by-side) impingen jets is found to be even more unsteady than that due to a single jet. The unsteadiness has been quantified by ground plane pressure measurements. These have shown the fluctuations to be broadband in nature but with a low frequency hump. This confirms results from elsewhere using low speed jets. The present work also suggests that the peak amplitudes in ground plane pressure fluctuations occur either side of the vortex. There is, as yet, no evidence that the vortex fluctuations are fed by impinging jet instabilities. Nor does nozzle pressure ratio seem to have a direct effect on the spectrum of fluctuations.
Ristau, Neil; Siden, Gunnar Leif
2015-07-21
An airfoil includes a leading edge, a trailing edge downstream from the leading edge, a pressure surface between the leading and trailing edges, and a suction surface between the leading and trailing edges and opposite the pressure surface. A first convex section on the suction surface decreases in curvature downstream from the leading edge, and a throat on the suction surface is downstream from the first convex section. A second convex section is on the suction surface downstream from the throat, and a first convex segment of the second convex section increases in curvature.
Differential pressure sensing system for airfoils usable in turbine engines
Yang, Wen-Ching; Stampahar, Maria E.
2005-09-13
A detection system for identifying airfoils having a cooling systems with orifices that are plugged with contaminants or with showerheads having a portion burned off. The detection system measures pressures at different locations and calculates or measures a differential pressure. The differential pressure may be compared with a known benchmark value to determine whether the differential pressure has changed. Changes in the differential pressure may indicate that one or more of the orifices in a cooling system of an airfoil are plugged or that portions of, or all of, a showerhead has burned off.
Technology for pressure-instrumented thin airfoil models
NASA Technical Reports Server (NTRS)
Wigley, David A.
1988-01-01
A novel method of airfoil model construction was developed. This Laminated Sheet technique uses 0.8 mm thick sheets of A286 containing a network of pre-formed channels which are vacuum brazed together to form the airfoil. A 6.25 percent model of the X29A canard, which has a 5 percent thick section, was built using this technique. The model contained a total of 96 pressure orifices, 56 in three chordwise rows on the upper surface and 37 in three similar rows on the lower surface. It was tested in the NASA Langley 0.3 m Transonic Cryogenic Tunnel. Unique aerodynamic data was obtained over the full range of temperature and pressure. Part of the data was at transonic Mach numbers and flight Reynolds number. A larger two dimensional model of the NACA 64a-105 airfoil section was also fabricated. Scale up presented some problems, but a testable airfoil was fabricated.
Unsteady separation experiments on 2-D airfoils, 3-D wings, and model helicopter rotors
NASA Technical Reports Server (NTRS)
Lorber, Peter F.; Carta, Franklin O.
1992-01-01
Information on unsteady separation and dynamic stall is being obtained from two experimental programs that have been underway at United Technologies Research Center since 1984. The first program is designed to obtain detailed surface pressure and boundary layer condition information during high amplitude pitching oscillations of a large (17.3 in. chord) model wing in a wind tunnel. The second program involves the construction and testing of a pressure-instrumented model helicopter rotor. This presentation describes some of the results of these experiments, and in particular compares the detailed dynamic stall inception information obtained from the oscillating wing with the unsteady separation and reverse flow results measured on the retreating blade side of the model rotor during wind tunnel testing.
NASA Technical Reports Server (NTRS)
Bartels, Robert E.; Edwards, John W.
1997-01-01
Steady and unsteady experimental data are presented for several fixed geometry conditions from a test in the NASA Langley 0.3-Meter Transonic Cryogenic Tunnel. The purpose of this test was to obtain unsteady data for transonic conditions on a fixed and pitching supercritical airfoil at high Reynolds numbers. Data and brief analyses for several of the fixed geometry test conditions will be presented here. These are at Reynolds numbers from 6 x 10(exp 6) to 35 x 10(exp 6) bases on chord length, and span a limited range of Mach numbers and angles of attack just below and at the onset of shock buffet. Reynolds scaling effects appear in both the steady pressure data and in the onset of shock buffet at Reynolds numbers of 15 x 10(exp 6) and 3O x 10(exp 6) per chord length.
Nonlinear switched models for control of unsteady forces on a rapidly pitching airfoil
NASA Astrophysics Data System (ADS)
Dawson, Scott; Brunton, Steven; Rowley, Clarence
2013-11-01
The unsteady aerodynamic forces incident on a pitching flat plate airfoil at a Reynolds number of 100 are investigated through direct numerical simulation. Linear state-space models, identified from impulse response data via the eigensystem realization algorithm, are used to accurately track rapid changes in lift coefficient through either feedback or feedforward control, even in the presence of gust disturbances. We develop a technique to project between states of linear models obtained at different angles of attack using primal and pseudo-adjoint balanced POD modes. This allows for the formation of a nonlinear switched model that is accurate over a wide range of angles of attack, in both pre- and post-stall regimes. We additionally investigate phenomena that are not captured by linear models, such as an increase in mean lift that occurs when vortex shedding frequencies are excited. The effect of changing the pitch axis is also investigated, where it is found that pitching aft of the mid-chord results in right half plane zeros that increase the difficulty of the control problem. This work was supported by AFOSR grant FA9550-12-1-0075.
Airfoil shape and thickness effects on transonic airloads and flutter
NASA Technical Reports Server (NTRS)
Bland, S. R.; Edwards, J. W.
1983-01-01
A transient pulse technique is used to obtain harmonic forces from a time-marching solution of the complete unsteady transonic small perturbation potential equation. The unsteady pressures and forces acting on a model of the NACA 64A010 conventional airfoil and the MBB A-3 supercritical airfoil over a range of Mach numbers are examined in detail. Flutter calculations at constant angle of attack show a similar flutter behavior for both airfoils, except for a boundary shift in Mach number associated with corresponding Mach number shift in the unsteady aerodynamic forces. Differences in the static aeroelastic twist behavior for the two airfoils are significant.
Airfoil shape and thickness effects on transonic airloads and flutter
NASA Technical Reports Server (NTRS)
Bland, S. R.; Edwards, J. W.
1983-01-01
A transient pulse technique is used to obtain harmonic forces from a time-marching solution of the complete unsteady transonic small perturbation potential evaluation. The unsteady pressures and forces acting on a model of the NACA 64A010 conventional airfoil and the MBB A-3 supercritical airfoil over a range of Mach numbers are examined in detail. Flutter calculations at constant angle of attack show a similar flutter behavior for both airfoils, except for a boundary shift in Mach number associated with a corresponding Mach number shift in the unsteady aerodynamic forces. Differences in the static aeroelastic twist behavior for the two airfoils are significant.
NASA Astrophysics Data System (ADS)
Disotell, Kevin J.; Nikoueeyan, Pourya; Naughton, Jonathan W.; Gregory, James W.
2016-05-01
Recognizing the need for global surface measurement techniques to characterize the time-varying, three-dimensional loading encountered on rotating wind turbine blades, fast-responding pressure-sensitive paint (PSP) has been evaluated for resolving unsteady aerodynamic effects in incompressible flow. Results of a study aimed at demonstrating the laser-based, single-shot PSP technique on a low Reynolds number wind turbine airfoil in static and dynamic stall are reported. PSP was applied to the suction side of a Delft DU97-W-300 airfoil (maximum thickness-to-chord ratio of 30 %) at a chord Reynolds number of 225,000 in the University of Wyoming open-return wind tunnel. Static and dynamic stall behaviors are presented using instantaneous and phase-averaged global pressure maps. In particular, a three-dimensional pressure topology driven by a stall cell pattern is detected near the maximum lift condition on the steady airfoil. Trends in the PSP-measured pressure topology on the steady airfoil were confirmed using surface oil visualization. The dynamic stall case was characterized by a sinusoidal pitching motion with mean angle of 15.7°, amplitude of 11.2°, and reduced frequency of 0.106 based on semichord. PSP images were acquired at selected phase positions, capturing the breakdown of nominally two-dimensional flow near lift stall, development of post-stall suction near the trailing edge, and a highly three-dimensional topology as the flow reattaches. Structural patterns in the surface pressure topologies are considered from the analysis of the individual PSP snapshots, enabled by a laser-based excitation system that achieves sufficient signal-to-noise ratio in the single-shot images. The PSP results are found to be in general agreement with observations about the steady and unsteady stall characteristics expected for the airfoil.
NASA Technical Reports Server (NTRS)
St.hilaire, A. O.; Carta, F. O.
1983-01-01
The unsteady chordwise force response on the airfoil surface was investigated and its sensitivity to the various system parameters was examined. A further examination of unsteady aerodynamic data on a tunnel spanning wing (both swept and unswept), obtained in a wind tunnel, was performed. The main body of this data analysis was carried out by analyzing the propagation speed of pressure disturbances along the chord and by studying the behavior of the unsteady part of the chordwise pressure distribution at various points of the airfoil pitching cycle. It was found that Mach number effects dominate the approach to and the inception of both static and dynamic stall. The stall angle decreases as the Mach number increases. However, sweep dominates the load behavior within the stall regime. Large phase differences between unswept and swept responses, that do not exist at low lift coefficient, appear once the stall boundary is penetrated. It was also found that reduced frequency is not a reliable indicator of the unsteady aerodynamic response in the high angle of attack regime.
Estimation of unsteady lift on a pitching airfoil from wake velocity surveys
NASA Technical Reports Server (NTRS)
Zaman, K. B. M. Q.; Panda, J.; Rumsey, C. L.
1993-01-01
The results of a joint experimental and computational study on the flowfield over a periodically pitched NACA0012 airfoil, and the resultant lift variation, are reported in this paper. The lift variation over a cycle of oscillation, and hence the lift hysteresis loop, is estimated from the velocity distribution in the wake measured or computed for successive phases of the cycle. Experimentally, the estimated lift hysteresis loops are compared with available data from the literature as well as with limited force balance measurements. Computationally, the estimated lift variations are compared with the corresponding variation obtained from the surface pressure distribution. Four analytical formulations for the lift estimation from wake surveys are considered and relative successes of the four are discussed.
NASA Technical Reports Server (NTRS)
Crimi, P.
1974-01-01
A method for analyzing unsteady airfoil stall was refined by including nonlinear effects in the representation of the inviscid flow. Certain other aspects of the potential-flow model were reexamined and the effects of varying Reynolds number on stall characteristics were investigated. Refinement of the formulation improved the representation of the flow and chordwise pressure distribution below stall, but substantial quantitative differences between computed and measured results are still evident for sinusoidal pitching through stall. Agreement is substantially improved by assuming the growth rate of the dead-air region at the onset of leading-edge stall is of the order of the component of the free stream normal to the airfoil chordline. The method predicts the expected increase in the resistance to stalling with increasing Reynolds number. Results indicate that a given airfoil can undergo both trailing-edge and leading-edge stall under unsteady conditions.
Non-linear unsteady wing theory, part 1. Quasi two-dimensional behavior: Airfoils and slender wings
NASA Technical Reports Server (NTRS)
Mccune, J. E.
1987-01-01
The initial phases of a study of the large-amplitude unsteady aerodynamics of wings in severe maneuver are reported. The research centers on vortex flows, their initiation at wing surfaces, their subsequent convection, and interaction dynamically with wings and control surfaces. The focus is on 2D and quasi-2D aspects of the problem and features the development of an exact nonlinear unsteady airfoil theory as well as an approach to the crossflow problem for slender wing applications including leading-edge separation. The effective use of interactive on-line computing in quantifying and visualizing the nonsteady effects of severe maneuver is demonstrated. Interactive computational work is now possible, in which a maneuver can be initiated and its effects observed and analyzed immediately.
NASA Technical Reports Server (NTRS)
Envia, Edmane
2000-01-01
The solution to the problem of unsteady response of an isolated finite span swept airfoil to an incident gust is most conveniently obtained by introducing an acoustic velocity potential and eliminating all primitive variables from the governing equations. The end result is the convected wave equation for the acoustic velocity potential, phi. For the unsteady response of a rectilinear swept cascade to an incident gust problem, the governing equation and the imposed boundary conditions are the same as those for the isolated airfoil problem.
Porous plug for reducing orifice induced pressure error in airfoils
NASA Technical Reports Server (NTRS)
Plentovich, Elizabeth B. (Inventor); Gloss, Blair B. (Inventor); Eves, John W. (Inventor); Stack, John P. (Inventor)
1988-01-01
A porous plug is provided for the reduction or elimination of positive error caused by the orifice during static pressure measurements of airfoils. The porous plug is press fitted into the orifice, thereby preventing the error caused either by fluid flow turning into the exposed orifice or by the fluid flow stagnating at the downstream edge of the orifice. In addition, the porous plug is made flush with the outer surface of the airfoil, by filing and polishing, to provide a smooth surface which alleviates the error caused by imperfections in the orifice. The porous plug is preferably made of sintered metal, which allows air to pass through the pores, so that the static pressure measurements can be made by remote transducers.
NASA Technical Reports Server (NTRS)
Rivera, Jose A., Jr.; Dansberry, Bryan E.; Farmer, Moses G.; Eckstrom, Clinton V.; Seidel, David A.; Bennett, Robert M.
1991-01-01
The Structural Dynamics Div. at NASA-Langley has started a wind tunnel activity referred to as the Benchmark Models Program. The objective is to acquire test data that will be useful for developing and evaluating aeroelastic type Computational Fluid Dynamics codes currently in use or under development. The progress is described which was achieved in testing the first model in the Benchmark Models Program. Experimental flutter boundaries are presented for a rigid semispan model (NACA 0012 airfoil section) mounted on a flexible mount system. Also, steady and unsteady pressure measurements taken at the flutter condition are presented. The pressure data were acquired over the entire model chord located at the 60 pct. span station.
NASA Technical Reports Server (NTRS)
Kleb, William L.; Batina, John T.; Williams, Marc H.
1990-01-01
A temporal adaptive algorithm for the time-integration of the two-dimensional Euler or Navier-Stokes equations is presented. The flow solver involves an upwind flux-split spatial discretization for the convective terms and central differencing for the shear-stress and heat flux terms on an unstructured mesh of triangles. The temporal adaptive algorithm is a time-accurate integration procedure which allows flows with high spatial and temporal gradients to be computed efficiently by advancing each grid cell near its maximum allowable time step. Results indicate that an appreciable computational savings can be achieved for both inviscid and viscous unsteady airfoil problems using unstructured meshes without degrading spatial or temporal accuracy.
Numerical Study of Unsteady Supercavitation Perturbed by a Pressure Wave
NASA Astrophysics Data System (ADS)
Zheng, J. G.; Khoo, B. C.
2016-06-01
The unsteady features of supercavitation disturbed by an introduced pressure wave are investigated numerically using a one-fluid cavitation model. The supercavitating flow is assumed to be the homogeneous mixture of liquid and vapour which are locally under both kinetic and thermodynamic equilibrium. The compressibility effects of liquid water are taken into account to model the propagation of pressure wave through flow and its interaction with supercavitation bubble. The interaction between supercavity enveloping an underwater flat-nose cylinder and pressure wave is simulated and the resulting unsteady behavior of supercavitation is illustrated. It is observed that the supercavity will become unstable under the impact of the pressure wave and may collapse locally, which depends on the strength of perturbation. The huge pressure surge accompanying the collapse of supercavitation may cause the material erosion, noise, vibration and efficiency loss of operating underwater devices.
Unsteady pressure loads in a generic high speed engine model
NASA Technical Reports Server (NTRS)
Parrott, Tony L.; Jones, Michael G.; Thurlow, Ernie M.
1992-01-01
Unsteady pressure loads were measured along the top interior wall of a generic high-speed engine (GHSE) model undergoing performance tests in the combustion-Heated Scramjet Test Facility at the Langley Research Center. Flow to the model inlet was simulated at 72000 ft and a flight Mach number of 4. The inlet Mach number was 3.5 with a total temperature and pressure of 1640 R and 92 psia. The unsteady pressure loads were measured with 5 piezoresistive gages, recessed into the wall 4 to 12 gage diameters to reduce incident heat flux to the diaphragms, and distributed from the inlet to the combustor. Contributors to the unsteady pressure loads included boundary layer turbulence, combustion noise, and transients generated by unstart loads. Typical turbulent boundary layer rms pressures in the inlet ranged from 133 dB in the inlet to 181 dB in the combustor over the frequency range from 0 to 5 kHz. Downstream of the inlet exist, combustion noise was shown to dominate boundary layer turbulence noise at increased heat release rates. Noise levels in the isolator section increased by 15 dB when the fuel-air ratio was increased from 0.37 to 0.57 of the stoichiometric ratio. Transient pressure disturbances associated with engine unstarts were measured in the inlet and have an upstream propagation speed of about 7 ft/sec and pressure jumps of at least 3 psia.
Unsteady flows in a two-dimensional linear cascade with low-pressure turbine blades
NASA Astrophysics Data System (ADS)
Murawski, Christopher Gabriel
Experimental studies of unsteady flow phenomena in a low pressure turbine linear cascade are presented. Turbine engine flow passages contain numerous loss mechanisms. The loss mechanisms investigated in this study are low Reynolds number and freestream turbulence effects, secondary flows and wake interactions. Also, a method is implemented which decreases the profile losses due to low Reynolds number effects. The results are presented in three segments. First, the effects of Reynolds number and freestream turbulence intensity on the low-pressure turbine cascade blade are investigated. The condition of the blade's boundary layer is the leading factor controlling the level of profile loss. The losses from the airfoil decrease as the Reynolds number and freestream turbulence increase due to a decrease in the size of the separation zone on the suction side of the turbine airfoil. Boundary layer separation occurs on the suction surface of the turbine. Changes to this region are achieved when attaching different length tail sets to the turbine airfoils which alters the axial chord of each blade. A clear improvement on suction side boundary layer behavior at low Reynolds numbers was seen when the tail extensions were shorter than about 9% of axial chord. Finally, the effect wake disturbance frequency on the secondary flow vortex structure in a turbine cascade is studied. Cylinders are traversed across the front of the blade row to simulate turbine blade disturbances. The response of the secondary flow structure to the movement of the wake generator shuttle with zero, one and multiple wake generator rods are presented. Multiple wake disturbance frequencies are varied from 12 Hz to 52 Hz. Multiple wake disturbance frequency below the axial chord flow frequency enable the secondary flow vortex structure to re-establish itself between each wake disturbance event. Axial chord flow frequency is defined as the axial velocity in the cascade divided by the axial chord length of
NASA Technical Reports Server (NTRS)
Mcalister, K. W.; Pucci, S. L.; Mccroskey, W. J.; Carr, L. W.
1982-01-01
Experimentally derived force and moment data are presented for eight airfoil sections that were tested at fixed and varying incidence in a subsonic two dimensional stream. Airfoil incidence was varied through sinusoidal oscillations in pitch over a wide range of amplitude and frequency. The surface pressure distribution, as well as the lift, drag, and pitching moment derived therefrom, are displayed in a uniform fashion to delineate the static and dynamic characteristics of each airfoil both in and out of stall.
Technology for pressure-instrumented thin airfoil models, phase 1
NASA Technical Reports Server (NTRS)
Wigley, D. A.
1985-01-01
A network of channels was chemically milled into one surface of a pair of matched plates having bond planes which were neither planar or profiled to match the contour of the trailing edge of a supercritical airfoil for testing in cryogenic wind tunnels. Vacuum brazing bonded the plates together to create a network of pressure passages without blockages or cross leaks. The greatest success was achieved with the smaller samples and planar bonding surfaces. In larger samples, problems were encountered due to warpage created by the relief of residual stresses. Successful bonds were formed by brazing A286, Nitronic 40 and 300 series stainless steels at 1065 C using AMS 4777B brazing alloy, but excessive grain growth occurred in samples of 200 grade 18 nickel maraging steels. Good bonds were obtained with maraging steel using a 47 percent Nickel-47 percent Palladium-6 percent Silicon alloy and brazing at 927 C. Electro-Discharge-Machining was an effective method of cutting profiled bond planes and airfoil contours. Orifices of good definition were obtained when the EDM wire cut passed through predrilled holes. Possible configurations for joints between small segments and the larger main wing were also studied.
Unsteady wing surface pressures in the wake of a propeller
NASA Technical Reports Server (NTRS)
Johnston, R. T.; Sullivan, J. P.
1992-01-01
The unsteady nature of the propeller slipstream interacting with a wing has been studied by flow visualization and unsteady wing surface pressure measurements. Flow visualization was performed by marking the propeller tip vortex with smoke. Unsteady wing surface pressures were measured by traversing a wing instrumented with a chordwise array of 16 microphones in a spanwise direction through the propeller wake. This work yielded information on the motion of the propeller wake as it passes over the wing. As the propeller wake passed over the wing: the propeller tip vortex experienced an inviscid interaction at the leading edge; viscous action at the leading edge severed the propeller tip vortex; the propeller tip vortex experienced significant spanwise and chordwise displacements and then deformed in order to reconnect at the trailing edge; axial velocity in the vortex core caused the helical vortex to thicken or stretch near the wing surface; and, the magnitude of the pressure fluctuations decreased in magnitude with distance traveled along the chord.
Flow Control on Low-Pressure Turbine Airfoils Using Vortex Generator Jets
NASA Technical Reports Server (NTRS)
Volino, Ralph J.; Ibrahim, Mounir B.; Kartuzova, Olga
2010-01-01
Motivation - Higher loading on Low-Pressure Turbine (LPT) airfoils: Reduce airfoil count, weight, cost. Increase efficiency, and Limited by suction side separation. Growing understanding of transition, separation, wake effects: Improved models. Take advantage of wakes. Higher lift airfoils in use. Further loading increases may require flow control: Passive: trips, dimples, etc. Active: plasma actuators, vortex generator jets (VGJs). Can increased loading offset higher losses on high lift airfoils. Objectives: Advance knowledge of boundary layer separation and transition under LPT conditions. Demonstrate, improve understanding of separation control with pulsed VGJs. Produce detailed experimental data base. Test and develop computational models.
NASA Technical Reports Server (NTRS)
Liu, Yi; Sankar, Lakshmi N.; Englar, Robert J.; Ahuja, Krishan K.
2003-01-01
The aerodynamic characteristics of a Circulation Control Wing (CCW) airfoil have been numerically investigated, and comparisons with experimental data have been made. The configuration chosen was a supercritical airfoil with a 30 degree dual-radius CCW flap. Steady and pulsed jet calculations were performed. It was found that the use of steady jets, even at very small mass flow rates, yielded a lift coefficient that is comparable or superior to conventional high-lift systems. The attached flow over the flap also gave rise to lower drag coefficients, and high L/D ratios. Pulsed jets with a 50% duty cycle were also studied. It was found that they were effective in generating lift at lower reduced mass flow rates compared to a steady jet, provided the pulse frequency was sufficiently high. This benefit was attributable to the fact that the momentum coefficient of the pulsed jet, during the portions of the cycle when the jet was on, was typically twice as much as that of a steady jet.
NASA Technical Reports Server (NTRS)
Cazier, F. W., Jr.; Watson, J. J.; Doggett, R. V., Jr.; Sandford, M. C.; Ricketts, R. H.
1981-01-01
Highlight results are presented from subsonic and transonic pressure measurement studies conducted in the Langley Transonic Dynamics Tunnel on a supercritical wing model representative of an energy efficient transport design. Steady- and unsteady-pressure data were acquired on the upper and lower wing surface at an off-design Mach number of 0.60 and at the design Mach number of 0.78, for a Reynolds number of 2.2 x 10(6) (based on the wing average chord). The model configuration consisted of a sidewall-Mounted half-body fuselage and a semi-span wing with an aspect ratio of 10.76, a leading-edge sweepback angle of 28.8 degrees, and supercritical airfoil sections. The wing is instrumented with 252 static pressure orifices and 164 dynamic pressure gages. Model test variables included wing angle of attack, control-surface mean deflection angle, control-surface oscillating deflection angle and frequency, and phasing between oscillating leading-edge and trailing-edge controls when used together.
NASA Astrophysics Data System (ADS)
Gordeyev, Stanislav; De Lucca, Nicholas; Jumper, Eric J.; Hird, Kyle; Juliano, Thomas J.; Gregory, James W.; Thordahl, James; Wittich, Donald J.
2014-01-01
Spatially temporally resolved unsteady pressure fields on a surface of a hemisphere-on-cylinder turret with either a flat or a conformal window with realistic features such as gaps and "smile" cutouts were characterized using fast-response pressure-sensitive paint at M = 0.33 for several window viewing angles. Various statistical properties of pressure fields were computed, and geometry effects on the unsteady pressure fields were analyzed and discussed. Proper orthogonal decomposition was also used to extract dominant pressure modes and corresponding temporal coefficients and to analyze and compare instantaneous pressure structures for different turret geometric features and the window viewing angles. An unsteady separation off the turret and a recirculation region downstream of the turret were identified as dominant sources of the unsteady pressure. It was found that while all geometric features affected the unsteady pressure field, the "smiles," positioned spanwise-symmetrically on both sides of the turret, were the leading cause of these changes, followed by the looking forward flat window. The gaps, the side- and the back-looking flat window introduced only small local changes.
Some observations of surface pressures and the near wake of a blunt trailing edge airfoil
NASA Technical Reports Server (NTRS)
Digumarthi, R. V.; Koutsoyannis, S. P.; Karamcheti, K.
1981-01-01
Experiments with a truncated and untruncated airfoils of profiles NACA 640A10, were carried out in subsonic wind tunnels in a velocity range of 19m/s to 54m/s corresponding to Reynolds numbers of 200,000 to 468,000 based on the chord. Airfoil spanned the test section to achieve two dimensionality of the model. Velocity measurements, pressure measurements, and vortex shedding in the wake were measured using a hotwire and pressure transducers. The measured chordwise static pressure distribution on the smooth trailing edge airfoil along the midspan plane, agreed with the theoretical results calculated on the basis of the potential flow for that airfoil. Boundary layer profiles measured in the midspan plane, behind the maximum thickness of the airfoil show no separation of the flow. Spanwise distribution of the measured static pressure on the upper surface of the airfoil shows uniformity for both configurations with and without the boundary layer trip. This uniformity of pressure distribution and separation indicates that the flow on the airfoil was uniform and two dimensional in character.
Cavity Unsteady-Pressure Measurements at Subsonic and Transonic Speeds
NASA Technical Reports Server (NTRS)
Tracy, Maureen B.; Plentovich, E. B.
1997-01-01
An experimental investigation was conducted in the Langley 8-Foot Transonic Pressure Tunnel to determine the flow characteristics of rectangular cavities with varying relative dimensions at subsonic and transonic speeds. Cavities were tested with width-to-depth ratios of 1, 4, 8, and 16 for length-to-depth ratios l/h of 1 through 17.5. The maximum cavity depth was 2.4 in., and the turbulent boundary layer approaching the cavity was approximately 0.5 in. thick. Unsteady- and mean static-pressure measurements were made at free-stream Mach numbers from 0.20 to 0.95 at a unit Reynolds number per foot of approximately 3 x 10(exp 6); however, only unsteady-pressure results are presented in this paper. Results indicate that as l/h increases, cavity flows changed from resonant to nonresonant with resonant amplitudes decreasing gradually. Resonant spectra are obtained largely in cavities with mean static-pressure distributions characteristic of open and transitional flows. Resonance sometimes occurred for closed flow. Increasing cavity width or decreasing cavity depth while holding l/h fixed had the effect of increasing resonant amplitudes and sometimes induced resonance. The effects due to changes in width are more pronounced. Decreasing Mach number has the effect of broadening the resonances.
Pressure distribution in unsteady sink and source flows.
Voropayev, S I
2015-05-01
Basic flow generated in a viscous incompressible fluid by a "point" sink (source) of mass is revised. In practice, such flow can be modeled by sucking (pushing) fluid from a thin tube with a small porous sphere at one end. Intuitively, by sucking (pushing) fluid, one creates low (high) pressure near the origin and a positive (negative) radial pressure gradient drives the fluid to (from) the origin. A simple analysis, however, shows that the pressure distribution for both steady flows is the same. Then a question arises: How does the fluid "know" in what direction to flow? To explain this "paradox" an unsteady flow is considered and the pressure terms responsible for the flow direction are derived. PMID:26066255
Fan Noise Source Diagnostic Test: Vane Unsteady Pressure Results
NASA Technical Reports Server (NTRS)
Envia, Edmane
2002-01-01
To investigate the nature of fan outlet guide vane pressure fluctuations and their link to rotor-stator interaction noise, time histories of vane fluctuating pressures were digitally acquired as part of the Fan Noise Source Diagnostic Test. Vane unsteady pressures were measured at seven fan tip speeds for both a radial and a swept vane configuration. Using time-domain averaging and spectral analysis, the blade passing frequency (BPF) harmonic and broadband contents of the vane pressures were individually analyzed. Significant Sound Pressure Level (SPL) reductions were observed for the swept vane relative to the radial vane for the BPF harmonics of vane pressure, but vane broadband reductions due to sweep turned out to be much smaller especially on an average basis. Cross-correlation analysis was used to establish the level of spatial coherence of broadband pressures between different locations on the vane and integral length scales of pressure fluctuations were estimated from these correlations. Two main results of this work are: (1) the average broadband level on the vane (in dB) increases linearly with the fan tip speed for both the radial and swept vanes, and (2) the broadband pressure distribution on the vane is nearly homogeneous and its integral length scale is a monotonically decreasing function of fan tip speed.
Initial Assesment of Space Launch System Transonic Unsteady Pressure Environment
NASA Technical Reports Server (NTRS)
Sekula, Martin K.; Piatak, David J.; Rausch, Russ D.; Florance, James R.; Ramey, James M.
2015-01-01
A series of wind tunnel tests were conducted at the NASA Langley Research Center Transonic Dynamics Tunnel to assess the transonic buffet environment for the Space Launch System (SLS) launch vehicle. An initial test, conducted in 2012, indicated an elevated buffet environment prompting a second test to provide further insight into the buffet phenomena and assess potential solutions to reduce the response levels of these environments. During the course of the test program, eight variants of the SLS-10000 configuration were examined. The effect of these configuration variants on the coefficient of the root-mean-square fluctuation of pressure about the mean as a function of test condition indicates that the maximum fluctuating pressure levels are extremely sensitive to the geometry of the forward attachment of the solid rocket boosters (SRBs) to the SLS Core. The addition of flow fences or changes to the SRB nose cone geometry can alleviate the unsteady pressure environment.
Pressure distribution over an NACA 23012 airfoil with a slotted and a plain flap
NASA Technical Reports Server (NTRS)
Wenzinger, Carl J; Delano, James B
1938-01-01
Report presents the results of pressure-distribution of an NACA 23012 airfoil equipped with a slotted flap and with a plain flap conducted in the 7 by 10-foot wind tunnel. A test installation was used in which the 7-foot-span airfoil was mounted vertically between the upper and lower sides of the closed test section so that two-dimensional flow was approximated. The pressures were measured on the upper and lower surfaces at one chord section both on the main airfoil and on the flaps for several different flap deflections and at several angles of attack. The data are presented in the form of pressure-distribution diagrams and as graphs of calculated section coefficients for the airfoil-and-flap combinations and also for the flaps alone. The results are useful for application to rib and flap structural design; in addition, the plain-flap data furnish considerable information applicable to the structural design of plain ailerons.
First-stage high pressure turbine bucket airfoil
Brown, Theresa A.; Ahmadi, Majid; Clemens, Eugene; Perry, II, Jacob C.; Holiday, Allyn K.; Delehanty, Richard A.; Jacala, Ariel Caesar
2004-05-25
The first-stage buckets have airfoil profiles substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in Table I wherein Z is a perpendicular distance from a plane normal to a radius of the turbine centerline and containing the X and Y values with the Z value commencing at zero in the X, Y plane at the radially innermost aerodynamic section of the airfoil and X and Y are coordinates defining the airfoil profile at each distance Z. The X, Y and Z values may be scaled as a function of the same constant or number to provide a scaled-up or scaled-down airfoil section for the bucket.
Pressure distribution over an airfoil section with a flap and tab
NASA Technical Reports Server (NTRS)
Wenzinger, Carl J
1937-01-01
Report presents the results of wind tunnel tests made in the NACA 7 by 10-foot wind tunnel of a Clark Y airfoil with a flap and an inset tab. The pressures were measured on both the upper and lower surfaces at one chord section. Calculations were made of the normal-force and pitching-moment coefficients of the airfoil section with flap section with tab, and the normal-force and hinge-moments coefficients of the tab alone. In addition, comparisons were made of the theoretical and experimental values for an airfoil with a multiply hinged flap system.
NASA Astrophysics Data System (ADS)
Ferreira, C.; Gonzalez, A.; Baldacchino, D.; Aparicio, M.; Gómez, S.; Munduate, X.; Garcia, N. R.; Sørensen, J. N.; Jost, E.; Knecht, S.; Lutz, T.; Chassapogiannis, P.; Diakakis, K.; Papadakis, G.; Voutsinas, S.; Prospathopoulos, J.; Gillebaart, T.; van Zuijlen, A.
2016-09-01
The FP7 AdVanced Aerodynamic Tools for lArge Rotors - Avatar project aims to develop and validate advanced aerodynamic models, to be used in integral design codes for the next generation of large scale wind turbines (10-20MW). One of the approaches towards reaching rotors for 10-20MW size is the application of flow control devices, such as flaps. In Task 3.2: Development of aerodynamic codes for modelling of flow devices on aerofoils and, rotors of the Avatar project, aerodynamic codes are benchmarked and validated against the experimental data of a DU95W180 airfoil in steady and unsteady flow, for different angle of attack and flap settings, including unsteady oscillatory trailing-edge-flap motion, carried out within the framework of WP3: Models for Flow Devices and Flow Control, Task 3.1: CFD and Experimental Database. The aerodynamics codes are: AdaptFoil2D, Foil2W, FLOWer, MaPFlow, OpenFOAM, Q3UIC, ATEFlap. The codes include unsteady Eulerian CFD simulations with grid deformation, panel models and indicial engineering models. The validation cases correspond to 18 steady flow cases, and 42 unsteady flow cases, for varying angle of attack, flap deflection and reduced frequency, with free and forced transition. The validation of the models show varying degrees of agreement, varying between models and flow cases.
Wall Pressure Unsteadiness and Side Loads in Overexpanded Rocket Nozzles
NASA Technical Reports Server (NTRS)
Baars, Woutijn J.; Tinney, Charles E.; Ruf, Joseph H.; Brown, Andrew M.; McDaniels, David M.
2012-01-01
Surveys of both the static and dynamic wall pressure signatures on the interior surface of a sub-scale, cold-flow and thrust optimized parabolic nozzle are conducted during fixed nozzle pressure ratios corresponding to FSS and RSS states. The motive is to develop a better understanding for the sources of off-axis loads during the transient start-up of overexpanded rocket nozzles. During FSS state, pressure spectra reveal frequency content resembling SWTBLI. Presumably, when the internal flow is in RSS state, separation bubbles are trapped by shocks and expansion waves; interactions between the separated flow regions and the waves produce asymmetric pressure distributions. An analysis of the azimuthal modes reveals how the breathing mode encompasses most of the resolved energy and that the side load inducing mode is coherent with the response moment measured by strain gauges mounted upstream of the nozzle on a flexible tube. Finally, the unsteady pressure is locally more energetic during RSS, albeit direct measurements of the response moments indicate higher side load activity when in FSS state. It is postulated that these discrepancies are attributed to cancellation effects between annular separation bubbles.
Ultrafast Time Response Pressure-Sensitive Paint for Unsteady Shock-Wave Research
NASA Astrophysics Data System (ADS)
Numata, Daiju; Asai, Keisuke
Pressure-Sensitive Paint (PSP) is an optical pressure measurement technique widely used in aerodynamic experiments, and has been applied to unsteady shock-wave phenomena [1, 2]. However, one of the largest problems to apply PSP to high-speed and unsteady phenomena is the response time of PSP.
NASA Technical Reports Server (NTRS)
Ricketts, R. H.; Watson, J. J.; Sandford, M. C.; Seidel, D. A.
1983-01-01
Wind-tunnel tests to measure unsteady aerodynamic data in the transonic region have been completed on an aspect ratio 2.0 rectangular wing with a supercritical airfoil. The geometric and structural properties of the wing are presented. (Other references contain the measured aerodynamic data.) Both measured and design airfoil coordinates are presented and compared. In addition, measured wing bending and torsional stiffness distributions and some trailing-edge flexibility influence coefficients are presented.
Broadband Noise Predictions for an Airfoil in a Turbulent Stream
NASA Technical Reports Server (NTRS)
Casper, J.; Farassat, F.; Mish, P. F.; Devenport, W. J.
2003-01-01
Loading noise is predicted from unsteady surface pressure measurements on a NACA 0015 airfoil immersed in grid-generated turbulence. The time-dependent pressure is obtained from an array of synchronized transducers on the airfoil surface. Far field noise is predicted by using the time-dependent surface pressure as input to Formulation 1A of Farassat, a solution of the Ffowcs Williams - Hawkings equation. Acoustic predictions are performed with and without the effects of airfoil surface curvature. Scaling rules are developed to compare the present far field predictions with acoustic measurements that are available in the literature.
Unsteady blade pressure measurements on a model counterrotation propeller
NASA Technical Reports Server (NTRS)
Heidelberg, Laurence J.; Woodward, Richard P.
1989-01-01
In an exploratory effort an advanced counterrotation propeller instrumented with blade-mounted pressure transducers was tested in the NASA Lewis 9- by 15-Foot Anechoic Wind Tunnel at a simulated takeoff and landing speed of Mach 0.20. The propeller's aft diameter was reduced to investigate possible noise reductions resulting from reduced blade row interaction with the tip vortex. The propeller was tested at three blade row spacings at fixed blade setting angles, at the maximum blade row spacing at higher blade setting angles and at propeller axis angles attack to the flow up to + or - 16 deg. A limited number of unsteady blade surface pressure measurements were made on both rotors of the model counterrotation propeller. Emphasis was placed on determining the effects of rotor-rotor interactions on the blade surface pressures. A unique method of processing the pressure signals was developed that enables even weak interaction waveforms and spectra to be separated from the total signal. The interaction on the aft rotor was many times stronger than that on the forward rotor. The fundamental rotor interaction tone exhibited complicated behavior but generally increased with rotational speed and blade setting angle and decreased with rotor spacing. With the propeller axis at an angle to the flow, the phase response of the aft rotor appeared to be significantly affected by the presence of the forward rotor.
Unsteady aerodynamics of an oscillating cascade in a compressible flow field
NASA Technical Reports Server (NTRS)
Buffum, Daniel H.; Boldman, Donald R.; Fleeter, Sanford
1987-01-01
Fundamental experiments were performed in the NASA Lewis Transonic Oscillating Cascade Facility to investigate and quantify the unsteady aerodynamics of a cascade of biconvex airfoils executing torsion-mode oscillations at realistic reduced frequencies. Flush-mounted, high-response miniature pressure transducers were used to measure the unsteady airfoil surface pressures. The pressures were measured for three interblade phase angles at two inlet Mach numbers, 0.65 and 0.80, and two incidence angles, 0 and 7 deg. The time-variant pressures were analyzed by means of discrete Fourier transform techniques, and these unique data were then compared with predictions from a linearized unsteady cascade model. The experimental results indicate that the interblade phase angle had a major effect on the chordwise distributions of the airfoil surface unsteady pressure, and that reduced frequency, incidence angle, and Mach number had a somewhat less significant effect.
NASA Technical Reports Server (NTRS)
Mcghee, Robert J.; Walker, Betty S.; Millard, Betty F.
1988-01-01
Experimental results were obtained for an Eppler 387 airfoil in the Langley Low Turbulence Pressure Tunnel. The tests were conducted over a Mach number range from 0.03 to 0.13 and a chord Reynolds number range for 60,000 to 460,000. Lift and pitching moment data were obtained from airfoil surface pressure measurements and drag data for wake surveys. Oil flow visualization was used to determine laminar separation and turbulent reattachment locations. Comparisons of these results with data on the Eppler 387 airfoil from two other facilities as well as the Eppler airfoil code are included.
NASA Technical Reports Server (NTRS)
Schuster, David M.; Scott, Robert C.; Bartels, Robert E.; Edwards, John W.; Bennett, Robert M.
2000-01-01
As computational fluid dynamics methods mature, code development is rapidly transitioning from prediction of steady flowfields to unsteady flows. This change in emphasis offers a number of new challenges to the research community, not the least of which is obtaining detailed, accurate unsteady experimental data with which to evaluate new methods. Researchers at NASA Langley Research Center (LaRC) have been actively measuring unsteady pressure distributions for nearly 40 years. Over the last 20 years, these measurements have focused on developing high-quality datasets for use in code evaluation. This paper provides a sample of unsteady pressure measurements obtained by LaRC and available for government, university, and industry researchers to evaluate new and existing unsteady aerodynamic analysis methods. A number of cases are highlighted and discussed with attention focused on the unique character of the individual datasets and their perceived usefulness for code evaluation. Ongoing LaRC research in this area is also presented.
Investigation of the unsteady pressure distribution on the blades of an axial flow fan
NASA Technical Reports Server (NTRS)
Henderson, R. E.; Franke, G. F.
1978-01-01
The unsteady response of a stator blade caused by the interaction of the stator with the wakes of an upstream rotor was investigated. Unsteady pressure distributions were measured using a blade instrumented with a series miniature pressure transducers. The influence of several geometrical and flow parameters - rotor/stator spacing, stator solidity and stator incidence angle - were studied to determine the unsteady response of the stator to these parameters. A major influence on the stator unsteady response is due to the stator solidity. At high solidities the blade-to-blade interference has a larger contribution. While the range of rotor/stator spacings investigated had a minor influence, the effect of stator incidence angle is significant. The data indicate the existence of an optimum positive incidence which minimizes the unsteady response.
NASA Technical Reports Server (NTRS)
Somers, Dan M. (Inventor)
2005-01-01
An airfoil having a fore airfoil element, an aft airfoil element, and a slot region in between them. These elements induce laminar flow over substantially all of the fore airfoil element and also provide for laminar flow in at least a portion of the slot region. The method of the invention is one for inducing natural laminar flow over an airfoil. In the method, a fore airfoil element, having a leading and trailing edge, and an aft airfoil element define a slot region. Natural laminar flow is induced over substantially all of the fore airfoil element, by inducing the pressures on both surfaces of the fore airfoil element to decrease to a location proximate the trailing edge of the fore airfoil element using pressures created by the aft airfoil element.
RSRM Chamber Pressure Oscillations: Transit Time Models and Unsteady CFD
NASA Technical Reports Server (NTRS)
Nesman, Tom; Stewart, Eric
1996-01-01
through an iterative structural and CFD analysis. The analysis domain ended just upstream of the nozzle throat. This is an acoustic boundary condition that caused the motor to behave as a closed-open organ pipe. This differs from the RSRM which behaves like a closed-closed organ pipe. The unsteady CFD solution shows RSRM chamber pressure oscillations predominately at the longitudinal acoustic mode frequencies of a closed-open organ pipe. Vortex shedding in the joint cavities and at the inhibitors contribute disturbances to the flow at the second longitudinal acoustic mode frequency. Further studies are planned using an analysis domain that extends downstream of the nozzle throat.
Theoretical Prediction of Pressure Distributions on Nonlifting Airfoils at High Subsonic Speeds
NASA Technical Reports Server (NTRS)
Spreiter, John R; Alksne, Alberta
1955-01-01
Theoretical pressure distributions on nonlifting circular-arc airfoils in two-dimensional flows with high subsonic free-stream velocity are found by determining approximate solutions, through an iteration process, of an integral equation for transonic flow proposed by Oswatitsch. The integral equation stems directly from the small-disturbance theory for transonic flow. This method of analysis possesses the advantage of remaining in the physical, rather than the hodograph, variable and can be applied in airfoils having curved surfaces. After discussion of the derivation of the integral equation and qualitative aspects of the solution, results of calculations carried out for circular-arc airfoils in flows with free-stream Mach numbers up to unity are described. These results indicate most of the principal phenomena observed in experimental studies.
NASA Technical Reports Server (NTRS)
Graham, Donald J; Nitzberg, Gerald E; Olson, Robert N
1945-01-01
Pressure distributions determined from high-speed wind-tunnel tests are presented for five NACA airfoil sections representative of both low-drag and conventional types. Section characteristics of lift, drag, and quarter-chord pitching moment are presented along with the measured pressure distributions for the NACA 65sub2-215 (a=0.5), 66sub2-215 (a=0.6), 0015, 23015, and 4415 airfoils for Mach numbers up to approximately 0.85. A critical study is made of the airfoil pressure distributions in an attempt to formulate a set of general criteria for defining the character of high speed flows over typical airfoil shapes. Comparisons are made of the relative characteristics of the low-drag and conventional airfoils investigated insofar as they would influence the high-speed performance and the high-speed stability and control characteristics of airplanes employing these wing sections.
Pressure Distribution on a Slotted R.A.F. 31 Airfoil in the Variable Density Wind Tunnel
NASA Technical Reports Server (NTRS)
Jacobs, Eastman N
1929-01-01
Measurements were made in the variable density wind tunnel to determine the pressure distribution over one section of a R.A.F. 31 airfoil with a leading edge slot fully open. To provide data for the study of scale effect on this type of airfoil, the tests were conducted with air densities of approximately one and twenty atmospheres.
Synthetic Vortex Generator Jets Used to Control Separation on Low-Pressure Turbine Airfoils
NASA Technical Reports Server (NTRS)
Ashpis, David E.; Volino, Ralph J.
2005-01-01
Low-pressure turbine (LPT) airfoils are subject to increasingly stronger pressure gradients as designers impose higher loading in an effort to improve efficiency and lower cost by reducing the number of airfoils in an engine. When the adverse pressure gradient on the suction side of these airfoils becomes strong enough, the boundary layer will separate. Separation bubbles, particularly those that fail to reattach, can result in a significant loss of lift and a subsequent degradation of engine efficiency. The problem is particularly relevant in aircraft engines. Airfoils optimized to produce maximum power under takeoff conditions may still experience boundary layer separation at cruise conditions because of the thinner air and lower Reynolds numbers at altitude. Component efficiency can drop significantly between takeoff and cruise conditions. The decrease is about 2 percent in large commercial transport engines, and it could be as large as 7 percent in smaller engines operating at higher altitudes. Therefore, it is very beneficial to eliminate, or at least reduce, the separation bubble.
Spatial Characteristics of the Unsteady Differential Pressures on 16 percent F/A-18 Vertical Tails
NASA Technical Reports Server (NTRS)
Moses, Robert W.; Ashley, Holt
1998-01-01
Buffeting is an aeroelastic phenomenon which plagues high performance aircraft at high angles of attack. For the F/A-18 at high angles of attack, vortices emanating from wing/fuselage leading edge extensions burst, immersing the vertical tails in their turbulent wake. The resulting buffeting of the vertical tails is a concern from fatigue and inspection points of view. Previous flight and wind-tunnel investigations to determine the buffet loads on the tail did not provide a complete description of the spatial characteristics of the unsteady differential pressures. Consequently, the unsteady differential pressures were considered to be fully correlated in the analyses of buffet and buffeting. The use of fully correlated pressures in estimating the generalized aerodynamic forces for the analysis of buffeting yielded responses that exceeded those measured in flight and in the wind tunnel. To learn more about the spatial characteristics of the unsteady differential pressures, an available 16%, sting-mounted, F-18 wind-tunnel model was modified and tested in the Transonic Dynamics Tunnel (TDT) at the NASA Langley Research Center as part of the ACROBAT (Actively Controlled Response Of Buffet-Affected Tails) program. Surface pressures were measured at high angles of attack on flexible and rigid tails. Cross-correlation and cross-spectral analyses of the pressure time histories indicate that the unsteady differential pressures are not fully correlated. In fact, the unsteady differential pressure resemble a wave that travels along the tail. At constant angle of attack, the pressure correlation varies with flight speed.
NASA Technical Reports Server (NTRS)
Ozturk, Burak; Schobeiri, Meinhard T.
2009-01-01
The present study, which is the first of a series of investigations of low pressure turbine (LPT) boundary layer aerodynamics, is aimed at providing detailed unsteady boundary layer flow information to understand the underlying physics of the inception, onset, and extent of the separation zone. A detailed experimental study on the behavior of the separation zone on the suction surface of a highly loaded LPT-blade under periodic unsteady wake flow is presented. Experimental investigations were performed on a large-scale, high-subsonic unsteady turbine cascade research facility with an integrated wake generator and test section unit. Blade Pak B geometry was used in the cascade. The wakes were generated by continuously moving cylindrical bars device. Boundary layer investigations were performed using hot wire anemometry at Reynolds number of 110,000, based on the blade suction surface length and the exit velocity, for one steady and two unsteady inlet flow conditions, with the corresponding passing frequencies, wake velocities, and turbulence intensities. The reduced frequencies cover the entire operation range of LP-turbines. In addition to the unsteady boundary layer measurements, blade surface pressure measurements were performed at Re = 50,000, 75,000, 100,000, 110,000, and 125,000. For each Reynolds number, surface pressure measurements are carried out at one steady and two periodic unsteady inlet flow conditions. Detailed unsteady boundary layer measurement identifies the onset and extension of the separation zone as well as its behavior under unsteady wake flow. The results, presented in ensemble-averaged and contour plot forms, help to understand the physics of the separation phenomenon under periodic unsteady wake flow.
NASA Astrophysics Data System (ADS)
Wolf, William R.; Azevedo, João L. F.; Lele, Sanjiva K.
2013-12-01
This paper presents a further analysis of results of airfoil self-noise prediction obtained in the previous work using large eddy simulation and acoustic analogy. The physical mechanisms responsible for airfoil noise generation in the aerodynamic flows analyzed are a combination of turbulent and laminar boundary layers, as well as vortex shedding (VS) originated due to trailing edge bluntness. The primary interest here consists of evaluating the effects of mean flow convection, quadrupole sources and vortex shedding tonal noise on the overall sound pressure level (OASPL) of a NACA0012 airfoil at low and moderate freestream Mach numbers. The overall sound pressure level is the measured quantity which eventually would be the main concern in terms of noise generation for aircraft and wind energy companies, and regulating agencies. The Reynolds number based on the airfoil chord is fixed at Rec=408,000 for all flow configurations studied. The results demonstrate that, for moderate Mach numbers, mean flow effects and quadrupole sources considerably increase OASPL and, therefore, should be taken into account in the acoustic prediction. For a low Mach number flow with vortex shedding, it is observed that OASPL is higher when laminar boundary layer separation is the VS driving mechanism compared to trailing edge bluntness.
NASA Technical Reports Server (NTRS)
Lawing, P. L.
1985-01-01
A method of constructing airfoils by inscribing pressure channels on the face of opposing plates, bonding them together to form one plate with integral channels, and contour machining this plate to form an airfoil model is described. The research and development program to develop the bonding technology is described as well as the construction and testing of an airfoil model. Sample aerodynamic data sets are presented and discussed. Also, work currently under way to produce thin airfoils with camber is presented. Samples of the aft section of a 6 percent airfoil with complete pressure instrumentation including the trailing edge are pictured and described. This technique is particularly useful in fabricating models for transonic cryogenic testing, but it should find application in a wide ange of model construction projects, as well as the fabrication of fuel injectors, space hardware, and other applications requiring advanced bonding technology and intricate fluid passages.
NASA Technical Reports Server (NTRS)
Reid, E G; Bamber, M J
1928-01-01
The tests described in this report constitute a preliminary investigation of airfoil boundary layer control, as carried out in the atmospheric wind tunnel of the Langley Memorial Aeronautical Laboratory, from February to August, 1927. Tests were made on a U.S.A. 27 airfoil section with various slot shapes and combinations, and at various amounts of pressure or suction on the slots. The lift of airfoils can be increased by removing or by accelerating the boundary layer. Removing the boundary layer by suction is more economical than to accelerate it by jet action. Gauze-covered suction slots apparently give the best results. When not in operation, all suction slots tested had a detrimental effect upon the aerodynamic characteristics of the airfoil which was not apparent with the backward-opening pressure slots. Thick, blunt-nose airfoils would seem to give best results with boundary layer control.
A fast approach to designing airfoils from given pressure distribution in compressible flows
NASA Technical Reports Server (NTRS)
Daripa, Prabir
1987-01-01
A new inverse method for aerodynamic design of airfols is presented for subcritical flows. The pressure distribution in this method can be prescribed as a function of the arc length of the as-yet unknown body. This inverse problem is shown to be mathematically equivalent to solving only one nonlinear boundary value problem subject to known Dirichlet data on the boundary. The solution to this problem determines the airfoil, the freestream Mach number, and the upstream flow direction. The existence of a solution to a given pressure distribution is discussed. The method is easy to implement and extremely efficient. A series of results for which comparisons are made with the known airfoils is presented.
Calculation and Correlation of the Unsteady Flowfield in a High Pressure Turbine
NASA Technical Reports Server (NTRS)
Bakhle, Milind A.; Liu, Jong S.; Panovsky, Josef; Keith, Theo G., Jr.; Mehmed, Oral
2002-01-01
Forced vibrations in turbomachinery components can cause blades to crack or fail due to high-cycle fatigue. Such forced response problems will become more pronounced in newer engines with higher pressure ratios and smaller axial gap between blade rows. An accurate numerical prediction of the unsteady aerodynamics phenomena that cause resonant forced vibrations is increasingly important to designers. Validation of the computational fluid dynamics (CFD) codes used to model the unsteady aerodynamic excitations is necessary before these codes can be used with confidence. Recently published benchmark data, including unsteady pressures and vibratory strains, for a high-pressure turbine stage makes such code validation possible. In the present work, a three dimensional, unsteady, multi blade-row, Reynolds-Averaged Navier Stokes code is applied to a turbine stage that was recently tested in a short duration test facility. Two configurations with three operating conditions corresponding to modes 2, 3, and 4 crossings on the Campbell diagram are analyzed. Unsteady pressures on the rotor surface are compared with data.
The variation with Reynolds number of pressure distribution over an airfoil section
NASA Technical Reports Server (NTRS)
Pinkerton, Robert M
1938-01-01
Pressures were simultaneously measured at 54 orifices distributed over the midspan section of a 5 by 30-inch rectangular model of the NACA 4412 airfoil in the variable-density tunnel. These measurements were made at 17 angles of attack from -20 degrees to 30 degrees for eight values of the effective Reynolds number form approximately 100,000 to 8,200,000. Accurate data were thus obtained for studying the variation of pressure distribution with Reynolds number. These results on the NACA 4412 section indicated that the pressure distribution is practically unaffected by changes in Reynolds number except where separation is involved.
An Overview of Unsteady Pressure Measurements in the Transonic Dynamics Tunnel
NASA Technical Reports Server (NTRS)
Schuster, David M.; Edwards, John W.; Bennett, Robert M.
2000-01-01
The NASA Langley Transonic Dynamics Tunnel has served as a unique national facility for aeroelastic testing for over forty years. A significant portion of this testing has been to measure unsteady pressures on models undergoing flutter, forced oscillations, or buffet. These tests have ranged from early launch vehicle buffet to flutter of a generic high-speed transport. This paper will highlight some of the test techniques, model design approaches, and the many unsteady pressure tests conducted in the TDT. The objectives and results of the data acquired during these tests will be summarized for each case and a brief discussion of ongoing research involving unsteady pressure measurements and new TDT capabilities will be presented.
NASA Astrophysics Data System (ADS)
Matsuda, Yu; Yorita, Daisuke; Egami, Yasuhiro; Kameya, Tomohiro; Kakihara, Noriaki; Yamaguchi, Hiroki; Asai, Keisuke; Niimi, Tomohide
2013-10-01
The pressure-sensitive paint technique based on the heterodyne method was proposed for the precise pressure measurement of unsteady flow fields. This measurement is realized by detecting the beat signal that results from interference between a modulating illumination light source and a pressure fluctuation. The beat signal is captured by a camera with a considerably lower frame rate than the frequency of the pressure fluctuation. By carefully adjusting the frequency of the light and the camera frame rate, the signal at the frequency of interest is detected, while the noise signals at other frequencies are eliminated. To demonstrate the proposed method, we measured the pressure fluctuations in a resonance tube at the fundamental, second, and third harmonics. The pressure fluctuation distributions were successfully obtained and were consistent with measurements from a pressure transducer. The proposed method is a useful technique for measuring unsteady phenomena.
Matsuda, Yu; Yorita, Daisuke; Egami, Yasuhiro; Kameya, Tomohiro; Kakihara, Noriaki; Yamaguchi, Hiroki; Asai, Keisuke; Niimi, Tomohide
2013-10-01
The pressure-sensitive paint technique based on the heterodyne method was proposed for the precise pressure measurement of unsteady flow fields. This measurement is realized by detecting the beat signal that results from interference between a modulating illumination light source and a pressure fluctuation. The beat signal is captured by a camera with a considerably lower frame rate than the frequency of the pressure fluctuation. By carefully adjusting the frequency of the light and the camera frame rate, the signal at the frequency of interest is detected, while the noise signals at other frequencies are eliminated. To demonstrate the proposed method, we measured the pressure fluctuations in a resonance tube at the fundamental, second, and third harmonics. The pressure fluctuation distributions were successfully obtained and were consistent with measurements from a pressure transducer. The proposed method is a useful technique for measuring unsteady phenomena.
Airfoil wake and linear theory gust response including sub and superresonant flow conditions
NASA Technical Reports Server (NTRS)
Henderson, Gregory H.; Fleeter, Sanford
1992-01-01
The unsteady aerodynamic gust response of a high solidity stator vane row is examined in terms of the fundamental gust modeling assumptions with particular attention given to the effects near an acoustic resonance. A series of experiments was performed with gusts generated by rotors comprised of perforated plates and airfoils. It is concluded that, for both the perforated plate and airfoil wake generated gusts, the unsteady pressure responses do not agree with the linear-theory gust predictions near an acoustic resonance. The effects of the acoustic resonance phenomena are clearly evident on the airfoil surface unsteady pressure responses. The transition of the measured lift coefficients across the acoustic resonance from the subresonant regime to the superresonant regime occurs in a simple linear fashion.
Unsteady blade pressure measurements for the SR-7A propeller at cruise conditions
NASA Technical Reports Server (NTRS)
Heidelberg, L. J.; Nallasamy, M.
1990-01-01
The unsteady blade surface pressures were measured on the SR-7A propeller. The freestream Mach number, inflow angle, and advance ratio were varied while measurements were made at nine blade stations. At a freestream Mach number of 0.8, the data in terms of unsteady pressure coefficient versus azimuth angle are compared to an unsteady three-dimensional Euler solution, yielding very encouraging results. The code predicts the shape (phase) of the waveform very well, while the magnitude is over-predicted in many cases. At tunnel Mach numbers below 0.6, an unusually large response on the suction surface at 0.15 chord and 0.88 radius was observed. The behavior of this response suggests the presence of a leading-edge vortex.
Unsteady blade pressure measurements for the SR-7A propeller at cruise conditions
NASA Technical Reports Server (NTRS)
Heidelberg, L. J.; Nallasamy, M.
1990-01-01
The unsteady blade surface pressures were measured on the SR-7A propeller. The freestream Mach no., inflow angle, and advance ratio were varied while measurements were made at nine blade stations. At a freestream Mach no. of 0.8, the data in terms of unsteady pressure coefficient vs. azimuth angle are compared to an unsteady 3-D Euler solution, yielding very encouraging results. The code predicts the shape (phase) of the waveform very well, while the magnitude is over-predicted in many cases. At tunnel Mach nos. below 0.6, an unusually large response on the suction surface at 0.15 chord and 0.88 radius was observed. The behavior of this response suggests the presence of a leading edge vortex. The midchord measuring stations on the suction surface exhibit a response that leads the forcing function while most other locations show a phase lag.
Aeroacoustic Study of a High-Fidelity Aircraft Model. Part 2; Unsteady Surface Pressures
NASA Technical Reports Server (NTRS)
Khorrami, Mehdi R.; Neuhart, Danny H.
2012-01-01
In this paper, we present unsteady surface pressure measurements for an 18%-scale, semi-span Gulfstream aircraft model. This high-fidelity model is being used to perform detailed studies of airframe noise associated with main landing gear, flap components, and gear-flap interaction noise, as well as to evaluate novel noise reduction concepts. The aerodynamic segment of the tests, conducted in the NASA Langley Research Center 14- by 22-Foot Subsonic Tunnel, was completed in November 2010. To discern the characteristics of the surface pressure fluctuations in the vicinity of the prominent noise sources, unsteady sensors were installed on the inboard and outboard flap edges, and on the main gear wheels, struts, and door. Various configurations were tested, including flap deflections of 0?, 20?, and 39?, with and without the main landing gear. The majority of unsteady surface pressure measurements were acquired for the nominal landing configuration where the main gear was deployed and the flap was deflected 39?. To assess the Mach number variation of the surface pressure amplitudes, measurements were obtained at Mach numbers of 0.16, 0.20, and 0.24. Comparison of the unsteady surface pressures with the main gear on and off shows significant interaction between the gear wake and the inboard flap edge, resulting in higher amplitude fluctuations when the gear is present.
Development of a nonlinear unsteady transonic flow theory
NASA Technical Reports Server (NTRS)
Stahara, S. S.; Spreiter, J. R.
1973-01-01
A nonlinear, unsteady, small-disturbance theory capable of predicting inviscid transonic flows about aerodynamic configurations undergoing both rigid body and elastic oscillations was developed. The theory is based on the concept of dividing the flow into steady and unsteady components and then solving, by method of local linearization, the coupled differential equation for unsteady surface pressure distribution. The equations, valid at all frequencies, were derived for two-dimensional flows, numerical results, were obtained for two classses of airfoils and two types of oscillatory motions.
Preliminary results of unsteady blade surface pressure measurements for the SR-3 propeller
NASA Technical Reports Server (NTRS)
Heidelberg, L. J.; Clark, B. J.
1986-01-01
Unsteady blade surface pressures were measured on an advanced, highly swept propeller known as SR-3. These measurements were obtained because the unsteady aerodynamics of these highly loaded transonic blades is important to noise generation and aeroelastic response. Specifically, the response to periodic angle-of-attack change was measured for both two- and eight-bladed configurations over a range of flight Mach numbers from 0.4 to 0.85. The periodic angle-of-attack change was obtained by placing the propeller axis at angles up to 4 deg to the flow. Most of the results are presented in terms of the unsteady pressure coefficient variation with Mach number. Both cascade and Mach number effects were largest on the suction surface near the leading edge. The results of a three-dimensional Euler code applied in a quasi-steady fashion were compared to measured data at the reduced frequency of 0.1 and showed relatively poor agreement. Pressure waveforms are shown that suggest shock phenomena may play an important part in the unsteady pressure response at some blade locations.
NASA Technical Reports Server (NTRS)
Dolling, David S.; Barter, John W.
1995-01-01
The focus was on developing means of controlling and reducing unsteady pressure loads in separated shock wave turbulent boundary layer interactions. Section 1 describes how vortex generators can be used to effectively reduce loads in compression ramp interaction, while Section 2 focuses on the effects of 'boundary-layer separators' on the same interaction.
On the Physics of Flow Separation Along a Low Pressure Turbine Blade Under Unsteady Flow Conditions
NASA Technical Reports Server (NTRS)
Schobeiri, Meinhard T.; Ozturk, Burak; Ashpis, David E.
2005-01-01
The present study, which is the first of a series of investigations dealing with specific issues of low pressure turbine (LPT) boundary layer aerodynamics, is aimed at providing detailed unsteady boundary flow information to understand the underlying physics of the inception, onset, and extent of the separation zone. A detailed experimental study on the behavior of the separation zone on the suction surface of a highly loaded LPT-blade under periodic unsteady wake flow is presented. Experimental investigations were performed at Texas A&M Turbomachinery Performance and Flow Research Laboratory using a large-scale unsteady turbine cascade research facility with an integrated wake generator and test section unit. To account for a high flow deflection of LPT-cascades at design and off-design operating points, the entire wake generator and test section unit including the traversing system is designed to allow a precise angle adjustment of the cascade relative to the incoming flow. This is done by a hydraulic platform, which simultaneously lifts and rotates the wake generator and test section unit. The unit is then attached to the tunnel exit nozzle with an angular accuracy of better than 0.05 , which is measured electronically. Utilizing a Reynolds number of 110,000 based on the blade suction surface length and the exit velocity, one steady and two different unsteady inlet flow conditions with the corresponding passing frequencies, wake velocities and turbulence intensities are investigated using hot-wire anemometry. In addition to the unsteady boundary layer measurements, blade surface pressure measurements were performed at Re=50,000, 75,000, 100,000, and 125,000 at one steady and two periodic unsteady inlet flow conditions. Detailed unsteady boundary layer measurement identifies the onset and extent of the separation zone as well as its behavior under unsteady wake flow. The results presented in ensemble-averaged and contour plot forms contribute to understanding the
On the Physics of Flow Separation Along a Low Pressure Turbine Blade Under Unsteady Flow Conditions
NASA Technical Reports Server (NTRS)
Schobeiri, Meinhard T.; Ozturk, Burak; Ashpis, David E.
2003-01-01
The present study, which is the first of a series of investigations dealing with specific issues of low pressure turbine (LPT) boundary layer aerodynamics, is aimed at providing detailed unsteady boundary flow information to understand the underlying physics of the inception, onset, and extent of the separation zone. A detailed experimental study on the behavior of the separation zone on the suction surface of a highly loaded LPT-blade under periodic unsteady wake flow is presented. Experimental investigations were performed at Texas A&M Turbomachinery Performance and Flow Research Laboratory using a large-scale unsteady turbine cascade research facility with an integrated wake generator and test section unit. To account for a high flow deflection of LPT-cascades at design and off-design operating points, the entire wake generator and test section unit including the traversing system is designed to allow a precise angle adjustment of the cascade relative to the incoming flow. This is done by a hydraulic platform, which simultaneously lifts and rotates the wake generator and test section unit. The unit is then attached to the tunnel exit nozzle with an angular accuracy of better than 0.05 , which is measured electronically. Utilizing a Reynolds number of 110,000 based on the blade suction surface length and the exit velocity, one steady and two different unsteady inlet flowconditions with the corresponding passing frequencies, wake velocities and turbulence intensities are investigated using hot-wire anemometry. In addition to the unsteady boundary layer measurements, blade surface pressure measurements were performed at Re=50,000, 75,000, 100,000, and 125,000 at one steady and two periodic unsteady inlet flow conditions. Detailed unsteady boundary layer measurement identifies the onset and extent of the separation zone as well as its behavior under unsteady wake flow. The results presented in ensemble-averaged and contour plot forms contribute to understanding the
The Influence of Unsteadiness on the Analysis of Pressure Gain Combustion Devices
NASA Technical Reports Server (NTRS)
Paxson, Daniel E.; Kaemming, Tom
2013-01-01
Pressure gain combustion (PGC) has been the object of scientific study for over a century due to its promise of improved thermodynamic efficiency. In many recent application concepts PGC is utilized as a component in an otherwise continuous, normally steady flow system, such as a gas turbine or ram jet engine. However, PGC is inherently unsteady. Failure to account for the effects of this periodic unsteadiness can lead to misunderstanding and errors in performance calculations. This paper seeks to provide some clarity by presenting a consistent method of thermodynamic cycle analysis for a device utilizing PGC technology. The incorporation of the unsteady PGC process into the conservation equations for a continuous flow device is presented. Most importantly, the appropriate method for computing the conservation of momentum is presented. It will be shown that proper, consistent analysis of cyclic conservation principles produces representative performance predictions.
The investigation of parachute fabric permeability under an unsteady pressure differential
NASA Astrophysics Data System (ADS)
Rondeau, Nichole C.
An apparatus for assessing permeability of textiles subjected to time-varying pressure differentials is presented. A Computer Numerically Controlled Piston Permeability Apparatus (CNC-PPA) that can control the volume flow rate through a fabric has been designed and built. This test device has been developed in an effort to improve the understanding and design choices for aerodynamic decelerators. Preliminary results for a low permeability fabric (PIA-C-44378, Type IV) under both steady and unsteady loads are presented. The results from this investigation do indicate a small effect of unsteady pressure differential on the fabric permeability. The fabric permeability is slightly higher than the static permeability when the pressure differential is increasing with respect to time and the opposite is true when the pressure differential is decreasing. This change in permeability is more pronounced as the pressure is higher and the pressure changes more rapidly with respect to time, suggesting dynamic permeability likely affects highly unsteady phenomena such as parachute opening.
NASA Technical Reports Server (NTRS)
Volino, Ralph J.; Hultgren, Lennart .
2000-01-01
Detailed velocity measurements were made along a flat plate subject to the same dimensionless pressure gradient as the suction side of a modern low-pressure turbine airfoil. Reynolds numbers based on wetted plate length and nominal exit velocity were varied from 50,000 to 300,000, covering cruise to takeoff conditions. Low and high inlet free-stream turbulence intensities (0.2% and 7%) were set using passive grids. The location of boundary-layer separation does not depend strongly on the free-stream turbulence level or Reynolds number, as long as the boundary layer remains non-turbulent prior to separation. Strong acceleration prevents transition on the upstream part of the plate in all cases. Both free-stream turbulence and Reynolds number have strong effects on transition in the adverse pressure gradient region. Under low free-stream turbulence conditions transition is induced by instability waves in the shear layer of the separation bubble. Reattachment generally occurs at the transition start. At Re = 50,000 the separation bubble does not close before the trailing edge of the modeled airfoil. At higher Re, transition moves upstream, and the boundary layer reattaches. With high free-stream turbulence levels, transition appears to occur in a bypass mode, similar to that in attached boundary layers. Transition moves upstream, resulting in shorter separation regions. At Re above 200,000, transition begins before separation. Mean velocity, turbulence and intermittency profiles are presented.
NASA Astrophysics Data System (ADS)
Gardner, A. D.; Klein, C.; Sachs, W. E.; Henne, U.; Mai, H.; Richter, K.
2014-09-01
Dynamic stall on a pitching OA209 airfoil in a wind tunnel is investigated at Mach 0.3 and 0.5 using high-speed pressure-sensitive paint (PSP) and pressure measurements. At Mach 0.3, the dynamic stall vortex was observed to propagate faster at the airfoil midline than at the wind-tunnel wall, resulting in a "bowed" vortex shape. At Mach 0.5, shock-induced stall was observed, with initial separation under the shock foot and subsequent expansion of the separated region upstream, downstream and along the breadth of the airfoil. No dynamic stall vortex could be observed at Mach 0.5. The investigation of flow control by blowing showed the potential advantages of PSP over pressure transducers for a complex three-dimensional flow.
NASA Technical Reports Server (NTRS)
Torres, Francisco J.
1987-01-01
Six airfoil interferograms were evaluated using a semiautomatic image-processor system which digitizes, segments, and extracts the fringe coordinates along a polygonal line. The resulting fringe order function was converted into density and pressure distributions and a comparison was made with pressure transducer data at the same wind tunnel test conditions. Three airfoil shapes were used in the evaluation to test the capabilities of the image processor with a variety of flows. Symmetric, supercritical, and circulation-control airfoil interferograms provided fringe patterns with shocks, separated flows, and high-pressure regions for evaluation. Regions along the polygon line with very clear fringe patterns yielded results within 1% of transducer measurements, while poorer quality regions, particularly near the leading and trailing edges, yielded results that were not as good.
Two-dimensional unsteady lift problems in supersonic flight
NASA Technical Reports Server (NTRS)
Heaslet, Max A; Lomax, Harvard
1949-01-01
The variation of pressure distribution is calculated for a two-dimensional supersonic airfoil either experiencing a sudden angle-of-attack change or entering a sharp-edge gust. From these pressure distributions the indicial lift functions applicable to unsteady lift problems are determined for two cases. Results are presented which permit the determination of maximum increment in lift coefficient attained by an unrestrained airfoil during its flight through a gust. As an application of these results, the minimum altitude for safe flight through a specific gust is calculated for a particular supersonic wing of given strength and wing loading.
Experience with transonic unsteady aerodynamic calculations
NASA Technical Reports Server (NTRS)
Edwards, J. W.; Bland, S. R.; Seidel, D. A.
1985-01-01
Comparisons of calculated and experimental transonic unsteady pressures and airloads for four of the AGARD Two Dimensional Aeroelastic Configurations and for a rectangular supercritical wing are presented. The two dimensional computer code, XTRAN2L, implementing the transonic small perturbation equation was used to obtain results for: (1) pitching oscillations of the NACA 64A010A; NLR 7301 and NACA 0012 airfoils; (2) flap oscillations for the NACA 64A006 and NRL 7301 airfoils; and (3) transient ramping motions for the NACA 0012 airfoils. Results from the three dimensional code XTRAN3S are compared with data from a rectangular supercritical wing oscillating in pitch. These cases illustrate the conditions under which the transonic inviscid small perturbation equation provides reasonable predictions.
Unsteady Cascade Aerodynamic Response Using a Multiphysics Simulation Code
NASA Technical Reports Server (NTRS)
Lawrence, C.; Reddy, T. S. R.; Spyropoulos, E.
2000-01-01
The multiphysics code Spectrum(TM) is applied to calculate the unsteady aerodynamic pressures of oscillating cascade of airfoils representing a blade row of a turbomachinery component. Multiphysics simulation is based on a single computational framework for the modeling of multiple interacting physical phenomena, in the present case being between fluids and structures. Interaction constraints are enforced in a fully coupled manner using the augmented-Lagrangian method. The arbitrary Lagrangian-Eulerian method is utilized to account for deformable fluid domains resulting from blade motions. Unsteady pressures are calculated for a cascade designated as the tenth standard, and undergoing plunging and pitching oscillations. The predicted unsteady pressures are compared with those obtained from an unsteady Euler co-de refer-red in the literature. The Spectrum(TM) code predictions showed good correlation for the cases considered.
Spectral analysis of unsteady surface pressure on a pusher propeller
NASA Technical Reports Server (NTRS)
Farokhi, Saeed
1992-01-01
A propeller of an advanced turboprop testbed aircraft in pusher configuration is instrumented with 22 miniature blade-mounted transducers (BMTs) at two radii. Upstream pylon wake interaction with the propeller is the source of a one-per-cycle excitation for the blades in flight. The time history of fluctuating pressure signals over 26 flight conditions is statistically analyzed in the frequency domain. The rms amplitude of fluctuating pressure signals measured by suction surface BMTs indicates a very strong presence of the fundamental frequency over most of the upper surface. The pylon wake pressure signature on the propeller trailing edge, i.e., x/c not less than 0.80, shows predominantly random turbulence; hence, the amplitude of the fundamental frequency wave is fairly small. The resurgence of a large amplitude fundamental harmonic with coherent pylon wake signature further downstream, say at 90 percent chord, is unexpected behavior. The appearance of a dominating second propeller shaft order in the spectra of the rms pressure in transonic flight conditions identifies the presence of a two-per-cycle excitation source in the azimuthal direction. This is due to the presence of a shock wave, as evidenced by the pressure-time history plots.
Application of the pressure sensitive paint technique to steady and unsteady flow
NASA Technical Reports Server (NTRS)
Shimbo, Y.; Mehta, R.; Cantwell, B.
1996-01-01
Pressure sensitive paint is a newly-developed optical measurement technique with which one can get a continuous pressure distribution in much shorter time and lower cost than a conventional pressure tap measurement. However, most of the current pressure sensitive paint applications are restricted to steady pressure measurement at high speeds because of the small signal-to-noise ratio at low speed and a slow response to pressure changes. In the present study, three phases of work have been completed to extend the application of the pressure sensitive paint technique to low-speed testing and to investigate the applicability of the paint technique to unsteady flow. First the measurement system using a commercially available PtOEP/GP-197 pressure sensitive paint was established and applied to impinging jet measurements. An in-situ calibration using only five pressure tap data points was applied and the results showed good repeatability and good agreement with conventional pressure tap measurements on the whole painted area. The overall measurement accuracy in these experiments was found to be within 0.1 psi. The pressure sensitive paint technique was then applied to low-speed wind tunnel tests using a 60 deg delta wing model with leading edge blowing slots. The technical problems encountered in low-speed testing were resolved by using a high grade CCD camera and applying corrections to improve the measurement accuracy. Even at 35 m/s, the paint data not only agreed well with conventional pressure tap measurements but also clearly showed the suction region generated by the leading edge vortices. The vortex breakdown was also detected at alpha=30 deg. It was found that a pressure difference of 0.2 psi was required for a quantitative pressure measurement in this experiment and that temperature control or a parallel temperature measurement is necessary if thermal uniformity does not hold on the model. Finally, the pressure sensitive paint was applied to a periodically
Numerical Simulations of Steady and Unsteady Transition in Low-Pressure Turbine Blade Rows
NASA Technical Reports Server (NTRS)
Dorney, Daniel
1998-01-01
Transition plays an important role in the prediction of losses and performance in low-pressure turbines. The transition location on a turbine blade may vary significantly because of the wakes from upstream blade rows, and intermittent flow separation can also affect the transition process in an unsteady flow environment. In the present investigation, an unsteady Navier-Stokes analysis is used to predict transition in a low-pressure turbine cascade and a low-pressure turbine stage. The numerical flow analysis is third-order spatially accurate and second-order temporally accurate, and the equations of motion are integrated using an implicit time-marching procedure. The Baldwin-Lomax and k-epsilon turbulence models, in conjunction with several algebraic transition models, have been used to predict the location of transition. Predicted results include unsteady blade loadings, time-histories of the pressure, transition locations and boundary layer quantities, as well as performance quantities and comparisons with the available experimental/design data.
Electro-Luminescence based Pressure-Sensitive Paint System for Unsteady Flow Field Measurements
NASA Astrophysics Data System (ADS)
Iijima, Yoshimi; Sakaue, Hirotaka
2011-11-01
Electro-luminescence (EL) based pressure-sensitive paint (PSP) system is developed for capturing unsteady flow fields. It has advantages in uniform distribution in the illumination without remotely apply the illumination source from the testing object. The resultant system can be applied directly onto a testing object surface. It consists of an inorganic EL and a PSP. The EL emits blue illumination uniformly applied onto the PSP layer. Because of a sheet illumination, the EL gives uniform distribution, while a point illumination gives a spot in illumination. The PSP is developed to provide a fast response to a change in pressure. It uses a porous particle and a polymer to create porous-polymer PSP. The response time characterization of the developed system is included in the presentation. The developed system is applied to an unsteady flow field, such as a sound field in a resonance tube.
Evaluation of the constant pressure panel method (CPM) for unsteady air loads prediction
NASA Technical Reports Server (NTRS)
Appa, Kari; Smith, Michael J. C.
1988-01-01
This paper evaluates the capability of the constant pressure panel method (CPM) code to predict unsteady aerodynamic pressures, lift and moment distributions, and generalized forces for general wing-body configurations in supersonic flow. Stability derivatives are computed and correlated for the X-29 and an Oblique Wing Research Aircraft, and a flutter analysis is carried out for a wing wind tunnel test example. Most results are shown to correlate well with test or published data. Although the emphasis of this paper is on evaluation, an improvement in the CPM code's handling of intersecting lifting surfaces is briefly discussed. An attractive feature of the CPM code is that it shares the basic data requirements and computational arrangements of the doublet lattice method. A unified code to predict unsteady subsonic or supersonic airloads is therefore possible.
Methodology of Blade Unsteady Pressure Measurement in the NASA Transonic Flutter Cascade
NASA Technical Reports Server (NTRS)
Lepicovsky, J.; McFarland, E. R.; Capece, V. R.; Jett, T. A.; Senyitko, R. G.
2002-01-01
In this report the methodology adopted to measure unsteady pressures on blade surfaces in the NASA Transonic Flutter Cascade under conditions of simulated blade flutter is described. The previous work done in this cascade reported that the oscillating cascade produced waves, which for some interblade phase angles reflected off the wind tunnel walls back into the cascade, interfered with the cascade unsteady aerodynamics, and contaminated the acquired data. To alleviate the problems with data contamination due to the back wall interference, a method of influence coefficients was selected for the future unsteady work in this cascade. In this approach only one blade in the cascade is oscillated at a time. The majority of the report is concerned with the experimental technique used and the experimental data generated in the facility. The report presents a list of all test conditions for the small amplitude of blade oscillations, and shows examples of some of the results achieved. The report does not discuss data analysis procedures like ensemble averaging, frequency analysis, and unsteady blade loading diagrams reconstructed using the influence coefficient method. Finally, the report presents the lessons learned from this phase of the experimental effort, and suggests the improvements and directions of the experimental work for tests to be carried out for large oscillation amplitudes.
Nearfield Unsteady Pressures at Cruise Mach Numbers for a Model Scale Counter-Rotation Open Rotor
NASA Technical Reports Server (NTRS)
Stephens, David B.
2012-01-01
An open rotor experiment was conducted at cruise Mach numbers and the unsteady pressure in the nearfield was measured. The system included extensive performance measurements, which can help provide insight into the noise generating mechanisms in the absence of flow measurements. A set of data acquired at a constant blade pitch angle but various rotor speeds was examined. The tone levels generated by the front and rear rotor were found to be nearly equal when the thrust was evenly balanced between rotors.
Advanced turboprop wing installation effects measured by unsteady blade pressure and noise
NASA Technical Reports Server (NTRS)
Heidelberg, Laurence J.; Woodward, Richard P.
1987-01-01
A single rotation model propeller (SR-7A) was tested at simulated takeoff/approach conditions (Mach 0.2), in the NASA Lewis 9- by 15-Ft Anechoic Wind Tunnel. Both unsteady blade surface pressures and noise measurements were made for a tractor configuration with propeller/straight wing and propeller alone configurations. The angle between the wing chord and propeller axis (droop angle) was varied along with the wing angle of attack to determine the effects on noise and unsteady loading. A method was developed that uses unsteady blade pressure measurements to provide a quantitative indication of propeller inflow conditions, at least for a uniform (across the propeller disk) inflow angle. The wing installation caused a nearly uniform upwash at the propeller inlet as evidenced by the domination of the pressure spectra by the first shaft order. This inflow angle increased at a rate of almost 150 percent of that of the wing angle-of-attack for a propeller-wing spacing of 0.54 wing chords at a constant droop angle. The flyover noise, as measured by the maximum blade passing frequency level, correlates closely with the propeller inflow angle (approx. 0.6 dB per degree of inflow angle) for all droop angles and wing angles of attack tested, including the propeller alone data. Large changes in the unsteady pressure responses on the suction surface of the blade were observed as the advance ratio was varied. The presence of a leading edge vortex may explain this behavior since changes in the location of this vortex would change with loading (advance ratio).
Single-shot temperature- and pressure-sensitive paint measurements on an unsteady helicopter blade
NASA Astrophysics Data System (ADS)
Disotell, Kevin J.; Peng, Di; Juliano, Thomas J.; Gregory, James W.; Crafton, Jim W.; Komerath, Narayanan M.
2014-02-01
Unsteady pressure-sensitive paint (PSP) measurements were acquired on an articulated model helicopter rotor of 0.26 m diameter in edgewise flow to simulate forward flight conditions. The rotor was operated at advance ratios (free stream velocity normalized by hover tip speed) of 0.15 and 0.30 at a cycle-averaged tip chord Reynolds number of 1.1 × 105, with collective and longitudinal cyclic pitch inputs of 10° and 2.5°, respectively. A single-shot data acquisition technique allowed a camera to record the paint luminescence after a single pulse of high-energy laser excitation, yielding sufficient signal-to-noise ratio to avoid image averaging. Platinum tetra(pentafluorophenyl) porphyrin (PtTFPP) in a porous polymer/ceramic binder served as the PSP. To address errors caused by image blurring and temperature sensitivity, a previously reported motion deblurring algorithm was implemented and the temperature correction was made using temperature-sensitive paint measurements on a second rotor blade. Instantaneous, unsteady surface pressure maps at a rotation rate of 82 Hz captured different aerodynamic responses between the two sides of the rotor disk and were compared to the nominally steady hover case. Cycle-to-cycle variations in tip unsteadiness on the retreating blade were also observed, causing oblique pressure features which may be linked to three-dimensional stall.
NASA Technical Reports Server (NTRS)
Riffel, R. E.; Rothrock, M. D.
1980-01-01
A two dimensional cascade of harmonically oscillating airfoils was designed to model a near tip section from a rotor which was known to have experienced supersonic translational model flutter. This five bladed cascade had a solidity of 1.52 and a setting angle of 0.90 rad. Unique graphite epoxy airfoils were fabricated to achieve the realistic high reduced frequency level of 0.15. The cascade was tested over a range of static pressure ratios approximating the blade element operating conditions of the rotor along a constant speed line which penetrated the flutter boundary. The time steady and time unsteady flow field surrounding the center cascade airfoil were investigated.
NASA Technical Reports Server (NTRS)
Kaattari, G. E.
1973-01-01
A method is presented for determining shock envelopes and pressure distributions for two-dimensional airfoils at angles of attack sufficiently large to cause shock detachment and subsonic flow over the windward surface of the airfoil. Correlation functions obtained from exact solutions are used to relate the shock standoff distance at the stagnation and sonic points of the body through a suitable choice for the shock shape. The necessary correlation functions were obtained from perfect gas solutions but may be extended to any gas flow for which the normal shock-density ratio can be specified.
Real-Time Unsteady Loads Measurements Using Hot-Film Sensors
NASA Technical Reports Server (NTRS)
Mangalam, Arun S.; Moes, Timothy R.
2004-01-01
Several flight-critical aerodynamic problems such as buffet, flutter, stall, and wing rock are strongly affected or caused by abrupt changes in unsteady aerodynamic loads and moments. Advanced sensing and flow diagnostic techniques have made possible simultaneous identification and tracking, in realtime, of the critical surface, viscosity-related aerodynamic phenomena under both steady and unsteady flight conditions. The wind tunnel study reported here correlates surface hot-film measurements of leading edge stagnation point and separation point, with unsteady aerodynamic loads on a NACA 0015 airfoil. Lift predicted from the correlation model matches lift obtained from pressure sensors for an airfoil undergoing harmonic pitchup and pitchdown motions. An analytical model was developed that demonstrates expected stall trends for pitchup and pitchdown motions. This report demonstrates an ability to obtain unsteady aerodynamic loads in real time, which could lead to advances in air vehicle safety, performance, ride-quality, control, and health management.
Real-Time Unsteady Loads Measurements Using Hot-Film Sensors
NASA Technical Reports Server (NTRS)
Mangalam, Arun S.; Moes, Timothy R.
2004-01-01
Several flight-critical aerodynamic problems such as buffet, flutter, stall, and wing rock are strongly affected or caused by abrupt changes in unsteady aerodynamic loads and moments. Advanced sensing and flow diagnostic techniques have made possible simultaneous identification and tracking, in real-time, of the critical surface, viscosity-related aerodynamic phenomena under both steady and unsteady flight conditions. The wind tunnel study reported here correlates surface hot-film measurements of leading edge stagnation point and separation point, with unsteady aerodynamic loads on a NACA 0015 airfoil. Lift predicted from the correlation model matches lift obtained from pressure sensors for an airfoil undergoing harmonic pitchup and pitchdown motions. An analytical model was developed that demonstrates expected stall trends for pitchup and pitchdown motions. This report demonstrates an ability to obtain unsteady aerodynamic loads in real-time, which could lead to advances in air vehicle safety, performance, ride-quality, control, and health management.
NASA Technical Reports Server (NTRS)
Jones, Gregory; Balakrishna, Sundareswara; DeMoss, Joshua; Goodliff, Scott; Bailey, Matthew
2015-01-01
Pressure fluctuations have been measured over the course of several tests in the National Transonic Facility to study unsteady phenomenon both with and without the influence of a model. Broadband spectral analysis will be used to characterize the length scales of the tunnel. Special attention will be given to the large-scale, low frequency data that influences the Mach number and force and moment variability. This paper will also discuss the significance of the vorticity and sound fields that can be related to the Common Research Model and will also highlight the comparisons to an empty tunnel configuration. The effectiveness of vortex generators placed at the interface of the test section and wind tunnel diffuser showed promise in reducing the empty tunnel unsteadiness, however, the vortex generators were ineffective in the presence of a model.
Numerical solution of periodic vortical flows about a thin airfoil
NASA Technical Reports Server (NTRS)
Scott, James R.; Atassi, Hafiz M.
1989-01-01
A numerical method is developed for computing periodic, three-dimensional, vortical flows around isolated airfoils. The unsteady velocity is split into a vortical component which is a known function of the upstream flow conditions and the Lagrangian coordinates of the mean flow, and an irrotational field whose potential satisfies a nonconstant-coefficient, inhomogeneous, convective wave equation. Solutions for thin airfoils at zero degrees incidence to the mean flow are presented in this paper. Using an elliptic coordinate transformation, the computational domain is transformed into a rectangle. The Sommerfeld radiation condition is applied to the unsteady pressure on the grid line corresponding to the far field boundary. The results are compared with a Possio solver, and it is shown that for maximum accuracy the grid should depend on both the Mach number and reduced frequency. Finally, in order to assess the range of validity of the classical thin airfoil approximation, results for airfoils with zero thickness are compared with results for airfoils with small thickness.
NASA Astrophysics Data System (ADS)
Fang, Shuo; Disotell, Kevin J.; Long, Samuel R.; Gregory, James W.; Semmelmayer, Frank C.; Guyton, Robert W.
2011-06-01
The current work focuses on the development and application of fast-responding polymer/ceramic pressure-sensitive paint (PSP) as an advanced surface pressure measurement technique for unsteady flow fields in large-scale wind tunnels. To demonstrate the unsteady PSP technique, the unsteady surface pressure distribution over a hemispherical dome placed in the United States Air Force Research Laboratory's Trisonic Gasdynamics Facility (TGF) was studied by phase-locking to the characteristic frequency in the flow caused by an unsteady separated shear layer shed from the dome. The wind tunnel was operated at stagnation pressures of 23.92 and 71.84 kPa, with the test section flow at Mach 0.6. Under the two operating conditions, the predominant shear layer frequency was measured to be 272 and 400 Hz, respectively. The quasi-periodic shear layer frequency enabled a phase-averaged method to be employed for capturing the unsteady shock motion on the hemisphere. Unsteady pressure data resulting from this technique are shown to correlate well with measurements acquired by conventional measurement techniques. Measurement uncertainty in the phase-averaging technique will be discussed. To address measurement uncertainties from temperature sensitivity and model movement, a new implementation of an AC-coupled data representation is offered.
Coupling of Low Speed Fan Stator Vane Unsteady Pressures to Duct Modes: Measured versus Predicted
NASA Technical Reports Server (NTRS)
Sutliff, Daniel L.; Heidelberg, Laurence J.; Envia, Edmane
1999-01-01
Uniform-flow annular-duct Green's functions are the essential elements of the classical acoustic analogy approach to the problem of computing the noise generated by rotor-stator interaction inside the fan duct. This paper investigates the accuracy of this class of Green's functions for predicting the duct noise levels when measured stator vane unsteady surface pressures are used as input to the theoretical formulation. The accuracy of the method is evaluated by comparing the predicted and measured acoustic power levels for the NASA 48 inch low speed Active Noise Control Fan. The unsteady surface pressures are measured,by an array of microphones imbedded in the suction and pressure sides of a single vane, while the duct mode levels are measured using a rotating rake system installed in the inlet and exhaust sections of the fan duct. The predicted levels are computed using properly weighted integrals of measured surface pressure distribution. The data-theory comparisons are generally quite good particularly when the mode cut-off criterion is carefully interpreted. This suggests that, at least for low speed fans, the uniform-flow annular-duct Green's function theory can be reliably used for prediction of duct mode levels if the cascade surface pressure distribution is accurately known.
Characterization of dynamic stall on 9-15 % thick airfoils using experiment and computation
NASA Astrophysics Data System (ADS)
Davidson, Phillip B.
In recent years, the blade geometry on wind turbines and helicopters has been optimized for a particular span location. Unsteady flow phenomena like dynamic stall limit these designs and need to be better understood and correctly simulated. Currently, empirical and computational fluid dynamics (CFD) methods are used to simulate rotating wind turbine or helicopter blades, but each of these methods has limitations in predicting unsteady separated flows. To address these needs, the present work investigated oscillating airfoils over a range of conditions with an approach that provided fast, low-cost unsteady pressure data combined with a highly resolved flow field to better understand the physics of dynamic stall. An additional objective was to show how such data may be used to assess CFD simulations. This research has yielded interesting results showing characteristics of thin airfoil stall, leading edge stall, and trailing edge stall that were sorted and classified. Classification of the oscillating airfoil behavior with or without dynamic stall was performed using previous definitions for stall regime, separation characteristics, and other qualitative differences in stall pattern. After classifying the unsteady flow for each of the cases, comparison of experimental results and results obtained using an unsteady Reynolds Averaged Navier-Stokes (URANS) solver was performed to assess the ability of the solver to produce the same unsteady effects. Although both experiment and computation produced similar flow features, the timing and magnitude of the features in the dynamic stall and re-attachment process of the pitching cycle exhibited some significant differences.
Calculation of steady and unsteady pressures at supersonic speeds with CAP-TSD
NASA Technical Reports Server (NTRS)
Bennett, Robert M.; Bland, Samuel R.; Batina, John T.; Gibbons, Michael D.; Mabey, Dennis G.
1989-01-01
A finite difference technique is used to solve the transonic small disturbance flow equation making use of shock capturing to treat wave discontinuities. Thus the nonlinear effects of thickness and angle of attack are considered. Such an approach is made feasible by the development of a new code called CAP-TSD (Computational Aeroelasticity Program - Transonic Small Disturbance), and is based on a fully implicit approximate factorization (AF) finite difference method to solve the time dependent transonic small disturbance equation. The application of the CAP-TSD code to the calculation of low to moderate supersonic steady and unsteady flows is presented. In particular, comparisons with exact linear theory solutions are made for steady and unsteady cases to evaluate shock capturing and other features of the current method. In addition, steady solutions obtained from an Euler code are used to evaluate the small disturbance aspects of the code. Steady and unsteady pressure comparisons are made with measurements for an F-15 wing model and for the RAE tailplane model.
Investigation of impinging jet resonant modes using unsteady pressure-sensitive paint measurements
NASA Astrophysics Data System (ADS)
Davis, Timothy; Edstrand, Adam; Alvi, Farrukh; Cattafesta, Louis; Yorita, Daisuke; Asai, Keisuke
2015-05-01
At given nozzle to plate spacings, the flow field of high-speed impinging jets is known to be characterized by a resonance phenomenon. Large coherent structures that convect downstream and impinge on the surface create strong acoustic waves that interact with the inherently unstable shear layer at the nozzle exit. This feedback mechanism, driven by the coherent structures in the jet shear layer, can either be axisymmetric or helical in nature. Fast-response pressure-sensitive paint (PSP) is applied to the impingement surface to map the unsteady pressure distribution associated with these resonant modes. Phase-averaged results acquired at several kHz are obtained using a flush mounted unsteady pressure transducer on the impingement plate as a reference signal. Tests are conducted on a Mach 1.5 jet at nozzle to plate spacings of . The resulting phase-averaged distribution reveals dramatically different flow fields at the corresponding impingement heights. The existence of a purely axisymmetric mode with a frequency of 6.3 kHz is identified at and is characterized by concentric rings of higher/lower pressure that propagate radially with increasing phase. Two simultaneous modes are observed at with one being a dominant symmetric mode at 7.1 kHz and the second a sub-dominant helical mode at 4.3 kHz. Complimentary phase-conditioned Schlieren images are also obtained visualizing the flow structures associated with each mode and are consistent with the PSP results.
NASA Technical Reports Server (NTRS)
Lakshminarayana, B.; Ho, Y.; Basson, A.
1993-01-01
The objective of this research is to simulate steady and unsteady viscous flows, including rotor/stator interaction and tip clearance effects in turbomachinery. The numerical formulation for steady flow developed here includes an efficient grid generation scheme, particularly suited to computational grids for the analysis of turbulent turbomachinery flows and tip clearance flows, and a semi-implicit, pressure-based computational fluid dynamics scheme that directly includes artificial dissipation, and is applicable to both viscous and inviscid flows. The values of these artificial dissipation is optimized to achieve accuracy and convergency in the solution. The numerical model is used to investigate the structure of tip clearance flows in a turbine nozzle. The structure of leakage flow is captured accurately, including blade-to-blade variation of all three velocity components, pitch and yaw angles, losses and blade static pressures in the tip clearance region. The simulation also includes evaluation of such quantities of leakage mass flow, vortex strength, losses, dominant leakage flow regions and the spanwise extent affected by the leakage flow. It is demonstrated, through optimization of grid size and artificial dissipation, that the tip clearance flow field can be captured accurately. The above numerical formulation was modified to incorporate time accurate solutions. An inner loop iteration scheme is used at each time step to account for the non-linear effects. The computation of unsteady flow through a flat plate cascade subjected to a transverse gust reveals that the choice of grid spacing and the amount of artificial dissipation is critical for accurate prediction of unsteady phenomena. The rotor-stator interaction problem is simulated by starting the computation upstream of the stator, and the upstream rotor wake is specified from the experimental data. The results show that the stator potential effects have appreciable influence on the upstream rotor wake
NASA Technical Reports Server (NTRS)
Shyam, Vikram; Ameri, Ali
2009-01-01
Unsteady 3-D RANS simulations have been performed on a highly loaded transonic turbine stage and results are compared to steady calculations as well as to experiment. A low Reynolds number k-epsilon turbulence model is employed to provide closure for the RANS system. A phase-lag boundary condition is used in the tangential direction. This allows the unsteady simulation to be performed by using only one blade from each of the two rows. The objective of this work is to study the effect of unsteadiness on rotor heat transfer and to glean any insight into unsteady flow physics. The role of the stator wake passing on the pressure distribution at the leading edge is also studied. The simulated heat transfer and pressure results agreed favorably with experiment. The time-averaged heat transfer predicted by the unsteady simulation is higher than the heat transfer predicted by the steady simulation everywhere except at the leading edge. The shock structure formed due to stator-rotor interaction was analyzed. Heat transfer and pressure at the hub and casing were also studied. Thermal segregation was observed that leads to the heat transfer patterns predicted by steady and unsteady simulations to be different.
NASA Technical Reports Server (NTRS)
Abbott, Ira H; Von Doenhoff, Albert E; Stivers, Louis, Jr
1945-01-01
The historical development of NACA airfoils is briefly reviewed. New data are presented that permit the rapid calculation of the approximate pressure distributions for the older NACA four-digit and five-digit airfoils by the same methods used for the NACA 6-series airfoils. The general methods used to derive the basic thickness forms for NACA 6 and 7-series airfoils together with their corresponding pressure distributions are presented. Detail data necessary for the application of the airfoils to wing design are presented in supplementary figures placed at the end of the paper. The report includes an analysis of the lift, drag, pitching-moment, and critical-speed characteristics of the airfoils, together with a discussion of the effects of surface conditions. Available data on high-lift devices are presented. Problems associated with lateral-control devices, leading-edge air intakes, and interference are briefly discussed, together with aerodynamic problems of application. (author)
NASA Technical Reports Server (NTRS)
Alter, Stephen J.; Brauckmann, Gregory J.; Kleb, William L.; Glass, Christopher E.; Streett, Craig L.; Schuster, David M.
2015-01-01
A transonic flow field about a Space Launch System (SLS) configuration was simulated with the Fully Unstructured Three-Dimensional (FUN3D) computational fluid dynamics (CFD) code at wind tunnel conditions. Unsteady, time-accurate computations were performed using second-order Delayed Detached Eddy Simulation (DDES) for up to 1.5 physical seconds. The surface pressure time history was collected at 619 locations, 169 of which matched locations on a 2.5 percent wind tunnel model that was tested in the 11 ft. x 11 ft. test section of the NASA Ames Research Center's Unitary Plan Wind Tunnel. Comparisons between computation and experiment showed that the peak surface pressure RMS level occurs behind the forward attach hardware, and good agreement for frequency and power was obtained in this region. Computational domain, grid resolution, and time step sensitivity studies were performed. These included an investigation of pseudo-time sub-iteration convergence. Using these sensitivity studies and experimental data comparisons, a set of best practices to date have been established for FUN3D simulations for SLS launch vehicle analysis. To the author's knowledge, this is the first time DDES has been used in a systematic approach and establish simulation time needed, to analyze unsteady pressure loads on a space launch vehicle such as the NASA SLS.
Unsteady surface pressure measurements on a slender delta wing undergoing limit cycle wing rock
NASA Technical Reports Server (NTRS)
Arena, Andrew S., Jr.; Nelson, Robert C.
1991-01-01
An experimental investigation of slender wing limit cycle motion known as wing rock was investigated using two unique experimental systems. Dynamic roll moment measurements and visualization data on the leading edge vortices were obtained using a free to roll apparatus that incorporates an airbearing spindle. In addition, both static and unsteady surface pressure data was measured on the top and bottom surfaces of the model. To obtain the unsteady surface pressure data a new computer controller drive system was developed to accurately reproduce the free to roll time history motions. The data from these experiments include, roll angle time histories, vortex trajectory data on the position of the vortices relative to the model's surface, and surface pressure measurements as a function of roll angle when the model is stationary or undergoing a wing rock motion. The roll time history data was numerically differentiated to determine the dynamic roll moment coefficient. An analysis of these data revealed that the primary mechanism for the limit cycle behavior was a time lag in the position of the vortices normal to the wing surface.
Effects of Transducer Installation on Unsteady Pressure Measurements on Oscillating Blades
NASA Technical Reports Server (NTRS)
Lepicovsky, Jan
2006-01-01
Unsteady pressures were measured above the suction side of a blade that was oscillated to simulate blade stall flutter. Measurements were made at blade oscillation frequencies up to 500 Hz. Two types of miniature pressure transducers were used: surface-mounted flat custom-made, and conventional miniature, body-mounted transducers. The signals of the surface-mounted transducers are significantly affected by blade acceleration, whereas the signals of body-mounted transducers are practically free of this distortion. A procedure was introduced to correct the signals of surface-mounted transducers to rectify the signal distortion due to blade acceleration. The signals from body-mounted transducers, and corrected signals from surface-mounted transducers represent true unsteady pressure signals on the surface of a blade subjected to forced oscillations. However, the use of body-mounted conventional transducers is preferred for the following reasons: no signal corrections are needed for blade acceleration, the conventional transducers are noticeably less expensive than custom-made flat transducers, the survival rate of body-mounted transducers is much higher, and finally installation of body-mounted transducers does not disturb the blade surface of interest.
Effect of Trailing Edge Shape on the Unsteady Aerodynamics of Reverse Flow Dynamic Stall
NASA Astrophysics Data System (ADS)
Lind, Andrew; Jones, Anya
2015-11-01
This work considers dynamic stall in reverse flow, where flow travels over an oscillating airfoil from the geometric trailing edge towards the leading edge. An airfoil with a sharp geometric trailing edge causes early formation of a primary dynamic stall vortex since the sharp edge acts as the aerodynamic leading edge in reverse flow. The present work experimentally examines the potential merits of using an airfoil with a blunt geometric trailing edge to delay flow separation and dynamic stall vortex formation while undergoing oscillations in reverse flow. Time-resolved and phase-averaged flow fields and pressure distributions are compared for airfoils with different trailing edge shapes. Specifically, the evolution of unsteady flow features such as primary, secondary, and trailing edge vortices is examined. The influence of these flow features on the unsteady pressure distributions and integrated unsteady airloads provide insight on the torsional loading of rotor blades as they oscillate in reverse flow. The airfoil with a blunt trailing edge delays reverse flow dynamic stall, but this leads to greater downward-acting lift and pitching moment. These results are fundamental to alleviating vibrations of high-speed helicopters, where much of the rotor operates in reverse flow.
Unsteady blade surface pressures on a large-scale advanced propeller - Prediction and data
NASA Technical Reports Server (NTRS)
Nallasamy, M.; Groeneweg, J. F.
1990-01-01
An unsteady three dimensional Euler analysis technique is employed to compute the flowfield of an advanced propeller operating at an angle of attack. The predicted blade pressure waveforms are compared with wind tunnel data at two Mach numbers, 0.5 and 0.2. The inflow angle is three degrees. For an inflow Mach number of 0.5, the predicted pressure response is in fair agreement with data: the predicted phases of the waveforms are in close agreement with data while the magnitudes are underpredicted. At the low Mach number of 0.2 (take-off) the numerical solution shows the formation of a leading edge vortex which is in qualitative agreement with measurements. However, the highly nonlinear pressure response measured on the blade suction surface is not captured in the present inviscid analysis.
Unsteady blade-surface pressures on a large-scale advanced propeller: Prediction and data
NASA Technical Reports Server (NTRS)
Nallasamy, M.; Groeneweg, J. F.
1990-01-01
An unsteady 3-D Euler analysis technique is employed to compute the flow field of an advanced propeller operating at an angle of attack. The predicted blade pressure waveforms are compared with wind tunnel data at two Mach numbers, 0.5 and 0.2. The inflow angle is three degrees. For an inflow Mach number of 0.5, the predicted pressure response is in fair agreement with data: the predicted phases of the waveforms are in close agreement with data while the magnitudes are underpredicted. At the low Mach number of 0.2 (takeoff), the numerical solution shows the formation of a leading edge vortex which is in qualitative agreement with measurements. However, the highly nonlinear pressure response measured on the blade suction surface is not captured in the present inviscid analysis.
Full-scale Force and Pressure-distribution Tests on a Tapered U.S.A. 45 Airfoil
NASA Technical Reports Server (NTRS)
Parsons, John F
1935-01-01
This report presents the results of force and pressure-distribution tests on a 2:1 tapered USA 45 airfoil as determined in the full-scale wind tunnel. The airfoil has a constant-chord center section and rounded tips and is tapered in thickness from 18 percent at the root to 9 percent at the tip. Force tests were made throughout a Reynolds Number range of approximately 2,000,000 to 8,000,000 providing data on the scale effect in addition to the conventional characteristics. Pressure-distribution data were obtained from tests at a Reynolds Number of approximately 4,000,000. The aerodynamic characteristics given by the usual dimensionless coefficients are presented graphically.
NASA Technical Reports Server (NTRS)
Alter, Stephen J.; Brauckmann, Gregory J.; Kleb, Bil; Streett, Craig L; Glass, Christopher E.; Schuster, David M.
2015-01-01
Using the Fully Unstructured Three-Dimensional (FUN3D) computational fluid dynamics code, an unsteady, time-accurate flow field about a Space Launch System configuration was simulated at a transonic wind tunnel condition (Mach = 0.9). Delayed detached eddy simulation combined with Reynolds Averaged Naiver-Stokes and a Spallart-Almaras turbulence model were employed for the simulation. Second order accurate time evolution scheme was used to simulate the flow field, with a minimum of 0.2 seconds of simulated time to as much as 1.4 seconds. Data was collected at 480 pressure taps at locations, 139 of which matched a 3% wind tunnel model, tested in the Transonic Dynamic Tunnel (TDT) facility at NASA Langley Research Center. Comparisons between computation and experiment showed agreement within 5% in terms of location for peak RMS levels, and 20% for frequency and magnitude of power spectral densities. Grid resolution and time step sensitivity studies were performed to identify methods for improved accuracy comparisons to wind tunnel data. With limited computational resources, accurate trends for reduced vibratory loads on the vehicle were observed. Exploratory methods such as determining minimized computed errors based on CFL number and sub-iterations, as well as evaluating frequency content of the unsteady pressures and evaluation of oscillatory shock structures were used in this study to enhance computational efficiency and solution accuracy. These techniques enabled development of a set of best practices, for the evaluation of future flight vehicle designs in terms of vibratory loads.
NASA Astrophysics Data System (ADS)
Izmaylov, R.; Lebedev, A.
2015-08-01
Centrifugal compressors are complex energy equipment. Automotive control and protection system should meet the requirements: of operation reliability and durability. In turbocompressors there are at least two dangerous areas: surge and rotating stall. Antisurge protecting systems usually use parametric or feature methods. As a rule industrial system are parametric. The main disadvantages of anti-surge parametric systems are difficulties in mass flow measurements in natural gas pipeline compressor. The principal idea of feature method is based on the experimental fact: as a rule just before the onset of surge rotating or precursor stall established in compressor. In this case the problem consists in detecting of unsteady pressure or velocity fluctuations characteristic signals. Wavelet analysis is the best method for detecting onset of rotating stall in spite of high level of spurious signals (rotating wakes, turbulence, etc.). This method is compatible with state of the art DSP systems of industrial control. Examples of wavelet analysis application for detecting onset of rotating stall in typical stages centrifugal compressor are presented. Experimental investigations include unsteady pressure measurement and sophisticated data acquisition system. Wavelet transforms used biorthogonal wavelets in Mathlab systems.
Causal mechanisms in airfoil-circulation formation
NASA Astrophysics Data System (ADS)
Zhu, J. Y.; Liu, T. S.; Liu, L. Q.; Zou, S. F.; Wu, J. Z.
2015-12-01
In this paper, we trace the dynamic origin, rather than any kinematic interpretations, of lift in two-dimensional flow to the physical root of airfoil circulation. We show that the key causal process is the vorticity creation by tangent pressure gradient at the airfoil surface via no-slip condition, of which the theoretical basis has been given by Lighthill ["Introduction: Boundary layer theory," in Laminar Boundary Layers, edited by L. Rosenhead (Clarendon Press, 1963), pp. 46-113], which we further elaborate. This mechanism can be clearly revealed in terms of vorticity formulation but is hidden in conventional momentum formulation, and hence has long been missing in the history of one's efforts to understand lift. By a careful numerical simulation of the flow around a NACA-0012 airfoil, and using both Eulerian and Lagrangian descriptions, we illustrate the detailed transient process by which the airfoil gains its circulation and demonstrate the dominating role of relevant dynamical causal mechanisms at the boundary. In so doing, we find that the various statements for the establishment of Kutta condition in steady inviscid flow actually correspond to a sequence of events in unsteady viscous flow.
Unsteady rarefied gas flow in a microchannel driven by a pressure difference
NASA Astrophysics Data System (ADS)
Konopel'ko, N. A.; Titarev, V. A.; Shakhov, E. M.
2016-03-01
The kinetic S-model is used to study the unsteady rarefied gas flow through a plane channel between two parallel infinite plates. Initially, the gas is at rest and is separated by the plane x = 0 with different pressure values on opposite sides. The gas deceleration effect of the channel walls is studied depending on the degree of gas rarefaction and the initial pressure drop, assuming that the molecules are diffusely reflected from the boundary. The decay of the shock wave and the disappearance of the uniform flow region behind the shock wave are monitored. Special attention is given to the gas mass flux through the cross section at x = 0, which is computed as a function of time. The asymptotic behavior of the solution at unboundedly increasing time is analyzed. The kinetic equation is solved numerically by applying a conservative finite-difference method of second-order accuracy in space.
On the attenuating effect of permeability on the low frequency sound of an airfoil
NASA Astrophysics Data System (ADS)
Weidenfeld, M.; Manela, A.
2016-08-01
The effect of structure permeability on the far-field radiation of a thin airfoil is studied. Assuming low-Mach and high-Reynolds number flow, the near- and far-field descriptions are investigated at flapping-flight and unsteady flow conditions. Analysis is carried out using thin-airfoil theory and compact-body-based calculations for the hydrodynamic and acoustic fields, respectively. Airfoil porosity is modeled via Darcy's law, governed by prescribed distribution of surface intrinsic permeability. Discrete vortex model is applied to describe airfoil wake evolution. To assess the impact of penetrability, results are compared to counterpart predictions for the sound of an impermeable airfoil. Considering the finite-chord airfoil as "acoustically transparent", the leading-order contribution of surface porosity is obtained in terms of an acoustic dipole. It is shown that, at all flow conditions considered, porosity causes attenuation in outcome sound level. This is accompanied by a time-delay in the pressure signal, reflecting the mediating effect of permeability on the interaction of fluid flow with airfoil edge points. To the extent that thin-airfoil theory holds (requiring small normal-to-airfoil flow velocities), the results indicate on a decrease of ~ 10 percent and more in the total energy radiated by a permeable versus an impermeable airfoil. This amounts to a reduction in system sound pressure level of 3 dB and above at pitching flight conditions, where the sound-reducing effect of the seepage dipole pressure becomes dominant. The applicability of Darcy's law to model the effect of material porosity is discussed in light of existing literature.
NASA Technical Reports Server (NTRS)
Weatherill, W. H.; Ehlers, F. E.
1979-01-01
The design and usage of a pilot program for calculating the pressure distributions over harmonically oscillating airfoils in transonic flow are described. The procedure used is based on separating the velocity potential into steady and unsteady parts and linearizing the resulting unsteady differential equations for small disturbances. The steady velocity potential which must be obtained from some other program, was required for input. The unsteady equation, as solved, is linear with spatially varying coefficients. Since sinusoidal motion was assumed, time was not a variable. The numerical solution was obtained through a finite difference formulation and either a line relaxation or an out of core direct solution method.
NASA Technical Reports Server (NTRS)
Jenkins, Renaldo V.; Hill, Acquilla S.; Ray, Edward J.
1988-01-01
This report presents in graphic and tabular forms the aerodynamic coefficient and surface pressure distribution data for a NASA SC(2)-0714 airfoil tested in the Langley 0.3-Meter Transonic Cryogenic Tunnel. The test was another in a series of tests involved in the joint NASA/U.S. Industry Advanced Technology Airfoil Tests program. This 14% thick supercritical airfoil was tested at Mach numbers from 0.6 to 0.76 and angles of attack from -2.0 to 6.0 degrees. The test Reynolds numbers were 4 million, 6 million, 10 million, 15 million, 30 million, 40 million, and 45 million.
NASA Technical Reports Server (NTRS)
Morgan, Harry L., Jr.
2002-01-01
This report describes the results of an experimental study conducted in the Langley Low-Turbulence Pressure Tunnel to determine the effects of Reynolds number and Mach number on the two-dimensional aerodynamic performance of the Langley Energy Efficient Transport (EET) High-Lift Airfoil. The high-lift airfoil was a supercritical-type airfoil with a thickness-to- chord ratio of 0.12 and was equipped with a leading-edge slat and a double-slotted trailing-edge flap. The leading-edge slat could be deflected -30 deg, -40 deg, -50 deg, and -60 deg, and the trailing-edge flaps could be deflected to 15 deg, 30 deg, 45 deg, and 60 deg. The gaps and overlaps for the slat and flaps were fixed at each deflection resulting in 16 different configurations. All 16 configurations were tested through a Reynolds number range of 2.5 to 18 million at a Mach number of 0.20. Selected configurations were also tested through a Mach number range of 0.10 to 0.35. The plotted and tabulated force, moment, and pressure data are available on the CD-ROM supplement L-18221.
Reduction of airfoil trailing edge noise by trailing edge blowing
NASA Astrophysics Data System (ADS)
Gerhard, T.; Erbslöh, S.; Carolus, T.
2014-06-01
The paper deals with airfoil trailing edge noise and its reduction by trailing edge blowing. A Somers S834 airfoil section which originally was designed for small wind turbines is investigated. To mimic realistic Reynolds numbers the boundary layer is tripped on pressure and suction side. The chordwise position of the blowing slot is varied. The acoustic sources, i.e. the unsteady flow quantities in the turbulent boundary layer in the vicinity of the trailing edge, are quantified for the airfoil without and with trailing edge blowing by means of a large eddy simulation and complementary measurements. Eventually the far field airfoil noise is measured by a two-microphone filtering and correlation and a 40 microphone array technique. Both, LES-prediction and measurements showed that a suitable blowing jet on the airfoil suction side is able to reduce significantly the turbulence intensity and the induced surface pressure fluctuations in the trailing edge region. As a consequence, trailing edge noise associated with a spectral hump around 500 Hz could be reduced by 3 dB. For that a jet velocity of 50% of the free field velocity was sufficient. The most favourable slot position was at 90% chord length.
Measurements in a Transitional Boundary Layer Under Low-Pressure Turbine Airfoil Conditions
NASA Technical Reports Server (NTRS)
Simon, Terrence W.; Qiu, Songgang; Yuan, Kebiao; Ashpis, David (Technical Monitor); Simon, Fred (Technical Monitor)
2000-01-01
This report presents the results of an experimental study of transition from laminar to turbulent flow in boundary layers or in shear layers over separation zones on a convex-curved surface which simulates the suction surface of a low-pressure turbine airfoil. Flows with various free-stream turbulence intensity (FSTI) values (0.5%, 2.5% and 10%), and various Reynolds numbers (50,000, 100,000 200,000 and 300,000) are investigated. Reynold numbers in the present study are based on suction surface length and passage exit mean velocity. Flow separation followed by transition within the separated flow region is observed for the lower-Re cases at each of the FSTI levels. At the highest Reynolds numbers and at elevated FSn, transition of the attached boundary layer begins before separation, and the separation zone is small. Transition proceeds in the shear layer over the separation bubble. For both the transitional boundary layer and the transitional shear layer, mean velocity, turbulence intensity and intermittency (the fraction of the time the flow is turbulent) distributions are presented. The present data are compared to published distribution models for bypass transition, intermittency distribution through transition, transition start position, and transition length. A model developed for transition of separated flows is shown to adequately predict the location of the beginning of transition, for these cases, and a model developed for transitional boundary layer flows seems to adequately predict the path of intermittency through transition when the transition start and end are known. These results are useful for the design of low-pressure turbine stages which are known to operate under conditions replicated by these tests.
Self-sustained shock oscillations on airfoils at transonic speeds
NASA Astrophysics Data System (ADS)
Lee, B. H. K.
2001-02-01
Self-sustained shock wave oscillations on airfoils at transonic flow conditions are associated with the phenomenon of buffeting. The physical mechanisms of the periodic shock motion are not yet fully understood even though experiments performed over fifty years ago have demonstrated the presence of oscillatory shock waves on the airfoil surfaces at high subsonic speeds. The unsteady pressure fluctuations generated by the low-frequency large-amplitude shock motions are highly undesirable from the structural integrity and aircraft maneuverability point of view. For modern supercritical wing design with thick profiles, the shock-induced fluctuations are particularly severe and methods to reduce the shock wave amplitudes to lower values or even to delay the oscillations to higher Mach numbers or incidence angles will result in expanding the buffet boundary of the airfoil. This review begins with a recapitulation of the classical work on shock-induced bubble separation and trailing edge separation of a turbulent boundary layer. The characteristics of the unsteady pressure fluctuations are used to classify the types of shock-boundary layer interaction. The various modes of shock wave motion for different flow conditions and airfoil configurations are described. The buffet boundaries obtained using the standard trailing edge pressure divergence technique and an alternative approach of measuring the divergence of normal fluctuating forces are compared to show the equivalence. The mechanisms of self-sustained shock oscillations are discussed for symmetrical circular-arc airfoils at zero incidence and for supercritical airfoils at high incidence angles with fully separated flows. The properties of disturbances in the wake are examined from linear stability analysis of two-dimensional compressible flows. The advances in high-speed computing make predictions of buffeting flows possible. Navier-Stokes solvers and approximate boundary layer-inviscid flow interaction methods are
Computational and Experimental Unsteady Pressures for Alternate SLS Booster Nose Shapes
NASA Technical Reports Server (NTRS)
Braukmann, Gregory J.; Streett, Craig L.; Kleb, William L.; Alter, Stephen J.; Murphy, Kelly J.; Glass, Christopher E.
2015-01-01
Delayed Detached Eddy Simulation (DDES) predictions of the unsteady transonic flow about a Space Launch System (SLS) configuration were made with the Fully UNstructured Three-Dimensional (FUN3D) flow solver. The computational predictions were validated against results from a 2.5% model tested in the NASA Ames 11-Foot Transonic Unitary Plan Facility. The peak C(sub p,rms) value was under-predicted for the baseline, Mach 0.9 case, but the general trends of high C(sub p,rms) levels behind the forward attach hardware, reducing as one moves away both streamwise and circumferentially, were captured. Frequency of the peak power in power spectral density estimates was consistently under-predicted. Five alternate booster nose shapes were assessed, and several were shown to reduce the surface pressure fluctuations, both as predicted by the computations and verified by the wind tunnel results.
Prediction of Unsteady Blade Surface Pressures on an Advanced Propeller at an Angle of Attack
NASA Technical Reports Server (NTRS)
Nallasamy, M.; Groeneweg, J. F.
1989-01-01
The numerical solution of the unsteady, three-dimensional, Euler equations is considered in order to obtain the blade surface pressures of an advanced propeller at an angle of attack. The specific configuration considered is the SR7L propeller at cruise conditions with a 4.6 deg inflow angle corresponding to the plus 2 deg nacelle tilt of the Propeller Test Assessment (PTA) flight test condition. The results indicate nearly sinusoidal response of the blade loading, with angle of attack. For the first time, detailed variations of the chordwise loading as a function of azimuthal angle are presented. It is observed that the blade is lightly loaded for part of the revolution and shocks appear from hub to about 80 percent radial station for the highly loaded portion of the revolution.
An experimental study of airfoil-spoiler aerodynamics
NASA Technical Reports Server (NTRS)
Mclachlan, B. G.; Karamcheti, K.
1985-01-01
The steady/unsteady flow field generated by a typical two dimensional airfoil with a statically deflected flap type spoiler was investigated. Subsonic wind tunnel tests were made over a range of parameters: spoiler deflection, angle of attack, and two Reynolds numbers; and comprehensive measurements of the mean and fluctuating surface pressures, velocities in the boundary layer, and velocities in the wake. Schlieren flow visualization of the near wake structure was performed. The mean lift, moment, and surface pressure characteristics are in agreement with previous investigations of spoiler aerodynamics. At large spoiler deflections, boundary layer character affects the static pressure distribution in the spoiler hingeline region; and, the wake mean velocity fields reveals a closed region of reversed flow aft of the spoiler. It is shown that the unsteady flow field characteristics are as follows: (1) the unsteady nature of the wake is characterized by vortex shedding; (2) the character of the vortex shedding changes with spoiler deflection; (3) the vortex shedding characteristics are in agreement with other bluff body investigations; and (4) the vortex shedding frequency component of the fluctuating surface pressure field is of appreciable magnitude at large spoiler deflections. The flow past an airfoil with deflected spoiler is a particular problem in bluff body aerodynamics is considered.
An algorithm to estimate unsteady and quasi-steady pressure fields from velocity field measurements.
Dabiri, John O; Bose, Sanjeeb; Gemmell, Brad J; Colin, Sean P; Costello, John H
2014-02-01
We describe and characterize a method for estimating the pressure field corresponding to velocity field measurements such as those obtained by using particle image velocimetry. The pressure gradient is estimated from a time series of velocity fields for unsteady calculations or from a single velocity field for quasi-steady calculations. The corresponding pressure field is determined based on median polling of several integration paths through the pressure gradient field in order to reduce the effect of measurement errors that accumulate along individual integration paths. Integration paths are restricted to the nodes of the measured velocity field, thereby eliminating the need for measurement interpolation during this step and significantly reducing the computational cost of the algorithm relative to previous approaches. The method is validated by using numerically simulated flow past a stationary, two-dimensional bluff body and a computational model of a three-dimensional, self-propelled anguilliform swimmer to study the effects of spatial and temporal resolution, domain size, signal-to-noise ratio and out-of-plane effects. Particle image velocimetry measurements of a freely swimming jellyfish medusa and a freely swimming lamprey are analyzed using the method to demonstrate the efficacy of the approach when applied to empirical data.
An algorithm to estimate unsteady and quasi-steady pressure fields from velocity field measurements.
Dabiri, John O; Bose, Sanjeeb; Gemmell, Brad J; Colin, Sean P; Costello, John H
2014-02-01
We describe and characterize a method for estimating the pressure field corresponding to velocity field measurements such as those obtained by using particle image velocimetry. The pressure gradient is estimated from a time series of velocity fields for unsteady calculations or from a single velocity field for quasi-steady calculations. The corresponding pressure field is determined based on median polling of several integration paths through the pressure gradient field in order to reduce the effect of measurement errors that accumulate along individual integration paths. Integration paths are restricted to the nodes of the measured velocity field, thereby eliminating the need for measurement interpolation during this step and significantly reducing the computational cost of the algorithm relative to previous approaches. The method is validated by using numerically simulated flow past a stationary, two-dimensional bluff body and a computational model of a three-dimensional, self-propelled anguilliform swimmer to study the effects of spatial and temporal resolution, domain size, signal-to-noise ratio and out-of-plane effects. Particle image velocimetry measurements of a freely swimming jellyfish medusa and a freely swimming lamprey are analyzed using the method to demonstrate the efficacy of the approach when applied to empirical data. PMID:24115059
Sting-free Unsteady Flowfield, Base Pressure and Force Measurements on Axisymmetric Bluff-Body
NASA Astrophysics Data System (ADS)
Higuchi, Hiroshi; Sawada, Hideo; Kato, Hiroyuki; Kunimasu, Tetsuya
2006-11-01
To avoid interference of model support, flowfields as well as aerodynamic force and base pressure on blunt short cylinders in axial flow were measured at Re=100,000 with the JAXA 60cm magnetic suspension and balance system. The fineness ratio ranged from 1.27 to 1.79. A digital telemeter system was developed for the base pressure measurement, and the velocity field was obtained using a PIV system. Vortices along separating shear layer and shear layer flappings with or without reattachment on the wall were observed. Downstream the cylinder in the azimuthal plane, PIV snapshots showed large-scale motion of longitudinal vortices. These instantaneous flowfields presented excellent axisymmetry when they were ensemble-averaged. Mean base pressure agreed with the drag variation at different fineness ratios. The present magnetic suspension and balance system allowed evaluation of low frequency unsteady aerodynamic force vector from feedback current to the coils and the detected small model movement. Base pressure fluctuations were compared with the drag fluctuations and discussed in light of overall flowfield phenomena.
NASA Astrophysics Data System (ADS)
Duquesne, P.; Desch"nes, C.; Iliescu, M.; Ciocan, G. D.
2010-03-01
Investigations of the flow behavior are currently carried out experimentally on models of hydraulic turbines. Quantities such as unsteady velocity can be acquired using PIV or LDV techniques, static wall pressure using steady or unsteady pressure transducers and wall shear stress using hot-film anemometry. More rarely acquired however, the unsteady total pressure at different locations in the flowstream would give more information on the flow dynamics and would be a key component for setting boundary conditions for CFD simulations. Following the example of classical Pitot tubes, which can only give averaged pressure values though, we have developed a five-hole pressure probe with embedded sensors that can measure unsteady values of total pressure, local flow velocity and direction. The probe head is designed to have a minimum impact on the flowstream, and the miniature sensors are placed in a cross configuration compared to the probe's support axis. This paper focuses on the utilization of normalized calibration coefficients and their use for unsteady values, and on the justification for using our cross sensor repartition. The calibration setup is presented briefly, including a water potential jet that requires the calculation of specific calibration coefficients. Different phenomena were observed during experimentation. Their impact on the accuracy of the probe is analyzed. The probe's operation range for this particular calibration setup is discussed. Finally, we focus on the influence of the sensors repartition on the tridimensional shape of the calibration coefficients, and we provide a way to calculate the first approximate solution for the reverse calculus while the sensors are not aligned with the probe's arm.
NASA Astrophysics Data System (ADS)
Duquesne, P.; Deschênes, C.; Iliescu, M.; Ciocan, G. D.
2009-12-01
Investigations of the flow behavior are currently carried out experimentally on models of hydraulic turbines. Quantities such as unsteady velocity can be acquired using PIV or LDV techniques, static wall pressure using steady or unsteady pressure transducers and wall shear stress using hot-film anemometry. More rarely acquired however, the unsteady total pressure at different locations in the flowstream would give more information on the flow dynamics and would be a key component for setting boundary conditions for CFD simulations. Following the example of classical Pitot tubes, which can only give averaged pressure values though, we have developed a five-hole pressure probe with embedded sensors that can measure unsteady values of total pressure, local flow velocity and direction. The probe head is designed to have a minimum impact on the flowstream, and the miniature sensors are placed in a cross configuration compared to the probe's support axis. This paper focuses on the utilization of normalized calibration coefficients and their use for unsteady values, and on the justification for using our cross sensor repartition. The calibration setup is presented briefly, including a water potential jet that requires the calculation of specific calibration coefficients. Different phenomena were observed during experimentation. Their impact on the accuracy of the probe is analyzed. The probe's operation range for this particular calibration setup is discussed. Finally, we focus on the influence of the sensors repartition on the tridimensional shape of the calibration coefficients, and we provide a way to calculate the first approximate solution for the reverse calculus while the sensors are not aligned with the probe's arm.
NASA Technical Reports Server (NTRS)
Johnson, W. G., Jr.; Hill, A. S.; Eichmann, O.
1985-01-01
A wind tunnel investigation of a NASA 12-percent-thick, advanced-technology supercritical airfoil was conducted in the Langley 0.3-Meter Transonic Cryogenic Tunnel (TCT). This investigation represents another in the series of NASA/U.S. industry two-dimensional airfoil studies to be completed in the Advanced Technology Airfoil Tests program. Test temperature was varied from 220 K to 96 K at pressures ranging from 1.2 to 4.3 atm. Mach number was varied from 0.50 to 0.80. This investigation was designed to: (1) test a NASA advanced-technology airfoil from low to flight equivalent Reynolds numbers, (2) provide experience in cryogenic wind-tunnel model design and testing techniques, and (3) demonstrate the suitability of the 0.3-m TCT as an airfoil test facility. All the test objectives were met. The pressure data are presented without analysis in tabulated format and as plots of pressure coefficient versus position on the airfoil. This report was prepared for use in conjunction with the aerodynamic coefficient data published in NASA-TM-86371. Data are included which demonstrate the effects of fixed transition. Also included are remarks on the model design and fabrication.
NASA Technical Reports Server (NTRS)
Bereketab, Semere; Wang, Hong-Wei; Mish, Patrick; Devenport, William J.
2000-01-01
Two grids have been developed for the Virginia Tech 6 ft x 6 ft Stability wind tunnel for the purpose of generating homogeneous isotropic turbulent flows for the study of unsteady airfoil response. The first, a square bi-planar grid with a 12" mesh size and an open area ratio of 69.4%, was mounted in the wind tunnel contraction. The second grid, a metal weave with a 1.2 in. mesh size and an open area ratio of 68.2% was mounted in the tunnel test section. Detailed statistical and spectral measurements of the turbulence generated by the two grids are presented for wind tunnel free stream speeds of 10, 20, 30 and 40 m/s. These measurements show the flows to be closely homogeneous and isotropic. Both grids produce flows with a turbulence intensity of about 4% at the location planned for the airfoil leading edge. Turbulence produced by the large grid has an integral scale of some 3.2 inches here. Turbulence produced by the small grid is an order of magnitude smaller. For wavenumbers below the upper limit of the inertial subrange, the spectra and correlations measured with both grids at all speeds can be represented using the von Karman interpolation formula with a single velocity and length scale. The spectra maybe accurately represented over the entire wavenumber range by a modification of the von Karman interpolation formula that includes the effects of dissipation. These models are most accurate at the higher speeds (30 and 40 m/s).
NASA Technical Reports Server (NTRS)
St.hilaire, A. O.; Carta, F. O.
1983-01-01
The analysis of the chordwise load distribution and its sensitivity to the various system parameters represents the next phase of the overall study and is the subject of the present two volume report. The present volume is a compilation of all of the time history response data obtained during the test program previously described. The data have been tabulated in the form of Fourier coefficients for reasons of compactness and for ease by the user to reproduce the unsteady component of the individual pressures and the complete (unsteady plus steady state components) integrated load results. This data volume contains the individual pressure response time histories along the chord followed by the corresponding integrated load results. A further description of these data tables can be found in the text that follows.
A supercritical airfoil experiment
NASA Technical Reports Server (NTRS)
Mateer, G. G.; Seegmiller, H. L.; Hand, L. A.; Szodruck, J.
1994-01-01
The purpose of this investigation is to provide a comprehensive data base for the validation of numerical simulations. The objective of the present paper is to provide a tabulation of the experimental data. The data were obtained in the two-dimensional, transonic flowfield surrounding a supercritical airfoil. A variety of flows were studied in which the boundary layer at the trailing edge of the model was either attached or separated. Unsteady flows were avoided by controlling the Mach number and angle of attack. Surface pressures were measured on both the model and wind tunnel walls, and the flowfield surrounding the model was documented using a laser Doppler velocimeter (LDV). Although wall interference could not be completely eliminated, its effect was minimized by employing the following techniques. Sidewall boundary layers were reduced by aspiration, and upper and lower walls were contoured to accommodate the flow around the model and the boundary-layer growth on the tunnel walls. A data base with minimal interference from a tunnel with solid walls provides an ideal basis for evaluating the development of codes for the transonic speed range because the codes can include the wall boundary conditions more precisely than interference connections can be made to the data sets.
NASA Technical Reports Server (NTRS)
Stivers, Louis S.; Abbott, Ira H.; von Doenhoff, Albert E.
1945-01-01
Recent airfoil data for both flight and wind-tunnel tests have been collected and correlated insofar as possible. The flight data consist largely of drag measurements made by the wake-survey method. Most of the data on airfoil section characteristics were obtained in the Langley two-dimensional low-turbulence pressure tunnel. Detail data necessary for the application of NACA 6-serles airfoils to wing design are presented in supplementary figures, together with recent data for the NACA 24-, 44-, and 230-series airfoils. The general methods used to derive the basic thickness forms for NACA 6- and 7-series airfoils and their corresponding pressure distributions are presented. Data and methods are given for rapidly obtaining the approximate pressure distributions for NACA four-digit, five-digit, 6-, and 7-series airfoils. The report includes an analysis of the lift, drag, pitching-moment, and critical-speed characteristics of the airfoils, together with a discussion of the effects of surface conditions. Available data on high-lift devices are presented. Problems associated with lateral-control devices, leading-edge air intakes, and interference are briefly discussed. The data indicate that the effects of surface condition on the lift and drag characteristics are at least as large as the effects of the airfoil shape and must be considered in airfoil selection and the prediction of wing characteristics. Airfoils permitting extensive laminar flow, such as the NACA 6-series airfoils, have much lower drag coefficients at high speed and cruising lift coefficients than earlier types-of airfoils if, and only if, the wing surfaces are sufficiently smooth and fair. The NACA 6-series airfoils also have favorable critical-speed characteristics and do not appear to present unusual problems associated with the application of high-lift and lateral-control devices. Much of the data given in the NACA Advance Confidential Report entitled "Preliminary Low-Drag-Airfoil and Flap Data from
An unsteady microfluidic T-form mixer perturbed by hydrodynamic pressure
Ma, Yanbao; Sun, Chien-Pin; Fields, Michael; Li, Yang; Haake, David A; Churchill, Bernard M; Ho, Chih-Ming
2009-01-01
An unsteady microfluidic T-form mixer driven by pressure disturbances was designed and investigated. The performance of the mixer was examined both through numerical simulation and experimentation. Linear Stokes equations were used for these low Reynolds number flows. Unsteady mixing in a micro-channel of two aqueous solutions differing in concentrations of chemical species was described using a convection-dominated diffusion equation. The task was greatly simplified by employing linear superimposition of a velocity field for solving a scalar species concentration equation. Low-order-based numerical codes were found not to be suitable for simulation of a convection-dominated mixing process due to erroneous computational dissipation. The convection-dominated diffusion problem was addressed by designing a numerical algorithm with high numerical accuracy and computational-cost effectiveness. This numerical scheme was validated by examining a test case prior to being applied to the mixing simulation. Parametric analysis was performed using this newly developed numerical algorithm to determine the best mixing conditions. Numerical simulation identified the best mixing condition to have a Strouhal number (St)of 0.42. For a T-junction mixer (with channel width = 196 μm), about 75% mixing can be finished within a mixing distance of less than 3 mm (i.e. 15 channel width) at St = 0.42 for flow with a Reynolds number less than 0.24. Numerical results were validated experimentally by mixing two aqueous solutions containing yellow and blue dyes. Visualization of the flow field under the microscope revealed a high level of agreement between numerical simulation and experimental results. PMID:19177174
NASA Technical Reports Server (NTRS)
Carta, F. O.
1981-01-01
Computer data are provided for tests conducted on a linear cascade of airfoils oscillating in pitch to measure the unsteady pressure response on selected blades along the leading edge plane of the cascade, over the chord of the center blade, and on the sidewall in the plane of the leading edge.
Water Flow Testing and Unsteady Pressure Analysis of a Two-Bladed Liquid Oxidizer Pump Inducer
NASA Technical Reports Server (NTRS)
Schwarz, Jordan B.; Mulder, Andrew; Zoladz, Thomas
2011-01-01
The unsteady fluid dynamic performance of a cavitating two-bladed oxidizer turbopump inducer was characterized through sub-scale water flow testing. While testing a novel inlet duct design that included a cavitation suppression groove, unusual high-frequency pressure oscillations were observed. With potential implications for inducer blade loads, these high-frequency components were analyzed extensively in order to understand their origins and impacts to blade loading. Water flow testing provides a technique to determine pump performance without the costs and hazards associated with handling cryogenic propellants. Water has a similar density and Reynolds number to liquid oxygen. In a 70%-scale water flow test, the inducer-only pump performance was evaluated. Over a range of flow rates, the pump inlet pressure was gradually reduced, causing the flow to cavitate near the pump inducer. A nominal, smooth inducer inlet was tested, followed by an inlet duct with a circumferential groove designed to suppress cavitation. A subsequent 52%-scale water flow test in another facility evaluated the combined inducer-impeller pump performance. With the nominal inlet design, the inducer showed traditional cavitation and surge characteristics. Significant bearing loads were created by large side loads on the inducer during synchronous cavitation. The grooved inlet successfully mitigated these loads by greatly reducing synchronous cavitation, however high-frequency pressure oscillations were observed over a range of frequencies. Analytical signal processing techniques showed these oscillations to be created by a rotating, multi-celled train of pressure pulses, and subsequent CFD analysis suggested that such pulses could be created by the interaction of rotating inducer blades with fluid trapped in a cavitation suppression groove. Despite their relatively low amplitude, these high-frequency pressure oscillations posed a design concern due to their sensitivity to flow conditions and
Measurement of Unsteady Pressure Data on a Large HSCT Semispan Wing and Comparison with Analysis
NASA Technical Reports Server (NTRS)
Scott, Robert C.; Silva, Walter A.; Florance, James R.; Keller, Donald F.
2002-01-01
Experimental data from wind-tunnel tests of the Rigid Semispan Model (RSM) performed at NASA Langley's Transonic Dynamics Tunnel (TDT) are presented. The primary focus of the paper is on data obtained from testing of the RSM on the Oscillating Turntable (OTT). The OTT is capable of oscillating models in pitch at various amplitudes and frequencies about mean angles of attack. Steady and unsteady pressure data obtained during testing of the RSM on the OTT is presented and compared to data obtained from previous tests of the RSM on a load balance and on a Pitch and Plunge Apparatus (PAPA). Testing of the RSM on the PAPA resulted in utter boundaries that were strongly dependent on angle of attack across the Mach number range. Pressure data from all three tests indicates the existence of vortical flows at moderate angles of attack. The correlation between the vortical flows and the unusual utter boundaries from the RSM/PAPA test is discussed. Comparisons of experimental data with analyses using the CFL3Dv6 computational fluid dynamics code are presented.
NASA Astrophysics Data System (ADS)
Masnadi, N.; Duncan, J. H.
2013-11-01
The non-linear response of a water surface to a slow-moving pressure distribution is studied experimentally using a vertically oriented carriage-mounted air-jet tube that is set to translate over the water surface in a long tank. The free surface deformation pattern is measured with a full-field refraction-based method that utilizes a vertically oriented digital movie camera (under the tank) and a random dot pattern (above the water surface). At towing speeds just below the minimum phase speed of gravity-capillary waves (cmin ~ 23 cm/s), an unsteady V-shaped pattern is formed behind the pressure source. Localized depressions are generated near the source and propagate in pairs along the two arms of the V-shaped pattern. These depressions are eventually shed from the tips of the pattern at a frequency of about 1 Hz. It is found that the shape and phase speeds of the first depressions shed in each run are quantitatively similar to the freely-propagating gravity-capillary lumps from potential flow calculations. In the experiments, the amplitudes of the depressions decrease by approximately 60 percent while travelling 12 wavelengths. The depressions shed later in each run behave in a less consistent manner, probably due to their interaction with neighboring depressions.
NASA Technical Reports Server (NTRS)
Allison, Dennis O.; Sewall, William G.
1995-01-01
Longitudinal characteristics and wing-section pressure distributions are compared for the EA-6B airplane with and without airfoil modifications. The airfoil modifications were designed to increase low-speed maximum lift for maneuvering, while having a minimal effect on transonic performance. Section contour changes were confined to the leading-edge slat and trailing-edge flap regions of the wing. Experimental data are analyzed from tests in the Langley 16-Foot Transonic Tunnel on the baseline and two modified wing-fuselage configurations with the slats and flaps in their retracted positions. Wing modification effects on subsonic and transonic performance are seen in wing-section pressure distributions of the various configurations at similar lift coefficients. The modified-wing configurations produced maximum lift coefficients which exceeded those of the baseline configuration at low-speed Mach numbers (0.300 and 0.400). This benefit was related to the behavior of the wing upper surface leading-edge suction peak and the behavior of the trailing-edge pressure. At transonic Mach numbers (0.725 to 0.900), the wing modifications produced a somewhat stronger nose-down pitching moment, a slightly higher drag at low-lift levels, and a lower drag at higher lift levels.
Transonic flow about a thick circular-arc airfoil
NASA Technical Reports Server (NTRS)
Mcdevitt, J. B.; Levy, L. L., Jr.; Deiwert, G. S.
1975-01-01
An experimental and theoretical study of transonic flow over a thick airfoil, prompted by a need for adequately documented experiments that could provide rigorous verification of viscous flow simulation computer codes, is reported. Special attention is given to the shock-induced separation phenomenon in the turbulent regime. Measurements presented include surface pressures, streamline and flow separation patterns, and shadowgraphs. For a limited range of free-stream Mach numbers the airfoil flow field is found to be unsteady. Dynamic pressure measurements and high-speed shadowgraph movies were taken to investigate this phenomenon. Comparisons of experimentally determined and numerically simulated steady flows using a new viscous-turbulent code are also included. The comparisons show the importance of including an accurate turbulence model. When the shock-boundary layer interaction is weak the turbulence model employed appears adequate, but when the interaction is strong, and extensive regions of separation are present, the model is inadequate and needs further development.
Control of Vortex Shedding on an Airfoil using Mini Flaps at Low Reynolds Number
NASA Astrophysics Data System (ADS)
Oshiyama, Daisuke; Numata, Daiju; Asai, Keisuke
2015-11-01
In this study, the effects of mini flaps (MFs) on a NACA0012 airfoil were investigated experimentally at low Reynolds number. MFs are small flat plates attached to the trailing edge of an airfoil perpendicularly. All the tests were conducted at the Tohoku-University Basic Aerodynamic Research Tunnel at the chord Reynolds number of 25,000. Aerodynamic forces were measured using a 3-component balance and the surface flow was visualized by luminescent oil film technique. The results of force measurement show that attachment of MFs enhances lift and the enhanced lift increases with MF height. On the other hand, the results of oil flow visualization show that attachment of MFs enlarges the separated region on the airfoil rather than diminishes it. To understand the physical mechanism of MFs for lift enhancement, the flow around the airfoil was visualized by the smoke-wire method and the wake profile behind the airfoil was measured using a hot wire anemometer. It was found that vortices shed periodically from the tip of the MFs and interact with the separated shear layer from the upper surface. This unsteady vortex shedding forms a low-pressure region on the upper surface, generating higher lift. These results suggest that the height of MFs controls the frequency of vortex shedding behind the MF, forcing the separated shear layer on the upper surface flow in unsteady manner.
The passage of a distorted velocity field through a cascade of airfoils
NASA Technical Reports Server (NTRS)
Adamczyk, J. J.
1976-01-01
An analysis has been developed to predict the unsteady force and moment generated by the passage of a timewise periodic total pressure distortion through an arbitrary cascade of airfoils. The mathematical formulation of this analysis is based on the assumption that the magnitudes of the timewise fluctuations of the variables which describe the flow field are small compared to their time average values. This assumption permits the development of a linear unsteady perturbation analysis about a steady flow field. In addition to this linearization assumption the fluid medium is assumed to be incompressible and inviscid. The mathematical development begins by decomposing the velocity field surrounding an infinite cascade of airfoils into its irrotational and rotational components. The rotational component is associated with an upstream unsteady total pressure distortion and is defined in terms of the vorticity field associated with the distortion pattern. The irrotational component is further decomposed into a steady and unsteady part. A combined analytical and numerical procedure has been developed to solve the field equations which govern the rotational and irrotational velocity fields. Results of this analysis show a strong influence of mean loading on the unsteady force generated by the passage of a one dimensional gust through a cascade of compressor blades.
Aerodynamic Flow Control of a Maneuvering Airfoil
NASA Astrophysics Data System (ADS)
Brzozowski, Daniel P.; Culp, John; Glezer, Ari
2010-11-01
The unsteady aerodynamic forces and moments on a maneuvering, free-moving airfoil are varied in wind tunnel experiments by controlling vorticity generation/accumulation near the surface using hybrid synthetic jet actuators. The dynamic characteristics of the airfoil that is mounted on a 2-DOF traverse are controlled using position and attitude feedback loops that are actuated by servo motors. Bi-directional changes in the pitching moment are induced using controllable trapped vorticity concentrations on the suction and pressure surfaces near the trailing edge. The dynamic coupling between the actuation and the time-dependent flow field is characterized using simultaneous force and velocity measurements that are taken phase-locked to the commanded actuation waveform. The time scales associated with the actuation process is determined from PIV measurements of vorticity flux downstream of the trailing edge. Circulation time history shows that the entire flow over the airfoil readjusts within about 1.5 TCONV, which is about two orders of magnitude shorter than the characteristic time associated with the controlled maneuver of the wind tunnel model. This illustrates that flow-control actuation can be typically effected on time scales commensurate with the flow's convective time scale, and that the maneuver response is only limited by the inertia of the platform. Supported by AFSOR.
NASA Technical Reports Server (NTRS)
Ramsey, John K.; Erwin, Dan
2004-01-01
An experimental influence coefficient technique was used to obtain unsteady aerodynamic influence coefficients and, consequently, unsteady pressures for a cascade of symmetric airfoils oscillating in pitch about mid-chord. Stagger angles of 0 deg and 10 deg were investigated for a cascade with a gap-to-chord ratio of 0.417 operating at an axial Mach number of 1.9, resulting in a supersonic leading-edge locus. Reduced frequencies ranged from 0.056 to 0.2. The influence coefficients obtained determine the unsteady pressures for any interblade phase angle. The unsteady pressures were compared with those predicted by several algorithms for interblade phase angles of 0 deg and 180 deg.
NASA Technical Reports Server (NTRS)
Wentz, W. H., Jr.; Fiscko, K. A.
1979-01-01
Force and surface pressure distributions were measured for the 21% LS(1)-0421 modified airfoil fitted with 20% aileron, 25% slotted flap and 10% slot lip spoiler. All tests were conducted at a Reynolds number of 2.2 x 10 to the 6th power and a Mach number of 0.13. The lift, drag, pitching moments, control surface normal force and hinge moments, and surface pressure distributions are included in the results. Incremental performance of flap and aileron are discussed and compared to the GA(W)-2 airfoil. Spoiler control which shows a slight reversal tendency at high alpha, is examined.
NASA Technical Reports Server (NTRS)
Suzen, Y. B.; Huang, P. G.; Hultgren, Lennart S.; Ashpis, David E.
2003-01-01
A new transport equation for the intermittency factor was proposed to predict separated and transitional boundary layers under low-pressure turbine airfoil conditions. The intermittent behavior of the transitional flows is taken into account and incorporated into computations by modifying the eddy viscosity, t , with the intermittency factor, y. Turbulent quantities are predicted by using Menter s two-equation turbulence model (SST). The intermittency factor is obtained from a transport equation model, which not only can reproduce the experimentally observed streamwise variation of the intermittency in the transition zone, but also can provide a realistic cross-stream variation of the intermittency profile. In this paper, the intermittency model is used to predict a recent separated and transitional boundary layer experiment under low pressure turbine airfoil conditions. The experiment provides detailed measurements of velocity, turbulent kinetic energy and intermittency profiles for a number of Reynolds numbers and freestream turbulent intensity conditions and is suitable for validation purposes. Detailed comparisons of computational results with experimental data are presented and good agreements between the experiments and predictions are obtained.
Multiple piece turbine airfoil
Kimmel, Keith D; Wilson, Jr., Jack W.
2010-11-02
A turbine airfoil, such as a rotor blade or a stator vane, for a gas turbine engine, the airfoil formed as a shell and spar construction with a plurality of dog bone struts each mounted within openings formed within the shell and spar to allow for relative motion between the spar and shell in the airfoil chordwise direction while also forming a seal between adjacent cooling channels. The struts provide the seal as well as prevent bulging of the shell from the spar due to the cooling air pressure.
Aerodynamic characteristics and pressure distributions for an executive-jet baseline airfoil section
NASA Technical Reports Server (NTRS)
Allison, Dennis O.; Mineck, Raymond E.
1993-01-01
A wind tunnel test of an executive-jet baseline airfoil model was conducted in the adaptive-wall test section of the NASA Langley 0.3-Meter Transonic Cryogenic Tunnel. The primary goal of the test was to measure airfoil aerodynamic characteristics over a wide range of flow conditions that encompass two design points. The two design Mach numbers were 0.654 and 0.735 with corresponding Reynolds numbers of 4.5 x 10(exp 6) and 8.9 x 10(exp 6) based on chord, respectively, and normal-force coefficients of 0.98 and 0.51, respectively. The tests were conducted over a Mach number range from 0.250 to 0.780 and a chord Reynolds number range from 3 x 10(exp 6) to 18 x 10(exp 6). The angle of attack was varied from -2 deg to a maximum below 10 deg with one exception in which the maximum was 14 deg for a Mach number of 0.250 at a chord Reynolds number of 4.5 x 10(exp 6). Boundary-layer transition was fixed at 5 percent of chord on both the upper and lower surfaces of the model for most of the test. The adaptive-wall test section had flexible top and bottom walls and rigid sidewalls. Wall interference was minimized by the movement of the adaptive walls, and the airfoil aerodynamic characteristics were corrected for any residual top and bottom wall interference.
NASA Technical Reports Server (NTRS)
Gregorek, G. M.; Kuniega, R. J.; Nyland, T. W.
1988-01-01
The aerodynamic similarity between a small (4-inch chord) wind tunnel model and a full-scale wind turbine blade (24-foot tip section with a 36-inch chord) was evaluated by comparing selected pressure distributions around the geometrically similar cross sections. The airfoils were NACA 64-621 sections, including trailing-edge ailerons with a width equal to 38 percent of the airfoil chord. The model airfoil was tested in the OSU 6- by 12-inch High Reynolds Number Wind Tunnel; the full-scale blade section was tested in the NASA Langley Research Center 30- by 60-foot Subsonic Wind Tunnel. The model airfoil contained 61 pressure taps connected by embedded tubes to pressure transducers. A belt containing 29 pressure taps was fixed to the full-scale section at midspan to obtain surface pressure data. Lift coefficients were obtained by integrating pressures, and corrections were made for the 3-D effects of blade twist and downwash in the blade tip section. The results of the two different experimental methods correlated well for angles of attack from minus 4 to 36 degrees and aileron reflections from 0 to 90 degrees.
Supercritical flow about a thick circular-arc airfoil
NASA Technical Reports Server (NTRS)
Mcdevitt, J. B.
1979-01-01
The supercritical flow about a biconvex circular-arc airfoil is being thoroughly documented at Ames Research Center in order to provide experimental test cases suitable for guiding and evaluating current and future computer codes. The effects of angle of attack, effects of leading and trailing-edge splitter plates, additional unsteady pressure fluctuation (buffeting) measurements and glow-field shadowgraphs, and application of an oil-film technique to display separated-wake streamlines were studied. Computed and measured pressure distributions for steady and unsteady flows, using a recent computer code representative of current methodology, are compared. It was found that the numerical solutions are often fundamentally incorrect in that only strong (shock-polar terminology) shocks are captured, whereas experimentally, both strong and weak shock waves appear.
NASA Technical Reports Server (NTRS)
Street, William G; Ames, Milton B
1939-01-01
Pressure-distribution tests of an N.A.C.A. 0009 airfoil with a 50-percent-chord plain flap and three plain tabs, having chords 10, 20, and 30 percent of the flap chord, were made in the N.A.C.A. 4- by 6- foot vertical tunnel. The tests supplied aerodynamic section data that may be applied to the design of horizontal and vertical tail surfaces. The results are presented as resultant-pressure diagrams for the airfoil with the flap and the 20-percent-chord tab. Plots are also given of increments of normal-force and hinge-moment coefficients for the airfoil, the flap, and the three tabs. The experimental results and values computed by analytical methods are in good agreement for small flap and tab deflections. The results of the tests indicated that the effectiveness of all three tab sizes in reducing flap hinge moments decreased with increasing flap deflection.
Static and unsteady pressure measurements on a 50 degree clipped delta wing at M = 0.9
NASA Technical Reports Server (NTRS)
Hess, R. W.; Wynne, E. C.; Cazier, F. W.
1982-01-01
Static and unsteady pressures were measured on a 50.45 degree clipped delta wing in the Langley Transonic Dynamics Tunnel with Freon as the test medium. Data taken at M = 0.9 is presented for static and oscillatory deflections of the trailing edge control surface and for the wing in pitch. Comparisons of the static measured data are made with results computed using the Bailey-Ballhaus small disturbance code.
Frey, Gary A.; Twardochleb, Christopher Z.
1998-01-01
Past airfoil configurations have been used to improve aerodynamic performance and engine efficiencies. The present airfoil configuration further increases component life and reduces maintenance by reducing internal stress within the airfoil itself. The airfoil includes a chord and a span. Each of the chord and the span has a bow being summed to form a generally "C" configuration of the airfoil. The generally "C" configuration includes a compound bow in which internal stresses resulting from a thermal temperature gradient are reduced. The structural configuration reduces internal stresses resulting from thermal expansion.
Frey, G.A.; Twardochleb, C.Z.
1998-01-13
Past airfoil configurations have been used to improve aerodynamic performance and engine efficiencies. The present airfoil configuration further increases component life and reduces maintenance by reducing internal stress within the airfoil itself. The airfoil includes a chord and a span. Each of the chord and the span has a bow being summed to form a generally ``C`` configuration of the airfoil. The generally ``C`` configuration includes a compound bow in which internal stresses resulting from a thermal temperature gradient are reduced. The structural configuration reduces internal stresses resulting from thermal expansion. 6 figs.
NASA Technical Reports Server (NTRS)
Wentz, W. H., Jr.; Fiscko, K. A.
1978-01-01
Surface pressure distributions were measured for the 13% thick GA(W)-2 airfoil section fitted with 20% aileron, 25% slotted flap and 30% Fowler flap. All tests were conducted at a Reynolds number of 2.2 x 10 to the 6th power and a Mach number of 0.13. Pressure distribution and force and moment coefficient measurements are compared with theoretical results for a number of cases. Agreement between theory and experiment is generally good for low angles of attack and small flap deflections. For high angles and large flap deflections where regions of separation are present, the theory is inadequate. Theoretical drag predictions are poor for all flap-extended cases.
An improved method for calculating flow past flapping and hovering airfoils
NASA Astrophysics Data System (ADS)
Sengupta, T. K.; Vikas, V.; Johri, A.
2005-12-01
A method is reported here for calculating unsteady aerodynamics of hovering and flapping airfoil for two-dimensional flow via the following improved methodologies: (a) a correct formulation of the problem using stream function (ψ) and vorticity (ω) as dependent variables; (b) calculating loads and moment by a new method to solve the governing pressure Poisson equation (PPE) in a truncated part of the computational domain on a nonstaggered grid; (c) accurate solution using high accuracy compact difference scheme for the vorticity transport equation (VTE) and (d) accelerating the computations by using a high-order filter after each time step of integration. These have been used to solve Navier-Stokes equation for flow past flapping and hovering NACA 0014 and 0015 airfoils at typical Reynolds numbers relevant to the study of unsteady aerodynamics of micro air vehicle (MAV) and insect/bird flight.
NASA Technical Reports Server (NTRS)
Mueller, T. J. (Editor)
1985-01-01
Topics of interest in the design, flow modeling and visualization, and turbulence and flow separation effects for low Reynolds number (Re) airfoils are discussed. Design methods are presented for Re from 50,000-500,000, including a viscous-inviscid coupling method and by using a constrained pitching moment. The effects of pressure gradients, unsteady viscous aerodynamics and separation bubbles are investigated, with particular note made of factors which most influence the size and location of separation bubbles and control their effects. Attention is also given to experimentation with low Re airfoils and to numerical models of symmetry breaking and lift hysteresis from separation. Both steady and unsteady flow experiments are reviewed, with the trials having been held in wind tunnels and the free atmosphere. The topics discussed are of interest to designers of RPVs, high altitude aircraft, sailplanes, ultralights and wind turbines.
Unsteady potential flow past a propeller blade section
NASA Technical Reports Server (NTRS)
Takallu, M. A.
1990-01-01
An analytical study was conducted to predict the effect of an oscillating stream on the time dependent sectional pressure and lift coefficients of a model propeller blade. The assumption is that as the blade sections encounter a wake, the actual angles of attack vary in a sinusoidal manner through the wake, thus each blade is exposed to an unsteady stream oscillating about a mean value at a certain reduced frequency. On the other hand, an isolated propeller at some angle of attack can experience periodic changes in the value of the flow angle causing unsteady loads on the blades. Such a flow condition requires the inclusion of new expressions in the formulation of the unsteady potential flow around the blade sections. These expressions account for time variation of angle of attack and total shed vortices in the wake of each airfoil section. It was found that the final expressions for the unsteady pressure distribution on each blade section are periodic and that the unsteady circulation and lift coefficients exhibit a hysteresis loop.
Three-dimensional unsteady flow calculations in an advanced gas generator turbine
NASA Technical Reports Server (NTRS)
Rangwalla, Akil A.
1993-01-01
This paper deals with the application of a three-dimensional, unsteady Navier-Stokes code for predicting the unsteady flow in a single stage of an advanced gas generator turbine. The numerical method solves the three-dimensional thin-layer Navier-Stokes equations, using a system of overlaid grids, which allow for relative motion between the rotor and stator airfoils. Results in the form of time averaged pressures and pressure amplitudes on the airfoil surfaces will be shown. In addition, instantaneous contours of pressure, Mach number, etc. will be presented in order to provide a greater understanding of the inviscid as well as the viscous aspects of the flowfield. Also, relevant secondary flow features such as cross-plane velocity vectors and total pressure contours will be presented. Prior work in two-dimensions has indicated that for the advanced designs, the unsteady interactions can play a significant role in turbine performance. These interactions affect not only the stage efficiency but can substantially alter the time-averaged features of the flow. This work is a natural extension of the work done in two-dimensions and hopes to address some of the issues raised by the two-dimensional calculations. These calculations are being performed as an integral part of an actual design process and demonstrate the value of unsteady rotor-stator interaction calculations in the design of turbomachines.
NASA Technical Reports Server (NTRS)
Sivells, James C; Spooner, Stanley H
1949-01-01
Report presents the results of an investigation conducted in the Langley 19-foot pressure tunnel to determine the maximum lift and stalling characteristics of two thin wings equipped with several types of flaps. Split, single slotted, and double slotted flaps were tested on one wing which had NACA 65-210 airfoil sections and split and double slotted flaps were tested on the other, which had NACA 64-210 airfoil sections. Both wings were zero sweep, an aspect ratio of 9, and a taper ratio of 0.4.
Two-dimensional wind tunnel test of an oscillating rotor airfoil, volume 1
NASA Technical Reports Server (NTRS)
Dadone, L. U.
1977-01-01
A two dimensional wind tunnel test was conducted to obtain the quasisteady and unsteady characteristics of an advanced airfoil designed for helicopter rotor applications. Differential pressures were measured at 17 locations along the chord of the airfoil model. The airfoil motions were sinusoidal forced-pitch oscillations about the quarter chord at amplitudes varying from 2.5 to 10.0 degrees and at frequencies from 23 Hz to 90 Hz. The quasisteady tests were conducted at Mach numbers from 0.2 to 0.9, and the oscillatory tests between M = 0.2 and M = 0.7. At quasisteady conditions a limited number of drag measurements were made with a wake-traversing probe.
NASA Technical Reports Server (NTRS)
Dussauge, J. P.; Debieve, J. F.
1980-01-01
The amplification or reduction of unsteady velocity perturbations under the influence of strong flow acceleration or deceleration was studied. Supersonic flows with large velocity, pressure gradients, and the conditions in which the velocity fluctuations depend on the action of the average gradients of pressure and velocity rather than turbulence, are described. Results are analyzed statistically and interpreted as a return to laminar process. It is shown that this return to laminar implies negative values in the turbulence production terms for kinetic energy. A simple geometrical representation of the Reynolds stress production is given.
NASA Technical Reports Server (NTRS)
Johnson, W. G., Jr.; Hill, A. S.
1985-01-01
A wind-tunnel investigation designed to test a Boeing advanced-technology airfoil from low to flight-equivalent Reynolds numbers has been completed in the Langley 0.3-Meter Transonic Cryogenic Tunnel. This investigation represents the first in a series of NASA/U.S. industry two-dimensional airfoil studies to be completed in the Advanced Technology Airfoil Test program. Test temperature was varied from ambient to about 100 K at pressures ranging from about 1.2 to 6.0 atm. Mach number was varied from about 0.40 to 0.80. These variables provided a Reynolds number (based on airfoil chord) range from 4.4 X 10 to the 6th power to 50.0 X 10 to the 6th power. All the test objectives were met. The pressure data are presented without analysis in plotted and tabulated formats for use in conjunction with the aerodynamic coefficient data published as NASA TM-81922. At the time of the test, these pressure data were considered proprietary and have only recently been made available by Boeing for general release. Data are included which demonstrate the effects of fixed transition. Also included are remarks on the model design, the model structural integrity, and the overall test experience.
Comparison of Analytical and CAA Solutions: Single Airfoil Gust Response
NASA Technical Reports Server (NTRS)
Scott, James R.
2000-01-01
The purpose of this problem is to test the ability of a computational fluid dynamics/computational aeroacoustics code to accurately predict the unsteady aerodynamic and aeroacoustic response of a single airfoil to a two-dimensional, periodic vortical gust.
Active Control of Flow Separation Over an Airfoil
NASA Technical Reports Server (NTRS)
Ravindran, S. S.
1999-01-01
Designing an aircraft without conventional control surfaces is of interest to aerospace community. In this direction, smart actuator devices such as synthetic jets have been proposed to provide aircraft maneuverability instead of control surfaces. In this article, a numerical study is performed to investigate the effects of unsteady suction and blowing on airfoils. The unsteady suction and blowing is introduced at the leading edge of the airfoil in the form of tangential jet. Numerical solutions are obtained using Reynolds-Averaged viscous compressible Navier-Stokes equations. Unsteady suction and blowing is investigated as a means of separation control to obtain lift on airfoils. The effect of blowing coefficients on lift and drag is investigated. The numerical simulations are compared with experiments from the Tel-Aviv University (TAU). These results indicate that unsteady suction and blowing can be used as a means of separation control to generate lift on airfoils.
NASA Technical Reports Server (NTRS)
Batina, John T.
1990-01-01
Improved algorithm for the solution of the time-dependent Euler equations are presented for unsteady aerodynamic analysis involving unstructured dynamic meshes. The improvements were developed recently to the spatial and temporal discretizations used by unstructured grid flow solvers. The spatial discretization involves a flux-split approach which is naturally dissipative and captures shock waves sharply with at most one grid point within the shock structure. The temporal discretization involves an implicit time-integration scheme using a Gauss-Seidel relaxation procedure which is computationally efficient for either steady or unsteady flow problems. For example, very large time steps may be used for rapid convergence to steady state, and the step size for unsteady cases may be selected for temporal accuracy rather than for numerical stability. Steady and unsteady flow results are presented for the NACA 0012 airfoil to demonstrate applications of the new Euler solvers. The unsteady results were obtained for the airfoil pitching harmonically about the quarter chord. The resulting instantaneous pressure distributions and lift and moment coefficients during a cycle of motion compare well with experimental data. A description of the Euler solvers is presented along with results and comparisons which assess the capability.
NASA Technical Reports Server (NTRS)
Batina, John T.
1990-01-01
Improved algorithms for the solution of the time-dependent Euler equations are presented for unsteady aerodynamic analysis involving unstructured dynamic meshes. The improvements have been developed recently to the spatial and temporal discretizations used by unstructured grid flow solvers. The spatial discretization involves a flux-split approach which is naturally dissipative and captures shock waves sharply with at most one grid point within the shock structure. The temporal discretization involves an implicit time-integration shceme using a Gauss-Seidel relaxation procedure which is computationally efficient for either steady or unsteady flow problems. For example, very large time steps may be used for rapid convergence to steady state, and the step size for unsteady cases may be selected for temporal accuracy rather than for numerical stability. Steady and unsteady flow results are presented for the NACA 0012 airfoil to demonstrate applications of the new Euler solvers. The unsteady results were obtained for the airfoil pitching harmonically about the quarter chord. The resulting instantaneous pressure distributions and lift and moment coefficients during a cycle of motion compare well with experimental data. The paper presents a description of the Euler solvers along with results and comparisons which assess the capability.
NASA Technical Reports Server (NTRS)
Riffel, R. E.; Rothrock, M. D.
1980-01-01
A two dimensional cascade of harmonically oscillating airfoils was designed to model a near tip section from a rotor which was known to have experienced supersonic torsional flutter. This five bladed cascade had a solidity of 1.17 and a setting angle of 1.07 rad. Graphite epoxy airfoils were fabricated to achieve the realistically high reduced frequency level of 0.44. The cascade was tested over a range of static pressure ratios approximating the blade element operating conditions of the rotor along a constant speed line which penetrated the flutter boundary. The time-steady and time-unsteady flow field surrounding the center cascade airfoil were investigated. The effects of reduced solidity and decreased setting angle on the flow field were also evaluated.
NASA Technical Reports Server (NTRS)
Whitcomb, R. T.
1983-01-01
This lecture consists of three parts, in which discussions are presented of the current state of development of transonic or supercritical airfoils designed for fully turbulent boundary layers on the surfaces, previous research on subcritical airfoils designed to achieve laminar boundary layers on all or parts of the surfaces, and current research on supercritical airfoils designed to achieve laminar boundary layers. In the first part the use of available two dimensional computer codes in the development of supercritical airfoils and the general trends in the design of such airfoils with turbulent boundary layers are discussed. The second part provides the necessary background on laminar boundary layer phenomena. The last part, which constitutes the major portion of the lecture, covers research by NASA on supercritical airfoils utilizing both decreasing pressure gradients and surface suction for stabilizing the laminar boundary layer. An investigation of the former has been recently conducted in fight using gloves on the wing panels of the U.S. Air Force F111 TACT airplane, research on the later is currently being conducted in a transonic wind tunnel which has been modified to greatly reduce the stream turbulence and noise levels in the tests section.
NASA Technical Reports Server (NTRS)
Shyam, Vikram; Ameri, Ali; Luk, Daniel F.; Chen, Jen-Ping
2010-01-01
Unsteady three-dimensional RANS simulations have been performed on a highly loaded transonic turbine stage and results are compared to steady calculations as well as experiment. A low Reynolds number k- turbulence model is employed to provide closure for the RANS system. A phase-lag boundary condition is used in the periodic direction. This allows the unsteady simulation to be performed by using only one blade from each of the two rows. The objective of this paper is to study the effect of unsteadiness on rotor heat transfer and to glean any insight into unsteady flow physics. The role of the stator wake passing on the pressure distribution at the leading edge is also studied. The simulated heat transfer and pressure results agreed favorably with experiment. The time-averaged heat transfer predicted by the unsteady simulation is higher than the heat transfer predicted by the steady simulation everywhere except at the leading edge. The shock structure formed due to stator-rotor interaction was analyzed. Heat transfer and pressure at the hub and casing were also studied. Thermal segregation was observed that leads to the heat transfer patterns predicted by steady and unsteady simulations to be different.
NASA Technical Reports Server (NTRS)
Ladson, Charles L.; Hill, Acquilla S.; Johnson, William G., Jr.
1987-01-01
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.
Generalized Theodorsen solution for singular integral equations of the airfoil class
NASA Technical Reports Server (NTRS)
Williams, M. H.
1977-01-01
A class of singular integral equations is considered which arise in various two-dimensional mixed boundary-value problems with simple harmonic time variation. A problem typical of this class is that of determining the lifting pressure distribution on an oscillating airfoil in an unbounded incompressible potential flow. It is shown that Theodorsen's (1935) solution to this problem, with some modification, is valid for a general class of unsteady kernel functions. The technique employed is to consider an equivalent steady problem and then show that the unsteady resolvent and unsteady homogeneous solution can be written directly in terms of the steady solutions and a single frequency-dependent function which reduces to the Theodorsen function for the steady kernel.
Udegbunam, E.O.
1991-01-01
This paper presents a FORTRAN program for the determination of two-phase relative permeabilities from unsteady-state displacement data with capillary pressure terms included. The interpretative model employed in this program combines the simultaneous solution of a variant of the fractional flow equation which includes a capillary pressure term and an integro-differential equation derived from Darcy's law without assuming the simplified Buckley-Leverett flow. The incorporation of capillary pressure in the governing equations dispenses with the high flowrate experimental requirements normally employed to overcome capillarity effects. An illustrative example is presented herein which implements this program for the determination of oil/water relative permeabilities from a sandstone core sample. Results obtained compares favorably with results previously given in the literature. ?? 1991.
Computation of airfoil buffet boundaries
NASA Technical Reports Server (NTRS)
Levy, L. L., Jr.; Bailey, H. E.
1981-01-01
The ILLIAC IV computer has been programmed with an implicit, finite-difference code for solving the thin layer compressible Navier-Stokes equation. Results presented for the case of the buffet boundaries of a conventional and a supercritical airfoil section at high Reynolds numbers are found to be in agreement with experimentally determined buffet boundaries, especially at the higher freestream Mach numbers and lower lift coefficients where the onset of unsteady flows is associated with shock wave-induced boundary layer separation.
NASA Technical Reports Server (NTRS)
Oeztuerk, B; Schobeiri, M. T.; Ashpis, David E.
2005-01-01
The paper experimentally and theoretically studies the effects of periodic unsteady wake flow and aerodynamic characteristics on boundary layer development, separation and re-attachment along the suction surface of a low pressure turbine blade. The experiments were carried out at Reynolds number of 110,000 (based on suction surface length and exit velocity). For one steady and two different unsteady inlet flow conditions with the corresponding passing frequencies, intermittency behaviors were experimentally and theoretically investigated. The current investigation attempts to extend the intermittency unsteady boundary layer transition model developed in previously to the LPT cases, where separation occurs on the suction surface at a low Reynolds number. The results of the unsteady boundary layer measurements and the intermittency analysis were presented in the ensemble-averaged and contour plot forms. The analysis of the boundary layer experimental data with the flow separation, confirms the universal character of the relative intermittency function which is described by a Gausssian function.
Anderson, E J; DeMont, M E
2000-09-01
High-speed, high-resolution digital video recordings of swimming squid (Loligo pealei) were acquired. These recordings were used to determine very accurate swimming kinematics, body deformations and mantle cavity volume. The time-varying squid profile was digitized automatically from the acquired swimming sequences. Mantle cavity volume flow rates were determined under the assumption of axisymmetry and the condition of incompressibility. The data were then used to calculate jet velocity, jet thrust and intramantle pressure, including unsteady effects. Because of the accurate measurements of volume flow rate, the standard use of estimated discharge coefficients was avoided. Equations for jet and whole-cycle propulsive efficiency were developed, including a general equation incorporating unsteady effects. Squid were observed to eject up to 94 % of their intramantle working fluid at relatively high swimming speeds. As a result, the standard use of the so-called large-reservoir approximation in the determination of intramantle pressure by the Bernoulli equation leads to significant errors in calculating intramantle pressure from jet velocity and vice versa. The failure of this approximation in squid locomotion also implies that pressure variation throughout the mantle cannot be ignored. In addition, the unsteady terms of the Bernoulli equation and the momentum equation proved to be significant to the determination of intramantle pressure and jet thrust. Equations of propulsive efficiency derived for squid did not resemble Froude efficiency. Instead, they resembled the equation of rocket motor propulsive efficiency. The Froude equation was found to underestimate the propulsive efficiency of the jet period of the squid locomotory cycle and to overestimate whole-cycle propulsive efficiency when compared with efficiencies calculated from equations derived with the squid locomotory apparatus in mind. The equations for squid propulsive efficiency reveal that the refill
Dynamic stall experiments on the NACA 0012 airfoil
NASA Technical Reports Server (NTRS)
Mcalister, K. W.; Carr, L. W.; Mccroskey, W. J.
1978-01-01
The flow over a NACA 0012 airfoil undergoing large oscillations in pitch was experimentally studied at a Reynolds number of and over a range of frequencies and amplitudes. Hot-wire probes and surface-pressure transducers were used to clarify the role of the laminar separation bubble, to delineate the growth and shedding of the stall vortex, and to quantify the resultant aerodynamic loads. In addition to the pressure distributions and normal force and pitching moment data that have often been obtained in previous investigations, estimates of the unsteady drag force during dynamic stall have been derived from the surface pressure measurements. Special characteristics of the pressure response, which are symptomatic of the occurrence and relative severity of moment stall, have also been examined.
NASA Astrophysics Data System (ADS)
Iijima, Yoshimi; Sakaue, Hirotaka
2012-11-01
Electro-luminescence based pressure-sensitive paint (EL-PSP) system uses an EL as an illumination source for a PSP measurement. EL can be directly applied onto a PSP model to eliminate a remote illumination. This gives a uniform illumination on a PSP model without moving/re-directing the illumination. The temperature dependency can be reduced by the opposite temperature dependency of the EL and PSP. At present, the system is demonstrated in a steady flow field. To extend the system for capturing an unsteady flow field, a fast responding PSP and the lifetime characterization of the system are required. The former can be achieved by using a porous PSP. The latter is discussed in the present presentation. The EL-PSP system needs an AC input to illuminate the EL, which gives a pulsed/periodic excitation to a PSP. This limits the acquisition timing of the flow field; a frequent timing can resolve a fast unsteady flow field. The lifetime of the PSP emission can be related to the pressure. The lifetime decays of the EL and PSP are measured to discuss the lifetime characterization of the system.
NASA Technical Reports Server (NTRS)
Mccain, W. E.
1982-01-01
The results of a comparative study using the unsteady aerodynamic lifting surface theory, known as the Doublet Lattice method, and experimental subsonic steady- and unsteady-pressure measurements, are presented for a high-aspect-ratio supercritical wing model. Comparisons of pressure distributions due to wing angle of attack and control-surface deflections were made. In general, good correlation existed between experimental and theoretical data over most of the wing planform. The more significant deviations found between experimental and theoretical data were in the vicinity of control surfaces for both static and oscillatory control-surface deflections.
Turbine airfoil to shround attachment
Campbell, Christian X; Morrison, Jay A; James, Allister W; Snider, Raymond G; Eshak, Daniel M; Marra, John J; Wessell, Brian J
2014-05-06
A turbine airfoil (31) with an end portion (42) that tapers (44) toward the end (43) of the airfoil. A ridge (46) extends around the end portion. It has proximal (66) and distal (67) sides. A shroud platform (50) is bi-cast onto the end portion around the ridge without bonding. Cooling shrinks the platform into compression (62) on the end portion (42) of the airfoil. Gaps between the airfoil and platform are formed using a fugitive material (56) in the bi-casting stage. These gaps are designed in combination with the taper angle (44) to accommodate differential thermal expansion while maintaining a gas seal along the contact surfaces. The taper angle (44) may vary from lesser on the pressure side (36) to greater on the suction side (38) of the airfoil. A collar portion (52) of the platform provides sufficient contact area for connection stability.
Multiple piece turbine airfoil
Kimmel, Keith D
2010-11-09
A turbine airfoil, such as a rotor blade or a stator vane, for a gas turbine engine, the airfoil formed as a shell and spar construction with a plurality of hook shaped struts each mounted within channels extending in a spanwise direction of the spar and the shell to allow for relative motion between the spar and shell in the airfoil chordwise direction while also fanning a seal between adjacent cooling channels. The struts provide the seal as well as prevent bulging of the shell from the spar due to the cooling air pressure. The hook struts have a hooked shaped end and a rounded shaped end in order to insert the struts into the spar.
NASA Technical Reports Server (NTRS)
Irani, E.; Snyder, M. H.
1988-01-01
An averaging total pressure wake rake used by the Cessna Aircraft Company in flight tests of a modified 210 airplane with a laminar flow wing was calibrated in wind tunnel tests against a five-tube pressure probe. The model generating the wake was a full-scale model of the Cessna airplane wing. Indications of drag trends were the same for both instruments.
NASA Astrophysics Data System (ADS)
Miyamoto, Kensuke; Miyazaki, Takeshi; Sakaue, Hirotaka
2009-11-01
Simultaneous reference- and signal-image acquisition for pressure-sensitive paint (PSP) measurement is presented in this paper. The system to acquire these images consists of a two-color PSP and a high-speed CCD camera. The two-color PSP provides a reference and a signal luminescence separated by their luminescent peak wavelengths of 520 nm and 620 nm, respectively. The reference luminescent image is acquired through a green filtered CCD, while the signal image through a red filtered CCD. This system can cancel a non-uniform illumination for exciting two-color PSP as well as a non-uniform image acquisition of a CCD camera due to the location between the camera and the testing object. This system can be applied to a PSP measurement, which includes translation, vibration, and deformation of the testing object. Our measurement system is thus advantage from a conventional PSP measurement, which requires stationary location among the testing object, camera, and illumination source. Unsteady motion of a plate coated with our two-color PSP is included as a demonstration. The developed system captured the unsteady motion as well as gas impingement on the coated plate with camera frame rate of 300 Hz.
NASA Astrophysics Data System (ADS)
Du, Yang; Tu, Shan; Wang, Hongjuan
Two-way sequential fluid-structure interaction method was used to analyze and discuss the characteristics of unsteady fluid-structure interaction of the complex flow channel of a steam turbine ball type control valve. Research indicates that when the pressure ratio changes as a sine wave, its flow rate occurs a sine wave change, and the maximum flow rate value of 57.46kg•s-1 occurs in the minimum pressure ratio condition. The longitudinal force of the structure domain decreases with the reduction of the pressure ratio, and points to the opposite direction of the flow. The lateral force increases with the decrease of the pressure ratio, and points to the opposite direction of the flow. The maximum value of deformation and force of the structure domain changes consistently with the pressure ratio fluctuation. The maximum value of the structure domain stress is 28.67MPa, which is far less than the yield strength of the structure material, and the maximum deformation value is 3.25um.
Faller, W E; Schreck, S J
1995-01-01
The capability to control unsteady separated flow fields could dramatically enhance aircraft agility. To enable control, however, real-time prediction of these flow fields over a broad parameter range must be realized. The present work describes real-time predictions of three-dimensional unsteady separated flow fields and aerodynamic coefficients using neural networks. Unsteady surface-pressure readings were obtained from an airfoil pitched at a constant rate through the static stall angle. All data sets were comprised of 15 simultaneously acquired pressure records and one pitch angle record. Five such records and the associated pitch angle histories were used to train the neural network using a time-series algorithm. Post-training, the input to the network was the pitch angle (alpha), the angular velocity (dalpha/dt), and the initial 15 recorded surface pressures at time (t (0)). Subsequently, the time (t+Deltat) network predictions, for each of the surface pressures, were fed back as the input to the network throughout the pitch history. The results indicated that the neural network accurately predicted the unsteady separated flow fields as well as the aerodynamic coefficients to within 5% of the experimental data. Consistent results were obtained both for the training set as well as for generalization to both other constant pitch rates and to sinusoidal pitch motions. The results clearly indicated that the neural-network model could predict the unsteady surface-pressure distributions and aerodynamic coefficients based solely on angle of attack information. The capability for real-time prediction of both unsteady separated flow fields and aerodynamic coefficients across a wide range of parameters in turn provides a critical step towards the development of control systems targeted at exploiting unsteady aerodynamics for aircraft manoeuvrability enhancement.
NASA Technical Reports Server (NTRS)
Bartels, Robert E.
1999-01-01
This paper presents a modification of the spring analogy scheme which uses axial linear spring stiffness with selective spring stiffening/relaxation. An alternate approach to solving the geometric conservation law is taken which eliminates the need for storage of metric Jacobians at previous time steps. Efficiency and verification are illustrated with several unsteady 2-D airfoil Euler computations. The method is next applied to the computation of the turbulent flow about a 2-D airfoil and wing with two and three- dimensional moving spoiler surfaces, and the results compared with Benchmark Active Controls Technology (BACT) experimental data. The aeroelastic response at low dynamic pressure of an airfoil to a single large scale oscillation of a spoiler surface is computed. This study confirms that it is possible to achieve accurate solutions with a very large time step for aeroelastic problems using the fluid solver and aeroelastic integrator as discussed in this paper.
Navier-Stokes simulations of WECS airfoil flowfields
Homicz, G.F.
1994-06-01
Sandia National Laboratories has initiated an effort to apply Computational Fluid Dynamics (CFD) to the study of WECS aerodynamics. Preliminary calculations are presented for the flow past a SAND 0018/50 airfoil. The flow solver used is F3D, an implicitly, finite-difference code which solves the Thin-Layer Navier-airfoil. The flow solver used is F3D, an implicit, finite-difference code which solves the Thin-Layer Navier-Stokes equations. 2D steady-state calculations are presented at various angles of attack, {alpha}. Sectional lift and drag coefficient, as well as surface pressure distributions, are compared with wind tunnel data, and exhibit reasonable agreement at low to moderate angles of attack. At high {alpha}, where the airfoil is stalled, a converged solution to the steady-state equations could not be obtained. The flowfield continued to change with successive iterations, which is consistent with the fact that the actual flow is inherently transient, and requires the solution of the full unsteady form of the equations.
NASA Astrophysics Data System (ADS)
Volkov, K. N.
2013-03-01
A turbulent flow of a suspension of solid particles in a gas is considered. The suspension is located in a channel with permeable walls (the pressure at the left end face of the channel follows a sinusoidal law). The flow considered here reflects the principal features of the flow in the combustion chamber of a solid-propellant rocket motor. The unsteady flow of the gas suspension is described by using the Eulerian-Lagrangian approach. A stochastic variant of the discrete-trajectory approach is used for modeling the particle motion. The influence of the condensed phase on the turbulence characteristics and acoustic oscillations of the parameters of the working medium in the channel in the case of injection is discussed. The calculated results are compared with data obtained in a physical experiment.
An alternating-direction implicit algorithm for unsteady potential flow
NASA Technical Reports Server (NTRS)
Chipman, R.; Jameson, A.
1981-01-01
An implicit finite-difference scheme is presented for the efficient computation of unsteady potential flow about airfoils. The formulation uses density and the velocity potential as dependent variables and is cast in conservation form to assure the theoretically correct determination of shockwave location and speed. To enable boundary conditions to be imposed directly on the airfoil surface, a time-varying sheared-rectilinear coordinate transformation is employed. Calculated time-history solutions on a pulsating airfoil are compared with the results of other unsteady transonic codes, including a previous method of the authors. The present method is demonstrated to be unconditionally stable and to give accurate solutions with sharply resolved shocks.
NASA Astrophysics Data System (ADS)
Gordon, Kathryn; Morris, Scott; Jemcov, Aleksandar; Cameron, Joshua
2013-11-01
The interaction of components in a compressible, internal flow often results in unsteady interactions between the wakes and moving blades. A prime example in which this flow feature is of interest is the interaction between the downstream rotor blades in a transonic axial compressor with the wake vortices shed from the upstream inlet guide vane (IGV). Previous work shows that a double row of counter-rotating vortices convects downstream into the rotor passage as a result of the rotor blade bow shock impinging on the IGV. The rotor-relative time-mean total pressure distribution has a region of high total pressure corresponding to the pathline of the vortices. The present work focuses on the relationship between the magnitude of the time-mean rotor-relative total pressure profile and the axial spacing between the IGV and the rotor. A survey of different axial gap sizes is performed in a two-dimensional computational study to obtain the sensitivity of the pressure profile amplitude to IGV-rotor axial spacing.
NASA Technical Reports Server (NTRS)
Dougherty, N. S.; Burnette, D. W.; Holt, J. B.; Matienzo, Jose
1993-01-01
Time-accurate unsteady flow simulations are being performed supporting the SRM T+68sec pressure 'spike' anomaly investigation. The anomaly occurred in the RH SRM during the STS-54 flight (STS-54B) but not in the LH SRM (STS-54A) causing a momentary thrust mismatch approaching the allowable limit at that time into the flight. Full-motor internal flow simulations using the USA-2D axisymmetric code are in progress for the nominal propellant burn-back geometry and flow conditions at T+68-sec--Pc = 630 psi, gamma = 1.1381, T(sub c) = 6200 R, perfect gas without aluminum particulate. In a cooperative effort with other investigation team members, CFD-derived pressure loading on the NBR and castable inhibitors was used iteratively to obtain nominal deformed geometry of each inhibitor, and the deformed (bent back) inhibitor geometry was entered into this model. Deformed geometry was computed using structural finite-element models. A solution for the unsteady flow has been obtained for the nominal flow conditions (existing prior to the occurrence of the anomaly) showing sustained standing pressure oscillations at nominally 14.5 Hz in the motor IL acoustic mode that flight and static test data confirm to be normally present at this time. Average mass flow discharged from the nozzle was confirmed to be the nominal expected (9550 lbm/sec). The local inlet boundary condition is being perturbed at the location of the presumed reconstructed anomaly as identified by interior ballistics performance specialist team members. A time variation in local mass flow is used to simulate sudden increase in burning area due to localized propellant grain cracks. The solution will proceed to develop a pressure rise (proportional to total mass flow rate change squared). The volume-filling time constant (equivalent to 0.5 Hz) comes into play in shaping the rise rate of the developing pressure 'spike' as it propagates at the speed of sound in both directions to the motor head end and nozzle. The
Numerical investigation of acoustic radiation from vortex-airfoil interaction
NASA Astrophysics Data System (ADS)
Legault, Anne; Ji, Minsuk; Wang, Meng
2012-11-01
Numerical simulations of vortices interacting with a NACA 0012 airfoil and a flat-plate airfoil at zero angle of attack are carried out to assess the applicability and accuracy of classical theories. Unsteady lift and sound are computed and compared with the predictions by theories of Sears and Amiet, which assume a thin-plate airfoil in an inviscid flow. A Navier-Stokes solver is used in the simulations, and therefore viscous effects are taken into consideration. For the thin-plate airfoil, the effect of viscosity is negligible. For a NACA 0012 airfoil, the viscous contribution to the unsteady lift and sound mainly comes from coherent vortex shedding in the wake of the airfoil and the interaction of the incoming vortices with the airfoil wake, which become stronger at higher Reynolds numbers for a 2-D laminar flow. When the flow is turbulent at chord Reynolds number of 4 . 8 ×105 , however, the viscous contribution becomes negligible as coherent vortex shedding is not present. Sound radiation from vortex-airfoil interaction at turbulent Reynolds numbers is computed numerically via Lighthill's theory and the result is compared with the predictions of Amiet and Curle. The effect of the airfoil thickness is also examined. Supported by ONR Grant N00014-09-1-1088.
Travel of the center of pressure of airfoils transversely to the air stream
NASA Technical Reports Server (NTRS)
Katzmayr, Richard
1929-01-01
The experiments here described were performed for the purpose of obtaining the essential facts concerning the distribution of the air force along the span. We did not follow, however, the time-consuming method of point-to-point measurements of the pressure distribution on the wing surfaces, but determined directly the moment of mean force about an axis passing through the middle of the span parallel to the direction of flight.
Reversed cowl flap inlet thrust augmentor. [with adjustable airfoil
NASA Technical Reports Server (NTRS)
Cheng, D. Y. (Inventor)
1975-01-01
An adjustable airfoil is described for varying the geometry of a jet inlet and an ejector inlet in a jet engine for providing thrust augmentation and noise reduction. The airfoil comprises essentially a plurality of segments which are extended radially outward and retracted relative to the longitudinal axis of the engine as a function of a change in the pressure differential between the upstream and downstream surfaces of the airfoil. A servo mechanism responsive to the change in the pressure differential is coupled to the airfoil to extend and retract the airfoil segments to maintain the pressure at a maximum on the downstream side of the airfoil relative to the pressure on the upstream side of the airfoil. At low speeds, such as at take-offs and landings, the airfoil is fully extended while at high speeds it is fully retracted.
Pressure Distribution Over a Rectangular Airfoil with a Partial-Span Split Flap
NASA Technical Reports Server (NTRS)
Wenzinger, Carl J; Harris, Thomas A
1937-01-01
This report presents the results of pressure-distribution tests of a Clark y wing model with a partial-span split flap made to determine the distribution of air loads over both the wing and the flap. The model was used in conjunction with a reflection plane in the NACA 7 by 10 foot wind tunnel. The 20-percent-chord split flap extended over the inboard 60 percent of the semispan. The tests were made at various flap deflections up to 45 degrees and covered a range of angles of attack from zero lift to approximately maximum lift for each deflection.
Theory and Experiment of Multielement Airfoils: A Comparison
NASA Technical Reports Server (NTRS)
Czerwiec, Ryan; Edwards, J. R.; Rumsey, C. L.; Hassan, H. A.
2000-01-01
A detailed comparison of computed and measured pressure distributions, velocity profiles, transition onset, and Reynolds shear stresses for multi-element airfoils is presented. It is shown that the transitional k-zeta model, which is implemented into CFL3D, does a good job of predicting pressure distributions, transition onset, and velocity profiles with the exception of velocities in the slat wake region. Considering the fact that the hot wire used was not fine enough to resolve Reynolds stresses in the boundary layer, comparisons of turbulence stresses varied from good to fair. It is suggested that the effects of unsteadiness be thoroughly evaluated before more complicated transition/turbulence models are used. Further, it is concluded that the present work presents a viable and economical method for calculating laminar/transitional/turbuient flows over complex shapes without user interface.
NASA Technical Reports Server (NTRS)
Sandford, M. C.; Ricketts, R. H.; Cazier, F. W., Jr.
1980-01-01
A supercritical wing with an aspect ratio of 10.76 and with two trailing-edge oscillating control surfaces is described. The semispan wing is instrumented with 252 static orifices and 164 in situ dynamic-pressure gages for studying the effects of control-surface position and motion on steady- and unsteady-pressures at transonic speeds. Results from initial tests conducted in the Langley Transonic Dynamics Tunnel at two Reynolds numbers are presented in tabular form.
NASA Technical Reports Server (NTRS)
Schobeiri, M. T.; Ozturk, B.; Ashpis, David E.
2007-01-01
The paper experimentally studies the effects of periodic unsteady wake flow and different Reynolds numbers on boundary layer development, separation and re-attachment along the suction surface of a low pressure turbine blade. The experimental investigations were performed on a large scale, subsonic unsteady turbine cascade research facility at Turbomachinery Performance and Flow Research Laboratory (TPFL) of Texas A&M University. The experiments were carried out at Reynolds numbers of 110,000 and 150,000 (based on suction surface length and exit velocity). One steady and two different unsteady inlet flow conditions with the corresponding passing frequencies, wake velocities, and turbulence intensities were investigated. The reduced frequencies chosen cover the operating range of LP turbines. In addition to the unsteady boundary layer measurements, surface pressure measurements were performed. The inception, onset, and the extent of the separation bubble information collected from the pressure measurements were compared with the hot wire measurements. The results presented in ensemble-averaged, and the contour plot forms help to understand the physics of the separation phenomenon under periodic unsteady wake flow and different Reynolds number. It was found that the suction surface displayed a strong separation bubble for these three different reduced frequencies. For each condition, the locations defining the separation bubble were determined carefully analyzing and examining the pressure and mean velocity profile data. The location of the boundary layer separation was dependent of the Reynolds number. It is observed that starting point of the separation bubble and the re-attachment point move further downstream by increasing Reynolds number from 110,000 to 150,000. Also, the size of the separation bubble is smaller when compared to that for Re=110,000.
URANS simulations of separated flow with stall cells over an NREL S826 airfoil
NASA Astrophysics Data System (ADS)
Sarlak, H.; Nishino, T.; Sørensen, J. N.
2016-06-01
A series of wind tunnel measurements and oil flow visualization was recently carried out at the Technical University of Denmark in order to investigate flow characteristics over a 14% thick NREL S826 airfoil at low Reynolds numbers. This paper aims at presenting numerical simulations of the same airfoil using unsteady Reynolds-averaged Navier-Stokes (URANS) approach. Results of the simulations are demonstrated in terms of mean flow velocity, lift and drag, as well as pressure distribution, and validated against available experimental data. The simulations are carried out with a wide computational domain (with a span-to-chord ratio of 5) and it is illustrated that the URANS approach is capable of predicting 3D spanwise structures, known as stall cells.
Flow control of a wind turbine airfoil under aero-elastic pitching
NASA Astrophysics Data System (ADS)
Nikoueeyan, Pourya
Active flow control has demonstrated promising advantages in addressing problems associated with dynamic stall in large wind turbine blades. The feasibility of employing Gurney flaps for active flow control of a pitching flatback airfoil has been assessed in this study. The effectiveness of the Gurney flap in different steady and unsteady inflow conditions from attached-flow to deep dynamic stall condition were characterized by acquiring surface pressure measurements and Particle Image Velocimetry data. The results indicate that the Gurney flap has significant authority in changing the aerodynamics that can affect the aeroelastic behavior of the pitching system used in this work. The fast transient aerodynamics of the airfoil during Gurney flap deployment is also considered as a necessary requirement for an effective closed-loop control scheme. The results show great potential in using the Gurney flap as an effective and fast responding flow control device that can regulate blade loading in a broad range of inflow conditions.
NASA Astrophysics Data System (ADS)
Tanasa, Constantin; Susan-Resiga, Romeo; Muntean, Sebastian; Stuparu, Adrian; Bosioc, Alin; Ciocan, Tiberiu
2015-12-01
Decelerated swirling flows in conical diffusers can develop self-induced instabilities for some configurations of incoming flow. As a result, an upstream steady and axis-symmetrical flow becomes unsteady and three-dimensional with associated pressure fluctuations that hinder the operation of tubomachines. This paper introduces a new approach for mitigating the swirling flow instabilities using a diaphragm into the cone. As a result, the severe flow deceleration is diminished, taming the precessing helical vortex and its associated pressure fluctuations. A full three-dimensional unsteady numerical study is performed for swirling flow apparatus developed at the Politehnica University of Timisoara. Fourier spectra corresponding to unsteady numerical data acquired without and with diaphragm are compared in several locations displaced along to the element of the cone in order to assess the novel concept. It is clearly revealed that the diaphragm mitigates the unsteadiness effects associated to the swirling flow with precessing helical vortex in the conical diffuser. The novel concept paves the way towards a new passive control technique.
NASA Technical Reports Server (NTRS)
Gross, A. R.; Steinle, F. W., Jr.
1975-01-01
A NACA 64A010 pressure-instrumented airfoil was tested at transonic speeds over a range of angle of attack from -1 to 12 degrees at various Reynolds numbers ranging from 2 to 6 million in air, argon, Freon 12, and a mixture of argon and Freon 12 having a ratio of specific heats corresponding to air. Good agreement of results is obtained for conditions where compressibility is not significant and for the air and comparable argon-Freon 12 mixture. Comparison of heavy gas results with air, when adjusted for transonic similarity, show improved, but less than desired agreement.
NASA Technical Reports Server (NTRS)
Lindsey, W F
1946-01-01
An investigation has been conducted in the Langley rectangular high-speed tunnel to determine the effect of compressibility on the pressure distribution for a modified NACA 65,3-019 airfoil having a 0.20-chord flap. The investigation was made for an angle-of-attack range extending from -2 to 12 deg at .20 flap deflections from 0 to -12 deg. Test data were obtained for Mach numbers from 0.28 to approximately 0.74. The results show that the effectiveness of the trailing-edge-type control surface rapidly decreased and approached zero as the Mach number increased above the critical value.
Separated-flow unsteady pressures and forces on elastically responding structures
NASA Technical Reports Server (NTRS)
Coke, C. F.; Riddle, D. W.; Hwang, C.
1977-01-01
Broadband rms, spectral density, and spatial correlation information that characterizes the fluctuating pressures and forces that cause aircraft buffet is presented. The main theme is to show the effects of elasticity. In order to do so, data are presented that were obtained in regions of separated flow on wings of wind-tunnel models of varying stiffness and on the wing of a full-scale aircraft. Reynolds number effects on the pressure fluctuations are also discussed.
NASA Technical Reports Server (NTRS)
Coe, C. F.; Riddle, D. W.; Hwang, C.
1977-01-01
Broadband rms, spectral density, and spatial correlation information that characterizes the fluctuating pressures and forces that cause aircraft buffet is presented. The main theme of the paper in describing buffet excitation is to show the effects of elasticity. Data are presented that were obtained in regions of separated flow on wings of wind-tunnel models of varying stiffness and on the wing of a full scale aircraft. Reynolds number effects on the pressure fluctuations are also discussed.
NASA Technical Reports Server (NTRS)
Mccain, W. E.
1984-01-01
The unsteady aerodynamic lifting surface theory, the Doublet Lattice method, with experimental steady and unsteady pressure measurements of a high aspect ratio supercritical wing model at a Mach number of 0.78 were compared. The steady pressure data comparisons were made for incremental changes in angle of attack and control surface deflection. The unsteady pressure data comparisons were made at set angle of attack positions with oscillating control surface deflections. Significant viscous and transonic effects in the experimental aerodynamics which cannot be predicted by the Doublet Lattice method are shown. This study should assist development of empirical correction methods that may be applied to improve Doublet Lattice calculations of lifting surface aerodynamics.
Parameter study of simplified dragonfly airfoil geometry at Reynolds number of 6000.
Levy, David-Elie; Seifert, Avraham
2010-10-21
Aerodynamic study of a simplified Dragonfly airfoil in gliding flight at Reynolds numbers below 10,000 is motivated by both pure scientific interest and technological applications. At these Reynolds numbers, the natural insect flight could provide inspiration for technology development of Micro UAV's and more. Insect wings are typically characterized by corrugated airfoils. The present study follows a fundamental flow physics study (Levy and Seifert, 2009), that revealed the importance of flow separation from the first corrugation, the roll-up of the separated shear layer to discrete vortices and their role in promoting flow reattachment to the aft arc, as the leading mechanism enabling high-lift, low drag performance of the Dragonfly gliding flight. This paper describes the effect of systematic airfoil geometry variations on the aerodynamic properties of a simplified Dragonfly airfoil at Reynolds number of 6000. The parameter study includes a detailed analysis of small variations of the nominal geometry, such as corrugation placement or height, rear arc and trailing edge shape. Numerical simulations using the 2D laminar Navier-Stokes equations revealed that the flow accelerating over the first corrugation slope is followed by an unsteady pressure recovery, combined with vortex shedding. The latter allows the reattachment of the flow over the rear arc. Also, the drag values are directly linked to the vortices' magnitude. This parametric study shows that geometric variations which reduce the vortices' amplitude, as reduction of the rear cavity depth or the reduction of the rear arc and trailing edge curvature, will reduce the drag values. Other changes will extend the flow reattachment over the rear arc for a larger mean lift coefficients range; such as the negative deflection of the forward flat plate. These changes consequently reduce the drag values at higher mean lift coefficients. The detailed geometry study enabled the definition of a corrugated airfoil
Parameter study of simplified dragonfly airfoil geometry at Reynolds number of 6000.
Levy, David-Elie; Seifert, Avraham
2010-10-21
Aerodynamic study of a simplified Dragonfly airfoil in gliding flight at Reynolds numbers below 10,000 is motivated by both pure scientific interest and technological applications. At these Reynolds numbers, the natural insect flight could provide inspiration for technology development of Micro UAV's and more. Insect wings are typically characterized by corrugated airfoils. The present study follows a fundamental flow physics study (Levy and Seifert, 2009), that revealed the importance of flow separation from the first corrugation, the roll-up of the separated shear layer to discrete vortices and their role in promoting flow reattachment to the aft arc, as the leading mechanism enabling high-lift, low drag performance of the Dragonfly gliding flight. This paper describes the effect of systematic airfoil geometry variations on the aerodynamic properties of a simplified Dragonfly airfoil at Reynolds number of 6000. The parameter study includes a detailed analysis of small variations of the nominal geometry, such as corrugation placement or height, rear arc and trailing edge shape. Numerical simulations using the 2D laminar Navier-Stokes equations revealed that the flow accelerating over the first corrugation slope is followed by an unsteady pressure recovery, combined with vortex shedding. The latter allows the reattachment of the flow over the rear arc. Also, the drag values are directly linked to the vortices' magnitude. This parametric study shows that geometric variations which reduce the vortices' amplitude, as reduction of the rear cavity depth or the reduction of the rear arc and trailing edge curvature, will reduce the drag values. Other changes will extend the flow reattachment over the rear arc for a larger mean lift coefficients range; such as the negative deflection of the forward flat plate. These changes consequently reduce the drag values at higher mean lift coefficients. The detailed geometry study enabled the definition of a corrugated airfoil
Virtual Shaping of a Two-dimensional NACA 0015 Airfoil Using Synthetic Jet Actuator
NASA Technical Reports Server (NTRS)
Chen, Fang-Jenq; Beeler, George B.
2002-01-01
The Aircraft Morphing Program at NASA Langley envisions an aircraft without conventional control surfaces. Instead of moving control surfaces, the vehicle control systems may be implemented with a combination of propulsive forces, micro surface effectors, and fluidic devices dynamically operated by an intelligent flight control system to provide aircraft maneuverability over each mission segment. As a part of this program, a two-dimensional NACA 0015 airfoil model was designed to test mild maneuvering capability of synthetic jets in a subsonic wind tunnel. The objective of the experiments is to assess the applicability of using unsteady suction and blowing to alter the aerodynamic shape of an airfoil with a purpose to enhance lift and/or to reduce drag. Synthetic jet actuation at different chordwise locations, different forcing frequencies and amplitudes, under different freestream velocities are investigated. The effect of virtual shape change is indicated by a localized increase of surface pressure in the neighborhood of synthetic jet actuation. That causes a negative lift to the airfoil with an upper surface actuation. When actuation is applied near the airfoil leading edge, it appears that the stagnation line is shifted inducing an effect similar to that caused by a small angle of attack to produce an overall lift change.
Garcia-Crespo, Andres Jose
2015-03-03
A composite blade assembly for mounting on a turbine wheel includes a ceramic airfoil and an airfoil platform. The ceramic airfoil is formed with an airfoil portion, a blade shank portion and a blade dovetail tang. The metal platform includes a platform shank and a radially inner platform dovetail. The ceramic airfoil is captured within the metal platform, such that in use, the ceramic airfoil is held within the turbine wheel independent of the metal platform.
Tangler, James L.; Somers, Dan M.
1996-01-01
Airfoils for the blade of a wind turbine wherein each airfoil is characterized by a thickness in a range from 16%-24% and a maximum lift coefficient designed to be largely insensitive to roughness effects. The airfoils include a family of airfoils for a blade 15 to 25 meters in length, a family of airfoils for a blade 1 to 5 meters in length, and a family of airfoils for a blade 5 to 10 meters in length.
Tangler, J.L.; Somers, D.M.
1996-10-08
Airfoils are disclosed for the blade of a wind turbine wherein each airfoil is characterized by a thickness in a range from 16%-24% and a maximum lift coefficient designed to be largely insensitive to roughness effects. The airfoils include a family of airfoils for a blade 15 to 25 meters in length, a family of airfoils for a blade 1 to 5 meters in length, and a family of airfoils for a blade 5 to 10 meters in length. 10 figs.
Inverse transonic airfoil design including viscous interaction
NASA Technical Reports Server (NTRS)
Carlson, L. A.
1976-01-01
A numerical technique was developed for the analysis of specified transonic airfoils or for the design of airfoils having a prescribed pressure distribution, including the effect of weak viscous interaction. The method uses the full potential equation, a stretched Cartesian coordinate system, and the Nash-MacDonald turbulent boundary layer method. Comparisons with experimental data for typical transonic airfoils show excellent agreement. An example shows the application of the method to design a thick aft-cambered airfoil, and the effects of viscous interaction on its performance are discussed.
Wavy flow cooling concept for turbine airfoils
Liang, George
2010-08-31
An airfoil including an outer wall and a cooling cavity formed therein. The cooling cavity includes a leading edge flow channel located adjacent a leading edge of the airfoil and a trailing edge flow channel located adjacent a trailing edge of the airfoil. Each of the leading edge and trailing edge flow channels define respective first and second flow axes located between pressure and suction sides of the airfoil. A plurality of rib members are located within each of the flow channels, spaced along the flow axes, and alternately extending from opposing sides of the flow channels to define undulating flow paths through the flow channels.
Clark, Edward L.; Henfling, John F.; McBride, Donald D.
1999-05-12
An inverse Fourier transform method for removing lag from pressure measurements has been used by various researchers, given an experimentally derived transfer function to characterize the pressure plumbing. This paper presents a Method of Characteristics (MOC) solution technique for predicting the transfer function and thus easily determining its sensitivity to various plumbing pammeters. The MOC solution has been used in the pipeline industry for some time for application to transient flow in pipelines, but it also lends itself well to this application. For highly nonsteady pressures frequency-dependent friction can cause significant distortion of the traveling waves. This is accounted for in the formulation. A simple bench experiment and proof-of-principle test provide evidence to establish the range of validity of the method.
Aerodynamic characteristics of airfoils with ice accretions
NASA Technical Reports Server (NTRS)
Bragg, M. B.; Gregorek, G. M.
1982-01-01
Results of a wind tunnel test to evaluate the performance of an airfoil with simulated rime ice are presented with theoretical comparisons. A NACA 65A413 airfoil was tested in the OSU 6 x 22 inch Transonic Airfoil Wind Tunnel at a Reynolds number near three million and Mach numbers from 0.20 to 0.80. The model was tested in four configurations to determine the aero-dynamic effects of the roughness and shape of a rime ice accretion. The simulated rime ice shape was obtained analytically using a time-stepping dry ice accretion computer code. Lift, drag, moment coefficients, and pressure distributions for the clean and simulated rime ice cases are reported. The measured degradation in airfoil performance is compared to an analytical method which uses existing airfoil analysis computer codes with empirical corrections for the surface roughness. A discussion of the empirical surface roughness correction and uses of other airfoil computer methods is included.
Calculation of unsteady fan rotor response caused by downstream flow distortions
NASA Technical Reports Server (NTRS)
Obrien, W. F.; Richardson, S. M.; Ng, W. F.
1984-01-01
The present model for fan rotor/support strut airfoil interaction uses a time-marching code for the rotor flow, coupled with a potential flow model for the stator-strut region. Study of the effect of strut design variables indicates that rotor flow disturbance is increased by the primary variables of larger strut thickness and circumferential spacing, while decreasing exponentially with increased rotor-strut separation. The time-marching code predicts local rotor pressure and flow perturbations in response to an unsteady downstream boundary condition.
Time-resolved stereo PIV measurements of shock-boundary layer interaction on a supercritical airfoil
NASA Astrophysics Data System (ADS)
Hartmann, Axel; Klaas, Michael; Schröder, Wolfgang
2012-03-01
Time-resolved stereo particle-image velocimetry (TR-SPIV) and unsteady pressure measurements are used to analyze the unsteady flow over a supercritical DRA-2303 airfoil in transonic flow. The dynamic shock wave-boundary layer interaction is one of the most essential features of this unsteady flow causing a distinct oscillation of the flow field. Results from wind-tunnel experiments with a variation of the freestream Mach number at Reynolds numbers ranging from 2.55 to 2.79 × 106 are analyzed regarding the origin and nature of the unsteady shock-boundary layer interaction. Therefore, the TR-SPIV results are analyzed for three buffet flows. One flow exhibits a sinusoidal streamwise oscillation of the shock wave only due to an acoustic feedback loop formed by the shock wave and the trailing-edge noise. The other two buffet flows have been intentionally influenced by an artificial acoustic source installed downstream of the test section to investigate the behavior of the interaction to upstream-propagating disturbances generated by a defined source of noise. The results show that such upstream-propagating disturbances could be identified to be responsible for the upstream displacement of the shock wave and that the feedback loop is formed by a pulsating separation of the boundary layer dependent on the shock position and the sound pressure level at the shock position. Thereby, the pulsation of the separation could be determined to be a reaction to the shock motion and not vice versa.
NASA Technical Reports Server (NTRS)
Sandford, M. C.; Ricketts, R. H.
1983-01-01
A high aspect ratio supercritical wing with oscillating control surfaces is described. The semispan wing model was instrumented with 252 static pressure orifices and 164 in situ dynamic pressure gages for studying the effects of control surface position and sinusoidal motion on steady and unsteady pressures. Results from the present test (the third in a series of tests on this model) were obtained in the Langley Transonic Dynamics Tunnel at Mach numbers of 0.60, 0.78, and 0.86 and are presented in tabular form.
Airfoil shape for flight at subsonic speeds
Whitcomb, Richard T.
1976-01-01
An airfoil having an upper surface shaped to control flow accelerations and pressure distribution over the upper surface and to prevent separation of the boundary layer due to shock wave formulation at high subsonic speeds well above the critical Mach number. A highly cambered trailing edge section improves overall airfoil lifting efficiency.
Unsteady transonic flow in cascades
NASA Technical Reports Server (NTRS)
Surampudi, S. P.; Adamczyk, J. J.
1984-01-01
There is a need for methods to predict the unsteady air loads associated with flutter of turbomachinery blading at transonic speeds. The results of such an analysis in which the steady relative flow approaching a cascade of thin airfoils is assumed to be transonic, irrotational, and isentropic is presented. The blades in the cascade are allowed to undergo a small amplitude harmonic oscillation which generates a small unsteady flow superimposed on the existing steady flow. The blades are assumed to oscillate with a prescribed motion of constant amplitude and interblade phase angle. The equations of motion are obtained by linearizing about a uniform flow the inviscid nonheat conducting continuity and momentum equations. The resulting equations are solved by employing the Weiner Hopf technique. The solution yields the unsteady aerodynamic forces acting on the cascade at Mach number equal to 1. Making use of an unsteady transonic similarity law, these results are compared with the results obtained from linear unsteady subsonic and supersonic cascade theories. A parametric study is conducted to find the effects of reduced frequency, solidity, stagger angle, and position of pitching axis on the flutter.
NASA Technical Reports Server (NTRS)
LOVE EUGENE S
1957-01-01
An analysis has been made of available experimental data to show the effects of most of the variables that are more predominant in determining base pressure at supersonic speeds. The analysis covers base pressures for two-dimensional airfoils and for bodies of revolution with and without stabilizing fins and is restricted to turbulent boundary layers. The present status of available experimental information is summarized as are the existing methods for predicting base pressure. A simple semiempirical method is presented for estimating base pressure. For two-dimensional bases, this method stems from an analogy established between the base-pressure phenomena and the peak pressure rise associated with the separation of the boundary layer. An analysis made for axially symmetric flow indicates that the base pressure for bodies of revolution is subject to the same analogy. Based upon the methods presented, estimations are made of such effects as Mach number, angle of attack, boattailing, fineness ratio, and fins. These estimations give fair predictions of experimental results. (author)
Trailing edge flow conditions as a factor in airfoil design
NASA Technical Reports Server (NTRS)
Ormsbee, A. I.; Maughmer, M. D.
1984-01-01
Some new developments relevant to the design of single-element airfoils using potential flow methods are presented. In particular, the role played by the non-dimensional trailing edge velocity in design is considered and the relationship between the specified value and the resulting airfoil geometry is explored. In addition, the ramifications of the unbounded trailing edge pressure gradients generally present in the potential flow solution of the flow over an airfoil are examined, and the conditions necessary to obtain a class of airfoils having finite trailing edge pressure gradients developed. The incorporation of these conditions into the inverse method of Eppler is presented and the modified scheme employed to generate a number of airfoils for consideration. The detailed viscous analysis of airfoils having finite trailing edge pressure gradients demonstrates a reduction in the strong inviscid-viscid interactions generally present near the trailing edge of an airfoil.
NASA Technical Reports Server (NTRS)
St.hilaire, A. O.; Carta, F. O.
1979-01-01
The effect of sweep on the dynamic response of the NACA 0012 airfoil was investigated. Unsteady chordwise distributed pressure data were obtained from a tunnel spanning wing equipped with 21 single surface transducers (13 on the suction side and 8 on the pressure side of the airfoil). The pressure data were obtained at pitching amplitudes of 8 and 10 degrees over a tunnel Mach number range of 0.10 to 0.46 and a pitching frequency range of 2.5 to 10.6 cycles per second. The wing was oscillated in the unswept and swept positions about the quarter-chord pivot axis relative to mean incidence angle settings of 0, 9, 12, and 15 degrees. A compilation of all the response data obtained during the test program is presented. These data are in the form of normal force, chord force, lift force, pressure drag, and moment hysteresis loops derived from chordwise integrations of the unsteady pressure distributions. The hysteresis loops are organized in two main sections. In the first section, the loop data are arranged to show the effect of sweep (lambda = 0 and 30 deg) for all available combinations of mean incidence angle, pitching amplitude, reduced frequency, and chordwise Mach number. The second section shows the effect of chordwise Mach number (MC = 0.30 and MC = 0.40) on the swept wing response for all available combinations of mean incidence angle, pitching amplitude, and reduced frequency.
NASA Technical Reports Server (NTRS)
Schobeiri, M. T.; Radke, R. E.
1996-01-01
Boundary layer transition and development on a turbomachinery blade is subjected to highly periodic unsteady turbulent flow, pressure gradient in longitudinal as well as lateral direction, and surface curvature. To study the effects of periodic unsteady wakes on the concave surface of a turbine blade, a curved plate was utilized. On the concave surface of this plate, detailed experimental investigations were carried out under zero and negative pressure gradient. The measurements were performed in an unsteady flow research facility using a rotating cascade of rods positioned upstream of the curved plate. Boundary layer measurements using a hot-wire probe were analyzed by the ensemble-averaging technique. The results presented in the temporal-spatial domain display the transition and further development of the boundary layer, specifically the ensemble-averaged velocity and turbulence intensity. As the results show, the turbulent patches generated by the wakes have different leading and trailing edge velocities and merge with the boundary layer resulting in a strong deformation and generation of a high turbulence intensity core. After the turbulent patch has totally penetrated into the boundary layer, pronounced becalmed regions were formed behind the turbulent patch and were extended far beyond the point they would occur in the corresponding undisturbed steady boundary layer.
NASA Astrophysics Data System (ADS)
Newman, James C., III
1995-01-01
The limiting factor in simulating flows past realistic configurations of interest has been the discretization of the physical domain on which the governing equations of fluid flow may be solved. In an attempt to circumvent this problem, many Computational Fluid Dynamic (CFD) methodologies that are based on different grid generation and domain decomposition techniques have been developed. However, due to the costs involved and expertise required, very few comparative studies between these methods have been performed. In the present work, the two CFD methodologies which show the most promise for treating complex three-dimensional configurations as well as unsteady moving boundary problems are evaluated. These are namely the structured-overlapped and the unstructured grid schemes. Both methods use a cell centered, finite volume, upwind approach. The structured-overlapped algorithm uses an approximately factored, alternating direction implicit scheme to perform the time integration, whereas, the unstructured algorithm uses an explicit Runge-Kutta method. To examine the accuracy, efficiency, and limitations of each scheme, they are applied to the same steady complex multicomponent configurations and unsteady moving boundary problems. The steady complex cases consist of computing the subsonic flow about a two-dimensional high-lift multielement airfoil and the transonic flow about a three-dimensional wing/pylon/finned store assembly. The unsteady moving boundary problems are a forced pitching oscillation of an airfoil in a transonic freestream and a two-dimensional, subsonic airfoil/store separation sequence. Accuracy was accessed through the comparison of computed and experimentally measured pressure coefficient data on several of the wing/pylon/finned store assembly's components and at numerous angles-of-attack for the pitching airfoil. From this study, it was found that both the structured-overlapped and the unstructured grid schemes yielded flow solutions of
Comparative Study of Airfoil Flow Separation Criteria
NASA Astrophysics Data System (ADS)
Laws, Nick; Kahouli, Waad; Epps, Brenden
2015-11-01
Airfoil flow separation impacts a multitude of applications including turbomachinery, wind turbines, and bio-inspired micro-aerial vehicles. In order to achieve maximum performance, some devices operate near the edge of flow separation, and others use dynamic flow separation advantageously. Numerous criteria exist for predicting the onset of airfoil flow separation. This talk presents a comparative study of a number of such criteria, with emphasis paid to speed and accuracy of the calculations. We evaluate the criteria using a two-dimensional unsteady vortex lattice method, which allows for rapid analysis (on the order of seconds instead of days for a full Navier-Stokes solution) and design of optimal airfoil geometry and kinematics. Furthermore, dynamic analyses permit evaluation of dynamic stall conditions for enhanced lift via leading edge vortex shedding, commonly present in small flapping-wing flyers such as the bumblebee and hummingbird.
NASA Technical Reports Server (NTRS)
Steger, J. L.; Caradonna, F. X.
1980-01-01
An implicit finite difference procedure is developed to solve the unsteady full potential equation in conservation law form. Computational efficiency is maintained by use of approximate factorization techniques. The numerical algorithm is first order in time and second order in space. A circulation model and difference equations are developed for lifting airfoils in unsteady flow; however, thin airfoil body boundary conditions have been used with stretching functions to simplify the development of the numerical algorithm.
A Wind Tunnel Model to Explore Unsteady Circulation Control for General Aviation Applications
NASA Technical Reports Server (NTRS)
Cagle, Christopher M.; Jones, Gregory S.
2002-01-01
Circulation Control airfoils have been demonstrated to provide substantial improvements in lift over conventional airfoils. The General Aviation Circular Control model is an attempt to address some of the concerns of this technique. The primary focus is to substantially reduce the amount of air mass flow by implementing unsteady flow. This paper describes a wind tunnel model that implements unsteady circulation control by pulsing internal pneumatic valves and details some preliminary results from the first test entry.
NASA Technical Reports Server (NTRS)
Bennett, Robert M.; Bland, Samuel R.; Batina, John T.; Gibbons, Michael D.; Mabey, Dennis G.
1987-01-01
A transonic unsteady aerodynamic and aeroelasticity code has been developed for application to realistic aircraft configurations. The new code is called CAP-TSD which is an acronym for Computational Aeroelasticity Program - Transonic Small Disturbance. The CAP-TSD code uses a time-accurate approximate factorization algorithm for solution of the unsteady transonic small-disturbance equation that is efficient for solution of steady and unsteady transonic flow problems including supersonic freestream flows. The new code can treat complete aircraft geometries with multiple lifting surfaces and bodies. Applications to wings in supersonic freestream flow are presented. Comparisons with selected exact solutions from linear theory are presented showing generally favorable results. Calculations for both steady and oscillatory cases for the F-5 and RAE tailplane models are compared with experimental data and also show good overall agreement. Selected steady calculations are further compared with a steady flow Euler code.
Airfoil Tonal Noise Generation in Resonant Environments
NASA Astrophysics Data System (ADS)
Atobe, Takashi; Tuinstra, Marthijn; Takagi, Shohei
To clarify tonal noise generation, an experimental study on airfoil tonal noise was undertaken using a conventional wind tunnel, which allows acoustic reflection on test section walls. A two-dimensional wing model with the NACA0015 cross-section was used at 5 degrees angle of attack. Most previous experiments conducted in anechoic environments commonly show that the tonal noise frequency is selected in an overall trend of U1.5 (U is uniform velocity) locally consisting of a step-like structure, and Tollmien-Schlichting disturbances are rapidly amplified in the backflow region near the trailing edge of the pressure surface. The present experiments in an acoustically resonant environment show that the tonal noise emanates in accordance with the aforementioned features. However, the ladder-like structure has a different local slope from that observed in anechoic flow. These characteristics suggest that acoustic resonance does not play a fundamental role in tonal noise generation. Observation by hot-wire and smoke visualization techniques shows that unsteady disturbances rather than Tollmien-Schlichting waves are rapidly magnified by the Kelvin-Helmholtz instability in the backflow region. The frequency selection mechanism at tonal noise generation still remains unsolved.
NASA Technical Reports Server (NTRS)
Haviland, J. K.; Schroeder, J. C.
1978-01-01
As part of an overall study of the scaling laws for the fluctuating pressures induced on the wings and flaps of STOL aircraft by jet engine impingement, an experimental investigation was made of the near field fluctuating pressures behind a cold circular jet, both when it was free and when it was impinging on a flat plate. Miniature static pressure probes were developed for measurements in the free jet and on the flat plate which were connected by plastic tubing to 1/8 inch microphones and acted as pressure transducers. Using a digital correlator together with an FFT program on the CDC 6400 computer, spectral densities, relative amplitudes, phase lags, and coherences were also obtained for the signals from pairs of these probes, and were used to calibrate these probes directly against microphones. This system of instrumentation was employed to obtain single point rms and third octave surveys of the static pressures in the free jet and on the surface of the plate.
Unsteady separated flows - Novel experimental approach
NASA Technical Reports Server (NTRS)
Krothapalli, A.; Lourenco, L.
1987-01-01
A novel experimental technique, commonly referred to as laser speckle velocimetry or particle image displacement velocimetry (PIDV), is developed for the measurement of instantaneous velocity fields in unsteady and steady flows. The main advantage of this technique is the fact that the velocity field is measured with sufficient accuracy so that the distribution of vorticity can be calculated with accuracy. The PIDV technique, which is ideally suited for the study of unsteady separated flows, has been utilized to measure the development of the separated flow field generated by a high angle-of-attack (alpha = 30 deg) NACA 0012 airfoil, started impulsively from rest.
NASA Technical Reports Server (NTRS)
Parzych, D.; Boyd, L.; Meissner, W.; Wyrostek, A.
1991-01-01
An experiment was performed by Hamilton Standard, Division of United Technologies Corporation, under contract by LeRC, to measure the blade surface pressure of a large scale, 8 blade model prop-fan in flight. The test bed was the Gulfstream 2 Prop-Fan Test Assessment (PTA) aircraft. The objective of the test was to measure the steady and periodic blade surface pressure resulting from three different Prop-Fan air inflow angles at various takeoff and cruise conditions. The inflow angles were obtained by varying the nacelle tilt angles, which ranged from -3 to +2 degrees. A range of power loadings, tip speeds, and altitudes were tested at each nacelle tilt angle over the flight Mach number range of 0.30 to 0.80. Unsteady blade pressure data tabulated as Fourier coefficients for the first 35 harmonics of shaft rotational frequency and the steady (non-varying) pressure component are presented.
Robust, optimal subsonic airfoil shapes
NASA Technical Reports Server (NTRS)
Rai, Man Mohan (Inventor)
2008-01-01
Method system, and product from application of the method, for design of a subsonic airfoil shape, beginning with an arbitrary initial airfoil shape and incorporating one or more constraints on the airfoil geometric parameters and flow characteristics. The resulting design is robust against variations in airfoil dimensions and local airfoil shape introduced in the airfoil manufacturing process. A perturbation procedure provides a class of airfoil shapes, beginning with an initial airfoil shape.
Unsteady transonic aerodynamics
Nixon, D.
1989-01-01
Various papers on unsteady transonic aerodynamics are presented. The topics addressed include: physical phenomena associated with unsteady transonic flows, basic equations for unsteady transonic flow, practical problems concerning aircraft, basic numerical methods, computational methods for unsteady transonic flows, application of transonic flow analysis to helicopter rotor problems, unsteady aerodynamics for turbomachinery aeroelastic applications, alternative methods for modeling unsteady transonic flows.
Unsteady aerodynamic modeling for arbitrary motions. [for active control techniques
NASA Technical Reports Server (NTRS)
Edwards, J. W.
1977-01-01
Results indicating that unsteady aerodynamic loads derived under the assumption of simple harmonic motions executed by airfoil or wing can be extended to arbitrary motions are summarized. The generalized Theodorsen (1953) function referable to loads due to simple harmonic oscillations of a wing section in incompressible flow, the Laplace inversion integral for unsteady aerodynamic loads, calculations of root loci of aeroelastic loads, and analysis of generalized compressible transient airloads are discussed.
An unsteady Euler scheme for the analysis of ducted propellers
NASA Technical Reports Server (NTRS)
Srivastava, R.
1992-01-01
An efficient unsteady solution procedure has been developed for analyzing inviscid unsteady flow past ducted propeller configurations. This scheme is first order accurate in time and second order accurate in space. The solution procedure has been applied to a ducted propeller consisting of an 8-bladed SR7 propeller with a duct of NACA 0003 airfoil cross section around it, operating in a steady axisymmetric flowfield. The variation of elemental blade loading with radius, compares well with other published numerical results.
NASA Technical Reports Server (NTRS)
Yoshikawa, K. K.
1982-01-01
The time-dependent, compressible, Reynolds-averaged, full Navier-Stokes equations are applied to solve an axisymmetric flow around a forward-facing stepbody (spikebody) at supersonic speeds and a stalling airfoil at transonic speeds. Important transient and unsteady phenomena, not yet well understood, are examined, and significant new findings of the present solution to the phenomena are discussed. The phenomena described in detail are as follows: The evolution of the shock wave pressure built up by the impact of the pressure waves, one from the trailing edge; the separation of the flow as influenced by the shock wave; the location of the reversed flow, the separation point, and the reattachment point; and the transient (or unsteady) phenomena of the flow pulsation, oscillation, and stalling of the body and airfoil wake flow. The numerical results show that the transient flow instability is caused by a supersonic jet induced in the separation bubble by the shock-bifurcation (lambda shock) mechanism between the separation shock and the reflected shock. Pulsation and stall phenomena are caused by a sudden increase in the leading-edge pressure due to the jet and the separation bubble interacting along the stagnation point flow.
An alternating-direction-implicit algorithm for the unsteady potential equation in conservation form
NASA Technical Reports Server (NTRS)
Chipman, R. R.
1980-01-01
An implicit finite difference scheme for an efficient computation of unsteady potential flow about airfoils is presented. The formulation uses density and velocity potential as dependent variables, and is cast in conservation form to assure the theoretically correct determination of shockwave location and speed. To enable boundary conditions to be imposed directly on the airfoil surface, a time varying sheared rectilinear coordinate transformation is employed. Calculated time history solutions on a pulsating airfoil are compared with the results of another unsteady transonic code. It is concluded that the method has excellent numerical stability and gives accurate solutions with sharply resolved shocks.
Multiple Solutions of Transonic Flow over NACA0012 Airfoil
NASA Astrophysics Data System (ADS)
Xiong, Juntao; Liu, Ya; Liu, Feng; Luo, Shijun; Zhao, Zijie; Ren, Xudong; Gao, Chao
2012-11-01
Multiple solutions of the small-disturbance potential equation and full potential equation were known for the NACA0012 airfoil in a certain range of transonic Mach numbers and at zero angle of attack. However the multiple solutions for this airfoil were not observed using Euler or Navier-Stokes equations under the above flow conditions. In the present work, both the Unsteady Reynolds-Averaged Navier-Stokes (URANS) computations and transonic wind tunnel experiments are performed under certain Reynolds numbers to further study the problem. The results of the two methods reveal that buffet appears in a narrow Mach number range where the potential flow methods predict multiple solutions. Boundary layer displacement thickness computed from URANS at the same flow condition is used to modify the geometry of the airfoil. Euler equations are then solved for the modified geometry. The results show that the addition of the boundary layer displacement thickness creates multiple solutions for the NACA0012 airfoil. Global linear stability analysis is also performed on the original and the modified airfoils. This shows a close relationship between the viscous unsteady shock buffet phenomenon of transonic airfoil flow and the existence of multiple solutions of the external inviscid flow. Postdoctoral Research Assistant.
Steady inviscid transonic flows over planar airfoils: A search for a simplified procedure
NASA Technical Reports Server (NTRS)
Magnus, R.; Yoshihara, H.
1973-01-01
A finite difference procedure based upon a system of unsteady equations in proper conservation form with either exact or small disturbance steady terms is used to calculate the steady flows over several classes of airfoils. The airfoil condition is fulfilled on a slab whose upstream extremity is a semi-circle overlaying the airfoil leading edge circle. The limitations of the small disturbance equations are demonstrated in an extreme example of a blunt-nosed, aft-cambered airfoil. The necessity of using the equations in proper conservation form to capture the shock properly is stressed. Ability of the steady relaxation procedures to capture the shock is briefly examined.
NASA Technical Reports Server (NTRS)
Schobeiri, M. T.; John, J.
1996-01-01
The turbomachinery wake flow development is largely influenced by streamline curvature and streamwise pressure gradient. The objective of this investigation is to study the development of the wake under the influence of streamline curvature and streamwise pressure gradient. The experimental investigation is carried out in two phases. The first phase involves the study of the wake behind a stationary circular cylinder (steady wake) in curved channels at positive, zero, and negative streamwise pressure gradients. The mean velocity and Reynolds stress components are measured using a X-hot-film probe. The measured quantities obtained in probe coordinates are transformed to a curvilinear coordinate system along the wake centerline and are presented in similarity coordinates. The results of the steady wakes suggest strong asymmetry in velocity and Reynolds stress components. However, the velocity defect profiles in similarity coordinates are almost symmetrical and follow the same distribution as the zero pressure gradient straight wake. The results of Reynolds stress distributions show higher values on the inner side of the wake than the outer side. Other quantities, including the decay of maximum velocity defect, growth of wake width, and wake integral parameters, are also presented for the three different pressure gradient cases of steady wake. The decay rate of velocity defect is fastest for the negative streamwise pressure gradient case and slowest for the positive pressure gradient case. Conversely, the growth of the wake width is fastest for the positive streamwise pressure gradient case and slowest for the negative streamwise pressure gradient. The second phase studies the development of periodic unsteady wakes generated by the circular cylinders of the rotating wake generator in a curved channel at zero streamwise pressure gradient. Instantaneous velocity components of the periodic unsteady wakes, measured with a stationary X-hot-film probe, are analyzed by the
NASA Astrophysics Data System (ADS)
Chen, G. H.; Y Wang, G.; Huang, B.; Huang, X.; Gao, Y.
2013-12-01
To address the unsteady cavity behaviour and induced pressure fluctuation in cloud cavitating flow, cavitation images and pressure fluctuation signals are simultaneously acquired by high speed visualization system and 4 piezo-electric transducers in a convergent-divergent channel. The cavitation images are processed by using a home-developed software to obtain the time evolutions of global cavity area. Frequency analysis is conducted for both global cavity area and pressure signal. Bubble dynamics is introduced to analyze the correlation between pressure fluctuation in the downstream and global cavity behaviour. Two conclusions are achieved: First, in cloud cavitating flow, the time evolution of both the cavity behaviour and pressure fluctuation are quasi-periodic, one quasi-period can be divided into three main stages: growth of attached cavity, shedding of attached cavity, coalescence and collapse of detached cavity. Second, the dominant frequency of global cavity area and pressure fluctuation on 4 transducers are the same, it's 20Hz in this study. Third, it's found that during the stage of growth of attached cavity and growth, collapse of detached cavity, the correlation between global cavity area and induced pressure in the downstream is similar with that of a single bubble; while, such correlation is not clear when several travelling cavities exist at the same time.
Trailing edge modifications for flatback airfoils.
Kahn, Daniel L.; van Dam, C.P.; Berg, Dale E.
2008-03-01
The adoption of blunt trailing edge airfoils (also called flatback airfoils) for the inboard region of large wind turbine blades has been proposed. Blunt trailing edge airfoils would not only provide a number of structural benefits, such as increased structural volume and ease of fabrication and handling, but they have also been found to improve the lift characteristics of thick airfoils. Therefore, the incorporation of blunt trailing edge airfoils would allow blade designers to more freely address the structural demands without having to sacrifice aerodynamic performance. These airfoils do have the disadvantage of generating high levels of drag as a result of the low-pressure steady or periodic flow in the near-wake of the blunt trailing edge. Although for rotors, the drag penalty appears secondary to the lift enhancement produced by the blunt trailing edge, high drag levels are of concern in terms of the negative effect on the torque and power generated by the rotor. Hence, devices are sought that mitigate the drag of these airfoils. This report summarizes the literature on bluff body vortex shedding and bluff body drag reduction devices and proposes four devices for further study in the wind tunnel.
Investigation of low-speed turbulent separated flow around airfoils
NASA Technical Reports Server (NTRS)
Wadcock, Alan J.
1987-01-01
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=0.85. 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.
On the Use of Surface Porosity to Reduce Unsteady Lift
NASA Technical Reports Server (NTRS)
Tinetti, Ana F.; Kelly, Jeffrey J.; Bauer, Steven X. S.; Thomas, Russell H.
2001-01-01
An innovative application of existing technology is proposed for attenuating the effects of transient phenomena, such as rotor-stator and rotor-strut interactions, linked to noise and fatigue failure in turbomachinery environments. A computational study was designed to assess the potential of passive porosity technology as a mechanism for alleviating interaction effects by reducing the unsteady lift developed on a stator airfoil subject to wake impingement. The study involved a typical high bypass fan Stator airfoil (solid baseline and several porous configurations), immersed in a free field and exposed to the effects of a transversely moving wake. It was found that, for the airfoil under consideration, the magnitude of the unsteady lift could be reduced more than 18% without incurring significant performance losses.
Unsteady turbulent boundary layer analysis
NASA Technical Reports Server (NTRS)
Singleton, R. E.; Nash, J. F.; Carl, L. W.; Patel, V. C.
1973-01-01
The governing equations for an unsteady turbulent boundary layer on a swept infinite cylinder, composed of a continuity equation, a pair of momentum equations and a pair of turbulent energy equations which include upstream history efforts, are solved numerically. An explicit finite difference analog to the partial differential equations is formulated and developed into a computer program. Calculations were made for a variety of unsteady flows in both two and three dimensions but primarily for two dimensional flow fields in order to first understand some of the fundamental physical aspects of unsteady turbulent boundary layers. Oscillating free stream flows without pressure gradient, oscillating retarded free stream flows and monotonically time-varying flows have all been studied for a wide frequency range. It was found that to the lowest frequency considered, the lower frequency bound being determined by economic considerations (machine time), there were significant unsteady effects on the turbulent boundary layer.
NASA Technical Reports Server (NTRS)
Scott, James R.
1991-01-01
A numerical method is developed for solving periodic, three-dimensional, vortical flows around lifting airfoils in subsonic flow. The first-order method that is presented fully accounts for the distortion effects of the nonuniform mean flow on the convected upstream vortical disturbances. The unsteady velocity is split into a vortical component which is a known function of the upstream flow conditions and the Lagrangian coordinates of the mean flow, and an irrotational field whose potential satisfies a nonconstant-coefficient, inhomogeneous, convective wave equation. Using an elliptic coordinate transformation, the unsteady boundary value problem is solved in the frequency domain on grids which are determined as a function of the Mach number and reduced frequency. The numerical scheme is validated through extensive comparisons with known solutions to unsteady vortical flow problems. In general, it is seen that the agreement between the numerical and analytical results is very good for reduced frequencies ranging from 0 to 4, and for Mach numbers ranging from .1 to .8. Numerical results are also presented for a wide variety of flow configurations for the purpose of determining the effects of airfoil thickness, angle of attack, camber, and Mach number on the unsteady lift and moment of airfoils subjected to periodic vortical gusts. It is seen that each of these parameters can have a significant effect on the unsteady airfoil response to the incident disturbances, and that the effect depends strongly upon the reduced frequency and the dimensionality of the gust. For a one-dimensional (transverse) or two-dimensional (transverse and longitudinal) gust, the results indicate that airfoil thickness increases the unsteady lift and moment at the low reduced frequencies but decreases it at the high reduced frequencies. The results show that an increase in airfoil Mach number leads to a significant increase in the unsteady lift and moment for the low reduced frequencies, but a
NASA Technical Reports Server (NTRS)
Ramsey, John K.; Erwin, Dan
2005-01-01
Experimental data were obtained to help validate analytical and computational fluid dynamics (CFD) codes used to compute unsteady cascade aerodynamics in a supersonicaxial- flow regime. Results from two analytical codes and one CFD code were compared with experimental data. One analytical code did not account for airfoil thickness or camber; another, using piston theory (piston code), accounted for thickness and camber upstream of the first shockwave/airfoil impingement locations. The Euler CFD code accounted fully for airfoil shape.
The pressure field of a gust interacting with a flat plate
NASA Technical Reports Server (NTRS)
Atassi, Hafiz M.; Patrick, Sheryl M.
1995-01-01
A semianalytical solution is presented for the unsteady pressure field of a vortical gust interacting with a flat-plate airfoil in subsonic flow. The solution will serve as a benchmark for evaluating the accuracy and efficiency of time dependent numerical schemes. The specific case considered corresponds to the ICASE benchmark problem number 6. The results are compared with those of asymptotic theories for high frequency and show excellent agreement.
NASA Technical Reports Server (NTRS)
Whitcomb, R. T. (Inventor)
1976-01-01
An airfoil is examined that has an upper surface shaped to control flow accelerations and pressure distribution over the upper surface and to prevent separation of the boundary layer due to shock wave formulation at high subsonic speeds well above the critical Mach number. A highly cambered trailing edge section improves overall airfoil lifting efficiency. Diagrams illustrating supersonic flow and shock waves over the airfoil are shown.
NASA Technical Reports Server (NTRS)
Batina, J. T.
1985-01-01
Unsteady transonic flow calculations for aerodynamically interfering airfoil configurations are performed as a first-step toward solving the three-dimensional canard-wing interaction problem. These calculations are performed by extending the XTRAN2L two-dimensional unsteady transonic small-disturbance code to include an additional airfoil. Unsteady transonic forces due to plunge and pitch motions of a two-dimensional canard and wing are presented. Results for a variety of canard-wing separation distances reveal the effects of aerodynamic interference on unsteady transonic airloads. Aeroelastic analyses employing these unsteady airloads demonstrate the effects of aerodynamic interference on aeroelastic stability and flutter. For the configurations studied, increases in wing flutter speed result with the inclusion of the aerodynamically interfering canard.
NASA Technical Reports Server (NTRS)
Batina, J. T.
1985-01-01
Unsteady transonic flow calculations for aerodynamically interfering airfoil configurations are performed as a first step toward solving the three dimensional canard wing interaction problem. These calculations are performed by extending the XTRAN2L two dimensional unsteady transonic small disturbance code to include an additional airfoil. Unsteady transonic forces due to plunge and pitch motions of a two dimensional canard and wing are presented. Results for a variety of canard wing separation distances reveal the effects of aerodynamic interference on unsteady transonic airloads. Aeroelastic analyses employing these unsteady airloads demonstrate the effects of aerodynamic interference on aeroelastic stability and flutter. For the configurations studied, increases in wing flutter speed result with the inclusion of the aerodynamically interfering canard.
Numerical study of porous airfoils in transonic flow
NASA Technical Reports Server (NTRS)
Chen, C. L.; Chow, C. Y.; Holst, T. L.; Vandalsem, W. R.
1985-01-01
A numerical study was made to examine the effect of a porous surface on the aerodynamic performance of a transonic airfoil. The pressure jump across the normal shock wave on the upper surface of the airfoil was reduced by making the surface below the shock porous. The weakened shock is preceded by an oblique shock at the upstream end of the porous surface where air is blown out of the cavity. The lambda shock structure shown in the numerical result qualitatively agrees with that observed in the wind tunnel. According to the present analysis, the porous airfoil has a smaller drag and a higher lift than the solid airfoil.
The effect of a cavity on airfoil tones
NASA Astrophysics Data System (ADS)
Schumacher, Karn L.; Doolan, Con J.; Kelso, Richard M.
2014-03-01
The presence of a cavity in the pressure surface of an airfoil has been found via experiment to play a role in the production of airfoil tones, which was attributed to the presence of an acoustic feedback loop. The cavity length was sufficiently small that cavity oscillation modes did not occur for most of the investigated chord-based Reynolds number range of 70,000-320,000. The airfoil tonal noise frequencies varied as the position of the cavity was moved along a parallel section at the airfoil's maximum thickness: specifically, for a given velocity, the frequency spacing of the tones was inversely proportional to the geometric distance between the cavity and the trailing edge. The boundary layer instability waves considered responsible for the airfoil tones were only detected downstream of the cavity. This may be the first experimental verification of these aspects of the feedback loop model for airfoil tonal noise.
Application of two procedures for dual-point design of transonic airfoils
NASA Technical Reports Server (NTRS)
Mineck, Raymond E.; Campbell, Richard L.; Allison, Dennis O.
1994-01-01
Two dual-point design procedures were developed to reduce the objective function of a baseline airfoil at two design points. The first procedure to develop a redesigned airfoil used a weighted average of the shapes of two intermediate airfoils redesigned at each of the two design points. The second procedure used a weighted average of two pressure distributions obtained from an intermediate airfoil redesigned at each of the two design points. Each procedure was used to design a new airfoil with reduced wave drag at the cruise condition without increasing the wave drag or pitching moment at the climb condition. Two cycles of the airfoil shape-averaging procedure successfully designed a new airfoil that reduced the objective function and satisfied the constraints. One cycle of the target (desired) pressure-averaging procedure was used to design two new airfoils that reduced the objective function and came close to satisfying the constraints.
NASA Technical Reports Server (NTRS)
Mutterperl, William
1944-01-01
A method of conformal transformation is developed that maps an airfoil into a straight line, the line being chosen as the extended chord line of the airfoil. The mapping is accomplished by operating directly with the airfoil ordinates. The absence of any preliminary transformation is found to shorten the work substantially over that of previous methods. Use is made of the superposition of solutions to obtain a rigorous counterpart of the approximate methods of thin-airfoils theory. The method is applied to the solution of the direct and inverse problems for arbitrary airfoils and pressure distributions. Numerical examples are given. Applications to more general types of regions, in particular to biplanes and to cascades of airfoils, are indicated. (author)
Near-wall serpentine cooled turbine airfoil
Lee, Ching-Pang
2013-09-17
A serpentine coolant flow path (54A-54G) formed by inner walls (50, 52) in a cavity (49) between pressure and suction side walls (22, 24) of a turbine airfoil (20A). A coolant flow (58) enters (56) an end of the airfoil, flows into a span-wise channel (54A), then flows forward (54B) over the inner surface of the pressure side wall, then turns behind the leading edge (26), and flows back along a forward part of the suction side wall, then follows a loop (54E) forward and back around an inner wall (52), then flows along an intermediate part of the suction side wall, then flows into an aft channel (54G) between the pressure and suction side walls, then exits the trailing edge (28). This provides cooling matched to the heating topography of the airfoil, minimizes differential thermal expansion, revives the coolant, and minimizes the flow volume needed.
Analytical studies of new airfoils for wind turbines
NASA Technical Reports Server (NTRS)
Wentz, W. H., Jr.; Calhoun, J. T.
1981-01-01
Computer studies were conducted to analyze the potential gains associated with utilizing new airfoils for large wind turbine rotor blades. Attempts to include 3-dimensional stalling effects were inconclusive. It is recommended that blade pressure measurements be made to clarify the nature of blade stalling. It is also recommended that new laminar flow airfoils be used as rotor blade sections.
Active Control of Separation From the Flap of a Supercritical Airfoil
NASA Technical Reports Server (NTRS)
Melton, La Tunia Pack; Yao, Chung-Sheng; Seifert, Avi
2003-01-01
Active flow control in the form of periodic zero-mass-flux excitation was applied at several regions on the leading edge and trailing edge flaps of a simplified high-lift system t o delay flow separation. The NASA Energy Efficient Transport (EET) supercritical airfoil was equipped with a 15% chord simply hinged leading edge flap and a 25% chord simply hinged trailing edge flap. Detailed flow features were measured in an attempt to identify optimal actuator placement. The measurements included steady and unsteady model and tunnel wall pressures, wake surveys, arrays of surface hot-films, flow visualization, and particle image velocimetry (PIV). The current paper describes the application of active separation control at several locations on the deflected trailing edge flap. High frequency (F(+) approx.= 10) and low frequency amplitude modulation (F(+)AM approx.= 1) of the high frequency excitation were used for control. Preliminary efforts to combine leading and trailing edge flap excitations are also reported.
Minnowbrook IV: 2003 Workshop on Transition and Unsteady Aspects of Turbomachinery Flows
NASA Technical Reports Server (NTRS)
LaGraff, John E. (Editor); Ashpis, David E.
2004-01-01
This Minnowbrook IV 2003 workshop on Transition and Unsteady Aspects of Turbomachinery Flows includes the following topics: 1) Current Issues in Unsteady Turbomachinery Flows; 2) Global Instability and Control of Low-Pressure Turbine Flows; 3) Influence of End Wall Leakage on Secondary Flow Development in Axial Turbines; 4) Active and Passive Flow Control on Low Pressure Turbine Airfoils; 5) Experimental and Numerical Investigation of Transitional Flows as Affected by Passing Wakes; 6) Effects of Freestream Turbulence on Turbine Blade Heat Transfer; 7) Bypass Transition Via Continuous Modes and Unsteady Effects on Film Cooling; 8) High Frequency Surface Heat Flux Imaging of Bypass Transition; 9) Skin Friction and Heat Flux Oscillations in Upstream Moving Wave Packets; 10) Transition Mechanisms and Use of Surface Roughness to Enhance the Benefits of Wake Passing in LP Turbines; 11) Transient Growth Approach to Roughness-Induced Transition; 12) Roughness- and Freestream-Turbulence-Induced Transient Growth as a Bypass Transition Mechanism; 13) Receptivity Calculations as a Means to Predicting Transition; 14) On Streamwise Vortices in a Curved Wall Jet and Their Effect on the Mean Flow; 15) Plasma Actuators for Separation Control of Low Pressure Turbine Blades; 16) Boundary-Layer Separation Control Under Low-Pressure-Turbine Conditions Using Glow-Discharge Plasma Actuators; 17) Control of Separation for Low Pressure Turbine Blades: Numerical Simulation; 18) Effects of Elevated Free-Stream Turbulence on Active Control of a Separation Bubble; 19) Wakes, Calming and Transition Under Strong Adverse Pressure Gradients; 20) Transitional Bubble in Periodic Flow Phase Shift; 21) Modelling Spots: The Calmed Region, Pressure Gradient Effects and Background; 22) Modeling of Unsteady Transitional Flow on Axial Compressor Blades; 23) Challenges in Predicting Component Efficiencies in Turbomachines With Low Reynolds Number Blading; 24) Observations on the Causal Relationship Between
Effect of cavity on shock oscillation in transonic flow over RAE2822 supercritical airfoil
NASA Astrophysics Data System (ADS)
Rahman, M. Rizwanur; Labib, Md. Itmam; Hasan, A. B. M. Toufique; Ali, M.; Mitsutake, Y.; Setoguchi, T.
2016-07-01
Transonic flow past a supercritical airfoil is strongly influenced by the interaction of shock wave with boundary layer. This interaction induces unsteady self-sustaining shock wave oscillation, flow instability, drag rise and buffet onset which limit the flight envelop. In the present study, a computational analysis has been carried out to investigate the flow past a supercritical RAE2822 airfoil in transonic speeds. To control the shock wave oscillation, a cavity is introduced on the airfoil surface where shock wave oscillates. Different geometric configurations have been investigated for finding optimum cavity geometry and dimension. Unsteady Reynolds averaged Navier-Stokes equations (RANS) are computed at Mach 0.729 with an angle of attack of 5°. Computed results are well validated with the available experimental data in case of baseline airfoil. However, in case of airfoil with control cavity; it has been observed that the introduction of cavity completely suppresses the unsteady shock wave oscillation. Further, significant drag reduction and successive improvement of aerodynamic performance have been observed in airfoil with shock control cavity.
Multi-pass cooling for turbine airfoils
Liang, George
2011-06-28
An airfoil for a turbine vane of a gas turbine engine. The airfoil includes an outer wall having pressure and suction sides, and a radially extending cooling cavity located between the pressure and suction sides. A plurality of partitions extend radially through the cooling cavity to define a plurality of interconnected cooling channels located at successive chordal locations through the cooling cavity. The cooling channels define a serpentine flow path extending in the chordal direction. Further, the cooling channels include a plurality of interconnected chambers and the chambers define a serpentine path extending in the radial direction within the serpentine path extending in the chordal direction.
NASA Technical Reports Server (NTRS)
Ozturk, B.; Schobeiri, M. T.; Ashpis, David E.
2005-01-01
The paper experimentally studies the effects of periodic unsteady wake flow and different Reynolds numbers on boundary layer development, separation and re-attachment along the suction surface of a low pressure turbine blade. The experimental investigations were performed on a large scale, subsonic unsteady turbine cascade research facility at Turbomachinery Performance and Flow Research Laboratory (TPFL) of Texas A&M University. The experiments were carried out at Reynolds numbers of 110,000 and 150,000 (based on suction surface length and exit velocity). One steady and two different unsteady inlet flow conditions with the corresponding passing frequencies, wake velocities, and turbulence intensities were investigated. The reduced frequencies chosen cover the operating range of LP turbines. In addition to the unsteady boundary layer measurements, surface pressure measurements were performed. The inception, onset, and the extent of the separation bubble information collected from the pressure measurements were compared with the hot wire measurements. The results presented in ensemble-averaged, and the contour plot forms help to understand the physics of the separation phenomenon under periodic unsteady wake flow and different Reynolds number. It was found that the suction surface displayed a strong separation bubble for these three different reduced frequencies. For each condition, the locations defining the separation bubble were determined carefully analyzing and examining the pressure and mean velocity profile data. The location of the boundary layer separation was dependent of the Reynolds number. It is observed that starting point of the separation bubble and the re-attachment point move further downstream by increasing Reynolds number from 110,000 to 150,000. Also, the size of the separation bubble is smaller when compared to that for Re=110,000.
Potential flow analysis of glaze ice accretions on an airfoil
NASA Technical Reports Server (NTRS)
Zaguli, R. J.
1984-01-01
The results of an analytical/experimental study of the flow fields about an airfoil with leading edge glaze ice accretion shapes are presented. Tests were conducted in the Icing Research Tunnel to measure surface pressure distributions and boundary layer separation reattachment characteristics on a general aviation wing section to which was affixed wooden ice shapes which approximated typical glaze ice accretions. Comparisons were made with predicted pressure distributions using current airfoil analysis codes as well as the Bristow mixed analysis/design airfoil panel code. The Bristow code was also used to predict the separation reattachment dividing streamline by inputting the appropriate experimental surface pressure distribution.
Unsteady aerodynamic modeling and active aeroelastic control
NASA Technical Reports Server (NTRS)
Edwards, J. W.
1977-01-01
Unsteady aerodynamic modeling techniques are developed and applied to the study of active control of elastic vehicles. The problem of active control of a supercritical flutter mode poses a definite design goal stability, and is treated in detail. The transfer functions relating the arbitrary airfoil motions to the airloads are derived from the Laplace transforms of the linearized airload expressions for incompressible two dimensional flow. The transfer function relating the motions to the circulatory part of these loads is recognized as the Theodorsen function extended to complex values of reduced frequency, and is termed the generalized Theodorsen function. Inversion of the Laplace transforms yields exact transient airloads and airfoil motions. Exact root loci of aeroelastic modes are calculated, providing quantitative information regarding subcritical and supercritical flutter conditions.
Study of Boundary Layer Development in a Two-Stage Low-Pressure Turbine
NASA Technical Reports Server (NTRS)
Dorney, Daniel J.; Ashpis, David E.; Halstead, David E.; Wisler, David C.
1999-01-01
Experimental data from jet-engine tests have indicated that unsteady blade row interactions and separation can have a significant impact on the efficiency of low-pressure turbine stages. Measured turbine efficiencies at takeoff can be as much as two points higher than those at cruise conditions. Several recent studies have revealed that Reynolds number effects may contribute to the lower efficiencies at cruise conditions. In the current study numerical simulations have been performed to study the boundary layer development in a two-stage low-pressure turbine, and to evaluate the transition models available for low Reynolds number flows in turbomachinery. The results of the simulations have been compared with experimental data, including airfoil loadings and integral boundary layer quantities. The predicted unsteady results display similar trends to the experimental data, but significantly overestimate the amplitude of the unsteadiness. The time-averaged results show close agreement with the experimental data.
Development and testing of airfoils for high-altitude aircraft
NASA Technical Reports Server (NTRS)
Drela, Mark (Principal Investigator)
1996-01-01
Specific tasks included airfoil design; study of airfoil constraints on pullout maneuver; selection of tail airfoils; examination of wing twist; test section instrumentation and layout; and integrated airfoil/heat-exchanger tests. In the course of designing the airfoil, specifically for the APEX test vehicle, extensive studies were made over the Mach and Reynolds number ranges of interest. It is intended to be representative of airfoils required for lightweight aircraft operating at extreme altitudes, which is the primary research objective of the APEX program. Also considered were thickness, pitching moment, and off-design behavior. The maximum ceiling parameter M(exp 2)C(sub L) value achievable by the Apex-16 airfoil was found to be a strong constraint on the pullout maneuver. The NACA 1410 and 2410 airfoils (inverted) were identified as good candidates for the tail, with predictable behavior at low Reynolds numbers and good tolerance to flap deflections. With regards to wing twist, it was decided that a simple flat wing was a reasonable compromise. The test section instrumentation consisted of surface pressure taps, wake rakes, surface-mounted microphones, and skin-friction gauges. Also, a modest wind tunnel test was performed for an integrated airfoil/heat-exchanger configuration, which is currently on Aurora's 'Theseus' aircraft. Although not directly related to the APEX tests, the aerodynamics or heat exchangers has been identified as a crucial aspect of designing high-altitude aircraft and hence is relevant to the ERAST program.
NASA Technical Reports Server (NTRS)
Lamah, C. A.; Harris, W. L.
1983-01-01
A novel analytical-numerical method for calculating unsteady small disturbance transonic flow over airfoils has been developed. The method uses an extended integral equation technique, based on both the velocity potential and the acceleration potential, to predict unsteady aerodynamic loading on airfoils oscillating in subcritical transonic free stream conditions. The formulation is an extension of the work of Sivaneri and Harris (1980) for steady, non-lifting flows and utilizes the linear theory of Landahl (1961) for decoupling of steady and unsteady components. The analytical-numerical procedure involves several intnegrating schemes and applies to general frequencies of oscillations. The technique is illustrated by computing the transonic flow about parabolic arc airfoils. Specific unsteady results for reduced frequencies based on semi-chord of 0.01, 0.1, 0.3, 0.4 and 0.6 are given. Comparison of results with those obtained by an ADI finite difference scheme is made.
An approach to the constrained design of natural laminar flow airfoils
NASA Technical Reports Server (NTRS)
Green, Bradford Earl
1995-01-01
A design method has been developed by which an airfoil with a substantial amount of natural laminar flow can be designed, while maintaining other aerodynamic and geometric constraints. After obtaining the initial airfoil's pressure distribution at the design lift coefficient using an Euler solver coupled with an integml turbulent boundary layer method, the calculations from a laminar boundary layer solver are used by a stability analysis code to obtain estimates of the transition location (using N-Factors) for the starting airfoil. A new design method then calculates a target pressure distribution that will increase the larninar flow toward the desired amounl An airfoil design method is then iteratively used to design an airfoil that possesses that target pressure distribution. The new airfoil's boundary layer stability characteristics are determined, and this iterative process continues until an airfoil is designed that meets the laminar flow requirement and as many of the other constraints as possible.
An Approach to the Constrained Design of Natural Laminar Flow Airfoils
NASA Technical Reports Server (NTRS)
Green, Bradford E.
1997-01-01
A design method has been developed by which an airfoil with a substantial amount of natural laminar flow can be designed, while maintaining other aerodynamic and geometric constraints. After obtaining the initial airfoil's pressure distribution at the design lift coefficient using an Euler solver coupled with an integral turbulent boundary layer method, the calculations from a laminar boundary layer solver are used by a stability analysis code to obtain estimates of the transition location (using N-Factors) for the starting airfoil. A new design method then calculates a target pressure distribution that will increase the laminar flow toward the desired amount. An airfoil design method is then iteratively used to design an airfoil that possesses that target pressure distribution. The new airfoil's boundary layer stability characteristics are determined, and this iterative process continues until an airfoil is designed that meets the laminar flow requirement and as many of the other constraints as possible.
Airfoil Vibration Dampers program
NASA Technical Reports Server (NTRS)
Cook, Robert M.
1991-01-01
The Airfoil Vibration Damper program has consisted of an analysis phase and a testing phase. During the analysis phase, a state-of-the-art computer code was developed, which can be used to guide designers in the placement and sizing of friction dampers. The use of this computer code was demonstrated by performing representative analyses on turbine blades from the High Pressure Oxidizer Turbopump (HPOTP) and High Pressure Fuel Turbopump (HPFTP) of the Space Shuttle Main Engine (SSME). The testing phase of the program consisted of performing friction damping tests on two different cantilever beams. Data from these tests provided an empirical check on the accuracy of the computer code developed in the analysis phase. Results of the analysis and testing showed that the computer code can accurately predict the performance of friction dampers. In addition, a valuable set of friction damping data was generated, which can be used to aid in the design of friction dampers, as well as provide benchmark test cases for future code developers.
On wall-pressure sources associated with the unsteady separation in a vortex-ring/wall interaction
NASA Astrophysics Data System (ADS)
Naguib, Ahmed M.; Koochesfahani, Manoochehr M.
2004-07-01
Molecular tagging velocimetery measurements from an earlier investigation are used to study the wall-pressure and the flow structures responsible for its generation in the flow field resulting from the impingement of an axisymmetric vortex ring on a flat wall. The velocity-field data are used to obtain the spatial distribution of the pressure sources, and the result is employed in conjunction with the solution to Poisson's equation to yield the wall-pressure information. The outcome reveals a characteristic wall-pressure signature that is produced repeatedly whenever the primary vortex ring interacts with the wall to form vortex rings with opposite sense of vorticity. The details of the signature are analyzed and related to specific flow features and their mutual interaction at different phases of the generation/evolution cycle of the new rings. Finally, the flow mechanisms leading to the generation of substantial positive and negative wall pressure in the characteristic signature are clarified.
Response of a thin airfoil encountering strong density discontinuity
Marble, F.E.
1993-12-01
Airfoil theory for unsteady motion has been developed extensively assuming the undisturbed medium to be of uniform density, a restriction accurate for motion in the atmosphere. In some instances, notably for airfoil comprising fan, compressor and turbine blade rows, the undisturbed medium may carry density variations or ``spots``, resulting from non-uniformities in temperature or composition, of a size comparable to the blade chord. This condition exists for turbine blades, immediately downstream of the main burner of a gas turbine engine where the density fluctuations of the order of 50 percent may occur. Disturbances of a somewhat smaller magnitude arise from the ingestion of hot boundary layers into fans, and exhaust into hovercraft. Because these regions of non-uniform density convect with the moving medium, the airfoil experiences a time varying load and moment which the authors calculate.
Numerical simulations of iced airfoils and wings
NASA Astrophysics Data System (ADS)
Pan, Jianping
A numerical study was conducted to understand the effects of simulated ridge and leading-edge ice shapes on the aerodynamic performance of airfoils and wings. In the first part of this study, a range of Reynolds numbers and Mach numbers, as well as ice-shape sizes and ice-shape locations were examined for various airfoils with the Reynolds-Averaged Navier-Stokes approach. Comparisons between simulation results and experimental force data showed favorable comparison up to stall conditions. At and past stall condition, the aerodynamic forces were typically not predicted accurately for large upper-surface ice shapes. A lift-break (pseudo-stall) condition was then defined based on the lift curve slope change. The lift-break angles compared reasonably with experimental stall angles, and indicated that the critical ice-shape location tended to be near the location of minimum pressure and the location of the most adverse pressure gradient. With the aim of improving the predictive ability of the stall behavior for iced airfoils, simulations using the Detached Eddy Simulation (DES) approach were conducted in the second part of this numerical investigation. Three-dimensional DES computations were performed for a series of angles of attack around stall for the iced NACA 23012 and NLF 0414 airfoils. The simulations for both iced airfoils provided the maximum lift coefficients and stall behaviors qualitatively consistent with experiments.
Aerodynamic Characteristics of Airfoils at High Speeds
NASA Technical Reports Server (NTRS)
Briggs, L J; Hull, G F; Dryden, H L
1925-01-01
This report deals with an experimental investigation of the aerodynamical characteristics of airfoils at high speeds. Lift, drag, and center of pressure measurements were made on six airfoils of the type used by the air service in propeller design, at speeds ranging from 550 to 1,000 feet per second. The results show a definite limit to the speed at which airfoils may efficiently be used to produce lift, the lift coefficient decreasing and the drag coefficient increasing as the speed approaches the speed of sound. The change in lift coefficient is large for thick airfoil sections (camber ratio 0.14 to 0.20) and for high angles of attack. The change is not marked for thin sections (camber ratio 0.10) at low angles of attack, for the speed range employed. At high speeds the center of pressure moves back toward the trailing edge of the airfoil as the speed increases. The results indicate that the use of tip speeds approaching the speed of sound for propellers of customary design involves a serious loss in efficiency.
Effect of an extendable slat on the stall behavior of a VR-12 airfoil
NASA Technical Reports Server (NTRS)
Dehugues, P. Plantin; Mcalister, K. W.; Tung, C.
1993-01-01
Experimental and computational tests were performed on a VR-12 airfoil to determine if the dynamic-stall behavior that normally accompanies high-angle pitch oscillations could be modified by segmenting the forward portion of the airfoil and extending it ahead of the main element. In the extended position the configuration would appear as an airfoil with a leading-edge slat, and in the retracted position it would appear as a conventional VR-12 airfoil. The calculations were obtained from a numerical code that models the vorticity transport equation for an incompressible fluid. These results were compared with test data from the water tunnel facility of the Aeroflightdynamics Directorate at Ames Research Center. Steady and unsteady flows around both airfoils were examined at angles of attack between 0 and 30 deg. The Reynolds number was fixed at 200,000 and the unsteady pitch oscillations followed a sinusoidal motion described by alpha = alpha(sub m) + 10 deg sin(omega t). The mean angle (alpha(sub m)) was varied from 10 to 20 deg and the reduced frequency from 0.05 to 0.20. The results from the experiment and the calculations show that the extended-slat VR-12 airfoil experiences a delay in both static and dynamic stall not experienced by the basic VR-12 airfoil.
Unsteady transonic flow over cascade blades
NASA Technical Reports Server (NTRS)
Surampudi, S. P.; Adamczyk, J. J.
1986-01-01
An attempt is made to develop an efficient staggered cascade blade unsteady aerodynamics model for the neighborhood of March 1, representing the blade row by a rectilinear two-dimensional cascade of thin, flat plate airfoils. The equations of motion are derived on the basis of linearized transonic small perturbation theory, and an analytical solution is obtained by means of the Wiener-Hopf procedure. Making use of the transonic similarity law, the results obtained are compared with those of other linearized cascade analyses. A parametric study is conducted to find the effects of reduced frequency, stagger angle, solidity, and the location of the pitching axis on cascade stability.
NASA Technical Reports Server (NTRS)
Bingham, G. J.; Noonan, K. W.
1974-01-01
An investigation was conducted in a low-turbulence pressure tunnel to determine the two-dimensional lift and pitching-moment characteristics of an NACA 6716 and an NACA 4416 airfoil with 35-percent-chord single-slotted flaps. Both models were tested with flaps deflected from 0 deg to 45 deg, at angles of attack from minus 6 deg to several degrees past stall, at Reynolds numbers from 3.0 million to 13.8 million, and primarily at a Mach number of 0.23. Tests were also made to determine the effect of several slot entry shapes on performance.
NASA Technical Reports Server (NTRS)
Catalano, G. D.; Morton, J. B.; Humphris, R. R.
1978-01-01
The effects of increasing the velocity ratio, lambda sub j were explored. The quantities measured include the width of the mixing region, the mean velocity field, turbulent intensities and time scales. In addition, wall and static pressure velocity correlations and coherences are presented. The velocity measurements were made using a laser Doppler velocimeter with a phase locked loop processor. The fluctuating pressures were monitored using condenser type microphones.
NASA Technical Reports Server (NTRS)
Bingham, G. J. (Inventor)
1984-01-01
An airfoil which has particular application to the blade or blades of rotor aircraft and aircraft propellers is presented. The airfoil thickness distribution, camber and leading edge radius are shaped to locate the airfoil crest at a more aft position along the chord, and to increase the freestream Mach number at which sonic flow is attained at the airfoil crest. The reduced slope of the airfoil causes a reduction in velocity at the airfoil crest at lift coefficients from zero to the maximum lift coefficient. The leading edge radius is adjusted so that the maximum local Mach number at 1.25 percent chord and at the designed maximum lift coefficient is limited to about 0.48 when the Mach number normal to the leading edge is approximately 0.20. The lower surface leading edge radius is shaped so that the maximum local Mach number at the leading edge is limited to about 0.29 when the Mach number normal to the leading edge is approximately 0.20. The drag divergence Mach number associated with the airfoil is moved to a higher Mach number over a range of lift coefficients resulting in superior aircraft performance.
Theoretical study on two-dimensional aerodynamic characteristics of unsteady wings.
Azuma, Akira; Okamoto, Masato
2005-05-01
A simple computing method based on a potential theory is developed for two-dimensional steady and unsteady deflected wings. This method of theoretical analysis is essentially related to thin and angular airfoils. Thus, the method is very simple but is effective to forecast aerodynamic forces for deflected or angular airfoils with a small camber operating in high Reynolds number flow, specifically in unsteady motion. The suction force acting on the leading edge of steady airfoils is theoretically obtained by using the Blasius formula. By Polhamus's leading edge suction analogy, the suction force is considered to be directed upward in partially separated flow for real thin airfoil with sharp leading edge. The theory can also be applied to obtain the aerodynamic characteristics of thin airfoils operating on low Reynolds number flow under some degree of approximation. This is very useful for the unsteady aerodynamic analysis because the Navier-Stokes equation can be solved by neither analytical nor numerical method for the thin and angular airfoils, which are common in the insect wing.
NASA Astrophysics Data System (ADS)
Ford, C. L.; Winroth, M.; Alfredsson, P. H.
2016-08-01
An entirely pressure-based vortex-shedding meter has been designed for use in practical time-dependent flows. The meter is capable of measuring mass-flow rate in variable density gases in spite of the fact that fluid temperature is not directly measured. Unlike other vortex meters, a pressure based meter is incredibly robust and may be used in industrial type flows; an environment wholly unsuitable for hot-wires for example. The meter has been tested in a number of static and dynamic flow cases, across a range of mass-flow rates and pressures. The accuracy of the meter is typically better than about 3% in a static flow and resolves the fluctuating mass-flow with an accuracy that is better than or equivalent to a hot-wire method.
Wind tunnel testing of low-drag airfoils
NASA Technical Reports Server (NTRS)
Harvey, W. Donald; Mcghee, R. J.; Harris, C. D.
1986-01-01
Results are presented for the measured performance recently obtained on several airfoil concepts designed to achieve low drag by maintaining extensive regions of laminar flow without compromising high-lift performance. The wind tunnel results extend from subsonic to transonic speeds and include boundary-layer control through shaping and suction. The research was conducted in the NASA Langley 8-Ft Transonic Pressure Tunnel (TPT) and Low Turbulence Pressure Tunnel (LTPT) which have been developed for testing such low-drag airfoils. Emphasis is placed on identifying some of the major factors influencing the anticipated performance of low-drag airfoils.
Near-wall serpentine cooled turbine airfoil
Lee, Ching-Pang
2014-10-28
A serpentine coolant flow path is formed by inner walls in a cavity between pressure and suction side walls of a turbine airfoil, the cavity partitioned by one or more transverse partitions into a plurality of continuous serpentine cooling flow streams each having a respective coolant inlet.
Blowing Circulation Control on a Seaplane Airfoil
NASA Astrophysics Data System (ADS)
Guo, B. D.; Liu, P. Q.; Qu, Q. L.
2011-09-01
RANS simulations are presented for blowing circulation control on a seaplane airfoil. Realizable k-epsilon turbulent model and pressure-based coupled algorithm with second-order discretization were adopted to simulate the compressible flow. Both clear and simple flap configuration were simulated with blowing momentum coefficient Cμ = 0, 0.15 and 0.30. The results show that blowing near the airfoil trailing edge could enhance the Coanda effect, delay the flow separation, and increase the lift coefficient dramatically. The blowing circulation control is promising to apply to taking off and landing of an amphibious aircraft or seaplane.
Closed loop steam cooled airfoil
Widrig, Scott M.; Rudolph, Ronald J.; Wagner, Gregg P.
2006-04-18
An airfoil, a method of manufacturing an airfoil, and a system for cooling an airfoil is provided. The cooling system can be used with an airfoil located in the first stages of a combustion turbine within a combined cycle power generation plant and involves flowing closed loop steam through a pin array set within an airfoil. The airfoil can comprise a cavity having a cooling chamber bounded by an interior wall and an exterior wall so that steam can enter the cavity, pass through the pin array, and then return to the cavity to thereby cool the airfoil. The method of manufacturing an airfoil can include a type of lost wax investment casting process in which a pin array is cast into an airfoil to form a cooling chamber.
Turbine airfoil with ambient cooling system
Campbell, Jr, Christian X.; Marra, John J.; Marsh, Jan H.
2016-06-07
A turbine airfoil usable in a turbine engine and having at least one ambient air cooling system is disclosed. At least a portion of the cooling system may include one or more cooling channels configured to receive ambient air at about atmospheric pressure. The ambient air cooling system may have a tip static pressure to ambient pressure ratio of at least 0.5, and in at least one embodiment, may include a tip static pressure to ambient pressure ratio of between about 0.5 and about 3.0. The cooling system may also be configured such that an under root slot chamber in the root is large to minimize supply air velocity. One or more cooling channels of the ambient air cooling system may terminate at an outlet at the tip such that the outlet is aligned with inner surfaces forming the at least one cooling channel in the airfoil to facilitate high mass flow.
Lift-Enhancing Tabs on Multielement Airfoils
NASA Technical Reports Server (NTRS)
Ross, James C.; Storms, Bruce L.; Carrannanto, Paul G.
1995-01-01
The use of flat-plate tabs (similar to Gurney flaps) to enhance the lift of multielement airfoils is extended here by placing them on the pressure side and near the trailing edge of the main element rather than just on the furthest downstream wing element. The tabs studied range in height from 0.125 to 1.25% of the airfoil reference chord. In practice, such tabs would be retracted when the high-lift system is stowed. The effectiveness of the concept was demonstrated experimentally and computationally on a two-dimensional NACA 63(sub 2)-215 Mod B airfoil with a single-slotted, 30%-chord flap. Both the experiments and computations showed that the tabs significantly increase the lift at a given angle of attack and the maximum lift coefficient of the airfoil. The computational results showed that the increased lift was a result of additional turning of the flow by the tab that reduced or eliminated now separation on the flap. The best configuration tested, a 0.5%-chord tab placed 0.5% chord upstream of the trailing edge of the main element, increased the maximum lift coefficient of the airfoil by 12% and the maximum lift-to-drag ratio by 40%.
Airfoil noise in a uniform flow
NASA Astrophysics Data System (ADS)
Garcia, P.
An experimental analysis was made of the noise radiated by a NACA 0012 airfoil in a uniform flow in the CEPRA 19 anechoic wind tunnel. The investigations concerned the estimate of the radiated noise from existing theories developed in particular by Chandiramani, Chase and Howe. They required experimental characterization of the pressure field induced by the turbulent boundary layer in the trailing edge region of the airfoil. This work is original in that it allows the noise to be predicted from wave number spectrum measurements made using a sensor array. The prediction is not limited to low frequencies as is the case for computations using the measured integral scales of Corcos. This approach was also applied to airfoils at an incidence.
Turbine airfoil fabricated from tapered extrusions
Marra, John J
2013-07-16
An airfoil (30) and fabrication process for turbine blades with cooling channels (26). Tapered tubes (32A-32D) are bonded together in a parallel sequence, forming a leading edge (21), a trailing edge (22), and pressure and suction side walls (23, 24) connected by internal ribs (25). The tapered tubes may be extruded without camber to simplify the extrusion process, then bonded along matching surfaces (34), forming a non-cambered airfoil (28), which may be cambered in a hot forming process and cut (48) to length. The tubes may have tapered walls that are thinner at the blade tip (T1) than at the base (T2), reducing mass. A cap (50) may be attached to the blade tip. A mounting lug (58) may be forged (60) on the airfoil base and then machined, completing the blade for mounting in a turbine rotor disk.
Dong, R.; Chu, S.; Katz, J.
1997-07-01
Particle Image Velocimetry (PIV), pressure, and noise measurements are used to study the effect of modifications to tongue and impeller geometries on the flow structure and resulting noise in a centrifugal pump. It is demonstrated that the primary sources of noise are associated with interactions of the nonuniform outflux from the impeller (jet/wake phenomenon) with the tongue. Consequently, significant reduction of noise is achieved by increasing the gap between the tongue and the impeller up to about 20% of the impeller radius. Further increase in the gap affects the performance adversely with minimal impact on the noise level. When the gap is narrow, the primary sources of noise are impingement of the wake on the tip of the tongue, and tongue oscillations when the pressure difference across it is high. At about 20% gap, the entire wake and its associated vorticity trains miss the tongue, and the only (quite weak) effect of nonuniform outflux is the impingement of the jet on the tongue. An attempt is also made to reduce the nonuniformity in outflux from the impeller by inserting short vanes between the blades. They cause reduction in the size of the original wakes, but generate an additional jet/wake phenomenon of their own. Both wakes are weak to a level that their impacts on local pressure fluctuations and noise are insignificant. The only remaining major contributor to noise is tongue oscillations. This effect is shown to be dependent on the stiffness of the tongue.
Turbine airfoil with controlled area cooling arrangement
Liang, George
2010-04-27
A gas turbine airfoil (10) includes a serpentine cooling path (32) with a plurality of channels (34,42,44) fluidly interconnected by a plurality of turns (38,40) for cooling the airfoil wall material. A splitter component (50) is positioned within at least one of the channels to bifurcate the channel into a pressure-side channel (46) passing in between the outer wall (28) and the inner wall (30) of the pressure side (24) and a suction-side channel (48) passing in between the outer wall (28) and the inner wall (30) of the suction side (26) longitudinally downstream of an intermediate height (52). The cross-sectional area of the pressure-side channel (46) and suction-side channel (48) are thereby controlled in spite of an increasing cross-sectional area of the airfoil along its longitudinal length, ensuring a sufficiently high mach number to provide a desired degree of cooling throughout the entire length of the airfoil.
NASA supercritical airfoils: A matrix of family-related airfoils
NASA Technical Reports Server (NTRS)
Harris, Charles D.
1990-01-01
The NASA supercritical airfoil development program is summarized in a chronological fashion. Some of the airfoil design guidelines are discussed, and coordinates of a matrix of family related supercritical airfoils ranging from thicknesses of 2 to 18 percent and over a design lift coefficient range from 0 to 1.0 are presented.
NASA Technical Reports Server (NTRS)
Moracz, Donald J. (Inventor); Cook, Charles R. (Inventor); Toth, Istvan J. (Inventor)
1984-01-01
An improved method of making an airfoil includes stacking plies in two groups. A separator ply is positioned between the two groups of plies. The groups of plies and the separator ply are interconnected to form an airfoil blank. The airfoil blank is shaped, by forging or other methods, to have a desired configuration. The material of the separator ply is then dissolved or otherwise removed from between the two sections of the airfoil blank to provide access to the interior of the airfoil blank. Material is removed from inner sides of the two separated sections to form core receiving cavities. After cores have been placed in the cavities, the two sections of the airfoil blank are interconnected and the shaping of the airfoil is completed. The cores are subsequently removed from the completed airfoil.
NASA Technical Reports Server (NTRS)
Moracz, Donald J. (Inventor); Cook, Charles R. (Inventor); Toth, Istvan J. (Inventor)
1986-01-01
An improved method of making an airfoil includes stacking plies in two groups. A separator ply is positioned between the two groups of plies. The groups of plies and the separator ply are interconnected to form an airfoil blank. The airfoil blank is shaped, by forging or other methods, to have a desired configuration. The material of the separator ply is then dissolved or otherwise removed from between the two sections of the airfoil blank to provide access to the interior of the airfoil blank. Material is removed from inner sides of the two separated sections to form core receiving cavities. After cores have been placed in the cavities, the two sections of the airfoil blank are interconnected and the shaping of the airfoil is completed. The cores are subsequently removed from the completed airfoil.
Steady pressure measurements on an Aeroelastic Research Wing (ARW-2)
NASA Technical Reports Server (NTRS)
Sandford, Maynard C.; Seidel, David A.; Eckstrom, Clinton V.
1994-01-01
Transonic steady and unsteady pressure tests have been conducted in the Langley transonic dynamics tunnel on a large elastic wing known as the DAST ARW-2. The wing has a supercritical airfoil, an aspect ratio of 10.3, a leading-edge sweep back angle of 28.8 degrees, and two inboard and one outboard trailing-edge control surfaces. Only the outboard control surface was deflected to generate steady and unsteady flow over the wing during this study. Only the steady surface pressure, control-surface hinge moment, wing-tip deflection, and wing-root bending moment measurements are presented. The results from this elastic wing test are in tabulated form to assist in calibrating advanced computational fluid dynamics (CFD) algorithms.
Tonal noise production from a wall-mounted finite airfoil
NASA Astrophysics Data System (ADS)
Moreau, Danielle J.; Doolan, Con J.
2016-02-01
This study is concerned with the flow-induced noise of a smooth wall-mounted finite airfoil with flat ended tip and natural boundary layer transition. Far-field noise measurements have been taken at a single observer location and with a microphone array in the Virginia Tech Stability Wind Tunnel for a wall-mounted finite airfoil with aspect ratios of L / C = 1 - 3, at a range of Reynolds numbers (ReC = 7.9 ×105 - 1.6 ×106, based on chord) and geometric angles of attack (α = 0 - 6 °). At these Reynolds numbers, the wall-mounted finite airfoil produces a broadband noise contribution with a number of discrete equispaced tones at non-zero angles of attack. Spectral data are also presented for the noise produced due to three-dimensional vortex flow near the airfoil tip and wall junction to show the contributions of these flow features to airfoil noise generation. Tonal noise production is linked to the presence of a transitional flow state to the trailing edge and an accompanying region of mildly separated flow on the pressure surface. The separated flow region and tonal noise source location shift along the airfoil trailing edge towards the free-end region with increasing geometric angle of attack due to the influence of the tip flow field over the airfoil span. Tonal envelopes defining the operating conditions for tonal noise production from a wall-mounted finite airfoil are derived and show that the domain of tonal noise production differs significantly from that of a two-dimensional airfoil. Tonal noise production shifts to lower Reynolds numbers and higher geometric angles of attack as airfoil aspect ratio is reduced.
Tangler, J.L.; Somers, D.M.
2000-05-30
Airfoils for the tip and mid-span regions of a wind turbine blade have upper surface and lower surface shapes and contours between a leading edge and a trailing edge that minimize roughness effects of the airfoil and provide maximum lift coefficients that are largely insensitive to roughness effects. The airfoil in one embodiment is shaped and contoured to have a thickness in a range of about fourteen to seventeen percent, a Reynolds number in a range of about 1,500,000 to 2,000,000, and a maximum lift coefficient in a range of about 1.4 to 1.5. In another embodiment, the airfoil is shaped and contoured to have a thickness in a range of about fourteen percent to sixteen percent, a Reynolds number in a range of about 1,500,000 to 3,000,000, and a maximum lift coefficient in a range of about 0.7 to 1.5. Another embodiment of the airfoil is shaped and contoured to have a Reynolds in a range of about 1,500,000 to 4,000,000, and a maximum lift coefficient in a range of about 1.0 to 1.5.
Tangler, James L.; Somers, Dan M.
2000-01-01
Airfoils for the tip and mid-span regions of a wind turbine blade have upper surface and lower surface shapes and contours between a leading edge and a trailing edge that minimize roughness effects of the airfoil and provide maximum lift coefficients that are largely insensitive to roughness effects. The airfoil in one embodiment is shaped and contoured to have a thickness in a range of about fourteen to seventeen percent, a Reynolds number in a range of about 1,500,000 to 2,000,000, and a maximum lift coefficient in a range of about 1.4 to 1.5. In another embodiment, the airfoil is shaped and contoured to have a thickness in a range of about fourteen percent to sixteen percent, a Reynolds number in a range of about 1,500,000 to 3,000,000, and a maximum lift coefficient in a range of about 0.7 to 1.5. Another embodiment of the airfoil is shaped and contoured to have a Reynolds in a range of about 1,500,000 to 4,000,000, and a maximum lift coefficient in a range of about 1.0 to 1.5.
Second-order subsonic airfoil theory including edge effects
NASA Technical Reports Server (NTRS)
Van Dyke, Milton D
1956-01-01
Several recent advances in plane subsonic flow theory are combined into a unified second-order theory for airfoil sections of arbitrary shape. The solution is reached in three steps: the incompressible result is found by integration, it is converted into the corresponding subsonic compressible result by means of the second-order compressibility rule, and it is rendered uniformly valid near stagnation points by further rules. Solutions for a number of airfoils are given and are compared with the results of other theories and of experiment. A straight-forward computing scheme is outlined for calculating the surface velocities and pressures on any airfoil at any angle of attack
Low-speed single-element airfoil synthesis
NASA Technical Reports Server (NTRS)
Mcmasters, J. H.; Henderson, M. L.
1979-01-01
The use of recently developed airfoil analysis/design computational tools to clarify, enrich and extend the existing experimental data base on low-speed, single element airfoils is demonstrated. A discussion of the problem of tailoring an airfoil for a specific application at its appropriate Reynolds number is presented. This problem is approached by use of inverse (or synthesis) techniques, wherein a desirable set of boundary layer characteristics, performance objectives, and constraints are specified, which then leads to derivation of a corresponding viscous flow pressure distribution. Examples are presented which demonstrate the synthesis approach, following presentation of some historical information and background data which motivate the basic synthesis process.
Robust, Optimal Subsonic Airfoil Shapes
NASA Technical Reports Server (NTRS)
Rai, Man Mohan
2014-01-01
A method has been developed to create an airfoil robust enough to operate satisfactorily in different environments. This method determines a robust, optimal, subsonic airfoil shape, beginning with an arbitrary initial airfoil shape, and imposes the necessary constraints on the design. Also, this method is flexible and extendible to a larger class of requirements and changes in constraints imposed.
Unsteady aerodynamics of reverse flow dynamic stall on an oscillating blade section
NASA Astrophysics Data System (ADS)
Lind, Andrew H.; Jones, Anya R.
2016-07-01
Wind tunnel experiments were performed on a sinusoidally oscillating NACA 0012 blade section in reverse flow. Time-resolved particle image velocimetry and unsteady surface pressure measurements were used to characterize the evolution of reverse flow dynamic stall and its sensitivity to pitch and flow parameters. The effects of a sharp aerodynamic leading edge on the fundamental flow physics of reverse flow dynamic stall are explored in depth. Reynolds number was varied up to Re = 5 × 105, reduced frequency was varied up to k = 0.511, mean pitch angle was varied up to 15∘, and two pitch amplitudes of 5∘ and 10∘ were studied. It was found that reverse flow dynamic stall of the NACA 0012 airfoil is weakly sensitive to the Reynolds numbers tested due to flow separation at the sharp aerodynamic leading edge. Reduced frequency strongly affects the onset and persistence of dynamic stall vortices. The type of dynamic stall observed (i.e., number of vortex structures) increases with a decrease in reduced frequency and increase in maximum pitch angle. The characterization and parameter sensitivity of reverse flow dynamic stall given in the present work will enable the development of a physics-based analytical model of this unsteady aerodynamic phenomenon.
Grid generation by elliptic partial differential equations for a tri-element Augmentor-Wing airfoil
NASA Technical Reports Server (NTRS)
Sorenson, R. L.
1982-01-01
Two efforts to numerically simulate the flow about the Augmentor-Wing airfoil in the cruise configuration using the GRAPE elliptic partial differential equation grid generator algorithm are discussed. The Augmentor-Wing consists of a main airfoil with a slotted trailing edge for blowing and two smaller airfoils shrouding the blowing jet. The airfoil and the algorithm are described, and the application of GRAPE to an unsteady viscous flow simulation and a transonic full-potential approach is considered. The procedure involves dividing a complicated flow region into an arbitrary number of zones and ensuring continuity of grid lines, their slopes, and their point distributions across the zonal boundaries. The method for distributing the body-surface grid points is discussed.
Circulation methods in unsteady and three-dimensional flows
NASA Astrophysics Data System (ADS)
Yuan, Jiankun
The largely unstudied extension of ultrasonic circulation measurement techniques (UCMT) to determine instantaneous lift in unsteady and three-dimensional flows has been addressed in this work. A combined analytical-numerical-experimental approach was undertaken with the goal of developing methods to properly convert the measurable time-dependent bound circulation to instantaneous lift force in unsteady flows. The measurement of mean sectional lift distribution along structure spans in three-dimensional flows was also studied. An unsteady correction method for thin airfoils was developed analytically and validated numerically (with finite element solutions) to properly convert bound circulation to instantaneous lift based on unsteady potential flow theory. Results show that the unsteady correction method can provide increased accuracy for unsteady lift prediction over the Kutta-Joukowski method used in previous unsteady flow studies. The unsteady correction model generally should be included for instantaneous lift prediction as long as the bound circulation is time-dependent. Using the same framework, we also studied determination of instantaneous lift forces on stationary bluff bodies (circular cylinders) at low Reynolds number (Re = 100). Various force models, including an approximate vortex force model, were studied. A new unsteady model, similar to that developed for the thin airfoils, using instantaneous bound circulation values, was proposed. Another important issue studied in this thesis is the effect of acoustic path sensitivity on bound circulation determination, which we found to be crucial for accurately predicting the instantaneous lift in both unsteady flat plate and cylinder flows. Proper path selection should take into account the location of boundary layers, attached and shed vortices. These findings will be useful in future experimental design of UCMT, PIV and LDV methods. Finally, we used the UCMT method to experimentally study the mean spatial
NASA Technical Reports Server (NTRS)
Carta, F. O.
1982-01-01
Tests were conducted on a linear cascade of airfoils oscillating in pitch to measure the unsteady pressure response on selected blades along the leading edge plane of the cascade, over the chord of the center blade, and on the sidewall in the plane of the leading edge. The pressure data were reduced to Fourier coefficient form for direct comparison, and were also processed to yield integrated loads and, particularly, the aerodynamic damping coefficient. Results from the unsteady Verdon/Caspar theory for cascaded blades with nonzero thickness and camber were compared with the experimental measurements. The three primary results are: (1) from the leading edge plane blade data, the cascade was judged to be periodic in unsteady flow over the range of parameters tested; (2) the interblade phase angle was found to be the single most important parameter affecting the stability of the oscillating cascade blades; and (3) the real blade theory and the experiment were in excellent agreement for the several cases chosen for comparison.
Predicting Unsteady Aeroelastic Behavior
NASA Technical Reports Server (NTRS)
Strganac, Thomas W.; Mook, Dean T.
1990-01-01
New method for predicting subsonic flutter, static deflections, and aeroelastic divergence developed. Unsteady aerodynamic loads determined by unsteady-vortex-lattice method. Accounts for aspect ratio and angle of attack. Equations for motion of wing and flow field solved iteratively and simultaneously. Used to predict transient responses to initial disturbances, and to predict steady-state static and oscillatory responses. Potential application for research in such unsteady structural/flow interactions as those in windmills, turbines, and compressors.
NASA Technical Reports Server (NTRS)
Turner, Travis L. (Inventor); Khorrami, Mehdi R. (Inventor); Lockard, David P. (Inventor); McKenney, Martin J. (Inventor); Atherley, Raymond D. (Inventor); Kidd, Reggie T. (Inventor)
2014-01-01
A multi-element airfoil system includes an airfoil element having a leading edge region and a skin element coupled to the airfoil element. A slat deployment system is coupled to the slat and the skin element, and is capable of deploying and retracting the slat and the skin element. The skin element substantially fills the lateral gap formed between the slat and the airfoil element when the slat is deployed. The system further includes an uncoupling device and a sensor to remove the skin element from the gap based on a critical angle-of-attack of the airfoil element. The system can alternatively comprise a trailing edge flap, where a skin element substantially fills the lateral gap between the flap and the trailing edge region of the airfoil element. In each case, the skin element fills a gap between the airfoil element and the deployed flap or slat to reduce airframe noise.
Unsteady lifting-line theory with applications
NASA Technical Reports Server (NTRS)
Ahmadi, A. R.; Widnall, S. E.
1982-01-01
Unsteady lifting-line theory is developed for a flexible unswept wing of large aspect ratio oscillating at low frequency in inviscid incompressible flow. The theory is formulated in terms of the acceleration potential and treated by the method of matched asymptotic expansions. The wing displacements are prescribed and the pressure field, airloads, and unsteady induced downwash are obtained in closed form. Sample numerical calculations are presented. The present work identifies and resolves errors in the unsteady lifting-line theory of James and points out a limitation in that of Van Holten. Comparison of the results of Reissner's approximate unsteady lifting-surface theory with those of the present work shows favorable agreement. The present work thus provides some formal justification for Reissner's ad hoc theory. For engineering purposes, the region of applicability of the theory in the reduced frequency-aspect ratio domain is identified approximately and found to cover most cases of practical interest.
Shock unsteadiness creation and propagation: experimental analysis
NASA Astrophysics Data System (ADS)
Benay, R.; Alaphilippe, M.; Severac, N.
2012-09-01
The possibility of creating unsteady distortions of the tip shock by waves emitted from an aircraft is assessed experimentally. The model chosen is a cylindrical fore body equipped with a spike. This configuration is known for generating an important level of unsteadiness around the spike in supersonic regime. The wind tunnel Mach number is equal to 2. The experiments show that waves emitted from this source propagate along the tip shock and interact with it. It is then assessed that this interaction produces a periodic distortion of the shock that propagates to the external flow. Unsteady pressure sensors, high speed schlieren films, hot wire probing and laser Doppler velocimetry are used as complementary experimental means. The final result is a coherent representation of the complex mechanism of wave propagation that has been evidenced. The principle of creating unsteady shock deformation by onboard equipments could be examined as a possibly promising method of sonic boom control.
Closed-Loop Aerodynamic Flow Control of a Maneuvering Airfoil
NASA Astrophysics Data System (ADS)
Brzozowski, Daniel P.; Culp, John R.; Glezer, Ari
2011-11-01
The unsteady interaction between trailing edge aerodynamic flow control and airfoil motion in pitch and plunge is investigated in wind tunnel experiments using a 2-DOF traverse which enables application of time-dependent external torque and forces by servo motors. The global aerodynamic forces and moments are regulated by controlling vorticity generation and accumulation near the surface using hybrid synthetic jet actuators. The dynamic coupling between the actuation and the time-dependent flow field is characterized using simultaneous force and velocity measurements that are taken phase-locked to the commanded actuation waveform. The effect of the unsteady motion on the model-embedded flow control is assessed in unsteady several maneuvers. Circulation time history that is estimated from a PIV wake survey shows that the entire flow over the airfoil readjusts within about 1.5 TCONV, which is about two orders of magnitude shorter than the characteristic time associated with the controlled maneuver of the wind tunnel model. This illustrates that flow-control actuation can be typically effected on time scales that are commensurate with the flow's convective time scale, and that the maneuver response is primarily limited by the inertia of the platform.
Dynamic Stall Characteristics of Drooped Leading Edge Airfoils
NASA Technical Reports Server (NTRS)
Sankar, Lakshmi N.; Sahin, Mehmet; Gopal, Naveen
2000-01-01
Helicopters in high-speed forward flight usually experience large regions of dynamic stall over the retreating side of the rotor disk. The rapid variations in the lift and pitching moments associated with the stall process can result in vibratory loads, and can cause fatigue and failure of pitch links. In some instances, the large time lag between the aerodynamic forces and the blade motion can trigger stall flutter. A number of techniques for the alleviation of dynamic stall have been proposed and studied by researchers. Passive and active control techniques have both been explored. Passive techniques include the use of high solidity rotors that reduce the lift coefficients of individual blades, leading edge slots and leading edge slats. Active control techniques include steady and unsteady blowing, and dynamically deformable leading edge (DDLE) airfoils. Considerable amount of experimental and numerical data has been collected on the effectiveness of these concepts. One concept that has not received as much attention is the drooped-leading edge airfoil idea. It has been observed in wind tunnel studies and flight tests that drooped leading edge airfoils can have a milder dynamic stall, with a significantly milder load hysteresis. Drooped leading edge airfoils may not, however, be suitable at other conditions, e.g. in hover, or in transonic flow. Work needs to be done on the analysis and design of drooped leading edge airfoils for efficient operation in a variety of flight regimes (hover, dynamic stall, and transonic flow). One concept that is worthy of investigation is the dynamically drooping airfoil, where the leading edge shape is changed roughly once-per-rev to mitigate the dynamic stall.
Aerodynamic Simulation of Ice Accretion on Airfoils
NASA Technical Reports Server (NTRS)
Broeren, Andy P.; Addy, Harold E., Jr.; Bragg, Michael B.; Busch, Greg T.; Montreuil, Emmanuel
2011-01-01
This report describes recent improvements in aerodynamic scaling and simulation of ice accretion on airfoils. Ice accretions were classified into four types on the basis of aerodynamic effects: roughness, horn, streamwise, and spanwise ridge. The NASA Icing Research Tunnel (IRT) was used to generate ice accretions within these four types using both subscale and full-scale models. Large-scale, pressurized windtunnel testing was performed using a 72-in.- (1.83-m-) chord, NACA 23012 airfoil model with high-fidelity, three-dimensional castings of the IRT ice accretions. Performance data were recorded over Reynolds numbers from 4.5 x 10(exp 6) to 15.9 x 10(exp 6) and Mach numbers from 0.10 to 0.28. Lower fidelity ice-accretion simulation methods were developed and tested on an 18-in.- (0.46-m-) chord NACA 23012 airfoil model in a small-scale wind tunnel at a lower Reynolds number. The aerodynamic accuracy of the lower fidelity, subscale ice simulations was validated against the full-scale results for a factor of 4 reduction in model scale and a factor of 8 reduction in Reynolds number. This research has defined the level of geometric fidelity required for artificial ice shapes to yield aerodynamic performance results to within a known level of uncertainty and has culminated in a proposed methodology for subscale iced-airfoil aerodynamic simulation.
NASA Technical Reports Server (NTRS)
1997-01-01
10' NACA 0015 with 30% chord trailing edge flap deflected 20 degrees. Used in 0.3 Meter Transonic Cryogenic Tunnel, this airfoil has a 0.44 mm slot at 70% chord. Oscillatory blowing out of slot used for separation control. Howard Price appears in side view shot, in building 1145, Studio.
Turbine airfoil manufacturing technology
Kortovich, C.
1995-12-31
The specific goal of this program is to define manufacturing methods that will allow single crystal technology to be applied to complex-cored airfoils components for power generation applications. Tasks addressed include: alloy melt practice to reduce the sulfur content; improvement of casting process; core materials design; and grain orientation control.
Numerical analysis of bio-inspired corrugated airfoil at low Reynolds number
NASA Astrophysics Data System (ADS)
Mondal, Partha Protim; Rahman, Md. Masudur; Hasan, A. B. M. Toufique
2016-07-01
A numerical study was conducted to investigate the aerodynamic performance of a bio-inspired corrugated airfoil at the chord Reynolds number of Rec=80,000 to explore the potential advantages of such airfoils at low Reynolds numbers. This study represents the transient nature of corrugated airfoils at low Reynolds number where flow is assumed to be laminar, unsteady, incompressible and two dimensional. The simulations include a sharp interface Cartesian grid based meshing employed with laminar viscous model. The flow field surrounding the corrugated airfoil has been analyzed using structured grid Finite Volume Method (FVM) based on Navier-Stokes equation. All parameters used in flow simulation are expressed in non-dimensional quantities for better understanding of flow behavior, regardless of dimensions or the fluid that is used. The simulated results revealed that the corrugated airfoil provides high lift with moderate drag and prevents large scale flow separation at higher angles of attack. This happens due to the negative shear drag produced by the recirculation zones which occurs in the valleys of the corrugated airfoils. The existence of small circulation bubbles sitting in the valleys prevents large scale flow separation thus increasing the aerodynamic performance of the corrugated airfoil.
A critical evaluation of the predictions of the NASA-Lockheed multielement airfoil computer program
NASA Technical Reports Server (NTRS)
Brune, G. W.; Manke, J. W.
1978-01-01
Theoretical predictions of several versions of the multielement airfoil computer program are evaluated. The computed results are compared with experimental high lift data of general aviation airfoils with a single trailing edge flap, and of airfoils with a leading edge flap and double slotted trailing edge flaps. Theoretical and experimental data include lift, pitching moment, profile drag and surface pressure distributions, boundary layer integral parameters, skin friction coefficients, and velocity profiles.
Control of boundary layer separation and the wake of an airfoil using ns-DBD plasma actuators
NASA Astrophysics Data System (ADS)
Ashcraft, Timothy
The efficacy of nanosecond pulse driven dielectric barrier discharge (ns-DBD) plasma actuators for boundary layer separation and wake control is investigated experimentally. A single ns-DBD plasma actuator is placed at the leading edge of a NACA 0012 airfoil model. Both baseline and controlled flow fields are studied using static pressure measurements, Particle Image Velocimetry (PIV) and Constant Temperature Anemometry (CTA). Experiments are primarily performed at Re = 0.74 x 106 and alpha = 18°. CP, PIV and CTA data show that a forcing frequency of F+ = 1.14 is optimal for separation control. CTA surveys of the wake at x/c = 7 indicate three approximate regimes of behavior. Forcing in the range 0.92< F+ < 1.52 results in the best conditions for separation control over the airfoil, but has no dominant signature in the wake at x/c = 7. Excitation in the range of 0.23 < F+ < 0.92 produces a single dominant frequency in the wake while F+ < 0.23 shows behavior representing a possible impulse response or nonlinear effects. PIV data confirm these observations in all three regimes. Cross-correlations of CTA data are also employed to evaluate the two-dimensionality of the excited wake. The initial results presented here are part of an ongoing effort to use active flow control (AFC), in the form of ns-DBDs, as an enabling technology for the study of unsteady aerodynamics and vortex-body interactions.
The Ultimate Flow Controlled Wind Turbine Blade Airfoil
NASA Astrophysics Data System (ADS)
Seifert, Avraham; Dolgopyat, Danny; Friedland, Ori; Shig, Lior
2015-11-01
Active flow control is being studied as an enabling technology to enhance and maintain high efficiency of wind turbine blades also with contaminated surface and unsteady winds as well as at off-design operating conditions. The study is focused on a 25% thick airfoil (DU91-W2-250) suitable for the mid blade radius location. Initially a clean airfoil was fabricated and tested, as well as compared to XFoil predictions. From these experiments, the evolution of the separation location was identified. Five locations for installing active flow control actuators are available on this airfoil. It uses both Piezo fluidic (``Synthetic jets'') and the Suction and Oscillatory Blowing (SaOB) actuators. Then we evaluate both actuation concepts overall energy efficiency and efficacy in controlling boundary layer separation. Since efficient actuation is to be found at low amplitudes when placed close to separation location, distributed actuation is used. Following the completion of the baseline studies the study has focused on the airfoil instrumentation and extensive wind tunnel testing over a Reynolds number range of 0.2 to 1.5 Million. Sample results will be presented and outline for continued study will be discussed.
Stability of Inviscid Flow over Airfoils Admitting Multiple Numerical Solutions
NASA Astrophysics Data System (ADS)
Liu, Ya; Xiong, Juntao; Liu, Feng; Luo, Shijun
2012-11-01
Multiple numerical solutions at the same flight condition are found of inviscid transonic flow over certain airfoils (Jameson et al., AIAA 2011-3509) within some Mach number range. Both symmetric and asymmetric solutions exist for a symmetric airfoil at zero angle of attack. Global linear stability analysis of the multiple solutions is conducted. Linear perturbation equations of the Euler equations around a steady-state solution are formed and discretized numerically. An eigenvalue problem is then constructed using the modal analysis approach. Only a small portion of the eigen spectrum is needed and thus can be found efficiently by using Arnoldi's algorithm. The least stable or unstable mode corresponds to the eigenvalue with the largest real part. Analysis of the NACA 0012 airfoil indicates stability of symmetric solutions of the Euler equations at conditions where buffet is found from unsteady Navier-Stokes equations. Euler solutions of the same airfoil but modified to include the displacement thickness of the boundary layer computed from the Navier-Stokes equations, however, exhibit instability based on the present linear stability analysis. Graduate Student.
NASA Technical Reports Server (NTRS)
Carta, F. O.
1981-01-01
Tests were conducted a linear cascade of airfoils oscillating in pitch to measure the unsteady pressure response on selected blade along the leading edge plane of the cascade, over the chord of the center blade, and on the sidewall in the plane of the leading edge. The tests were conducted for all 96 combinations 2 mean camberline incidence angles 2 pitching amplitudes 3 reduced frequencies and 8 interblade phase angles. The pressure data were reduced to Fourier coefficient form for direct comparison, and were also processed to yield integrated loads and particularly, the aerodynamic damping coefficient. Data obtained during the test program, reproduced from the printout of the data reduction program are complied. A further description of the contents of this report is found in the text that follows.
Experimental investigation of the flowfield of an oscillating airfoil
NASA Technical Reports Server (NTRS)
Panda, J.; Zaman, K. B. M. Q.
1992-01-01
The flowfield of an airfoil oscillated periodically over a wide range of reduced frequencies, 0 less than or = k less than or = 1.6 is studied experimentally at chord Reynolds numbers of R sub c = 22,000 and 44,000. The NACA0012 airfoil is pitched sinusoidally about one quarter chord between angles of attack (alpha) of 5 and 25 degrees. Detailed flow visualization and phase averaged vorticity measurements are carried out for k = 0.2 to document the evolution and the shedding of the dynamic stall vortex (DSV). In addition to the DSV, an intense vortex of opposite sign originates from the trailing edge just when the DSV is shed. After being shed into the wake, the two together take the shape of a large 'mushroom' while being convected away from the airfoil. The unsteady circulation around the airfoil and, therefore, the time varying component of the lift is estimated in a novel way from the shed vorticity flux and is found to be in good agreement with the lift variation reported by others. The delay in the shedding of the DSV with increasing k, as observed by previous researchers, is documented for the full range of k. The DSV, for example, is shed nearly at the maximum alpha of 25 degrees at k = 0.2, but is shed at the minimum alpha of 5 degrees at k = 0.8. At low k, the flowfield appears quasi-steady and the bluff body shedding corresponding to the maximum alpha (25 degrees) dominates the unsteady fluctuations in the wake.
Experimental investigation of the flowfield of an oscillating airfoil
NASA Technical Reports Server (NTRS)
Panda, J.; Zaman, K. B. M. Q.
1992-01-01
The flow field of an airfoil oscillated periodically over a wide range of reduced frequencies, 0 less than k less than 1.6, is studied experimentally at chord Reynolds numbers of R sub c = 22,000 and 44,000. The NACA0012 airfoil is pitched sinusoidally about one quarter chord between alpha of 5 deg and 25 deg. Detailed flow visualization and phase averaged vorticity measurements are carried out for k = 0.2 to document the evolution and the shedding of the dynamic stall vortex (DSV). In addition to the DSV, an intense vortex of opposite sign originates from the trailing edge just when the DSV is shed. After being shed into the wake, the two together take the shape of a large 'mushroom' while being convected away from the airfoil. The unsteady circulation around the airfoil and, therefore, the time varying component of the lift is estimated in a novel way from the shed vorticity flux and is found to be in good agreement with the lift variation reported by others. The delay in the shedding of the DSV with increasing k, as observed by previous researchers, is documented for the full range of k. The DSV, for example, is shed nearly at the maximum alpha of 25 deg at k = 0.2, but is shed at the minimum alpha of 5 deg at k = 0.8. At low k, the flowfield appears quasi-steady and the bluff body shedding corresponding to the maximum alpha (25 deg) dominates the unsteady fluctuations in the wake.
Rotor-generated unsteady aerodynamic interactions in a 1½ stage compressor
NASA Astrophysics Data System (ADS)
Papalia, John J.
Because High Cycle Fatigue (HCF) remains the predominant surprise failure mode in gas turbine engines, HCF avoidance design systems are utilized to identify possible failures early in the engine development process. A key requirement of these analyses is accurate determination of the aerodynamic forcing function and corresponding airfoil unsteady response. The current study expands the limited experimental database of blade row interactions necessary for calibration of predictive HCF analyses, with transonic axial-flow compressors of particular interest due to the presence of rotor leading edge shocks. The majority of HCF failures in aircraft engines occur at off-design operating conditions. Therefore, experiments focused on rotor-IGV interactions at off-design are conducted in the Purdue Transonic Research Compressor. The rotor-generated IGV unsteady aerodynamics are quantified when the IGV reset angle causes the vane trailing edge to be nearly aligned with the rotor leading edge shocks. A significant vane response to the impulsive static pressure perturbation associated with a shock is evident in the point measurements at 90% span, with details of this complex interaction revealed in the corresponding time-variant vane-to-vane flow field data. Industry wide implementation of Controlled Diffusion Airfoils (CDA) in modern compressors motivated an investigation of upstream propagating CDA rotor-generated forcing functions. Whole field velocity measurements in the reconfigured Purdue Transonic Research Compressor along the design speedline reveal steady loading had a considerable effect on the rotor shock structure. A detached rotor leading edge shock exists at low loading, with an attached leading edge and mid-chord suction surface normal shock present at nominal loading. These CDA forcing functions are 3--4 times smaller than those generated by the baseline NACA 65 rotor at their respective operating points. However, the IGV unsteady aerodynamic response to the CDA
Vertical axis wind turbine airfoil
Krivcov, Vladimir; Krivospitski, Vladimir; Maksimov, Vasili; Halstead, Richard; Grahov, Jurij Vasiljevich
2012-12-18
A vertical axis wind turbine airfoil is described. The wind turbine airfoil can include a leading edge, a trailing edge, an upper curved surface, a lower curved surface, and a centerline running between the upper surface and the lower surface and from the leading edge to the trailing edge. The airfoil can be configured so that the distance between the centerline and the upper surface is the same as the distance between the centerline and the lower surface at all points along the length of the airfoil. A plurality of such airfoils can be included in a vertical axis wind turbine. These airfoils can be vertically disposed and can rotate about a vertical axis.
User's manual for airfoil flow field computer code SRAIR
NASA Technical Reports Server (NTRS)
Shamroth, S. J.
1985-01-01
A two dimensional unsteady Navier-Stokes calculation procedure with specific application to the isolated airfoil problem is presented. The procedure solves the full, ensemble averaged Navier-Stokes equations with turbulence represented by a mixing length model. The equations are solved in a general nonorthogonal coordinate system which is obtained via an external source. Specific Cartesian locations of grid points are required as input for this code. The method of solution is based upon the Briley-McDonald LBI procedure. The manual discusses the analysis, flow of the program, control steam, input and output.
An Exploratory Investigation of a Slotted, Natural-Laminar-Flow Airfoil
NASA Technical Reports Server (NTRS)
Somers, Dan M.
2012-01-01
A 15-percent-thick, slotted, natural-laminar-flow (SNLF) airfoil, the S103, for general aviation applications has been designed and analyzed theoretically and verified experimentally in the Langley Low-Turbulence Pressure Tunnel. The two primary objectives of high maximum lift and low profile drag have been achieved. The constraints on the pitching moment and the airfoil thickness have been satisfied. The airfoil exhibits a rapid stall, which does not meet the design goal. Comparisons of the theoretical and experimental results show good agreement. Comparison with the baseline, NASA NLF(1)-0215F airfoil confirms the achievement of the objectives.
An investigation of the aerodynamic characteristics of a new general aviation airfoil in flight
NASA Technical Reports Server (NTRS)
Gregorek, G. M.; Hoffmann, M. J.; Weislogel, G. S.
1982-01-01
A low speed airfoil, the GA(W)-2, - a 13% thickness to chord ratio airfoil was evaluated. The wing of a Beech Sundowner was modified at by adding balsa ribs and covered with aluminum skin, to alter the existing airfoil shape to that of the GA(W)-2 airfoil. The aircraft was flown in a flight test program that gathered wing surface pressures and wake data from which the lift drag, and pitching moment of the airfoil could be determined. After the base line performance of the airfoil was measured, the drag due to surface irregularities such as steps, rivets and surface waviness was determined. The potential reduction of drag through the use of surface coatings such as KAPTON was also investigated.
Airfoil-Shaped Fluid Flow Tool for Use in Making Differential Measurements
NASA Technical Reports Server (NTRS)
England, John Dwight (Inventor); Kelley, Anthony R. (Inventor); Cronise, Raymond J. (Inventor)
2014-01-01
A fluid flow tool includes an airfoil structure and a support arm. The airfoil structure's high-pressure side and low-pressure side are positioned in a conduit by the support arm coupled to the conduit. The high-pressure and low-pressure sides substantially face opposing walls of the conduit. At least one measurement port is formed in the airfoil structure at each of its high-pressure side and low-pressure side. A first manifold, formed in the airfoil structure and in fluid communication with each measurement port so-formed at the high-pressure side, extends through the airfoil structure and support arm to terminate and be accessible at the exterior wall of the conduit. A second manifold, formed in the airfoil structure and in fluid communication with each measurement port so-formed at the low-pressure side, extends through the airfoil structure and support arm to terminate and be accessible at the exterior wall of the conduit.
Turbulence modelling for unsteady separated flows
NASA Technical Reports Server (NTRS)
Bradshaw, Peter
1992-01-01
The exact transport equations for turbulent (Reynolds) stresses have left-hand sides representing the 'substantial derivatives' of the Reynolds stresses, i.e., the rates of change of stress with respect to time, as seen by an observer following the mean motion of the fluid. Here the 'mean' is a statistical average for the turbulent motion, distinguished from the ordered unsteadiness on which it is superimposed: for a turbomachine blade or a cyclically-pitching airfoil, the mean is a phase average. Written in coordinates fixed with respect to a solid surface, the substantial derivative appears partly as an Eulerian time derivative at given spatial coordinate position and partly as a spatial derivative. Separation presents two specific problems to a turbulence model: (1) prediction of the flow near separation depends critically on the 'near-wall' part of the turbulence model, and (2) downstream of separation, a boundary layer changes gradually to a mixing layer.
NASA Technical Reports Server (NTRS)
Ott, Eric A.
2005-01-01
Scoping of shape changing airfoil concepts including both aerodynamic analysis and materials-related technology assessment effort was performed. Three general categories of potential components were considered-fan blades, booster and compressor blades, and stator airfoils. Based on perceived contributions to improving engine efficiency, the fan blade was chosen as the primary application for a more detailed assessment. A high-level aerodynamic assessment using a GE90-90B Block 4 engine cycle and fan blade geometry indicates that blade camber changes of approximately +/-4deg would be sufficient to result in fan efficiency improvements nearing 1 percent. Constraints related to flight safety and failed mode operation suggest that use of the baseline blade shape with actuation to the optimum cruise condition during a portion of the cycle would be likely required. Application of these conditions to the QAT fan blade and engine cycle was estimated to result in an overall fan efficiency gain of 0.4 percent.
Slat tonal noise mechanisms in a two-dimensional multi-element airfoil configuration
NASA Astrophysics Data System (ADS)
Olson, Susan Denise
Airframe noise from commercial aircraft rivals engine noise in the landing phase of flight. Unsteady flow around the leading edge slat of the high-lift system contributes significantly to airframe noise. This research experimentally investigated tonal noise sources arising from the unsteady flow around the leading edge slat from a flow physics standpoint. Using this approach, efficient acoustic control strategies naturally follow. This study resolved two types of acoustic tones, trailing edge vortex shedding noise and acoustic resonance, originating from the slat component of a two-dimensional, three-element airfoil model. The latter noise mechanism had not been documented until the current study. Sharpening the slat trailing edge reduced the overall sound pressure level from the shedding noise by 12dB. The reduction in radiated acoustic pressure was directly proportional to the reduction in the strength of the vortex shedding. The acoustic power was conceptually related to the strength of the global instability in the wake. The onset of the slat aeroacoustic resonance corresponded to the formation of a separation bubble on the upper surface of the slat. The peak acoustic frequency matched the frequency of large-scale structures ejected from the separation bubble. A feedback loop between the near-wake region and the separation point sustained the resonance. A model based on a phase matching condition between the acoustic waves and the generation of the separated shear layer instability waves successfully predicted the acoustic frequencies. The acoustic tone followed the modal variation closest to the subharmonic of the natural instability frequency, which caused the discontinuous jumps in frequency with increasing velocity.
Research on unsteady transonic flow theory
NASA Technical Reports Server (NTRS)
Revell, J. D.
1973-01-01
A two-dimensional theory is considered for the unsteady flow disturbances caused by aeroelastic deformations of a thick wing at high subsonic freestream Mach numbers, having a single, internally embedded supercritical (locally supersonic) steady flow region adjacent to the low pressure side of the wing. The theory develops a matrix of unsteady aerodynamic influence coefficients (AICs) suitable as a strip theory for aeroelastic analysis of large aspect ratio thick wings of moderate sweep, typical of a wide class of current and future aircraft. The theory derives the linearized unsteady flow solutions separately for both the subcritical and supercritical regions. These solutions are coupled together to give the requisite (wing pressure-downwash) AICs by the intermediate step of defining flow disturbances on the sonic line, and at the shock wave; these intermediate quantities are then algebraically eliminated by expressing them in terms of the wing surface downwash.
Numerical simulations of unsteady flows in turbomachines
NASA Astrophysics Data System (ADS)
Dorney, Daniel Joseph
The performance of axial and centrifugal turbomachines is significantly affected by the presence of unsteady and viscous flow mechanisms. Most contemporary design systems, however, use steady or linearized unsteady inviscid flow analyses to generate new blade shapes. In an effort to increase the understanding of unsteady viscous flows in turbomachinery blade rows, and to determine the limitations of linearized inviscid flow analyses, a two-part investigation was conducted. In the first portion of this investigation, a nonlinear viscous flow analysis was developed for the prediction of unsteady flows in two dimensional axial turbomachinery blade rows. The boundary conditions were formulated to allow the specification of vortical, entropic and acoustic excitations at the inlet, and acoustic excitations at exit, of a cascade. Numerical simulations were performed for flat plate and compressor exit guide vane cascades, and the predicted results were compared with solutions from classical linearized theory and linearized inviscid flow analysis. The unsteady pressure fields predicted with the current analysis showed close agreement with the linearized solutions for low to moderate temporal frequency vortical and acoustic excitations. As the temporal frequency of the excitations was increased, nonlinear effects caused discrepancies to develop between the linearized and Navier-Stokes solution sets. The inclusion of viscosity had a significant impact on the unsteady vorticity field, but only a minimal effect on the unsteady pressure field. In the second part of this investigation, a quasi-three-dimensional Navier-Stokes analysis was modified and applied to flows in centrifugal turbomachinery blade rows. Inviscid and viscous flow simulations were performed for a centrifugal impeller at three operating conditions. By comparing the predicted and experimental circumferential distributions of the relative frame velocity and flow angle downstream of the impeller, it was
NASA Technical Reports Server (NTRS)
Fromme, J.; Golberg, M.; Werth, J.
1979-01-01
The numerical computation of unsteady airloads acting upon thin airfoils with multiple leading and trailing-edge controls in two-dimensional ventilated subsonic wind tunnels is studied. The foundation of the computational method is strengthened with a new and more powerful mathematical existence and convergence theory for solving Cauchy singular integral equations of the first kind, and the method of convergence acceleration by extrapolation to the limit is introduced to analyze airfoils with flaps. New results are presented for steady and unsteady flow, including the effect of acoustic resonance between ventilated wind-tunnel walls and airfoils with oscillating flaps. The computer program TWODI is available for general use and a complete set of instructions is provided.
LES tests on airfoil trailing edge serration
NASA Astrophysics Data System (ADS)
Zhu, Wei Jun; Shen, Wen Zhong
2016-09-01
In the present study, a large number of acoustic simulations are carried out for a low noise airfoil with different Trailing Edge Serrations (TES). The Ffowcs Williams-Hawkings (FWH) acoustic analogy is used for noise prediction at trailing edge. The acoustic solver is running on the platform of our in-house incompressible flow solver EllipSys3D. The flow solution is first obtained from the Large Eddy Simulation (LES), the acoustic part is then carried out based on the instantaneous hydrodynamic pressure and velocity field. To obtain the time history data of sound pressure, the flow quantities are integrated around the airfoil surface through the FWH approach. For all the simulations, the chord based Reynolds number is around 1.5x106. In the test matrix, the effects from angle of attack, the TE flap angle, the length/width of the TES are investigated. Even though the airfoil under investigation is already optimized for low noise emission, most numerical simulations and wind tunnel experiments show that the noise level is further decreased by adding the TES device.
Airfoil-shaped micro-mixers for reducing fouling on membrane surfaces
Ho, Clifford K; Altman, Susan J; Clem, Paul G; Hibbs, Michael; Cook, Adam W
2012-10-23
An array of airfoil-shaped micro-mixers that enhances fluid mixing within permeable membrane channels, such as used in reverse-osmosis filtration units, while minimizing additional pressure drop. The enhanced mixing reduces fouling of the membrane surfaces. The airfoil-shaped micro-mixer can also be coated with or comprised of biofouling-resistant (biocidal/germicidal) ingredients.
Application of numerical optimization to the design of low speed airfoils
NASA Technical Reports Server (NTRS)
Hicks, R. M.; Vanderplaats, G. N.
1975-01-01
A practical procedure for the optimum design of low-speed airfoils is demonstrated. The procedure uses an optimization program based on the method of feasible directions coupled with an aerodynamic analysis program that uses a relaxation solution of the inviscid, full potential equation. Results are presented for airfoils designed to have small adverse pressure gradients, high maximum lift, and low pitching moment.
NREL airfoil families for HAWTs
NASA Astrophysics Data System (ADS)
Tangler, J. L.; Somers, D. M.
1995-01-01
The development of special-purpose airfoils for horizontal-axis wind turbines (HAWTs) began in 1984 as a joint effort between the National Renewable Energy Laboratory (NREL), formerly the Solar Energy Research Institute (SERI), and Airfoils, Incorporated. Since that time seven airfoil families have been designed for various size rotors using the Eppler Airfoil Design and Analysis Code. A general performance requirement of the new airfoil families is that they exhibit a maximum lift coefficient (c(sub l,max)) which is relatively insensitive to roughness effects. The airfoil families address the needs of stall-regulated, variable-pitch, and variable-rpm wind turbines. For stall-regulated rotors, better peak-power control is achieved through the design of tip airfoils that restrain the maximum lift coefficient. Restrained maximum lift coefficient allows the use of more swept disc area for a given generator size. Also, for stall-regulated rotors, tip airfoils with high thickness are used to accommodate overspeed control devices. For variable-pitch and variable-rpm rotors, tip airfoils having a high maximum lift coefficient lend themselves to lightweight blades with low solidity. Tip airfoils having low thickness result in less drag for blades having full-span pitch control. Annual energy improvements from the NREL airfoil families are projected to be 23% to 35% for stall-regulated turbines, 8% to 20% for variable-pitch turbines, and 8% to 10% for variable-rpm turbines. The improvement for stall-regulated turbines has been verified in field tests.
NREL airfoil families for HAWTs
Tangler, J L; Somers, D M
1995-01-01
The development of special-purpose airfoils for horizontal-axis wind turbines (HAWTs) began in 1984 as a joint effort between the National Renewable Energy Laboratory (NREL), formerly the Solar Energy Research Institute (SERI), and Airfoils, Incorporated. Since that time seven airfoil families have been designed for various size rotors using the Eppler Airfoil Design and Analysis Code. A general performance requirement of the new airfoil families is that they exhibit a maximum lift coefficient (c{sub l,max}) which is relatively insensitive to roughness effects. The airfoil families address the needs of stall-regulated, variable-pitch, and variable-rpm wind turbines. For stall-regulated rotors, better peak-power control is achieved through the design of tip airfoils that restrain the maximum lift coefficient. Restrained maximum lift coefficient allows the use of more swept disc area for a given generator size. Also, for stall-regulated rotors, tip airfoils with high thickness are used to accommodate overspeed control devices. For variable-pitch and variable-rpm rotors, tip airfoils having a high maximum lift coefficient lend themselves to lightweight blades with low solidity. Tip airfoils having low thickness result in less drag for blades having full-span pitch control. Annual energy improvements from the NREL airfoil families are projected to be 23% to 35% for stall-regulated turbines, 8% to 20% for variable-pitch turbines, and 8% to 10% for variable-rpm turbines. The improvement for stall-regulated turbines has been verified in field tests.
Suppression of dynamic stall with a leading-edge slat on a VR-7 airfoil
NASA Technical Reports Server (NTRS)
Mcalister, K. W.; Tung, C.
1993-01-01
The VR-7 airfoil was experimentally studied with and without a leading-edge slat at fixed angles of attack from 0 deg to 30 deg at Re = 200,000 and for unsteady pitching motions described by alpha equals alpha(sub m) + 10 deg(sin(wt)). The models were two dimensional, and the test was performed in a water tunnel at Ames Research Center. The unsteady conditions ranged over Re equals 100,000 to 250,000, k equals 0.001 to 0.2, and alpha(sub m) = 10 deg to 20 deg. Unsteady lift, drag, and pitching-moment measurements were obtained along with fluorescent-dye flow visualizations. The addition of the slat was found to delay the static-drag and static-moment stall by about 5 degrees and to eliminate completely the development of a dynamic-stall vortex during unsteady motions that reached angles as high as 25 degrees. In all of the unsteady cases studied, the slat caused a significant reduction in the force and moment hysteresis amplitudes. The reduced frequency was found to have the greatest effect on the results, whereas the Reynolds number had little effect on the behavior of either the basic or the slatted airfoil. The slat caused a slight drag penalty at low angles of attack, but generally increased the lift/drag ratio when averaged over the full cycle of oscillation.
A two dimensional study of rotor/airfoil interaction in hover
NASA Technical Reports Server (NTRS)
Lee, Chyang S.
1988-01-01
A two dimensional model for the chordwise flow near the wing tip of the tilt rotor in hover is presented. The airfoil is represented by vortex panels and the rotor is modeled by doublet panels. The rotor slipstream and the airfoil wake are simulated by free point vortices. Calculations on a 20 percent thick elliptical airfoil under a uniform rotor inflow are performed. Variations on rotor size, spacing between the rotor and the airfoil, ground effect, and the influence upper surface blowing in download reduction are analyzed. Rotor size has only a minor influence on download when it is small. Increase of the rotor/airfoil spacing causes a gradual decrease on download. Proximity to the ground effectively reduces the download and makes the wake unsteady. The surface blowing changes the whole flow structure and significantly reduces the download within the assumption of a potential solution. Improvement on the present model is recommended to estimate the wall jets induced suction on the airfoil lower surface.
Some aspects regarding unsteady fluid flow with time dependent boundary conditions
NASA Astrophysics Data System (ADS)
Rotaru, Constantin; CÃ®rciu, Ionicǎ
2016-02-01
This study explores the dynamic characteristics of the unsteady fluid flow around an airfoil where the time dependency is introduced through the boundary conditions. The models of the vortex wake's shape and strength were based on the time dependency history of the motion and the methods of solution that were developed for these models included the treatment of the zero normal flow on a solid surface. A numerical investigation has been performed using Maplesoft environment. As a result of the nonuniform motion, the path along which the airfoil moves was assumed to be prescribed. In the paper the wake shed from the trailing edge of the lifting surfaces was modeled by vortex distribution.
NASA Technical Reports Server (NTRS)
Han, J. C.; Chandra, P. R.
1987-01-01
The heat transfer characteristics of turbulent air flow in a multipass channel were studied via the naphthalene sublimation technique. The naphthalene-coated test section, consisting of two straight, square channels joined by a 180 deg turn, resembled the internal cooling passages of gas turbine airfoils. The top and bottom surfaces of the test channel were roughened by rib turbulators. The rib height-to-hydraulic diameter ratio (e/D) were 0.063 and 0.094, and the rib pitch-to-height ratio (P/e) were 10 and 20. The local heat/mass transfer coefficients on the roughened top wall and on the smooth divider and side walls of the test channel were determined for three Reynolds numbers of 15, 30, and 60, thousand, and for three angles of attack (alpha) of 90, 60, and 45 deg. Results showed that the local Sherwood numbers on the ribbed walls were 1.5 to 6.5 times those for a fully developed flow in a smooth square duct. The average ribbed-wall Sherwood numbers were 2.5 to 3.5 times higher than the fully developed values, depending on the rib angle of attack and the Reynolds number. The results also indicated that, before the turn, the heat/mass transfer coefficients in the cases of alpha = 60 and 45 deg were higher than those in the case of alpha=90 deg. However, after the turn, the heat/mass transfer coefficients in the oblique-rib cases were lower than those in the transverse rib case. Correlations for the average Sherwood number ratios for individual channel surfaces and for the overall Sherwood number ratios are reported. Correlations for the fully developed friction factors and for the loss coefficients are also provided.
NASA Technical Reports Server (NTRS)
Nagamatsu, H. T.; Ficarra, R.; Orozco, R.
1983-01-01
The optimization of passive shock wave/boundary layer control for supercritical airfoil drag reduction was investigated in a 3 in. x 15.4 in. Transonic Blowdown Wind Tunnel. A 14% thick supercritical airfoil was tested with 0%, 1.42% and 2.8% porosities at Mach numbers of .70 to .83. The 1.42% case incorporated a linear increase in porosity with the flow direction while the 2.8% case was uniform porosity. The static pressure distributions over the airfoil, the wake impact pressure data for determining the profile drag, and the Schlieren photographs for porous surface airfoils are presented and compared with the results for solid-surface airfoils. While the results show that linear 1.42% porosity actually led to a slight increase in drag it was found that the uniform 2.8% porosity can lead to a drag reduction of 46% at M = .81.
Fluid mechanics of dynamic stall. I - Unsteady flow concepts
NASA Technical Reports Server (NTRS)
Ericsson, L. E.; Reding, J. P.
1988-01-01
Advanced military aircraft 'supermaneuverability' requirements entail the sustained operation of airfoils at stalled flow conditions. The present work addresses the effects of separated flow on vehicle dynamics; an analytic method is presented which employs static experimental data to predict the separated flow effect on incompressible unsteady aerodynamics. The key parameters in the analytic relationship between steady and nonsteady aerodynamics are the time-lag before a change of flow conditions can affect the separation-induced aerodynamic loads, the accelerated flow effect, and the moving wall effect.
Control of Flow Separation Using Adaptive Airfoils
NASA Technical Reports Server (NTRS)
Chandrasekhara, M. S.; Wilder, M. C.; Carr, L. W.; Davis, Sanford S. (Technical Monitor)
1996-01-01
A novel way of controlling flow separation is reported. The approach involves using an adaptive airfoil geometry that changes its leading edge shape to adjust to the instantaneous flow at high angles of attack such that the flow over it remains attached. In particular, a baseline NACA 0012 airfoil, whose leading edge curvature could be changed dynamically by 400% was tested under quasi-steady compressible flow conditions. A mechanical drive system was used to produce a rounded leading edge to reduce the strong local flow acceleration around its nose and thus reduce the strong adverse pressure gradient that follows such a rapid acceleration. Tests in steady flow showed that at M = 0.3, the flow separated at about 14 deg. angle of attack for the NACA 0012 profile but could be kept attached up to an angle of about 18 deg by changing the nose curvature. No significant hysteresis effects were observed; the flow could be made to reattach from its separated state at high angles by changing the leading edge curvature. Interestingly, the flow over a nearly semicircular nosed airfoil was separated even at low angles.
Current Issues in Unsteady Turbomachinery Flows (Images)
NASA Technical Reports Server (NTRS)
Povinelli, Louis
2004-01-01
Among the numerous causes for unsteadiness in turbo machinery flows are turbulence and flow environment, wakes from stationary and rotating vanes, boundary layer separation, boundary layer/shear layer instabilities, presence of shock waves and deliberate unsteadiness for flow control purposes. These unsteady phenomena may lead to flow-structure interactions such as flutter and forced vibration as well as system instabilities such as stall and surge. A major issue of unsteadiness relates to the fact that a fundamental understanding of unsteady flow physics is lacking and requires continued attention. Accurate simulations and sufficient high fidelity experimental data are not available. The Glenn Research Center plan for Engine Component Flow Physics Modeling is part of the NASA 21st Century Aircraft Program. The main components of the plan include Low Pressure Turbine National Combustor Code. The goals, technical output and benefits/impacts of each element are described in the presentation. The specific areas selected for discussion in this presentation are blade wake interactions, flow control, and combustor exit turbulence and modeling.
Wave propagation in beams with periodic arrays of airfoil-shaped resonating units
NASA Astrophysics Data System (ADS)
Casadei, Filippo; Bertoldi, Katia
2014-12-01
This paper presents an analytical and numerical study on the dispersion properties of an Euler-Bernoulli beam immersed in a steady fluid flow with periodic arrays of airfoil-shaped vibration absorbers attached to it. The resonance characteristics of the airfoils generate strong attenuation of flexural waves in the beam occurring at frequencies defined by the properties of the airfoils and the speed of the incident fluid. Analytical and numerical tools are developed to investigate the effects of the incident flow on the dispersion properties and the bandgaps of the system. Both steady and unsteady aerodynamic models are used to model the lift force and the pitching moment acting on the resonators and their effect on the dispersion relations of the system is evaluated. Finally, an effective medium description of the beam is developed to capture its behavior at long-wavelengths. In this regime, the system can be effectively considered as an acoustic metamaterial with adaptive dispersion properties.
Calculation of boundary layers near the stagnation point of an oscillating airfoil
NASA Technical Reports Server (NTRS)
Cebeci, T.; Carr, L. W.
1983-01-01
The results of an investigation of boundary layers close to the stagnation point of an oscillating airfoil are reported. Two procedures for generating initial conditions, the characteristics box scheme and a quasi-static approach, were investigated, and the quasi-static approach was shown to be appropriate provided the initial region was far from any flow separation. With initial conditions generated in this way, the unsteady boundary layer equations were solved for the flow in the leading edge region of a NACA 0012 airfoil oscillating from 0 to 5 deg. Results were obtained for both laminar and turbulent flow, and, in the latter case, the effect of transition was assessed by specifying its occurrence at different locations. The results demonstrate the validity of the numerical scheme and suggest that the procedures should be applied to calculation of the entire flow around oscillating airfoils.
Water-tunnel experiments on an oscillating airfoil at RE equals 21,000
NASA Technical Reports Server (NTRS)
Mcalister, K. W.; Carr, L. W.
1978-01-01
Flow visualization experiments were performed in a water tunnel on a modified NACA 0012 airfoil undergoing large amplitude harmonic oscillations in pitch. Hydrogen bubbles were used to: (1) create a conveniently striated and well preserved set of inviscid flow markers; and (2) to expose the succession of events occurring within the viscous domain during the onset of dynamic stall. Unsteady effects were shown to have an important influence on the progression of flow reversal along the airfoil surface prior to stall. A region of reversed flow underlying a free shear layer was found to momentarily exist over the entire upper surface without any appreciable disturbance of the viscous-inviscid boundary. A flow protuberance was observed to develop near the leading edge, while minor vortices evolve from an expanding instability of the free shear layer over the rear portion of the airfoil. The complete breakdown of this shear layer culminates in the successive formation of two dominant vortices.
Tangler, J. L.; Kocurek, J. D.
2004-10-01
The objective of this study was to provide post-stall airfoil data input guidelines for the prediction of peak and post-peak rotor power when using blade-element momentum theory. A steady-state data set from the Unsteady Aerodynamic Experiment (UAE) rotor test was used to provide guidelines for the development of a global post-stall method for the prediction of post-stall 3-D airfoil characteristics to be used with 2-D airfoil data. Based on these UAE data, methods to emulate the 3-D aerodynamics in the post-stall region were explored. Also suggested are experimental tests needed to better understand the 3-D flow physics and to quantify needed theory or empirical factors for a global post-stall approach to support blade-element momentum methods.
A natural low-frequency oscillation of the flow over an airfoil near stalling conditions
NASA Technical Reports Server (NTRS)
Zaman, K. B. M. Q.; Mckinzie, D. J.; Rumsey, C. L.
1989-01-01
An experimental and computational study of the low-frequency oscillation observed in the flow over an airfoil at the onset of static stall is presented. Wind-tunnel results obtained with two-dimensional airfoil models show that this phenomena takes place only with a transitional state of the separating boundary layer. It is noted that the flowfield does not involve a Karman vortex street. The experimental results agree well with the results of a two-dimensional Navier-Stokes code. The present study demonstrates that the low-frequency oscillations produce intense flow fluctuations which impart much larger unsteady forces to the airfoil than experienced by bluff-body shedding and which may represent the primary aerodynamics of stall flutter of blades and wings.
Effects of a ground vortex on the aerodynamics of an airfoil
NASA Technical Reports Server (NTRS)
Krothapalli, A.; Leopold, D.
1988-01-01
An experimental investigation was carried out to study the aerodynamics of an airfoil with a rectangular jet exiting from its lower surface at fifty percent of the chord. The airfoil was tested with and without the influence of a ground plane. Surface static pressures were measured on the airfoil at jet to free stream velocity ratios ranging from 0 to 9. From these pressures, the variation of C sub L with velocity ratio was easily determined. The measurements indicated significant positive and negative pressure regions on the lower surface of the airfoil ahead of and after the nozzle exit respectively. The presence of a ground plane enhanced these pressure regions at low velocity ratios, but at a particular ratio for each plane location, a recirculation zone or vortex formed ahead of the jet resulting in decreased pressures and a drop in C sub L.
Temperature Cancellation Method of Motion-Capturing Pressure-Sensitive Paint System
NASA Astrophysics Data System (ADS)
Sakaue, Hirotaka; Yamada, Yuki; Okabe, Taika; Miyazaki, Takeshi
2013-11-01
Motion-capturing pressure-sensitive paint system uses two luminescent outputs to extract the pressure field on an aerodynamic object. This uses a luminescent imaging technique to relate the luminescent output to the pressure. In the previous study, this system is applied to capture the time-resolved unsteady pressure fields on a fluttering airfoil, and a bullet-shaped model. Pressure-sensitive paint (PSP) has a temperature dependency, which is a major error source for the PSP measurement. Motion-capturing PSP system also involves the temperature dependency of PSP itself. In the presentation, we propose a temperature-cancellation method of the motion-capturing PSP system. This method does not require a separate temperature measurement for the temperature correction that is advantage for capturing the pressure field on a moving object.
Aerodynamics Characteristics of Multi-Element Airfoils at -90 Degrees Incidence
NASA Technical Reports Server (NTRS)
Stremel, Paul M.; Schmitz, Fredric H. (Technical Monitor)
1994-01-01
A developed method has been applied to calculate accurately the viscous flow about airfoils normal to the free-stream flow. This method has special application to the analysis of tilt rotor aircraft in the evaluation of download. In particular, the flow about an XV-15 airfoil with and without deflected leading and trailing edge flaps at -90 degrees incidence is evaluated. The multi-element aspect of the method provides for the evaluation of slotted flap configurations which may lead to decreased drag. The method solves for turbulent flow at flight Reynolds numbers. The flow about the XV-15 airfoil with and without flap deflections has been calculated and compared with experimental data at a Reynolds number of one million. The comparison between the calculated and measured pressure distributions are very good, thereby, verifying the method. The aerodynamic evaluation of multielement airfoils will be conducted to determine airfoil/flap configurations for reduced airfoil drag. Comparisons between the calculated lift, drag and pitching moment on the airfoil and the airfoil surface pressure will also be presented.
A General Algorithm for Reusing Krylov Subspace Information. I. Unsteady Navier-Stokes
NASA Technical Reports Server (NTRS)
Carpenter, Mark H.; Vuik, C.; Lucas, Peter; vanGijzen, Martin; Bijl, Hester
2010-01-01
A general algorithm is developed that reuses available information to accelerate the iterative convergence of linear systems with multiple right-hand sides A x = b (sup i), which are commonly encountered in steady or unsteady simulations of nonlinear equations. The algorithm is based on the classical GMRES algorithm with eigenvector enrichment but also includes a Galerkin projection preprocessing step and several novel Krylov subspace reuse strategies. The new approach is applied to a set of test problems, including an unsteady turbulent airfoil, and is shown in some cases to provide significant improvement in computational efficiency relative to baseline approaches.
Airfoil for a gas turbine engine
Liang, George
2011-05-24
An airfoil is provided for a turbine of a gas turbine engine. The airfoil comprises: an outer structure comprising a first wall including a leading edge, a trailing edge, a pressure side, and a suction side; an inner structure comprising a second wall spaced from the first wall and at least one intermediate wall; and structure extending between the first and second walls so as to define first and second gaps between the first and second walls. The second wall and the at least one intermediate wall define at least one pressure side supply cavity and at least one suction side supply cavity. The second wall may include at least one first opening near the leading edge of the first wall. The first opening may extend from the at least one pressure side supply cavity to the first gap. The second wall may further comprise at least one second opening near the trailing edge of the outer structure. The second opening may extend from the at least one suction side supply cavity to the second gap. The first wall may comprise at least one first exit opening extending from the first gap through the pressure side of the first wall and at least one second exit opening extending from the second gap through the suction side of the second wall.
Effects of grit roughness and pitch oscillations on the NACA 4415 airfoil
Hoffmann, M.J.; Reuss Ramsay, R.; Gregorek, G.M.
1996-07-01
A NACA 4415 airfoil model was tested in The Ohio State University Aeronautical and Astronautical Research Laboratory 3 x 5 subsonic wind tunnel under steady state and unsteady conditions. The test defined baseline conditions for steady state angles of attack from {minus}10{degree} to +40{degree} and examined unsteady behavior by oscillating the model about its pitch axis for three mean angles, three frequencies, and two amplitudes. For all cases, Reynolds numbers of 0.75, 1, 1.25, and 1.5 million were used. In addition, these were repeated after the application of leading edge grit roughness (LEGR) to determine contamination effects on the airfoil performance. Steady state results of the NACA 4415 testing at Reynolds number of 1.25 million showed a baseline maximum lift coefficient of 1.30 at 12.3{degree} angle of attack. The application of LEGR reduced the maximum lift coefficient by 20% and increased the 0.0090 minimum drag coefficient value by 62%. The zero lift pitching moment of {minus}0.0967 showed a 13% reduction in magnitude to {minus}0.0842 with LEGR applied. Data were also obtained for two pitch oscillation amplitudes: {+-}5.5{degree} and {+-}10{degree}. The larger amplitude consistently gave a higher maximum lift coefficient than the smaller amplitude, and both unsteady maximum lift coefficients were greater than the steady state values. Stall is delayed on the airfoil while the angle of attack is increasing, thereby causing an increase in maximum lift coefficient. A hysteresis behavior was exhibited for all the unsteady test cases. The hysteresis loops were larger for the higher reduced frequencies and for the larger amplitude oscillations. As in the steady case, the effect of LEGR in the unsteady case was to reduce the lift coefficient at high angles of attack. In addition, with LEGR, the hysteresis behavior persisted into lower angles of attack than for the clean case.
Effects of grit roughness and pitch oscillations on the S810 airfoil
Ramsay, R.R.; Hoffman, M.J.; Gregorek, G.M.
1996-01-01
An S810 airfoil model was tested in The Ohio State University Aeronautical and Astronautical Research Laboratory 3 x 5 subsonic wind tunnel under steady state and unsteady conditions. The test defined baseline conditions for steady state angles of attack from -20{degrees} to +40{degrees} and examined unsteady behavior by oscillating the model about its pitch axis for three mean angles, three frequencies, and two amplitudes. For all cases, Reynolds numbers of 0.75, 1, 1.25, and 1.5 million were used. In addition, the above conditions were repeated after the application of leading edge grit roughness (LEGR) to determine contamination effects on the airfoil performance. Baseline steady state results of the S810 testing showed a maximum lift coefficient of 1.15 at 15.2{degrees}angle of attack. The application of LEGR reduced the maximum lift coefficient by 12% and increased the 0.0085 minimum drag coefficient value by 88%. The zero lift pitching moment of -0.0286 showed a 16% reduction in magnitude to -0.0241 with LEGR applied. Data were also obtained for two pitch oscillation amplitudes: {plus_minus}5.5{degrees} and {plus_minus}10{degrees}. The larger amplitude consistently gave a higher maximum lift coefficient than the smaller amplitude and both sets of unsteady maximum lift coefficients were greater than the steady state values. Stall was delayed on the airfoil while the angle of attack was increasing, thereby causing an increase in maximum lift coefficient. A hysteresis behavior was exhibited for all the unsteady test cases. The hysteresis loops were larger for the higher reduced frequencies and for the larger amplitude oscillations. In addition to the hysteresis behavior, an unusual feature of these data were a sudden increase in the lift coefficient where the onset of stall was expected. As in the steady case, the effect of LEGR in the unsteady case was to reduce the lift coefficient at high angles of attack.
Airfoil nozzle and shroud assembly
Shaffer, James E.; Norton, Paul F.
1997-01-01
An airfoil and nozzle assembly including an outer shroud having a plurality of vane members attached to an inner surface and having a cantilevered end. The assembly further includes a inner shroud being formed by a plurality of segments. Each of the segments having a first end and a second end and having a recess positioned in each of the ends. The cantilevered end of the vane member being positioned in the recess. The airfoil and nozzle assembly being made from a material having a lower rate of thermal expansion than that of the components to which the airfoil and nozzle assembly is attached.
Airfoil nozzle and shroud assembly
Shaffer, J.E.; Norton, P.F.
1997-06-03
An airfoil and nozzle assembly are disclosed including an outer shroud having a plurality of vane members attached to an inner surface and having a cantilevered end. The assembly further includes a inner shroud being formed by a plurality of segments. Each of the segments having a first end and a second end and having a recess positioned in each of the ends. The cantilevered end of the vane member being positioned in the recess. The airfoil and nozzle assembly being made from a material having a lower rate of thermal expansion than that of the components to which the airfoil and nozzle assembly is attached. 5 figs.
The effect of undulating leading-edge modifications on NACA 0021 airfoil characteristics
NASA Astrophysics Data System (ADS)
Rostamzadeh, N.; Kelso, R. M.; Dally, B. B.; Hansen, K. L.
2013-11-01
In spite of its mammoth physical size, the humpback whale's manoeuvrability in hunting has captured the attention of biologists as well as fluid mechanists. It has now been established that the protrusions on the leading-edges of the humpback's pectoral flippers, known as tubercles, account for this species' agility and manoeuvrability. In the present work, Prandtl's nonlinear lifting-line theory was employed to propose a hypothesis that the favourable traits observed in the performance of tubercled lifting bodies are not exclusive to this form of leading-edge configuration. Accordingly, a novel alternative to tubercles was introduced and incorporated into the design of four airfoils that underwent wind tunnel force and pressure measurement tests in the transitional flow regime. In addition, a Computation Fluid Dynamics study was performed using the Shear Stress Transport transitional model in the context of unsteady Reynolds-Averaged Navier-Stokes at several attack angles. The results from the numerical investigation are in reasonable agreement with those of the experiments, and suggest the presence of features that are also observed in flows over tubercled foils, most notably a distinct pair of streamwise vortices for each wavelength of the tubercle-like feature.
Active Control of Separation From the Flap of a Supercritical Airfoil
NASA Technical Reports Server (NTRS)
Melton, LaTunia Pack; Yao, Chung-Sheng; Seifert, Avi
2006-01-01
Zero-mass-flux periodic excitation was applied at several regions on a simplified high-lift system to delay the occurrence of flow separation. The NASA Energy Efficient Transport (EET) supercritical airfoil was equipped with a 15% chord simply hinged leading edge flap and a 25% chord simply hinged trailing edge flap. Detailed flow features were measured in an attempt to identify optimal actuator placement. The measurements included steady and unsteady model and tunnel wall pressures, wake surveys, arrays of surface hot-films, flow visualization, and particle image velocimetry (PIV). The current paper describes the application of active separation control at several locations on the deflected trailing edge flap. High frequency (F(+) approximately equal to 10) and low frequency amplitude modulation (F(+) sub AM approximately equal to 1) of the high frequency excitation were used for control. It was noted that the same performance gains were obtained with amplitude modulation and required only 30% of the momentum input required by pure sine excitation.
Bifurcations in unsteady aerodynamics
NASA Technical Reports Server (NTRS)
Tobak, M.; Unal, A.
1986-01-01
Nonlinear algebraic functional expansions are used to create a form for the unsteady aerodynamic response that is consistent with solutions of the time dependent Navier-Stokes equations. An enumeration of means of invalidating Frechet differentiability of the aerodynamic response, one of which is aerodynamic bifurcation, is proposed as a way of classifying steady and unsteady aerodynamic phenomena that are important in flight dynamics applications. Accomodating bifurcation phenomena involving time dependent equilibrium states within a mathematical model of the aerodynamic response raises an issue of memory effects that becomes more important with each successive bifurcation.
Thin airfoil theory based on approximate solution of the transonic flow equation
NASA Technical Reports Server (NTRS)
Spreiter, John R; Alksne, Alberta Y
1957-01-01
A method is presented for the approximate solution of the nonlinear equations transonic flow theory. Solutions are found for two-dimensional flows at a Mach number of 1 and for purely subsonic and purely supersonic flows. Results are obtained in closed analytic form for a large and significant class of nonlifting airfoils. At a Mach number of 1 general expressions are given for the pressure distribution on an airfoil of specified geometry and for the shape of an airfoil having a prescribed pressure distribution. Extensive comparisons are made with available data, particularly for a Mach number of 1, and with existing solutions.
Thin airfoil theory based on approximate solution of the transonic flow equation
NASA Technical Reports Server (NTRS)
Spreiter, John R; Alksne, Alberta Y
1958-01-01
A method is presented for the approximate solution of the nonlinear equations of transonic flow theory. Solutions are found for two-dimensional flows at a Mach number of 1 and for purely subsonic and purely supersonic flows. Results are obtained in closed analytic form for a large and significant class of nonlifting airfoils. At a Mach number of 1 general expressions are given for the pressure distribution on an airfoil of specified geometry and for the shape of an airfoil having a prescribed pressure distribution. Extensive comparisons are made with available data, particularly for a Mach number of 1, and with existing solutions.
Aspects of Numerical Simulation of Circulation Control Airfoils
NASA Technical Reports Server (NTRS)
Swanson, R. C.; Rumsey, C. L.; Anders, S. G.
2005-01-01
The mass-averaged compressible Navier-Stokes equations are solved for circulation control airfoils. Numerical solutions are computed with a multigrid method that uses an implicit approximate factorization smoother. The effects of flow conditions (e.g., free-stream Mach number, angle of attack, momentum coefficient) and mesh on the prediction of circulation control airfoil flows are considered. In addition, the impact of turbulence modeling, including curvature effects and modifications to reduce eddy viscosity levels in the wall jet (i.e., Coanda flow), is discussed. Computed pressure distributions are compared with available experimental data.
NASA Astrophysics Data System (ADS)
Zhang, Qiang
The effects of surface roughness, turbulence intensity, Mach number, and streamline curvature-airfoil shape on the aerodynamic performance of turbine airfoils are investigated in compressible, high speed flows. The University of Utah Transonic Wind Tunnel is employed for the experimental part of the study. Two different test sections are designed to produce Mach numbers, Reynolds numbers, passage mass flow rates, and physical dimensions, which match values along turbine blades in operating engines: (i) a nonturning test section with a symmetric airfoil, and (ii) a cascade test section with a cambered turbine vane. The nonuniform, irregular, three-dimensional surface roughness is characterized using the equivalent sand grain roughness size. Changing the airfoil surface roughness condition has a substantial effect on wake profiles of total pressure loss coefficients, normalized Mach number, normalized kinetic energy, and on the normalized and dimensional magnitudes of Integrated Aerodynamic Losses produced by the airfoils. Comparisons with results for a symmetric airfoil and a cambered vane show that roughness has more substantial effects on losses produced by the symmetric airfoil than the cambered vane. Data are also provided that illustrate the larger loss magnitudes are generally present with flow turning and cambered airfoils, than with symmetric airfoils. Wake turbulence structure of symmetric airfoils and cambered vanes are also studied experimentally. The effects of surface roughness and freestream turbulence levels on wake distributions of mean velocity, turbulence intensity, and power spectral density profiles and vortex shedding frequencies are quantified one axial chord length downstream of the test airfoils. As the level of surface roughness increases, all wake profile quantities broaden significantly and nondimensional vortex shedding frequencies decrease. Wake profiles produced by the symmetric airfoil are more sensitive to variations of surface
Flatback airfoil wind tunnel experiment.
Mayda, Edward A.; van Dam, C.P.; Chao, David D.; Berg, Dale E.
2008-04-01
A computational fluid dynamics study of thick wind turbine section shapes in the test section of the UC Davis wind tunnel at a chord Reynolds number of one million is presented. The goals of this study are to validate standard wind tunnel wall corrections for high solid blockage conditions and to reaffirm the favorable effect of a blunt trailing edge or flatback on the performance characteristics of a representative thick airfoil shape prior to building the wind tunnel models and conducting the experiment. The numerical simulations prove the standard wind tunnel corrections to be largely valid for the proposed test of 40% maximum thickness to chord ratio airfoils at a solid blockage ratio of 10%. Comparison of the computed lift characteristics of a sharp trailing edge baseline airfoil and derived flatback airfoils reaffirms the earlier observed trend of reduced sensitivity to surface contamination with increasing trailing edge thickness.
Second Stage Turbine Bucket Airfoil.
Xu, Liming; Ahmadi, Majid; Humanchuk, David John; Moretto, Nicholas; Delehanty, Richard Edward
2003-05-06
The second-stage buckets have airfoil profiles substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in inches in Table I wherein Z is a perpendicular distance from a plane normal to a radius of the turbine centerline and containing the X and Y values with the Z value commencing at zero in the X, Y plane at the radially innermost aerodynamic section of the airfoil and X and Y are coordinate values defining the airfoil profile at each distance Z. The X, Y and Z values may be scaled as a function of the same constant or number to provide a scaled-up or scaled-down airfoil section for the bucket.
Boundary Layer Control on Airfoils.
ERIC Educational Resources Information Center
Gerhab, George; Eastlake, Charles
1991-01-01
A phenomena, boundary layer control (BLC), produced when visualizing the fluidlike flow of air is described. The use of BLC in modifying aerodynamic characteristics of airfoils, race cars, and boats is discussed. (KR)
Analysis of a theoretically optimized transonic airfoil
NASA Technical Reports Server (NTRS)
Lores, M. E.; Burdges, K. P.; Shrewsbury, G. D.
1978-01-01
Numerical optimization was used in conjunction with an inviscid, full potential equation, transonic flow analysis computer code to design an upper surface contour for a conventional airfoil to improve its supercritical performance. The modified airfoil was tested in a compressible flow wind tunnel. The modified airfoil's performance was evaluated by comparison with test data for the baseline airfoil and for an airfoil developed by optimization of leading edge of the baseline airfoil. While the leading edge modification performed as expected, the upper surface re-design did not produce all of the expected performance improvements. Theoretical solutions computed using a full potential, transonic airfoil code corrected for viscosity were compared to experimental data for the baseline airfoil and the upper surface modification. These correlations showed that the theory predicted the aerodynamics of the baseline airfoil fairly well, but failed to accurately compute drag characteristics for the upper surface modification.
Nozzle airfoil having movable nozzle ribs
Yu, Yufeng Phillip; Itzel, Gary Michael
2002-01-01
A nozzle vane or airfoil structure is provided in which the nozzle ribs are connected to the side walls of the vane or airfoil in such a way that the ribs provide the requisite mechanical support between the concave side and convex side of the airfoil but are not locked in the radial direction of the assembly, longitudinally of the airfoil. The ribs may be bi-cast onto a preformed airfoil side wall structure or fastened to the airfoil by an interlocking slide connection and/or welding. By attaching the nozzle ribs to the nozzle airfoil metal in such a way that allows play longitudinally of the airfoil, the temperature difference induced radial thermal stresses at the nozzle airfoil/rib joint area are reduced while maintaining proper mechanical support of the nozzle side walls.
Investigation of the Kline-Fogleman airfoil section for rotor blade applications
NASA Technical Reports Server (NTRS)
Lumsdaine, E.; Johnson, W. S.; Fletcher, L. M.; Peach, J. E.
1974-01-01
Wind tunnel tests of a wedgeshaped airfoil with sharp leading edge and a spanwise step were conducted. The airfoil was tested with variations of the following parameters: (1) Reynolds number, (2) step location, (3) step shape, (4) apex angle, and (5) with the step on either the upper or lower surface. The results are compared with a flat plate and with wedge airfoils without a step having the same aspect ratio. Water table tests were conducted for flow visualization and it was determined that the flow separates from the upper surface at low angles of attack. The wind tunnel tests show that the lift/drag ratio of the airfoil is lower than for a flat plate and the pressure data show that the airfoil derives its lift in the same manner as a flat plate.
NASA Technical Reports Server (NTRS)
Harris, Charles D.; Harvey, William D.; Brooks, Cuyler W., Jr.
1988-01-01
A large-chord, swept, supercritical, laminar-flow-control (LFC) airfoil was designed and constructed and is currently undergoing tests in the Langley 8 ft Transonic Pressure Tunnel. The experiment was directed toward evaluating the compatibility of LFC and supercritical airfoils, validating prediction techniques, and generating a data base for future transport airfoil design as part of NASA's ongoing research program to significantly reduce drag and increase aircraft efficiency. Unique features of the airfoil included a high design Mach number with shock free flow and boundary layer control by suction. Special requirements for the experiment included modifications to the wind tunnel to achieve the necessary flow quality and contouring of the test section walls to simulate free air flow about a swept model at transonic speeds. Design of the airfoil with a slotted suction surface, the suction system, and modifications to the tunnel to meet test requirements are discussed.
Effects of finite aspect ratio on wind turbine airfoil measurements
NASA Astrophysics Data System (ADS)
Kiefer, Janik; Miller, Mark A.; Hultmark, Marcus; Hansen, Martin O. L.
2016-09-01
Wind turbines partly operate in stalled conditions within their operational cycle. To simulate these conditions, it is also necessary to obtain 2-D airfoil data in terms of lift and drag coefficients at high angles of attack. Such data has been obtained previously, but often at low aspect ratios and only barely past the stall point, where strong wall boundary layer influence is expected. In this study, the influence of the wall boundary layer on 2D airfoil data, especially in the post stall domain, is investigated. Here, a wind turbine airfoil is tested at different angles of attack and with two aspect ratios of AR = 1 and AR = 2. The tests are conducted in a wind tunnel that is pressurized up to 150 bar in order to achieve a constant Reynolds number of Rec = 3 • 106, despite the variable chord length.
NASA Technical Reports Server (NTRS)
Prandtl, L
1940-01-01
The basic ideas of a new method for treating the problem of the airfoil are presented, and a review is given of the problems thus far computed for incompressible and supersonic flows. Test results are reported for the airfoil of circular plan form and the results are shown to agree well with the theory. As a supplement, a theory based on the older methods is presented for the rectangular of small aspect ratio.
NASA Technical Reports Server (NTRS)
Hathaway, M. D.; Suder, K. L.; Strazisar, A. J.; Adamczyk, J. J.; Okiishi, T. H.
1987-01-01
Unsteady velocity field measurements made within the stator row of a transonic axial-flow fan are presented. Measurements were obtained at midspan for two different stator blade rows using a laser anemometer. The first stator row consists of double circular-arc airfoils with a solidity of 1.68. The second features controlled-diffusion airfoils with a solidity of 0.85. Both were tested at design-speed peak efficiency conditions. In addition, the controlled-diffusion stator was also tested at near stall conditions. The procedures developed here are used to identify the rotor wake generated and unresolved unsteadiness from the velocity measurements (rotor wake generated unsteadiness refers to the unsteadiness generated by the rotor wake velocity deficit and unresolved unsteadiness refers to all remaining unsteadiness which contributes to the spread in the distribution of velocities such as vortex shedding, turbulence, etc.). Auto and cross correlations of these unsteady velocity fluctuations are presented to show their relative magnitude and spatial distributions. Amplification and attenuation of both rotor wake generated and unresolved unsteadiness are shown to occur within the stator blade passage.
NASA Technical Reports Server (NTRS)
Hathaway, M. D.; Suder, K. L.; Okiishi, T. H.; Strazisar, A. J.; Adamczyk, J. J.
1987-01-01
Unsteady velocity field measurements made within the stator row of a transonic axial-flow fan are presented. Measurements were obtained at midspan for two different stator blade rows using a laser anemometer. The first stator row consists of double circular-arc airfoils with a solidity of 1.68. The second features controlled-diffusion airfoils with a solidity of 0.85. Both were tested at design-speed peak efficiency conditions. In addition, the controlled-diffusion stator was also tested at near stall conditions. The procedures developed here are used to identify the rotor wake generated and unresolved unsteadiness from the velocity measurements (rotor wake generated unsteadiness refers to the unsteadiness generated by the rotor wake velocity deficit and unresolved unsteadiness refers to all remaining unsteadiness which contributes to the spread in the distribution of velocities such as vortex shedding, turbulence, etc.). Auto and cross correlations of these unsteady velocity fluctuations are presented to show their relative magnitude and spatial distributions. Amplification and attenuation of both rotor wake generated and unresolved unsteadiness are shown to occur within the stator blade passage.
Nature-Inspired Airfoils for Environmental Noise Reduction
NASA Astrophysics Data System (ADS)
Han, Suyeong; Kyung, Richard
2013-11-01
Recently, study on the insects' flapping flight became one of the challenging research subjects in the field of environmental engineering and aeronautics because of its potential applicability to intelligent micro-robots capable of autonomous flight and the next generation aerial-vehicles. In order to uncover its curious unsteady characteristics, many researchers have conducted experimental and computational studies on the unsteady aerodynamics of insects' flapping flight. In the present work, the unsteady flow physics around insect wings are conducted by utilizing numerical and computational simulation. The e-AIRS [6] (e-Science Aerospace Integrated Research System) gives a balanced service between computational and experimental aerodynamics, along with integrated research process of these two research activities. This paper presents the wing motions and their aerodynamics with a two dimensional approach to reduce environmental noise during the airflight. Also this paper shows an optimal phase angle, where the thrust is maximized at the position of minimized drag, which occurs when noise is minimized. Aside from the two-dimensional approach, stroke angles and phase angles of the airfoils are set as parameters, to determine which motion yields the best aerodynamic characteristics.
NASA Technical Reports Server (NTRS)
Adamczyk, J. L.
1974-01-01
An approximate solution is reported for the unsteady aerodynamic response of an infinite swept wing encountering a vertical oblique gust in a compressible stream. The approximate expressions are of closed form and do not require excessive computer storage or computation time, and further, they are in good agreement with the results of exact theory. This analysis is used to predict the unsteady aerodynamic response of a helicopter rotor blade encountering the trailing vortex from a previous blade. Significant effects of three dimensionality and compressibility are evident in the results obtained. In addition, an approximate solution for the unsteady aerodynamic forces associated with the pitching or plunging motion of a two dimensional airfoil in a subsonic stream is presented. The mathematical form of this solution approaches the incompressible solution as the Mach number vanishes, the linear transonic solution as the Mach number approaches one, and the solution predicted by piston theory as the reduced frequency becomes large.
Numerical and experimental investigations on unsteady aerodynamics of flapping wings
NASA Astrophysics Data System (ADS)
Yu, Meilin
The development of a dynamic unstructured grid high-order accurate spectral difference (SD) method for the three dimensional compressible Navier-Stokes (N-S) equations and its applications in flapping-wing aerodynamics are carried out in this work. Grid deformation is achieved via an algebraic blending strategy to save computational cost. The Geometric Conservation Law (GCL) is imposed to ensure that grid deformation will not contaminate the flow physics. A low Mach number preconditioning procedure is conducted in the developed solver to handle the bio-inspired flow. The capability of the low Mach number preconditioned SD solver is demonstrated by a series of two dimensional (2D) and three dimensional (3D) simulations of the unsteady vortex dominated flow. Several topics in the flapping wing aerodynamics are numerically and experimentally investigated in this work. These topics cover some of the cutting-edge issues in flapping wing aerodynamics, including the wake structure analysis, airfoil thickness and kinematics effects on the aerodynamic performances, vortex structure analysis around 3D flapping wings and the kinematics optimization. Wake structures behind a sinusoidally pitching NACA0012 airfoil are studied with both experimental and numerical approaches. The experiments are carried out with Particle Image Velocimetry (PIV) and two types of wake transition processes, namely the transition from a drag-indicative wake to a thrust-indicative wake and that from the symmetric wake to the asymmetric wake are distinguished. The numerical results from the developed SD solver agree well with the experimental results. It is numerically found that the deflective direction of the asymmetric wake is determined by the initial conditions, e.g. initial phase angle. As most insects use thin wings (i. e., wing thickness is only a few percent of the chord length) in flapping flight, the effects of airfoil thickness on thrust generation are numerically investigated by simulating
NASA Technical Reports Server (NTRS)
Harris, C. D.
1974-01-01
Refinements in a 10 percent thick supercritical airfoil produced improvements in the overall drag characteristics at normal force coefficients from about 0.30 to 0.65 compared with earlier supercritical airfoils which were developed for a normal force coefficient of 0.7. The drag divergence Mach number of the improved supercritical airfoil (airfoil 26a) varied from approximately 0.82 at a normal force coefficient to of 0.30, to 0.78 at a normal force coefficient of 0.80 with no drag creep evident. Integrated section force and moment data, surface pressure distributions, and typical wake survey profiles are presented.
Analysis of high Reynolds numbers effects on a wind turbine airfoil using 2D wind tunnel test data
NASA Astrophysics Data System (ADS)
Pires, O.; Munduate, X.; Ceyhan, O.; Jacobs, M.; Snel, H.
2016-09-01
The aerodynamic behaviour of a wind turbine airfoil has been measured in a dedicated 2D wind tunnel test at the DNW High Pressure Wind Tunnel in Gottingen (HDG), Germany. The tests have been performed on the DU00W212 airfoil at different Reynolds numbers: 3, 6, 9, 12 and 15 million, and at low Mach numbers (below 0.1). Both clean and tripped conditions of the airfoil have been measured. An analysis of the impact of a wide Reynolds number variation over the aerodynamic characteristics of this airfoil has been performed.
NASA Technical Reports Server (NTRS)
Flemming, Robert J.
1984-01-01
Five full scale rotorcraft airfoils were tested in the NASA Ames Eleven-Foot Transonic Wind Tunnel for full scale Reynolds numbers at Mach numbers from 0.3 to 1.07. The models, which spanned the tunnel from floor to ceiling, included two modern baseline airfoils, the SC1095 and SC1094 R8, which have been previously tested in other facilities. Three advanced transonic airfoils, designated the SSC-A09, SSC-A07, and SSC-B08, were tested to confirm predicted performance and provide confirmation of advanced airfoil design methods. The test showed that the eleven-foot tunnel is suited to two-dimensional airfoil testing. Maximum lift coefficients, drag coefficients, pitching moments, and pressure coefficient distributions are presented. The airfoil analysis codes agreed well with the data, with the Grumman GRUMFOIL code giving the best overall performance correlation.
Cooled airfoil in a turbine engine
Vitt, Paul H; Kemp, David A; Lee, Ching-Pang; Marra, John J
2015-04-21
An airfoil in a gas turbine engine includes an outer wall and an inner wall. The outer wall includes a leading edge, a trailing edge opposed from the leading edge in a chordal direction, a pressure side, and a suction side. The inner wall is coupled to the outer wall at a single chordal location and includes portions spaced from the pressure and suction sides of the outer wall so as to form first and second gaps between the inner wall and the respective pressure and suction sides. The inner wall defines a chamber therein and includes openings that provide fluid communication between the respective gaps and the chamber. The gaps receive cooling fluid that provides cooling to the outer wall as it flows through the gaps. The cooling fluid, after traversing at least substantial portions of the gaps, passes into the chamber through the openings in the inner wall.
Unsteady Flows in Axial Turbomachines
NASA Technical Reports Server (NTRS)
Marble, F. E.; Rannie, W. D.
1957-01-01
Of the various unsteady flows that occur in axial turbomachines certain asymmetric disturbances, of wave length large in comparison with blade spacing, have become understood to a certain extent. These disturbances divide themselves into two categories: self-induced oscillations and force disturbances. A special type of propagating stall appears as a self-induced disturbance; an asymmetric velocity profile introduced at the compressor inlet constitutes a forced disturbance. Both phenomena have been treated from a unified theoretical point of view in which the asymmetric disturbances are linearized and the blade characteristics are assumed quasi-steady. Experimental results are in essential agreement with this theory wherever the limitations of the theory are satisfied. For the self-induced disturbances and the more interesting examples of the forced disturbances, the dominant blade characteristic is the dependence of total pressure loss, rather than the turning angle, upon the local blade inlet angle.
Unsteady jet flow computation towards noise prediction
NASA Technical Reports Server (NTRS)
Soh, Woo-Yung
1994-01-01
An attempt has been made to combine a wave solution method and an unsteady flow computation to produce an integrated aeroacoustic code to predict far-field jet noise. An axisymmetric subsonic jet is considered for this purpose. A fourth order space accurate Pade compact scheme is used for the unsteady Navier-Stokes solution. A Kirchhoff surface integral for the wave equation is employed through the use of an imaginary surface which is a circular cylinder enclosing the jet at a distance. Information such as pressure and its time and normal derivatives is provided on the surface. The sound prediction is performed side by side with the jet flow computation. Retarded time is also taken into consideration since the cylinder body is not acoustically compact. The far-field sound pressure has the directivity and spectra show that low frequency peaks shift toward higher frequency region as the observation angle increases from the jet flow axis.
Detailed transonic flow field measurements about a supercritical airfoil section
NASA Technical Reports Server (NTRS)
Hurley, F. X.; Spaid, F. W.; Roos, F. W.; Stivers, L. S., Jr.; Bandettini, A.
1975-01-01
The transonic flow field about a Whitcomb-type supercritical airfoil profile was measured in detail. In addition to the usual surface pressure distributions and wake surveys, schlieren photographs were taken and velocity vector profiles were determined in the upper surface boundary layer and in the near wake. Spanwise variations in the measured pressures were also determined. The data are analyzed with the aid of an inviscid transonic finite-difference computer program as well as with boundary layer modeling and calculation schemes.
Simulation of shock wave buffet and its suppression on an OAT15A supercritical airfoil by IDDES
NASA Astrophysics Data System (ADS)
Huang, JingBo; Xiao, ZhiXiang; Liu, Jian; Fu, Song
2012-02-01
In the present paper, extremely unsteady shock wave buffet induced by strong shock wave/boundary-layer interactions (SWBLI) on the upper surface of an OAT15A supercritical airfoil at Mach number of 0.73 and angle of attack of 3.5 degrees is first numerically simulated by IDDES, one of the most advanced RANS/LES hybrid methods. The results imply that conventional URANS methods are unable to effectively predict the buffet phenomenon on the wing surface; IDDES, which involves more flow physics, predicted buffet phenomenon. Some complex flow phenomena are predicted and demonstrated, such as periodical oscillations of shock wave in the streamwise direction, strong shear layer detached from the shock wave due to SWBLI and plenty of small scale structures broken down by the shear layer instability and in the wake. The root mean square (RMS) of fluctuating pressure coefficients and streamwise range of shock wave oscillation reasonably agree with experimental data. Then, two vortex generators (VG) both with an inclination angle of 30 degrees to the main flow directions are mounted in front of the shock wave region on the upper surface to suppress shock wave buffet. The results show that shock wave buffet can be significantly suppressed by VGs, the RMS level of pressure in the buffet region is effectively reduced, and averaged shock wave position is obviously pushed downstream, resulting in increased total lift.
Effects of grit roughness and pitch oscillations on the S814 airfoil
Janiszewska, J.M.; Ramsay, R.R.; Hoffmann, M.J.; Gregorek, G.M.
1996-07-01
Horizontal-axis wind turbine rotors experience unsteady aerodynamics when the rotor is yawed, when rotor blades pass through the support tower wake, and when the wind is gusting. An understanding of this unsteady behavior is necessary to assist in the design of new rotor airfoils. The rotors also experience performance degradation due to surface roughness. These surface irregularities are due to the accumulation of insect debris, ice, and/or the aging process. Wind tunnel studies that examine both the steady and unsteady behavior of airfoils can help define pertinent flow phenomena, and the resultant data can also be used to validate analytical computer codes. An S814 airfoil model was tested in The Ohio State University Aeronautical and Astronautical Research Laboratory (OSU/AARL) 3 X 5 subsonic wind tunnel (3 X 5) under steady flow with both stationary model conditions and pitch oscillations. To study the extent of performance loss due to surface roughness, a leading edge grit roughness pattern (LEGR) was used to simulate leading edge contamination. After baseline cases were completed, the LEGR was applied for both steady state and model pitch oscillation cases. The Reynolds numbers for steady state conditions were 0.75, 1, 1.25 and 1.5 million, while the angle of attack ranged from -20{degrees} to +40{degrees}. While the model underwent pitch oscillations, data were acquired at Reynolds numbers of 0.75, 1, 1.25, and 1.5 million, at frequencies of 0.6, 1.2, and 1.8 Hz. Two sine wave forcing functions {+-}5.5{degrees} and {+-}10{degrees}, were used; at mean angles of attack of 8{degrees}, 14{degrees}, and 20{degrees}. For purposes herein, any reference to unsteady conditions means the model was in pitch oscillation.
Effects of grit roughness and pitch oscillations on the S801 airfoil
Ramsay, R.R.; Hoffman, M.J.; Gregorek, G.M.
1996-01-01
Horizontal axis wind turbine rotors experience unsteady aerodynamics due to wind shear when the rotor is yawed, when rotor blades pass through the support tower wake, and when the wind is gusting. An understanding of this unsteady behavior is necessary to assist in the calculation of rotor performance and loads. The rotors also experience performance degradation due to surface roughness. These surface irregularities are due to the accumulation of insect debris, ice, and the aging process. Wind tunnel studies that examine both the steady and unsteady behavior of airfoils can help define pertinent flow phenomena, and the resultant data can be used to validate analytical computer codes. A S801 airfoil model was tested in The Ohio State University Aeronautical and Astronautical Research Laboratory (OSU/AARL) 3x5 subsonic wind tunnel (3x5) under steady flow and stationary model conditions, as well as with the model undergoing pitch oscillations. To study the possible extent of performance loss due to surface roughness, a standard grit pattern (LEGR) was used to simulate leading edge contamination. After baseline cases were completed, the LEGR was applied for both steady state and model pitch oscillation cases. The Reynolds numbers used for steady state conditions were 0.75, 1, 1.25, and 1.5 million, while the angle of attack ranged from -20{degrees} to +40{degrees}. With the model undergoing pitch oscillations, data were acquired at Reynolds numbers of 0.75, 1, 1.25, and 1.4 million, at frequencies of 0.6, 1.2, and 1.8 Hz. Two sine wave forcing functions were used, {plus_minus} 5.5 {degrees}and {plus_minus} 10{degrees}, at mean angles of attack of 8{degrees} 14{degrees} and 20{degrees} For purposes herein, any reference to unsteady conditions means that the airfoil model was in pitch oscillation about the quarter chord.
NASA Technical Reports Server (NTRS)
Kohl, F. J.
1982-01-01
The methodology to predict deposit evolution (deposition rate and subsequent flow of liquid deposits) as a function of fuel and air impurity content and relevant aerodynamic parameters for turbine airfoils is developed in this research. The spectrum of deposition conditions encountered in gas turbine operations includes the mechanisms of vapor deposition, small particle deposition with thermophoresis, and larger particle deposition with inertial effects. The focus is on using a simplified version of the comprehensive multicomponent vapor diffusion formalism to make deposition predictions for: (1) simple geometry collectors; and (2) gas turbine blade shapes, including both developing laminar and turbulent boundary layers. For the gas turbine blade the insights developed in previous programs are being combined with heat and mass transfer coefficient calculations using the STAN 5 boundary layer code to predict vapor deposition rates and corresponding liquid layer thicknesses on turbine blades. A computer program is being written which utilizes the local values of the calculated deposition rate and skin friction to calculate the increment in liquid condensate layer growth along a collector surface.
Transonic pressure distributions on a rectangular supercritical wing oscillating in pitch
NASA Technical Reports Server (NTRS)
Ricketts, R. H.; Sandford, M. C.; Seidel, D. A.; Watson, J. J.
1983-01-01
Steady and unsteady aerodynamic data were measured on a rectangular wing with a 12 percent thick supercritical airfoil mounted in the NASA Langley Transonic Dynamics Tunnel. The wing was oscillated in pitch to generate the unsteady aerodynamic data. The purpose of the wind-tunnel test was to measure data for use in the development and assessment of transonic analytical codes. The effects on the wing pressure distributions of Mach number, mean angle of attack, and oscillation frequency and amplitude were measured. Results from the newly-developed XTRAN3S program (a nonlinear transonic small disturbance code) and from the RHOIV program (a linear lifting surface kernel function code) were compared to measured data for a Mach number of 0.7 and for oscillation frequencies ranging from 0 to 20 Hz. The XTRAN3S steady and unsteady results agreed fairly well with the measured results. The RHOIV unsteady-result agreement was fair but, of course, did not predict shock effects.
Transonic pressure distributions on a rectangular supercritical wing oscillating in pitch
NASA Technical Reports Server (NTRS)
Ricketts, R. H.; Sandford, M. C.; Seidel, D. A.; Watson, J. J.
1983-01-01
Steady and unsteady aerodynamic data were measured on a rectangular wing with a 12 percent thick supercritical airfoil mounted in the NASA Langley Transonic Dynamics Tunnel. The wing was oscillated in pitch to generate the unsteady aerodynamic data. The purpose of the wind-tunnel test was to measure data for use in the development and assessment of transonic analytical codes. The effects on the wing pressure distributions of Mach number, mean angle of attack, and oscillation frequency and amplitude were measured. Results from the newly-developed XTRAN3S program (a non-linear transonic small disturbance code) and from the RHOIV program (a linear lifting surface kernel function code) were compared to measured data for a Mach number of 0.7 and for oscillation frequencies ranging from 0 to 20 Hz. The XTRAN3S steady and unsteady results agreed fairly well with the measured data. The RHOIV unsteady-result agreement was fair but, of course, did not predict shock effects.
Bayesian inference of nonlinear unsteady aerodynamics from aeroelastic limit cycle oscillations
NASA Astrophysics Data System (ADS)
Sandhu, Rimple; Poirel, Dominique; Pettit, Chris; Khalil, Mohammad; Sarkar, Abhijit
2016-07-01
A Bayesian model selection and parameter estimation algorithm is applied to investigate the influence of nonlinear and unsteady aerodynamic loads on the limit cycle oscillation (LCO) of a pitching airfoil in the transitional Reynolds number regime. At small angles of attack, laminar boundary layer trailing edge separation causes negative aerodynamic damping leading to the LCO. The fluid-structure interaction of the rigid, but elastically mounted, airfoil and nonlinear unsteady aerodynamics is represented by two coupled nonlinear stochastic ordinary differential equations containing uncertain parameters and model approximation errors. Several plausible aerodynamic models with increasing complexity are proposed to describe the aeroelastic system leading to LCO. The likelihood in the posterior parameter probability density function (pdf) is available semi-analytically using the extended Kalman filter for the state estimation of the coupled nonlinear structural and unsteady aerodynamic model. The posterior parameter pdf is sampled using a parallel and adaptive Markov Chain Monte Carlo (MCMC) algorithm. The posterior probability of each model is estimated using the Chib-Jeliazkov method that directly uses the posterior MCMC samples for evidence (marginal likelihood) computation. The Bayesian algorithm is validated through a numerical study and then applied to model the nonlinear unsteady aerodynamic loads using wind-tunnel test data at various Reynolds numbers.
An experimental study of transonic flow about a supercritical airfoil
NASA Technical Reports Server (NTRS)
Spaid, F. W.; Dahlin, J. A.; Bachalo, W. D.; Stivers, L. S., Jr.
1983-01-01
A series of experiments was conducted on flow fields about two airfoil models whose sections are slight modifications of the original Whitcomb supercritical airfoil section. Data obtained include surface static-pressure distributions, far-wake surveys, oil-flow photographs, pitot-pressure surveys in the viscous regions, and holographic interferograms. These data were obtained for different combinations of lift coefficient and free-stream Mach number, which included both subcritical cases and flows with upper-surface shock waves. The availability of both pitot-pressure data and density data from interferograms allowed determination of flow-field properties in the vicinity of the trailing edge and in the wake without recourse to any assumptions about the local static pressure. The data show that significant static-pressure gradients normal to viscous layers exist in this region, and that they persist to approximately 10% chord downstream of the trailing edge. Comparisons are made between measured boundary-layer properties and results from boundary-layer computations that employed measured static-pressure distributions, as well as comparisons between data and results of airfoil flow-field computations.
Shape optimization of corrugated airfoils
NASA Astrophysics Data System (ADS)
Jain, Sambhav; Bhatt, Varun Dhananjay; Mittal, Sanjay
2015-12-01
The effect of corrugations on the aerodynamic performance of a Mueller C4 airfoil, placed at a 5° angle of attack and Re=10{,}000, is investigated. A stabilized finite element method is employed to solve the incompressible flow equations in two dimensions. A novel parameterization scheme is proposed that enables representation of corrugations on the surface of the airfoil, and their spontaneous appearance in the shape optimization loop, if indeed they improve aerodynamic performance. Computations are carried out for different location and number of corrugations, while holding their height fixed. The first corrugation causes an increase in lift and drag. Each of the later corrugations leads to a reduction in drag. Shape optimization of the Mueller C4 airfoil is carried out using various objective functions and optimization strategies, based on controlling airfoil thickness and camber. One of the optimal shapes leads to 50 % increase in lift coefficient and 23 % increase in aerodynamic efficiency compared to the Mueller C4 airfoil.
Experimental studies of flow separation and stalling on a two-dimensional airfoil at low speeds
NASA Technical Reports Server (NTRS)
Seetharam, H. C.; Wentz, W. H., Jr.
1975-01-01
Detailed measurements of flow fields associated with low-speed turbulent boundary layers were made for the 17% thick GA(W)-1 airfoil section at nominal angles of attack of 10 deg, 14 deg, and 18 deg, Reynolds number 2.2 x 10, and Mach number 0.13. The data include pressure and velocity surveys of the pre- and post-separated regions on the airfoil and the associated wake. The boundary layer characteristics including regions of separation on the airfoil are also presented.
NASA Technical Reports Server (NTRS)
Harvey, William D.; Harris, Charles D.; Brooks, Cuyler W., Jr.
1989-01-01
A swept, supercritical laminar flow control (LFC) airfoil designated NASA SCLFC(1)-0513F was tested at subsonic and transonic speeds in the NASA Langley eight-foot Transonic Pressure Tunnel. This paper examines Tollmien-Schlichting and crossflow disturbance amplification for this airfoil using the linear stability method. The design methodology using linear stability analysis is evaluated and the results of the incompressible and compressible methods are compared. Experimental data on the swept, supercritical LFC airfoil and reference wind tunnel and flight results are used to correlate and evaluate the N-factor method for transition prediction over a speed range M(infinity) from zero to one.
Advanced technology airfoil research, volume 2. [conferences
NASA Technical Reports Server (NTRS)
1979-01-01
A comprehensive review of airfoil research is presented. The major thrust of the research is in three areas: development of computational aerodynamic codes for airfoil analysis and design, development of experimental facilities and test techniques, and all types of airfoil applications.
Root region airfoil for wind turbine
Tangler, James L.; Somers, Dan M.
1995-01-01
A thick airfoil for the root region of the blade of a wind turbine. The airfoil has a thickness in a range from 24%-26% and a Reynolds number in a range from 1,000,000 to 1,800,000. The airfoil has a maximum lift coefficient of 1.4-1.6 that has minimum sensitivity to roughness effects.
Application of an unsteady Navier-Stokes solver to transonic turbine design
NASA Technical Reports Server (NTRS)
Rangwalla, Akil A.; Madavan, Nateri K.; Johnson, Paul D.
1991-01-01
This study presents a numerical evaluation of the performance of the first stage of a new-generation turbine design. The numerical method solves the two-dimensional Navier-Stokes equations using a system of patched grids. Three-dimensional effects of stream-tube contraction are also modeled. The study focuses on the effects of axial gap variation on the unsteady rotor-stator interactions and on stage performance. Results are presented for three different axial gaps. The results indicate that the unsteady interactions can be very large in this design. These interactions affect not only the stage efficiency but also substantially alter the time-averaged features of the flow. In particular, for the case of the smallest axial gap, it was found that there was an unsteady shock on the stator suction surface which spanned the gap region and impinged upon the moving rotor airfoils.
Development of a linearized unsteady aerodynamic analysis for cascade gust response predictions
NASA Technical Reports Server (NTRS)
Verdon, Joseph M.; Hall, Kenneth C.
1990-01-01
A method for predicting the unsteady aerodynamic response of a cascade of airfoils to entropic, vortical, and acoustic gust excitations is being developed. Here, the unsteady flow is regarded as a small perturbation of a nonuniform isentropic and irrotational steady background flow. A splitting technique is used to decompose the linearized unsteady velocity into rotational and irrotational parts leading to equations for the complex amplitudes of the linearized unsteady entropy, rotational velocity, and velocity potential that are coupled only sequentially. The entropic and rotational velocity fluctuations are described by transport equations for which closed-form solutions in terms of the mean-flow drift and stream functions can be determined. The potential fluctuation is described by an inhomogeneous convected wave equation in which the source term depends on the rotational velocity field, and is determined using finite-difference procedures. The analytical and numerical techniques used to determine the linearized unsteady flow are outlined. Results are presented to indicate the status of the solution procedure and to demonstrate the impact of blade geometry and mean blade loading on the aerodynamic response of cascades to vortical gust excitations. The analysis described herein leads to very efficient predictions of cascade unsteady aerodynamic response phenomena making it useful for turbomachinery aeroelastic and aeroacoustic design applications.
The importance of the unsteady Kutta condition when modelling gust-aerofoil interaction
NASA Astrophysics Data System (ADS)
Ayton, Lorna J.; Gill, J. R.; Peake, N.
2016-09-01
The Kutta condition is applied to aerofoils with sharp trailing edges to allow for viscous effects to be considered within a simplified system of equations that are inviscid. This paper discusses in detail the inclusion of an unsteady Kutta condition at a sharp trailing edge during gust-aerofoil interaction and illustrates how the analytic solution for the far-field noise generated by this interaction changes if the unsteady Kutta condition is neglected, or more precisely, if the unsteady pressure is permitted to be singular at the trailing edge. The analytic solution, both with and without the unsteady Kutta condition, is compared with numerical results that have no imposed unsteady Kutta condition. Importantly the results agree well only when the unsteady Kutta condition is neglected in the analytic solution. This paper highlights where the far-field acoustics are most affected by neglecting the unsteady Kutta condition for a variety of singularities that can occur in the unsteady pressure at the trailing edge and shows that results permitting different behaviour in the unsteady surface pressure at the trailing edge could give significantly different far-field noise predictions, even though the surface pressure elsewhere along the aerofoil surface agrees with benchmark solutions.
Performance measurements of an airfoil at low Reynolds numbers
NASA Technical Reports Server (NTRS)
Mcghee, Robert J.; Walker, Betty S.
1989-01-01
Performance characteristics of an Eppler 387 airfoil using both direct (force) and indirect (pressure) measurement techniques have been obtained at Reynolds numbers from 60,000 to 460,000 in the Langley Low-Turbulence Pressure Tunnel. Lift, drag, and pitching-moment data were obtained from two internally-mounted strain-gage balances specifically designed for small aerodynamic loads. Comparisons of these results with data from a pressure model of an Eppler 387 airfoil are included. Drag data for both models using the wake traverse method are compared with the balance data. Oil flow visualization and surface mounted hot-film sensors were used to determine laminar-separation and turbulent-reattachment locations. Problems associated with obtaining accurate wind-tunnel data at low Reynolds numbers are discussed.
Advanced Airfoils Boost Helicopter Performance
NASA Technical Reports Server (NTRS)
2007-01-01
Carson Helicopters Inc. licensed the Langley RC4 series of airfoils in 1993 to develop a replacement main rotor blade for their Sikorsky S-61 helicopters. The company's fleet of S-61 helicopters has been rebuilt to include Langley's patented airfoil design, and the helicopters are now able to carry heavier loads and fly faster and farther, and the main rotor blades have twice the previous service life. In aerial firefighting, the performance-boosting airfoils have helped the U.S. Department of Agriculture's Forest Service control the spread of wildfires. In 2003, Carson Helicopters signed a contract with Ducommun AeroStructures Inc., to manufacture the composite blades for Carson Helicopters to sell
Dynamic stall on a pitching and surging airfoil
NASA Astrophysics Data System (ADS)
Dunne, Reeve; McKeon, Beverley J.
2015-08-01
Vertical axis wind turbine blades undergo dynamic stall due to the large angle of attack variation they experience during a turbine rotation. The flow over a single blade was modeled using a sinusoidally pitching and surging airfoil in a non-rotating frame with a constant freestream flow at a mean chord Reynolds number of . Two-dimensional, time-resolved velocity fields were acquired using particle image velocimetry. Vorticity contours were used to visualize shear layer and vortex activity. A low-order model of dynamic stall was developed using dynamic mode decomposition, from which primary and secondary dynamic separation modes were identified. The interaction between these two modes was able to capture the physics of dynamic stall and as such can be extended to other turbine configurations and problems in unsteady aerodynamics. Results from the linear pitch/surge frame are extrapolated to the rotating VAWT frame to investigate the behavior of identified flow structures.
NASA Technical Reports Server (NTRS)
Crivellini, A.; Golubev, V.; Mankbadi, R.; Scott, J. R.; Hixon, R.; Povinelli, L.; Kiraly, L. James (Technical Monitor)
2002-01-01
The nonlinear response of symmetric and loaded airfoils to an impinging vortical gust is investigated in the parametric space of gust dimension, intensity, and frequency. The study, which was designed to investigate the validity limits for a linear analysis, is implemented by applying a nonlinear high-order prefactored compact code and comparing results with linear solutions from the GUST3D frequency-domain solver. Both the unsteady aerodynamic and acoustic gust responses are examined.
Hook nozzle arrangement for supporting airfoil vanes
Shaffer, James E.; Norton, Paul F.
1996-01-01
A gas turbine engine's nozzle structure includes a nozzle support ring, a plurality of shroud segments, and a plurality of airfoil vanes. The plurality of shroud segments are distributed around the nozzle support ring. Each airfoil vane is connected to a corresponding shroud segment so that the airfoil vanes are also distributed around the nozzle support ring. Each shroud segment has a hook engaging the nozzle support ring so that the shroud segments and corresponding airfoil vanes are supported by the nozzle support ring. The nozzle support ring, the shroud segments, and the airfoil vanes may be ceramic.
Airfoil shape for a turbine nozzle
Burdgick, Steven Sebastian; Patik, Joseph Francis; Itzel, Gary Michael
2002-01-01
A first-stage nozzle vane includes an airfoil having a profile according to Table I. The annulus profile of the hot gas path is defined in conjunction with the airfoil profile and the profile of the inner and outer walls by the Cartesian coordinate values given in Tables I and II, respectively. The airfoil is a three-dimensional bowed design, both in the airfoil body and in the trailing edge. The airfoil is steam and air-cooled by flowing cooling mediums through cavities extending in the vane between inner and outer walls.
Hook nozzle arrangement for supporting airfoil vanes
Shaffer, J.E.; Norton, P.F.
1996-02-20
A gas turbine engine`s nozzle structure includes a nozzle support ring, a plurality of shroud segments, and a plurality of airfoil vanes. The plurality of shroud segments are distributed around the nozzle support ring. Each airfoil vane is connected to a corresponding shroud segment so that the airfoil vanes are also distributed around the nozzle support ring. Each shroud segment has a hook engaging the nozzle support ring so that the shroud segments and corresponding airfoil vanes are supported by the nozzle support ring. The nozzle support ring, the shroud segments, and the airfoil vanes may be ceramic. 8 figs.
Aeroelastic airfoil smart spar
NASA Technical Reports Server (NTRS)
Greenhalgh, Skott; Pastore, Christopher M.; Garfinkle, Moishe
1993-01-01
Aircraft wings and rotor-blades are subject to undesirable bending and twisting excursions that arise from unsteady aerodynamic forces during high speed flight, abrupt maneuvers, or hard landings. These bending excursions can range in amplitude from wing-tip flutter to failure. A continuous-filament construction 'smart' laminated composite box-beam spar is described which corrects itself when subject to undesirable bending excursions or flutter. The load-bearing spar is constructed so that any tendency for the wing or rotor-blade to bend from its normal position is met by opposite twisting of the spar to restore the wing to its normal position. Experimental and theoretical characterization of these spars was made to evaluate the torsion-flexure coupling associated with symmetric lay-ups. The materials used were uniweave AS-4 graphite and a matrix comprised of Shell 8132 resin and U-40 hardener. Experimental tests were conducted on five spars to determine spar twist and bend as a function of load for 0, 17, 30, 45 and 60 deg fiber angle lay-ups. Symmetric fiber lay-ups do exhibit torsion-flexure couplings. Predictions of the twist and bend versus load were made for different fiber orientations in laminated spars using a spline function structural analysis. The analytical results were compared with experimental results for validation. Excellent correlation between experimental and analytical values was found.
NASA Astrophysics Data System (ADS)
Sahoo, Dipankar
Improved basic understanding, predictability, and controllability of vortex-dominated and unsteady aerodynamic flows are important in enhancement of the performance of next generation helicopters. The primary objective of this research project was improved understanding of the fundamental vorticity and turbulent flow physics for a dynamically stalling airfoil at realistic helicopter flight conditions. An experimental program was performed on a large-scale (C = 0.45 m) dynamically pitching NACA 0012 wing operating in the Texas A&M University large-scale wind tunnel. High-resolution particle image velocimetry data were acquired on the first 10-15% of the wing. Six test cases were examined including the unsteady (k>0) and steady (k=0) conditions. The relevant mechanical, shear and turbulent time-scales were all of comparable magnitude, which indicated that the flow was in a state of mechanical non-equilibrium, and the expected flow separation and reattachment hystersis was observed. Analyses of the databases provided new insights into the leading-edge Reynolds stress structure and the turbulent transport processes. Both of which were previously uncharacterized. During the upstroke motion of the wing, a bubble structure formed in the leading-edge Reynolds shear stress. The size of the bubble increased with increasing angle-of-attack before being diffused into a shear layer at full separation. The turbulent transport analyses indicated that the axial stress production was positive, where the transverse production was negative. This implied that axial turbulent stresses were being produced from the axial component of the mean flow. A significant portion of the energy was transferred to the transverse stress through the pressure-strain redistribution, and then back to the transverse mean flow through the negative transverse production. An opposite trend was observed further downstream of this region.
Control of Flow Separation Using Adaptive Airfoils
NASA Technical Reports Server (NTRS)
Chandrasekhara, M. S.; Wilder, M. C.; Carr, L. W.; Davis, Sanford S. (Technical Monitor)
1996-01-01
A novel way of controlling flow separation is reported. The approach involves using an adaptive airfoil geometry that changes its leading edge shape to adjust to the instantaneous flow at high angles of attack such that the flow over it remains attached. In particular, a baseline NACA 0012 airfoil, whose leading edge curvature could be changed dynamically by 400% was tested under quasi-steady compressible flow conditions. A mechanical drive system was used to produce a rounded leading edge to reduce the strong local flow acceleration around its nose and thus reduce the strong adverse pressure gradient that follows such a rapid acceleration. Tests in steady flow showed that at M = 0.3, the flow separated at about 14 deg. angle of attack for the NACA 0012 profile but could be kept attached up to an angle of about 18 deg by changing the nose curvature. No significant hysteresis effects were observed; the flow could be made to reattach from its separated state at high angles by changing the leading edge curvature.
Aerodynamic characteristics of an oscillating airfoil
NASA Astrophysics Data System (ADS)
Wickens, R. H.
1986-03-01
Results are reported from wind tunnel tests to study the effects of dynamic aerodynamics on the efficiency of a NACA 0018 airfoil used on a Darreius vertical axis wind turbine (VAWT). The topic is of interest because of uncontrolled pitching which occurs during operation and which produces stall, turbulence and separation effects that reduce efficiency. Present stream-tube theory and axial momentum models are not applicable in the unstable regimes. The wind tunnel tests were conducted with a 45 m/sec flow with an Re of 1.5 million. The situation mimicked typical wind turbine operational conditions. The airfoil was mounted on a hydraulic actuator to allow it to rotate about its quarter-chord location and to control the extent and frequency of oscillations. Data were also gathered on the performance in a steady flow for comparative purposes. Summary data are provided on the static and total pressures over a complete cycle of oscillation, and related to the angles of attack, time of onset of stall, and the lift and drag coefficients. The limitations of the study with regard to the absence of consideration of the flow acceleration experienced by an advancing blade are noted.
Flow-induced noise of a wall-mounted finite airfoil at low-to-moderate Reynolds number
NASA Astrophysics Data System (ADS)
Moreau, Danielle J.; Prime, Zebb; Porteous, Ric; Doolan, Con J.; Valeau, Vincent
2014-12-01
This paper presents an experimental investigation of the flow-induced noise created by a wall-mounted finite airfoil at low-to-moderate Reynolds number and zero angle of attack. Far-field noise measurements have been taken at a single observer location and with two perpendicular microphone arrays in an anechoic wind tunnel at Reynolds numbers of Rec=9.2×104-1.6×105, based on chord, and for a variety of airfoil aspect ratios (length to chord ratio of L/C=0.2-2, corresponding to length to thickness ratio of L/T=1.7-16.7). Additionally, surface oil-film visualisation images and unsteady velocity measurements taken in the near trailing edge wake are related to far-field noise measurements to determine the flow mechanisms responsible for noise generation. The results show that the wall-mounted finite airfoil radiates noise similar to a two-dimensional airfoil when L/T>8.3. Despite the incoming boundary layer height at the junction being 1.30≤δ/T≤1.46, junction and tip flow suppresses tonal noise production for airfoil's up to L/T=8.3 at Rec=9.2×104-1.2×105. Trailing edge noise is found to be the dominant airfoil noise generation mechanism at frequencies above 1 kHz with the position of the noise source along the trailing edge determined by the proportion of the airfoil span influenced by flow at the airfoil-wall junction.
A two element laminar flow airfoil optimized for cruise. M.S. Thesis
NASA Technical Reports Server (NTRS)
Steen, Gregory Glen
1994-01-01
Numerical and experimental results are presented for a new two-element, fixed-geometry natural laminar flow airfoil optimized for cruise Reynolds numbers on the order of three million. The airfoil design consists of a primary element and an independent secondary element with a primary to secondary chord ratio of three to one. The airfoil was designed to improve the cruise lift-to-drag ratio while maintaining an appropriate landing capability when compared to conventional airfoils. The airfoil was numerically developed utilizing the NASA Langley Multi-Component Airfoil Analysis computer code running on a personal computer. Numerical results show a nearly 11.75 percent decrease in overall wing drag with no increase in stall speed at sailplane cruise conditions when compared to a wing based on an efficient single element airfoil. Section surface pressure, wake survey, transition location, and flow visualization results were obtained in the Texas A&M University Low Speed Wind Tunnel. Comparisons between the numerical and experimental data, the effects of the relative position and angle of the two elements, and Reynolds number variations from 8 x 10(exp 5) to 3 x 10(exp 6) for the optimum geometry case are presented.
SiC/SiC Leading Edge Turbine Airfoil Tested Under Simulated Gas Turbine Conditions
NASA Technical Reports Server (NTRS)
Robinson, R. Craig; Hatton, Kenneth S.
1999-01-01
Silicon-based ceramics have been proposed as component materials for use in gas turbine engine hot-sections. A high pressure burner rig was used to expose both a baseline metal airfoil and ceramic matrix composite leading edge airfoil to typical gas turbine conditions to comparatively evaluate the material response at high temperatures. To eliminate many of the concerns related to an entirely ceramic, rotating airfoil, this study has focused on equipping a stationary metal airfoil with a ceramic leading edge insert to demonstrate the feasibility and benefits of such a configuration. Here, the idea was to allow the SiC/SiC composite to be integrated as the airfoil's leading edge, operating in a "free-floating" or unrestrained manner. and provide temperature relief to the metal blade underneath. The test included cycling the airfoils between simulated idle, lift, and cruise flight conditions. In addition, the airfoils were air-cooled, uniquely instrumented, and exposed to the same internal and external conditions, which included gas temperatures in excess of 1370 C (2500 F). Results show the leading edge insert remained structurally intact after 200 simulated flight cycles with only a slightly oxidized surface. The instrumentation clearly suggested a significant reduction (approximately 600 F) in internal metal temperatures as a result of the ceramic leading edge. The object of this testing was to validate the design and analysis done by Materials Research and Design of Rosemont, PA and to determine the feasibility of this design for the intended application.
NASA Technical Reports Server (NTRS)
Hess, R. W.; Wynne, E. C.; Cazier, F. W.
1982-01-01
Pressures were measured with Freon as the test medium. Data taken at M = 0.9 is presented for static and oscillatory deflections of the trailing edge control surface and for the wing in pitch. Comparisons of the static measured data are made with results computed using the Bailey-Ballhaus small disturbance code.
Unsteady Turbopump Flow Simulations
NASA Technical Reports Server (NTRS)
Centin, Kiris C.; Kwak, Dochan
2001-01-01
The objective of the current effort is two-fold: 1) to provide a computational framework for design and analysis of the entire fuel supply system of a liquid rocket engine; and 2) to provide high-fidelity unsteady turbopump flow analysis capability to support the design of pump sub-systems for advanced space transportation vehicle. Since the space launch systems in the near future are likely to involve liquid propulsion system, increasing the efficiency and reliability of the turbopump components is an important task. To date, computational tools for design/analysis of turbopump flow are based on relatively lower fidelity methods. Unsteady, three-dimensional viscous flow analysis tool involving stationary and rotational components for the entire turbopump assembly has not been available, at least, for real-world engineering applications. Present effort is an attempt to provide this capability so that developers of the vehicle will be able to extract such information as transient flow phenomena for start up, impact of non-uniform inflow, system vibration and impact on the structure. Those quantities are not readily available from simplified design tools. In this presentation, the progress being made toward complete turbo-pump simulation capability for a liquid rocket engine is reported. Space Shuttle Main Engine (SSME) turbo-pump is used as a test case for the performance evaluation of the hybrid MPI/Open-MP and MLP versions of the INS3D code. Relative motion of the grid system for rotor-stator interaction was obtained by employing overset grid techniques. Time-accuracy of the scheme has been evaluated by using simple test cases. Unsteady computations for SSME turbopump, which contains 106 zones with 34.5 Million grid points, are currently underway on Origin 2000 systems at NASA Ames Research Center. Results from these time-accurate simulations with moving boundary capability and the performance of the parallel versions of the code will be presented.
An unsteady rotor/fuselage interaction method
NASA Technical Reports Server (NTRS)
Egolf, T. Alan; Lorber, Peter F.
1987-01-01
An analytical method has been developed to treat unsteady helicopter rotor, wake, and fuselage interaction aerodynamics. An existing lifting line/prescribed wake rotor analysis and a source panel fuselage analysis were modified to predict vibratory fuselage airloads. The analyses were coupled through the induced flow velocities of the rotor and wake on the fuselage and the fuselage on the rotor. A prescribed displacement technique was used to distort the rotor wake about the fuselage. Sensitivity studies were performed to determine the influence of wake and body geometry on the computed airloads. Predicted and measured mean and unsteady pressures on a cylindrical body in the wake of a two-bladed rotor were compared. Initial results show good qualitative agreement.
Wind tunnel evaluation of a truncated NACA 64-621 airfoil for wind turbine applications
Law, S.P.; Gregorek, G.M.
1987-07-01
An experimental program to measure the aerodynamic performance of a NACA 64-621 airfoil with a truncated trailing edge for wind turbine applications has been conducted in the Ohio State University Aeronautical and Astronautical Research Laboratory 6 in. x 22 in. pressurized wind tunnel. The blunted or trailing edge truncated (TET) airfoil has an advantage over similar sharp trailing edge airfoils because it is able to streamline a larger spar structure, while also providing aerodynamic properties that are quite good. Surface pressures were measured and integrated to determine the lift, pressure drag, and moment coefficients over angles of attack ranging from -14 to +90 at Mach 0.2 and Reynolds numbers of 1,000,000 and 600,000. Results are compared to the NACA 0025, 0030, and 0035 thick airfoils with sharp trailing edges. Comparison shows that the 30% thick NACA 64-621-TET airfoil has higher maximum lift, higher lift curve slope, lower drag at higher lift coefficients, and higher chordwise force coefficient than similar thick airfoils with sharp trailing edges.
Wind tunnel evaluation of a truncated NACA 64-621 airfoil for wind turbine applications
NASA Astrophysics Data System (ADS)
Law, S. P.; Gregorek, G. M.
1987-07-01
An experimental program to measure the aerodynamic performance of a NACA 64-621 airfoil with a truncated trailing edge for wind turbine applications has been conducted in the Ohio State University Aeronautical and Astronautical Research Laboratory 6 in. by 21 in. pressurized wind tunnel. The blunted or trailing edge truncated (TET) airfoil has an advantage over similar trailing edge airfoils because it is able to streamline a larger spar structure, while also providing aerodynamic properties that are quite good. Surface pressures were measured and integrated to determine the lift, pressure drag, and moment coefficients over angles of attack ranging from -14 to +90 deg at Mach 0.2 and Reynolds numbers of 1,000,000 and 600,000. Results are compared to the NACA 0025, 0030, and 0035 thick airfoils with sharp trailing edges. Comparison shows that the 30 percent thick NACA 64-621-TET airfoil has higher maximum lift, higher lift curve slope, lower drag at higher lift coefficients, and higher chordwise force coefficient than similar thick airfoils with sharp trailing edges.
Wind tunnel evaluation of a truncated NACA 64-621 airfoil for wind turbine applications
NASA Technical Reports Server (NTRS)
Law, S. P.; Gregorek, G. M.
1987-01-01
An experimental program to measure the aerodynamic performance of a NACA 64-621 airfoil with a truncated trailing edge for wind turbine applications has been conducted in the Ohio State University Aeronautical and Astronautical Research Laboratory 6 in. by 21 in. pressurized wind tunnel. The blunted or trailing edge truncated (TET) airfoil has an advantage over similar trailing edge airfoils because it is able to streamline a larger spar structure, while also providing aerodynamic properties that are quite good. Surface pressures were measured and integrated to determine the lift, pressure drag, and moment coefficients over angles of attack ranging from -14 to +90 deg at Mach 0.2 and Reynolds numbers of 1,000,000 and 600,000. Results are compared to the NACA 0025, 0030, and 0035 thick airfoils with sharp trailing edges. Comparison shows that the 30 percent thick NACA 64-621-TET airfoil has higher maximum lift, higher lift curve slope, lower drag at higher lift coefficients, and higher chordwise force coefficient than similar thick airfoils with sharp trailing edges.
Porosity effect on supercritical airfoil drag reduction by shock wave/boundary layer control
NASA Technical Reports Server (NTRS)
Nagamatsu, H. T.; Orozco, R. D.; Ling, D. C.
1984-01-01
An investigation of the passive shock wave/boundary layer control for reducing the drag of 14 percent-thick supercritical airfoil was conducted in the 3 in. x 15.4 in. RPI Transonic Wind Tunnel at transonic Mach numbers. Various porous surfaces with a cavity beneath it was positioned on the area of the airfoil, mounted on the test section bottom wall, where the shock wave occurs. The static pressure distributions over the airfoil, the wake impact pressure survey for determining the profile drag and the Schlieren photographs for porous surfaces are presented and compared with the results for solid surface airfoil. With a uniform porosity surface the normal shock wave for solid surface was changed to a lambda shock wave, and the wake impact pressure data indicated an appreciable drag reduction at transonic Mach numbers. For a free stream Mach number of 0.81 the profile drag coefficient for the airfoil top surface with uniform porosity was 46 percent lower than for the solid surface airfoil.
Preparing and Analyzing Iced Airfoils
NASA Technical Reports Server (NTRS)
Vickerman, Mary B.; Baez, Marivell; Braun, Donald C.; Cotton, Barbara J.; Choo, Yung K.; Coroneos, Rula M.; Pennline, James A.; Hackenberg, Anthony W.; Schilling, Herbert W.; Slater, John W.; Burke, Kevin M.; Nolan, Gerald J.; Brown, Dennis
2004-01-01
SmaggIce version 1.2 is a computer program for preparing and analyzing iced airfoils. It includes interactive tools for (1) measuring ice-shape characteristics, (2) controlled smoothing of ice shapes, (3) curve discretization, (4) generation of artificial ice shapes, and (5) detection and correction of input errors. Measurements of ice shapes are essential for establishing relationships between characteristics of ice and effects of ice on airfoil performance. The shape-smoothing tool helps prepare ice shapes for use with already available grid-generation and computational-fluid-dynamics software for studying the aerodynamic effects of smoothed ice on airfoils. The artificial ice-shape generation tool supports parametric studies since ice-shape parameters can easily be controlled with the artificial ice. In such studies, artificial shapes generated by this program can supplement simulated ice obtained from icing research tunnels and real ice obtained from flight test under icing weather condition. SmaggIce also automatically detects geometry errors such as tangles or duplicate points in the boundary which may be introduced by digitization and provides tools to correct these. By use of interactive tools included in SmaggIce version 1.2, one can easily characterize ice shapes and prepare iced airfoils for grid generation and flow simulations.
Wind-tunnel Tests of the NACA 45-125 Airfoil: A Thick Airfoil for High-Speed Airplanes
NASA Technical Reports Server (NTRS)
Delano, James B.
1940-01-01
Investigations of the pressure distribution, the profile drag, and the location of transition for a 30-inch-chord 25-percent-thick N.A,C.A. 45-125 airfoil were made in the N.A.C.A 8-foot high-speed wind tunnel for the purpose of aiding in the development of a thick wing for high-speed airplanes. The tests were made at a lift coefficient of 0.1 for Reynolds Numbers from 1,750,000 to 8,690,000, corresponding to speeds from 80 to 440 miles per hour at 59 F. The effect on the profile drag of fixing the transition point was also investigated. The effect of compressibility on the rate of increase of pressure coefficients was found to be greater than that predicted by a simplified theoretical expression for thin wings. The results indicated that, for a lift coefficient of 0.1, the critical speed of the N.A.C,A. 45-125 airfoil was about 460 miles per hour at 59 F,. The value of the profile-drag coefficient at a Reynolds Number of 4,500,000 was 0.0058, or about half as large as the value for the N.A,C,A. 0025 airfoil. The increase in the profile-drag coefficient for a given movement of the transition point was about three times as large as the corresponding increase for the N.A.C,A. 0012 airfoil. Transition determinations indicated that, for Reynolds Numbers up to ?,000,000, laminar boundary 1ayers were maintained over approximately 40 percent of the upper and the lower surfaces of the airfoil.
NASA Astrophysics Data System (ADS)
Gardner, A. D.; Richter, K.
2015-06-01
A method of boundary layer transition measurement is presented for wind tunnel models instrumented with surface pressure taps. The measurement relies on taking a number of theoretically identical measurements at different times and then analysing the standard deviation of the pressures. Due to the slight unsteady movement of the transition position, a peak in the standard deviation of pressure is found at the transition position, and this is correlated with measurements of the transition position with an infrared camera and hot-film anemometers. In contrast to microphone measurements, it is shown that the transition detection works for data which have been low-pass filtered with a cut-off of 1 Hz. The application to static and dynamic transition measurements on static and periodically pitching helicopter rotor blade airfoils at Mach 0.3-0.5 is demonstrated.
NASA Technical Reports Server (NTRS)
Weinberg, B. C.; Mcdonald, H.
1982-01-01
A numerical scheme is developed for solving the time dependent, three dimensional compressible viscous flow equations to be used as an aid in the design of helicopter rotors. In order to further investigate the numerical procedure, the computer code developed to solve an approximate form of the three dimensional unsteady Navier-Stokes equations employing a linearized block implicit technique in conjunction with a QR operator scheme is tested. Results of calculations are presented for several two dimensional boundary layer flows including steady turbulent and unsteady laminar cases. A comparison of fourth order and second order solutions indicate that increased accuracy can be obtained without any significant increases in cost (run time). The results of the computations also indicate that the computer code can be applied to more complex flows such as those encountered on rotating airfoils. The geometry of a symmetric NACA four digit airfoil is considered and the appropriate geometrical properties are computed.
Unsteady aerodynamics of blade rows
NASA Technical Reports Server (NTRS)
Verdon, Joseph M.
1989-01-01
The requirements placed on an unsteady aerodynamic theory intended for turbomachinery aeroelastic or aeroacoustic applications are discussed along with a brief description of the various theoretical models that are available to address these requirements. The major emphasis is placed on the description of a linearized inviscid theory which fully accounts for the affects of a nonuniform mean or steady flow on unsteady aerodynamic response. Although this linearization was developed primarily for blade flutter prediction, more general equations are presented which account for unsteady excitations due to incident external aerodynamic disturbances as well as those due to prescribed blade motions. The motivation for this linearized unsteady aerodynamic theory is focused on, its physical and mathematical formulation is outlined and examples are presented to illustrate the status of numerical solution procedures and several effects of mean flow nonuniformity on unsteady aerodynamic response.
Design considerations of advanced supercritical low drag suction airfoils
NASA Technical Reports Server (NTRS)
Pfenninger, W.; Reed, H. L.; Dagenhart, J. R.
1980-01-01
Supercritical low drag suction laminar flow airfoils were laid out for shock-free flow at design freestream Mach = 0.76, design lift coefficient = 0.58, and t/c = 0.13. The design goals were the minimization of suction laminarization problems and the assurance of shock-free flow at freestream Mach not greater than design freestream Mach (for design lift coefficient) as well as at lift coefficient not greater than design lift coefficient (for design freestream Mach); this involved limiting the height-to-length ratio of the supersonic zone at design to 0.35. High design freestream Mach numbers result with extensive supersonic flow (over 80% of the chord) on the upper surface, with a steep Stratford-type rear pressure rise with suction, as well as by carrying lift essentially in front- and rear-loaded regions of the airfoil with high static pressures on the carved out front and rear lower surface.
A natural low frequency oscillation in the wake of an airfoil near stalling conditions
NASA Technical Reports Server (NTRS)
Zaman, K. B. M. Q.; Mckinzie, D. J.
1988-01-01
An unusually low frequency oscillation in the flow over an airfoil was explored experimentally. Wind tunnel measurements were carried out with a two dimensional airfoil model at a chord Reynolds number of 100,000. During deep stall the usual bluff-body shedding occurred at a Strouhal number. But at the onset of stall a low frequency periodic oscillation occurred, the corresponding Strouhal number being an order of magnitude lower. The phenomenon occurred in relatively unclean flow when the freestream turbulence was raised to 0.4 percent, but did not in the cleaner flow with turbulence intensity of 0.1 percent. It could also be produced by certain high frequency acoustic excitation. Details of the flow field are compared between a case of low frequency oscillation at alpha = 15 deg and a case of bluff-body shedding at alpha = 22.5 deg. The origin of the low frequency oscillation traces to the upper surface of the airfoil and is seemingly associated with the periodic formation and breakdown of a large separation bubble. The intense flow fluctuations impart significant unsteady forces to the airfoil but diminish rapidly within a distance of one chord from the trailing edge.
A natural low frequency oscillation in the wake of an airfoil near stalling conditions
NASA Technical Reports Server (NTRS)
Zaman, K. B. M. Q.; Mckinzie, D. J.
1987-01-01
An unusually low frequency oscillation in the flow over an airfoil was explored experimentally. Wind tunnel measurements were carried out with a two dimensional airfoil model at a chord Reynolds number of 100,000. During deep stall the usual bluff-body shedding occurred at a Strouhal number. But at the onset of stall a low frequency periodic oscillation occurred, the corresponding Strouhal number being an order of magnitude lower. The phenomenon occurred in relatively unclean flow when the freestream turbulence was raised to 0.4 percent, but did not in the cleaner flow with turbulence intensity of 0.1 percent. It could also be produced by certain high frequency acoustic excitation. Details of the flow field are compared between a case of low frequency oscillation at alpha = 15 deg and a case of bluff-body shedding at alpha = 22.5 deg. The origin of the low frequency oscillation traces to the upper surface of the airfoil and is seemingly associated with the periodic formation and breakdown of a large separation bubble. The intense flow fluctuations impart significant unsteady forces to the airfoil but diminish rapidly within a distance of one chord from the trailing edge.
NASA Technical Reports Server (NTRS)
Goradia, S. H.; Mehta, J. M.; Shrewsbury, G. S.
1977-01-01
The viscous flow phenomena associated with sharp and blunt trailing edge airfoils were investigated. Experimental measurements were obtained for a 17 percent thick, high performance GAW-1 airfoil. Experimental measurements consist of velocity and static pressure profiles which were obtained by the use of forward and reverse total pressure probes and disc type static pressure probes over the surface and in the wake of sharp and blunt trailing edge airfoils. Measurements of the upper surface boundary layer were obtained in both the attached and separated flow regions. In addition, static pressure data were acquired, and skin friction on the airfoil upper surface was measured with a specially constructed device. Comparison of the viscous flow data with data previously obtained elsewhere indicates reasonable agreement in the attached flow region. In the separated flow region, considerable differences exist between these two sets of measurements.
CFD aerodynamic analysis of non-conventional airfoil sections for very large rotor blades
NASA Astrophysics Data System (ADS)
Papadakis, G.; Voutsinas, S.; Sieros, G.; Chaviaropoulos, T.
2014-12-01
The aerodynamic performance of flat-back and elliptically shaped airfoils is analyzed on the basis of CFD simulations. Incompressible and low-Mach preconditioned compressible unsteady simulations have been carried out using the k-w SST and the Spalart Allmaras turbulence models. Time averaged lift and drag coefficients are compared to wind tunnel data for the FB 3500-1750 flat back airfoil while amplitudes and frequencies are also recorded. Prior to separation averaged lift is well predicted while drag is overestimated keeping however the trend in the tests. The CFD models considered, predict separation with a 5° delay which is reflected on the load results. Similar results are provided for a modified NACA0035 with a rounded (elliptically shaped) trailing edge. Finally as regards the dynamic characteristics in the load signals, there is fair agreement in terms of Str number but significant differences in terms of lift and drag amplitudes.
Comparison of sound power radiation from isolated airfoils and cascades in a turbulent flow.
Blandeau, Vincent P; Joseph, Phillip F; Jenkins, Gareth; Powles, Christopher J
2011-06-01
An analytical model of the sound power radiated from a flat plate airfoil of infinite span in a 2D turbulent flow is presented. The effects of stagger angle on the radiated sound power are included so that the sound power radiated upstream and downstream relative to the fan axis can be predicted. Closed-form asymptotic expressions, valid at low and high frequencies, are provided for the upstream, downstream, and total sound power. A study of the effects of chord length on the total sound power at all reduced frequencies is presented. Excellent agreement for frequencies above a critical frequency is shown between the fast analytical isolated airfoil model presented in this paper and an existing, computationally demanding, cascade model, in which the unsteady loading of the cascade is computed numerically. Reasonable agreement is also observed at low frequencies for low solidity cascade configurations.
Navier-Stokes computations for circulation control airfoils
NASA Technical Reports Server (NTRS)
Pulliam, Thomas H.; Jespersen, Dennis C.; Barth, Timothy J.
1987-01-01
Navier-Stokes computations of subsonic to transonic flow past airfoils with augmented lift due to rearward jet blowing over a curved trailing edge are presented. The approach uses a spiral grid topology. Solutions are obtained using a Navier-Stokes code which employs an implicit finite difference method, an algebraic turbulence model, and developments which improve stability, convergence, and accuracy. Results are compared against experiments for no jet blowing and moderate jet pressures and demonstrate the capability to compute these complicated flows.
Navier-Stokes computations for circulation controlled airfoils
NASA Technical Reports Server (NTRS)
Pulliam, T. H.; Jesperen, D. C.; Barth, T. J.
1986-01-01
Navier-Stokes computations of subsonic to transonic flow past airfoils with augmented lift due to rearward jet blowing over a curved trailing edge are presented. The approach uses a spiral grid topology. Solutions are obtained using a Navier-Stokes code which employs an implicit finite difference method, an algebraic turbulence model, and developments which improve stability, convergence, and accuracy. Results are compared against experiments for no jet blowing and moderate jet pressures and demonstrate the capability to compute these complicated flows.
Calculation of the rotor induced download on airfoils
NASA Technical Reports Server (NTRS)
Lee, C. S.
1989-01-01
Interactions between the rotors and wing of a rotary wing aircraft in hover have a significant detrimental effect on its payload performance. The reduction of payload results from the wake of lifting rotors impinging on the wing, which is at 90 deg angle of attack in hover. This vertical drag, often referred as download, can be as large as 15 percent of the total rotor thrust in hover. The rotor wake is a three-dimensional, unsteady flow with concentrated tip vortices. With the rotor tip vortices impinging on the upper surface of the wing, the flow over the wing is not only three-dimensional and unsteady, but also separated from the leading and trailing edges. A simplified two-dimensional model was developed to demonstrate the stability of the methodology. The flow model combines a panel method to represent the rotor and the wing, and a vortex method to track the wing wake. A parametric study of the download on a 20 percent thick elliptical airfoil below a rotor disk of uniform inflow was performed. Comparisons with experimental data are made where the data are available. This approach is now being extended to three-dimensional flows. Preliminary results on a wing at 90 deg angle of attack in free stream is presented.
NASA Technical Reports Server (NTRS)
Cole, Gregory M.; Mueller, Thomas J.
1990-01-01
An experimental investigation was conducted to measure the flow velocity in the boundary layer of an Eppler 387 airfoil. In particular, the laminar separation bubble that this airfoil exhibits at low Reynolds numbers was the focus. Single component laser Doppler velocimetry data were obtained at a Reynolds number of 100,000 at an angle of attack of 2.0 degree. Static Pressure and flow visualization data for the Eppler 387 airfoil were also obtained. The difficulty in obtaining accurate experimental measurements at low Reynolds numbers is addressed. Laser Doppler velocimetry boundary layer data for the NACA 663-018 airfoil at a Reynolds number of 160,000 and angle of attack of 12 degree is also presented.
Application of shock tubes to transonic airfoil testing at high Reynolds numbers
NASA Technical Reports Server (NTRS)
Cook, W. J.; Chaney, M. J.; Presley, L. L.; Chapman, G. T.
1978-01-01
Performance analysis of a gas-driven shock tube shows that transonic airfoil flows with chord Reynolds numbers of the order of 100 million can be produced, with limitations being imposed by the structural integrity of the facility or the model. A study of flow development over a simple circular arc airfoil at zero angle of attack was carried out in a shock tube at low and intermediate Reynolds numbers to assess the testing technique. Results obtained from schlieren photography and airfoil pressure measurements show that steady transonic flows similar to those produced for the same airfoil in a wind tunnel can be generated within the available testing time in a shock tube with properly contoured test section walls.
A Method for the Constrained Design of Natural Laminar Flow Airfoils
NASA Technical Reports Server (NTRS)
Green, Bradford E.; Whitesides, John L.; Campbell, Richard L.; Mineck, Raymond E.
1996-01-01
A fully automated iterative design method has been developed by which an airfoil with a substantial amount of natural laminar flow can be designed, while maintaining other aerodynamic and geometric constraints. Drag reductions have been realized using the design method over a range of Mach numbers, Reynolds numbers and airfoil thicknesses. The thrusts of the method are its ability to calculate a target N-Factor distribution that forces the flow to undergo transition at the desired location; the target-pressure-N-Factor relationship that is used to reduce the N-Factors in order to prolong transition; and its ability to design airfoils to meet lift, pitching moment, thickness and leading-edge radius constraints while also being able to meet the natural laminar flow constraint. The method uses several existing CFD codes and can design a new airfoil in only a few days using a Silicon Graphics IRIS workstation.
Unsteady transonic flow calculations for realistic aircraft configurations
NASA Technical Reports Server (NTRS)
Batina, John T.; Seidel, David A.; Bland, Samuel R.; Bennett, Robert M.
1987-01-01
A transonic unsteady aerodynamic and aeroelasticity code has been developed for application to realistic aircraft configurations. The new code is called CAP-TSD which is an acronym for Computational Aeroelasticity Program - Transonic Small Disturbance. The CAP-TSD code uses a time-accurate approximate factorization (AF) algorithm for solution of the unsteady transonic small-disturbance equation. The AF algorithm is very efficient for solution of steady and unsteady transonic flow problems. It can provide accurate solutions in only several hundred time steps yielding a significant computational cost savings when compared to alternative methods. The new code can treat complete aircraft geometries with multiple lifting surfaces and bodies including canard, wing, tail, control surfaces, launchers, pylons, fuselage, stores, and nacelles. Applications are presented for a series of five configurations of increasing complexity to demonstrate the wide range of geometrical applicability of CAP-TSD. These results are in good agreement with available experimental steady and unsteady pressure data. Calculations for the General Dynamics one-ninth scale F-16C aircraft model are presented to demonstrate application to a realistic configuration. Unsteady results for the entire F-16C aircraft undergoing a rigid pitching motion illustrated the capability required to perform transonic unsteady aerodynamic and aeroelastic analyses for such configurations.
NASA Astrophysics Data System (ADS)
Hounjet, M. H. L.
1988-10-01
A description of the method CAR88 for the calculation of steady and time linearized unsteady subsonic and supersonic flow about complex two dimensional and three dimensional configurations is given. The method belongs to the category of the so called panel methods. Results of applications in supersonic flow made to a two dimensional oscillating flat plate, a two dimensional double and single edge airfoil, three dimensional cones, and to single and multiple interfering wing surface are shown.