Some aspects of flow control over a NACA0015 airfoil using synthetic jets
NASA Astrophysics Data System (ADS)
Parthasarathy, T.; Das, S. P.
2017-04-01
Flow control for performance enhancement over airfoils has become an increasingly important topic. This work details the characteristics of flow control using synthetic jets over a NACA0015 airfoil at a Reynolds number of 896,000 based on the chord length and free stream velocity, and at 20° angle of attack wherein the flow is separated. Numerical simulations were performed to help understand the behaviour of the controlled flow for a range of synthetic jet parameters. Analysis of key flow parameters such as phase averaged pressure and streamline profiles indicate that the synthetic jet is efficient in increasing the lift coefficient; more so for larger jet amplitudes and at smaller angles of jet injection. Behaviour of the flow characteristics for controlled cases has been analysed from the flow structures obtained from the same. This work serves as a platform to qualitatively and quantitatively understand the effects of the jet parameters on the separated flow over the 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.
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)
Active Flow Control at Low Reynolds Numbers on a NACA 0015 Airfoil
NASA Technical Reports Server (NTRS)
Melton, LaTunia Pack; Hannon, Judith; Yao, Chung-Sheng; Harris, Jerome
2008-01-01
Results from a low Reynolds number wind tunnel experiment on a NACA 0015 airfoil with a 30% chord trailing edge flap tested at deflection angles of 0, 20, and 40 are presented and discussed. Zero net mass flux periodic excitation was applied at the ap shoulder to control flow separation for flap deflections larger than 0. The primary objective of the experiment was to compare force and moment data obtained from integrating surface pressures to data obtained from a 5-component strain-gage balance in preparation for additional three-dimensional testing of the model. To achieve this objective, active flow control is applied at an angle of attack of 6 where published results indicate that oscillatory momentum coefficients exceeding 1% are required to delay separation. Periodic excitation with an oscillatory momentum coefficient of 1.5% and a reduced frequency of 0.71 caused a significant delay of separation on the airfoil with a flap deflection of 20. Higher momentum coefficients at the same reduced frequency were required to achieve a similar level of flow attachment on the airfoil with a flap deflection of 40. There was a favorable comparison between the balance and integrated pressure force and moment results.
NASA Technical Reports Server (NTRS)
Hassan, Ahmed
1999-01-01
Using the two-dimensional ARC2D Navier-Stokes flow solver analyses were conducted to predict the sectional aerodynamic characteristics of the flapped NACA-0015 airfoil section. To facilitate the analyses and the generation of the computational grids, the airfoil with the deflected trailing edge flap was treated as a single element airfoil with no allowance for a gap between the flap's leading edge and the base of the forward portion of the airfoil. Generation of the O-type computational grids was accomplished using the HYGRID hyperbolic grid generation program. Results were obtained for a wide range of Mach numbers, angles of attack and flap deflections. The predicted sectional lift, drag and pitching moment values for the airfoil were then cast in tabular format (C81) to be used in lifting-line helicopter rotor aerodynamic performance calculations. Similar were also generated for the flap. Mathematical expressions providing the variation of the sectional lift and pitching moment coefficients for the airfoil and for the flap as a function of flap chord length and flap deflection angle were derived within the context of thin airfoil theory. The airfoil's sectional drag coefficient were derived using the ARC2D drag predictions for equivalent two dimensional flow conditions.
NASA Astrophysics Data System (ADS)
Boukenkoul, M. A.; Li, F. C.; Aounallah, M.
2017-03-01
The present study aims to investigate numerically the flow control possibility using a synthetic jet actuation over a bi-dimensional NACA0015 airfoil manoeuvring at a highly turbulent flow (8.9e105 Reynolds to chord number). The 2-D flow behaviour was computed using the ANSYS Fluent commercial code. The so-called Reynolds Averaged Navier-Stocks (RANS) approach has been tested for one (Spalat-Allmaras S-A) and two (K-ε) transport equations for the turbulence modelling. Both present a weakness to predict the stall angle effectively. The S-A lift coefficient slope seems to be the closest to the experimental data. The synthetic jet control exhibits an extraordinary lift coefficient enhancement at high Angles Of Attack (AOA) but seems to be less obvious at low AOA, where the flow is still attached. A synthetic jet of a Strouhal (St = 2) and momentum (Cμ of 0.56%), delays the stall onset from 15 to 19deg with enhancing the lift coefficient by 40%. The actual work has been enriched by studying the effect of the jet’s frequency and momentum on the lift temporal signal. Also, the interaction between the mean flow and the synthetic jet structures topology was undertaken.
Sheldahl, R.E.; Klimas, P.C.; Feltz, L.V.
1980-03-01
A 5-metre-diameter vertical-axis wind turbine has undergone continued testing since 1976 at the Sandia Laboratories Wind Turbine site. The latest tests of this machine have been with extruded aluminum blades of NACA-0015 airfoil cross section. The results of these tests at several turbine rotational speeds are presented and compared with earlier test results. A performance comparison is made with a vortex/lifting line computational code. The performance of the turbine with the extruded blades met all expectations.
NACA 0015 wing pressure and trailing vortex measurements
NASA Technical Reports Server (NTRS)
Mcalister, K. W.; Takahashi, R. K.
1991-01-01
A NACA 0015 semispan wing was placed in a low-speed wind tunnel, and measurements were made of the pressure on the upper and lower surface of the wing and of velocity across the vortex trailing downstream from the tip of the wing. Pressure data were obtained for both 2-D and 3-D configurations. These data feature a detailed comparison between wing tips with square and round lateral edges. A two-component laser velocimeter was used to measure velocity profiles across the vortex at numerous stations behind the wing and for various combinations of conditions. These conditions include three aspect ratios, three chord lengths, a square- and a round lateral-tip, presence or absence of a boundary-layer trip, and three image plane positions located opposite the wing tip. Both pressure and velocity measurements were made for the angles of attack 4 deg less than or equal to alpha less than or equal to 12 deg and for Reynolds numbers 1 x 10(exp 6) less than or equal to Re less than or equal to 3 x 10(exp 6).
Flow analysis of NACA0015 and its application on tidal power
NASA Astrophysics Data System (ADS)
Liu, D. M.; Chen, H.; Yuan, L. L.; Zhao, Y. Z.; Liu, X. B.
2015-01-01
NACA0015 is a commonly used hydrofoil for hydraulic machinery. Because the calculated flow characteristic is sensitive to the mesh density and distribution, a careful check of mesh independence should be performed before the CFD calculation. For this purpose, flow comparison was done between calculation and experimental test at the same operating condition with the Reynolds number equal to 130,000. When the span length is two times of the chord length, the ratio of lift to drag is 34.97, which is close to the test result of 36.47. The pressure coefficient and cavitation characteristic were also calculated. For example, the pressure coefficient was compared in different flow conditions with Reynolds number equal to 130,000, 200,000 and 400,000. The cavitation characteristic was calculated at different cavitation numbers, e.g. 0.5, 1.0 and 1.5. Based on these hydraulic and cavitation characteristic, a tidal turbine runner was developed by using hydrofoil NACA0015. According to our studies, the flow characteristic and cavitation index can satisfy the engineering requirements.
Study of Cavitation Shedding Dynamics on a NACA0015 Hydrofoil Using X-Ray Densitometry
NASA Astrophysics Data System (ADS)
Ganesh, Harish; Wu, Juliana; Ceccio, Steven
2016-11-01
Cavitation dynamics on the NACA0015 hydrofoil at several attack angles are found to be spectrally rich, being multi-modal with abrupt changes in Strouhal number with change in cavitation number. Present study focusses on identifying the physical mechanisms responsible for the change in cavitation dynamics on a NACA0015 hydrofoil in a re-circulating water tunnel using time resolved X-ray densitometry. Time-resolved void fraction flow fields obtained using X-ray densitometry, synchronized with acoustic noise measurements using a hydrophone, are used to identify different flow features and mechanisms that are responsible for the change in the observed spectral behavior. It is shown that under higher cavitation numbers, the shedding mechanism is predominantly re-entrant liquid flow based, but as the cavitation number drops many different processes are at play. At lower cavitation numbers, the shed cavity cloud collapse arrests cavity growth and this results in altered cycle dynamics and hence the Strouhal number. In addition, propagation bubbly shock waves are also found to be a dominant mechanism of shedding for certain conditions. The multi-modal nature of the acoustic pressure signature is explained by presence of different flow features, which could be concurrent or alternating. Office of Naval Research.
Numerical simulation of cavitation erosion on a NACA0015 hydrofoil based on bubble collapse strength
NASA Astrophysics Data System (ADS)
Hidalgo, V.; Luo, X.; Escaler, X.; Huang, R.; Valencia, E.
2015-12-01
The prediction of erosion under unsteady cavitation is crucial to prevent damage in hydraulic machinery. The present investigation deals with the numerical simulation of erosive partial cavitation around a NACA0015 hydrofoil. The study presents the calculation of the bubble collapse strength, Sb, based on the bubble potential energy to identify the surface areas with highest risk of damage. The results are obtained with a numerical scheme assuming homogeneous mixture flow, implicit LES and Zwart cavitation model. The 3D unsteady flow simulation has been solved using OpenFOAM. Python language and OpenFOAM calculator (foamCalcEx) have been used to obtain and represent Sb. The obtained results clearly show the instants of erosive bubble collapse and the affected surface areas.
NASA Astrophysics Data System (ADS)
Kravtsova, A. Yu.; Markovich, D. M.; Pervunin, K. S.; Timoshevskii, M. V.; Hanjalić, K.
2014-12-01
Using high-speed visualization and particle image velocimetry (PIV), cavitating flows near a plane plate with a rounded leading edge and NACA0015 hydrofoil at angles of attack from 0° to 9° are studied. In the experiments, several known types of cavitation, as well as some differences, were detected with variation of the cavitation number. In particular, at small angles of attack (up to 3°), cavitation on the plate appears in the form of a streak array; on the hydrofoil, it appears in the form of individual bubbles. For the NACA0015 hydrofoil, isolated and intermittent streaks are divided and grow in regimes with developed cavitation; then, however, they merge in bubble clouds and form an extremely regular cellular structure. With an increase in the angle of attack to 9°, the structure of the cavitation cavity on the hydrofoil is changed by the streak structure, like in the case with the plate. In this work, it is shown that PIV permits one to measure the velocity in cavitating flows, in particular, within the gas-vapor phase. It was established from the analysis of distributions of the average flow velocity and moments of velocity fluctuations that the cavitation generation is caused by the development of the carrier fluid flow near the leading edge of the hydrofoil. Down the stream, however, the flow structure strongly depends on the cavitation regime, which is seen from the comparison of the distributions with the case of a single-phase flow. The presented measurements qualitatively verify general trends and show some quantitative distinctions for the two considered flowpast bodies.
An Investigation of Physics and Control of Flow Passing a NACA 0015 in Fully-Reversed Condition
NASA Astrophysics Data System (ADS)
Clifford, Christopher J.
Flow control experiments were performed on a NACA 0015 airfoil in fully-reversed condition, which is anticipated to occur on the retreating blade side of advanced helicopters such as slowed-rotor compound rotorcraft. Control was achieved using nanosecond dielectric barrier discharge (NS-DBD) plasma actuators. The Reynolds number based on a chord length of 203 mm was fixed at 5.0 · 105, corresponding to a freestream velocity of ˜38 m/s. Two angles of attack were considered: α = 0° and 15°, each of which is relevant to a particular implementation of slowed-rotor technology. At α = 0°, the flow resembles that of a flow behind a cylinder. A von Karman vortex street formed in the wake where alternating vortex shedding occurred at a Strouhal number of 0.12. Excitation was performed using an NS-DBD on one side of the airfoil, with plasma formation just upstream of the separation line. However, there was no discernible influence upon the baseline behavior. At α = 15°, fully separated flow on the suction side extended well beyond the airfoil with naturally shed vortices at a Strouhal number of 0.19. Plasma actuation was evaluated at both the aerodynamic leading-edge (ALE) and aerodynamic trailing-edge (ATE) of the airfoil. The flow responded to the plasma actuation at the ALE by generating organized coherent structures in the shear layer over the separated region. Moderate excitation around the natural shedding Strouhal number had the most significant effects: synchronizing the shedding from the ALE and ATE, creating moderately sized structures that convected far downstream, greatly reducing the separation area, increasing lift, and decreasing drag. Excitation at much higher Strouhal numbers resulted in the flow returning to its natural shedding state, but with less coherent structures that diffused in the wake. This reduced the separation area and significantly reduced drag. Plasma actuation at the ATE caused a reduction in the magnitude of the fundamental and
Cavitation dynamics on a NACA0015 hydrofoil using time resolved X-ray densitometry
NASA Astrophysics Data System (ADS)
Ganesh, Harish; Wu, Juliana; Ceccio, Steven
2015-11-01
Recent investigations of partial cavitation have shown that the transition from stable to shedding cavities can be related to the presence of both propagating bubbly shocks and re-entrant liquid jets originating in the cavity closure region. In the present study, formation of sheet cavitation and its transition to periodically shedding cavities is studied on a NACA0015 hydrofoil in a recirculating water tunnel at different attack angles. Using high-speed videos and time resolved X-ray densitometry, the instantaneous void fraction flow fields are obtained to identify the principal mechanism responsible for transition from stable to shedding cavities over a range of attack angles and cavitation numbers. The role of attack angle is of particular interest, since is it related to the pressure gradient at cavity enclosure, and can lead to the formation of stronger reentrant flows. The relative importance of reentrant liquid flow and bubbly shock wave propagation will be discussed This work is supported by Office of Naval Research.
A study of wakes trailing a cavitating and ventilated NACA 0015 hydrofoil
NASA Astrophysics Data System (ADS)
Time, Rune W.; Kjeldsen, Morten; Arndt, Roger E.
2007-11-01
In certain types of pumping applications oscillations are induced by operation with liquids containing immiscible gas in solution. To understand the physics of this process, a series of tests with a ventilated 2D NACA 0015 hydrofoil were performed in the water tunnel at the St. Anthony Falls Laboratory of the University of Minnesota. A relation between the cavitation index for both ventilated and naturally cavitating flows has been established. This allows direct comparison between the two types of two-phase flows. A good correspondence is found when comparing drag, lift break-down and lift/drag oscillation amplitudes for a properly defined cavitation index. A careful analysis is required to compare the spectral content of oscillatory flow. This study concentrates on analyzing wake structure for selected ventilation rates and cavitation numbers. The wake structure is mapped by conventional PIV techniques. Results for high cavitation/ventilation number flows are presented. Since conventional PIV represents snap shots of a time evolving phenomenon special care needs to be taken for the analysis.
NASA Astrophysics Data System (ADS)
Prothin, Sebastien; Billard, Jean-Yves; Djeridi, Henda
2016-10-01
The purpose of the present study is to get a better understanding of the hydrodynamic instabilities of sheet cavities which develop along solid walls. The main objective is to highlight the spatial and temporal behavior of such a cavity when it develops on a NACA0015 foil at high Reynolds number. Experimental results show a quasi-steady, periodic, bifurcation domain, with aperiodic cavity behavior corresponding to σ/2 α values of 5.75, 5, 4.3 and 3.58. Robust mathematical methods of signal postprocessing (proper orthogonal decomposition and dynamic mode decomposition) were applied in order to emphasize the spatio-temporal nature of the flow. These new techniques put in evidence the 3D effects due to the reentrant jet instabilities or due to propagating shock wave mechanism at the origin of the shedding process of the cavitation cloud.
NASA Astrophysics Data System (ADS)
Prothin, Sebastien; Djeridi, Henda; Billard, Jean-Yves
2014-05-01
In this paper, the influence of a single tip vortex on boundary layer detachment is studied. This study offers a preliminary approach in order to better understand the interaction between a propeller hub vortex and the rudder installed in its wake. This configuration belongs to the field of marine propulsion and encompasses such specific problem as cavitation inception, modification of propulsive performances and induced vibrations. To better understand the complex mechanisms due to propeller-rudder interactions it was decided to emphasize configurations where the hub vortex is generated by an elliptical 3-D foil and is located upstream of a 2-D NACA0015 foil at high incidences for a Reynolds number of 5×105. The physical mechanisms were studied using Time Resolved Stereoscopic Particle Image Velocimetry (TR-SPIV) techniques. Particular attention was paid to the detachment at 25° incidence and a detailed cartography of the mean and turbulent properties of the wake is presented. Proper Orthogonal Decomposition (POD) analysis was applied in order to highlight the unsteady nature of the flow using phase averaging based on the first POD coefficients to characterize the turbulent and coherent process in the near wake of the rudder.
NASA Astrophysics Data System (ADS)
Adjali, Saadia; Belkadi, Mustapha; Aounallah, Mohammed; Imine, Omar
2015-05-01
Accurate simulation of turbulent free surface flows around surface ships has a central role in the optimal design of such naval vessels. The flow problem to be simulated is rich in complexity and poses many modeling challenges because of the existence of breaking waves around the ship hull, and because of the interaction of the two-phase flow with the turbulent boundary layer. In this paper, our goal is to estimate the lift and drag coefficients for NACA 0012 of hydrofoil advancing in calm water under steady conditions with free surface and emerged NACA 0015. The commercial CFD software FLUENT version 14 is used for the computations in the present study. The calculated grid is established using the code computer GAMBIT 2.3.26.The shear stress k-ωSST model is used for turbulence modeling and the volume of fluid technique is employed to simulate the free-surface motion. In this computation, the second order upwind scheme is used for discretizing the convection terms in the momentum transport equations, the Modified HRIC scheme for VOF discretisation. The results obtained compare well with the experimental data.
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.
NASA Technical Reports Server (NTRS)
Derkacs, Thomas (Inventor); Fetheroff, Charles W. (Inventor); Matay, Istvan M. (Inventor); Toth, Istvan J. (Inventor)
1983-01-01
Although the method and apparatus of the present invention can be utilized to apply either a uniform or a nonuniform covering of material over many different workpieces, the apparatus (20) is advantageously utilized to apply a thermal barrier covering (64) to an airfoil (22) which is used in a turbine engine. The airfoil is held by a gripper assembly (86) while a spray gun (24) is effective to apply the covering over the airfoil. When a portion of the covering has been applied, a sensor (28) is utilized to detect the thickness of the covering. A control apparatus (32) compares the thickness of the covering of material which has been applied with the desired thickness and is subsequently effective to regulate the operation of the spray gun to adaptively apply a covering of a desired thickness with an accuracy of at least plus or minus 0.0015 of an inch (1.5 mils) despite unanticipated process variations.
Optimization of a synthetic jet actuator for flow control around an airfoil
NASA Astrophysics Data System (ADS)
Montazer, E.; Mirzaei, M.; Salami, E.; Ward, T. A.; Romli, F. I.; Kazi, S. N.
2016-10-01
This paper deals with the optimization of a synthetic jet actuator parameters in the control flow around the NACA0015 airfoil at two angles of attack: 13° (i.e. the stall angle of NACA0015) and 16° (i.e. the post stall angle of NACA0015) to maximize the aerodynamic performance of the airfoil. Synthetic jet actuator is a zero mass flux-active flow control device that alternately injects and removes fluid through a small slot at the input movement frequency of a diaphragm. The movement of the diaphragm and also the external flow around the airfoil were simulated using numerical approach. The objective of the optimization process function was maximum lift-drag ratio (L/D) and the optimization variables were jet frequency, length of the jet slot and jet location along the chord. The power coefficient of the jet was considered as a constraint. The response surface optimization method was employed to achieve the optimal parameters. The results showed that the actuator is more effective for post stall angles of attack that can lead to an enhancement of 66% in L/D.
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.
Darrieus wind-turbine airfoil configurations
Migliore, P.G.; Fritschen, J.R.
1982-06-01
The purpose of this study was to determine what aerodynamic performance improvement, if any, could be achieved by judiciously choosing the airfoil sections for Darrieus wind turbine blades. Analysis was limited to machines using two blades of infinite aspect ratio, having rotor solidites from seven to twenty-one percent, and operating at maximum Reynolds numbers of approximately three million. Ten different airfoils, having thickness to chord ratios of twelve, fifteen and eighteen percent, were investigated. Performance calculations indicated that the NACA 6-series airfoils yield peak power coefficients at least as great as the NACA four-digit airfoils which have historically been chosen for Darrieus turbines. Furthermore, the power coefficient-tip speed ratio curves were broader and flatter for the 6-series airfoils. Sample calculations for an NACA 63/sub 2/-015 airfoil showed an annual energy output increase of 17 to 27% depending upon rotor solidity, compared to an NACA 0015 airfoil. An attempt was made to account for the flow curvature effects associated with Darrieus turbines by transforming the NACA 63/sub 2/-015 airfoil to an appropriate shape.
Comparison of NACA 6-series and 4-digit airfoils for Darrieus wind turbines
Migliore, P.G.
1983-07-01
The aerodynamic efficiency of Darrieus wind turbines as effected by blade airfoil geometry was investigated. Analysis was limited to curved-bladed machines having rotor solidities of 7 to 21% and operating at a Reynolds number of 3 X 10/sup 6/, Ten different airfoils, having thickness-to-chord ratios of 12, 15, and 18%, were studied. Performance estimates were made using a blade element/momentum theory approach. Results indicated that NACA 6-series airfoils yeild peak power coefficients as great as NACA 4-digit airfoils and have broader and flatter power coefficient-tip speed ratio curves. Sample calculations for an NACA 63/sub 2/-015 airfoil showed an annual energy output increase of 17-27%, depending on rotor solidity, compared to an NACA 0015 airfoil.
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.
Tomographic-PIV and TSP for Airfoil Flow-Control Studies
NASA Astrophysics Data System (ADS)
Stolt, Adam; Otte, John; Krech, Jesstin; Estevadeordal, Jordi
2016-11-01
Airfoil blades can experience a significant change of angle of attack during operation cycles that can lead to boundary layer separation and dynamic stall. It is unclear how elements distributed at the leading edge would affect the aerodynamic performance, boundary layer separation and transition, and stall behaviors. In the present study, various passive flow control structures, such as distributed dimples and bumps have been investigated and compared to airfoil geometries including the baseline smooth NACA0015 airfoil. Along with standard particle image velocimetry (PIV), a curved-laser sheet PIV, tomographic PIV, and Temperature Sensitive Paint (TSP) techniques have been combined to reveal spanwise flow information in the curved surface of the airfoil. Results show the effects and induced flow patterns of the various elements on boundary layer separation and stall at various angles of attack and compare them with the smooth models.
NASA Astrophysics Data System (ADS)
Hirata, Katsuya; Kawakita, Masatoshi; Iijima, Takayoshi; Koga, Mitsuhiro; Kihira, Mitsuhiko; Funaki, Jiro
The aerodynamic characteristics of airfoils have been researched in higher Reynolds-number ranges more than 106, in a historic context closely related with the developments of airplanes and fluid machineries in the last century. However, our knowledge is not enough at low and middle Reynolds-number ranges. So, in the present study, we investigate such basic airfoils as a NACA0015, a flat plate and the flat plates with modified fore-face and after-face geometries at Reynolds number Re < 1.0×105, using two- and three-dimensional computations together with wind-tunnel and water-tank experiments. As a result, we have revealed the effect of the Reynolds number Re upon the minimum drag coefficient CDmin. Besides, we have shown the effects of attack angle α upon various aerodynamic characteristics such as the lift coefficient CL, the drag coefficient CD and the lift-to-drag ratio CL/CD at Re = 1.0×102, discussing those effects on the basis of both near-flow-field information and surface-pressure profiles. Such results suggest the importance of sharp leading edges, which implies the possibility of an inversed NACA0015. Furthermore, concerning the flat-plate airfoil, we investigate the influences of fore-face and after-face geometries upon such effects.
CFD simulation of flow-induced vibration of an elastically supported airfoil
NASA Astrophysics Data System (ADS)
Šidlof, Petr
2016-03-01
Flow-induced vibration of lifting or control surfaces in aircraft may lead to catastrophic consequences. Under certain circumstances, the interaction between the airflow and the elastic structure may lead to instability with energy transferred from the airflow to the structure and with exponentially increasing amplitudes of the structure. In the current work, a CFD simulation of an elastically supported NACA0015 airfoil with two degrees of freedom (pitch and plunge) coupled with 2D incompressible airflow is presented. The geometry of the airfoil, mass, moment of inertia, location of the centroid, linear and torsional stiffness was matched to properties of a physical airfoil model used for wind-tunnel measurements. The simulations were run within the OpenFOAM computational package. The results of the CFD simulations were compared with the experimental data.
Experimental investigation on the unsteady lift of an airfoil in a sinusoidal streamwise gust
NASA Astrophysics Data System (ADS)
Yang, Yang; Li, Mingshui; Ma, Cunming; Li, Shaopeng
2017-05-01
The unsteady lift of a stationary airfoil encountering a sinusoidal streamwise gust is experimentally investigated under attached flow conditions. The airfoil has an NACA 0015 profile with a chord length c = 0.5 m. The Reynolds number of the main flow is in the range of 198 000 ⩽ Re ⩽ 331 000. The reduced frequency of the perturbation ω ranges from 0.047 to 0.392. The results show that the fluctuation in the lift force reduces very much at the high frequency. It can be attributed to a significant reduction in the fluctuating pressure difference over the front half of the airfoil. Thus experimental aerodynamic admittance functions for the unsteady lift are lower than theoretical predictions relating to this case with increasing reduced frequencies, though experimental values are close to theoretical values at low reduced frequencies.
Flow past a self-oscillating airfoil with two degrees of freedom: measurements and simulations
NASA Astrophysics Data System (ADS)
Šidlof, Petr; Štěpán, Martin; Vlček, Václav; Řidký, Václav; Šimurda, David; Horáček, Jaromír
2014-03-01
The paper focuses on investigation of the unsteady subsonic airflow past an elastically supported airfoil for subcritical flow velocities and during the onset of the flutter instability. A physical model of the NACA0015 airfoil has been designed and manufactured, allowing motion with two degrees of freedom: pitching (rotation about the elastic axis) and plunging (vertical motion). The structural mass and stiffness matrix can be tuned to certain extent, so that the natural frequencies of the two modes approach as needed. The model was placed in the measuring section of the wind tunnel in the aerodynamic laboratory of the Institute of Thermomechanics in Nový Knín, and subjected to low Mach number airflow up to the flow velocities when self-oscillation reach amplitudes dangerous for the structural integrity of the model. The motion of the airfoil was registered by a high-speed camera, with synchronous measurement of the mechanic vibration and discrete pressure sensors on the surface of the airfoil. The results of the measurements are presented together with numerical simulation results, based on a finite volume CFD model of airflow past a vibrating airfoil.
NASA Technical Reports Server (NTRS)
Addy, Harold E., Jr.; Potapczuk, Mark G.; Sheldon, David W.
1997-01-01
This report presents results from the first icing tests performed in the Modem Airfoils program. Two airfoils have been subjected to icing tests in the NASA Lewis Icing Research Tunnel (IRT). Both airfoils were two dimensional airfoils; one was representative of a commercial transport airfoil while the other was representative of a business jet airfoil. The icing test conditions were selected from the FAR Appendix C envelopes. Effects on aerodynamic performance are presented including the effects of varying amounts of glaze ice as well as the effects of approximately the same amounts of glaze, mixed, and rime ice. Actual ice shapes obtained in these tests are also presented for these cases. In addition, comparisons are shown between ice shapes from the tests and ice shapes predicted by the computer code, LEWICE for similar conditions. Significant results from the tests are that relatively small amounts of ice can have nearly as much effect on airfoil lift coefficient as much greater amounts of ice and that glaze ice usually has a more detrimental effect than either rime or mixed ice. LEWICE predictions of ice shapes, in general, compared reasonably well with ice shapes obtained in the IRT, although differences in details of the ice shapes were observed.
Numerical simulation of dynamic stall around an airfoil in Darrieus motion
Allet, A.; Halle, S.; Paraschivoiu, I.
1999-02-01
The objective of this study is to investigate the two-dimensional unsteady flow around an airfoil undergoing a Darrieus motion in dynamic stall conditions. For this purpose, a numerical solver based on the solution of the Reynolds-averaged Navier-Stokes equations expressed in a streamfunction-vorticity formulation in a non-inertial frame of reference was developed. The governing equations are solved by the streamline upwind Petrov-Galerkin finite element method (FEM). Temporal discretization is achieved by second-order-accurate finite differences. The resulting global matrix system is linearized by the Newton method and solved by the generalized minimum residual method (GMRES) with an incomplete triangular factorization preconditioning (ILU). Turbulence effects are introduced in the solver by an eddy viscosity model. The investigation centers on an evaluation of the algebraic Cebeci-Smith model (CSM) and the nonequilibrium Johnson-King model (JKM). In an effort to predict dynamic stall features on rotating airfoils, first the authors present some testing results concerning the performance of both turbulence models for the flat plate case. Then, computed flow structure together with aerodynamic coefficients for a NACA 0015 airfoil in Darrieus motion under dynamic stall conditions are presented.
NASA Astrophysics Data System (ADS)
Řidký, V.; Šidlof, P.; Vlček, V.
2013-04-01
The work is devoted to comparing measured data with the results of numerical simulations. As mathematical model was used mathematical model whitout turbulence for incompressible flow In the experiment was observed the behavior of designed NACA0015 airfoil in airflow. For the numerical solution was used OpenFOAM computational package, this is open-source software based on finite volume method. In the numerical solution is prescribed displacement of the airfoil, which corresponds to the experiment. The velocity at a point close to the airfoil surface is compared with the experimental data obtained from interferographic measurements of the velocity field. Numerical solution is computed on a 3D mesh composed of about 1 million ortogonal hexahedron elements. The time step is limited by the Courant number. Parallel computations are run on supercomputers of the CIV at Technical University in Prague (HAL and FOX) and on a computer cluster of the Faculty of Mechatronics of Liberec (HYDRA). Run time is fixed at five periods, the results from the fifth periods and average value for all periods are then be compared with experiment.
Ghodoosian, N.
1984-05-01
An analytical model leading to the pressure distribution on the cross section of a Darrieus Rotor Blade (airfoil) has veen constructed. The model is based on the inviscid flow theory and the contribution of the nonsteady wake vortices was neglected. The analytical model was translated into a computer code in order to study a variety of boundary conditions encountered by the rotating blades of the Darrieus Rotor. Results indicate that, for a pitching airfoil, lift can be adequately approximated by the Kutta-Joukowski forces, despite notable deviations in the pressure distribution on the airfoil. These deviations are most significant at the upwind half of the Darrieus Rotor where higher life is accompanied by increased adverse pressure gradients. The effect of pitching on lift can be approximated by a linear shift in the angle of attack proportional to the blade angular velocity. Tabulation of the fluid velocity about the pitching-only NACA 0015 allowed the principle of superposition to be used to determine the fluid velocity about a translating and pitching airfoil.
NASA Astrophysics Data System (ADS)
Komuro, Atsushi; Takashima, Keisuke; Konno, Kaiki; Tanaka, Naoki; Nonomura, Taku; Kaneko, Toshiro; Ando, Akira; Asai, Keisuke
2017-06-01
Gas-density perturbations near an airfoil surface generated by a nanosecond dielectric-barrier-discharge plasma actuator (ns-DBDPA) are visualized using a high-speed Schlieren imaging method. Wind-tunnel experiments are conducted for a wind speed of 20 m s-1 with an NACA0015 airfoil whose chord length is 100 mm. The results show that the ns-DBDPA first generates a pressure wave and then stochastic perturbations of the gas density near the leading edge of the airfoil. Two structures with different characteristics are observed in the stochastic perturbations. One structure propagates along the boundary between the shear layer and the main flow at a speed close to that of the main flow. The other propagates more slowly on the surface of the airfoil and causes mixing between the main and shear flows. It is observed that these two heated structures interact with each other, resulting in a recovery in the negative pressure coefficient at the leading edge of the airfoil.
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.
Navier-Stokes simulation of the flow around an airfoil in Darrieus motion
Tchon, K.F.; Paraschivoiu, I. . Dept. of Mechanical Engineering)
1994-12-01
In order to study the dynamic stall phenomenon on a Darrieus wind turbine, the incompressible flow field around a moving airfoil is simulated using a noninertial stream function-vorticity formulation of the two-dimensional unsteady navier-Stokes equations. Spatial discretization is achieved by the streamline upwind Petrov-Galerkin finite element method on a hybrid mesh composed of a structured region of quadrilateral elements in the vicinity of solid boundaries, an unstructured region of triangular elements elsewhere, and a layer of infinite elements surrounding the domain and projecting the external boundary to infinity. Temporal discretization is achieved by an implicit second order finite difference scheme. At each time step, a nonlinear algebraic system is solved by a Newton method. To accelerate computations, the generalized minimum residual method with an incomplete triangular factorization preconditioning is used to solve the linearized Newton systems. The solver is applied to simulate the flow around a NACA 0015 airfoil in Darrieus motion and the results are compared to experimental observations. To the authors' knowledge, it is the first time that the simulation of such a motion has been performed using the Navier-Stokes equations.
Investigation of airfoil leading edge separation control with nanosecond plasma actuator
NASA Astrophysics Data System (ADS)
Zheng, J. G.; Cui, Y. D.; Zhao, Z. J.; Li, J.; Khoo, B. C.
2016-11-01
A combined numerical and experimental investigation of airfoil leading edge flow separation control with a nanosecond dielectric barrier discharge (DBD) plasma actuator is presented. Our study concentrates on describing dynamics of detailed flow actuation process and elucidating the nanosecond DBD actuation mechanism. A loose coupling methodology is employed to perform simulation, which consists of a self-similar plasma model for the description of pulsed discharge and two-dimensional Reynolds averaged Navier-Stokes (RANS) equations for the calculation of external airflow. A series of simulations of poststall flows around a NACA0015 airfoil is conducted with a Reynolds number range covering both low and high Re at Re=(0.05 ,0.15 ,1.2 ) ×106 . Meanwhile, wind-tunnel experiment is performed for two low Re flows to measure aerodynamic force on airfoil model and transient flow field with time-resolved particle image velocimetry (PIV). The PIV measurement provides possibly the clearest view of flow reattachment process under the actuation of a nanosecond plasma actuator ever observed in experiments, which is highly comparable to that predicted by simulation. It is found from the detailed simulation that the discharge-induced residual heat rather than shock wave plays a dominant role in flow control. For any leading edge separations, the preliminary flow reattachment is realized by residual heat-induced spanwise vortices. After that, the nanosecond actuator functions by continuing exciting flow instability at poststall attack angles or acting as an active trip near stall angle. As a result, the controlled flow is characterized by a train of repetitive, downstream moving vortices over suction surface or an attached turbulent boundary layer, which depends on both angle of attack and Reynolds number. The advection of residual temperature with external flow offers a nanosecond plasma actuator a lot of flexibility to extend its influence region. Animations are provided for
Evaluation of Turbulence Models for Unsteady Flows of an Oscillating Airfoil
NASA Technical Reports Server (NTRS)
Srinivasan, G. R.; Ekaterinaris, J. A.; McCroskey, W. J.
1995-01-01
Unsteady flowfields of a two-dimensional oscillating airfoil are calculated using an implicit, finite-difference, Navier Stokes numerical scheme. Five widely used turbulence models are used with the numerical scheme to assess the accuracy and suitability of the models for simulating the retreating blade stall of helicopter rotor in forward flight. Three unsteady flow conditions corresponding to an essentially attached flow, light-stall, and deep-stall cases of an oscillating NACA 0015 wing experiment were chosen as test cases for computations. Results of unsteady airloads hysteresis curves, harmonics of unsteady pressures, and instantaneous flowfield patterns are presented. Some effects of grid density, time-step size, and numerical dissipation on the unsteady solutions relevant to the evaluation of turbulence models are examined. Comparison of unsteady airloads with experimental data show that all models tested are deficient in some sense and no single model predicts airloads consistently and in agreement with experiment for the three flow regimes. The chief findings are that the simple algebraic model based on the renormalization group theory (RNG) offers some improvement over the Baldwin Lomax model in all flow regimes with nearly same computational cost. The one-equation models provide significant improvement over the algebraic and the half-equation models but have their own limitations. The Baldwin-Barth model overpredicts separation and underpredicts reattachment. In contrast, the Spalart-Allmaras model underpredicts separation and overpredicts reattachment.
A computer program for the design and analysis of low-speed airfoils
NASA Technical Reports Server (NTRS)
Eppler, R.; Somers, D. M.
1980-01-01
A conformal mapping method for the design of airfoils with prescribed velocity distribution characteristics, a panel method for the analysis of the potential flow about given airfoils, and a boundary layer method have been combined. With this combined method, airfoils with prescribed boundary layer characteristics can be designed and airfoils with prescribed shapes can be analyzed. All three methods are described briefly. The program and its input options are described. A complete listing is given as an appendix.
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, 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.
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.
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
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.
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.
Active Flow Separation Control on a NACA 0015 Wing Using Fluidic Actuators
NASA Technical Reports Server (NTRS)
Melton, Latunia P.
2014-01-01
Results are presented from a recent set of wind tunnel experiments using sweeping jet actuators to control ow separation on the 30% chord trailing edge ap of a 30 deg. swept wing model with an aspect ratio (AR) of 4.35. Two sweeping jet actuator locations were examined, one on the flap shoulder and one on the trailing edge flap. The parameters that were varied included actuator momentum, freestream velocity, and trailing edge flap deflection (Delta f ) angle. The primary focus of this set of experiments was to determine the mass flow and momentum requirements for controlling separation on the flap, especially at large flap deflection angles which would be characteristic of a high lift system. Surface pressure data, force and moment data, and stereoscopic particle image velocimetry (PIV) data were acquired to evaluate the performance benefits due to applying active flow control. Improvements in lift over the majority of the wing span were obtained using sweeping jet actuator control. High momentum coefficient, Cu, levels were needed when using the actuators on the ap because they were located downstream of separation. Actuators on the flap shoulder performed slightly better but actuator size, orientation, and spacing still need to be optimized.
Optimal Control of Airfoil Flow Separation using Fluidic Excitation
NASA Astrophysics Data System (ADS)
Shahrabi, Arireza F.
This thesis deals with the control of flow separation around a symmetric airfoils with the aid of multiple synthetic jet actuators (SJAs). CFD simulation methods have been implemented to uncover the flow separation regimes and associated properties such as frequencies and momentum ratio. In the first part of the study, the SJA was studied thoroughly. Large Eddy Simulations (LES) were performed for one individual cavity; the time history of SJA of the outlet velocity profile and the net momentum imparted to the flow were analyzed. The studied SJA is asymmetrical and operates with the aid of a piezoelectric (PZT) ceramic circular plate actuator. A three-dimensional mesh for the computational domain of the SJA and the surrounding volume was developed and was used to evaluate the details of the airflow conditions inside the SJA as well as at the outlet. The vibration of the PZT ceramic actuator was used as a boundary condition in the computational model to drive the SJA. Particular attention was given to developing a predictive model of the SJA outlet velocity. Results showed that the SJA velocity output is correlated to the PZT ceramic plate vibration, especially for the first frequency mode. SJAs are a particular class of zero net mass flux (ZNMF) fluidic devices with net imparted momentum to the flow. The net momentum imparted to the flow in the separated region is such that positive enhancement during AFC operations is achieved. Flows around the NACA 0015 airfoil were simulated for a range of operating conditions. Attention was given to the active open and closed loop control solutions for an airfoil with SJA at different angles of attack and flap angles. A large number of simulations using RANS & LES models were performed to study the effects of the momentum ratio (Cμ) in the range of 0 to 11% and of the non-dimensional frequency, F+, in the range of 0 to 2 for the control of flow separation at a practical angle of attack and flap angle. The optimum value of C
1980-02-01
the elliptic cross section is considered to be more representative of the NACA 64A010 airfoil with boundary layer displacement thickness added on than...section and the flat plate airfoil with Kutta condition. The experimental results are for the NACA 64A010 airfoil at M = 0.5 and Reynolds number between...practice for actual airfoils. The experimental data shown in Fig. 3.5 are for the NACA 4 and 5 digit series airfoils (Ref. 17). The lift curve slope is
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)
Ormsbee, A. I.
1977-01-01
Airfoil geometries were developed for low speed high lift applications, such as general aviation aircraft, propellers and helicopter rotors. The primary effort was to determine the extent to which the application of turbulent boundary layer separation criteria, plus manipulation of other input parameters, specifically trailing edging velocity ratio, could be utilized to achieve high C sub Lmax airfoils with relatively low drag at C sub Lmax. Both single-element and double-element airfoils were considered. Wind tunnel testing of some airfoils was included.
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
Airfoil System for Cruising Flight
NASA Technical Reports Server (NTRS)
Shams, Qamar A. (Inventor); Liu, Tianshu (Inventor)
2014-01-01
An airfoil system includes an airfoil body and at least one flexible strip. The airfoil body has a top surface and a bottom surface, a chord length, a span, and a maximum thickness. Each flexible strip is attached along at least one edge thereof to either the top or bottom surface of the airfoil body. The flexible strip has a spanwise length that is a function of the airfoil body's span, a chordwise width that is a function of the airfoil body's chord length, and a thickness that is a function of the airfoil body's maximum thickness.
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.
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.
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.
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.
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.
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.
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 Astrophysics Data System (ADS)
Bragg, M. B.; Broeren, A. P.; Blumenthal, L. A.
2005-07-01
Past research on airfoil aerodynamics in icing are reviewed. This review emphasizes the time period after the 1978 NASA Lewis workshop that initiated the modern icing research program at NASA and the current period after the 1994 ATR accident where aerodynamics research has been more aircraft safety focused. Research pre-1978 is also briefly reviewed. Following this review, our current knowledge of iced airfoil aerodynamics is presented from a flowfield-physics perspective. This article identifies four classes of ice accretions: roughness, horn ice, streamwise ice, and spanwise-ridge ice. For each class, the key flowfield features such as flowfield separation and reattachment are discussed and how these contribute to the known aerodynamic effects of these ice shapes. Finally Reynolds number and Mach number effects on iced-airfoil aerodynamics are summarized.
NASA Technical Reports Server (NTRS)
Hylton, Larry D.
1986-01-01
Emphasis is placed on developing more accurate analytical models for predicting turbine airfoil external heat transfer rates. Performance goals of new engines require highly refined, accurate design tools to meet durability requirements. In order to obtain improvements in analytical capabilities, programs are required which focus on enhancing analytical techniques through verification of new models by comparison with relevant experimental data. The objectives of the current program are to develop an analytical approach, based on boundary layer theory, for predicting the effects of airfoil film cooling on downstream heat transfer rates and to verify the resulting analytical method by comparison of predictions with hot cascade data obtained under this program.
NASA Technical Reports Server (NTRS)
Garabedian, P. R.
1979-01-01
Computer codes for the design and analysis of transonic airfoils are considered. The design code relies on the method of complex characteristics in the hodograph plane to construct shockless airfoil. The analysis code uses artificial viscosity to calculate flows with weak shock waves at off-design conditions. Comparisons with experiments show that an excellent simulation of two dimensional wind tunnel tests is obtained. The codes have been widely adopted by the aircraft industry as a tool for the development of supercritical wing technology.
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.
Airfoil Design and Rotorcraft Performance
NASA Technical Reports Server (NTRS)
Bousman, William G.
2003-01-01
The relationship between global performance of a typical helicopter and the airfoil environment, as represented by the airfoil angles of attack and Mach number, has been examined using the comprehensive analysis CAMRAD II. A general correspondence is observed between global performance parameters, such as rotor L/D, and airfoil performance parameters, such as airfoil L/D, the drag bucket boundaries, and the divergence Mach number. Effects of design parameters such as blade twist and rotor speed variation have been examined and, in most cases, improvements observed in global performance are also observed in terms of airfoil performance. The relations observed between global Performance and the airfoil environment suggests that the emphasis in airfoil design should be for good L/D, while the maximum lift coefficient performance is less important.
Ice Roughness and Thickness Evolution on a Swept NACA 0012 Airfoil
NASA Technical Reports Server (NTRS)
McClain, Stephen T.; Vargas, Mario; Tsao, Jen-Ching
2017-01-01
Several recent studies have been performed in the Icing Research Tunnel (IRT) at NASA Glenn Research Center focusing on the evolution, spatial variations, and proper scaling of ice roughness on airfoils without sweep exposed to icing conditions employed in classical roughness studies. For this study, experiments were performed in the IRT to investigate the ice roughness and thickness evolution on a 91.44-cm (36-in.) chord NACA 0012 airfoil, swept at 30-deg with 0deg angle of attack, and exposed to both Appendix C and Appendix O (SLD) icing conditions. The ice accretion event times used in the study were less than the time required to form substantially three-dimensional structures, such as scallops, on the airfoil surface. Following each ice accretion event, the iced airfoils were scanned using a ROMER Absolute Arm laser-scanning system. The resulting point clouds were then analyzed using the self-organizing map approach of McClain and Kreeger to determine the spatial roughness variations along the surfaces of the iced airfoils. The resulting measurements demonstrate linearly increasing roughness and thickness parameters with ice accretion time. Further, when compared to dimensionless or scaled results from unswept airfoil investigations, the results of this investigation indicate that the mechanisms for early stage roughness and thickness formation on swept wings are similar to those for unswept wings.
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.
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.
NREL airfoil families for HAWTs
Tangler, J.L.; Somers, D.M.
1995-12-31
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 nine 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 1,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.
Lift enhancing tabs for airfoils
NASA Technical Reports Server (NTRS)
Ross, James C. (Inventor)
1994-01-01
A tab deployable from the trailing edge of a main airfoil element forces flow onto a following airfoil element, such as a flap, to keep the flow attached and thus enhance lift. For aircraft wings with high lift systems that include leading edge slats, the slats may also be provided with tabs to turn the flow onto the following main element.
Measuremants in the wake of an infinite swept airfoil
NASA Technical Reports Server (NTRS)
Novak, C. J.; Ramaprian, B. R.
1982-01-01
This is a report of the measurements in the trailing edge region as well as in the report of the developing wake behind a swept NACA 0012 airfoil at zero incidence and a sweep angle of 30 degrees. The measurements include both the mean and turbulent flow properties. The mean flow velocities, flow inclination and static pressure are measured using a calibrated three-hole yaw probe. The measurements of all the relevant Reynolds stress components in the wake are made using a tri-axial hot-wire probe and a digital data processing technique developed by the authors. The development of the three dimensional near-wake into a nearly two dimensional far-wake is discussed in the light of the experimental data. A complete set of wake data along with the data on the initial boundary layer in the trailing edge region of the airfoil are tabulated in an appendix to the report.
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.
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.
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.
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)
Numerical design of shockless airfoils
NASA Technical Reports Server (NTRS)
Garabedian, P. R.
1979-01-01
An attempt is made to indicate and briefly discuss only the most significant achievements of the research. The most successful contribution from the contract was the code for two dimensional analysis of airfoils in transonic flow.
Langley airfoil-research program
NASA Technical Reports Server (NTRS)
Bobbitt, P. J.
1979-01-01
An overview of past, present, and future airfoil research activities at the Langley Research Center is given. The immediate past and future occupy most of the discussion; however, past accomplishments and milestones going back to the early NACA years are dealt with in a broad-brush way to give a better perspective of current developments and programs. In addition to the historical perspective, a short description of the facilities which are now being used in the airfoil program is given. This is followed by a discussion of airfoil developments, advances in airfoil design and analysis tools (mostly those that have taken place over the past 5 or 6 years), and tunnel-wall-interference predictive methods and measurements. Future research requirements are treated.
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.
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.
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)
Dynamic airfoil stall investigations
NASA Technical Reports Server (NTRS)
Platzer, M. F.; Chandrasekhara, M. S.; Ekaterinaris, J. A.; Carr, L. W.
1992-01-01
Experimental and computational investigations of the dynamic stall phenomenon continue to attract the attention of various research groups in the major aeronautical research laboratories. There are two reasons for this continued research interest. First, the occurrence of dynamic stall on the retreating blade of helicopters imposes a severe performance limitation and thus suggests to search for ways to delay the onset of dynamic stall. Second, the lift enhancement prior to dynamic stall presents an opportunity to achieve enhanced maneuverability of fighter aircraft. A description of the major parameters affecting dynamic stall and lift and an evaluation of research efforts prior to 1988 has been given by Carr. In this paper the authors' recent progress in the development of experimental and computational methods to analyze the dynamic stall phenomena occurring on NACA 0112 airfoils is reviewed. First, the major experimental and computational approaches and results are summarized. This is followed by an assessment of our results and an outlook toward the future.
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.
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.
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.
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.
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.
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.
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.
Ice accretion simulations on airfoils
NASA Astrophysics Data System (ADS)
Özgen, S.; Uğur, N.; Görgülü, I.; Tatar, V.
2017-06-01
Ice shape predictions for a NACA0012 airfoil and collection efficiency predictions for the Twin Otter airfoil are obtained and presented. The results are validated with reference numerical and experimental data. Ice accretion modeling mainly consists of four steps: flow field solution; droplet trajectory calculations; thermodynamic analyses; and ice accretion simulation with the Extended Messinger Model. The models are implemented in a FORTRAN code to perform icing analyses for twodimensional (2D) geometries. The results are in good agreement with experimental and numerical reference data. It is deduced that increasing computational layers in calculations improves the ice shape predictions. The results indicate that collection efficiencies and impingement zone increase with increasing droplet diameter.
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.
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.
Pneumatic Spoiler Controls Airfoil Lift
NASA Technical Reports Server (NTRS)
Hunter, D.; Krauss, T.
1991-01-01
Air ejection from leading edge of airfoil used for controlled decrease of lift. Pneumatic-spoiler principle developed for equalizing lift on helicopter rotor blades. Also used to enhance aerodynamic control of short-fuselage or rudderless aircraft such as "flying-wing" airplanes. Leading-edge injection increases maneuverability of such high-performance fixed-wing aircraft as fighters.
Supercritical Flow Past Symmetrical Airfoils.
1980-12-01
about quasi-elliptic airfoil sections. The method was later extended by Boerstoel [1967] to present a catalog of solutions for certain body shapes. Bauer...Lecture Notes in Economics and Mathematical Systems, Springer- Verlag, New York, 1972. Boerstoel , J. W., "A Survey of Symmetrical Transonic Potential
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.;
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.
Pneumatic Spoiler Controls Airfoil Lift
NASA Technical Reports Server (NTRS)
Hunter, D.; Krauss, T.
1991-01-01
Air ejection from leading edge of airfoil used for controlled decrease of lift. Pneumatic-spoiler principle developed for equalizing lift on helicopter rotor blades. Also used to enhance aerodynamic control of short-fuselage or rudderless aircraft such as "flying-wing" airplanes. Leading-edge injection increases maneuverability of such high-performance fixed-wing aircraft as fighters.
The prediction of airfoil characteristics
NASA Technical Reports Server (NTRS)
Higgins, George J
1930-01-01
This report describes and develops methods by which the aerodynamic characteristics of an airfoil may be calculated with sufficient accuracy for use in airplane design. These methods for prediction are based on the present aerodynamic theory and on empirical formulas derived from data obtained in the N. A. C. A. variable density wind tunnel at a Reynolds number corresponding approximately to full scale. (author)
OUT Success Stories: Advanced Airfoils for Wind Turbines
DOE R&D Accomplishments Database
Jones, J.; Green, B.
2000-08-01
New airfoils have substantially increased the aerodynamic efficiency of wind turbines. It is clear that these new airfoils substantially increased energy output from wind turbines. Virtually all new blades built in this country today use these advanced airfoil designs.
Thin oblique airfoils at supersonic speed
NASA Technical Reports Server (NTRS)
Jone, Robert T
1946-01-01
The well-known methods of thin-airfoil theory have been extended to oblique or sweptback airfoils of finite aspect ratio moving at supersonic speeds. The cases considered thus far are symmetrical airfoils at zero lift having plan forms bounded by straight lines. Because of the conical form of the elementary flow fields, the results are comparable in simplicity to the results of the two-dimensional thin-airfoil theory for subsonic speeds. In the case of untapered airfoils swept back behind the Mach cone the pressure distribution at the center section is similar to that given by the Ackeret theory for a straight airfoil. With increasing distance from the center section the distribution approaches the form given by the subsonic-flow theory. The pressure drag is concentrated chiefly at the center section and for long wings a slight negative drag may appear on outboard sections. (author)
Boundary-layer stability and airfoil design
NASA Technical Reports Server (NTRS)
Viken, Jeffrey K.
1986-01-01
Several different natural laminar flow (NLF) airfoils have been analyzed for stability of the laminar boundary layer using linear stability codes. The NLF airfoils analyzed come from three different design conditions: incompressible; compressible with no sweep; and compressible with sweep. Some of the design problems are discussed, concentrating on those problems associated with keeping the boundary layer laminar. Also, there is a discussion on how a linear stability analysis was effectively used to improve the design for some of the airfoils.
Evaluation of a stalled airfoil analysis program
NASA Technical Reports Server (NTRS)
Rumsey, C. L.
1985-01-01
The Stalled Airfoil Analysis Program (SAAP) is a computer code for predicting the aerodynamic characteristics of an airfoil up to, and beyond, stall. SAAP is presently evaluated through comparisons with experiments and with two other theoretical methods over an extensive range of airfoils and Reynolds number conditions. SAAP modeled drag more accurately than either of the other methods, and at angles of attack below stall yielded a smoother lift variation with angle of attack.
Airfoil seal system for gas turbine engine
None, None
2013-06-25
A turbine airfoil seal system of a turbine engine having a seal base with a plurality of seal strips extending therefrom for sealing gaps between rotational airfoils and adjacent stationary components. The seal strips may overlap each other and may be generally aligned with each other. The seal strips may flex during operation to further reduce the gap between the rotational airfoils and adjacent stationary components.
Root region airfoil for wind turbine
Tangler, J.L.; Somers, D.M.
1995-05-23
A thick airfoil is described 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. 3 Figs.
Airfoil Theory at Supersonic Speed
NASA Technical Reports Server (NTRS)
Schlichting, H
1939-01-01
A theory is developed for the airfoil of finite span at supersonic speed analogous to the Prandtl airfoil theory of 1918-1919 for incompressible flow. In addition to the profile and induced drags, account must be taken at supersonic flow of still another drag, namely, the wave drag, which is independent of the wing aspect ratio. Both wave and induced drags are proportional to the square of the lift and depend on the Mach number, that is, the ratio of flight to sound speed. In general, in the case of supersonic flow, the drag-lift ratio is considerably less favorable than is the case for incompressible flow. Among others the following examples are considered: 1) lifting line with constant lift distribution (horseshoe vortex); 2) computation of wave and induced drag and the twist of a trapezoidal wing of constant lift density; 3) computation of the lift distribution and drag of an untwisted rectangular wing.
Schooling behavior of heaving flexible airfoils
NASA Astrophysics Data System (ADS)
Im, Sunghyuk; Sung, Hyung Jin
2016-11-01
The schooling behavior of rigid and flexible NACA0017 airfoils in the heaving motion is experimentally explored in a merry-go-round equipment. The airfoil was attached to the end of a horizontal support bar whose other end was connected to the freely rotating vertical axis. The axis was forced to undergo a sinusoidal motion in the vertical direction to make a pure heaving motion of the airfoils in the frequency range of 0.5 to 5 Hz. The propulsion due to the heaving airfoils is expressed by a horizontally rotating speed of the support bar. This experimental setup is simulating infinite schooling situations of airfoils in an in-phase heaving motion with the streamwise distance d. The ratio of the distance to the chord length d/ c was determined by the number of airfoils (1 <= n <= 8) . The rotational frequency F according to the heaving frequency f was measured with different experimental parameters. The schooling number S = f /(nF), representing the number of heaving oscillations between each airfoil, was introduced to explain the schooling behavior of the airfoils. The effects of the flexibility, d/ c and f on the propulsive performance were examined with the schooling behavior of the airfoils. This work was supported by the Creative Research Initiatives (No. 2016-004749) program of the National Research Foundation of Korea (MSIP).
The further development of circulation control airfoils
NASA Technical Reports Server (NTRS)
Wood, N. J.
1987-01-01
The performance trends of circulation control airfoils are reviewed and observations are made as to where improvements in performance and expansion of the flight envelope may be feasible. A new analytically defined family of airfoils is suggested, all of which maintain the fore and aft symmetry required for stopped rotor application. It is important to recognize that any improvements in section capabilities may not be totally applicable to the present vehicle operation. It remains for the designers of the rotor system to reappraise the three dimensional operating environment in view of the different airfoil operating characteristics and for the airfoil definitions to be flexible while maintaining satisfactory levels of performance.
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.
Transonic airfoil flowfield analysis using Cartesian coordinates
NASA Technical Reports Server (NTRS)
Carlson, L. A.
1975-01-01
A numerical technique for analyzing transonic airfoils is presented. The method employs the basic features of Jameson's iterative solution for the full potential equation, except that Cartesian coordinates are used rather than a grid which fits the airfoil, such as the conformal circle-plane or 'sheared parabolic' coordinates which were used previously. Comparison with previous results shows that it is not necessary to match the computational grid to the airfoil surface, and that accurate results can be obtained with a Cartesian grid for lifting supercritical airfoils.
Generalized multi-point inverse airfoil design
NASA Technical Reports Server (NTRS)
Selig, Michael S.; Maughmer, Mark D.
1991-01-01
In a rather general sense, inverse airfoil design can be taken to mean the problem of specifying a desired set of airfoil characteristics, such as the airfoil maximum thickness ratio, pitching moment, part of the velocity distribution or boundary-layer development, etc., then from this information determine the corresponding airfoil shape. This paper presents a method which approaches the design problem from this perspective. In particular, the airfoil is divided into segments along which, together with the design conditions, either the velocity distribution or boundary-layer development may be prescribed. In addition to these local desired distributions, single parameters like the airfoil thickness can be specified. The problem of finding the airfoil shape is determined by coupling an incompressible, inviscid, inverse airfoil design method with a direct integral boundary-layer analysis method and solving the resulting nonlinear equations via a multidimensional Newton iteration technique. The approach is fast and easily allows for interactive design. It is also flexible and could be adapted to solving compressible, inverse airfoil design problems.
High Lift, Low Pitching Moment Airfoils
NASA Technical Reports Server (NTRS)
Noonan, Kevin W. (Inventor)
1988-01-01
Two families of airfoil sections which can be used for helicopter/rotorcraft rotor blades or aircraft propellers of a particular shape are prepared. An airfoil of either family is one which could be produced by the combination of a camber line and a thickness distribution or a thickness distribution which is scaled from these. An airfoil of either family has a unique and improved aerodynamic performance. The airfoils of either family are intended for use as inboard sections of a helicopter rotor blade or an aircraft propeller.
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.
Ice Accretions and Icing Effects for Modern Airfoils
NASA Technical Reports Server (NTRS)
Addy, Harold E., Jr.
2000-01-01
Icing tests were conducted to document ice shapes formed on three different two-dimensional airfoils and to study the effects of the accreted ice on aerodynamic performance. The models tested were representative of airfoil designs in current use for each of the commercial transport, business jet, and general aviation categories of aircraft. The models were subjected to a range of icing conditions in an icing wind tunnel. The conditions were selected primarily from the Federal Aviation Administration's Federal Aviation Regulations 25 Appendix C atmospheric icing conditions. A few large droplet icing conditions were included. To verify the aerodynamic performance measurements, molds were made of selected ice shapes formed in the icing tunnel. Castings of the ice were made from the molds and placed on a model in a dry, low-turbulence wind tunnel where precision aerodynamic performance measurements were made. Documentation of all the ice shapes and the aerodynamic performance measurements made during the icing tunnel tests is included in this report. Results from the dry, low-turbulence wind tunnel tests are also presented.
AFSMO/AFSCL- AIRFOIL SMOOTHING AND SCALING
NASA Technical Reports Server (NTRS)
Morgan, H. L
1994-01-01
Since its early beginnings, NASA has been actively involved in the design and testing of airfoil sections for a wide variety of applications. Recently a set of programs has been developed to smooth and scale arbitrary airfoil coordinates. The smoothing program, AFSMO, utilizes both least-squares polynomial and least-squares cubic-spline techniques to iteratively smooth the second derivatives of the y-axis airfoil coordinates with respect to a transformed x-axis system which unwraps the airfoil and stretches the nose and trailing-edge regions. The corresponding smooth airfoil coordinates are then determined by solving a tridiagonal matrix of simultaneous cubic-spline equations relating the y-axis coordinates and their corresponding second derivatives. The camber and thickness distribution of the smooth airfoil are also computed. The scaling program, AFSCL, may then be used to scale the thickness distribution generated by the smoothing program to a specified maximum thickness. Once the thickness distribution has been scaled, it is combined with the camber distribution to obtain the final scaled airfoil contour. The airfoil smoothing and scaling programs are written in FORTRAN IV for batch execution and have been implemented on a CDC CYBER 170 series computer with a central memory requirement of approximately 70K (octal) of 60 bit words. Both programs generate plotted output via CALCOMP type plotting calls. These programs were developed in 1983.
Measuring Lift with the Wright Airfoils
ERIC Educational Resources Information Center
Heavers, Richard M.; Soleymanloo, Arianne
2011-01-01
In this laboratory or demonstration exercise, we mount a small airfoil with its long axis vertical at one end of a nearly frictionless rotating platform. Air from a leaf blower produces a sidewise lift force L on the airfoil and a drag force D in the direction of the air flow (Fig. 1). The rotating platform is kept in equilibrium by adding weights…
AFSMO/AFSCL- AIRFOIL SMOOTHING AND SCALING
NASA Technical Reports Server (NTRS)
Morgan, H. L
1994-01-01
Since its early beginnings, NASA has been actively involved in the design and testing of airfoil sections for a wide variety of applications. Recently a set of programs has been developed to smooth and scale arbitrary airfoil coordinates. The smoothing program, AFSMO, utilizes both least-squares polynomial and least-squares cubic-spline techniques to iteratively smooth the second derivatives of the y-axis airfoil coordinates with respect to a transformed x-axis system which unwraps the airfoil and stretches the nose and trailing-edge regions. The corresponding smooth airfoil coordinates are then determined by solving a tridiagonal matrix of simultaneous cubic-spline equations relating the y-axis coordinates and their corresponding second derivatives. The camber and thickness distribution of the smooth airfoil are also computed. The scaling program, AFSCL, may then be used to scale the thickness distribution generated by the smoothing program to a specified maximum thickness. Once the thickness distribution has been scaled, it is combined with the camber distribution to obtain the final scaled airfoil contour. The airfoil smoothing and scaling programs are written in FORTRAN IV for batch execution and have been implemented on a CDC CYBER 170 series computer with a central memory requirement of approximately 70K (octal) of 60 bit words. Both programs generate plotted output via CALCOMP type plotting calls. These programs were developed in 1983.
Airfoil Dynamic Stall and Rotorcraft Maneuverability
NASA Technical Reports Server (NTRS)
Bousman, William G.
2000-01-01
The loading of an airfoil during dynamic stall is examined in terms of the augmented lift and the associated penalties in pitching moment and drag. It is shown that once stall occurs and a leading-edge vortex is shed from the airfoil there is a unique relationship between the augmented lift, the negative pitching moment, and the increase in drag. This relationship, referred to here as the dynamic stall function, shows limited sensitivity to effects such as the airfoil section profile and Mach number, and appears to be independent of such parameters as Reynolds number, reduced frequency, and blade sweep. For single-element airfoils there is little that can be done to improve rotorcraft maneuverability except to provide good static C(l(max)) characteristics and the chord or blade number that is required to provide the necessary rotor thrust. However, multi-element airfoils or airfoils with variable geometry features can provide augmented lift in some cases that exceeds that available from a single-element airfoil. The dynamic stall function is shown to be a useful tool for the evaluation of both measured and calculated dynamic stall characteristics of single element, multi-element, and variable geometry airfoils.
Measuring Lift with the Wright Airfoils
ERIC Educational Resources Information Center
Heavers, Richard M.; Soleymanloo, Arianne
2011-01-01
In this laboratory or demonstration exercise, we mount a small airfoil with its long axis vertical at one end of a nearly frictionless rotating platform. Air from a leaf blower produces a sidewise lift force L on the airfoil and a drag force D in the direction of the air flow (Fig. 1). The rotating platform is kept in equilibrium by adding weights…
Propulsion of a flapping and oscillating airfoil
NASA Technical Reports Server (NTRS)
Garrick, I E
1937-01-01
Formulas are given for the propelling or drag force experience in a uniform air stream by an airfoil or an airfoil-aileron combination, oscillating in any of three degrees of freedom; vertical flapping, torsional oscillations about a fixed axis parallel to the span, and angular oscillations of the aileron about a hinge.
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.
Airfoil profile in a nonuniform flow
NASA Technical Reports Server (NTRS)
Polasek, J.
1978-01-01
A theory of airfoil section past two dimensional nonuniform flow is developed. The theory is based on representation of airfoil section by vortex and source distributions and it can be used for calculation of aircraft wings in homogeneous and inhomogeneous flow, as well as for calculation of straight and radial blade and vane-cascades.
NASA Astrophysics Data System (ADS)
Manela, A.
2016-07-01
The acoustic signature of an acoustically compact tandem airfoil setup in uniform high-Reynolds number flow is investigated. The upstream airfoil is considered rigid and is actuated at its leading edge with small-amplitude harmonic pitching motion. The downstream airfoil is taken passive and elastic, with its motion forced by the vortex-street excitation of the upstream airfoil. The non-linear near-field description is obtained via potential thin-airfoil theory. It is then applied as a source term into the Powell-Howe acoustic analogy to yield the far-field dipole radiation of the system. To assess the effect of downstream-airfoil elasticity, results are compared with counterpart calculations for a non-elastic setup, where the downstream airfoil is rigid and stationary. Depending on the separation distance between airfoils, airfoil-motion and airfoil-wake dynamics shift between in-phase (synchronized) and counter-phase behaviors. Consequently, downstream airfoil elasticity may act to amplify or suppress sound through the direct contribution of elastic-airfoil motion to the total signal. Resonance-type motion of the elastic airfoil is found when the upstream airfoil is actuated at the least stable eigenfrequency of the downstream structure. This, again, results in system sound amplification or suppression, depending on the separation distance between airfoils. With increasing actuation frequency, the acoustic signal becomes dominated by the direct contribution of the upstream airfoil motion, whereas the relative contribution of the elastic airfoil to the total signature turns negligible.
Manela, A.
2016-07-15
The acoustic signature of an acoustically compact tandem airfoil setup in uniform high-Reynolds number flow is investigated. The upstream airfoil is considered rigid and is actuated at its leading edge with small-amplitude harmonic pitching motion. The downstream airfoil is taken passive and elastic, with its motion forced by the vortex-street excitation of the upstream airfoil. The non-linear near-field description is obtained via potential thin-airfoil theory. It is then applied as a source term into the Powell-Howe acoustic analogy to yield the far-field dipole radiation of the system. To assess the effect of downstream-airfoil elasticity, results are compared with counterpart calculations for a non-elastic setup, where the downstream airfoil is rigid and stationary. Depending on the separation distance between airfoils, airfoil-motion and airfoil-wake dynamics shift between in-phase (synchronized) and counter-phase behaviors. Consequently, downstream airfoil elasticity may act to amplify or suppress sound through the direct contribution of elastic-airfoil motion to the total signal. Resonance-type motion of the elastic airfoil is found when the upstream airfoil is actuated at the least stable eigenfrequency of the downstream structure. This, again, results in system sound amplification or suppression, depending on the separation distance between airfoils. With increasing actuation frequency, the acoustic signal becomes dominated by the direct contribution of the upstream airfoil motion, whereas the relative contribution of the elastic airfoil to the total signature turns negligible.
Study of a new airfoil used in reversible axial fans
NASA Technical Reports Server (NTRS)
Li, Chaojun; Wei, Baosuo; Gu, Chuangang
1991-01-01
The characteristics of the reverse ventilation of axial flow are analyzed. An s shaped airfoil with a double circular arc was tested in a wind tunnel. The experimental results showed that the characteristics of this new airfoil in reverse ventilation are the same as those in normal ventilation, and that this airfoil is better than the existing airfoils used on reversible axial fans.
Airfoil flutter model suspension system
NASA Technical Reports Server (NTRS)
Reed, Wilmer H. (Inventor)
1987-01-01
A wind tunnel suspension system for testing flutter models under various loads and at various angles of attack is described. The invention comprises a mounting bracket assembly affixing the suspension system to the wind tunnel, a drag-link assembly and a compound spring arrangement comprises a plunge spring working in opposition to a compressive spring so as to provide a high stiffness to trim out steady state loads and simultaneously a low stiffness to dynamic loads. By this arrangement an airfoil may be tested for oscillatory response in both plunge and pitch modes while being held under high lifting loads in a wind tunnel.
Design optimization of transonic airfoils
NASA Technical Reports Server (NTRS)
Joh, C.-Y.; Grossman, B.; Haftka, R. T.
1991-01-01
Numerical optimization procedures were considered for the design of airfoils in transonic flow based on the transonic small disturbance (TSD) and Euler equations. A sequential approximation optimization technique was implemented with an accurate approximation of the wave drag based on the Nixon's coordinate straining approach. A modification of the Euler surface boundary conditions was implemented in order to efficiently compute design sensitivities without remeshing the grid. Two effective design procedures producing converged designs in approximately 10 global iterations were developed: interchanging the role of the objective function and constraint and the direct lift maximization with move limits which were fixed absolute values of the design variables.
NASA low- and medium-speed airfoil development
NASA Technical Reports Server (NTRS)
Mcghee, R. J.; Beasley, W. D.; Whitcomb, R. T.
1979-01-01
The status of NASA low and medium speed airfoil research is discussed. Effects of airfoil thickness-chord ratios varying from 9 percent to 21 percent on the section characteristics for a design lift coefficient of 0.40 are presented for the initial low speed family of airfoils. Also, modifications to the 17-percent low-speed airfoil to reduce the pitching-moment coefficient and to the 21-percent low speed airfoil results are shown for two new medium speed airfoils with thickness ratios of 13 percent and 17 percent and design-lift coefficients of 0.30. Applications of NASA-developed airfoils to general aviation aircraft are summarized.
Spline-Based Smoothing of Airfoil Curvatures
NASA Technical Reports Server (NTRS)
Li, W.; Krist, S.
2008-01-01
Constrained fitting for airfoil curvature smoothing (CFACS) is a splinebased method of interpolating airfoil surface coordinates (and, concomitantly, airfoil thicknesses) between specified discrete design points so as to obtain smoothing of surface-curvature profiles in addition to basic smoothing of surfaces. CFACS was developed in recognition of the fact that the performance of a transonic airfoil is directly related to both the curvature profile and the smoothness of the airfoil surface. Older methods of interpolation of airfoil surfaces involve various compromises between smoothing of surfaces and exact fitting of surfaces to specified discrete design points. While some of the older methods take curvature profiles into account, they nevertheless sometimes yield unfavorable results, including curvature oscillations near end points and substantial deviations from desired leading-edge shapes. In CFACS as in most of the older methods, one seeks a compromise between smoothing and exact fitting. Unlike in the older methods, the airfoil surface is modified as little as possible from its original specified form and, instead, is smoothed in such a way that the curvature profile becomes a smooth fit of the curvature profile of the original airfoil specification. CFACS involves a combination of rigorous mathematical modeling and knowledge-based heuristics. Rigorous mathematical formulation provides assurance of removal of undesirable curvature oscillations with minimum modification of the airfoil geometry. Knowledge-based heuristics bridge the gap between theory and designers best practices. In CFACS, one of the measures of the deviation of an airfoil surface from smoothness is the sum of squares of the jumps in the third derivatives of a cubicspline interpolation of the airfoil data. This measure is incorporated into a formulation for minimizing an overall deviation- from-smoothness measure of the airfoil data within a specified fitting error tolerance. CFACS has been
Uncertainty Quantification for Airfoil Icing
NASA Astrophysics Data System (ADS)
DeGennaro, Anthony Matteo
Ensuring the safety of airplane flight in icing conditions is an important and active arena of research in the aerospace community. Notwithstanding the research, development, and legislation aimed at certifying airplanes for safe operation, an analysis of the effects of icing uncertainties on certification quantities of interest is generally lacking. The central objective of this thesis is to examine and analyze problems in airfoil ice accretion from the standpoint of uncertainty quantification. We focus on three distinct areas: user-informed, data-driven, and computational uncertainty quantification. In the user-informed approach to uncertainty quantification, we discuss important canonical icing classifications and show how these categories can be modeled using a few shape parameters. We then investigate the statistical effects of these parameters. In the data-driven approach, we build statistical models of airfoil ice shapes from databases of actual ice shapes, and quantify the effects of these parameters. Finally, in the computational approach, we investigate the effects of uncertainty in the physics of the ice accretion process, by perturbing the input to an in-house numerical ice accretion code that we develop in this thesis.
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.
Airfoil shape for a turbine bucket
Hyde, Susan Marie; By, Robert Romany; Tressler, Judd Dodge; Schaeffer, Jon Conrad; Sims, Calvin Levy
2005-06-28
Third stage turbine buckets have airfoil profiles substantially in accordance with Cartesian coordinate values of X, Y and Z set forth Table I wherein X and Y values are in inches and the Z values are non-dimensional values from 0 to 0.938 convertible to Z distances in inches by multiplying the Z values by the height of the airfoil in inches. The X and Y values are distances which, when connected by smooth continuing arcs, define airfoil profile sections at each distance Z. The profile sections at each distance Z are joined smoothly to one another to form a complete airfoil shape. The X and Y distances may be scalable as a function of the same constant or number to provide a scaled up or scaled down airfoil section for the bucket. The nominal airfoil given by the X, Y and Z distances lies within an envelop of .+-.0.150 inches in directions normal to the surface of the airfoil.
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.
Turbine airfoil with outer wall thickness indicators
Marra, John J; James, Allister W; Merrill, Gary B
2013-08-06
A turbine airfoil usable in a turbine engine and including a depth indicator for determining outer wall blade thickness. The airfoil may include an outer wall having a plurality of grooves in the outer surface of the outer wall. The grooves may have a depth that represents a desired outer surface and wall thickness of the outer wall. The material forming an outer surface of the outer wall may be removed to be flush with an innermost point in each groove, thereby reducing the wall thickness and increasing efficiency. The plurality of grooves may be positioned in a radially outer region of the airfoil proximate to the tip.
NASA Astrophysics Data System (ADS)
Liu, Yixiong; Yang, Ce; Song, Xiancheng
2015-04-01
A new airfoil shape parameterization method is developed, which extended the Bezier curve to the generalized form with adjustable shape parameters. The local control parameters at airfoil leading and trailing edge regions are enhanced, where have significant effect on the aerodynamic performance of wind turbine. The results show this improved parameterization method has advantages in the fitting characteristics of geometry shape and aerodynamic performance comparing with other three common airfoil parameterization methods. The new parameterization method is then applied to airfoil shape optimization for wind turbine using Genetic Algorithm (GA), and the wind turbine special airfoil, DU93-W-210, is optimized to achieve the favorable Cl/Cd at specified flow conditions. The aerodynamic characteristic of the optimum airfoil is obtained by solving the RANS equations in computational fluid dynamics (CFD) method, and the optimization convergence curves show that the new parameterization method has good convergence rate in less number of generations comparing with other methods. It is concluded that the new method not only has well controllability and completeness in airfoil shape representation and provides more flexibility in expressing the airfoil geometry shape, but also is capable to find efficient and optimal wind turbine airfoil. Additionally, it is shown that a suitable parameterization method is helpful for improving the convergence rate of the optimization algorithm.
Turbine airfoil to shroud attachment method
Campbell, Christian X; Kulkarni, Anand A; James, Allister W; Wessell, Brian J; Gear, Paul J
2014-12-23
Bi-casting a platform (50) onto an end portion (42) of a turbine airfoil (31) after forming a coating of a fugitive material (56) on the end portion. After bi-casting the platform, the coating is dissolved and removed to relieve differential thermal shrinkage stress between the airfoil and platform. The thickness of the coating is varied around the end portion in proportion to varying amounts of local differential process shrinkage. The coating may be sprayed (76A, 76B) onto the end portion in opposite directions parallel to a chord line (41) of the airfoil or parallel to a mid-platform length (80) of the platform to form respective layers tapering in thickness from the leading (32) and trailing (34) edges along the suction side (36) of the airfoil.
Airfoil self-noise and prediction
NASA Technical Reports Server (NTRS)
Brooks, Thomas F.; Pope, D. Stuart; Marcolini, Michael A.
1989-01-01
A prediction method is developed for the self-generated noise of an airfoil blade encountering smooth flow. The prediction methods for the individual self-noise mechanisms are semiempirical and are based on previous theoretical studies and data obtained from tests of two- and three-dimensional airfoil blade sections. The self-noise mechanisms are due to specific boundary-layer phenomena, that is, the boundary-layer turbulence passing the trailing edge, separated-boundary-layer and stalled flow over an airfoil, vortex shedding due to laminar boundary layer instabilities, vortex shedding from blunt trailing edges, and the turbulent vortex flow existing near the tip of lifting blades. The predictions are compared successfully with published data from three self-noise studies of different airfoil shapes. An application of the prediction method is reported for a large scale-model helicopter rotor, and the predictions compared well with experimental broadband noise measurements. A computer code of the method is given.
Third-stage turbine bucket airfoil
Pirolla, Peter Paul; Siden, Gunnar Leif; Humanchuk, David John; Brassfield, Steven Robert; Wilson, Paul Stuart
2002-01-01
The third-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 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.
Second-stage turbine bucket airfoil
Wang, John Zhiqiang; By, Robert Romany; Sims, Calvin L.; Hyde, Susan Marie
2002-01-01
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 and Y 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. The second-stage wheel has sixty buckets.
Low speed airfoil design and analysis
NASA Technical Reports Server (NTRS)
Eppler, R.; Somers, D. M.
1979-01-01
A low speed airfoil design and analysis program was developed which contains several unique features. In the design mode, the velocity distribution is not specified for one but many different angles of attack. Several iteration options are included which allow the trailing edge angle to be specified while other parameters are iterated. For airfoil analysis, a panel method is available which uses third-order panels having parabolic vorticity distributions. The flow condition is satisfied at the end points of the panels. Both sharp and blunt trailing edges can be analyzed. The integral boundary layer method with its laminar separation bubble analog, empirical transition criterion, and precise turbulent boundary layer equations compares very favorably with other methods, both integral and finite difference. Comparisons with experiment for several airfoils over a very wide Reynolds number range are discussed. Applications to high lift airfoil design are also demonstrated.
Ice Accretions on Modern Airfoils Investigated
NASA Technical Reports Server (NTRS)
Addy, Harold E., Jr.
2000-01-01
The Icing Branch at the NASA Glenn Research Center at Lewis Field initiated and conducted the Modern Airfoils Ice Accretions project to identify ice shapes and determine their effects on the aerodynamic performance of aircraft, particularly on lift and drag. Previous aircraft ice shape and performance documentation focused on a few, older airfoils. This permitted more basic studies of the ice accretion process to be undertaken. However, having established both a working data base of ice shapes and the capability to predict these shapes for basic airfoils, questions arose about how ice might accrete differently on airfoils more representative of those being designed and flown on various aircraft today. Similarly, information about how these ice shapes would affect aerodynamic performance was needed.
Measuring Lift with the Wright Airfoils
NASA Astrophysics Data System (ADS)
Heavers, Richard M.; Soleymanloo, Arianne
2011-11-01
In this laboratory or demonstration exercise, we mount a small airfoil with its long axis vertical at one end of a nearly frictionless rotating platform. Air from a leaf blower2 produces a sidewise lift force L on the airfoil and a drag force D in the direction of the air flow (Fig. 1). The rotating platform is kept in equilibrium by adding weights (the measured values of L) to the lower end of a string passing over a pulley and connected to the other end of the rotating platform (Fig. 2). Our homemade airfoils are similar to those tested by the Wright brothers in 1901. From our lift plots in Fig. 3, we can draw the same conclusions as the Wrights about the influence of an airfoil's curvature and shape on lift.
Pressure Distribution Over Airfoils with Fowler Flaps
NASA Technical Reports Server (NTRS)
Wenzinger, Carl J; Anderson, Walter B
1938-01-01
Report presents the results of tests made of a Clark y airfoil with a Clark y Fowler flap and of an NACA 23012 airfoil with NACA Fowler flaps. Some of the tests were made in the 7 by 10-foot wind tunnel and others in the 5-foot vertical wind tunnel. The pressures were measured on the upper and lower surfaces at one chord section both on the main airfoils and on the flaps for several angles of attack with the flaps located at the maximum-lift settings. A test installation was used in which the model was mounted in the wind tunnel between large end planes so that two-dimensional flow was approximated. The data are given in the form of pressure-distribution diagrams and as plots of calculated coefficients for the airfoil-and-flap combinations and for the flaps alone.
Liang, George [Palm City, FL
2011-01-18
An airfoil is provided for a gas turbine comprising an outer structure comprising a first wall, an inner structure comprising a second wall spaced relative to the first wall such that a cooling gap is defined between at least portions of the first and second walls, and seal structure provided within the cooling gap between the first and second walls for separating the cooling gap into first and second cooling fluid impingement gaps. An inner surface of the second wall may define an inner cavity. The inner structure may further comprise a separating member for separating the inner cavity of the inner structure into a cooling fluid supply cavity and a cooling fluid collector cavity. The second wall may comprise at least one first impingement passage, at least one second impingement passage, and at least one bleed passage.
Modeling and Grid Generation of Iced Airfoils
NASA Technical Reports Server (NTRS)
Vickerman, Mary B.; Baez, Marivell; Braun, Donald C.; Hackenberg, Anthony W.; Pennline, James A.; Schilling, Herbert W.
2007-01-01
SmaggIce Version 2.0 is a software toolkit for geometric modeling and grid generation for two-dimensional, singleand multi-element, clean and iced airfoils. A previous version of SmaggIce was described in Preparing and Analyzing Iced Airfoils, NASA Tech Briefs, Vol. 28, No. 8 (August 2004), page 32. To recapitulate: Ice shapes make it difficult to generate quality grids around airfoils, yet these grids are essential for predicting ice-induced complex flow. This software efficiently creates high-quality structured grids with tools that are uniquely tailored for various ice shapes. SmaggIce Version 2.0 significantly enhances the previous version primarily by adding the capability to generate grids for multi-element airfoils. This version of the software is an important step in streamlining the aeronautical analysis of ice airfoils using computational fluid dynamics (CFD) tools. The user may prepare the ice shape, define the flow domain, decompose it into blocks, generate grids, modify/divide/merge blocks, and control grid density and smoothness. All these steps may be performed efficiently even for the difficult glaze and rime ice shapes. Providing the means to generate highly controlled grids near rough ice, the software includes the creation of a wrap-around block (called the "viscous sublayer block"), which is a thin, C-type block around the wake line and iced airfoil. For multi-element airfoils, the software makes use of grids that wrap around and fill in the areas between the viscous sub-layer blocks for all elements that make up the airfoil. A scripting feature records the history of interactive steps, which can be edited and replayed later to produce other grids. Using this version of SmaggIce, ice shape handling and grid generation can become a practical engineering process, rather than a laborious research effort.
Transonic Airfoils with a Given Pressure Distribution,
1981-06-01
erovse sidst necesosar mod Ideatify b lock mmb)L An inverse design procedure for airfoils, based on hodograph techniques, has been developed. For...w L-:- " " -- - r- L i -- _ 9 ABSTRACT An inverse design procedure for airfoils, based on hodograph tech...generated in the hodograph plane by Nieuwand,5 Bauer, Garabedian and Korn,6 Boerstoel and Huizing,7 and Sobieczky.8 More recently, the development of
Transonic airfoil design using Cartesian coordinates
NASA Technical Reports Server (NTRS)
Carlson, L. A.
1976-01-01
A numerical technique for designing transonic airfoils having a prescribed pressure distribution (the inverse problem) is presented. The method employs the basic features of Jameson's iterative solution for the full potential equation, except that inverse boundary conditions and Cartesian coordinates are used. The method is a direct-inverse approach that controls trailing-edge closure. Examples show the application of the method to design aft-cambered and other airfoils specifically for transonic flight.
Unsteady Pressure Distributions on Airfoils in Cascade.
1980-04-01
of thin airfoil theory has been used by Henderson (-ftj’ and Bruce (1-7-)’to derive expressions for the unsteady response which includes the cascade...model in conjunction with the assumptions of thin airfoil theory has been used by Henderson (16) and Bruce (17) to derive expressions for the unsteady...effect, that is, a sharp change in the unsteady lift when the disturbance wavelength equals the blade spacing. Bruce (19) further extends this theory to
Propulsion by active and passive airfoil oscillation
NASA Astrophysics Data System (ADS)
Mackowski, A. W.; Williamson, C. H. K.
2013-11-01
Oscillating airfoils have been the subject of much research both as a mechanism of propulsion in engineering devices as well as a model of understanding how fish, birds, and insects produce thrust and maneuvering forces. Additionally, the jet or wake generated by an oscillating airfoil exhibits a multitude of vortex patterns, which are an interesting study in their own right. We present PIV measurements of the vortex flow behind an airfoil undergoing controlled pitching oscillations at moderate Reynolds number. As a method of propulsion, oscillating foils have been found to be capable performers when undergoing both pitching and heaving motions [Anderson et al. 1998]. While an airfoil undergoing only pitching motion is a relatively inefficient propulsor, we examine the effect of adding passive dynamics to the system: for example, actuated pitching with a passive spring in the heave direction. Practically speaking, a mechanical system with such an arrangement has the potential to reduce the cost and complexity of an oscillating airfoil propulsor. To study an airfoil undergoing both active and passive motion, we employ our ``cyber-physical fluid dynamics'' technique [Mackowski & Williamson, 2011] to simulate the effects of passive dynamics in a physical experiment.
Numerical investigation of multi-element airfoils
NASA Technical Reports Server (NTRS)
Cummings, Russell M.
1993-01-01
The flow over multi-element airfoils with flat-plate lift-enhancing tabs was numerically investigated. Tabs ranging in height from 0.25 percent to 1.25 percent of the reference airfoil chord were studied near the trailing edge of the main-element. This two-dimensional numerical simulation employed an incompressible Navier-Stokes solver on a structured, embedded grid topology. New grid refinements were used to improve the accuracy of the solution near the overlapping grid boundaries. The effects of various tabs were studied at a constant Reynolds number on a two-element airfoil with a slotted flap. Both computed and measured results indicated that a tab in the main-element cove improved the maximum lift and lift-to-drag ratio relative to the baseline airfoil without a tab. Computed streamlines revealed that the additional turning caused by the tab may reduce the amount of separated flow on the flap. A three-element airfoil was also studied over a range of Reynolds numbers. For the optimized flap rigging, the computed and measured Reynolds number effects were similar. When the flap was moved from the optimum position, numerical results indicated that a tab may help to reoptimize the airfoil to within 1 percent of the optimum flap case.
Analysis of Non-symmetrical Flapping Airfoils
NASA Astrophysics Data System (ADS)
Beng Tay, Wee; Lim, Kah Bin
2007-11-01
Simulations have been done to assess the performance of different types of non-symmetrical airfoils on lift, thrust and propulsive efficiency under different flapping configurations at a Reynolds number of 10,000. The variables studied include the Stroudal number, reduced frequency, pitch angle and phase angle difference. In order to analyze the variables more efficiently, the Design of Experiments using the response surface methodology is applied. The simulation results show that besides the flapping configuration, airfoil shape also has a profound effect on the efficiency, thrust and lift production. The 4 factors have different levels of significance on the responses, indicating the shape of the airfoil plays a part as well. Thrust production depends more heavily on these parameters, rather than the shape of the airfoil. On the other hand, lift production is primarily dominated by its airfoil shape. Efficiency falls somewhere in between. Two-factor interactions among the variables also exist in efficiency and thrust production. Vorticity plots are analyzed to explain some of the results. Overall, the s1020 airfoil is able to provide relatively good efficiency and at the same time generate high thrust and lift force. These results can be used to help in the design of a better ornithopter's wing.
Simulation of a Controlled Airfoil with Jets
NASA Technical Reports Server (NTRS)
Allan, Brian G.; Holt, Maurice; Packard, Andrew
1997-01-01
Numerical simulations of a two-dimensional airfoil, controlled by an applied moment in pitch and an airfoil controlled by jets, were investigated. These simulations couple the Reynolds-averaged Navier-Stokes equations and Euler's equations of rigid body motion, with an active control system. Controllers for both systems were designed to track altitude commands and were evaluated by simulating a closed-loop altitude step response using the coupled system. The airfoil controlled by a pitching moment used an optimal state feedback controller. A closed-loop simulation, of the airfoil with an applied moment, showed that the trajectories compared very well with quasi-steady aerodynamic theory, providing a measure of validation. The airfoil with jets used a controller designed by robust control methods. A linear plant model for this system was identified using open-loop data generated by the nonlinear coupled system. A closed-loop simulation of the airfoil with jets, showed good tracking of an altitude command. This simulation also showed oscillations in the control input as a result of dynamics not accounted for in the control design. This research work demonstrates how computational fluid dynamics, coupled with rigid body dynamics, and a control law can be used to prototype control systems in problematic nonlinear flight regimes.
Isolated and cascade airfoils with prescribed velocity distribution
NASA Technical Reports Server (NTRS)
Goldstein, Arthur W; Jerison, Meyer
1947-01-01
An exact solution of the problem of designing an airfoil with a prescribed velocity distribution on the suction surface in a given uniform flow of an incompressible perfect fluid is obtained by replacing the boundary of the airfoil by vortices. By this device, a method of solution is developed that is applicable both to isolated airfoils and to airfoils in cascade. The conformal transformation of the designed airfoil into a circle can then be obtained and the velocity distribution at any angle of attack computed. Numerical illustrations of the method are given for the airfoil in cascade.
Dragan C. Curcija
2006-09-15
This is the summary page for the technical and other reports on the DOE Cooperative Agreement DE-FC36-94CH10604 for the period of January 1, 2000 to December 31, 2004. The progress, technical and other reports and publications are consolidated by the contracting year and also by the cooperative agreement tasks. The listing sorted by tasks is also sub-sorted by fiscal year for easier navigation. These listings are given in appendix A and Appendix B of this summary report. Individual report files are located in each fiscal year directory (i.e., FY00, FY01, etc. up to FY04). The complete listing and report files are also posted on the web site and is fully navigable by these two criteria. The web site is at: http://www.ceere.org/beep/beep{_}pubsanddownloads.html. More significant and less obvious part of deliverables are applications of this research, which are used in everyday operations of NFRC, software tools and manufacturers design practice, which has significantly changed as a result of this and related research efforts.
Program manual for the Eppler airfoil inversion program
NASA Technical Reports Server (NTRS)
Thomson, W. G.
1975-01-01
A computer program is described for calculating the profile of an airfoil as well as the boundary layer momentum thickness and energy form parameter. The theory underlying the airfoil inversion technique developed by Eppler is discussed.
AirfoilPrep.py Documentation: Release 0.1.0
Ning, S. A.
2013-09-01
AirfoilPrep.py provides functionality to preprocess aerodynamic airfoil data. Essentially, the module is an object oriented version of the AirfoilPrep spreadsheet with additional functionality and is written in the Python language. It allows the user to read in two-dimensional aerodynamic airfoil data, apply three-dimensional rotation corrections for wind turbine applications, and extend the datato very large angles of attack. This document discusses installation, usage, and documentation of the module.
Critical Mach Numbers of Thin Airfoil Sections with Plain Flaps
NASA Technical Reports Server (NTRS)
Pardee, Otway O'm.; Heaslet, Max A.
1946-01-01
Critical Mach number as function of lift coefficient is determined for certain moderately thick NACA low-drag airfoils. Results, given graphically, included calculations on same airfoil sections with plain flaps for small flap deflections. Curves indicate optimum critical conditions for airfoils with flaps in such form that they can be compared with corresponding results for zero flap deflections. Plain flaps increase life-coefficient range for which critical Mach number is in region of high values characteristic of low-drag airfoils.
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%.
Aerodynamic Analysis Over Double Wedge Airfoil
NASA Astrophysics Data System (ADS)
Prasad, U. S.; Ajay, V. S.; Rajat, R. H.; Samanyu, S.
2017-05-01
Aeronautical studies are being focused more towards supersonic flights and methods to attain a better and safer flight with highest possible performance. Aerodynamic analysis is part of the whole procedure, which includes focusing on airfoil shapes which will permit sustained flight of aircraft at these speeds. Airfoil shapes differ based on the applications, hence the airfoil shapes considered for supersonic speeds are different from the ones considered for Subsonic. The present work is based on the effects of change in physical parameter for the Double wedge airfoil. Mach number range taken is for transonic and supersonic. Physical parameters considered for the Double wedge case with wedge angle (ranging from 5 degree to 15 degree. Available Computational tools are utilized for analysis. Double wedge airfoil is analysed at different Angles of attack (AOA) based on the wedge angle. Analysis is carried out using fluent at standard conditions with specific heat ratio taken as 1.4. Manual calculations for oblique shock properties are calculated with the help of Microsoft excel. MATLAB is used to form a code for obtaining shock angle with Mach number and wedge angle at the given parameters. Results obtained from manual calculations and fluent analysis are cross checked.
Coating-Substrate Systems for Thermomechanically Durable Turbine Airfoils
2015-06-30
Technical Report 4. TITLE AND SUBTITLE Coating - Substrate Systems for Thermomechanically Durable Turbine Airfoils 6. AUTHOR(S) Dr. Tresa Pollock 3...Thermomechanically Durable Turbine Airfoils Final Report ONRGrant#N00014-l 1-1-0616 Technical Contact (Principal Investigator) Tresa M. Pollock Materials...Substrate Systems for Thermomechanically Durable Turbine Airfoils 1. Summary In the severe operating environments encountered in Naval ship
Compressor airfoil tip clearance optimization system
Little, David A.; Pu, Zhengxiang
2015-08-18
A compressor airfoil tip clearance optimization system for reducing a gap between a tip of a compressor airfoil and a radially adjacent component of a turbine engine is disclosed. The turbine engine may include ID and OD flowpath boundaries configured to minimize compressor airfoil tip clearances during turbine engine operation in cooperation with one or more clearance reduction systems that are configured to move the rotor assembly axially to reduce tip clearance. The configurations of the ID and OD flowpath boundaries enhance the effectiveness of the axial movement of the rotor assembly, which includes movement of the ID flowpath boundary. During operation of the turbine engine, the rotor assembly may be moved axially to increase the efficiency of the turbine engine.
Options for Robust Airfoil Optimization under Uncertainty
NASA Technical Reports Server (NTRS)
Padula, Sharon L.; Li, Wu
2002-01-01
A robust optimization method is developed to overcome point-optimization at the sampled design points. This method combines the best features from several preliminary methods proposed by the authors and their colleagues. The robust airfoil shape optimization is a direct method for drag reduction over a given range of operating conditions and has three advantages: (1) it prevents severe degradation in the off-design performance by using a smart descent direction in each optimization iteration, (2) it uses a large number of spline control points as design variables yet the resulting airfoil shape does not need to be smoothed, and (3) it allows the user to make a tradeoff between the level of optimization and the amount of computing time consumed. For illustration purposes, the robust optimization method is used to solve a lift-constrained drag minimization problem for a two-dimensional (2-D) airfoil in Euler flow with 20 geometric design variables.
Vortex noise from nonrotating cylinders and airfoils
NASA Technical Reports Server (NTRS)
Schlinker, R. H.; Amiet, R. K.; Fink, M. R.
1976-01-01
An experimental study of vortex-shedding noise was conducted in an acoustic research tunnel over a Reynolds-number range applicable to full-scale helicopter tail-rotor blades. Two-dimensional tapered-chord nonrotating models were tested to simulate the effect of spanwise frequency variation on the vortex-shedding mechanism. Both a tapered circular cylinder and tapered airfoils were investigated. The results were compared with data for constant-diameter cylinder and constant-chord airfoil models also tested during this study. Far-field noise, surface pressure fluctuations, and spanwise correlation lengths were measured for each configuration. Vortex-shedding noise for tapered cylinders and airfoils was found to contain many narrowband-random peaks which occurred within a range of frequencies corresponding to a predictable Strouhal number referenced to the maximum and minimum chord. The noise was observed to depend on surface roughness and Reynolds number.
Feedback in separated flows over symmetric airfoils
NASA Technical Reports Server (NTRS)
Atassi, H. M.
1984-01-01
For a flow over an airfoil with laminar separation, a feedback cycle may exist whereby a Kelvin-Helmholtz instability wave emanating from the separation point on the airfoil surface grows along the shear layer and is diffracted as it interacts with the sharp trailing edge of the airfoil, causing acoustic radiation which, in turn, propagates upstream and regenerates the initial instability wave. The analysis is restricted to the high frequency limit. Solutions to the boundary-value problem are obtained using the slowly varying approximation and the method of matched asymptotic expansions. Resonant solutions exist for certain discrete values of the Reynolds and Strouhal numbers. The results are discussed and compared with available data.
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.
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.
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.
Modifying Airfoils for Low Reynolds Flight
NASA Astrophysics Data System (ADS)
Ong, Christopher; Carnasciali, Maria-Isabel
2015-11-01
There has been increased interest in Micro Air Vehicles (MAV) by both the private and government sectors. MAVs are miniature classed-UAVs that can operate in tighter spaces in urban or wooded regions. Sizes vary - from that of an insect to that of small bird - depending on intended functionality and usually operate at much lower speeds. Studies have shown that the aerodynamic performance of well-known airfoils can change significantly at low Reynolds numbers. In this work, we examine via parametric CFD analysis tools the behavior of airfoils at low Reynolds values. Furthermore, we investigate the impact of adding bio-inspired features to the airfoils such as humps or dimples. Results will be presented in comparison to established values.
TAIR- TRANSONIC AIRFOIL ANALYSIS COMPUTER CODE
NASA Technical Reports Server (NTRS)
Dougherty, F. C.
1994-01-01
The Transonic Airfoil analysis computer code, TAIR, was developed to employ a fast, fully implicit algorithm to solve the conservative full-potential equation for the steady transonic flow field about an arbitrary airfoil immersed in a subsonic free stream. The full-potential formulation is considered exact under the assumptions of irrotational, isentropic, and inviscid flow. These assumptions are valid for a wide range of practical transonic flows typical of modern aircraft cruise conditions. The primary features of TAIR include: a new fully implicit iteration scheme which is typically many times faster than classical successive line overrelaxation algorithms; a new, reliable artifical density spatial differencing scheme treating the conservative form of the full-potential equation; and a numerical mapping procedure capable of generating curvilinear, body-fitted finite-difference grids about arbitrary airfoil geometries. Three aspects emphasized during the development of the TAIR code were reliability, simplicity, and speed. The reliability of TAIR comes from two sources: the new algorithm employed and the implementation of effective convergence monitoring logic. TAIR achieves ease of use by employing a "default mode" that greatly simplifies code operation, especially by inexperienced users, and many useful options including: several airfoil-geometry input options, flexible user controls over program output, and a multiple solution capability. The speed of the TAIR code is attributed to the new algorithm and the manner in which it has been implemented. Input to the TAIR program consists of airfoil coordinates, aerodynamic and flow-field convergence parameters, and geometric and grid convergence parameters. The airfoil coordinates for many airfoil shapes can be generated in TAIR from just a few input parameters. Most of the other input parameters have default values which allow the user to run an analysis in the default mode by specifing only a few input parameters
Transonic flow theory of airfoils and wings
NASA Technical Reports Server (NTRS)
Garabedian, P. R.
1976-01-01
There are plans to use the supercritical wing on the next generation of commercial aircraft so as to economize on fuel consumption by reducing drag. Computer codes have served well in meeting the consequent demand for new wing sections. The possibility of replacing wind tunnel tests by computational fluid dynamics is discussed. Another approach to the supercritical wing is through shockless airfoils. A novel boundary value problem in the hodograph plane is studied that enables one to design a shockless airfoil so that its pressure distribution very nearly takes on data that are prescribed.
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.
An efficient algorithm for numerical airfoil optimization
NASA Technical Reports Server (NTRS)
Vanderplaats, G. N.
1979-01-01
A new optimization algorithm is presented. The method is based on sequential application of a second-order Taylor's series approximation to the airfoil characteristics. Compared to previous methods, design efficiency improvements of more than a factor of 2 are demonstrated. If multiple optimizations are performed, the efficiency improvements are more dramatic due to the ability of the technique to utilize existing data. The method is demonstrated by application to subsonic and transonic airfoil design but is a general optimization technique and is not limited to a particular application or aerodynamic analysis.
TAIR- TRANSONIC AIRFOIL ANALYSIS COMPUTER CODE
NASA Technical Reports Server (NTRS)
Dougherty, F. C.
1994-01-01
The Transonic Airfoil analysis computer code, TAIR, was developed to employ a fast, fully implicit algorithm to solve the conservative full-potential equation for the steady transonic flow field about an arbitrary airfoil immersed in a subsonic free stream. The full-potential formulation is considered exact under the assumptions of irrotational, isentropic, and inviscid flow. These assumptions are valid for a wide range of practical transonic flows typical of modern aircraft cruise conditions. The primary features of TAIR include: a new fully implicit iteration scheme which is typically many times faster than classical successive line overrelaxation algorithms; a new, reliable artifical density spatial differencing scheme treating the conservative form of the full-potential equation; and a numerical mapping procedure capable of generating curvilinear, body-fitted finite-difference grids about arbitrary airfoil geometries. Three aspects emphasized during the development of the TAIR code were reliability, simplicity, and speed. The reliability of TAIR comes from two sources: the new algorithm employed and the implementation of effective convergence monitoring logic. TAIR achieves ease of use by employing a "default mode" that greatly simplifies code operation, especially by inexperienced users, and many useful options including: several airfoil-geometry input options, flexible user controls over program output, and a multiple solution capability. The speed of the TAIR code is attributed to the new algorithm and the manner in which it has been implemented. Input to the TAIR program consists of airfoil coordinates, aerodynamic and flow-field convergence parameters, and geometric and grid convergence parameters. The airfoil coordinates for many airfoil shapes can be generated in TAIR from just a few input parameters. Most of the other input parameters have default values which allow the user to run an analysis in the default mode by specifing only a few input parameters
Analysis of airfoil leading edge separation bubbles
NASA Technical Reports Server (NTRS)
Carter, J. E.; Vatsa, V. N.
1982-01-01
A local inviscid-viscous interaction technique was developed for the analysis of low speed airfoil leading edge transitional separation bubbles. In this analysis an inverse boundary layer finite difference analysis is solved iteratively with a Cauchy integral representation of the inviscid flow which is assumed to be a linear perturbation to a known global viscous airfoil analysis. Favorable comparisons with data indicate the overall validity of the present localized interaction approach. In addition numerical tests were performed to test the sensitivity of the computed results to the mesh size, limits on the Cauchy integral, and the location of the transition region.
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.
Stiffness characteristics of airfoils under pulse loading
NASA Astrophysics Data System (ADS)
Turner, Kevin Eugene
The turbomachinery industry continually struggles with the adverse effects of contact rubs between airfoils and casings. The key parameter controlling the severity of a given rub event is the contact load produced when the airfoil tips incur into the casing. These highly non-linear and transient forces are difficult to calculate and their effects on the static and rotating components are not well understood. To help provide this insight, experimental and analytical capabilities have been established and exercised through an alliance between GE Aviation and The Ohio State University Gas Turbine Laboratory. One of the early findings of the program is the influence of blade flexibility on the physics of rub events. The core focus of the work presented in this dissertation is to quantify the influence of airfoil flexibility through a novel modeling approach that is based on the relationship between applied force duration and maximum tip deflection. This relationship is initially established using a series of forward, non-linear and transient analyses in which simulated impulse rub loads are applied. This procedure, although effective, is highly inefficient and costly to conduct by requiring numerous explicit simulations. To alleviate this issue, a simplified model, named the pulse magnification model, is developed that only requires a modal analysis and a static analyses to fully describe how the airfoil stiffness changes with respect to load duration. Results from the pulse magnification model are compared to results from the full transient simulation method and to experimental results, providing sound verification for the use of the modeling approach. Furthermore, a unique and highly efficient method to model airfoil geometries was developed and is outlined in this dissertation. This method produces quality Finite Element airfoil definitions directly from a fully parameterized mathematical model. The effectiveness of this approach is demonstrated by comparing modal
Multiple piece turbine engine airfoil with a structural spar
Vance, Steven J.
2011-10-11
A multiple piece turbine airfoil having an outer shell with an airfoil tip that is attached to a root with an internal structural spar is disclosed. The root may be formed from first and second sections that include an internal cavity configured to receive and secure the one or more components forming the generally elongated airfoil. The internal structural spar may be attached to an airfoil tip and place the generally elongated airfoil in compression. The configuration enables each component to be formed from different materials to reduce the cost of the materials and to optimize the choice of material for each component.
Advanced technology airfoil research, volume 1, part 2
NASA Technical Reports Server (NTRS)
1978-01-01
This compilation contains papers presented at the NASA Conference on Advanced Technology Airfoil Research held at Langley Research Center on March 7-9, 1978, which have unlimited distribution. This conference provided a comprehensive review of all NASA airfoil research, conducted in-house and under grant and contract. A broad spectrum of airfoil research outside of NASA was also reviewed. The major thrust of the technical sessions were 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.
... become infected. Although most people are familiar with appendicitis, it is a relatively rare disease. It is ... of the appendix (appendectomy). Recovery time for uncomplicated appendicitis is usually just three days.
On the acoustic radiation of a pitching airfoil
NASA Astrophysics Data System (ADS)
Manela, A.
2013-07-01
We examine the acoustic far field of a thin elastic airfoil, immersed in low-Mach non-uniform stream flow, and actuated by small-amplitude sinusoidal pitching motion. The near-field fluid-structure interaction problem is analyzed using potential thin-airfoil theory, combined with a discrete vortex model to describe the evolution of airfoil trailing edge wake. The leading order dipole-sound signature of the system is investigated using Powell-Howe acoustic analogy. Compared with a pitching rigid airfoil, the results demonstrate a two-fold effect of structure elasticity on airfoil acoustic field: at actuation frequencies close to the system least stable eigenfrequency, elasticity amplifies airfoil motion amplitude and associated sound levels; however, at frequencies distant from this eigenfrequency, structure elasticity acts to absorb system kinetic energy and reduce acoustic radiation. In the latter case, and with increasing pitching frequency ωp, a rigid-airfoil setup becomes significantly noisier than an elastic airfoil, owing to an ω _p^{5/2} increase of its direct motion noise component. Unlike rigid airfoil signature, it is shown that wake sound contribution to elastic airfoil radiation is significant for all ωp. Remarkably, this contribution contains, in addition to the fundamental pitching frequency, its odd multiple harmonics, which result from nonlinear interactions between the airfoil and the wake. The results suggest that structure elasticity may serve as a viable means for design of flapping flight noise control methodologies.
Performance predictions of VAWTs with NLF airfoil blades
Masson, C.; Leclerc, C.; Paraschivoiu, I.
1997-02-01
The successful design of an efficient Vertical Axis Wind Turbine (VAWT) can be obtained only when appropriate airfoil sections have been selected. Most VAWTs currently operating worldwide use blades of symmetrical NACA airfoil series. As these blades were designed for aviation applications, Sandia National Laboratories developed a family of airfoils specifically designed for VAWTs in order to decrease the Cost of Energy (COE) of the VAWT (Berg, 1990). Objectives formulated for the blade profile were: modest values of maximum lift coefficient, low drag at low angle of attack, high drag at high angle of attack, sharp stall, and low thickness-to-chord ratio. These features are similar to those of Natural Laminar Flow airfoils (NLF) and gave birth to the SNLA airfoil series. This technical brief illustrates the benefits and losses resulting from using NLF airfoils on VAWT blades. To achieve this goal, the streamtube model of Paraschivoiu (1988) is used to predict the performance of VAWTs equipped with blades of various airfoil shapes. The airfoil shapes considered are the conventional airfoils NACA 0018 and NACA 0021, and the SNLA 0018/50 airfoil designed at Sandia. Furthermore, the potential benefit of reducing the airfoil drag is clearly illustrated by the presentation of the individual contributions of lift and drag to power.
User's manual for ADAM (Advanced Dynamic Airfoil Model)
Oler, J.W.; Strickland, J.H.; Im, B.J.
1987-06-01
The computer code for an advanced dynamic airfoil model (ADAM) is described. The code is capable of calculating steady or unsteady flow over two-dimensional airfoils with allowances for boundary layer separation. Specific types of airfoil motions currently installed are steady rectilinear motion, impulsively started rectilinear motion, constant rate pitching, sinusoidal pitch oscillations, sinusoidal lateral plunging, and simulated Darrieus turbine motion. Other types of airfoil motion may be analyzed through simple modifications of a single subroutine. The code has a built-in capability to generate the geometric parameters for a cylinder, the NACA four-digit series of airfoils, and a NASA NLF-0416 laminar airfoil. Other types of airfoils are easily incorporated. The code ADAM is currently in a state of development. It is theoretically consistent and complete. However, further work is needed on the numerical implementation of the method.
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.
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.
Exact solutions in oscillating airfoil theory
NASA Technical Reports Server (NTRS)
Williams, M. H.
1977-01-01
A result obtained by Williams (1977) for two-dimensional airfoils oscillating in an arbitrary subsonic parallel flowfield is reformulated to show that the pressure distribution induced by any deformation can be construed from the particular solutions for heaving and pitching motions. Specific formulas are presented for an oscillating control surface with a sealed gap.
Analysis of non-symmetrical flapping airfoils
NASA Astrophysics Data System (ADS)
Tay, W. B.; Lim, K. B.
2009-08-01
Simulations have been done to assess the lift, thrust and propulsive efficiency of different types of non-symmetrical airfoils under different flapping configurations. The variables involved are reduced frequency, Strouhal number, pitch amplitude and phase angle. In order to analyze the variables more efficiently, the design of experiments using the response surface methodology is applied. Results show that both the variables and shape of the airfoil have a profound effect on the lift, thrust, and efficiency. By using non-symmetrical airfoils, average lift coefficient as high as 2.23 can be obtained. The average thrust coefficient and efficiency also reach high values of 2.53 and 0.61, respectively. The lift production is highly dependent on the airfoil’s shape while thrust production is influenced more heavily by the variables. Efficiency falls somewhere in between. Two-factor interactions are found to exist among the variables. This shows that it is not sufficient to analyze each variable individually. Vorticity diagrams are analyzed to explain the results obtained. Overall, the S1020 airfoil is able to provide relatively good efficiency and at the same time generate high thrust and lift force. These results aid in the design of a better ornithopter’s wing.
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.
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 Technical Reports Server (NTRS)
Harris, C. D.
1975-01-01
The aerodynamic characteristics of several supercritical airfoils interim to the improved 10-percent thick NASA supercritical airfoil 26a are discussed. The airfoils have related slope and curvature distributions over the rear which result in different aft camber. For identification, the airfoils are designated supercritical airfoils 12, 13, 21, 22, and 24. Data is presented without analysis.
Airfoil Ice-Accretion Aerodynamics Simulation
NASA Technical Reports Server (NTRS)
Bragg, Michael B.; Broeren, Andy P.; Addy, Harold E.; Potapczuk, Mark G.; Guffond, Didier; Montreuil, E.
2007-01-01
NASA Glenn Research Center, ONERA, and the University of Illinois are conducting a major research program whose goal is to improve our understanding of the aerodynamic scaling of ice accretions on airfoils. The program when it is completed will result in validated scaled simulation methods that produce the essential aerodynamic features of the full-scale iced-airfoil. This research will provide some of the first, high-fidelity, full-scale, iced-airfoil aerodynamic data. An initial study classified ice accretions based on their aerodynamics into four types: roughness, streamwise ice, horn ice, and spanwise-ridge ice. Subscale testing using a NACA 23012 airfoil was performed in the NASA IRT and University of Illinois wind tunnel to better understand the aerodynamics of these ice types and to test various levels of ice simulation fidelity. These studies are briefly reviewed here and have been presented in more detail in other papers. Based on these results, full-scale testing at the ONERA F1 tunnel using cast ice shapes obtained from molds taken in the IRT will provide full-scale iced airfoil data from full-scale ice accretions. Using these data as a baseline, the final step is to validate the simulation methods in scale in the Illinois wind tunnel. Computational ice accretion methods including LEWICE and ONICE have been used to guide the experiments and are briefly described and results shown. When full-scale and simulation aerodynamic results are available, these data will be used to further develop computational tools. Thus the purpose of the paper is to present an overview of the program and key results to date.
Transonic airfoil analysis and design in nonuniform flow
NASA Technical Reports Server (NTRS)
Chang, J. F.; Lan, C. E.
1986-01-01
A nonuniform transonic airfoil code is developed for applications in analysis, inverse design and direct optimization involving an airfoil immersed in propfan slipstream. Problems concerning the numerical stability, convergence, divergence and solution oscillations are discussed. The code is validated by comparing with some known results in incompressible flow. A parametric investigation indicates that the airfoil lift-drag ratio can be increased by decreasing the thickness ratio. A better performance can be achieved if the airfoil is located below the slipstream center. Airfoil characteristics designed by the inverse method and a direct optimization are compared. The airfoil designed with the method of direct optimization exhibits better characteristics and achieves a gain of 22 percent in lift-drag ratio with a reduction of 4 percent in thickness.
Quiet airfoils for small and large wind turbines
Tangler, James L [Boulder, CO; Somers, Dan L [Port Matilda, PA
2012-06-12
Thick airfoil families with desirable aerodynamic performance with minimal airfoil induced noise. The airfoil families are suitable for a variety of wind turbine designs and are particularly well-suited for use with horizontal axis wind turbines (HAWTs) with constant or variable speed using pitch and/or stall control. In exemplary embodiments, a first family of three thick airfoils is provided for use with small wind turbines and second family of three thick airfoils is provided for use with very large machines, e.g., an airfoil defined for each of three blade radial stations or blade portions defined along the length of a blade. Each of the families is designed to provide a high maximum lift coefficient or high lift, to exhibit docile stalls, to be relatively insensitive to roughness, and to achieve a low profile drag.
A simplified method for the calculation of airfoil pressure distribution
NASA Technical Reports Server (NTRS)
Allen, H Julian
1939-01-01
A method is presented for the rapid calculation of the pressure distribution over an airfoil section when the normal-force distribution and the pressure distribution over the "base profile" (i.e., the profile of the same airfoil were the camber line straight and the resulting airfoil at zero angle of attack) are known. This note is intended as a supplement to N.A.C.A. Report Nos. 631 and 634 wherein methods are presented for the calculation of the normal-force distribution over plain and flapped airfoils, respectively, but not of the pressures on the individual surfaces. Base-profile pressure-coefficient distributions for the usual N.A.C.A. family of airfoils, which are also suitable for several other commonly employed airfoils, are included in tabular form. With these tabulated base-profile pressures and the computed normal-force distributions, pressure distributions adequate for most engineering purposes can be obtained.
Aerodynamic shape optimization of Airfoils in 2-D incompressible flow
NASA Astrophysics Data System (ADS)
Rangasamy, Srinivethan; Upadhyay, Harshal; Somasekaran, Sandeep; Raghunath, Sreekanth
2010-11-01
An optimization framework was developed for maximizing the region of 2-D airfoil immersed in laminar flow with enhanced aerodynamic performance. It uses genetic algorithm over a population of 125, across 1000 generations, to optimize the airfoil. On a stand-alone computer, a run takes about an hour to obtain a converged solution. The airfoil geometry was generated using two Bezier curves; one to represent the thickness and the other the camber of the airfoil. The airfoil profile was generated by adding and subtracting the thickness curve from the camber curve. The coefficient of lift and drag was computed using potential velocity distribution obtained from panel code, and boundary layer transition prediction code was used to predict the location of onset of transition. The objective function of a particular design is evaluated as the weighted-average of aerodynamic characteristics at various angles of attacks. Optimization was carried out for several objective functions and the airfoil designs obtained were analyzed.
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.
New airfoils for small horizontal axis wind turbines
Giguere, P.; Selig, M.S.
1997-12-31
In a continuing effort to enhance the performance of small energy systems, one root airfoil and three primary airfoils were specifically designed for small horizontal axis wind turbines. These airfoils are intended primarily for 1-10 kW variable-speed wind turbines for both conventional (tapered/twisted) or pultruded blades. The four airfoils were wind-tunnel tested at Reynolds numbers between 100,000 and 500,000. Tests with simulated leading-edge roughness were also conducted. The results indicate that small variable-speed wind turbines should benefit from the use of the new airfoils which provide enhanced lift-to-drag ratio performance as compared with previously existing airfoils.
Turbine airfoil with laterally extending snubber having internal cooling system
Scribner, Carmen Andrew; Messmann, Stephen John; Marsh, Jan H.
2016-09-06
A turbine airfoil usable in a turbine engine and having at least one snubber with a snubber cooling system positioned therein and in communication with an airfoil cooling system is disclosed. The snubber may extend from the outer housing of the airfoil toward an adjacent turbine airfoil positioned within a row of airfoils. The snubber cooling system may include an inner cooling channel separated from an outer cooling channel by an inner wall. The inner wall may include a plurality of impingement cooling orifices that direct impingement fluid against an outer wall defining the outer cooling channel. In one embodiment, the cooling fluids may be exhausted from the snubber, and in another embodiment, the cooling fluids may be returned to the airfoil cooling system. Flow guides may be positioned in the outer cooling channel, which may reduce cross-flow by the impingement orifices, thereby increasing effectiveness.
Separated transonic airfoil flow calculations with a nonequilibrium turbulence model
NASA Technical Reports Server (NTRS)
King, L. S.; Johnson, D. A.
1985-01-01
Navier-Stokes transonic airfoil calculations based on a recently developed nonequilibrium, turbulence closure model are presented for a supercritical airfoil section at transonic cruise conditions and for a conventional airfoil section at shock-induced stall conditions. Comparisons with experimental data are presented which show that this nonequilibrium closure model performs significantly better than the popular Baldwin-Lomax and Cebeci-Smith equilibrium algebraic models when there is boundary-layer separation that results from the inviscid-viscous interactions.
Unsteady airfoil flow solutions on moving zonal grids
NASA Technical Reports Server (NTRS)
Cricelli, Antonio S.; Ekaterinaris, John A.; Platzer, Max F.
1992-01-01
Euler and Navier-Stokes solutions for airfoil flows on zonal grids are presented. The governing equations are solved with an implicit, iterative, factorized numerical scheme. The inviscid fluxes are examined with a third-order accurate upwind method. Zonal grid solutions are compared with experimental measurements for flows over airfoils at fixed angles of incidence. The computed unsteady solutions for rapidly pitching and oscillating airfoils are in good agreement with experiments.
S825 and S826 Airfoils: 1994--1995
Somers, D. M.
2005-01-01
A family of airfoils, the S825 and S826, for 20- to 40-meter, variable-speed and variable-pitch (toward feather), horizontal-axis wind turbines has been designed and analyzed theoretically. The two primary objectives of high maximum lift, insensitive to roughness, and low profile drag have been achieved. The constraints on the pitching moments and the airfoil thicknesses have been satisfied. The airfoils should exhibit docile stalls.
Analysis of airfoil transitional separation bubbles
NASA Technical Reports Server (NTRS)
Davis, R. L.; Carter, J. E.
1984-01-01
A previously developed local inviscid-viscous interaction technique for the analysis of airfoil transitional separation bubbles, ALESEP (Airfoil Leading Edge Separation) has been modified to utilize a more accurate windward finite difference procedure in the reversed flow region, and a natural transition/turbulence model has been incorporated for the prediction of transition within the separation bubble. Numerous calculations and experimental comparisons are presented to demonstrate the effects of the windward differencing scheme and the natural transition/turbulence model. Grid sensitivity and convergence capabilities of this inviscid-viscous interaction technique are briefly addressed. A major contribution of this report is that with the use of windward differencing, a second, counter-rotating eddy has been found to exist in the wall layer of the primary separation bubble.
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.
Turbine engine airfoil and platform assembly
Campbell, Christian X [Oviedo, FL; James, Allister W [Chuluota, FL; Morrison, Jay A [Oviedo, FL
2012-07-31
A turbine airfoil (22A) is formed by a first process using a first material. A platform (30A) is formed by a second process using a second material that may be different from the first material. The platform (30A) is assembled around a shank (23A) of the airfoil. One or more pins (36A) extend from the platform into holes (28) in the shank (23A). The platform may be formed in two portions (32A, 34A) and placed around the shank, enclosing it. The two platform portions may be bonded to each other. Alternately, the platform (30B) may be cast around the shank (23B) using a metal alloy with better castability than that of the blade and shank, which may be specialized for thermal tolerance. The pins (36A-36D) or holes for them do not extend to an outer surface (31) of the platform, avoiding stress concentrations.
Turbomachinery Airfoil Design Optimization Using Differential Evolution
NASA Technical Reports Server (NTRS)
Madavan, Nateri K.; Biegel, Bryan (Technical Monitor)
2002-01-01
An aerodynamic design optimization procedure that is based on a evolutionary algorithm known at Differential Evolution is described. Differential Evolution is a simple, fast, and robust evolutionary strategy that has been proven effective in determining the global optimum for several difficult optimization problems, including highly nonlinear systems with discontinuities and multiple local optima. The method is combined with a Navier-Stokes solver that evaluates the various intermediate designs and provides inputs to the optimization procedure. An efficient constraint handling mechanism is also incorporated. Results are presented for the inverse design of a turbine airfoil from a modern jet engine and compared to earlier methods. The capability of the method to search large design spaces and obtain the optimal airfoils in an automatic fashion is demonstrated. Substantial reductions in the overall computing time requirements are achieved by using the algorithm in conjunction with neural networks.
Turbomachinery Airfoil Design Optimization Using Differential Evolution
NASA Technical Reports Server (NTRS)
Madavan, Nateri K.; Biegel, Bryan A. (Technical Monitor)
2002-01-01
An aerodynamic design optimization procedure that is based on a evolutionary algorithm known at Differential Evolution is described. Differential Evolution is a simple, fast, and robust evolutionary strategy that has been proven effective in determining the global optimum for several difficult optimization problems, including highly nonlinear systems with discontinuities and multiple local optima. The method is combined with a Navier-Stokes solver that evaluates the various intermediate designs and provides inputs to the optimization procedure. An efficient constraint handling mechanism is also incorporated. Results are presented for the inverse design of a turbine airfoil from a modern jet engine. The capability of the method to search large design spaces and obtain the optimal airfoils in an automatic fashion is demonstrated. Substantial reductions in the overall computing time requirements are achieved by using the algorithm in conjunction with neural networks.
Aerodynamic properties of thick airfoils II
NASA Technical Reports Server (NTRS)
Norton, F H; Bacon, D L
1923-01-01
This investigation is an extension of NACA report no. 75 for the purpose of studying the effect of various modifications in a given wing section, including changes in thickness, height of lower camber, taper in thickness, and taper in plan form with special reference to the development of thick, efficient airfoils. The method consisted in testing the wings in the NACA 5-foot wind tunnel at speeds up to 50 meters (164 feet) per second while they were being supported on a new type of wire balance. Some of the airfoils developed showed results of great promise. For example, one wing (no. 81) with a thickness in the center of 4.5 times that of the U. S. A. 16 showed both uniformly high efficiency and a higher maximum lift than this excellent section. These thick sections will be especially useful on airplanes with cantilever construction. (author)
Low Reynolds number airfoil survey, volume 1
NASA Technical Reports Server (NTRS)
Carmichael, B. H.
1981-01-01
The differences in flow behavior two dimensional airfoils in the critical chordlength Reynolds number compared with lower and higher Reynolds number are discussed. The large laminar separation bubble is discussed in view of its important influence on critical Reynolds number airfoil behavior. The shortcomings of application of theoretical boundary layer computations which are successful at higher Reynolds numbers to the critical regime are discussed. The large variation in experimental aerodynamic characteristic measurement due to small changes in ambient turbulence, vibration, and sound level is illustrated. The difficulties in obtaining accurate detailed measurements in free flight and dramatic performance improvements at critical Reynolds number, achieved with various types of boundary layer tripping devices are discussed.
2014-12-31
YOUR FORM TO THE ABOVE ADDRESS. Ohio State University 1960 Kenny Road Columbus , OH 43210 -1016 30-Sep-2014 ABSTRACT Number of Papers published in peer...Jensen, Christopher D., Gregory, James W., Gompertz, Kyle A., Bons, Jeffrey P.. Compressible Dynamic Stall Mechanisms Due to Airfoil Pitching and...Characteristics of a NACA 0015 Airfoil,” Data Report, Aeronautical and Astronautical Research Laboratory, Ohio State University, Columbus , Ohio. Hird, K
Damping element for reducing the vibration of an airfoil
Campbell, Christian X; Marra, John J
2013-11-12
An airfoil (10) is provided with a tip (12) having an opening (14) to a center channel (24). A damping element (16) is inserted within the opening of the center channel, to reduce an induced vibration of the airfoil. The mass of the damping element, a spring constant of the damping element within the center channel, and/or a mounting location (58) of the damping element within the center channel may be adjustably varied, to shift a resonance frequency of the airfoil outside a natural operating frequency of the airfoil.
An analytical study for the design of advanced rotor airfoils
NASA Technical Reports Server (NTRS)
Kemp, L. D.
1973-01-01
A theoretical study has been conducted to design and evaluate two airfoils for helicopter rotors. The best basic shape, designed with a transonic hodograph design method, was modified to meet subsonic criteria. One airfoil had an additional constraint for low pitching-moment at the transonic design point. Airfoil characteristics were predicted. Results of a comparative analysis of helicopter performance indicate that the new airfoils will produce reduced rotor power requirements compared to the NACA 0012. The hodograph design method, written in CDC Algol, is listed and described.
Investigation of the Boundary Layer Behavior on Turbine Airfoils.
1979-08-01
turbine airfoil cascade . The airfoil profile was based on a turbine blade design used by Lander ’’4 and employed in previous wake studies by Cox and...simulate the wake from upstream turning vanes or blades , a circular cylinder was placed upstream of the centra l or test airfoil . The displacement of this...of turbine airfoil cascade model s by Cox and Han 15 are very much evident in the graph . It might be noted that the blade stag- nation points are at
Comparisons of Theoretical Methods for Predicting Airfoil Aerodynamic Characteristics
2010-08-01
Airfoil ,” Airfoils , U.S. Army Aviation Research, Development and Engineering Command, RDECOM TR 10-D-107, August 2010. [2] Somers, D.M. and...Maughmer, M.D., “Design and Experimental Results for the S407 Airfoil ,” U.S. Army Aviation Research, Development and Engineering Command, RDECOM TR 10-D...S414 Airfoil ,” U.S. Army Aviation Research, Development and Engineering Command, RDECOM TR 10-D-112, August 2010. [5] Somers, D.M. and Maughmer
Transonic airfoil analysis and design using Cartesian coordinates
NASA Technical Reports Server (NTRS)
Carlson, L. A.
1975-01-01
An inverse numerical technique for designing transonic airfoils having a prescribed pressure distribution is presented. The method uses the full potential equation, inverse boundary conditions, and Cartesian coordinates. It includes simultaneous airfoil update and utilizes a direct-inverse approach that permits a logical method for controlling trailing edge closure. The method can also be used for the analysis of flowfields about specified airfoils. Comparison with previous results shows that accurate results can be obtained with a Cartesian grid. Examples show the application of the method to design aft-cambered and other airfoils specifically for transonic flight.
Experiments on airfoils with trailing edge cut away
NASA Technical Reports Server (NTRS)
Ackeret, J
1927-01-01
Airfoils with their trailing edge cut away are often found on aircraft, as the fins on the hulls of flying boats and the central section of the wings for affording better visibility. It was therefore of some interest to discover the effect of such cutaways on the lift and drag and on the position of the center of pressure. For this purpose, systematic experiments were performed on two different airfoils, a symmetrical airfoil and an airfoil of medium thickness, with successive shortenings of their chords.
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.
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)
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)
Transonic airfoil and axial flow rotary machine
Nagai, Naonori; Iwatani, Junji
2015-09-01
Sectional profiles close to a tip 124 and a part between a midportion 125 and a hub 123 are shifted to the upstream of an operating fluid flow in a sweep direction. Accordingly, an S shape is formed in which the tip 124 and the part between the midportion 125 and the hub 123 protrude. As a result, it is possible reduce various losses due to shook, waves, thereby forming a transonic airfoil having an excellent aerodynamic characteristic.
Turbulent Flow over Rough Turbine Airfoils.
1985-08-01
SUBJECT TERMS (Continue on reverse if necessary and identify by block number) FIELD GROUP SUB. GR. Turbine blades ’ vanes ; surface roughness...turbulent boundary layer over rough turbine vanes or blades is developed. A new formulation of the mixing length model, expressed in the velocity-space...A-163 005 TURBULENT FLOW OVER ROUGH TURBINE AIRFOILS (U) OHIO 1/ STATE UNIV RESEARCH FOUNDATION COLUMBUS L S HAN AUG B5 OSURF-76357/?i4467 AFWL-TR-95
Linearized propulsion theory of flapping airfoils revisited
NASA Astrophysics Data System (ADS)
Fernandez-Feria, R.
2016-12-01
A vortical impulse theory is used to compute the thrust force of a plunging and pitching airfoil in forward flight at high Reynolds numbers within the framework of linear potential flow theory. The result is significantly different from the classical one of Garrick, which considered only two effects, the leading-edge suction and the projection in the flight direction of the pressure force on the airfoil. By taking into account the complete vorticity distribution on the airfoil and the wake the mean thrust coefficient contains, in addition to the pressure force projection term, a new term that generalizes the leading-edge suction term in Garrick's theory. This term depends on Theodorsen function C (k ) and on a new complex function C1(k ) of the reduced frequency k . The main qualitative difference with Garrick's theory is that the propulsive efficiency, or ratio of the mean thrust power and the mean input power required to drive the airfoil, tends to zero as the reduced frequency increases to infinity (as k-1), in contrast to Garrick's propulsive efficiency that tends to a constant (1 /2 ). Consequently, for pure pitching and combined pitching and plunging motions, the maximum of the propulsive efficiency is not reached as k →∞ like in Garrick's theory, but at a finite value of the reduced frequency that depends on the remaining nondimensional parameters. The present analytical results are in good agreement, for small amplitude oscillations, with numerical results from unsteady panel methods, and with experimental data and numerical results from the Navier-Stokes equations, except for small reduced frequencies where viscous effects are obviously important.
Tail Rotor Airfoils Stabilize Helicopters, Reduce Noise
NASA Technical Reports Server (NTRS)
2010-01-01
Founded by former Ames Research Center engineer Jim Van Horn, Van Horn Aviation of Tempe, Arizona, built upon a Langley Research Center airfoil design to create a high performance aftermarket tail rotor for the popular Bell 206 helicopter. The highly durable rotor has a lifetime twice that of the original equipment manufacturer blade, reduces noise by 40 percent, and displays enhanced performance at high altitudes. These improvements benefit helicopter performance for law enforcement, military training, wildfire and pipeline patrols, and emergency medical services.
Streamwise Oscillation of Airfoils into Reverse Flow
NASA Astrophysics Data System (ADS)
Granlund, Kenneth; Jones, Anya; Ol, Michael
2015-11-01
An airfoil in freestream is oscillated in streamwise direction to cyclically enter reverse flow. Measured lift is compared to analytical blade element theories. Advance ratio, reduced frequency and angle of attack is varied within those typical for helicopters. Experimental results reveal that lift does not become negative in the flow reversal part, contradicting one theory and supported by another. Flow visualization reveal the leading edge vortex advecting against the freestream to a point in front of the leading edge.
2010-08-01
airfoils where the drag coefficient within the laminar bucket is nearly constant. (See, for example, ref. 4.) This characteristic is related to the...edge with increasing (decreasing) lift coefficient. This feature results in a leading-edge shape that produces a suction peak at higher lift...should look like sketch 3. Sketch 3 1Director, Institute for Aerodynamics and Gas Dynamics, University of Stuttgart, Germany, 1974–1985.5 No suction
An Experimental Study of Airfoil Icing Characteristics
NASA Technical Reports Server (NTRS)
Shaw, R. J.; Sotos, R. G.; Solano, F. R.
1982-01-01
A full scale general aviation wing with a NACA 63 sub 2 A415 airfoil section was tested to determine icing characteristics for representative rime and glaze icing conditions. Measurements were made of ice accretion shapes and resultant wing section drag coefficient levels. It was found that the NACA 63 sub 2 A415 wing section was less sensitive to rime and glaze icing encounters for climb conditions.
Wake structure of a deformable Joukowski airfoil
NASA Astrophysics Data System (ADS)
Ysasi, Adam; Kanso, Eva; Newton, Paul K.
2011-10-01
We examine the vortical wake structure shed from a deformable Joukowski airfoil in an unbounded volume of inviscid and incompressible fluid. The deformable airfoil is considered to model a flapping fish. The vortex shedding is accounted for using an unsteady point vortex model commonly referred to as the Brown-Michael model. The airfoil’s deformations and rotations are prescribed in terms of a Jacobi elliptic function which exhibits, depending on a dimensionless parameter m, a range of periodic behaviors from sinusoidal to a more impulsive type flapping. Depending on the parameter m and the Strouhal number, one can identify five distinct wake structures, ranging from arrays of isolated point vortices to vortex dipoles and tripoles shed into the wake with every half-cycle of the airfoil flapping motion. We describe these regimes in the context of other published works which categorize wake topologies, and speculate on the importance of these wake structures in terms of periodic swimming and transient maneuvers of fish.
Ordered roughness effects on NACA 0026 airfoil
NASA Astrophysics Data System (ADS)
Harun, Z.; Abbas, A. A.; Dheyaa, R. Mohammed; Ghazali, M. I.
2016-10-01
The effects of highly-ordered rough surface - riblets, applied onto the surface of a NACA 0026 airfoil, are investigated experimentally using wind tunnel. The riblets are arranged in directionally converging - diverging pattern with dimensions of height, h = 1 mm, pitch or spacing, s = 1 mm, yaw angle α = 0o and 10o The airfoil with external geometry of 500 mm span, 600 mm chord and 156 mm thickness has been built using mostly woods and aluminium. Turbulence quantities are collected using hotwire anemometry. Hotwire measurements show that flows past converging and diverging pattern inherit similar patterns in the near-wall region for both mean velocity and turbulence intensities profiles. The mean velocity profiles in logarithmic regions for both flows past converging and diverging riblet pattern are lower than that with yaw angle α = 0o. Converging riblets cause the boundary layer to thicken and the flow with yaw angle α = 0o produces the thinnest boundary layer. Both the converging and diverging riblets cause pronounced outer peaks in the turbulence intensities profiles. Most importantly, flows past converging and diverging pattern experience 30% skin friction reductions. Higher order statistics show that riblet surfaces produce similar effects due to adverse pressure gradient. It is concluded that a small strip of different ordered roughness features applied at a leading edge of an airfoil can change the turbulence characteristics dramatically.
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.
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.
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.
47 CFR Appendix - Technical Appendix 2
Code of Federal Regulations, 2010 CFR
2010-10-01
... 47 Telecommunication 5 2010-10-01 2010-10-01 false Technical Appendix 2 Telecommunication NATIONAL TELECOMMUNICATIONS AND INFORMATION ADMINISTRATION, DEPARTMENT OF COMMERCE DIGITAL-TO-ANALOG CONVERTER BOX COUPON PROGRAM Waiver of household eligibility. Pt. 301, App. 2 Technical Appendix 2 TECHNICAL APPENDIX 2—NTIA...
Sealing apparatus for airfoils of gas turbine engines
Jones, R.B.
1998-05-19
An improved airfoil tip sealing apparatus is disclosed wherein brush seals are attached to airfoil tips with the distal ends of the brush seal fibers sealingly contacting opposing wall surfaces. Embodiments for variable vanes, stators and both cooled and uncooled turbine blade applications are disclosed. 17 figs.
Sealing apparatus for airfoils of gas turbine engines
Jones, Russell B.
1998-01-01
An improved airfoil tip sealing apparatus is disclosed wherein brush seals are attached to airfoil tips with the distal ends of the brush seal fibers sealingly contacting opposing wall surfaces. Embodiments for variable vanes, stators and both cooled and uncooled turbine blade applications are disclosed.
Design procedure for low-drag subsonic airfoils
NASA Technical Reports Server (NTRS)
Peterson, J. B.; Chen, A. B.
1975-01-01
Airfoil has least amount of drag under given restrictions of boundary layer transition position, lift coefficient, thickness ratio, and Reynolds number based on airfoil chord. It is suitable for use as wing and propeller aircraft sections operating at subsonic speeds and for hydrofoil sections and blades for fans, compressors, turbines, and windmills.
Effects of Compressibility on the Flow Past Thick Airfoil Sections
1948-07-01
Tests of Airfoils Designed to Delay the Compressibility Burble . NACA Rep. No. 763, 19^3• 3. Stack, John, Lindsey, W. F., and Littell, Robert E.: The...Compressi- bility Burble and the Effect of Compressibility on Pressures and Forces Acting on an Airfoil. NACA Rep. No. 6k6, 1938. k. Byrne, Robert
Numerical Airfoil Optimization Using a Reduced Number of Design Coordinates
NASA Technical Reports Server (NTRS)
Vanderplaats, G. N.; Hicks, R. M.
1976-01-01
A method is presented for numerical airfoil optimization whereby a reduced number of design coordinates are used to define the airfoil shape. The approach is to define the airfoil as a linear combination of shapes. These basic shapes may be analytically or numerically defined, allowing the designer to use his insight to propose candidate designs. The design problem becomes one of determining the participation of each such function in defining the optimum airfoil. Examples are presented for two-dimensional airfoil design and are compared with previous results based on a polynomial representation of the airfoil shape. Four existing NACA airfoils are used as basic shapes. Solutions equivalent to previous results are achieved with a factor of more than 3 improvements in efficiency, while superior designs are demonstrated with an efficiency greater than 2 over previous methods. With this shape definition, the optimization process is shown to exploit the simplifying assumptions in the inviscid aerodynamic analysis used here, thus demonstrating the need to use more advanced aerodynamics for airfoil optimization.
TRANDES: A FORTRAN program for transonic airfoil analysis or design
NASA Technical Reports Server (NTRS)
Carlson, L. A.
1977-01-01
A program called TRANDES is presented that is used for the analysis of steady, irrotational transonic flow over specified two-dimensional airfoils in free air or for the design of airfoils having a prescribed pressure distribution, including the effects of weak viscous interaction. Instructions on program usage, listings of the program, and sample cases are given.
Development of drive mechanism for an oscillating airfoil
NASA Technical Reports Server (NTRS)
Sticht, Clifford D.
1988-01-01
The design and development of an in-draft wind tunnel test section which will be used to study the dynamic stall of airfoils oscillating in pitch is described. The hardware developed comprises a spanned airfoil between schleiren windows, a four bar linkage, flywheels, a drive system and a test section structure.
The compressibility rule for drag of airfoil noses
NASA Technical Reports Server (NTRS)
Jones, R. T.; Vandyke, M. D.
1976-01-01
It is shown that the drag of any semi-infinite airfoil section in purely subsonic inviscid flow follows precisely the Prandtl-Glauert compressibility rule. The result for the parabola has application to leading edge corrections in thin airfoil theory.
Numerical computation of aeroacoustic transfer functions for realistic airfoils
NASA Astrophysics Data System (ADS)
Miotto, Renato Fuzaro; Wolf, William Roberto; de Santana, Leandro Dantas
2017-10-01
Based on Amiet's theory formalism, we propose a numerical framework to compute the aeroacoustic transfer function of realistic airfoil geometries. The aeroacoustic transfer function relates the amplitude and phase of an incoming periodic gust to the respective unsteady lift response permitting, therefore, the application of Curle's analogy to compute the radiated noise. The methodology is focused on the airfoil leading-edge noise problem being able to also consider the trailing-edge back-scattering and, consequently, airfoil compactness effects. The approach is valid for compressible subsonic flows and the airfoil blade is assumed of large aspect ratio subjected to three-dimensional periodic gusts with supersonic velocity trace at the airfoil leading edge (i.e. supercritical gusts). This work proposes the iterative application of the boundary element method to numerically solve the boundary value problem prescribed by the linearized airfoil theory. Details of the numerical implementation are discussed and include the application of boundary conditions in different steps of the iterative procedure, treatment of derivatives in the implementation of the Kutta condition and accurate representation of singularities present at the leading- and trailing-edges. This study validates the numerical approach by comparing results with Amiet's theory obtained analytically. Subsequently, effects of realistic airfoil geometries on the leading-edge airfoil radiated noise are presented.
The Aerodynamic Characteristics of Airfoils as Affected by Surface Roughness
NASA Technical Reports Server (NTRS)
HOCKER RAY W
1933-01-01
The effect on airfoil characteristics of surface roughness of varying degrees and types at different locations on an airfoil was investigated at high values of the Reynolds number in a variable density wind tunnel. Tests were made on a number of National Advisory Committee for Aeronautics (NACA) 0012 airfoil models on which the nature of the surface was varied from a rough to a very smooth finish. The effect on the airfoil characteristics of varying the location of a rough area in the region of the leading edge was also investigated. Airfoils with surfaces simulating lap joints were also tested. Measurable adverse effects were found to be caused by small irregularities in airfoil surfaces which might ordinarily be overlooked. The flow is sensitive to small irregularities of approximately 0.0002c in depth near the leading edge. The tests made on the surfaces simulating lap joints indicated that such surfaces cause small adverse effects. Additional data from earlier tests of another symmetrical airfoil are also included to indicate the variation of the maximum lift coefficient with the Reynolds number for an airfoil with a polished surface and with a very rough one.
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.
Airfoil family design for large offshore wind turbine blades
NASA Astrophysics Data System (ADS)
Méndez, B.; Munduate, X.; San Miguel, U.
2014-06-01
Wind turbine blades size has scaled-up during last years due to wind turbine platform increase especially for offshore applications. The EOLIA project 2007-2010 (Spanish Goverment funded project) was focused on the design of large offshore wind turbines for deep waters. The project was managed by ACCIONA Energia and the wind turbine technology was designed by ACCIONA Windpower. The project included the design of a wind turbine airfoil family especially conceived for large offshore wind turbine blades, in the order of 5MW machine. Large offshore wind turbines suffer high extreme loads due to their size, in addition the lack of noise restrictions allow higher tip speeds. Consequently, the airfoils presented in this work are designed for high Reynolds numbers with the main goal of reducing blade loads and mantainig power production. The new airfoil family was designed in collaboration with CENER (Spanish National Renewable Energy Centre). The airfoil family was designed using a evolutionary algorithm based optimization tool with different objectives, both aerodynamic and structural, coupled with an airfoil geometry generation tool. Force coefficients of the designed airfoil were obtained using the panel code XFOIL in which the boundary layer/inviscid flow coupling is ineracted via surface transpiration model. The desing methodology includes a novel technique to define the objective functions based on normalizing the functions using weight parameters created from data of airfoils used as reference. Four airfoils have been designed, here three of them will be presented, with relative thickness of 18%, 21%, 25%, which have been verified with the in-house CFD code, Wind Multi Block WMB, and later validated with wind tunnel experiments. Some of the objectives for the designed airfoils concern the aerodynamic behavior (high efficiency and lift, high tangential coefficient, insensitivity to rough conditions, etc.), others concern the geometry (good for structural design
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.
Multiple element airfoils optimized for maximum lift coefficient.
NASA Technical Reports Server (NTRS)
Ormsbee, A. I.; Chen, A. W.
1972-01-01
Optimum airfoils in the sense of maximum lift coefficient are obtained for incompressible fluid flow at large Reynolds number. The maximum lift coefficient is achieved by requiring that the turbulent skin friction be zero in the pressure rise region on the airfoil upper surface. Under this constraint, the pressure distribution is optimized. The optimum pressure distribution is a function of Reynolds number and the trailing edge velocity. Geometries of those airfoils which will generate these optimum pressure distributions are obtained using a direct-iterative method which is developed in this study. This method can be used to design airfoils consisting of any number of elements. Numerical examples of one- and two-element airfoils are given. The maximum lift coefficients obtained range from 2 to 2.5.
Reversible airfoils for stopped rotors in high speed flight
NASA Astrophysics Data System (ADS)
Niemiec, Robert; Jacobellis, George; Gandhi, Farhan
2014-10-01
This study starts with the design of a reversible airfoil rib for stopped-rotor applications, where the sharp trailing-edge morphs into the rounded leading-edge, and vice-versa. A NACA0012 airfoil is approximated in a piecewise linear manner and straight, rigid outer profile links used to define the airfoil contour. The end points of the profile links connect to control links, each set on a central actuation rod via an offset. Chordwise motion of the actuation rod moves the control and the profile links and reverses the airfoil. The paper describes the design methodology and evolution of the final design, based on which two reversible airfoil ribs were fabricated and used to assemble a finite span reversible rotor/wing demonstrator. The profile links were connected by Aluminum strips running in the spanwise direction which provided stiffness as well as support for a pre-tensioned elastomeric skin. An inter-rib connector with a curved-front nose piece supports the leading-edge. The model functioned well and was able to reverse smoothly back-and-forth, on application and reversal of a voltage to the motor. Navier-Stokes CFD simulations (using the TURNS code) show that the drag coefficient of the reversible airfoil (which had a 13% maximum thickness due to the thickness of the profile links) was comparable to that of the NACA0013 airfoil. The drag of a 16% thick elliptical airfoil was, on average, about twice as large, while that of a NACA0012 in reverse flow was 4-5 times as large, even prior to stall. The maximum lift coefficient of the reversible airfoil was lower than the elliptical airfoil, but higher than the NACA0012 in reverse flow operation.
Wind tunnel test of the S814 thick root airfoil
Somers, D.M.; Tangler, J.L.
1996-11-01
The objective of this wind-tunnel test was to verify the predictions of the Eppler Airfoil Design and Analysis Code for a very thick airfoil having a high maximum lift coefficient designed to be largely insensitive to leading-edge roughness effects. The 24 percent thick S814 airfoil was designed with these characteristics to accommodate aerodynamic and structural considerations for the root region of a wind-turbine blade. In addition, the airfoil`s maximum lift-to-drag ratio was designed to occur at a high lift coefficient. To accomplish the objective, a two-dimensional wind tunnel test of the S814 thick root airfoil was conducted in January 1994 in the low-turbulence wind tunnel of the Delft University of Technology Low Speed Laboratory, The Netherlands. Data were obtained with transition free and transition fixed for Reynolds numbers of 0.7, 1.0, 1.5, 2.0, and 3.0 {times} 10{sup 6}. For the design Reynolds number of 1.5 {times} 10{sup 6}, the maximum lift coefficient with transition free is 1.32, which satisfies the design specification. However, this value is significantly lower than the predicted maximum lift coefficient of almost 1.6. With transition fixed at the leading edge, the maximum lift coefficient is 1.22. The small difference in maximum lift coefficient between the transition-free and transition-fixed conditions demonstrates the airfoil`s minimal sensitivity to roughness effects. The S814 root airfoil was designed to complement existing NREL low maximum-lift-coefficient tip-region airfoils for rotor blades 10 to 15 meters in length.
Aerodynamic tests of Darrieus wind turbine blades
Migliore, P.G.; Walters, R.E.; Wolfe, W.P.
1983-03-01
An indoor facility for the aerodynamic testing of Darrieus turbine blades was developed. Lift, drag, and moment coefficients were measured for two blades whose angle of attack and chord-to-radius ratio were varied. The first blade used an NACA 0015 airfoil section; the second used a 15% elliptical cross section with a modified circular arc trailing edge. Blade aerodynamic coefficients were corrected to section coefficients for comparison to published rectilinear flow data. Although the airfoil sections were symmetrical, moment coefficients were not zero and the lift and drag curves were asymmetrical about zero lift coefficient and angle of attack. These features verified the predicted virtual camber and incidence phenomena. Boundary-layer centrifugal effects were manifested by discontinuous lift curves and large differences in the angle of zero lift between th NACA 0015 and elliptical airfoils. It was concluded that rectilinear flow aerodynamic data are not applicable to Darrieus turbine blades, even for small chord-to-radius ratios.
Turbine airfoil having outboard and inboard sections
Mazzola, Stefan; Marra, John J
2015-03-17
A turbine airfoil usable in a turbine engine and formed from at least an outboard section and an inboard section such that an inner end of the outboard section is attached to an outer end of the inboard section. The outboard section may be configured to provide a tip having adequate thickness and may extend radially inward from the tip with a generally constant cross-sectional area. The inboard section may be configured with a tapered cross-sectional area to support the outboard section.
Approximation concepts for numerical airfoil optimization
NASA Technical Reports Server (NTRS)
Vanderplaats, G. N.
1979-01-01
An efficient algorithm for airfoil optimization is presented. The algorithm utilizes approximation concepts to reduce the number of aerodynamic analyses required to reach the optimum design. Examples are presented and compared with previous results. Optimization efficiency improvements of more than a factor of 2 are demonstrated. Improvements in efficiency are demonstrated when analysis data obtained in previous designs are utilized. The method is a general optimization procedure and is not limited to this application. The method is intended for application to a wide range of engineering design problems.
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.
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.
The modelling of symmetric airfoil vortex generators
NASA Technical Reports Server (NTRS)
Reichert, B. A.; Wendt, B. J.
1996-01-01
An experimental study is conducted to determine the dependence of vortex generator geometry and impinging flow conditions on shed vortex circulation and crossplane peak vorticity for one type of vortex generator. The vortex generator is a symmetric airfoil having a NACA 0012 cross-sectional profile. The geometry and flow parameters varied include angle-of-attack alfa, chordlength c, span h, and Mach number M. The vortex generators are mounted either in isolation or in a symmetric counter-rotating array configuration on the inside surface of a straight pipe. The turbulent boundary layer thickness to pipe radius ratio is delta/R = 0. 17. Circulation and peak vorticity data are derived from crossplane velocity measurements conducted at or about 1 chord downstream of the vortex generator trailing edge. Shed vortex circulation is observed to be proportional to M, alfa, and h/delta. With these parameters held constant, circulation is observed to fall off in monotonic fashion with increasing airfoil aspect ratio AR. Shed vortex peak vorticity is also observed to be proportional to M, alfa, and h/delta. Unlike circulation, however, peak vorticity is observed to increase with increasing aspect ratio, reaching a peak value at AR approx. 2.0 before falling off.
FLEET Velocimetry Measurements on a Transonic Airfoil
NASA Technical Reports Server (NTRS)
Burns, Ross A.; Danehy, Paul M.
2017-01-01
Femtosecond laser electronic excitation tagging (FLEET) velocimetry was used to study the flowfield around a symmetric, transonic airfoil in the NASA Langley 0.3-m TCT facility. A nominal Mach number of 0.85 was investigated with a total pressure of 125 kPa and total temperature of 280 K. Two-components of velocity were measured along vertical profiles at different locations above, below, and aft of the airfoil at angles of attack of 0 deg, 3.5 deg, and 7deg. Measurements were assessed for their accuracy, precision, dynamic range, spatial resolution, and overall measurement uncertainty in the context of the applied flowfield. Measurement precisions as low as 1 m/s were observed, while overall uncertainties ranged from 4 to 5 percent. Velocity profiles within the wake showed sufficient accuracy, precision, and sensitivity to resolve both the mean and fluctuating velocities and general flow physics such as shear layer growth. Evidence of flow separation is found at high angles of attack.
Linearized propulsion theory of flapping airfoils revisited
NASA Astrophysics Data System (ADS)
Fernandez-Feria, Ramon
2016-11-01
A vortical impulse theory is used to compute the thrust of a plunging and pitching airfoil in forward flight within the framework of linear potential flow theory. The result is significantly different from the classical one of Garrick that considered the leading-edge suction and the projection in the flight direction of the pressure force. By taking into account the complete vorticity distribution on the airfoil and the wake the mean thrust coefficient contains a new term that generalizes the leading-edge suction term and depends on Theodorsen function C (k) and on a new complex function C1 (k) of the reduced frequency k. The main qualitative difference with Garrick's theory is that the propulsive efficiency tends to zero as the reduced frequency increases to infinity (as 1 / k), in contrast to Garrick's efficiency that tends to a constant (1 / 2). Consequently, for pure pitching and combined pitching and plunging motions, the maximum of the propulsive efficiency is not reached as k -> ∞ like in Garrick's theory, but at a finite value of the reduced frequency that depends on the remaining non-dimensional parameters. The present analytical results are in good agreement with experimental data and numerical results for small amplitude oscillations. Supported by the Ministerio de Economia y Competitividad of Spain Grant No. DPI2013-40479-P.
Inverse design of airfoils using a flexible membrane method
NASA Astrophysics Data System (ADS)
Thinsurat, Kamon
The Modified Garabedian Mc-Fadden (MGM) method is used to inversely design airfoils. The Finite Difference Method (FDM) for Non-Uniform Grids was developed to discretize the MGM equation for numerical solving. The Finite Difference Method (FDM) for Non-Uniform Grids has the advantage of being used flexibly with an unstructured grids airfoil. The commercial software FLUENT is being used as the flow solver. Several conditions are set in FLUENT such as subsonic inviscid flow, subsonic viscous flow, transonic inviscid flow, and transonic viscous flow to test the inverse design code for each condition. A moving grid program is used to create a mesh for new airfoils prior to importing meshes into FLUENT for the analysis of flows. For validation, an iterative process is used so the Cp distribution of the initial airfoil, the NACA0011, achieves the Cp distribution of the target airfoil, the NACA2315, for the subsonic inviscid case at M=0.2. Three other cases were carried out to validate the code. After the code validations, the inverse design method was used to design a shock free airfoil in the transonic condition and to design a separation free airfoil at a high angle of attack in the subsonic condition.
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.
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.
Efficiency of an auto-propelled flapping airfoil
NASA Astrophysics Data System (ADS)
Benkherouf, T.; Mekadem, M.; Oualli, H.; Hanchi, S.; Keirsbulck, L.; Labraga, L.
2011-05-01
The present study deals with an investigation of the flow aerodynamic characteristics and the propulsive velocity of a system equipped with a nature inspired propulsion system. In particular, the study is aimed at studying the effect of the flapping frequency on the flow behavior. We consider a NACA0014 airfoil undergoing a vertical sinusoidal flapping motion. In contrast to nearly all previous studies in the literature, the present work does not impose any velocity on the inlet flow. During each iteration the outer flow velocity is computed after having determined the forces exerted on the airfoil. Forward motion may only be produced by flapping motion of the airfoil. This is more consistent with the physical phenomenon. The non-stationary viscous flow around the flapping airfoil is simulated using Ansys-Fluent 12.0.7. The airfoil movement is achieved using the deformable mesh technique and an in-house developed User Define Function (UDF). Our results show the influence of flapping frequency and amplitude on both the airfoil velocity and the propulsive efficiency. The resulting motion is contrasts to the applied forces. In the present study, the frequency ranges from 0.1 to 20 Hz while the airfoil amplitude values considered are: 10%, 17.5%, 25% and 40%.
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.
Laser velocimetry measurements of oscillating airfoil dynamic stall flow field
NASA Technical Reports Server (NTRS)
Chandrasekhara, M. S.; Ahmed, S.
1991-01-01
Ensemble-averaged two-component velocity measurements over an airfoil experiencing oscillatory dynamic stall under compressibility conditions were obtained. The measurements show the formation of a separation bubble over the airfoil that persists till angles of attack close to when the dynamic stall vortex forms and convects. The fluid attains mean velocities as large as 1.6 times the free stream velocity with instantaneous values of 1.8 times the free stream velocity. The airfoil motion induces these large velocities in regions that are far removed from the surface.
Influence of airfoil thickness on convected gust interaction noise
NASA Technical Reports Server (NTRS)
Kerschen, E. J.; Tsai, C. T.
1989-01-01
The case of a symmetric airfoil at zero angle of attack is considered in order to determine the influence of airfoil thickness on sound generated by interaction with convected gusts. The analysis is based on a linearization of the Euler equations about the subsonic mean flow past the airfoil. Primary sound generation is found to occur in a local region surrounding the leading edge, with the size of the local region scaling on the gust wavelength. For a parabolic leading edge, moderate leading edge thickness is shown to decrease the noise level in the low Mach number limit.
Characteristics of an Airfoil as Affected by Fabric Sag
NASA Technical Reports Server (NTRS)
Ward, Kenneth E
1932-01-01
This report presents the results of tests made at a high value of the Reynolds Number in the N.A.C.A. variable-density wind tunnel to determine the aerodynamic characteristics of an airfoil as affected by fabric sag. Tests were made of two Gottingen 387 airfoils, one having the usual smooth surface and the other having a surface modified to simulate two types of fabric sag. The results of these tests indicate that the usual sagging of the wind covering between ribs has a very small effect on the aerodynamic characteristics of an airfoil.
Approximate method of designing a two-element airfoil
NASA Astrophysics Data System (ADS)
Abzalilov, D. F.; Mardanov, R. F.
2011-09-01
An approximate method is proposed for designing a two-element airfoil. The method is based on reducing an inverse boundary-value problem in a doubly connected domain to a problem in a singly connected domain located on a multisheet Riemann surface. The essence of the method is replacement of channels between the airfoil elements by channels of flow suction and blowing. The shape of these channels asymptotically tends to the annular shape of channels passing to infinity on the second sheet of the Riemann surface. The proposed method can be extended to designing multielement airfoils.
An abbreviated Reynolds stress turbulence model for airfoil flows
NASA Technical Reports Server (NTRS)
Gaffney, R. L., Jr.; Hassan, H. A.; Salas, M. D.
1990-01-01
An abbreviated Reynolds stress turbulence model is presented for solving turbulent flow over airfoils. The model consists of two partial differential equations, one for the Reynolds shear stress and the other for the turbulent kinetic energy. The normal stresses and the dissipation rate of turbulent kinetic energy are computed from algebraic relationships having the correct asymptotic near wall behavior. This allows the model to be integrated all the way to the wall without the use of wall functions. Results for a flat plate at zero angle of attack, a NACA 0012 airfoil and a RAE 2822 airfoil are presented.
Design and experimental results for the S809 airfoil
Somers, D M
1997-01-01
A 21-percent-thick, laminar-flow airfoil, the S809, for horizontal-axis wind-turbine applications, has been designed and analyzed theoretically and verified experimentally in the low-turbulence wind tunnel of the Delft University of Technology Low Speed Laboratory, The Netherlands. The two primary objectives of restrained maximum lift, insensitive to roughness, and low profile drag have been achieved. The airfoil also exhibits a docile stall. Comparisons of the theoretical and experimental results show good agreement. Comparisons with other airfoils illustrate the restrained maximum lift coefficient as well as the lower profile-drag coefficients, thus confirming the achievement of the primary objectives.
Design and experimental results for the S805 airfoil
Somers, D.M.
1997-01-01
An airfoil for horizontal-axis wind-turbine applications, the S805, has been designed and analyzed theoretically and verified experimentally in the low-turbulence wind tunnel of the Delft University of Technology Low Speed Laboratory, The Netherlands. The two primary objectives of restrained maximum lift, insensitive to roughness, and low profile drag have been achieved. The airfoil also exhibits a docile stall. Comparisons of the theoretical and experimental results show good agreement. Comparisons with other airfoils illustrate the restrained maximum lift coefficient as well as the lower profile-drag coefficients, thus confirming the achievement of the primary objectives.
Navier-Stokes analysis of blunt trailing edge airfoils
NASA Technical Reports Server (NTRS)
Stanaway, Sharon; Mccroskey, W. J.; Kroo, Ilan
1992-01-01
The flow around blunt trailing edge airfoils was studied by solving the Reynolds-averaged Navier-Stokes equations. The solution procedure combines a grid around the airfoil with a second grid for the wake so that the time advancement over the domain is fully implicit. This is not only very efficient for the algorithm but also allows implicit solutions of a one equation turbulence model appropriate for both boundary layers and wakes. An algebraic and two one-equation turbulence models are tested for a blunt RAE 2822 airfoil section and detailed comparisons with experimental data are presented in the trailing edge region.
Optimum Transonic Airfoils Based on the Euler Equations
NASA Technical Reports Server (NTRS)
Iollo, Angelo; Salas, Manuel, D.
1996-01-01
We solve the problem of determining airfoils that approximate, in a least square sense, given surface pressure distributions in transonic flight regimes. The flow is modeled by means of the Euler equations and the solution procedure is an adjoint- based minimization algorithm that makes use of the inverse Theodorsen transform in order to parameterize the airfoil. Fast convergence to the optimal solution is obtained by means of the pseudo-time method. Results are obtained using three different pressure distributions for several free stream conditions. The airfoils obtained have given a trailing edge angle.
Transonic airfoil calculations using solution-adaptive grids
NASA Technical Reports Server (NTRS)
Holst, T. L.; Brown, D.
1981-01-01
A new algorithm for generating solution-adaptive grids (SAG) about airfoil configurations embedded in transonic flow is presented. The present SAG approach uses only the airfoil surface solution to recluster grid points on the airfoil surface, i.e., the reclustering problem is one dimension smaller than the flow-field calculation problem. Special controls automatically built into the elliptic grid generation procedure are then used to obtain grids with suitable interior behavior. This concept of redistributing grid points greatly simplifies the idea of solution-adaptive grids. Numerical results indicate significant improvements in accuracy for SAG grids relative to standard grids using the same number of points.
Performance of two transonic airfoil wind tunnels utilizing limited ventilation
NASA Technical Reports Server (NTRS)
Lee, J. D.; Gregorek, G. M.
1984-01-01
A limited-zone ventilated wall panel was developed for a closed-wall icing tunnel which permitted correct simulation of transonic flow over model rotor airfoil sections with and without ice accretions. Candidate porous panels were tested in the Ohio State University 6- x 12-inch transonic airfoil tunnel and result in essentially interference-free flow, as evidenced by pressure distributions over a NACA 0012 airfoil for Mach numbers up to 0.75. Application to the NRC 12- x 12-inch icing tunnel showed a similar result, which allowed proper transonic flow simulation in that tunnel over its full speed range.
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
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.
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.
Application of a Slotted Airfoil for UH-60A Helicopter Performance
NASA Technical Reports Server (NTRS)
Yeo, Hyeonsoo; Lim, Joon W.
2002-01-01
Recently developed forward slotted airfoils were applied to a UH-60A helicopter, and the performance has been estimated. Baseline SC1095 and SC1094 R8 airfoil characteristics were modified based on CFD calculations of an A3c slotted airfoil to incorporate aerodynamic characteristics of a high-lifting airfoil. The slotted airfoil increases maximum thrust of the UH-60A helicopter by up to 25%, but a significant penalty is observed at CTI9 less than 0.11. This penalty results from higher drag than the baseline airfoil at low angles of attack. A drag reduction at high Mach numbers is necessary to fully exploit the airfoil capability in the rotorcraft application. Preliminary comparison of the slotted airfoil with the wide chord blade shows that the slotted airfoil has limited advantages over the wide chord blade.
Aeroacoustics and aerodynamic performance of a rotor with flatback airfoils.
Paquette, Joshua A.; Barone, Matthew Franklin; Christiansen, Monica; Simley, Eric
2010-06-01
The aerodynamic performance and aeroacoustic noise sources of a rotor employing flatback airfoils have been studied in field test campaign and companion modeling effort. The field test measurements of a sub-scale rotor employing nine meter blades include both performance measurements and acoustic measurements. The acoustic measurements are obtained using a 45 microphone beamforming array, enabling identification of both noise source amplitude and position. Semi-empirical models of flatback airfoil blunt trailing edge noise are developed and calibrated using available aeroacoustic wind tunnel test data. The model results and measurements indicate that flatback airfoil noise is less than drive train noise for the current test turbine. It is also demonstrated that the commonly used Brooks, Pope, and Marcolini model for blunt trailing edge noise may be over-conservative in predicting flatback airfoil noise for wind turbine applications.
Grid Sensitivity and Aerodynamic Optimization of Generic Airfoils
NASA Technical Reports Server (NTRS)
Sadrehaghighi, Ideen; Smith, Robert E.; Tiwari, Surendra N.
1995-01-01
An algorithm is developed to obtain the grid sensitivity with respect to design parameters for aerodynamic optimization. The procedure is advocating a novel (geometrical) parameterization using spline functions such as NURBS (Non-Uniform Rational B- Splines) for defining the airfoil geometry. An interactive algebraic grid generation technique is employed to generate C-type grids around airfoils. The grid sensitivity of the domain with respect to geometric design parameters has been obtained by direct differentiation of the grid equations. A hybrid approach is proposed for more geometrically complex configurations such as a wing or fuselage. The aerodynamic sensitivity coefficients are obtained by direct differentiation of the compressible two-dimensional thin-layer Navier-Stokes equations. An optimization package has been introduced into the algorithm in order to optimize the airfoil surface. Results demonstrate a substantially improved design due to maximized lift/drag ratio of the airfoil.
Tracking of raindrops in flow over an airfoil
NASA Technical Reports Server (NTRS)
Valentine, James R.; Decker, Rand
1993-01-01
The splashback that occurs when raindrops impact an airfoil results in an 'ejecta fog' of small droplets around the leading edge. Acceleration of these droplets by the air flow field is a momentum sink for the air flow and has been hypothesized to contribute to the degradation of airfoil performance in heavy rain. Presented here is a one-way coupled Eulerian-Lagrangian particle tracking scheme to evaluate droplet number densities and momentum sink terms around a NACA 64-210 airfoil section for three rainfall rates. A laminar air flow field is determined with a standard CFD code and input to the particle tracking algorithm. Raindrops are assumed to be non-interacting, non-deforming, non-evaporating, and non-spinning spheres and are tracked through the curvilinear grid used by the air flow code. A simple model is used to simulate impacts and the resulting splashback on the airfoil surface.
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.
Study of the TRAC Airfoil Table Computational System
NASA Technical Reports Server (NTRS)
Hu, Hong
1999-01-01
The report documents the study of the application of the TRAC airfoil table computational package (TRACFOIL) to the prediction of 2D airfoil force and moment data over a wide range of angle of attack and Mach number. The TRACFOIL generates the standard C-81 airfoil table for input into rotorcraft comprehensive codes such as CAM- RAD. The existing TRACFOIL computer package is successfully modified to run on Digital alpha workstations and on Cray-C90 supercomputers. A step-by-step instruction for using the package on both computer platforms is provided. Application of the newer version of TRACFOIL is made for two airfoil sections. The C-81 data obtained using the TRACFOIL method are compared with those of wind-tunnel data and results are presented.
High-flaps for natural laminar flow airfoils
NASA Technical Reports Server (NTRS)
Morgan, Harry L.
1986-01-01
A review of the NACA and NASA low-drag airfoil research is presented with particular emphasis given to the development of mechanical high-lift flap systems and their application to general aviation aircraft. These flap systems include split, plain, single-slotted, and double-slotted trailing-edge flaps plus slat and Krueger leading-edge devices. The recently developed continuous variable-camber high-lift mechanism is also described. The state-of-the-art of theoretical methods for the design and analysis of multi-component airfoils in two-dimensional subsonic flow is discussed, and a detailed description of the Langley MCARF (Multi-Component Airfoil Analysis Program) computer code is presented. The results of a recent effort to design a single- and double-slotted flap system for the NASA high speed natural laminar flow (HSNLF) (1)-0213 airfoil using the MCARF code are presented to demonstrate the capabilities and limitations of the code.
Design and experimental results for the S814 airfoil
Somers, D.M.
1997-01-01
A 24-percent-thick airfoil, the S814, for the root region of a horizontal-axis wind-turbine blade has been designed and analyzed theoretically and verified experimentally in the low-turbulence wind tunnel of the Delft University of Technology Low Speed Laboratory, The Netherlands. The two primary objectives of high maximum lift, insensitive to roughness, and low profile drag have been achieved. The constraints on the pitching moment and the airfoil thickness have been satisfied. Comparisons of the theoretical and experimental results show good agreement with the exception of maximum lift which is overpredicted. Comparisons with other airfoils illustrate the higher maximum lift and the lower profile drag of the S814 airfoil, thus confirming the achievement of the objectives.
Analysis of high-incidence separated flow past airfoils
NASA Technical Reports Server (NTRS)
Chia, K. N.; Osswald, G. A.; Chia, U.
1989-01-01
An unsteady Navier-Stokes (NS) analysis is developed and used to carefully examine high-incidence aerodynamic separated flows past airfoils. Clustered conformal C-grids are employed for the 12 percent thick symmetric Joukowski airfoil as well as for the NACA 0012 airfoil with a sharp trailing edge. The clustering is controlled by appropriate one-dimensional stretching transformations. An attempt is made to resolve many of the dominant scales of an unsteady flow with massive separation, while maintaining the transformation metrics to be smooth and continuous in the entire flow field. A fully implicit time-marching alternating-direction implicit-block Gaussian elimination (ADI-BGE) method is employed, in which no use is made of any explicit artificial dissipation. Detailed results are obtained for massively separated, unsteady flow past symmetric Joukowski and NACA 0012 airfoils.
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.
Boundary layer development on turbine airfoil suction surfaces
NASA Technical Reports Server (NTRS)
Sharma, O. P.; Wells, R. A.; Schlinker, R. H.; Bailey, D. A.
1981-01-01
The results of a study supported by NASA under the Energy Efficient Engine Program, conducted to investigate the development of boundary layers under the influence of velocity distributions that simulate the suction sides of two state-of-the-art turbine airfoils, are presented. One velocity distribution represented a forward loaded airfoil ('squared-off' design), while the other represented an aft loaded airfoil ('aft loaded' design). These velocity distributions were simulated in a low-speed, high-aspect-ratio wind tunnel specifically designed for boundary layer investigations. It is intended that the detailed data presented in this paper be used to develop improved turbulence model suitable for application to turbine airfoil design.
Computational design and analysis of flatback airfoil wind tunnel experiment.
Mayda, Edward A.; van Dam, C.P.; Chao, David D.; Berg, Dale E.
2008-03-01
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.
Transonic airfoil and wing design using Navier-Stokes codes
NASA Technical Reports Server (NTRS)
Yu, N. J.; Campbell, R. L.
1992-01-01
An iterative design method has been implemented into 2D and 3D Navier-Stokes codes for the design of airfoils or wings with given target pressure distributions. The method begins with the analysis of an initial geometry, and obtains the analysis pressure distributions of that geometry. The differences between analysis pressures and target pressures are used to drive geometry changes through the use of a streamline curvature method. This paper describes the procedure that makes the iterative design method work for Navier-Stokes codes. Examples of 2D airfoil design, and 3D wing design are included. It is demonstrated that the method is highly effective for airfoil or wing design at flow conditions where no substantial separation occurs. Problems encountered in the airfoil design with shock induced flow separations are discussed.
Viscous effects on transonic airfoil stability and response
NASA Technical Reports Server (NTRS)
Berry, H. M.; Batina, J. T.; Yang, T. Y.
1985-01-01
Viscous effects on transonic airfoil stability and response are investigated using an integral boundary layer model coupled to the inviscid XTRAN2L transonic small disturbance code. Unsteady transonic airloads required for stability analyses are computed using a pulse transfer function analysis including viscous effects. The pulse analysis provides unsteady aerodynamic forces for a wide range of reduced frequency in a single flow field computation. Nonlinear time marching aeroelastic solutions are presented which show the effects of viscosity on airfoil response behavior and flutter. Effects of amplitude on time marching responses are demonstrated. A state space aeroelastic model employing Pade approximants to describe the unsteady airloads is used to study the effects of viscosity on transonic airfoil stability. State space dynamic pressure root loci are in good overall agreement with time marching damping and frequency estimates. Parallel sets of results with and without viscous effects reveal the effects of viscosity on transonic unsteady airloads and aeroelastic characteristics of airfoils.
Turbine airfoil with a compliant outer wall
Campbell, Christian X [Oviedo, FL; Morrison, Jay A [Oviedo, FL
2012-04-03
A turbine airfoil usable in a turbine engine with a cooling system and a compliant dual wall configuration configured to enable thermal expansion between inner and outer layers while eliminating stress formation in the outer layer is disclosed. The compliant dual wall configuration may be formed a dual wall formed from inner and outer layers separated by a support structure. The outer layer may be a compliant layer configured such that the outer layer may thermally expand and thereby reduce the stress within the outer layer. The outer layer may be formed from a nonplanar surface configured to thermally expand. In another embodiment, the outer layer may be planar and include a plurality of slots enabling unrestricted thermal expansion in a direction aligned with the outer layer.
Experimental airfoil characterization under tailored turbulent conditions
NASA Astrophysics Data System (ADS)
Heißelmann, Hendrik; Peinke, Joachim; Hölling, Michael
2016-09-01
Studies of the impact of turbulent inflow conditions on the airfoil characteristics were performed within the EU FP7 project AVATAR. The aim of this study is to provide data for the validation of simulations and the improvement of engineering tools. Chord-wise pressure distributions and highly-resolved force data of the wind turbine dedicated DU 00-W-212 profile were measured in the wind tunnel in two tailored turbulent inflow conditions generated with an active grid. A sinusoidal and an intermittent pattern with customized inflow angle fluctuations were generated providing two significantly different distributions of reduced frequencies. The obtained pressure distributions and polars from the unsteady patterns are compared to the laminar baseline case.
Aerodynamic sound of flow past an airfoil
NASA Technical Reports Server (NTRS)
Wang, Meng
1995-01-01
The long term objective of this project is to develop a computational method for predicting the noise of turbulence-airfoil interactions, particularly at the trailing edge. We seek to obtain the energy-containing features of the turbulent boundary layers and the near-wake using Navier-Stokes Simulation (LES or DNS), and then to calculate the far-field acoustic characteristics by means of acoustic analogy theories, using the simulation data as acoustic source functions. Two distinct types of noise can be emitted from airfoil trailing edges. The first, a tonal or narrowband sound caused by vortex shedding, is normally associated with blunt trailing edges, high angles of attack, or laminar flow airfoils. The second source is of broadband nature arising from the aeroacoustic scattering of turbulent eddies by the trailing edge. Due to its importance to airframe noise, rotor and propeller noise, etc., trailing edge noise has been the subject of extensive theoretical (e.g. Crighton & Leppington 1971; Howe 1978) as well as experimental investigations (e.g. Brooks & Hodgson 1981; Blake & Gershfeld 1988). A number of challenges exist concerning acoustic analogy based noise computations. These include the elimination of spurious sound caused by vortices crossing permeable computational boundaries in the wake, the treatment of noncompact source regions, and the accurate description of wave reflection by the solid surface and scattering near the edge. In addition, accurate turbulence statistics in the flow field are required for the evaluation of acoustic source functions. Major efforts to date have been focused on the first two challenges. To this end, a paradigm problem of laminar vortex shedding, generated by a two dimensional, uniform stream past a NACA0012 airfoil, is used to address the relevant numerical issues. Under the low Mach number approximation, the near-field flow quantities are obtained by solving the incompressible Navier-Stokes equations numerically at chord
Three-dimensional effects on airfoils
NASA Technical Reports Server (NTRS)
Chevallier, J. P.
1983-01-01
The effects of boundary layer flows along the walls of wind tunnels were studied to validate the transfer of two dimensional calculations to three dimensional transonic flowfield calculations. Results from trials in various wind tunnels were examind to determine the effects of the wall boundary flow on the control surfaces of an airfoil. Models sliding along a groove in the wall of a channel at sub- and transonic speeds were examined, with the finding that with either nonuniformities in the groove, or even if the channel walls are uniform, the lateral boundary layer can cause variations in the central flow region or alter the onset of shock at the transition point. Models for the effects in both turbulence and in the absence of turbulence are formulated, and it is noted that the characteristics of individual wind tunnels must be studied to quantify any existing three dimensional effects.
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.
Computation of unsteady flows over airfoils
NASA Technical Reports Server (NTRS)
Ekaterinaris, J. A.; Platzer, M. F.
1992-01-01
Two methods are described for calculating unsteady flows over rapidly pitching airfoils. The first method is based on an interactive scheme in which the inviscid flow is obtained by a panel method. The boundary layer flow is computed by an interactive method that makes use of the Hilbert integral to couple the solutions of the inviscid and viscous flow equations. The second method is based on the solution of the compressible Navier-Stokes equations. The solution of these equations is obtained with an approximately factorized numerical algorithm, and with single block or multiple grids which enable grid embedding to enhance the resolution at isolated flow regions. In addition, the attached flow region can be computed by the numerical solution of compressible boundary layer equations. Unsteady pressure distributions obtained with both methods are compared with available experimental data.
Heat Transfer of Airfoils and Plates
NASA Technical Reports Server (NTRS)
Seibert, Otto
1943-01-01
The few available test data on the heat dissipation of wholly or partly heated airfoil models are compared with the corresponding data for the flat plate as obtained by an extension of Prandtl's momentum theory, with differentiation between laminar and turbulent boundary layer and transitional region between both, the extent and appearance of which depend upon certain critical factors. The satisfactory agreement obtained justifies far-reaching conclusions in respect to other profile forms and arrangements of heated surface areas. The temperature relationship of the material quantities in its effect on the heat dissipation is discussed as far as is possible at tk.e present state of research, and it is shown that the profile drag of heated wing surfaces can increase or decrease with the temperature increase depending upon the momentarily existent structure of the boundary layer.
Natural laminar flow airfoil analysis and trade studies
NASA Technical Reports Server (NTRS)
1979-01-01
An analysis of an airfoil for a large commercial transport cruising at Mach 0.8 and the use of advanced computer techniques to perform the analysis are described. Incorporation of the airfoil into a natural laminar flow transport configuration is addressed and a comparison of fuel requirements and operating costs between the natural laminar flow transport and an equivalent turbulent flow transport is addressed.
Flow Visualization of Dynamic Stall on an Oscillating Airfoil
1989-09-01
Dynamic Stall; Dynamic lift, ’Unsteady lift; Helicopter retreating blade stall; Oscillating airfoil ; Flow visualization,’Schlieren method ;k ez.S-,’ .0...the degree of MASTER OF SCIENCE IN AERONAUTICAL ENGINEERING from the NAVAL POSTGRADUATE SCHOOL September 1989 Author...and moment behavior is quite different from the static stall associated with fixed-wing airfoils . Helicopter retreating blade stall is a dynamic
TRANSEP: A program for high lift separated flow about airfoils
NASA Technical Reports Server (NTRS)
Carlson, L. A.
1980-01-01
A method and program called TRANSEP is presented that can be used for the analysis of the flow about a low speed airfoil under high lift, massive separation conditions. Since the present program is a modification of the direct-inverse TRANDES code, it can also be used for the design and analysis of transonic airfoils, including the effects of weak viscous interaction. Interactions on program usage, program modifications to convert TRANDES to TRANSEP, and sample cases and results are given.
Active aerodynamic control of wake-airfoil interaction noise - Experiment
NASA Astrophysics Data System (ADS)
Simonich, J. C.; Lavrich, P. L.; Sofrin, T. G.; Topol, D. A.
A proof of concept experiment is conducted that shows the potential for active aerodynamic control of rotor wake/stator interaction noise in a simplified manner. A single airfoil model representing the stator was fitted with a moveable trailing edge flap controlled by a servo motor. The control system moves the motor driven flap in the correct angular displacement phase and rate to reduce the unsteady load on the airfoil during the wake interaction.
Computation of turbulent near wake for asymmetric airfoils
NASA Technical Reports Server (NTRS)
Deiwert, G. S.
1979-01-01
A numerical procedure for studying the turbulent near wake of two dimensional airfoil sections is presented. The Reynolds Navier-Stokes equations were written for flow about bodies of arbitrary geometry and solved on an arbitrary nonuniform curvilinear computational mesh. Eddy viscosity and Reynolds stress turbulence transport models are considered. Specific examples are shown for airfoil section by using an algebraic viscosity model with streamwise relaxation and the interactive Reynolds stress model.
An assessment of airfoil design by numerical optimization
NASA Technical Reports Server (NTRS)
Hicks, R. M.; Murman, E. M.; Vanderplaats, G. N.
1974-01-01
A practical procedure for optimum design of aerodynamic shapes is demonstrated. The proposed procedure uses an optimization program based on the method of feasible directions coupled with an analysis program that uses a relaxation solution of the inviscid, transonic, small-disturbance equations. Results are presented for low-drag, nonlifting transonic airfoils. Extension of the method to lifting airfoils, other speed regimes, and to three dimensions if feasible.
Symmetric airfoil geometry effects on leading edge noise.
Gill, James; Zhang, X; Joseph, P
2013-10-01
Computational aeroacoustic methods are applied to the modeling of noise due to interactions between gusts and the leading edge of real symmetric airfoils. Single frequency harmonic gusts are interacted with various airfoil geometries at zero angle of attack. The effects of airfoil thickness and leading edge radius on noise are investigated systematically and independently for the first time, at higher frequencies than previously used in computational methods. Increases in both leading edge radius and thickness are found to reduce the predicted noise. This noise reduction effect becomes greater with increasing frequency and Mach number. The dominant noise reduction mechanism for airfoils with real geometry is found to be related to the leading edge stagnation region. It is shown that accurate leading edge noise predictions can be made when assuming an inviscid meanflow, but that it is not valid to assume a uniform meanflow. Analytic flat plate predictions are found to over-predict the noise due to a NACA 0002 airfoil by up to 3 dB at high frequencies. The accuracy of analytic flat plate solutions can be expected to decrease with increasing airfoil thickness, leading edge radius, gust frequency, and Mach number.
Computer programs for smoothing and scaling airfoil coordinates
NASA Technical Reports Server (NTRS)
Morgan, H. L., Jr.
1983-01-01
Detailed descriptions are given of the theoretical methods and associated computer codes of a program to smooth and a program to scale arbitrary airfoil coordinates. The smoothing program utilizes both least-squares polynomial and least-squares cubic spline techniques to smooth interatively the second derivatives of the y-axis airfoil coordinates with respect to a transformed x-axis system which unwraps the airfoil and stretches the nose and trailing-edge regions. The corresponding smooth airfoil coordinates are then determined by solving a tridiagonal matrix of simultaneous cubic-spline equations relating the y-axis coordinates and their corresponding second derivatives. A technique for computing the camber and thickness distribution of the smoothed airfoil is also discussed. The scaling program can then be used to scale the thickness distribution generated by the smoothing program to a specific maximum thickness which is then combined with the camber distribution to obtain the final scaled airfoil contour. Computer listings of the smoothing and scaling programs are included.
Incidence angle effects on convected gust airfoil noise
NASA Technical Reports Server (NTRS)
Kerschen, E. J.; Myers, M. R.
1983-01-01
An analysis is developed which predicts the influence of airfoil mean loading on noise generation due to convected gusts. The theory is based on a linearization of the exact inviscid equations about a nonuniform compressible mean flow and the solution is developed using singular perturbation techniques. The case of a flat plate airfoil, at incidence angle alpha, interacting with three-dimensional disturbances is analyzed. It is found that in the vicinity of the airfoil leading and trailing edges, local regions are present which scale on the disturbance wavelength, with the noise generation concentrated in these regions. Away from the airfoil edges, the mean flow variation is found to be slow compared to the disturbance wavelength and no significant noise generation occurs. The mean flow variation near the leading edge generates additional noise by distorting the convected gust. The cumulative effect of the airfoil mean loading in the trailing edge region produces a 0(1) phase shift between the disturbances on the upper and lower surfaces of the airfoil. A corresponding 0(1) decrease, compared to the alpha = 0 case, is found in the noise generated at the trailing edge.
Analysis of crossover between local and massive separation on airfoils
NASA Technical Reports Server (NTRS)
Barnett, Mark
1987-01-01
The occurrence of massive separation on airfoils operating at high Reynolds number poses an important problem to the aerodynamicist. In the present study, the phenomenon of crossover, induced by airfoil thickness, between local separation and massive separation is investigated for low speed (incompressible), symmetric flow past realistic airfoil geometries. This problem is studied both for the infinite Reynolds number asymptotic limit using triple-deck theory and for finite Reynolds number using interacting boundary-layer theory. Numerical results are presented which illustrate how the flow evolves from local to massive separation as the airfoil thickness is increased. The results of the triple-deck and the interacting boundary-layer analyses are found to be in qualitative agreement for the NACA four digit series and an uncambered supercritical airfoil. The effect of turbulence on the evolution of the flow is also considered. Solutions are presented for turbulent flows past a NACA 0014 airfoil and a circular cylinder. For the latter case, the calculated surface pressure distribution is found to agree well with experimental data if the proper eddy pressure level is specified.
Design of advanced airfoil for stall-regulated wind turbines
NASA Astrophysics Data System (ADS)
Grasso, F.; Coiro, D. P.; Bizzarrini, N.; Calise, G.
2016-09-01
Nowadays, all the modern MW-class wind turbines make use of pitch control to optimize the rotor performance and control the turbine. However, for kW-range machines, stall-regulated solutions are still attractive and largely used for their simplicity and robustness. On the design phase, the aerodynamics plays a crucial role, especially concerning the selection/design of the necessary airfoils. This is because the airfoil performance should guarantee high wind turbine performance, but also the needed machine control capabilities. In the present work, the design of a new airfoil dedicated for stall machines is discussed. The design strategy makes use of numerical optimization scheme where a gradient-based algorithm is coupled with XFOIL code and an original Bezier-curves-based parameterization to describe the airfoil shape. The performances of the new airfoil are compared in free and fixed transition conditions. In addition, the performance of the rotor is analysed comparing the impact of the new geometry with alternative candidates. The results show that the new airfoil offers better performance and control than existing candidates do.
NASA Technical Reports Server (NTRS)
Barger, R. L.
1974-01-01
A method has been developed for designing families of airfoils in which the members of a family have the same basic type of pressure distribution but vary in thickness ratio or lift, or both. Thickness ratio and lift may be prescribed independently. The method which is based on the Theodorsen thick-airfoil theory permits moderate variations from the basic shape on which the family is based.
First order analysis of the effect of pitching on the drag coefficient
Wilson, R.E.; Neff, J.A.
1985-11-01
A first order analysis was made for the drag coefficient of a pitching NACA 0015 airfoil below stall. The inviscid velocity ditribution for a translating NACA 0015 airfoil was superimposed with the additional circulation velocity for a pitching ellipse. The resulting velocity distribution was used to numerically integrate a momentum/boundary layer formulation to obtain the drag coefficient. For both laminar and turbulent boundary layers it was found that the effect of pitching on the drag coefficient can be approximated by a shift in angle of attack. The shift angle was found to be a linear function of the pitching velocity and to be less than the induced angle of attack caused by the pitching.
NASA Technical Reports Server (NTRS)
Craig, Anthony P.; Hansman, R. John
1987-01-01
Wind tunnel experiments were conducted on Wortmann FX67-K170, NACA 0012, and NACA 64-210 airfoils at rain rates of 1000 mm/hr and Reynolds numbers of 310,000 to compare the aerodynamic performance degradation of the airfoils and to attempt to identify the various mechanisms which affect performance in heavy rain conditions. Lift and drag were measured in dry and wet conditions, a variety of flow visualization techniques were employed, and a computational code which predicted airfoil boundary layer behavior was used. At low angles of attack, the lift degradation in wet conditions varied significantly between the airfoils. The Wortmann section had the greatest overall lift degradation and the NACA 64-210 airfoil had the smallest. At high angles of attack, the NACA 64-210 and 0012 airfoils had improved aerodynamic performance in rain conditions due to an apparent reduction of the boundry layer separation. Performance degradation in heavy rain for all three airfoils at low angles of attack could be emulated by forced boundary layer transition near the leading edge. The secondary effect occurs at time scales consistent with top surface water runback times. The runback layer is thought to effectively alter the airfoil geometry. The severity of the performance degradation for the airfoils varied. The relative differences appeared to be related to the susceptibility of each airfoil to premature boundary layer transition.
On the general theory of thin airfoils for nonuniform motion
NASA Technical Reports Server (NTRS)
Reissner, Eric
1944-01-01
General thin-airfoil theory for a compressible fluid is formulated as boundary problem for the velocity potential, without recourse to the theory of vortex motion. On the basis of this formulation the integral equation of lifting-surface theory for an incompressible fluid is derived with the chordwise component of the fluid velocity at the airfoil as the function to be determined. It is shown how by integration by parts this integral equation can be transformed into the Biot-Savart theorem. A clarification is gained regarding the use of principal value definitions for the integral which occur. The integral equation of lifting-surface theory is used a s the starting point for the establishment of a theory for the nonstationary airfoil which is a generalization of lifting-line theory for the stationary airfoil and which might be called "lifting-strip" theory. Explicit expressions are given for section lift and section moment in terms of the circulation function, which for any given wing deflection is to be determined from an integral equation which is of the type of the equation of lifting-line theory. The results obtained are for airfoils of uniform chord. They can be extended to tapered airfoils. One of the main uses of the results should be that they furnish a practical means for the analysis of the aerodynamic span effect in the problem of wing flutter. The range of applicability of "lifting-strip" theory is the same as that of lifting-line theory so that its results may be applied to airfoils with aspect ratios as low as three.
Tests of Airfoils Designed to Delay the Compressibility Burble
NASA Technical Reports Server (NTRS)
Stack, John
1939-01-01
Development of airfoil sections suitable for high-speed applications has generally been difficult because little was known of the flow phenomenon that occurs at high speeds. A definite critical speed has been found at which serious detrimental flow changes occur that lead to serious losses in lift and large increases in drag. This flow phenomenon, called the compressibility burble, was originally a propeller problem, but with the development of higher speed aircraft serious consideration must be given to other parts of the airplane. Fundamental investigations of high-speed airflow phenomenon have provided new information. An important conclusion of this work has been the determination of the critical speed, that is, the speed at which the compressibility burble occurs. The critical speed was shown to be the translational velocity at which the sum of the translational velocity and the maximum local induced velocity at the surface of the airfoil or other body equals the local speed of sound. Obviously then higher critical speeds can be attained through the development of airfoils that have minimum induced velocity for any given value of the lift coefficient. Presumably, the highest critical speed will be attained by an airfoil that has uniform chordwise distribution of induced velocity or, in other words, a flat pressure distribution curve. The ideal airfoil for any given high-speed application is, then, that form which at its operating lift coefficient has uniform chordwise distribution of induced velocity. Accordingly, an analytical search for such airfoil forms has been conducted and these forms are now being investigated experimentally in the 23-inch high-speed wind tunnel. The first airfoils investigated showed marked improvement over those forms already available, not only as to critical speed buy also the drag at low speeds is decreased considerably. Because of the immediate marked improvement, it was considered desirable to extend the thickness and lift
New Application of Principle of Variable-camber Airfoil : Lachassagne System
NASA Technical Reports Server (NTRS)
Toussaint, A
1924-01-01
In studying the application of his system of varying the camber of airfoil sections, Mr. Lachassagne has just obtained a series of airfoil sections whose polar envelope presents truly remarkable aerodynamic properties.
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.
Two experimental supercritical laminar-flow-control swept-wing airfoils
NASA Technical Reports Server (NTRS)
Allison, Dennis O.; Dagenhart, J. Ray
1987-01-01
Two supercritical laminar-flow-control airfoils were designed for a large-chord swept-wing experiment in the Langley 8-Foot Transonic Pressure Tunnel where suction was provided through most of the model surface for boundary-layer control. The first airfoil was derived from an existing full-chord laminar airfoil by extending the trailing edge and making changes in the two lower-surface concave regions. The second airfoil differed from the first one in that it was designed for testing without suction in the forward concave region of the lower surface. Differences between the first airfoil and the one from which it was derived as well as between the first and second airfoils are discussed. Airfoil coordinates and predicted pressure distributions for the design normal Mach number of 0.755 and section lift coefficient of 0.55 are given for the three airfoils.
Preliminary study of a wing-tip vortex using laser velocimetry
NASA Technical Reports Server (NTRS)
Takahashi, R. K.; Mcalister, K. W.
1987-01-01
Measurements have been made in the wake of a semi-span NACA 0015 airfoil with emphasis on the region of the wing tip vortex. The spanwise and streamwise velocity components were measured using a two-component laser Doppler velocimeter. The purpose of the study was to initiate the operation of a laser velocimeter system and to perform preliminary wake measurements in preparation for a more extensive study of the structure and near field development of a tip vortex.
2011-12-05
Report: Grant N00014-08-0331 Technical Objectives As critical components of advanced aircraft engines , turbine airfoils require coatings for...advanced aircrafi engines , turbine airfoils require coatings for enhancement of oxidation, corrosion and thermal capabilities . Airfoil coatings ofien...Oxidation and Corrosion Protection Coatings for Enhanced Thermo-Mechanical Durability of Turbine Airfoils 5b. GRANT NUMBER N00014-08-l-0331 5c
Wind tunnel results of the high-speed NLF(1)-0213 airfoil
NASA Technical Reports Server (NTRS)
Sewall, William G.; Mcghee, Robert J.; Hahne, David E.; Jordan, Frank L., Jr.
1987-01-01
Wind tunnel tests were conducted to evaluate a natural laminar flow airfoil designed for the high speed jet aircraft in general aviation. The airfoil, designated as the High Speed Natural Laminar Flow (HSNLF)(1)-0213, was tested in two dimensional wind tunnels to investigate the performance of the basic airfoil shape. A three dimensional wing designed with this airfoil and a high lift flap system is also being evaluated with a full size, half span model.
A study of compressibility effects on dynamic stall of rapidly pitching airfoils
NASA Technical Reports Server (NTRS)
Carr, Lawrence W.; Chandrasekhara, M. S.
1991-01-01
Results of recent experimental studies into the effect of compressibility on dynamic stall of oscillating airfoils are reviewed. Stroboscopic schlieren images of the strongly unsteady flow field are presented, showing the development of the dynamic stall vortex, and its progression down the airfoil. The effect of varying free-stream Mach number and frequency of oscillation of the airfoil are demonstrated, and examples of local supersonic flow are presented, including the presence of a shock near the leading edge of the airfoil.
Finite Difference Calculation of an Inviscid Transonic Flow over Oscillating Airfoils,
1980-10-01
solutions for the NLR 7301 airfoil and the NACA 64A010 airfoil also will be obtained. They will be compared with the results obtained by Tijdeman and...and the NACA 0012 airfoil oscillating in pitch, in order to obtain several individual flow patterns. The resulting unsteady pressure distributions...shock wave locations, etc, are presented. Furthermore, the unsteady numerical results obtained by this procedure for the NLR 7301 airfoil and the NACA
Uncertainty Analysis for a Jet Flap Airfoil
NASA Technical Reports Server (NTRS)
Green, Lawrence L.; Cruz, Josue
2006-01-01
An analysis of variance (ANOVA) study was performed to quantify the potential uncertainties of lift and pitching moment coefficient calculations from a computational fluid dynamics code, relative to an experiment, for a jet flap airfoil configuration. Uncertainties due to a number of factors including grid density, angle of attack and jet flap blowing coefficient were examined. The ANOVA software produced a numerical model of the input coefficient data, as functions of the selected factors, to a user-specified order (linear, 2-factor interference, quadratic, or cubic). Residuals between the model and actual data were also produced at each of the input conditions, and uncertainty confidence intervals (in the form of Least Significant Differences or LSD) for experimental, computational, and combined experimental / computational data sets were computed. The LSD bars indicate the smallest resolvable differences in the functional values (lift or pitching moment coefficient) attributable solely to changes in independent variable, given just the input data points from selected data sets. The software also provided a collection of diagnostics which evaluate the suitability of the input data set for use within the ANOVA process, and which examine the behavior of the resultant data, possibly suggesting transformations which should be applied to the data to reduce the LSD. The results illustrate some of the key features of, and results from, the uncertainty analysis studies, including the use of both numerical (continuous) and categorical (discrete) factors, the effects of the number and range of the input data points, and the effects of the number of factors considered simultaneously.
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.
Recent progress in the analysis of iced airfoils and wings
NASA Technical Reports Server (NTRS)
Cebeci, Tuncer; Chen, Hsun H.; Kaups, Kalle; Schimke, Sue
1992-01-01
Recent work on the analysis of iced airfoils and wings is described. Ice shapes for multielement airfoils and wings are computed using an extension of the LEWICE code that was developed for single airfoils. The aerodynamic properties of the iced wing are determined with an interactive scheme in which the solutions of the inviscid flow equations are obtained from a panel method and the solutions of the viscous flow equations are obtained from an inverse three-dimensional finite-difference boundary-layer method. A new interaction law is used to couple the inviscid and viscous flow solutions. The newly developed LEWICE multielement code is amplified to a high-lift configuration to calculate the ice shapes on the slat and on the main airfoil and on a four-element airfoil. The application of the LEWICE wing code to the calculation of ice shapes on a MS-317 swept wing shows good agreement with measurements. The interactive boundary-layer method is applied to a tapered iced wing in order to study the effect of icing on the aerodynamic properties of the wing at several angles of attack.
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.
Supercritical flow past a symmetrical bicircular arc airfoil
NASA Technical Reports Server (NTRS)
Holt, Maurice; Yew, Khoy Chuah
1989-01-01
A numerical scheme is developed for computing steady supercritical flow about symmetrical airfoils, applying it to an ellipse for zero angle of attack. An algorithmic description of this new scheme is presented. Application to a symmetrical bicircular arc airfoil is also proposed. The flow field before the shock is region 1. For transonic flow, singularity can be avoided by integrating the resulting ordinary differential equations away from the body. Region 2 contains the shock which will be located by shock fitting techniques. The shock divides region 2 into supersonic and subsonic regions and there is no singularity problem in this case. The Method of Lines is used in this region and it is advantageous to integrate the resulting ordinary differential equation along the body for shock fitting. Coaxial coordinates have to be used for the bicircular arc airfoil so that boundary values on the airfoil body can be taken with one direction of the coaxial coordinates fixed. To avoid taking boundary values at + or - infinity in the coaxial co-ordinary system, approximate analytical representation of the flow field near the tips of the airfoil is proposed.
Numerical computation of viscous flow about unconventional airfoil shapes
NASA Technical Reports Server (NTRS)
Ahmed, S.; Tannehill, J. C.
1990-01-01
A new two-dimensional computer code was developed to analyze the viscous flow around unconventional airfoils at various Mach numbers and angles of attack. The Navier-Stokes equations are solved using an implicit, upwind, finite-volume scheme. Both laminar and turbulent flows can be computed. A new nonequilibrium turbulence closure model was developed for computing turbulent flows. This two-layer eddy viscosity model was motivated by the success of the Johnson-King model in separated flow regions. The influence of history effects are described by an ordinary differential equation developed from the turbulent kinetic energy equation. The performance of the present code was evaluated by solving the flow around three airfoils using the Reynolds time-averaged Navier-Stokes equations. Excellent results were obtained for both attached and separated flows about the NACA 0012 airfoil, the RAE 2822 airfoil, and the Integrated Technology A 153W airfoil. Based on the comparison of the numerical solutions with the available experimental data, it is concluded that the present code in conjunction with the new nonequilibrium turbulence model gives excellent results.
Computational Analysis of Dual Radius Circulation Control Airfoils
NASA Technical Reports Server (NTRS)
Lee-Rausch, E. M.; Vatsa, V. N.; Rumsey, C. L.
2006-01-01
The goal of the work is to use multiple codes and multiple configurations to provide an assessment of the capability of RANS solvers to predict circulation control dual radius airfoil performance and also to identify key issues associated with the computational predictions of these configurations that can result in discrepancies in the predicted solutions. Solutions were obtained for the Georgia Tech Research Institute (GTRI) dual radius circulation control airfoil and the General Aviation Circulation Control (GACC) dual radius airfoil. For the GTRI-DR airfoil, two-dimensional structured and unstructured grid computations predicted the experimental trend in sectional lift variation with blowing coefficient very well. Good code to code comparisons between the chordwise surface pressure coefficients and the solution streamtraces also indicated that the detailed flow characteristics were matched between the computations. For the GACC-DR airfoil, two-dimensional structured and unstructured grid computations predicted the sectional lift and chordwise pressure distributions accurately at the no blowing condition. However at a moderate blowing coefficient, although the code to code variation was small, the differences between the computations and experiment were significant. Computations were made to investigate the sensitivity of the sectional lift and pressure distributions to some of the experimental and computational parameters, but none of these could entirely account for the differences in the experimental and computational results. Thus, CFD may indeed be adequate as a prediction tool for dual radius CC flows, but limited and difficult to obtain two-dimensional experimental data prevents a confident assessment at this time.
Design analysis of vertical wind turbine with airfoil variation
NASA Astrophysics Data System (ADS)
Maulana, Muhammad Ilham; Qaedy, T. Masykur Al; Nawawi, Muhammad
2016-03-01
With an ever increasing electrical energy crisis occurring in the Banda Aceh City, it will be important to investigate alternative methods of generating power in ways different than fossil fuels. In fact, one of the biggest sources of energy in Aceh is wind energy. It can be harnessed not only by big corporations but also by individuals using Vertical Axis Wind Turbines (VAWT). This paper presents a three-dimensional CFD analysis of the influence of airfoil design on performance of a Darrieus-type vertical-axis wind turbine (VAWT). The main objective of this paper is to develop an airfoil design for NACA 63-series vertical axis wind turbine, for average wind velocity 2,5 m/s. To utilize both lift and drag force, some of designs of airfoil are analyzed using a commercial computational fluid dynamics solver such us Fluent. Simulation is performed for this airfoil at different angles of attach rearranging from -12°, -8°, -4°, 0°, 4°, 8°, and 12°. The analysis showed that the significant enhancement in value of lift coefficient for airfoil NACA 63-series is occurred for NACA 63-412.
Design of high lift airfoils with a Stratford distribution by the Eppler method
NASA Technical Reports Server (NTRS)
Thomson, W. G.
1975-01-01
Airfoils having a Stratford pressure distribution, which has zero skin friction in the pressure recovery area, were investigated in an effort to develop high lift airfoils. The Eppler program, an inverse conformal mapping technique where the x and y coordinates of the airfoil are developed from a given velocity distribution, was used.
General theory of airfoil sections having arbitrary shape or pressure distribution
NASA Technical Reports Server (NTRS)
Allen, H Julian
1945-01-01
In this report a theory of thin airfoils of small camber is developed which permits either the velocity distribution corresponding to a given airfoil shape, or the airfoil shape corresponding to a given velocity distribution to be calculated. The procedures to be employed in these calculations are outlined and illustrated with suitable examples.
Experimental analysis of the aeroacoustics of cascaded airfoils
NASA Astrophysics Data System (ADS)
Perry, L. A.; Lauchle, G. C.
1992-11-01
The purpose of this study is to investigate the noise radiated by various louver designs. A louver is essentially a cascade of small airfoils, operating at the same angle of attack. Louvers are commonly used in heating, ventilation, and air conditioning (HVAC) systems to provide directional control of the exit airflow. The HVAC system of an automobile can be a significant source of acoustic annoyance. Louvers are typically placed in the near field of the driver and passenger and can be major contributors to the overall interior noise level. In this study, thirteen representative dashboard registers used in automotive HVAC applications are considered. These registers vary in airfoil shape, number of airfoils, number of support struts, inlet and outlet sizes, and other physical parameters. The research documented in this thesis is directed toward a better understanding of the parameters that significantly affect the amount of noise generated by a louver.
Aeroelastic response of an airfoil with structural and aerodynamic nonlinearities
NASA Astrophysics Data System (ADS)
Kholodar, Denis Borisovich
2002-04-01
The effects of structural and aerodynamic nonlinearities on the response of aeroelastic systems are examined for a typical airfoil section. Control surface freeplay is the source of the structural nonlinearity. Classical Theodorsen aerodynamic theory is combined with an appropriate structural theory to construct a nonlinear aeroelastic model that is used to simulate responses due to a prescribed angle of attack or gust loads. A companion wing tunnel experimental program using an aeroelastic airfoil model and a rotating slotted cylinder gust generator is used to validate the theoretical results. To study nonlinear aerodynamic effects, a computational fluid dynamics (CFD) method based upon Euler equations and developed at Duke is combined with a linear structural airfoil model to determine flutter and limit cycle oscillation (LCO) response for a wide range of structural parameters and Mach number. Finally, a system identification based reduced order aerodynamic model (ROM) is also presented. It will be useful when considering flutter and LCO of more complex wing planforms.
Unsteady flow model for circulation-control airfoils
NASA Technical Reports Server (NTRS)
Rao, B. M.
1979-01-01
An analysis and a numerical lifting surface method are developed for predicting the unsteady airloads on two-dimensional circulation control airfoils in incompressible flow. The analysis and the computer program are validated by correlating the computed unsteady airloads with test data and also with other theoretical solutions. Additionally, a mathematical model for predicting the bending-torsion flutter of a two-dimensional airfoil (a reference section of a wing or rotor blade) and a computer program using an iterative scheme are developed. The flutter program has a provision for using the CC airfoil airloads program or the Theodorsen hard flap solution to compute the unsteady lift and moment used in the flutter equations. The adopted mathematical model and the iterative scheme are used to perform a flutter analysis of a typical CC rotor blade reference section. The program seems to work well within the basic assumption of the incompressible flow.
Inverse boundary-layer technique for airfoil design
NASA Technical Reports Server (NTRS)
Henderson, M. L.
1979-01-01
A description is presented of a technique for the optimization of airfoil pressure distributions using an interactive inverse boundary-layer program. This program allows the user to determine quickly a near-optimum subsonic pressure distribution which meets his requirements for lift, drag, and pitching moment at the desired flow conditions. The method employs an inverse turbulent boundary-layer scheme for definition of the turbulent recovery portion of the pressure distribution. Two levels of pressure-distribution architecture are used - a simple roof top for preliminary studies and a more complex four-region architecture for a more refined design. A technique is employed to avoid the specification of pressure distributions which result in unrealistic airfoils, that is, those with negative thickness. The program allows rapid evaluation of a designed pressure distribution off-design in Reynolds number, transition location, and angle of attack, and will compute an airfoil contour for the designed pressure distribution using linear theory.
Ice Accretions on a Swept GLC-305 Airfoil
NASA Technical Reports Server (NTRS)
Vargas, Mario; Papadakis, Michael; Potapczuk, Mark; Addy, Harold; Sheldon, David; Giriunas, Julius
2002-01-01
An experiment was conducted in the Icing Research Tunnel (IRT) at NASA Glenn Research Center to obtain castings of ice accretions formed on a 28 deg. swept GLC-305 airfoil that is representative of a modern business aircraft wing. Because of the complexity of the casting process, the airfoil was designed with three removable leading edges covering the whole span. Ice accretions were obtained at six icing conditions. After the ice was accreted, the leading edges were detached from the airfoil and moved to a cold room. Molds of the ice accretions were obtained, and from them, urethane castings were fabricated. This experiment is the icing test of a two-part experiment to study the aerodynamic effects of ice accretions.
Integral Textile Structure for 3-D CMC Turbine Airfoils
NASA Technical Reports Server (NTRS)
Marshall, David B. (Inventor); Cox, Brian N. (Inventor); Sudre, Olivier H. (Inventor)
2017-01-01
An integral textile structure for 3-D CMC turbine airfoils includes top and bottom walls made from an angle-interlock weave, each of the walls comprising warp and weft fiber tows. The top and bottom walls are merged on a first side parallel to the warp fiber tows into a single wall along a portion of their widths, with the weft fiber tows making up the single wall interlocked through the wall's thickness such that delamination of the wall is inhibited. The single wall suitably forms the trailing edge of an airfoil; the top and bottom walls are preferably joined along a second side opposite the first side and parallel to the radial fiber tows by a continuously curved section in which the weave structure remains continuous with the weave structure in the top and bottom walls, the continuously curved section being the leading edge of the airfoil.
High-fidelity modeling of airfoil interaction with upstream turbulence
NASA Astrophysics Data System (ADS)
Brodnick, Jacob
To supplement past research on low speed unsteady airfoil responses to upstream disturbances, this work proposes and investigates a method to generate a turbulent momentum source to be convected downstream and interact with an SD7003 airfoil in a high-fidelity numerical simulation. A perturbation velocity field is generated from a summation of Fourier harmonics and applied to the forcing function in the momentum terms of the Navier Stokes Equations. The result is a three-dimensional, divergence-free, convected turbulent gust with applied statistical parameters. A parametric study has been done in 2D and 3D comparing the resultant flow fields and airfoil interactions for various numerical and physical parameters.
Aerodynamic performance of an annular classical airfoil cascade
NASA Technical Reports Server (NTRS)
Bergsten, D. E.; Stauter, R. C.; Fleeter, S.
1983-01-01
Results are presented for a series of experiments that were performed in a large-scale subsonic annular cascade facility that was specifically designed to provide three-dimensional aerodynamic data for the verification of numerical-calculation codes. In particular, the detailed three-dimensional aerodynamic performance of a classical flat-plate airfoil cascade is determined for angles of incidence of 0, 5, and 10 deg. The resulting data are analyzed and are correlated with predictions obtained from NASA's MERIDL and TSONIC numerical programs. It is found that: (1) at 0 and 5 deg, the airfoil surface data show a good correlation with the predictions; (2) at 10 deg, the data are in fair agreement with the numerical predictions; and (3) the two-dimensional Gaussian similarity relationship is appropriate for the wake velocity profiles in the mid-span region of the airfoil.
Computation of viscous transonic flow about a lifting airfoil
NASA Technical Reports Server (NTRS)
Walitt, L.; Liu, C. Y.
1976-01-01
The viscous transonic flow about a stationary body in free air was numerically investigated. The geometry chosen was a symmetric NACA 64A010 airfoil at a freestream Mach number of 0.8, a Reynolds number of 4 million based on chord, and angles of attack of 0 and 2 degrees. These conditions were such that, at 2 degrees incidence unsteady periodic motion was calculated along the aft portion of the airfoil and in its wake. Although no unsteady measurements were made for the NACA 64A010 airfoil at these flow conditions, interpolated steady measurements of lift, drag, and surface static pressures compared favorably with corresponding computed time-averaged lift, drag, and surface static pressures.
Improved algorithms for circulation-control airfoils in transonic flow
NASA Technical Reports Server (NTRS)
Dvorak, Frank A.; Strash, Daniel J.; York, Brian J.; Dash, Sanford M.
1987-01-01
A zonal model for aerodynamic analysis of two-dimensional transonic circulation control airfoils has been developed. The present approach combines a transonic full potential method for the global flow field and an integral boundary layer method for regions of the airfoil excluding the wall jet with a parabolized Navier-Stokes code for resolving the wall jet region beyond the slot. Existing methods suffer from two deficiencies: the insensitivity of the calculation to small changes in the Coanda surface geometry; and the inability to predict the shock structure of the underexpanded supersonic wall jets. The present wall jet procedure involves a pressure-split approach in the streamwise sense to enable noniterative solution of the coupled continuity and normal momentum equations for increased surface sensitivity and allows for expansion of applications to sonic slot exit conditions. Encouraging results are obtained in comparison with experimental data for two circulation airfoils with subsonic wall jets.
Automated CFD for Generation of Airfoil Performance Tables
NASA Technical Reports Server (NTRS)
Strawn, Roger; Mayda, E. Q.; vamDam, C. P.
2009-01-01
A method of automated computational fluid dynamics (CFD) has been invented for the generation of performance tables for an object subject to fluid flow. The method is applicable to the generation of tables that summarize the effects of two-dimensional flows about airfoils and that are in a format known in the art as C81. (A C81 airfoil performance table is a text file that lists coefficients of lift, drag, and pitching moment of an airfoil as functions of angle of attack for a range of Mach numbers.) The method makes it possible to efficiently generate and tabulate data from simulations of flows for parameter values spanning all operational ranges of actual or potential interest. In so doing, the method also enables filling of gaps and resolution of inconsistencies in C81 tables generated previously from incomplete experimental data or from theoretical calculations that involved questionable assumptions.
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.
Numerical analysis for comparison of aerodynamic characteristics of six airfoils
NASA Astrophysics Data System (ADS)
Saad, Magedi Moh M.; Mohd, Sofian Bin; Zulkafli, Mohd Fadhli; Shibani, Wanis Mustafa E.
2017-04-01
Comparison of six airfoils; FX 63-137, FX76-100, S835, S809, NACA63415, and NACA63215, have been performed using commercial software, FLUENT and XFOIL, in order to choose the best maximum lift to drag ratio in the region of 4×10∧6 Reynolds number, and also to specify their aerodynamic coefficients in Blade Element Momentum theory (BEM). These airfoils are candidates for the use as blade turbines and another application of flying object operating at high Reynolds number. In this study, the two-dimensional model of airfoils was established through Gambit software. The range of angles of attack is from -150 to 150. Comparisons of the numerical results generally show good agreements. Numerical results of FX 63-137 were also validated with experimental data. A good agreement of lift and drag coefficient from numerical simulations of FLUENT and experimental data was obtained at 10∧6 Reynolds number.
Improved algorithms for circulation-control airfoils in transonic flow
NASA Technical Reports Server (NTRS)
Dvorak, Frank A.; Strash, Daniel J.; York, Brian J.; Dash, Sanford M.
1987-01-01
A zonal model for aerodynamic analysis of two-dimensional transonic circulation control airfoils has been developed. The present approach combines a transonic full potential method for the global flow field and an integral boundary layer method for regions of the airfoil excluding the wall jet with a parabolized Navier-Stokes code for resolving the wall jet region beyond the slot. Existing methods suffer from two deficiencies: the insensitivity of the calculation to small changes in the Coanda surface geometry; and the inability to predict the shock structure of the underexpanded supersonic wall jets. The present wall jet procedure involves a pressure-split approach in the streamwise sense to enable noniterative solution of the coupled continuity and normal momentum equations for increased surface sensitivity and allows for expansion of applications to sonic slot exit conditions. Encouraging results are obtained in comparison with experimental data for two circulation airfoils with subsonic wall jets.
Numerical simulation of the vortical flow around a pitching airfoil
NASA Astrophysics Data System (ADS)
Fu, Xiang; Li, Gaohua; Wang, Fuxin
2017-04-01
In order to study the dynamic behaviors of the flapping wing, the vortical flow around a pitching NACA0012 airfoil is investigated. The unsteady flow field is obtained by a very efficient zonal procedure based on the velocity-vorticity formulation and the Reynolds number based on the chord length of the airfoil is set to 1 million. The zonal procedure divides up the whole computation domain in to three zones: potential flow zone, boundary layer zone and Navier-Stokes zone. Since the vorticity is absent in the potential flow zone, the vorticity transport equation needs only to be solved in the boundary layer zone and Navier-Stokes zone. Moreover, the boundary layer equations are solved in the boundary layer zone. This arrangement drastically reduces the computation time against the traditional numerical method. After the flow field computation, the evolution of the vortices around the airfoil is analyzed in detail.
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.
Evaluation of Airfoil Dynamic Stall Characteristics for Maneuverability
NASA Technical Reports Server (NTRS)
Bousman, William G.; Aiken, Edwin W. (Technical Monitor)
2000-01-01
In severe maneuvers, out of necessity for a military aircraft or inadvertently for a civil aircraft, a helicopter airfoil will stall in a dynamic manner and provide lift beyond what would be calculated based on static airfoil tests. The augmented lift that occurs in dynamic stall is related to a vortex that is shed near the leading edge of the airfoil. However, directly related to the augmented lift that results from the dynamic stall vortex are significant penalties in pitching moment and drag. An understanding of the relationship between the augmented lift in dynamic stall and the associated moment and drag penalties is the purpose of this paper. This relationship is characterized using data obtained in two-dimensional wind tunnel tests and related to the problem of helicopter maneuverability.
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.
Aerodynamics of a Flapping Airfoil with a Flexible Tail
NASA Astrophysics Data System (ADS)
Lai, Alan Kai San
This dissertation presents computational solutions to an airfoil in a oscillatory heaving motion with a aeroelastically flexible tail attachment. An unsteady potential flow solver is coupled to a structural solver to obtain the aeroelastic flow solution over an inviscid fluid to investigate the propulsive performance of such a configuration. The simulation is then extended to a two-dimensional viscous solver by coupling NASA's CFL3D solver to the structural solver to study how the flow is altered by the presence of viscosity. Finally, additional simulations are done in three dimensions over wings with varying aspect ratio to study the three-dimensional effects on the propulsive performance of an airfoil with an aeroelastic tail. The computation reveals that the addition of the aeroelastic trailing edge improved the thrust generated by a heaving airfoil significantly. As the frequency of the heaving motion increases, the thrust generated by the airfoil with the tail increases exponentially. In an inviscid fluid, the increase in thrust is insufficient to overcome the increase in power required to maintain the motion and as a result the overall propulsive efficiency is reduced. When the airfoil is heaving in a viscous fluid, the presence of a suction boundary layer and the appearance of leading edge vortex increase the thrust generated to such an extent that the propulsive efficiency is increased by about 3% when compared to the same airfoil with a rigid tail. The three-dimensional computations shows that the presence of the tip vorticies suppress some of the increase in thrust observed in the two-dimensional viscous computations for short span wings. For large span wings, the overall thrust enhancing capabilities of the aeroelastic tail is preserved.
NASA Technical Reports Server (NTRS)
1979-01-01
A comprehensive review of all NASA airfoil research, conducted both in-house and under grant and contract, as well as a broad spectrum of airfoil research outside of NASA is presented. Emphasis is placed on the development of computational aerodynamic codes for airfoil analysis and design, the development of experimental facilities and test techniques, and all types of airfoil applications.
Evolving aerodynamic airfoils for wind turbines through a genetic algorithm
NASA Astrophysics Data System (ADS)
Hernández, J. J.; Gómez, E.; Grageda, J. I.; Couder, C.; Solís, A.; Hanotel, C. L.; Ledesma, JI
2017-01-01
Nowadays, genetic algorithms stand out for airfoil optimisation, due to the virtues of mutation and crossing-over techniques. In this work we propose a genetic algorithm with arithmetic crossover rules. The optimisation criteria are taken to be the maximisation of both aerodynamic efficiency and lift coefficient, while minimising drag coefficient. Such algorithm shows greatly improvements in computational costs, as well as a high performance by obtaining optimised airfoils for Mexico City's specific wind conditions from generic wind turbines designed for higher Reynolds numbers, in few iterations.
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.
Feasibility of predicting performance degradation of airfoils in heavy rain
NASA Technical Reports Server (NTRS)
Bilanin, A. J.; Quackenbush, T. R.; Feo, A.
1989-01-01
The heavy rain aerodynamic performance penalty program is detailed. This effort supported the design of a fullscale test program as well as examined the feasibility of estimating the degradation of performance of airfoils from first principles. The analytic efforts were supplemented by a droplet splashback test program in an attempt to observe the physics of impact and generation of ejecta. These tests demonstrated that the interaction of rain with an airfoil is a highly complex phenomenon and this interaction is not likely to be analyzed analytically with existing tools.
Dynamic Stall of a Pitching and Horizontally Oscillating Airfoil
NASA Astrophysics Data System (ADS)
Martinat, G.; Braza#, M.; Harran, G.; Sevrain, A.; Tzabiras, G.; Hoarau, Y.; Favier, D.
This paper provides a study of the dynamic stall of a pitching airfoil and of a pitching and horizontally oscillating airfoil at 105 Reynolds number by means of nbumerical simulation. Three turbulence models are compared in both cases: URANS Spalart-Allmaras model, URANS k—ɛ Chien model and URANS=OES model. results are in accordance with experimental data but spalart model seems to be too much viscous to provide good results and overpredict hysteresis cycle observed where URANS=OES seems to be viscousless. URANS k—ɛ Chien model is providing the best results.
A hybrid algorithm for transonic airfoil and wing design
NASA Technical Reports Server (NTRS)
Campbell, Richard L.; Smith, Leigh A.
1987-01-01
The present method for the design of transonic airfoils and wings employs a predictor/corrector approach in which an analysis code calculates the flowfield for an initial geometry, then modifies it on the basis of the difference between calculated and target pressures. This allows the design method to be straightforwardly coupled with any existing analysis code, as presently undertaken with several two- and three-dimensional potential flow codes. The results obtained indicate that the method is robust and accurate, even in the cases of airfoils with strongly supercritical flow and shocks. The design codes are noted to require computational resources typical of current pure-inverse methods.
Euler computations of AGARD Working Group 07 airfoil test cases
NASA Technical Reports Server (NTRS)
Pulliam, T. H.; Barton, J. T.
1985-01-01
In an attempt to provide a set of accurate standard test problems for computational code developers, a series of inviscid airfoil test cases were chosen by the AGARD Working Group 07, a subpanel of the AGARD Fluid Dynamics Panel. The cases include three different airfoils at transonic to supersonic conditions. A large number of international experts responded in this effort with computations that have been contrasted for accuracy and consistency. This paper is a summary of the authors' contribution to this study. In particular, the important aspects of the solution process that made it possible to obtain the high level of accuracy needed in this study are stressed.
CFD Analysis of Circulation Control Airfoils Using Fluent
NASA Technical Reports Server (NTRS)
McGowan, Gregory; Gopalarathnam, Ashok
2005-01-01
In an effort to validate computational fluid dynamics procedures for calculating flows around circulation control airfoils, the commercial flow solver FLUENT was utilized to study the flow around a general aviation circulation control airfoil. The results were compared to experimental and computational fluid dynamics results conducted at the NASA Langley Research Center. The current effort was conducted in three stages: 1. A comparison of the results for free-air conditions to those from experiments. 2. A study of wind-tunnel wall effects. and 3. A study of the stagnation-point behavior.
Airfoil design by numerical optimization using a minicomputer
NASA Technical Reports Server (NTRS)
Hicks, R. M.; Szelazek, C. A.
1978-01-01
A computer program developed for the automated design of low speed airfoils utilizes a generalized Joukowski method for aerodynamic analysis coupled with a conjugate gradient, penalty function, numerical optimization algorithm to give an efficient calculation technique for use with minicomputers. The program designs airfoils with a prescribed pressure distribution as well as those which minimize or maximize some aerodynamic force coefficient. At present the method is restricted to inviscid, incompressible flow. A typical design problem will execute in 4.5 hr on an HP 9830 minicomputer.
Modeling of heavy-gas effects on airfoil flows
NASA Technical Reports Server (NTRS)
Drela, Mark
1992-01-01
Thermodynamic models were constructed for a calorically imperfect gas and for a non-ideal gas. These were incorporated into a quasi one dimensional flow solver to develop an understanding of the differences in flow behavior between the new models and the perfect gas model. The models were also incorporated into a two dimensional flow solver to investigate their effects on transonic airfoil flows. Specifically, the calculations simulated airfoil testing in a proposed high Reynolds number heavy gas test facility. The results indicate that the non-idealities caused significant differences in the flow field, but that matching of an appropriate non-dimensional parameter led to flows similar to those in air.
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 Technical Reports Server (NTRS)
Graham, Donald J
1949-01-01
Several groups of new airfoil sections, designated as the NACA 8-series, are derived analytically to have lift characteristics at supercritical Mach numbers which are favorable in the sense that the abrupt loss of lift, characteristic of the usual airfoil section at Mach numbers above the critical, is avoided. Aerodynamic characteristics determined from two-dimensional wind-tunnel tests at Mach numbers up to approximately 0.9 are presented for each of the derived airfoils. Comparisons are made between the characteristics of these airfoils and the corresponding characteristics of representative NACA 6-series airfoils.
Effects of Airfoil Thickness and Maximum Lift Coefficient on Roughness Sensitivity: 1997--1998
Somers, D. M.
2005-01-01
A matrix of airfoils has been developed to determine the effects of airfoil thickness and the maximum lift to leading-edge roughness. The matrix consists of three natural-laminar-flow airfoils, the S901, S902, and S903, for wind turbine applications. The airfoils have been designed and analyzed theoretically and verified experimentally in the Pennsylvania State University low-speed, low-turbulence wind tunnel. The effect of roughness on the maximum life increases with increasing airfoil thickness and decreases slightly with increasing maximum lift. Comparisons of the theoretical and experimental results generally show good agreement.
NASA Technical Reports Server (NTRS)
Jacobs, Eastman N
1932-01-01
Report presents the results of wind tunnel tests on a group of eight very thick airfoils having sections of the same thickness as those used near the roots of tapered airfoils. The tests were made to study certain discontinuities in the characteristic curves that have been obtained from previous tests of these airfoils, and to compare the characteristics of the different sections at values of the Reynolds number comparable with those attained in flight. The discontinuities were found to disappear as the Reynolds number was increased. The results obtained from the large-scale airfoil, a symmetrical airfoil having a thickness ratio of 21 per cent, has the best general characteristics.
Effect of pivot location and passive heave on propulsion from a pitching airfoil
NASA Astrophysics Data System (ADS)
Mackowski, A. W.; Williamson, C. H. K.
2017-01-01
We experimentally investigate the propulsive characteristics of a pitching NACA 0012 airfoil section, with emphasis on thrust and propulsive efficiency, at a Reynolds number of 1.7 ×104 . For the sake of mechanical simplicity, we consider an airfoil restricted to a single actuator in the pitching direction. We examine the effect of changing the airfoil's axis of rotation, finding that contrary to Garrick's linear theory, there exists a pitching axis near the airfoil that maximizes propulsive efficiency. Next, we examine the effect of placing passive springs on the airfoil in the heave (transverse) direction using our Cyber-Physical Fluid Dynamics technique. This elastic heaving motion allows the airfoil to combine pitching and heaving modes while being actuated only in the pitching direction. Two sets of dynamics are considered: one case where the airfoil is weighted unevenly and pitched about its center of mass (so that the resulting heaving motion is independent of inertial forces), and another case where the airfoil's center of mass is fixed at its centroid. For pitching at an amplitude of 8∘ and a reduced frequency k of two, we find that elastic heave produces a maximum propulsive efficiency of 35%, compared to 25% without any heave motion. Further, while operating at the same efficiency as the static-pivot case, we find that passive heaving greatly increases the magnitude of the airfoil's thrust. The airfoil configurations with highest propulsive efficiency generally involve pitching near or ahead of the airfoil's leading edge.
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.
Wind Tunnel Aeroacoustic Tests of Six Airfoils for Use on Small Wind Turbines: Preprint
Migliore, P.; Oerlemans, S.
2003-12-01
Aeroacoustic tests of seven airfoils were performed in an open jet anechoic wind tunnel. Six of the airfoils are candidates for use on small wind turbines operating at low Reynolds number. One airfoil was tested for comparison to benchmark data. Tests were conducted with and without boundary layer tripping. In some cases a turbulence grid was placed upstream in the test section to investigate inflow turbulence noise. An array of 48 microphones was used to locate noise sources and separate airfoil noise from extraneous tunnel noise. Trailing edge noise was dominant for all airfoils in clean tunnel flow. With the boundary layer untripped, several airfoils exhibited pure tones that disappeared after proper tripping was applied. In the presence of inflow turbulence, leading edge noise was dominant for all airfoils.
Design of a shape adaptive airfoil actuated by a Shape Memory Alloy strip for airplane tail
NASA Astrophysics Data System (ADS)
Shirzadeh, R.; Raissi Charmacani, K.; Tabesh, M.
2011-04-01
Of the factors that mainly affect the efficiency of the wing during a special flow regime, the shape of its airfoil cross section is the most significant. Airfoils are generally designed for a specific flight condition and, therefore, are not fully optimized in all flight conditions. It is very desirable to have an airfoil with the ability to change its shape based on the current regime. Shape memory alloy (SMA) actuators activate in response to changes in the temperature and can recover their original configuration after being deformed. This study presents the development of a method to control the shape of an airfoil using SMA actuators. To predict the thermomechanical behaviors of an SMA thin strip, 3D incremental formulation of the SMA constitutive model is implemented in FEA software package ABAQUS. The interactions between the airfoil structure and SMA thin strip actuator are investigated. Also, the aerodynamic performance of a standard airfoil with a plain flap is compared with an adaptive airfoil.
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.
Wind tunnel tests of two airfoils for wind turbines operating at high reynolds numbers
Sommers, D.; Tangler, J.
2000-06-29
The objectives of this study were to verify the predictions of the Eppler Airfoil Design and Analysis Code for Reynolds numbers up to 6 x 106 and to acquire the section characteristics of two airfoils being considered for large, megawatt-size wind turbines. One airfoil, the S825, was designed to achieve a high maximum lift coefficient suitable for variable-speed machines. The other airfoil, the S827, was designed to achieve a low maximum lift coefficient suitable for stall-regulated machines. Both airfoils were tested in the NASA Langley Low-Turbulence Pressure Tunnel (LTPT) for smooth, fixed-transition, and rough surface conditions at Reynolds numbers of 1, 2, 3, 4, and 6 x 106. The results show the maximum lift coefficient of both airfoils is substantially underpredicted for Reynolds numbers over 3 x 106 and emphasized the difficulty of designing low-lift airfoils for high Reynolds numbers.
Cantrell, Kirk J.; Serne, R. Jeffrey; Zachara, John M.; Krupka, Kenneth M.; Dresel, P. Evan; Brown, Christopher F.; Freshley, Mark D.
2008-01-17
This appendix discusses the geology of the Hanford Site and singe-shell tank (SST) waste management areas (WMAs). The purpose is to provide the most recent geochemical information available for the SST WMAs and the Integrated Disposal Facility. This appendix summarizes the information in the geochemistry data package for the SST WMAs.
Experimental Investigation on Airfoil Shock Control by Plasma Aerodynamic Actuation
NASA Astrophysics Data System (ADS)
Sun, Quan; Cheng, Bangqin; Li, Yinghong; Cui, Wei; Jin, Di; Li, Jun
2013-11-01
An experimental investigation on airfoil (NACA64—215) shock control is performed by plasma aerodynamic actuation in a supersonic tunnel (Ma = 2). The results of schlieren and pressure measurement show that when plasma aerodynamic actuation is applied, the position moves forward and the intensity of shock at the head of the airfoil weakens. With the increase in actuating voltage, the total pressure measured at the head of the airfoil increases, which means that the shock intensity decreases and the control effect increases. The best actuation effect is caused by upwind-direction actuation with a magnetic field, and then downwind-direction actuation with a magnetic field, while the control effect of aerodynamic actuation without a magnetic field is the most inconspicuous. The mean intensity of the normal shock at the head of the airfoil is relatively decreased by 16.33%, and the normal shock intensity is relatively reduced by 27.5% when 1000 V actuating voltage and upwind-direction actuation are applied with a magnetic field. This paper theoretically analyzes the Joule heating effect generated by DC discharge and the Lorentz force effect caused by the magnetic field. The discharge characteristics are compared for all kinds of actuation conditions to reveal the mechanism of shock control by plasma aerodynamic actuation.
Effects of changing airfoil aerodynamic characteristics on turning diffuser performances
NASA Astrophysics Data System (ADS)
Noh@Seth, Nur Hazirah; Isa, Norasikin Mat
2017-04-01
Combining both turning and diffusing activities by using 3-dimensional turning diffuser offer more advantages as compared to bend-diffuser systems. However, adverse pressure gradient and curvature design of turning diffuser itself will result in existence of secondary flow at the inner wall and both left and right wall region, which will disrupt turning diffuser performance. Introduction of baffle has successfully proven able to improve the performance of 3-dimensional turning diffuser in terms of both pressure recovery and flow uniformity using experimental approach. Preliminary design airfoil referred to previous study was used, and the results were used to validate present study simulation work. Aerodynamic characteristic of the airfoil were varied and series of simulation were conducted to study the effects of changing aerodynamic characteristics of an airfoil on turning diffuser performance. Optimum parameters proposed in this study have successfully improved 3-dimensional turning diffuser performance by 7.20% in terms of flow uniformity and 6.16% in terms of pressure recovery. Turning diffuser efficiency was also improved with increment of 6.12%. These parameters can be used in the future for reference in the design of airfoil baffle especially for usage involving 3-dimensional turning diffuser.
Two-Dimensional Grids About Airfoils and Other Shapes
NASA Technical Reports Server (NTRS)
Sorenson, R.
1982-01-01
GRAPE computer program generates two-dimensional finite-difference grids about airfoils and other shapes by use of Poisson differential equation. GRAPE can be used with any boundary shape, even one specified by tabulated points and including limited number of sharp corners. Numerically stable and computationally fast, GRAPE provides aerodynamic analyst with efficient and consistant means of grid generation.
The potential influence of rain on airfoil performance
NASA Technical Reports Server (NTRS)
Dunham, R. Earl, Jr.
1987-01-01
The potential influence of heavy rain on airfoil performance is discussed. Experimental methods for evaluating rain effects are reviewed. Important scaling considerations for extrapolating model data are presented. It is shown that considerable additional effort, both analytical and experimental, is necessary to understand the degree of hazard associated with flight operations in rain.
Cost Effective Repair Techniques for Turbine Airfoils. Volume I
1978-11-01
Turbine blades and vanes in current engines are subjected to the most hostile environment...payoff potential in turbine vanes / blades . The criteria used included: • Incidence of damage - Scrapped or damaged turbine airfoils at the ALC centers...Corporate Author: GENERAL ELECTRIC CO CINCINNATI OHIO AIRCRAFT ENGINE GROUP Unclassified Title: (U) Cost Effective Repair Techniques for Turbine
Cost Effective Repair Techniques for Turbine Airfoils. Volume 2
1979-04-01
BLADES , *GUIDE VANES , *REPAIR, TURBOFAN ENGINES , DIFFUSION BONDING, COST EFFECTIVENESS Identifiers: (U) * Turbine vanes , TF-39 engines , Activated...REPAIR TECHNIQUES FOR TURBINE AIRFOILS J. A. WEIN W. R. YOUNG GENERAL ELECTRIC COMPANY AIRCRAFT ENGINE GROUP CINCINNATI, OHIO 45215 APRIL 1979...Author: GENERAL ELECTRIC CO CINCINNATI OH AIRCRAFT ENGINE BUSINESS GROUP Unclassified Title: (U) Cost Effective Repair Techniques for
Decomposing the aerodynamic forces of low-Reynolds flapping airfoils
NASA Astrophysics Data System (ADS)
Moriche, Manuel; Garcia-Villalba, Manuel; Flores, Oscar
2016-11-01
We present direct numerical simulations of flow around flapping NACA0012 airfoils at relatively small Reynolds numbers, Re = 1000 . The simulations are carried out with TUCAN, an in-house code that solves the Navier-Stokes equations for an incompressible flow with an immersed boundary method to model the presence of the airfoil. The motion of the airfoil is composed of a vertical translation, heaving, and a rotation about the quarter of the chord, pitching. Both motions are prescribed by sinusoidal laws, with a reduced frequency of k = 1 . 41 , a pitching amplitude of 30deg and a heaving amplitude of one chord. Both, the mean pitch angle and the phase shift between pitching and heaving motions are varied, to build a database with 18 configurations. Four of these cases are analysed in detail using the force decomposition algorithm of Chang (1992) and Martín Alcántara et al. (2015). This method decomposes the total aerodynamic force into added-mass (translation and rotation of the airfoil), a volumetric contribution from the vorticity (circulatory effects) and a surface contribution proportional to viscosity. In particular we will focus on the second, analysing the contribution of the leading and trailing edge vortices that typically appear in these flows. This work has been supported by the Spanish MINECO under Grant TRA2013-41103-P. The authors thankfully acknowledge the computer resources provided by the Red Española de Supercomputacion.
Computer Program to Obtain Ordinates for NACA Airfoils
NASA Technical Reports Server (NTRS)
Ladson, Charles L.; Brooks, Cuyler W., Jr.; Hill, Acquilla S.; Sproles, Darrell W.
1996-01-01
Computer programs to produce the ordinates for airfoils of any thickness, thickness distribution, or camber in the NACA airfoil series were developed in the early 1970's and are published as NASA TM X-3069 and TM X-3284. For analytic airfoils, the ordinates are exact. For the 6-series and all but the leading edge of the 6A-series airfoils, agreement between the ordinates obtained from the program and previously published ordinates is generally within 5 x 10(exp -5) chord. Since the publication of these programs, the use of personal computers and individual workstations has proliferated. This report describes a computer program that combines the capabilities of the previously published versions. This program is written in ANSI FORTRAN 77 and can be compiled to run on DOS, UNIX, and VMS based personal computers and workstations as well as mainframes. An effort was made to make all inputs to the program as simple as possible to use and to lead the user through the process by means of a menu.
Plastic covering on airfoil structure provides smooth uninterrupted surface
NASA Technical Reports Server (NTRS)
Kinzler, J. A.; Fehrenkamp, L. G.; Heffernam, J. T.; Lee, W. S.
1975-01-01
Primed surface is covered with adhesive. Sheet of plastic film is stretched over adhesive and mechanical holder is used to apply tension to ends of sheet to make it conform to surface of airfoil. After adhesive cures, plastic can be trimmed with sharp cutting tool.
Effectiveness of Thermal-Pneumatic Airfoil-Ice-Protection System
NASA Technical Reports Server (NTRS)
Gowan, William H., Jr.; Mulholland, Donald R.
1951-01-01
Icing and drag investigations were conducted in the NACA Lewis icing research tunnel employing a combination thermal-pneumatic de-icer mounted on a 42-inch-chord NACA 0018 airfoil. The de-icer consisted of a 3-inch-wide electrically heated strip symmetrically located about the leading edge with inflatable tubes on the upper and lower airfoil surfaces aft of the heated area. The entire de-icer extended to approximately 25 percent of chord. A maximum power density of 9.25 watts per square inch was required for marginal ice protection on the airfoil leading edge at an air temperature of 00 F and an airspeed of 300 miles per hour. Drag measurements indicated, that without icing, the de-icer installation increased the section drag to approximately 140 percent of that of the bare airfoil; with the tubes inflated, this value increased to a maximum of approximately 620 percent. A 2-minute tube-inflation cycle prevented excessive ice formation on the inflatable area although small scattered residual Ice formations remained after inflation and were removed intermittently during later cycles. Effects of the time lag of heater temperatures after initial application of power and the insulating effect of ice formations on heater temperatures were also determined.
Numerical modeling of the vortex/airfoil interaction
NASA Technical Reports Server (NTRS)
Panaras, Argyris G.
1987-01-01
A modeling of the vortex-airfoil interaction is presented in which the finite-area of the real vortices is taken into consideration. Two vortex models are used. In the first, a disturbed piece of vorticity layer is simulated by four rows of discrete vortices of small strength. In the second, a number of discrete vortices is arranged within a circle. The first model may simulate a shear layer or a wake, while the second, a well-formed vortex. The method was applied to the calculation of the pressure induced on the surface of the airfoil by the interacting vortex. Both models give similar results. It was found that for large distances of the vortex from the surface of the airfoil, the consideration or not of the finite-area of the vortex is not a significant factor in determining the induced pressure field. However, when the distance of the vortex from the surface is reduced, its shape is distorted and the induced pressure pulses have lower amplitude than the ones induced by an equivalent point vortex. In the limit, where the vortex impinges on the leading edge of the airfoil, it is split into two and the time dependent pressure coefficient takes even negative values at some time intervals.
Advanced turbine study. [airfoil coling in rocket turbines
NASA Technical Reports Server (NTRS)
1982-01-01
Experiments to determine the available increase in turbine horsepower achieved by increasing turbine inlet temperature over a range of 1800 to 2600 R, while applying current gas turbine airfoil cling technology are discussed. Four cases of rocket turbine operating conditions were investigated. Two of the cases used O2/H2 propellant, one with a fuel flowrate of 160 pps, the other 80 pps. Two cases used O2/CH4 propellant, each having different fuel flowrates, pressure ratios, and inlet pressures. Film cooling was found to be the required scheme for these rocket turbine applications because of the high heat flux environments. Conventional convective or impingement cooling, used in jet engines, is inadequate in a rocket turbine environment because of the resulting high temperature gradients in the airfoil wall, causing high strains and low cyclic life. The hydrogen-rich turbine environment experienced a loss, or no gain, in delivered horsepower as turbine inlet temperature was increased at constant airfoil life. The effects of film cooling with regard to reduced flow available for turbine work, dilution of mainstream gas temperature and cooling reentry losses, offset the relatively low specific work capability of hydrogen when increasing turbine inlet temperature over the 1800 to 2600 R range. However, the methane-rich environment experienced an increase in delivered horsepower as turbine inlet temperature was increased at constant airfoil life. The results of a materials survey and heat transfer and durability analysis are 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.
3D scanning and printing of airfoils for modular UAS
NASA Astrophysics Data System (ADS)
Dahlgren, Robert P.; Pinsker, Ethan A.; Dary, Omar G.; Ogunbiyi, Joab A.; Mazhari, Arash Alex
2017-02-01
The NASA Ames Research Center has been developing small unmanned airborne systems (UAS) based upon remotecontrolled military aircraft such as the RQ-14 DragonEye and RQ-11 Raven manufactured by AeroVironment. The first step is replacing OEM avionics with COTS avionics that do not use military frequencies for command and control. 3D printing and other rapid prototyping techniques are used to graft RQ-14 components into new "FrankenEye" aircraft and RQ-11 components into new "FrankenRaven" airframes. To that end, it is necessary to design new components to concatenate wing sections into elongated wingspans, construct biplane architectures, attach payload pods, and add control surfaces. When making components such as wing splices it is critical that the curvature and angles of the splice identically match the existing wing at the mating surfaces. The RQ-14 has a thick, simple airfoil with a rectangular planform and no twist or dihedral which make splice development straightforward. On the other hand the RQ-11 has a much thinner sailplane-type airfoil having a tapered polyhedral planform. 3D scanning of the Raven wings with a NextEngine scanner could not capture the complex curvature of the high-performance RQ-11 airfoil, resulting in non-matching and even misshapen splice prototypes. To characterize the airfoil a coordinate measuring machine (CMM) was employed to measure the wing's shape, fiducials and mounting features, enabling capture of the subtle curves of the airfoil and the leading and trailing edges with high fidelity. In conclusion, both rapid and traditional techniques are needed to precisely measure and fabricate wing splice components.
Adjoint-based airfoil shape optimization in transonic flow
NASA Astrophysics Data System (ADS)
Gramanzini, Joe-Ray
The primary focus of this work is efficient aerodynamic shape optimization in transonic flow. Adjoint-based optimization techniques are employed on airfoil sections and evaluated in terms of computational accuracy as well as efficiency. This study examines two test cases proposed by the AIAA Aerodynamic Design Optimization Discussion Group. The first is a two-dimensional, transonic, inviscid, non-lifting optimization of a Modified-NACA 0012 airfoil. The second is a two-dimensional, transonic, viscous optimization problem using a RAE 2822 airfoil. The FUN3D CFD code of NASA Langley Research Center is used as the ow solver for the gradient-based optimization cases. Two shape parameterization techniques are employed to study their effect and the number of design variables on the final optimized shape: Multidisciplinary Aerodynamic-Structural Shape Optimization Using Deformation (MASSOUD) and the BandAids free-form deformation technique. For the two airfoil cases, angle of attack is treated as a global design variable. The thickness and camber distributions are the local design variables for MASSOUD, and selected airfoil surface grid points are the local design variables for BandAids. Using the MASSOUD technique, a drag reduction of 72.14% is achieved for the NACA 0012 case, reducing the total number of drag counts from 473.91 to 130.59. Employing the BandAids technique yields a 78.67% drag reduction, from 473.91 to 99.98. The RAE 2822 case exhibited a drag reduction from 217.79 to 132.79 counts, a 39.05% decrease using BandAids.
Single Airfoil Gust Response Problem: Category 3, Problem 1
NASA Technical Reports Server (NTRS)
Scott, James R.
2004-01-01
An unsteady aerodynamic code, called GUST3D (ref. 3), has been developed to solve equation (8) for flows with periodic vortical disturbances. The code uses a frequency-domain approach with second-order central differences and a pressure radiation condition in the far field. GUST3D requires as input certain mean flow quantities which are calculated separately by a potential flow solver. The solver calculates the mean ow using a Gothert's Rule approximation (ref. 3). On the airfoil surface, it uses the solution calculated by the potential code FLO36 (ref. 4). Figures 1-2 show the mean pressure along the airfoil surface for the two airfoil geometries. In Figures 3 - 8, we present the RMS pressure on the airfoil surface. Each figure shows three GUST3D solutions (calculated on grids with different far-field boundary locations). Three solutions are shown to provide some indication of the numerical uncertainty in the results. Figures 9 - 13 present the acoustic intensity. We again show three solutions per case. Note that no results are presented for the k1 = k2 = 2.0 loaded airfoil case, as an acceptable solution could not be obtained. A few comments need to be made about the results shown. First, since the last Workshop, the GUST3D code has been substantially upgraded. This includes implementing a more accurate far-field boundary condition (ref. 5) and developing improved gridding capabilities. This is the reason for any differences that may exist between the present results and results from the last Workshop. Second, the intensity results on the circle R = 4C were obtained using a Kirchoff method (ref. 6). The Kirchoff surface was the circle R = 2C. Finally, the GUST3D code is most accurate for low reduced frequencies. A new domain decomposition approach (ref. 7) has been developed to improve accuracy. Both the single domain and domain decomposition approaches were used in generating the present results.
Wind-Tunnel Investigation of an NACA 23021 Airfoil with a 0.32-Airfoil-Chord Double Slotted Flap
NASA Technical Reports Server (NTRS)
Fischel, Jack; Riebe, John M
1944-01-01
An investigation was made in the LMAL 7- by 10-foot wind tunnel of a NACA 23021 airfoil with a double slotted flap having a chord 32 percent of the airfoil chord (0.32c) to determine the aerodynamic section characteristics with the flaps deflected at various positions. The effects of moving the fore flap and rear flap as a unit and of deflecting or removing the lower lip of the slot were also determined. Three positions were selected for the fore flap and at each position the maximum lift of the airfoil was obtained with the rear flap at the maximum deflection used at that fore-flap position. The section lift of the airfoil increased as the fore flap was extended and maximum lift was obtained with the fore flap deflected 30 deg in the most extended position. This arrangement provided a maximum section lift coefficient of 3.31, which was higher than the value obtained with either a 0.2566c or a 0.40c single-slotted-flap arrangement and 0.25 less than the value obtained with a 0.4c double-slotted-flap arrangement on the same airfoil. The values of the profile-drag coefficient obtained with the 0.32c double slotted flap were larger than those for the 0.2566c or 0.40c single slotted flaps for section lift coefficients between 1.0 and approximately 2.7. At all values of the section lift coefficient above 1.0, the 0.40c double slotted flap had a lower profile drag than the 0.32c double slotted flap. At various values of the maximum section lift coefficient produced by various flap defections, the 0.32c double slotted flap gave negative section pitching-moment coefficients that were higher than those of other slotted flaps on the same airfoil. The 0.32c double slotted flap gave approximately the same maximum section lift coefficient as, but higher profile-drag coefficients over the entire lift range than, a similar arrangement of a 0.30c double slotted flap on an NACA 23012 airfoil.
The Effect of Aerodynamic Evaluators on the Multi-Objective Optimization of Flatback Airfoils
NASA Astrophysics Data System (ADS)
Miller, M.; Slew, K. Lee; Matida, E.
2016-09-01
With the long lengths of today's wind turbine rotor blades, there is a need to reduce the mass, thereby requiring stiffer airfoils, while maintaining the aerodynamic efficiency of the airfoils, particularly in the inboard region of the blade where structural demands are highest. Using a genetic algorithm, the multi-objective aero-structural optimization of 30% thick flatback airfoils was systematically performed for a variety of aerodynamic evaluators such as lift-to-drag ratio (Cl/Cd), torque (Ct), and torque-to-thrust ratio (Ct/Cn) to determine their influence on airfoil shape and performance. The airfoil optimized for Ct possessed a 4.8% thick trailing-edge, and a rather blunt leading-edge region which creates high levels of lift and correspondingly, drag. It's ability to maintain similar levels of lift and drag under forced transition conditions proved it's insensitivity to roughness. The airfoil optimized for Cl/Cd displayed relatively poor insensitivity to roughness due to the rather aft-located free transition points. The Ct/Cn optimized airfoil was found to have a very similar shape to that of the Cl/Cd airfoil, with a slightly more blunt leading-edge which aided in providing higher levels of lift and moderate insensitivity to roughness. The influence of the chosen aerodynamic evaluator under the specified conditions and constraints in the optimization of wind turbine airfoils is shown to have a direct impact on the airfoil shape and performance.
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.
Advancements in adaptive aerodynamic technologies for airfoils and wings
NASA Astrophysics Data System (ADS)
Jepson, Jeffrey Keith
Although aircraft operate over a wide range of flight conditions, current fixed-geometry aircraft are optimized for only a few of these conditions. By altering the shape of the aircraft, adaptive aerodynamics can be used to increase the safety and performance of an aircraft by tailoring the aircraft for multiple flight conditions. Of the various shape adaptation concepts currently being studied, the use of multiple trailing-edge flaps along the span of a wing offers a relatively high possibility of being incorporated on aircraft in the near future. Multiple trailing-edge flaps allow for effective spanwise camber adaptation with resulting drag benefits over a large speed range and load alleviation at high-g conditions. The research presented in this dissertation focuses on the development of this concept of using trailing-edge flaps to tailor an aircraft for multiple flight conditions. One of the major tasks involved in implementing trailing-edge flaps is in designing the airfoil to incorporate the flap. The first part of this dissertation presents a design formulation that incorporates aircraft performance considerations in the inverse design of low-speed laminar-flow adaptive airfoils with trailing-edge cruise flaps. The benefit of using adaptive airfoils is that the size of the low-drag region of the drag polar can be effectively increased without increasing the maximum thickness of the airfoil. Two aircraft performance parameters are considered: level-flight maximum speed and maximum range. It is shown that the lift coefficients for the lower and upper corners of the airfoil low-drag range can be appropriately adjusted to tailor the airfoil for these two aircraft performance parameters. The design problem is posed as a part of a multidimensional Newton iteration in an existing conformal-mapping based inverse design code, PROFOIL. This formulation automatically adjusts the lift coefficients for the corners of the low-drag range for a given flap deflection as
Aerodynamic performance of a small vertical axis wind turbine using an overset grid method
NASA Astrophysics Data System (ADS)
Bangga, Galih; Solichin, Mochammad; Daman, Aida; Sa'adiyah, Devy; Dessoky, Amgad; Lutz, Thorsten
2017-08-01
The present paper aims to asses the aerodynamic performance of a small vertical axis wind turbine operating at a small wind speed of 5 m/s for 6 different tip speed ratios (λ=2-7). The turbine consists of two blades constructed using the NACA 0015 airfoil. The study is carried out using computational fluid dynamics (CFD) methods employing an overset grid approach. The (URANS) SST k - ω is used as the turbulence model. For the preliminary study, simulations of the NACA 0015 under static conditions for a broad range of angle of attack and a rotating two-bladed VAWT are carried out. The results are compared with available measurement data and a good agreement is obtained. The simulations demonstrate that the maximum power coefficient attained is 0.45 for λ=4. The aerodynamic loads hysteresis are presented showing that the dynamic stall effect decreases with λ.
NASA Astrophysics Data System (ADS)
Manela, A.; Halachmi, M.
2015-06-01
The acoustic signature of side-by-side airfoils, subject to small-amplitude harmonic pitching and incoming flow unsteadiness, is investigated. The two-dimensional near-field problem is formulated using thin-airfoil theory, where flow unsteadiness is modeled as a passing line vortex, and wake evolution is calculated via the Brown and Michael formula. Assuming that the setup is acoustically compact, acoustic radiation is obtained by means of the Powell-Howe acoustic analogy. The associated compact Green's function is calculated numerically using potential-flow analysis of the fluid-structure flow domain. Results, comparing the acoustic radiation of the double-airfoil system to a reference case of a single airfoil, point to several mechanisms of sound attenuation and sound amplification, caused by airfoil-airfoil and airfoils-wake interactions. It is found that counter-phase pitching of the airfoils results in effective cloaking of the system, which otherwise becomes significantly noisy (as a 5/2-power of the pitching frequency) at large frequencies. In addition, depending on the distance between airfoils, in-phase pitching may result in an acoustic signature equivalent to a single airfoil (when the airfoils are adjacent) or to two separate airfoils (when the airfoils are far apart). In general, flow unsteadiness produces more sound when interacting with a double (compared with a single) airfoil setup. However, airfoils' nonlinear wake-wake interactions give rise to a sound reduction mechanism, which becomes most efficient at times when incoming vorticity passes above airfoils' leading and trailing edges. The present scheme can be readily extended to consider the acoustic properties of various double-airfoil configurations, as well as multiple (> 2) airfoil setups.
Method for forming a liquid cooled airfoil for a gas turbine
Grondahl, Clayton M.; Willmott, Leo C.; Muth, Myron C.
1981-01-01
A method for forming a liquid cooled airfoil for a gas turbine is disclosed. A plurality of holes are formed at spaced locations in an oversized airfoil blank. A pre-formed composite liquid coolant tube is bonded into each of the holes. The composite tube includes an inner member formed of an anti-corrosive material and an outer member formed of a material exhibiting a high degree of thermal conductivity. After the coolant tubes have been bonded to the airfoil blank, the airfoil blank is machined to a desired shape, such that a portion of the outer member of each of the composite tubes is contiguous with the outer surface of the machined airfoil blank. Finally, an external skin is bonded to the exposed outer surface of both the machined airfoil blank and the composite tubes.
Influence of airfoil thickness on sound generated by high-frequency gust interactions
NASA Technical Reports Server (NTRS)
Tsai, C. T.; Kerschen, E. J.
1992-01-01
The sound radiated by interaction of a short wavelength gust with a symmetric thin airfoil is analyzed. The theory is based on a linearization of the Euler equations about the subsonic mean flow past the airfoil. The sound generation mechanism is found to be concentrated in a local region surrounding the parabolic nose of the airfoil; the size of this local region scales on the gust wavelength. At low Mach numbers, moderate values of airfoil thickness decrease the sound power, while at higher Mach numbers the sound power tends to increase with airfoil thickness. Airfoil thickness produces dramatic changes in the far field directivity. Both the sound power and the directivity are strong functions of the gust orientation.
Low-speed wind-tunnel results for symmetrical NASA LS(1)-0013 airfoil
NASA Technical Reports Server (NTRS)
Ferris, James C.; Mcghee, Robert J.; Barnwell, Richard W.
1987-01-01
A wind-tunnel test has been conducted in the Langley Low-Turbulence Pressure Tunnel to evaluate the performance of a symmetrical NASA LS(1)-0013 airfoil which is a 13-percent-thick, low-speed airfoil. The airfoil contour was obtained from the thickness distribution of a 13-percent-thick, high-performance airfoil developed for general aviation airplanes. The tests were conducted at Mach numbers from 0.10 tp 0.37 over a Reynolds number range from about 0.6 to 12.0 X 10 to the 6th power. The angle of attack varied from about -8 to 20 degrees. The results indicate that the aerodynamic characteristics of the present airfoil are similar to, but slightly better than, those of the NACA 0012 airfoil.
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.
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)
Maresh, J. L.; Bragg, M. B.
1984-01-01
A method has been developed to predict the contamination of an airfoil by insects and the resultant performance penalty. Insect aerodynamics have been modeled and the impingement of insects on an airfoil are solved by calculating their trajectories. Upon impact, insect rupture and the resulting height of the debris is determined based on experimental data. A boundary layer analysis is performed to determine which insects cause boundary layer transition and the resultant drag penalty. A contaminated airfoil figure of merit is presented to be used to compare airfoil susceptibility. Results show that the insect contamination effects depend on accretion conditions, airfoil angle of attack and Reynolds number. The importance of the stagnation region to designing airfoils for minimum drag penalties is discussed.
The Effects of the Critical Ice Accretion on Airfoil and Wing Performance
NASA Technical Reports Server (NTRS)
Selig, Michael S.; Bragg, Michael B.; Saeed, Farooq
1998-01-01
In support of the NASA Lewis Modern Airfoils Ice Accretion Test Program, the University of Illinois at Urbana-Champaign provided expertise in airfoil design and aerodynamic analysis to determine the aerodynamic effect of ice accretion on modern airfoil sections. The effort has concentrated on establishing a design/testing methodology for "hybrid airfoils" or "sub-scale airfoils," that is, airfoils having a full-scale leading edge together with a specially designed and foreshortened aft section. The basic approach of using a full-scale leading edge with a foreshortened aft section was considered to a limited extent over 40 years ago. However, it was believed that the range of application of the method had not been fully exploited. Thus a systematic study was being undertaken to investigate and explore the range of application of the method so as to determine its overall potential.
Influence of airfoil thickness on sound generated by high-frequency gust interactions
NASA Technical Reports Server (NTRS)
Tsai, C. T.; Kerschen, E. J.
1992-01-01
The sound radiated by interaction of a short wavelength gust with a symmetric thin airfoil is analyzed. The theory is based on a linearization of the Euler equations about the subsonic mean flow past the airfoil. The sound generation mechanism is found to be concentrated in a local region surrounding the parabolic nose of the airfoil; the size of this local region scales on the gust wavelength. At low Mach numbers, moderate values of airfoil thickness decrease the sound power, while at higher Mach numbers the sound power tends to increase with airfoil thickness. Airfoil thickness produces dramatic changes in the far field directivity. Both the sound power and the directivity are strong functions of the gust orientation.
Tests of related forward-camber airfoils in the variable-density wind tunnel
NASA Technical Reports Server (NTRS)
Jacobs, Eastman N; Pinkerton, Robert M; Greenberg, Harry
1937-01-01
A recent investigation of numerous related airfoils indicated that positions of camber forward of the usual location resulted in an increase of the maximum lift. As an extension of this investigation, a series of forward-camber airfoils has been developed, the members of which show airfoil characteristics superior to those of the airfoils previously investigated. The primary object of this report is to present fully corrected results for airfoils in the useful range of shapes. With the data thus made available, an airplane designer may intelligently choose the best possible airfoil-section shape for a given application and may predict to a reasonable degree the aerodynamic characteristics to be expected in flight from the section shape chosen.
NASA Technical Reports Server (NTRS)
Addy, Harold E., Jr.; Broeren, Andy P.; Potapczuk, Mark G.; Lee, Sam; Guffond, Didier; Montreuil, Emmanuel; Moens, Frederic
2016-01-01
This report documents the data collected during the large wind tunnel campaigns conducted as part of the SUNSET project (StUdies oN Scaling EffecTs due to ice) also known as the Ice-Accretion Aerodynamics Simulation study: a joint effort by NASA, the Office National d'Etudes et Recherches Aérospatiales (ONERA), and the University of Illinois. These data form a benchmark database of full-scale ice accretions and corresponding ice-contaminated aerodynamic performance data for a two-dimensional (2D) NACA 23012 airfoil. The wider research effort also included an analysis of ice-contaminated aerodynamics that categorized ice accretions by aerodynamic effects and an investigation of subscale, low- Reynolds-number ice-contaminated aerodynamics for the NACA 23012 airfoil. The low-Reynolds-number investigation included an analysis of the geometric fidelity needed to reliably assess aerodynamic effects of airfoil icing using artificial ice shapes. Included herein are records of the ice accreted during campaigns in NASA Glenn Research Center's Icing Research Tunnel (IRT). Two different 2D NACA 23012 airfoil models were used during these campaigns; an 18-in. (45.7-cm) chord (subscale) model and a 72-in. (182.9-cm) chord (full-scale) model. The aircraft icing conditions used during these campaigns were selected from the Federal Aviation Administration's (FAA's) Code of Federal Regulations (CFR) Part 25 Appendix C icing envelopes. The records include the test conditions, photographs of the ice accreted, tracings of the ice, and ice depth measurements. Model coordinates and pressure tap locations are also presented. Also included herein are the data recorded during a wind tunnel campaign conducted in the F1 Subsonic Pressurized Wind Tunnel of ONERA. The F1 tunnel is a pressured, high- Reynolds-number facility that could accommodate the full-scale (72-in. (182.9-cm) chord) 2D NACA 23012 model. Molds were made of the ice accreted during selected test runs of the full-scale model
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.
Aerodynamic Characteristics of a Number of Modified NACA Four-Digit-Series Airfoil Sections
NASA Technical Reports Server (NTRS)
Loftin, Laurence K., Jr.; Cohen, Kenneth G.
1947-01-01
Theoretical pressure distributions and measured lift, drag, and pitching moment characteristics at three values of Reynolds number are presented for a group of NACA four-digit-series airfoil sections modified for high-speed applications. The effectiveness of flaps applied to these airfoils and the effect of standard leading-edge roughness were also investigated at one value of Reynolds number. Results are also presented of tests of three conventional NACA four-digit-series airfoil sections.
Peak power and blade loads on stall-regulated rotors as influenced by different airfoil families
Tangler, J.L.; Tu, P.K.C.
1988-08-01
At the Solar Energy Research Institute (SERI), new airfoils have been developed to help improve the performance and economics of horizontal-axis wind turbines (HAWTS). The objective of this study was to compare the performance characteristics of one of these airfoil families to other commonly used airfoil series for a typical three-bladed, stall-regulated HAWT. The traditional airfoil series chosen for comparison with SERI's new thin airfoil family were the NACA 23XXX, NACA 44XX, and NASA LS(1). The Micon 110 wind turbine was chosen because it is a typical three-bladed, stall-regulated rigid rotor system. The performance characteristics of the different airfoil series were derived analytically using the Eppler airfoil design code in the analysis mode. On a relative basis, this approach to comparing airfoils was considered more accurate than using airfoil performance characteristics based on wind-tunnel test data. After generating the performance characteristics for each airfoil series, the subsequent rotor performance and blade loads were calculated using SERI's PROPSH computer code. Resulting annual energy output, which is dependent on the wind-speed distribution, was calculated using SERI's Systems Engineering and Analysis Computer Code (SEACC). The results of the study show that fixed-wing airfoils generally result in excessive peak power for stall regulated, rigid rotors. By operating the wind turbine at a less desirable blade pitch angle, peak power can be reduced at the expense of higher mean blade loads and lower annual energy output. In contrast, the thin airfoil family was designed to reduce peak power at optimum blade pitch to minimize blade loads and maximize annual energy output. 7 refs., 12 figs.
S830, S831, and S832 Airfoils: November 2001-November 2002
Somers, D. M.
2005-08-01
A family of quiet, thick, natural-laminar-flow airfoils, the S830, S831, and S832, for 40 - 50-meter-diameter, variable-speed/variable-pitch, horizontal-axis wind turbines has been designed and analyzed theoretically. The two primary objectives of high maximum lift, relatively insensitive to roughness, and low profile drag have been achieved. The airfoils should exhibit docile stalls, which meet the design goal. The constraints on the pitching moment and the airfoils thicknesses have been satisfied.
S833, S834, and S835 Airfoils: November 2001--November 2002
Somers, D. M.
2005-08-01
A family of quiet, thick, natural-laminar-flow airfoils, the S833, S834, and S835, for 1 - 3-meter-diameter, variable-speed/variable-pitch, horizontal-axis wind turbines has been designed and analyzed theoretically. The two primary objectives of high maximum lift, relatively insensitive to roughness, and low profile drag have been achieved. The airfoils should exhibit docile stalls, which meet the design goal. The constraints on the pitching moment and the airfoils thicknesses have been satisfied.
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.
Wind Noise Measurements of Microphones Embedded in the Airfoil of an UAV
2012-10-01
Embedded in the Airfoil of an UAV October 2012 Wayne E. Prather and William G. Frazier National Center for Physical Acoustics...platforms has been increasing in recent years. Embedding microphones in the skin of an airfoil is one option for how acoustic sensors might be...employed on a small Unmanned Aerial Vehicle (UAV). The analysis of wind noise measurements made by an array of microphones embedded in the airfoil of a
NASA Technical Reports Server (NTRS)
Nicks, Oran W.; Korkan, Kenneth D.
1991-01-01
Two reports on student activities to determine the properties of a new laminar airfoil which were delivered at a conference on soaring technology are presented. The papers discuss a wind tunnel investigation and analysis of the SM701 airfoil and verification of the SM701 airfoil aerodynamic charcteristics utilizing theoretical techniques. The papers are based on a combination of analytical design, hands-on model fabrication, wind tunnel calibration and testing, data acquisition and analysis, and comparison of test results and theory.
Systematic Airfoil Tests in the Large Wind Tunnel of the DVL
NASA Technical Reports Server (NTRS)
Doetsch, H; Kramer, M
1938-01-01
The present report is a description of systematic tests at maximum lift on airfoils with and without split flap and of profile drag at low lift. In order to obtain an opinion as to the suitability of the airfoils with flaps, the maximum-lift measurements were repeated on airfoils with split flaps. The profile drag at low lift was arrived at by direct weighing and momentum measurements and, since the profiles were of unusual depth, extended to large Reynolds numbers.
Airfoil geometry converter: From Selig and Lednicer to GEO and mesh formats
NASA Astrophysics Data System (ADS)
Vogeltanz, Tomáš
2017-07-01
In this paper, a utility for the conversion of the airfoil geometry is described and its function is demonstrated. The utility can convert a Selig, Lednicer, or VSP airfoil file to the universal GEO format, and then it may mesh the geometry for a following CFD analysis. The first section lists the formats which may be used for the storage of airfoil data and for meshing. The main section introduces the utility, briefly describes the code, and shows an example of use.
Transition and Turbulence Structure in the Boundary Layers of an Oscillating Airfoil
1989-12-31
measurements near the trailing edge of an oscillating NACA 64A010 airfoil operating at various fiequencies and in unstalled condition. It was...to identify: (i) the conditions under which a sinusoidally oscillating NACA 0012 airfoil operates with a leading edge separation bubble; (ii) the...vortex formation. These experiments were conducted using an NACA 0012 airfoil with a tripping wire to promote transition of the boundary layer. The
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
PROFILE: Airfoil Geometry Manipulation and Display. User's Guide
NASA Technical Reports Server (NTRS)
Collins, Leslie; Saunders, David
1997-01-01
This report provides user information for program PROFILE, an aerodynamics design utility for plotting, tabulating, and manipulating airfoil profiles. A dozen main functions are available. The theory and implementation details for two of the more complex options are also presented. These are the REFINE option, for smoothing curvature in selected regions while retaining or seeking some specified thickness ratio, and the OPTIMIZE option, which seeks a specified curvature distribution. Use of programs QPLOT and BPLOT is also described, since all of the plots provided by PROFILE (airfoil coordinates, curvature distributions, pressure distributions)) are achieved via the general-purpose QPLOT utility. BPLOT illustrates (again, via QPLOT) the shape functions used by two of PROFILE's options. These three utilities should be distributed as one package. They were designed and implemented for the Applied Aerodynamics Branch at NASA Ames Research Center, Moffett Field, California. They are all written in FORTRAN 77 and run on DEC and SGI systems under OpenVMS and IRIX.
Supersonic flow past pointed-nose thin airfoils
NASA Astrophysics Data System (ADS)
Hemdan, Hamdi T.; Selim, Abdulfatah A.
1992-05-01
Accurate results for steady inviscid supersonic flow past any pointed-nose thin airfoil at small angle of attack are obtained in this paper. Straightforward perturbation from some basic wedge flow is applied to a recently developed hypersonic theory with second-order terms in the longitudinal curvature included. The hypersonic equations are thus reduced to two systems of simple linear ordinary differential equations for which polynomial solutions of the first and second degree are obtained. Further, the hypersonic theory is extended to include second-order effects on the basic wedge flow. This resulted in great increase in the accuracy of the results for curved airfoils besides, the range of validity of the theory extended to cover moderate supersonic speeds with quite reasonable accuracy.
A Computational Modeling Mystery Involving Airfoil Trailing Edge Treatments
NASA Astrophysics Data System (ADS)
Choo, Yeunun; Epps, Brenden
2015-11-01
In a curious result, Fairman (2002) observed that steady RANS calculations predicted larger lift than the experimentally-measured data for six different airfoils with non-traditional trailing edge treatments, whereas the time average of unsteady RANS calculations matched the experiments almost exactly. Are these results reproducible? If so, is the difference between steady and unsteady RANS calculations a numerical artifact, or is there a physical explanation? The goals of this project are to solve this thirteen year old mystery and further to model viscous/load coupling for airfoils with non-traditional trailing edges. These include cupped, beveled, and blunt trailing edges, which are common anti-singing treatments for marine propeller sections. In this talk, we present steady and unsteady RANS calculations (ANSYS Fluent) with careful attention paid to the possible effects of asymmetric unsteady vortex shedding and the modeling of turbulence anisotropy. The effects of non-traditional trailing edge treatments are visualized and explained.
Aerodynamic coefficients in generalized unsteady thin airfoil theory
NASA Technical Reports Server (NTRS)
Williams, M. H.
1980-01-01
Two cases are considered: (1) rigid body motion of an airfoil-flap combination consisting of vertical translation of given amplitude, rotation of given amplitude about a specified axis, and rotation of given amplitude of the control surface alone about its hinge; the upwash for this problem is defined mathematically; and (2) sinusoidal gust of given amplitude and wave number, for which the upwash is defined mathematically. Simple universal formulas are presented for the most important aerodynamic coefficients in unsteady thin airfoil theory. The lift and moment induced by a generalized gust are evaluated explicitly in terms of the gust wavelength. Similarly, in the control surface problem, the lift, moment, and hinge moments are given as explicit algebraic functions of hinge location. These results can be used together with any of the standard numerical inversion routines for the elementary loads (pitch and heave).
Airfoil for a turbine of a gas turbine engine
Liang, George
2010-12-21
An airfoil for a turbine of a gas turbine engine is provided. The airfoil comprises a main body comprising a wall structure defining an inner cavity adapted to receive a cooling air. The wall structure includes a first diffusion region and at least one first metering opening extending from the inner cavity to the first diffusion region. The wall structure further comprises at least one cooling circuit comprising a second diffusion region and at least one second metering opening extending from the first diffusion region to the second diffusion region. The at least one cooling circuit may further comprise at least one third metering opening, at least one third diffusion region and a fourth diffusion region.
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.
An Experimental Study of Airfoil Self-Noise
NASA Astrophysics Data System (ADS)
Shannon, Daniel W.; Morris, Scott C.; Ross, Mark
2007-11-01
The acoustic signature of airfoils at low Mach number is typically dominated by noise generated at the trailing edge. This ``self-noise'' is dipole-like in nature, and results from the scattering of the turbulent pressure fluctuations in proximity to a sharp edge. Predicting the radiated sound requires a detailed understanding of the stochastic properties of the turbulence, as well as knowledge of the acoustic Green's function for a particular geometry. The present research has focused on the direct measurement of edge scattering acoustics for a number of geometries. The experiments were conducted in an Anechoic Wind Tunnel using a pair of 40 microphone phased arrays. The test specimens include compact and non-compact airfoils, circular cylinders, blunt and sharp trailing edges, and a blown slot edge used for circulation control. The discussion will focus on the comparison of the various results in order to derive an improved understanding of the sound generation physics.
Predicting aerodynamic characteristic of typical wind turbine airfoils using CFD
Wolfe, W.P.; Ochs, S.S.
1997-09-01
An investigation was conducted into the capabilities and accuracy of a representative computational fluid dynamics code to predict the flow field and aerodynamic characteristics of typical wind-turbine airfoils. Comparisons of the computed pressure and aerodynamic coefficients were made with wind tunnel data. This work highlights two areas in CFD that require further investigation and development in order to enable accurate numerical simulations of flow about current generation wind-turbine airfoils: transition prediction and turbulence modeling. The results show that the laminar-to turbulent transition point must be modeled correctly to get accurate simulations for attached flow. Calculations also show that the standard turbulence model used in most commercial CFD codes, the k-e model, is not appropriate at angles of attack with flow separation. 14 refs., 28 figs., 4 tabs.
Horizontal axis wind turbine post stall airfoil characteristics synthesization
NASA Technical Reports Server (NTRS)
Tangler, James L.; Ostowari, Cyrus
1995-01-01
Blade-element/momentum performance prediction codes are routinely used for wind turbine design and analysis. A weakness of these codes is their inability to consistently predict peak power upon which the machine structural design and cost are strongly dependent. The purpose of this study was to compare post-stall airfoil characteristics synthesization theory to a systematically acquired wind tunnel data set in which the effects of aspect ratio, airfoil thickness, and Reynolds number were investigated. The results of this comparison identified discrepancies between current theory and the wind tunnel data which could not be resolved. Other factors not previously investigated may account for these discrepancies and have a significant effect on peak power prediction.
Airfoil in sinusoidal motion in a pulsating stream
NASA Technical Reports Server (NTRS)
Greenberg, J Mayo
1947-01-01
The forces and moments on a two-dimensional airfoil executing harmonic motions in a pulsating stream are derived on the basis of non-stationary incompressible potential flow theory, with the inclusion of the effect of the continuous sheet of vortices shed from the trailing edge. An assumption as to the form of the wake is made with a certain degree of approximation. A comparison with previous work applicable only to the special case of a stationary airfoil is made by means of a numerical example, and the excellent agreement obtained shows that the wake approximation is quite sufficient. The results obtained are expected to be useful in considerations of forced vibrations and flutter of rotary wing aircraft.
Control theory based airfoil design using the Euler equations
NASA Technical Reports Server (NTRS)
Jameson, Antony; Reuther, James
1994-01-01
This paper describes the implementation of optimization techniques based on control theory for airfoil design. In our previous work it was shown that control theory could be employed to devise effective optimization procedures for two-dimensional profiles by using the potential flow equation with either a conformal mapping or a general coordinate system. The goal of our present work is to extend the development to treat the Euler equations in two-dimensions by procedures that can readily be generalized to treat complex shapes in three-dimensions. Therefore, we have developed methods which can address airfoil design through either an analytic mapping or an arbitrary grid perturbation method applied to a finite volume discretization of the Euler equations. Here the control law serves to provide computationally inexpensive gradient information to a standard numerical optimization method. Results are presented for both the inverse problem and drag minimization 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.
A Theory of Unstaggered Airfoil Cascades in Compressible Flow
NASA Technical Reports Server (NTRS)
Spurr, Robert A.; Allen, H. Julian
1947-01-01
By use of the methods of thin airfoil theory, which include effects of compressibility, rela.tio^as are developed which permit the rapid determination of the pressure distribution over an unstaggered cascade of airfoils of a given profile, and the determination of the profile shape necessary to yield a given pressure distribution for small chord gap ratios, For incompressible flow the results of the theory are compared with available examples obtained by the more exact method of conformal transformation. Although the theory is developed for small chord/gap ratios, these comparisons show that it may be extended to chord/gap ratios of order unity, at least for low speed flows. Choking of cascades, a phenomenon of particular importance in compressor design, is considered.
A theory of unstaggered airfoil cascades in compressible flow
NASA Technical Reports Server (NTRS)
Spurr, Robert A; Allen, H Julian
1947-01-01
By use of the methods of thin airfoil theory, which include effects of compressibility, relations are developed which permit the rapid determination of the pressure distribution over an unstaggered cascade of airfoils of a given profile, and the determination of the profile shape necessary to yield a given pressure distribution for small chord/gap ratios. For incompressible flow the results of the theory are compared with available examples obtained by the more exact method of conformal transformation. Although the theory is developed for small chord/gap ratios, these comparisons show that it may be extended to chord/gap ratios of order unity, at least for low-speed flows. Choking cascades, a phenomenon of particular importance in compressor design, is considered.
Horizontal axis wind turbine post stall airfoil characteristics synthesization
Tangler, J.L. . Wind Energy Research Center); Ostowari, C. )
1991-06-01
Blade-element/momentum performance prediction codes are routinely used for wind turbine design and analysis. A weakness of these codes is their inability to consistently predict peak power upon which the machine structural design and cost are strongly dependent. The purpose of this study was to compare post-stall airfoil characteristics synthesization theory to a systematically acquired wind tunnel data set in which the effects of aspect ratio, airfoil thickness, and Reynolds number were investigated. The results of this comparison identified discrepancies between current theory and the wind tunnel data which could not be resolved. Other factors not previously investigated may account for these discrepancies and have a significant effect on peak power prediction. 5 refs., 3 figs.
Effect of oscillation frequency on wind turbine airfoil dynamic stall
NASA Astrophysics Data System (ADS)
Zhou, Z.; Li, C.; Nie, J. B.; Chen, Y.
2013-12-01
At the same oscillation amplitude, Reynolds Number, mean angle of attack, the dynamic stall characteristics of the NREL S809 airfoil undergoing sinusoidal pitch oscillations of different oscillation frequencies were investigated with modified k-ω SST turbulence model of CFD solution for two-dimensional numerical simulation. The predicted lift, drag coefficients and moment coefficients were compared with the Ohio State University wind tunnel test results, which showed a good agreement. The birth, development and breaking off of eddies were analyzed through streamline distribution around airfoil and the influence of oscillation frequencies on dynamic stall characteristics was also described and analyzed in detail, which enrich the database of dynamic stall characteristics needed by the quantization of oscillation frequencies on dynamic characteristics and prove that sliding mesh method is reliable when dealing with dynamic stall problems.
Improving turbine engine compressor performance retention through airfoil coatings
NASA Technical Reports Server (NTRS)
Friedrich, L. A.
1981-01-01
In order to evaluate the potential effectiveness of coatings in limiting erosive damage to compressor airfoils, an effort was initiated to evaluate candidate coatings for substrate alloys typically used in commercial engine high compressor blades. Laboratory and rig erosion testing of plasma deposited and diffusion coatings described in this paper have shown the potential of a two to four fold improvement in erosion life. The selective application of these coatings to approximately the outer third of the airfoil avoids coating the fatigue critical region of the blade, thus providing erosion resistance potentially without compromising the fatigue strength of the blade. Both the plasma and the diffusion coatings also offer the advantage of low initial cost and a multi-source production base.
An airfoil pitch apparatus-modeling and control design
NASA Astrophysics Data System (ADS)
Andrews, Daniel R.
1989-03-01
The study of dynamic stall of rapidly pitching airfoils is being conducted at NASA Ames Research Center. Understanding this physical phenomenon will aid in improving the maneuverability of fighter aircraft as well as civilian aircraft. A wind tunnel device which can linearly pitch and control an airfoil with rapid dynamic response is needed for such tests. To develop a mechanism capable of high accelerations, an accurate model and control system is created. The model contains mathematical representations of the mechanical system, including mass, spring, and damping characteristics for each structural element, as well as coulomb friction and servovalve saturation. Electrical components, both digital and analog, linear and nonlinear, are simulated. The implementation of such a high-performance system requires detailed control design as well as state-of-the-art components. This paper describes the system model, states the system requirements, and presents results of its theoretical performance which maximizes the structural and hydraulic aspects of this system.
CAST-10-2/DOA 2 Airfoil Studies Workshop Results
NASA Technical Reports Server (NTRS)
Ray, Edward J. (Compiler); Hill, Acquilla S. (Compiler)
1989-01-01
During the period of September 23 through 27, 1988, the Transonic Aerodynamics Division at the Langely Research Center hosted an International Workshop on CAST-10-2/DOA 2 Airfoil Studies. The CAST-10 studies were the outgrowth of several cooperative study agreements among the NASA, the NAE of Canada, the DLR of West Germany, and the ONERA of France. Both theoretical and experimental CAST-10 airfoil results that were obtained form an extensive series of tests and studies, were reviewed. These results provided an opportunity to make direct comparisons of adaptive wall test section (AWTS) results from the NASA 0.3-meter Transonic Cryogenic Tunnel and ONERA T-2 AWTS facilities with conventional ventilated wall wind tunnel results from the Canadian high Reynolds number two-dimensional test facility. Individual papers presented during the workshop are included.
Kasprzyk airfoil. The first wind-tunnel tests
NASA Technical Reports Server (NTRS)
Wusatowski, T.
1984-01-01
The Kasprzyk slotted flap glider airfoil (the Kasper wing) enabling glider flight at 32 km/h and 0.5 m/sec descent speed was wind tunnel tested in the U.S. The test layout is described and reasons offered for discrepancies between wind tunnel results and Polish in flight data: high induced drag caused by relative size of model wing span and tunnel, by vortex attenuators on the model and their proximity to the tunnel wall, nonsimilarity between flow over a smooth wing and flow over the Kasprzyk wing with bound vortices, obstruction of the tunnel test chamber cross section by the model wing, discrepant Reynolds numbers, and model airfoil aspect ratio much smaller than the prototype. The overall results offer partial confirmation of the Kasprzyk theory, but further in tunnel and in flight studies are recommended.
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
PIV and Acoustic Investigation for a 2D Wind Turbine Airfoil
NASA Astrophysics Data System (ADS)
Wang, Guannan; Glauser, Mark
2012-11-01
This study investigated the aerodynamic characteristics of a 2D airfoil designed for wind turbine applications using PIV and surface pressure measurements. The experiments were carried out in a low speed wind tunnel with/without large scale unsteadiness in the flow and with/without active closed loop blowing control on the suction surface of the airfoil. This study also measured the acoustic signal emitted from the same type of airfoil with six far field microphones in an anechoic chamber and the results indicated that the unsteadiness in the freestream affected the noise characteristics of the airfoil significantly. Funded by DoE through University of Minnesota Wind Energy Consortium.
Initial Circulation and Peak Vorticity Behavior of Vortices Shed from Airfoil Vortex Generators
NASA Technical Reports Server (NTRS)
Wendt, Bruce J.; Biesiadny, Tom (Technical Monitor)
2001-01-01
An extensive parametric study of vortices shed from airfoil vortex generators has been conducted to determine the dependence of initial vortex circulation and peak vorticity on elements of the airfoil geometry and impinging flow conditions. These elements include the airfoil angle of attack, chord length, span, aspect ratio, local boundary layer thickness, and free stream Mach number. In addition, the influence of airfoil-to-airfoil spacing on the circulation and peak vorticity has been examined for pairs of co-rotating and counter-rotating vortices. The vortex generators were symmetric airfoils having a NACA-0012 cross-sectional profile. These airfoils were mounted either in isolation, or in pairs, on the surface of a straight pipe. The turbulent boundary layer thickness to pipe radius ratio was about 17 percent. The circulation and peak vorticity data were derived from cross-plane velocity measurements acquired with a seven-hole probe at one chord-length downstream of the airfoil trailing edge location. The circulation is observed to be proportional to the free-stream Mach number, the angle-of-attack, and the span-to-boundary layer thickness ratio. With these parameters held constant, the circulation is observed to fall off in monotonic fashion with increasing airfoil aspect ratio. The peak vorticity is also observed to be proportional to the free-stream Mach number, the airfoil angle-of-attack, and the span-to-boundary layer thickness ratio. Unlike circulation, however, the peak vorticity is observed to increase with increasing aspect ratio, reaching a peak value at an aspect ratio of about 2.0 before falling off again at higher values of aspect ratio. Co-rotating vortices shed from closely spaced pairs of airfoils have values of circulation and peak vorticity under those values found for vortices shed from isolated airfoils of the same geometry. Conversely, counter-rotating vortices show enhanced values of circulation and peak vorticity when compared to values
Compilation of Information on the Transonic Attachment of Flows at the Leading Edges of Airfoils
NASA Technical Reports Server (NTRS)
Lindsey, Walter F; Landrum, Emma Jean
1958-01-01
Schlieren photographs have been compiled of the two-dimensional flow at transonic speeds past 37 airfoils. These airfoils have variously shaped profiles, and some are related in thickness and camber. The data for these airfoils were analyzed to provide basic information on the flow changes involved and to determine factors affecting transonic-flow attachment, which is a transition from separated to unseparated flow at the leading edges of two-dimensional airfoils at fixed angles as the subsonic Mach number is increased.
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.
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.
Impingement of Water Droplets on NACA 65A004 Airfoil at 8 deg Angle of Attack
NASA Technical Reports Server (NTRS)
Brun, R. J.; Gallagher, H. M.; Vogt, D. E.
1954-01-01
The trajectories of droplets in the air flowing past an NACA 65AO04 airfoil at an angle of attack of 8 deg were determined.. The amount of water in droplet form impinging on the airfoil, the area of droplet impingement, and the rate of droplet impingement per unit area on the airfoil surface were calculated from the trajectories and presented to cover a large range of flight and atmospheric conditions. These impingement characteristics are compared briefly with those previously reported for the same airfoil at an angle of attack of 4 deg.
Inverse airfoil design procedure using a multigrid Navier-Stokes method
NASA Technical Reports Server (NTRS)
Malone, J. B.; Swanson, R. C.
1991-01-01
The Modified Garabedian McFadden (MGM) design procedure was incorporated into an existing 2-D multigrid Navier-Stokes airfoil analysis method. The resulting design method is an iterative procedure based on a residual correction algorithm and permits the automated design of airfoil sections with prescribed surface pressure distributions. The new design method, Multigrid Modified Garabedian McFadden (MG-MGM), is demonstrated for several different transonic pressure distributions obtained from both symmetric and cambered airfoil shapes. The airfoil profiles generated with the MG-MGM code are compared to the original configurations to assess the capabilities of the inverse design method.
Wind tunnel wall interference effects on a supercritical airfoil at transonic speeds
NASA Technical Reports Server (NTRS)
Blackwell, J. A., Jr.; Pounds, G. A.
1976-01-01
Wind tunnel tests of a 10% supercritical airfoil have been conducted in the Lockheed Compressible Flow Facility at transonic speeds to determine the effects of varying wind tunnel wall porosity on airfoil performance. Wall configurations ranging in porosity from 1.3% to 10% were investigated at Reynolds numbers of 7 to 30 million. Experimental data presented to show the effect of varying wall porosity include airfoil surface pressures, airfoil forces, and wind tunnel wall pressures. Utilizing the experimental results, an assessment of the applicability of current subcritical theoretical methods to predict wall interference corrections in subsonic and transonic flows is made.
Design and Experimental Results for a Natural-Laminar-Flow Airfoil for General Aviation Applications
NASA Technical Reports Server (NTRS)
Somers, D. M.
1981-01-01
A natural-laminar-flow airfoil for general aviation applications, the NLF(1)-0416, was designed and analyzed theoretically and verified experimentally in the Langley Low-Turbulence Pressure Tunnel. The basic objective of combining the high maximum lift of the NASA low-speed airfoils with the low cruise drag of the NACA 6-series airfoils was achieved. The safety requirement that the maximum lift coefficient not be significantly affected with transition fixed near the leading edge was also met. Comparisons of the theoretical and experimental results show excellent agreement. Comparisons with other airfoils, both laminar flow and turbulent flow, confirm the achievement of the basic objective.
Wind-tunnel test results of airfoil modifications for the EA-6B
NASA Technical Reports Server (NTRS)
Sewall, W. G.; Mcghee, R. J.; Ferris, J. C.
1987-01-01
Wind-tunnel tests have been conducted (to determine the effects on airfoil performance for several airfoil modifications) for the EA-6B Wing Improvement Program. The modifications consist of contour changes to the leading-edge slat and trailing-edge flap to provide a higher low-speed maximum lift with no high-speed cruise-drag penalty. Airfoil sections from the 28- and 76-percent span stations were selected as baseline shapes with the major testing devoted to the inboard airfoil section (28-percent span station). The airfoil modifications increased the low-speed maximum lift coefficient between 20 and 35 percent over test conditions of 3 to 14 million chord Reynolds number and 0.14 to 0.34 Mach number. At the high-speed test conditions of 0.4 to 0.80 Mach number and 10 million chord Reynolds number, the modified airfoils had either matched or had lower drag coefficients for all normal-force coefficients above 0.2 as compared to the baseline airfoil. At normal-force coefficients less than 0.2, the baseline (original) airfoil had lower drag coefficients than any of the modified airfoils.
Airfoil lance apparatus for homogeneous humidification and sorbent dispersion in a gas stream
Myers, Robert B.; Yagiela, Anthony S.
1990-12-25
An apparatus for spraying an atomized mixture into a gas stream comprises a stream line airfoil member having a large radius leading edge and a small radius trailing edge. A nozzle assembly pierces the trailing edge of the airfoil member and is concentrically surrounded by a nacelle which directs shielding gas from the interior of the airfoil member around the nozzle assembly. Flowable medium to be atomized and atomizing gas for atomizing the medium are supplied in concentric conduits to the nozzle. A plurality of nozzles each surrounded by a nacelle are spaced along the trailing edge of the airfoil member.
Airfoil lance apparatus for homogeneous humidification and sorbent dispersion in a gas stream
Myers, R.B.; Yagiela, A.S.
1990-12-25
An apparatus for spraying an atomized mixture into a gas stream comprises a stream line airfoil member having a large radius leading edge and a small radius trailing edge. A nozzle assembly pierces the trailing edge of the airfoil member and is concentrically surrounded by a nacelle which directs shielding gas from the interior of the airfoil member around the nozzle assembly. Flowable medium to be atomized and atomizing gas for atomizing the medium are supplied in concentric conduits to the nozzle. A plurality of nozzles each surrounded by a nacelle are spaced along the trailing edge of the airfoil member. 3 figs.
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.
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.
Unsteady separation process and vorticity balance on unsteady airfoils
NASA Technical Reports Server (NTRS)
Ho, Chih-Ming; Gursul, Ismet; Shih, Chiang; Lin, Hank
1992-01-01
Low momentum fluid erupts at the unsteady separation region and forms a local shear layer at the viscous-inviscid interface. At the shear layer, the vorticity lumps into a vortex and protrudes into the inviscid region. This process initiates the separation process. The response of airfoils in unsteady free stream was investigated based on this vortex generation and convection concept. This approach enabled us to understand the complicated unsteady aerodynamics from a fundamental point of view.
25. "GAFFTC 19 OCT 60, BLAST EFFECTS ON AIRFOILS, STATIC ...
25. "G-AFFTC 19 OCT 60, BLAST EFFECTS ON AIRFOILS, STATIC RUN 5." View of track rail mounting and the water brake trough at 20,000-foot track. Looking northeast. File no. 12,358-60. - Edwards Air Force Base, South Base Sled Track, Edwards Air Force Base, North of Avenue B, between 100th & 140th Streets East, Lancaster, Los Angeles County, CA
Turbine blade having a constant thickness airfoil skin
Marra, John J
2012-10-23
A turbine blade is provided for a gas turbine comprising: a support structure comprising a base defining a root of the blade and a framework extending radially outwardly from the base, and an outer skin coupled to the support structure framework. The skin has a generally constant thickness along substantially the entire radial extent thereof. The framework and the skin define an airfoil of the blade.
Effects of Compressibility on the Characteristics of Five Airfoils
1947-04-25
f -• 1. Stack, John, Lindney, W. F., and LiU.el, Bobert E.: The Compreaaibility Burble tnd tne. Effect of CcmpreuclMlity on Preaaurea and...e, no. 9, J’Oy l^’il, pp. 337-356. 11. Stack, John: Teata of Airfoils Pejjgue-? to Itolny -.he Co.Tipre?i3lbility Burble . NACA T!J No. 5/6
Effects of enviromentally imposed roughness on airfoil performance
NASA Technical Reports Server (NTRS)
Cebeci, Tuncer
1987-01-01
The experimental evidence for the effects of rain, insects, and ice on airfoil performance are examined. The extent to which the available information can be incorporated in a calculation method in terms of change of shape and surface roughness is discussed. The methods described are based on the interactive boundary layer procedure of Cebeci or on the thin layer Navier Stokes procedure developed at NASA. Cases presented show that extensive flow separation occurs on the rough surfaces.
Numerical Grid Generation and Potential Airfoil Analysis and Design
1988-01-01
contents include a discussion of grid generation concepts and schemes in the literature , iter- ative methods for numerical grid generation and two...to multicomponent rN airfoils, arbitrary multiple bodies and cascade grid generation for the first time. We present representative results as a... literature , iter- ative methods for numerical grid generation and two differential grid gen- eration schemes: (1) an elliptic, and (2) a parabolic scheme. A
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.
The S411, S412, and S413 Airfoils
2010-08-01
not as low as at point A, unlike the polars of many laminar-flow airfoils where the drag coefficient within the laminar bucket is nearly constant...in a leading edge that produces a suction peak at higher lift coefficients, which ensures that tran- sition on the upper surface will occur very near...This concept allows a wide low-drag range to be achieved and increases the loading in the leading-edge region. The forward loading serves to balance
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.
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.
Laser Velocimetry Measurements of Oscillating Airfoil Dynamic Stall Flow Field
1991-06-01
Velocimetry Measurements of Oscillating Airfoil Dynamic Stall Flow Field By M.S.Chandrasekharal Navy-NASA Joint Institute of Aeronautics and Fluid Mechanics ...tunnel of the Fluid Mechanics Laboratory(FML) angle information. The other could be used for the at NASA Ames Research Center (ARC). It is one of...were on throat is always kept choked so that no disturbances a different traverse mechanism , but this was driven as can propagate upstream into the
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.
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.
Design and Experimental Results for the S415 Airfoil
2010-08-01
polars of many laminar-flow airfoils where the drag coefficient within the laminar bucket is nearly constant. (See, for example, ref. 8.) This... suction peak at higher lift coefficients, which ensures that transition on the upper surface will occur very near the leading edge. Thus, the...pressure distribution should look like sketch 3. Sketch 3 No suction peak exists at the leading edge. Instead, a moderately adverse pressure
Design and Experimental Results for the S411 Airfoil
2010-08-01
unlike the polars of many laminar-flow airfoils where the drag coefficient within the laminar bucket is nearly constant. (See, for example, ref. 8...produces a suction peak at higher lift coefficients, which ensures that tran- sition on the upper surface will occur very near the leading edge. Thus...like sketch 3. Sketch 3 No suction peak exists at the leading edge. Instead, a rounded peak occurs aft of the leading edge, which allows some laminar
Unsteady Flow Separation and Attachment Induced by Pitching Airfoils
1983-01-13
dynamic stall. The oc- er may remain totally laminar prior :o seoaration , curronce and severity of dyvnamic stall is directly and this, by definition...in **Professor, iep.Artrent of *..rospace Lnziner- pitch direction of the airfoil were achieved v’ine in Scicnces. ,mal I os, ill.it ion anrles ( .-5...Variables for Hotwire Study The creat ion and f.te of the highly synchro- nized, strcn; vor:ccit’-.,-rc carefully charictrrizzed Feynulds Reduced Mean
Leading-Edge "Pop-Up" Spoiler For Airfoil
NASA Technical Reports Server (NTRS)
Wilson, John C.; Lance, Michael B.
1991-01-01
New concept places spoiler in leading edge of airfoil, hinged along its trailing edge, so airflow helps to deploy it and force it against mechanical stop. Deployed "pop-up" spoiler quickly eliminates almost all aerodynamic lift of stabilator. Designed to be added to leading edge of existing stabilator, without major rework. Though initial application to be on helicopter stabilators, equally applicable to wings or winglike components.
Multi-element airfoil viscous-inviscid interactions
NASA Technical Reports Server (NTRS)
Gross, L. W.
1979-01-01
Subsonic viscous-inviscid interactions for multi-element airfoils are predicted by iterating between inviscid and viscous solutions until the performance coefficients converge. Inviscid flow is modelled by using distributed source-vortex singularities on configuration surface panels. Viscous effects are calculated by an existing laminar separation bubble model and a NASA-Lockheed boundary layer-wake method. Numerical formulations and example calculations are presented.
Design and Experimental Results for the S414 Airfoil
2010-08-01
Design Integration. AIAA Paper 88-4416, Sept. 1988. 2. Harris , Franklin D.: Rotary Wing Aerodynamics – Historical Perspective and Important Issues...PENNSYLVANIA STATE UNIVERSITY AUGUST 2010 ABSTRACT A 14.22-percent-thick, slotted, natural-laminar-flow (SNLF) airfoil, the S414, intended for rotorcraft ...is a major contributor to the total vehicle drag for most rotorcraft . In general, to maximize rotor lift-to-drag ratio for low-speed flight, the
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.
Effects of environmentally imposed roughness on airfoil performance
NASA Technical Reports Server (NTRS)
Cebeci, Tuncer
1987-01-01
The experimental evidence for the effects of rain, insects, and ice on airfoil performance are examined. The extent to which the available information can be incorporated in a calculation method in terms of change of shape and surface roughness is discussed. The methods described are based on the interactive boundary procedure of Cebeci or on the thin layer Navier Stokes procedure developed at NASA. Cases presented show that extensive flow separation occurs on the rough surfaces.
Macro-Fiber Composite actuated simply supported thin airfoils
NASA Astrophysics Data System (ADS)
Bilgen, Onur; Kochersberger, Kevin B.; Inman, Daniel J.; Ohanian, Osgar J., III
2010-05-01
A piezoceramic composite actuator known as Macro-Fiber Composite (MFC) is used for actuation in the design of a variable camber airfoil intended for a ducted fan aircraft. The study focuses on response characterization under aerodynamic loads for circular arc airfoils with variable pinned boundary conditions. A parametric study of fluid-structure interaction is employed to find pin locations along the chordwise direction that result in high lift generation. Wind tunnel experiments are conducted on a 1.0% thick, 127 mm chord MFC actuated bimorph airfoil that is simply supported at 5% and 50% of the chord. Aerodynamic and structural performance results are presented for a flow rate of 15 m s - 1 and a Reynolds number of 127 000. Non-linear effects due to aerodynamic and piezoceramic hysteresis are identified and discussed. A lift coefficient change of 1.46 is observed, purely due to voltage actuation. A maximum 2D L/D ratio of 17.8 is recorded through voltage excitation.
Performance Trades Study for Robust Airfoil Shape Optimization
NASA Technical Reports Server (NTRS)
Li, Wu; Padula, Sharon
2003-01-01
From time to time, existing aircraft need to be redesigned for new missions with modified operating conditions such as required lift or cruise speed. This research is motivated by the needs of conceptual and preliminary design teams for smooth airfoil shapes that are similar to the baseline design but have improved drag performance over a range of flight conditions. The proposed modified profile optimization method (MPOM) modifies a large number of design variables to search for nonintuitive performance improvements, while avoiding off-design performance degradation. Given a good initial design, the MPOM generates fairly smooth airfoils that are better than the baseline without making drastic shape changes. Moreover, the MPOM allows users to gain valuable information by exploring performance trades over various design conditions. Four simulation cases of airfoil optimization in transonic viscous ow are included to demonstrate the usefulness of the MPOM as a performance trades study tool. Simulation results are obtained by solving fully turbulent Navier-Stokes equations and the corresponding discrete adjoint equations using an unstructured grid computational fluid dynamics code FUN2D.
Wind turbine airfoil investigations in customized turbulent inflow
NASA Astrophysics Data System (ADS)
Heisselmann, Hendrik; Peinke, Joachim; Hoelling, Michael
2016-11-01
Experimental airfoil characterizations are usually performed in laminar or unsteady periodical flows. Neither of these matches the flow conditions of natural atmospheric flows as experienced by wind turbine blades. In the presented experimental study, an active grid is used to generate turbulent inflow with customized properties, like reduced frequencies or inflow angles. This is used not only to tune flow properties, but also to mimic time series of measured atmospheric wind speeds and inflow angles in the wind tunnel. Experiments were performed on a wind turbine dedicated DU 00-W-212 airfoil to obtain highly resolved force data and chord-wise pressure distributions at Re=500,000 and Re=900,000. Additional to a laminar baseline case, unsteady sinusoidal inflow fluctuations were applied as well as three different turbulent inflows with comparable turbulence intensity, but different inflow angle fluctuations to grasp the impact of inflow characteristics on the airfoil performance. In comparison with the laminar inflow case, the lift peak of the polar is shifted to higher angles of attack in the turbulent flows. While the laminar lift polars show a rather sudden transition to stall, a softer transition with an extended stall region is found for all turbulent cases. The presented work was performed within the project AVATAR and is funded from the European Unions Seventh Program for research, technological development and demonstration under Grand Agreement No FP7-ENERGY-2013-1/n 608396.
Flow control at low Reynolds numbers using periodic airfoil morphing
NASA Astrophysics Data System (ADS)
Jones, Gareth; Santer, Matthew; Papadakis, George; Bouremel, Yann; Debiasi, Marco; Imperial-NUS Joint PhD Collaboration
2014-11-01
The performance of airfoils operating at low Reynolds numbers is known to suffer from flow separation even at low angles of attack as a result of their boundary layers remaining laminar. The lack of mixing---a characteristic of turbulent boundary layers---leaves laminar boundary layers with insufficient energy to overcome the adverse pressure gradient that occurs in the pressure recovery region. This study looks at periodic surface morphing as an active flow control technique for airfoils in such a flight regime. It was discovered that at sufficiently high frequencies an oscillating surface is capable of not only reducing the size of the separated region---and consequently significantly reducing drag whilst simultaneously increasing lift---but it is also capable of delaying stall and as a result increasing CLmax. Furthermore, by bonding Macro Fiber Composite actuators (MFCs) to the underside of an airfoil skin and driving them with a sinusoidal frequency, it is shown that this control technique can be practically implemented in a lightweight, energy efficient way. Imperial-NUS Joint Ph.D. Programme.
Heat transfer measurements from a smooth NACA 0012 airfoil
NASA Technical Reports Server (NTRS)
Poinsatte, Philip E.; Van Fossen, G. J.; Newton, James E.; De Witt, Kenneth J.
1991-01-01
Local convective heat transfer coefficients were measured from a smooth NACA 0012 airfoil having a chord length of 0.533 m. Flight data were taken for the smooth airfoil at Reynolds numbers based on chord in the range 1.24 to 2.50 million and at various angles of attack up to 4 deg. During these flight tests, the freestream velocity turbulence intensity was found to be very low. Wind tunnel data were acquired in the Reynolds number range 1.20 to 4.52 million and at angles of attack from -4 to +8 deg. The turbulence intensity in the IRT was 0.5-0.7 percent with the cloud-generating sprays off. A direct comparison between the results obtained in flight and in the IRT showed that the higher level of turbulence intensity in the IRT had little effect on the heat transfer for the lower Reynolds numbers but caused a moderate increase in heat transfer at the higher Reynolds numbers. Turning on the cloud-generating spray nozzle atomizing air in the IRT did not alter the heat transfer. The present data were compared with leading-edge cylinder and flat plate heat transfer correlations that are often used to estimate airfoil heat transfer in computer codes.
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.
Calculation of real-gas effects on airfoil aerodynamic characteristics
NASA Technical Reports Server (NTRS)
Park, Chul; Yoon, Seokkwan
1990-01-01
The effects of high temperature thermochemical phenomena on the aerodynamic characteristics at hypersonic speeds are calculated for two-dimensional airfoils in air. The calculations are performed on an airfoil similar to that used for the Space Shuttle Orbiter, and ellipses of thickness ratios varying between 5 and 15 percent. For the airfoil, one flight condition is considered. For the ellipses, the calculations are carried out over a range of chord lengths, flight velocities, flight altitudes, and angles of attack. It is shown that the lift and drag coefficients are consistently reduced by the thermochemical phenomena, and that the behavior can be represented by a specific heat ratio value less than 1.4. The center of pressure shifts forward due to the thermochemical phenomena, but its extent is sensitively affected by the geometry and angle of attack and cannot be represented by a fixed specific heat ratio. The calculated results are in qualitative agreement with the data obtained during the entry flights of the Space Shuttle vehicle.
Prediction of Film Cooling on Gas Turbine Airfoils
NASA Technical Reports Server (NTRS)
Garg, Vijay K.; Gaugler, Raymond E.
1994-01-01
A three-dimensional Navier-Stokes analysis tool has been developed in order to study the effect of film cooling on the flow and heat transfer characteristics of actual turbine airfoils. An existing code (Arnone et al., 1991) has been modified for the purpose. The code is an explicit, multigrid, cell-centered, finite volume code with an algebraic turbulence model. Eigenvalue scaled artificial dissipation and variable-coefficient implicit residual smoothing are used with a full-multigrid technique. Moreover, Mayle's transition criterion (Mayle, 1991) is used. The effects of film cooling have been incorporated into the code in the form of appropriate boundary conditions at the hole locations on the airfoil surface. Each hole exit is represented by several control volumes, thus providing an ability to study the effect of hole shape on the film-cooling characteristics. Comparison is fair with near mid-span experimental data for four and nine rows of cooling holes, five on the shower head, and two rows each on the pressure and suction surfaces. The computations, however, show a strong spanwise variation of the heat transfer coefficient on the airfoil surface, specially with shower-head cooling.
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.
Aerodynamic Control of a Pitching Airfoil by Distributed Bleed Actuation
NASA Astrophysics Data System (ADS)
Kearney, John; Glezer, Ari
2013-11-01
The aerodynamic forces and moments on a dynamically pitching 2-D airfoil model are controlled in wind tunnel experiments using distributed active bleed. Bleed flow on the suction surface downstream of the leading edge is driven by pressure differences across the airfoil and is regulated by low-power louver actuators. The bleed interacts with cross flows to effect time-dependent variations of the vorticity flux and thereby alters the local flow attachment, resulting in significant changes in pre- and post-stall lift and pitching moment (over 50% increase in baseline post-stall lift). The flow field over the airfoil is measured using high-speed (2000 fps) PIV, resolving the dynamics and characteristic time-scales of production and advection of vorticity concentrations that are associated with transient variations in the aerodynamic forces and moments. In particular, it is shown that the actuation improves the lift hysteresis and pitch stability during the oscillatory pitching by altering the evolution of the dynamic stall vortex and the ensuing flow attachment during the downstroke. Supported by the Rotorcraft Center (VLRCOE) at Georgia Tech.
CFD Study of NACA 0018 Airfoil with Flow Control
NASA Technical Reports Server (NTRS)
Eggert, Christopher A.; Rumsey, Christopher L.
2017-01-01
The abilities of two different Reynolds-Averaged Navier-Stokes codes to predict the effects of an active flow control device are evaluated. The flow control device consists of a blowing slot located on the upper surface of an NACA 0018 airfoil, near the leading edge. A second blowing slot present on the airfoil near mid-chord is not evaluated here. Experimental results from a wind tunnel test show that a slot blowing with high momentum coefficient will increase the lift of the airfoil (compared to no blowing) and delay flow separation. A slot with low momentum coefficient will decrease the lift and induce separation even at low angles of attack. Two codes, CFL3D and FUN3D, are used in two-dimensional computations along with several different turbulence models. Two of these produced reasonable results for this flow, when run fully turbulent. A more advanced transition model failed to predict reasonable results, but warrants further study using different inputs. Including inviscid upper and lower tunnel walls in the simulations was found to be important in obtaining pressure distributions and lift coefficients that best matched experimental data. A limited number of three-dimensional computations were also performed.
Low-Reynolds number compressible flow around a triangular airfoil
NASA Astrophysics Data System (ADS)
Munday, Phillip; Taira, Kunihiko; Suwa, Tetsuya; Numata, Daiju; Asai, Keisuke
2013-11-01
We report on the combined numerical and experimental effort to analyze the nonlinear aerodynamics of a triangular airfoil in low-Reynolds number compressible flow that is representative of wings on future Martian air vehicles. The flow field around this airfoil is examined for a wide range of angles of attack and Mach numbers with three-dimensional direct numerical simulations at Re = 3000 . Companion experiments are conducted in a unique Martian wind tunnel that is placed in a vacuum chamber to simulate the Martian atmosphere. Computational findings are compared with pressure sensitive paint and direct force measurements and are found to be in agreement. The separated flow from the leading edge is found to form a large leading-edge vortex that sits directly above the apex of the airfoil and provides enhanced lift at post stall angles of attack. For higher subsonic flows, the vortical structures elongate in the streamwise direction resulting in reduced lift enhancement. We also observe that the onset of spanwise instability for higher angles of attack is delayed at lower Mach numbers. Currently at Mitsubishi Heavy Industries, Ltd., Nagasaki.
1949-06-01
airfoil and boundary-laysr investigations * ACR-L-^5 «—• —r *-»«. te, JM -ta> toUojtai< D> o_ ».s. Mi Zrzvr.-.zls division, T-2 AR’I:, Wr> v.t r...s, guide vanes, etc., is Xi/D for the section. Ids Number range ts a reasonable lift at which berty to meet to which the by the results s
Aerodynamics Investigation of Faceted Airfoils at Low Reynolds Number
NASA Astrophysics Data System (ADS)
Napolillo, Zachary G.
The desire and demand to fly farther and faster has progressively integrated the concept of optimization with airfoil design, resulting in increasingly complex numerical tools pursuing efficiency often at diminishing returns; while the costs and difficulty associated with fabrication increases with design complexity. Such efficiencies may often be necessary due to the power density limitations of certain aircraft such as small unmanned aerial vehicles (UAVs) and micro air vehicles (MAVs). This research, however, focuses on reducing the complexity of airfoils for applications where aerodynamic performance is less important than the efficiency of manufacturing; in this case a Hybrid Projectile. By employing faceted sections to approximate traditional contoured wing sections it may be possible to expedite manufacturing and reduce costs. We applied this method to the development of a low Reynolds number, disposable Hybrid Projectile requiring a 4.5:1 glide ratio, resulting in a series of airfoils which are geometric approximations to highly contoured cross-sections called ShopFoils. This series of airfoils both numerically and experimentally perform within a 10% margin of the SD6060 airfoil at low Re. Additionally, flow visualization has been conducted to qualitatively determine what mechanisms, if any, are responsible for the similarity in performance between the faceted ShopFoil sections and the SD6060. The data obtained by these experiments did not conclusively reveal how the faceted surfaces may influence low Re flow but did indicate that the ShopFoil s did not maintain flow attachment at higher angles of attack than the SD6060. Two reasons are provided for the unexpected performance of the ShopFoil: one is related to downwash effects, which are suspected of placing the outer portion of the span at an effective angle of attack where the ShopFoils outperform the SD6060; the other is the influence of the tip vortex on separation near the wing tips, which possibly
NASA Technical Reports Server (NTRS)
Spaid, F. W.; Dahlin, J. A.; Roos, F. W.; Stivers, L. S., Jr.
1983-01-01
Surface static-pressure and drag data obtained from tests of two slightly modified versions of the original NASA Whitcomb airfoil and a model of the NACA 0012 airfoil section are presented. Data for the supercritical airfoil were obtained for a free-stream Mach number range of 0.5 to 0.9, and a chord Reynolds number range of 2 x 10 to the 6th power to 4 x 10 to the 6th power. The NACA 0012 airfoil was tested at a constant chord Reynolds number of 2 x 10 to the 6th power and a free-stream Mach number range of 0.6 to 0.8.
Thick airfoil designs for the root of the 10MW INNWIND.EU wind turbine
NASA Astrophysics Data System (ADS)
Mu≁oz, A.; Méndez, B.; Munduate, X.
2016-09-01
The main objective of the “INNWIND.EU” project is to investigate and demonstrate innovative designs for 10-20MW offshore wind turbines and their key components, such as lightweight rotors. In this context, the present paper describes the development of two new airfoils for the blade root region. From the structural point of view, the root is the region in charge of transmitting all the loads of the blade to the hub. Thus, it is very important to include airfoils with adequate structural properties in this region. The present article makes use of high-thickness and blunt trailing edge airfoils to improve the structural characteristics of the airfoils used to build this blade region. CENER's (National Renewable Energy Center of Spain) airfoil design tool uses the airfoil software XFOIL to compute the aerodynamic characteristics of the designed airfoils. That software is based on panel methods which show some problems with the calculation of airfoils with thickness bigger than 35% and with blunt trailing edge. This drawback has been overcome with the development of an empirical correction for XFOIL lift and drag prediction based on airfoil experiments. From the aerodynamic point of view, thick airfoils are known to be very sensitive to surface contamination or turbulent inflow conditions. Consequently, the design optimization takes into account the aerodynamic torque in both clean and contaminated conditions. Two airfoils have been designed aiming to improve the structural and the aerodynamic behaviour of the blade in clean and contaminated conditions. This improvement has been corroborated with Blade Element Momentum (BEM) computations.
NASA Astrophysics Data System (ADS)
Chougule, Prasad; Nielsen, Søren R. K.
2014-06-01
Nowadays, small vertical axis wind turbines are receiving more attention due to their suitability in micro-electricity generation. There are few vertical axis wind turbine designs with good power curve. However, the efficiency of power extraction has not been improved. Therefore, an attempt has been made to utilize high lift technology for vertical axis wind turbines in order to improve power efficiency. High lift is obtained by double-element airfoil mainly used in aeroplane wing design. In this current work a low Reynolds number airfoil is selected to design a double-element airfoil blade for use in vertical axis wind turbine to improve the power efficiency. Double-element airfoil blade design consists of a main airfoil and a slat airfoil. Orientation of slat airfoil is a parameter of investigation in this paper and air flow simulation over double-element airfoil. With primary wind tunnel test an orientation parameter for the slat airfoil is initially obtained. Further a computational fluid dynamics (CFD) has been used to obtain the aerodynamic characteristics of double-element airfoil. The CFD simulations were carried out using ANSYS CFX software. It is observed that there is an increase in the lift coefficient by 26% for single-element airfoil at analysed conditions. The CFD simulation results were validated with wind tunnel tests. It is also observe that by selecting proper airfoil configuration and blade sizes an increase in lift coefficient can further be achieved.
Self-sustained Flow-acoustic Interactions in Airfoil Transitional Boundary Layers
2015-07-09
the free-jet anechoic wind tunnel of Ecole Centrale de Lyon. Measurements of wall pressure, far -field acoustic pressure and velocity fluctuations...airfoil and the slightly cambered SD7003 airfoil at moderate incidence were tested in the free-jet anechoic wind tunnel of Ecole Centrale de Lyon. A
NASA Technical Reports Server (NTRS)
Goradia, S. H.; Lilley, D. E.
1975-01-01
Theoretical and experimental studies are described which were conducted for the purpose of developing a new generalized method for the prediction of profile drag of single component airfoil sections with sharp trailing edges. This method aims at solution for the flow in the wake from the airfoil trailing edge to the large distance in the downstream direction; the profile drag of the given airfoil section can then easily be obtained from the momentum balance once the shape of velocity profile at a large distance from the airfoil trailing edge has been computed. Computer program subroutines have been developed for the computation of the profile drag and flow in the airfoil wake on CDC6600 computer. The required inputs to the computer program consist of free stream conditions and the characteristics of the boundary layers at the airfoil trailing edge or at the point of incipient separation in the neighborhood of airfoil trailing edge. The method described is quite generalized and hence can be extended to the solution of the profile drag for multi-component airfoil sections.
NASA Technical Reports Server (NTRS)
Harris, C. D.
1974-01-01
A lifting airfoil theoretically designed for shockless supercritical flow utilizing a complex hodograph method has been evaluated in the Langley 8-foot transonic pressure tunnel at design and off-design conditions. The experimental results are presented and compared with those of an experimentally designed supercritical airfoil which were obtained in the same tunnel.
1980-09-01
developed a complex hodograph solution satisfying the two-dimensional potential equation for supercritical flows over is- olated airfoils. By using...this hodograph technique, an isolated semi- infinite displacement body containing the airfoil, plus boundary layer and wake displacement thickness, could
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.
Study of laminar boundary layer instability noise study on a controlled diffusion airfoil
NASA Astrophysics Data System (ADS)
Jaiswal, Prateek; Sanjose, Marlene; Moreau, Stephane
2016-11-01
Detailed experimental study has been carried out on a Controlled Diffusion (CD) airfoil at 5° angle of attack and at chord based Reynolds number of 1 . 5 ×105 . All the measurements were done in an open-jet anechoic wind tunnel. The airfoil mock-up is held between two side plates, to keep the flow two-dimensional. PIV measurements have been performed in the wake and on the boundary layer of the airfoil. Pressure sensor probes on the airfoil were used to detect mean airfoil loading and remote microphone probes were used to measure unsteady pressure fluctuations on the surface of the airfoil. Furthermore the far field acoustic pressure was measured using an 1/2 inch ICP microphone. The results confirm very later transition of a laminar boundary layer to a turbulent boundary layer on the suction side of the airfoil. The process of transition of laminar to turbulent boundary layer comprises of turbulent reattachment of a separated shear layer. The pressure side of the boundary layer is found to be laminar and stable. Therefore tonal noise generated is attributed to events on suction side of the airfoil. The flow transition and emission of tones are further investigated in detail thanks to the complementary DNS study.
An analytic study of nonsteady two-phase laminar boundary layer around an airfoil
NASA Technical Reports Server (NTRS)
Hsu, Yu-Kao
1989-01-01
Recently, NASA, FAA, and other organizations have focused their attention upon the possible effects of rain on airfoil performance. Rhode carried out early experiments and concluded that the rain impacting the aircraft increased the drag. Bergrum made numerical calculation for the rain effects on airfoils. Luers and Haines did an analytic investigation and found that heavy rain induces severe aerodynamic penalties including both a momentum penalty due to the impact of the rain and a drag and lift penalty due to rain roughening of the airfoil and fuselage. More recently, Hansman and Barsotti performed experiments and declared that performance degradation of an airfoil in heavy rain is due to the effective roughening of the surface by the water layer. Hansman and Craig did further experimental research at low Reynolds number. E. Dunham made a critical review for the potential influence of rain on airfoil performance. Dunham et al. carried out experiments for the transport type airfoil and concluded that there is a reduction of maximum lift capability with increase in drag. There is a scarcity of published literature in analytic research of two-phase boundary layer around an airfoil. Analytic research is being improved. The following assumptions are made: the fluid flow is non-steady, viscous, and incompressible; the airfoil is represented by a two-dimensional flat plate; and there is only a laminar boundary layer throughout the flow region. The boundary layer approximation is solved and discussed.
Aerodynamic Characteristics of SC1095 and SC1094 R8 Airfoils
2003-12-01
Development, and Engineering Command Ames Research Center Moffett Field, California December 2003 National Aeronautics and Space Administration Ames...60A ROTOR BLADE AND AIRFOILS ................................................................................... 2 EVALUATION OF SECTION CHARACTERISTICS...Characteristics of SC1095 and SC1094 R8 Airfoils WILLIAM G. BOUSMAN Aeroflightdynamics Directorate U.S. Army Research, Development, and Engineering Command Ames
Airfoil gust response and the sound produced by airifoil-vortex interaction
NASA Technical Reports Server (NTRS)
Amiet, R. K.
1986-01-01
This paper contributes to the understanding of the noise generation process of an airfoil encountering an unsteady upwash. By using a fast Fourier transform together with accurate airfoil response functions, the lift-time waveform for an airfoil encountering a delta function gust (the indicial function) is calculated for a flat plate airfoil in a compressible flow. This shows the interesting property that the lift is constant until the generated acoustic wave reaches the trailing edge. Expressions are given for the magnitude of this constant and for the pressure distribution on the airfoil during this time interval. The case of an airfoil cutting through a line vortex is also analyzed. The pressure-time waveform in the far field is closely related to the left-time waveform for the above problem of an airfoil entering a delta function gust. The effects of varying the relevant parameters in the problem are studied, including the observed position, the core diameter of the vortex, the vortex orientation and the airfoil span. The far field sound varies significantly with observer position, illustrating the importance of non-compactness effects. Increasing the viscous core diameter tends to smooth the pressure-time waveform. For small viscous core radius and infinite span, changing the vortex orientation changes only the amplitude of the pressure-time waveform, and not the shape.
Main results of CAST-10 airfoil tested in T2 cryogenic wind tunnel
NASA Technical Reports Server (NTRS)
Blanchard, A.; Seraudie, A.; Breil, J. F.
1989-01-01
The aim of this cooperative NASA/DFVLR/ONERA project was to examine the performance of the CAST-10 airfoil in the T2 cryogenic wind tunnel. Tests included general characteristics of the CAST-10 airfoil and fundamental studies on Reynold number effects. Good T2 cryogenic operation was observed. Improvements should be done for moisture elimination and for side wall boundary layer effects.
A nonlinear theory for airfoils with trailing-edge jet flap
NASA Technical Reports Server (NTRS)
Barger, R. L.
1976-01-01
A nonlinear procedure for computing the pressure distribution on an airfoil with a trailing edge jet flap is described. The method is not restricted to thin airfoils or shallow jet-deflection angles. Correlation with experiment indicates that the characteristics of the pressure distribution are predicted by the theory, but the effect of entrainment is overpredicted with the entrainment coefficient used.
Profile Optimization Method for Robust Airfoil Shape Optimization in Viscous Flow
NASA Technical Reports Server (NTRS)
Li, Wu
2003-01-01
Simulation results obtained by using FUN2D for robust airfoil shape optimization in transonic viscous flow are included to show the potential of the profile optimization method for generating fairly smooth optimal airfoils with no off-design performance degradation.
Numerical Solution of Compressible Steady Flows around the RAE 2822 Airfoil
NASA Astrophysics Data System (ADS)
Kryštůfek, P.; Kozel, K.
2014-03-01
The article presents results of a numerical solution of subsonic, transonic and supersonic flows described by the system of Navier-Stokes equations in 2D laminar compressible flows around the RAE 2822 airfoil. Authors used FVM multistage Runge-Kutta method to numerically solve the flows around the RAE 2822 airfoil.