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.
On the Design of Lifting Airfoils with High Critical Mach Number Using Full Potential Theory
NASA Astrophysics Data System (ADS)
Kropinski, M. C. A.
We wish to construct airfoils that have the highest free-stream Mach number for a given set of geometric constraints for which the flow is nowhere supersonic. Nonlifting airfoils that maximize the critical Mach number for a given cross-sectional area are known to possess long sonic segments at their critical speed. To construct lifting airfoils, we proceed under the conjecture that an airfoil with a high value of has the longest possible arc length of sonic velocity over its upper and lower surface. In Kropinski etal. (1995) the lifting problem was tackled in transonic small-disturbance theory. In this paper we numerically construct lifting airfoils with high using the full potential theory and we show that these airfoils have significantly higher than some standard airfoils. We also construct airfoils with higher values of the lift coefficient, by relaxing the speed constraint on the lower surface of the airfoil to have a value less than sonic.
P. B. S. Lissaman; R. E. Wilson; R. W. Thresher; S. N. Walker
1979-01-01
The effect of high lift airfoils on the cost effectiveness of HAWT and VAWT (horizontal and vertical axis wind turbine) machines is studied. The scope involved first studying modern two dimensional airfoils, and developing a generalized formulation for their performance in terms of lift, drag, and thickness at appropriate Reynolds numbers. Single element, multi-element, symmetrical, extra thick airfoils and jet
Hodograph design of lifting airfoils with high critical mach numbers
NASA Astrophysics Data System (ADS)
Kropinski, M. C. A.; Schwendeman, D. W.; Cole, J. D.
1995-05-01
We wish to construct airfoils that have the highest free-stream Mach number M ? for a given set of geometric constraints for which the flow is nowhere supersonic. Nonlifting airfoils which maximize M ? for a given thickness ratio ? are known to possess long sonic segments at their critical speed. To construct lifting airfoils, we proceed under the conjecture that the optimal airfoil satisfying a given set of constraints is the one possessing the longest possible arc length of sonic velocity. A boundary-value problem is formulated in the hodograph plane using transonic small-disturbance theory whose solution determines an airfoil with long sonic arcs. For small lift coefficients, the hodograph domain covers two Riemann sheets and a finite-difference method is used to solve the boundary-value problem on this domain. A numerical integration of the solution around the boundary yields an airfoil shape, and three examples are discussed. The performance of these airfoils is compared with standard airfoils having the same lift coefficient and ?, and it is shown that the calculated airfoils have a 6% 10% increase in critical M ?.
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.
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.
Stephan Barth
2005-01-01
We present initial wind tunnel measurements which investigate the dynamic stall effect as it is caused by fluctuations of the wind direction in turbulent wind. In order to quantify this effect, the lift of an FX79-W-151A airfoil is determined by the integral of pressure distribution at the wind tunnel walls while rotating the airfoil with defined angular velocity. The rotation
An interactive boundary-layer approach to multielement airfoils at high lift
NASA Technical Reports Server (NTRS)
Cebeci, Tuncer
1992-01-01
A calculation method based on an interactive boundary-layer approach to multielement airfoils is described and is applied to three types of airfoil configurations with and without flap-wells in order to demonstrate the applicability of the method to general high-lift configurations. This method, well tested for single airfoils as a function of shape, angle of attack, and Reynolds number, is here shown to apply equally well to two-element airfoils and their wakes, to a flap-well region, and to a three-element arrangement which includes the effects of co-flowing regions, a flap well, and the wake of the elements. In addition to providing accurate representation of these flows, the method is general so that its extension to three-dimensional arrangements is likely to provide a practical, accurate and efficient tool to assist the design process.
Separation Control from the Flap of a High-Lift Airfoil Using DBD Plasma Actuators
NASA Astrophysics Data System (ADS)
Little, Jesse; Nishihara, Munetake; Adamovich, Igor; Samimy, Mo
2008-11-01
Control of separation from the flap of a high-lift airfoil using a single dielectric barrier discharge (DBD) plasma actuator has been investigated experimentally. This project is motivated by the desire to replace existing multi-element flap configurations with a single simple flap to allow more efficient high-lift generation. The results show that a single DBD plasma actuator located at the flap shoulder can increase or reduce the size of the time-averaged separation bubble over the flap depending on the frequency of actuation. In the latter case, the lift on the airfoil is increased due to improved circulation around the model, but it does not result in full reattachment on the deflected flap. These findings are consistent with previous research on high-lift airfoil configurations. The work will be expanded by exploring the effect of multiple actuators as well as their geometry and location on the size and structure of the separated region over the flap. This portion of the work will be done with an emphasis on optimizing the relative phase of each actuator and its effect on the separated flow region.
Effect of wakes from moving upstream rods on boundary layer separation from a high lift airfoil
NASA Astrophysics Data System (ADS)
Volino, Ralph J.
2011-11-01
Highly loaded airfoils in turbines allow power generation using fewer airfoils. High loading, however, can cause boundary layer separation, resulting in reduced lift and increased aerodynamic loss. Separation is affected by the interaction between rotating blades and stationary vanes. Wakes from upstream vanes periodically impinge on downstream blades, and can reduce separation. The wakes include elevated turbulence, which can induce transition, and a velocity deficit, which results in an impinging flow on the blade surface known as a ``negative jet.'' In the present study, flow through a linear cascade of very high lift airfoils is studied experimentally. Wakes are produced with moving rods which cut through the flow upstream of the airfoils, simulating the effect of upstream vanes. Pressure and velocity fields are documented. Wake spacing and velocity are varied. At low Reynolds numbers without wakes, the boundary layer separates and does not reattach. At high wake passing frequencies separation is largely suppressed. At lower frequencies, ensemble averaged velocity results show intermittent separation and reattachment during the wake passing cycle. Supported by NASA.
Jesse Little; Munetake Nishihara; Igor Adamovich; Mo Samimy
2010-01-01
Control of flow separation from the deflected flap of a high-lift airfoil up to Reynolds numbers of 240,000 (15 m\\/s) is explored\\u000a using a single dielectric barrier discharge (DBD) plasma actuator near the flap shoulder. Results show that the plasma discharge\\u000a can increase or reduce the size of the time-averaged separated region over the flap depending on the frequency of actuation.
Advanced natural laminar flow airfoil with high lift to drag ratio
NASA Technical Reports Server (NTRS)
Viken, Jeffrey K.; Pfenninger, Werner; Mcghee, Robert J.
1986-01-01
An experimental verification of a high performance natural laminar flow (NLF) airfoil for low speed and high Reynolds number applications was completed in the Langley Low Turbulence Pressure Tunnel (LTPT). Theoretical development allowed for the achievement of 0.70 chord laminar flow on both surfaces by the use of accelerated flow as long as tunnel turbulence did not cause upstream movement of transition with increasing chord Reynolds number. With such a rearward pressure recovery, a concave type deceleration was implemented. Two-dimensional theoretical analysis indicated that a minimum profile drag coefficient of 0.0026 was possible with the desired laminar flow at the design condition. With the three-foot chord two-dimensional model constructed for the LTPT experiment, a minimum profile drag coefficient of 0.0027 was measured at c sub l = 0.41 and Re sub c = 10 x 10 to the 6th power. The low drag bucket was shifted over a considerably large c sub l range by the use of the 12.5 percent chord trailing edge flap. A two-dimensional lift to drag ratio (L/D) was 245. Surprisingly high c sub l max values were obtained for an airfoil of this type. A 0.20 chort split flap with 60 deg deflection was also implemented to verify the airfoil's lift capabilities. A maximum lift coefficient of 2.70 was attained at Reynolds numbers of 3 and 6 million.
P. B. S. Lissaman; R. E. Wilson; R. W. Thresher; S. N. Walker
1979-01-01
The purpose of this project was to study the effect of high lift airfoils on the cost-effectiveness of HAWT and VAWT (horizontal and vertical axis wind turbine) machines. The scope involved first studying modern two-dimensional airfoils, and developing a generalized formulation for their performance in terms of lift, drag, and thickness at appropriate Reynolds numbers. Single element, multi-element, symmetrical, extra
NASA Technical Reports Server (NTRS)
Omar, E.; Zierten, T.; Hahn, M.; Szpiro, E.; Mahal, A.
1977-01-01
Three high lift systems for a 9.3 percent blunt based, supercritical airfoil were designed, fabricated, and wind tunnel tested. A method for calculating the viscous flow about two dimensional multicomponent airfoils was evaluated by comparing its predictions with test data. A comparison of high lift systems derived from supercritical airfoils with high lift systems derived from conventional airfoils is presented. The high lift systems for the supercritical airfoil were designed to achieve maximum lift and consisted of: (1) a single slotted flap, (2) a double slotted flap and a leading edge slat, and (3) a triple slotted flap and a leading edge slat. Aerodynamic force and moment data and surface pressure data are presented for all configurations and boundary layer and wake profiles for the single slotted flap configuration. The wind-tunnel models, test facilities and instrumentation, and data reduction are described.
NASA Technical Reports Server (NTRS)
Viken, Jeffrey K.; Watson-Viken, Sally A.; Pfenninger, Werner; Morgan, Harry L., Jr.; Campbell, Richard L.
1987-01-01
The design and testing of Natural Laminar Flow (NLF) airfoils is examined. The NLF airfoil was designed for low speed, having a low profile drag at high chord Reynolds numbers. The success of the low speed NLF airfoil sparked interest in a high speed NLF airfoil applied to a single engine business jet with an unswept wing. Work was also conducted on the two dimensional flap design. The airfoil was decambered by removing the aft loading, however, high design Mach numbers are possible by increasing the aft loading and reducing the camber overall on the airfoil. This approach would also allow for flatter acceleration regions which are more stabilizing for cross flow disturbances. Sweep could then be used to increase the design Mach number to a higher value also. There would be some degradation of high lift by decambering the airfoil overall, and this aspect would have to be considered in a final design.
Experimental Results for a Flapped Natural-laminar-flow Airfoil with High Lift/drag Ratio
NASA Technical Reports Server (NTRS)
Mcghee, R. J.; Viken, J. K.; Pfenninger, W.; Beasley, W. D.; Harvey, W. D.
1984-01-01
Experimental results have been obtained for a flapped natural-laminar-flow airfoil, NLF(1)-0414F, in the Langley Low-Turbulence Pressure Tunnel. The tests were conducted over a Mach number range from 0.05 to 0.40 and a chord Reynolds number range from about 3.0 x 10(6) to 22.0 x 10(6). The airfoil was designed for 0.70 chord laminar flow on both surfaces at a lift coefficient of 0.40, a Reynolds number of 10.0 x 10(6), and a Mach number of 0.40. A 0.125 chord simple flap was incorporated in the design to increase the low-drag, lift-coefficient range. Results were also obtained for a 0.20 chord split-flap deflected 60 deg.
An Experimntal Investigation of the 30P30N Multi-Element High-Lift Airfoil
NASA Technical Reports Server (NTRS)
Pascioni, Kyle A.; Cattafesta, Louis N.; Choudhari, Meelan M.
2014-01-01
High-lift devices often generate an unsteady flow field producing both broadband and tonal noise which radiates from the aircraft. In particular, the leading edge slat is often a dominant contributor to the noise signature. An experimental study of a simplified unswept high-lift configuration, the 30P30N, has been conducted to understand and identify the various flow-induced noise sources around the slat. Closed-wall wind tunnel tests are performed in the Florida State Aeroacoustic Tunnel (FSAT) to characterize the slat cove flow field using a combination of surface and off-body measurements. Mean surface pressures compare well with numerical predictions for the free-air configuration. Consistent with previous measurements and computations for 2D high-lift configurations, the frequency spectra of unsteady surface pressures on the slat surface display several narrowband peaks that decrease in strength as the angle of attack is increased. At positive angles of attack, there are four prominent peaks. The three higher frequency peaks correspond, approximately, to a harmonic sequence related to a feedback resonance involving unstable disturbances in the slat cove shear layer. The Strouhal numbers associated with these three peaks are nearly insensitive to the range of flow speeds (41-58 m/s) and the angles of attack tested (3-8.5 degrees). The first narrow-band peak has an order of magnitude lower frequency than the remaining peaks and displays noticeable sensitivity to the angle of attack. Stereoscopic particle image velocimetry (SPIV) measurements provide supplementary information about the shear layer characteristics and turbulence statistics that may be used for validating numerical simulations.
Airfoil optimization for unsteady flows with application to high-lift noise reduction
NASA Astrophysics Data System (ADS)
Rumpfkeil, Markus Peer
The use of steady-state aerodynamic optimization methods in the computational fluid dynamic (CFD) community is fairly well established. In particular, the use of adjoint methods has proven to be very beneficial because their cost is independent of the number of design variables. The application of numerical optimization to airframe-generated noise, however, has not received as much attention, but with the significant quieting of modern engines, airframe noise now competes with engine noise. Optimal control techniques for unsteady flows are needed in order to be able to reduce airframe-generated noise. In this thesis, a general framework is formulated to calculate the gradient of a cost function in a nonlinear unsteady flow environment via the discrete adjoint method. The unsteady optimization algorithm developed in this work utilizes a Newton-Krylov approach since the gradient-based optimizer uses the quasi-Newton method BFGS, Newton's method is applied to the nonlinear flow problem, GMRES is used to solve the resulting linear problem inexactly, and last but not least the linear adjoint problem is solved using Bi-CGSTAB. The flow is governed by the unsteady two-dimensional compressible Navier-Stokes equations in conjunction with a one-equation turbulence model, which are discretized using structured grids and a finite difference approach. The effectiveness of the unsteady optimization algorithm is demonstrated by applying it to several problems of interest including shocktubes, pulses in converging-diverging nozzles, rotating cylinders, transonic buffeting, and an unsteady trailing-edge flow. In order to address radiated far-field noise, an acoustic wave propagation program based on the Ffowcs Williams and Hawkings (FW-H) formulation is implemented and validated. The general framework is then used to derive the adjoint equations for a novel hybrid URANS/FW-H optimization algorithm in order to be able to optimize the shape of airfoils based on their calculated far-field pressure fluctuations. Validation and application results for this novel hybrid URANS/FW-H optimization algorithm show that it is possible to optimize the shape of an airfoil in an unsteady flow environment to minimize its radiated far-field noise while maintaining good aerodynamic performance.
Airfoil design: Finding the balance between design lift and structural stiffness
NASA Astrophysics Data System (ADS)
Bak, Christian; Gaudern, Nicholas; Zahle, Frederik; Vronsky, Tomas
2014-06-01
When upscaling wind turbine blades there is an increasing need for high levels of structural efficiency. In this paper the relationships between the aerodynamic characteristics; design lift and lift-drag ratio; and the structural characteristics were investigated. Using a unified optimization setup, airfoils were designed with relative thicknesses between 18% and 36%, a structural box height of 85% of the relative thickness, and varying box widths in chordwise direction between 20% and 40% of the chord length. The results from these airfoil designs showed that for a given flapwise stiffness, the design lift coefficient increases if the box length reduces and at the same time the relative thickness increases. Even though the conclusions are specific to the airfoil design approach used, the study indicated that an increased design lift required slightly higher relative thickness compared to airfoils with lower design lift to maintain the flapwise stiffness. Also, the study indicated that the lift-drag ratio as a function of flapwise stiffness was relatively independent of the airfoil design with a tendency that the lift-drag ratio decreased for large box lengths. The above conclusions were supported by an analysis of the three airfoil families Riso-C2, DU and FFA, where the lift-drag ratio as a function of flapwise stiffness was decreasing, but relatively independent of the airfoil design, and the design lift coefficient was varying depending on the design philosophy. To make the analysis complete also design lift and lift- drag ratio as a function of edgewise and torsional stiffness were shown.
The effect of sound on the lift of an airfoil
Kitowski, John Victor
1962-01-01
THE EFFECT OF SOUND ON THE LIFT OF AN AIRFOIL By John Victor Kitowski A Thesis Submitted to the Graduate School of the Agricultural and Mechanical College of Texas in partial fulfillment of the requirements for the degree of MASTER... OF SCIENCE May, 1962 Qepartment of Aeronautical Engineering Major Subject: Aeronautical Engineering THE EFFECT OF SOUND ON THE LIFT OF AN AIRFOIL A Thesis By John Victor Kitowski Approved as to style and content by: Chairman of Committee Head...
Impact of Airfoils on Aerodynamic Optimization of Heavy Lift Rotorcraft
C. W. Acree; Preston B. Martin; Ethan A. Romander
Rotor airfoils were developed for two large tiltrotor designs, the Large Civil Tilt Rotor (LCTR) and the Military Heavy Tilt Rotor (MHTR). The LCTR was the most promising of several rotorcraft concepts produced by the NASA Heavy Lift Rotorcraft Systems Investigation. It was designed to carry 120 passengers for 1200 nm, with performance of 350 knots cruise at 30,000 ft
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.
Optimization of an Advanced Design Three-Element Airfoil at High Reynolds Numbers
NASA Technical Reports Server (NTRS)
Lin, John C.; Dominik, Chet J.
1995-01-01
New high-lift components have been designed for a three-element advanced high-lift research airfoil using a state-of-the-art computational method. The new components were designed with the aim to provide high maximum-lift values while maintaining attached flow on the single-segment flap at approach conditions. This three-element airfoil has been tested in the NASA Langley Low-Turbulence Pressure Tunnel at chord Reynolds number up to 16 million. The performance of the NASA research airfoil is compared to a reference advanced high-lift research airfoil. Effects of Reynolds number on slat and flap rigging have been studied experimentally. The performance trend of this new high-lift design is comparable to that predicted by the computational method over much of the angle of attack range. Nevertheless, the method did not accurately predict the airfoil performance or the configuration-based trends near maximum lift.
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.
Aeroacoustic interaction of a distributed vortex with a lifting Joukowski airfoil
J. C. Hardin; S. L. Lamkin
1984-01-01
A first principles computational aeroacoustics calculation of the flow and noise fields produced by the interaction of a distributed vortex with a lifting Joukowski airfoil is accomplished at the Reynolds number of 200. The case considered is that where the circulations of the vortex and the airfoil are of opposite sign, corresponding to blade vortex interaction on the retreating side
An application of active surface heating for augmenting lift and reducing drag of an airfoil
NASA Technical Reports Server (NTRS)
Maestrello, Lucio; Badavi, Forooz F.; Noonan, Kevin W.
1988-01-01
Application of active control to separated flow on the RC(6)-08 airfoil at high angle of attack by localized surface heating is numerically simulated by integrating the compressible 2-D nonlinear Navier-Stokes equation solver. Active control is simulated by local modification of the temperature boundary condition over a narrow strip of the upper surface of the airfoil. Both mean and perturbed profiles are favorably altered when excited with the same natural frequency of the shear layer by moderate surface heating for both laminar and turbulent separation. The shear layer is found to be very sensitive to localized surface heating in the vicinity of the separation point. The excitation field at the surface sufficiently altered both the local as well as the global circulation to cause a significant increase in lift and reduction in drag.
The effect of sound on the lift of an airfoil
Kitowski, John Victor
1962-01-01
was unique in that it was the first to be conducted in a quantitative manner. The effect of sound on the wakes of flat plates, double wedge airfoils with parallel midbodies, a NACA 23012 airfoil, and a NACA 63006 laminar flow airfoil was examined. Quinn...
Lift enhancement and flow structure of airfoil with joint trailing-edge flap and Gurney flap
NASA Astrophysics Data System (ADS)
Lee, T.; Su, Y. Y.
2011-06-01
The impact of Gurney flaps (GF), of different heights and perforations, on the aerodynamic and wake characteristics of a NACA 0015 airfoil equipped with a trailing-edge flap (TEF) was investigated experimentally at Re = 2.54 × 105. The addition of the Gurney flap to the TEF produced a further increase in the downward turning of the mean flow (increased aft camber), leading to a significant increase in the lift, drag, and pitching moment compared to that produced by independently deployed TEF or GF. The maximum lift increased with flap height, with the maximum lift-enhancement effectiveness exhibited at the smallest flap height. The near wake behind the joint TEF and GF became wider and had a larger velocity deficit and fluctuations compared to independent GF and TEF deployment. The Gurney flap perforation had only a minor impact on the wake and aerodynamics characteristics compared to TEF with a solid GF. The rapid rise in lift generation of the joint TEF and GF application, compared to conventional TEF deployment, could provide an improved off-design high-lift device during landing and takeoff.
NASA Technical Reports Server (NTRS)
Nixon, D.
1978-01-01
The linear transonic perturbation integral equation previously derived for nonlifting airfoils is formulated for lifting cases. In order to treat shock wave motions, a strained coordinate system is used in which the shock location is invariant. The tangency boundary conditions are either formulated using the thin airfoil approximation or by using the analytic continuation concept. A direct numerical solution to this equation is derived in contrast to the iterative scheme initially used, and results of both lifting and nonlifting examples indicate that the method is satisfactory.
NASA Technical Reports Server (NTRS)
Page, F Handley
1926-01-01
The results that are described in this article form a complete series of tests on an airfoil fitted with front and rear slots, the rear slot being formed between the portion of the wing aft of the rear spar and the forward portion of the flap.
Clegg, J.D.
1988-03-01
This paper summarizes the major considerations in the selection, design, installation, operation, or repair of high-rate artificial-lift systems. The major types of artificial lift - sucker-rod pumps, gas-lift systems, electrical submersible pumps, hydraulic pumps and jets, and hydraulic turbine-driven pumps - will be discussed. An extensive bibliography of artificial-lift papers is included.
Aeroacoustic interaction of a distributed vortex with a lifting Joukowski airfoil
NASA Technical Reports Server (NTRS)
Hardin, J. C.; Lamkin, S. L.
1984-01-01
A first principles computational aeroacoustics calculation of the flow and noise fields produced by the interaction of a distributed vortex with a lifting Joukowski airfoil is accomplished at the Reynolds number of 200. The case considered is that where the circulations of the vortex and the airfoil are of opposite sign, corresponding to blade vortex interaction on the retreating side of a single helicopter rotor. The results show that the flow is unsteady, even in the absence of the incoming vortex, resulting in trailing edge noise generation. After the vortex is input, it initially experiences a quite rapid apparent diffusion rate produced by stretching in the airfoil velocity gradients. Consideration of the effects of finite vortex size and viscosity causes the noise radiation during the encounter to be much less impulsive than predicted by previous analyses.
The interdependence of profile drag and lift with Joukowski type and related airfoils
NASA Technical Reports Server (NTRS)
Muttray, H
1935-01-01
On the basis of a systematic investigation of Gottingen wind-tunnel data on Joukowski type and related airfoils, it is shown in what manner the profile drag coefficient is dependent on the lift coefficient. The individual factors for the construction of the profile drag polars are given. They afford a more accurate calculation of the performance coefficients of airplane designs than otherwise attainable with the conventional assumption of constant drag coefficient.
Experimental Study of Lift-Enhancing Tabs on a Two-Element Airfoil
NASA Technical Reports Server (NTRS)
Storms, Bruce L.; Ross, James C.
1995-01-01
The results of a wind-tunnel test are presented for a two-dimensional NASA 63(sub 2)-215 Mod B airfoil with a 30% chord single-slotted flap. The use of lift-enhancing tabs (similar to Gurney flaps) on the lower surface near the trailing edge of both elements was investigated on four nap configurations. A combination of vortex generators on the flap and lift-enhancing tabs was also investigated. Measurements of surface-pressure distributions and wake profiles were used to determine the aerodynamic performance of each configuration. By reducing flow separation on the flap, a lift-enhancing tab at the main-element trailing edge increased the maximum lift by 10.3% for the 42-deg flap case. The tab had a lesser effect at a moderate flap deflection (32 deg) and adversely affected the performance at the smallest flap deflection (22 deg). A tab located near the flap trailing edge produced an additional lift increment for all flap deflections. The application of vortex generators to the flap eliminated lift-curve hysteresis and reduced flow separation on two configurations with large flap deflections (greater than 40 deg). A maximum-lift coefficient of 3.32 (17% above the optimum baseline) was achieved with the combination of lift-enhancing tabs on both elements and vortex generators on the flap.
Low-speed aerodynamic characteristics of an airfoil optimized for maximum lift coefficient
NASA Technical Reports Server (NTRS)
Bingham, G. J.; Chen, A. W.
1972-01-01
An investigation has been conducted in the Langley low-turbulence pressure tunnel to determine the two-dimensional characteristics of an airfoil optimized for maximum lift coefficient. The design maximum lift coefficient was 2.1 at a Reynolds number of 9.7 million. The airfoil with a smooth surface and with surface roughness was tested at angles of attack from 6 deg to 26 deg, Reynolds numbers (based on airfoil chord) from 2.0 million to 12.9 million, and Mach numbers from 0.10 to 0.35. The experimental results are compared with values predicted by theory. The experimental pressure distributions observed at angles of attack up to at least 12 deg were similar to the theoretical values except for a slight increase in the experimental upper-surface pressure coefficients forward of 26 percent chord and a more severe gradient just behind the minimum-pressure-coefficient location. The maximum lift coefficients were measured with the model surface smooth and, depending on test conditions, varied from 1.5 to 1.6 whereas the design value was 2.1.
A quantitative study of the lift-enhancing flow field generated by an airfoil with a Gurney flap
NASA Astrophysics Data System (ADS)
Troolin, Daniel Ryan
2009-12-01
Although it is well known that a Gurney flap affects the lift, drag, and pressure distribution along an airfoil, the mechanisms behind the changes are still not well understood. The following research seeks to understand what is causing the effects of a Gurney flap through quantitative measurements of the spatial and temporal flow details, including force balance measurements, hotwire anemometry (HWA), high resolution particle image velocimetry (HRPIV), and time resolved particle image velocimetry (TRPIV). TRPIV is used to broaden the understanding of the interaction between the various vortex shedding modes which are elicited from the regions upstream and downstream of the flap. The HWA technique is useful for its very high frequency response, and is used in the wake of the airfoil in order to gain valuable insight into the nature of the vortex shedding frequencies. Vortices generated both upstream and downstream of the Gurney flap have been observed, and the vortex interactions, which occur due to the non-phase-locked nature of the shedding modes, are analyzed. The results are interpreted in terms of the known lift increment.
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)
A quantitative study of the lift-enhancing flow field generated by an airfoil with a Gurney flap
Daniel Ryan Troolin
2009-01-01
Although it is well known that a Gurney flap affects the lift, drag, and pressure distribution along an airfoil, the mechanisms behind the changes are still not well understood. The following research seeks to understand what is causing the effects of a Gurney flap through quantitative measurements of the spatial and temporal flow details, including force balance measurements, hotwire anemometry
Development of high-lift laminar wing using steady active flow control
NASA Astrophysics Data System (ADS)
Clayton, Patrick J.
Fuel costs represent a large fraction of aircraft operating costs. Increased aircraft fuel efficiency is thus desirable. Laminar airfoils have the advantage of reduced cruise drag and increased fuel efficiency. Unfortunately, they cannot perform adequately during high-lift situations (i.e. takeoff and landing) due to low stall angles and low maximum lift caused by flow separation. Active flow control has shown the ability to prevent or mitigate separation effects, and increase maximum lift. This fact makes AFC technology a fitting solution for improving high-lift systems and reducing the need for slats and flap elements. This study focused on experimentally investigating the effects of steady active flow control from three slots, located at 1%, 10%, and 80% chord, respectively, over a laminar airfoil with 45 degree deflected flap. A 30-inch-span airfoil model was designed, fabricated, and then tested in the Bill James 2.5'x3' Wind Tunnel at Iowa State University. Pressure data were collected along the mid-span of the airfoil, and lift and drag were calculated. Five test cases with varying injection locations and varying C? were chosen: baseline, blown flap, leading edge blowing, equal blowing, and unequal blowing. Of these cases, unequal blowing achieved the greatest lift enhancement over the baseline. All cases were able to increase lift; however, gains were less than anticipated.
NASA Technical Reports Server (NTRS)
Graham, Donald J; Nitzberg, Gerald E; Olson, Robert N
1945-01-01
Pressure distributions determined from high-speed wind-tunnel tests are presented for five NACA airfoil sections representative of both low-drag and conventional types. Section characteristics of lift, drag, and quarter-chord pitching moment are presented along with the measured pressure distributions for the NACA 65sub2-215 (a=0.5), 66sub2-215 (a=0.6), 0015, 23015, and 4415 airfoils for Mach numbers up to approximately 0.85. A critical study is made of the airfoil pressure distributions in an attempt to formulate a set of general criteria for defining the character of high speed flows over typical airfoil shapes. Comparisons are made of the relative characteristics of the low-drag and conventional airfoils investigated insofar as they would influence the high-speed performance and the high-speed stability and control characteristics of airplanes employing these wing sections.
NASA Technical Reports Server (NTRS)
Spera, David A.
2008-01-01
Equations are developed with which to calculate lift and drag coefficients along the spans of torsionally-stiff rotating airfoils of the type used in wind turbine rotors and wind tunnel fans, at angles of attack in both the unstalled and stalled aerodynamic regimes. Explicit adjustments are made for the effects of aspect ratio (length to chord width) and airfoil thickness ratio. Calculated lift and drag parameters are compared to measured parameters for 55 airfoil data sets including 585 test points. Mean deviation was found to be -0.4 percent and standard deviation was 4.8 percent. When the proposed equations were applied to the calculation of power from a stall-controlled wind turbine tested in a NASA wind tunnel, mean deviation from 54 data points was -1.3 percent and standard deviation was 4.0 percent. Pressure-rise calculations for a large wind tunnel fan deviated by 2.7 percent (mean) and 4.4 percent (standard). The assumption that a single set of lift and drag coefficient equations can represent the stalled aerodynamic behavior of a wide variety of airfoils was found to be satisfactory.
Summary of section data on trailing-edge high-lift devices
NASA Technical Reports Server (NTRS)
Cahill, Jones F
1949-01-01
A summary has been made of available data on the characteristics of airfoil sections with trailing-edge high-lift devices. Data for plain, split, and slotted flaps are collected and analyzed. The effects of each of the variables involved in the design of the various types of flap are examined and, in cases where sufficient data are given, optimum configurations are deduced. Wherever possible, the effects of airfoil section, Reynolds number, and leading-edge roughness are shown. For single and double slotted flaps, where a large amount of unrelated data are available, maximum lift coefficients of many configurations are presented in tables.
Robust Airfoil Optimization in High Resolution Design Space
NASA Technical Reports Server (NTRS)
Li, Wu; Padula, Sharon L.
2003-01-01
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 B-spline control points as design variables yet the resulting airfoil shape is fairly smooth, and (3) it allows the user to make a trade-off between the level of optimization and the amount of computing time consumed. The robust optimization method is demonstrated by solving a lift-constrained drag minimization problem for a two-dimensional airfoil in viscous flow with a large number of geometric design variables. Our experience with robust optimization indicates that our strategy produces reasonable airfoil shapes that are similar to the original airfoils, but these new shapes provide drag reduction over the specified range of Mach numbers. We have tested this strategy on a number of advanced airfoil models produced by knowledgeable aerodynamic design team members and found that our strategy produces airfoils better or equal to any designs produced by traditional design methods.
NASA Technical Reports Server (NTRS)
Schwier, W.
1947-01-01
The NACA 23012-4 airfoil was investigated for the purpose of increasing lift by means of blowing out air from the wing, in conjunction with the effect of plain flap of variable contour and slotted flap of 25 percent chord length. The wing also was provided with a hinged nose, to be deflected at will. Air was blown out frcm the wing immediately in front of the flap; also at the opening between wing and hinged nose,tangentially to the surface of the wing. Another device employed to increase maximum lift was a movable slat, to be opened to form a clot. Lift was measured in relation to the volume of blown-out air and considerable increases were observed with increasing volume.
NREL airfoil families for HAWTs
Tangler, J L [National Renewable Energy Lab., Golden, CO (United States); Somers, D M [Airfoils Inc., State College, PA (United States)
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.
Design of the LRP airfoil series using 2D CFD
NASA Astrophysics Data System (ADS)
Zahle, Frederik; Bak, Christian; Sørensen, Niels N.; Vronsky, Tomas; Gaudern, Nicholas
2014-06-01
This paper describes the design and wind tunnel testing of a high-Reynolds number, high lift airfoil series designed for wind turbines. The airfoils were designed using direct gradient- based numerical multi-point optimization based on a Bezier parameterization of the shape, coupled to the 2D Navier-Stokes flow solver EllipSys2D. The resulting airfoils, the LRP2-30 and LRP2-36, achieve both higher operational lift coefficients and higher lift to drag ratios compared to the equivalent FFA-W3 airfoils.
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.
Tangler, James L. (Boulder, CO); Somers, Dan M. (State College, PA)
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.
Non-Equilibrium Turbulence Modeling for High Lift Aerodynamics
NASA Technical Reports Server (NTRS)
Durbin, P. A.
1998-01-01
This phase is discussed in ('Non linear kappa - epsilon - upsilon(sup 2) modeling with application to high lift', Application of the kappa - epsilon -upsilon(sup 2) model to multi-component airfoils'). Further results are presented in 'Non-linear upsilon(sup 2) - f modeling with application to high-lift' The ADI solution method in the initial implementation was very slow to converge on multi-zone chimera meshes. I modified the INS implementation to use GMRES. This provided improved convergence and less need for user intervention in the solution process. There were some difficulties with implementation into the NASA compressible codes, due to their use of approximate factorization. The Helmholtz equation for f is not an evolution equation, so it is not of the form assumed by the approximate factorization method. Although The Kalitzin implementation involved a new solution algorithm ('An implementation of the upsilon(sup 2) - f model with application to transonic flows'). The algorithm involves introducing a relaxation term in the f-equation so that it can be factored. The factorization can be into a plane and a line, with GMRES used in the plane. The NASA code already evaluated coefficients in planes, so no additional memory is required except that associated the the GMRES algorithm. So the scope of this project has expanded via these interactions. . The high-lift work has dovetailed into turbine applications.
Key Topics for High-Lift Research: A Joint Wind Tunnel/Flight Test Approach
NASA Technical Reports Server (NTRS)
Fisher, David; Thomas, Flint O.; Nelson, Robert C.
1996-01-01
Future high-lift systems must achieve improved aerodynamic performance with simpler designs that involve fewer elements and reduced maintenance costs. To expeditiously achieve this, reliable CFD design tools are required. The development of useful CFD-based design tools for high lift systems requires increased attention to unresolved flow physics issues. The complex flow field over any multi-element airfoil may be broken down into certain generic component flows which are termed high-lift building block flows. In this report a broad spectrum of key flow field physics issues relevant to the design of improved high lift systems are considered. It is demonstrated that in-flight experiments utilizing the NASA Dryden Flight Test Fixture (which is essentially an instrumented ventral fin) carried on an F-15B support aircraft can provide a novel and cost effective method by which both Reynolds and Mach number effects associated with specific high lift building block flows can be investigated. These in-flight high lift building block flow experiments are most effective when performed in conjunction with coordinated ground based wind tunnel experiments in low speed facilities. For illustrative purposes three specific examples of in-flight high lift building block flow experiments capable of yielding a high payoff are described. The report concludes with a description of a joint wind tunnel/flight test approach to high lift aerodynamics research.
MEASUREMENT OF LIFT AND DRAG ON MEMS SCALE AIRFOILS IN SLIP FLOW
Michael J. Martin; Katsuo Kurabayashi; Iain D. Boyd
2001-01-01
Based on boundary layer theory, non-equilibrium effects will cause a reduction in drag on airfoils with chord lengths ranging from 10 to 100 microns. A novel facility is being designed at the University of Michigan to study these effects as a function of Knudsen number and Reynolds number. To measure the reduction in drag, and to study the effect on
FROM ANALYSIS TO DESIGN OF HIGH-LIFT CONFIGURATIONS USING A NEWTON-KRYLOV ALGORITHM
Marian Nemec; David W. Zingg
2002-01-01
An efficient multi-block Newton-Krylov algo- rithm using the compressible Navier-Stokes equations is presented for the analysis and de- sign of high-lift airfoil configurations. The pre- conditioned generalized minimum residual (GM- RES) method is applied to solve the discrete- adjoint equation, leading to a fast computation of accurate objective function gradients. Further- more, the GMRES method is used in conjunc- tion
Lift enhancement and flow structure of airfoil with joint trailing-edge flap and Gurney flap
T. Lee; Y. Y. Su
2011-01-01
The impact of Gurney flaps (GF), of different heights and perforations, on the aerodynamic and wake characteristics of a NACA\\u000a 0015 airfoil equipped with a trailing-edge flap (TEF) was investigated experimentally at Re = 2.54 × 105. The addition of the Gurney flap to the TEF produced a further increase in the downward turning of the mean flow (increased\\u000a aft camber), leading to a
Detached-Eddy Simulation of a Co-Flow Jet Airfoil at High Angle of Attack
Zha, Gecheng
Detached-Eddy Simulation of a Co-Flow Jet Airfoil at High Angle of Attack Baoyuan Wang , Ge) airfoil at high angles of attack (AoA) is conducted in this paper. A low diffusion E-CUSP (LDE) scheme leading edge and a suction slot near trailing edge on the airfoil suction surface. The slots are opened
NASA Technical Reports Server (NTRS)
Cone, Clarence D., Jr.
1961-01-01
The induced drag polar is developed for wt-ngs capable of attaining extremely high loadings while possessing an elliptical distribution of circulation. This development is accomplished through a theoretical investigation of the vortex-wake deformation process and the deduction of the airfoil forces from the impulse and kinetic energy contents of the ultimate wake form. The investigation shows that the induced velocities of the wake limit the maximum lift coefficient to a value of 1.94 times the wing aspect ratio, for aspect ratios equal to or less than 6.5, and that the section properties of the airfoil limit the lift coefficient to 12.6 for aspect ratios greater than 6.5. Relations are developed for the rate of deformation of the vortex wake. It is also shown that linear wing theory is app1icable up to lift coefficients equal to 1.1 times the aspect ratio.
NASA Technical Reports Server (NTRS)
Bristow, D. R.; Grose, G. G.
1978-01-01
The Douglas Neumann method for low-speed potential flow on arbitrary three-dimensional lifting bodies was modified by substituting the combined source and doublet surface paneling based on Green's identity for the original source panels. Numerical studies show improved accuracy and stability for thin lifting surfaces, permitting reduced panel number for high-lift devices and supercritical airfoil sections. The accuracy of flow in concave corners is improved. A method of airfoil section design for a given pressure distribution, based on Green's identity, was demonstrated. The program uses panels on the body surface with constant source strength and parabolic distribution of doublet strength, and a doublet sheet on the wake. The program is written for the CDC CYBER 175 computer. Results of calculations are presented for isolated bodies, wings, wing-body combinations, and internal flow.
Experiments on the flow field physics of confluent boundary layers for high-lift systems
NASA Technical Reports Server (NTRS)
Nelson, Robert C.; Thomas, F. O.; Chu, H. C.
1994-01-01
The use of sub-scale wind tunnel test data to predict the behavior of commercial transport high lift systems at in-flight Reynolds number is limited by the so-called 'inverse Reynolds number effect'. This involves an actual deterioration in the performance of a high lift device with increasing Reynolds number. A lack of understanding of the relevant flow field physics associated with numerous complicated viscous flow interactions that characterize flow over high-lift devices prohibits computational fluid dynamics from addressing Reynolds number effects. Clearly there is a need for research that has as its objective the clarification of the fundamental flow field physics associated with viscous effects in high lift systems. In this investigation, a detailed experimental investigation is being performed to study the interaction between the slat wake and the boundary layer on the primary airfoil which is known as a confluent boundary layer. This little-studied aspect of the multi-element airfoil problem deserves special attention due to its importance in the lift augmentation process. The goal of this research is is to provide an improved understanding of the flow physics associated with high lift generation. This process report will discuss the status of the research being conducted at the Hessert Center for Aerospace Research at the University of Notre Dame. The research is sponsored by NASA Ames Research Center under NASA grant NAG2-905. The report will include a discussion of the models that have been built or that are under construction, a description of the planned experiments, a description of a flow visualization apparatus that has been developed for generating colored smoke for confluent boundary layer studies and some preliminary measurements made using our new 3-component fiber optic LDV system.
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.
Advances in Pneumatic-Controlled High-Lift Systems Through Pulsed Blowing
NASA Technical Reports Server (NTRS)
Jones, Gregory S.; Englar, Robet J.
2003-01-01
Circulation Control technologies have been around for 65 years, and have been successfully demonstrated in laboratories and flight vehicles alike. Yet there are few production aircraft flying today that implement these advances. Circulation Control techniques may have been overlooked due to perceived unfavorable trade offs of mass flow, pitching moment, cruise drag, noise, etc. Improvements in certain aspects of Circulation Control technology are the focus of this paper. This report will describe airfoil and blown high lift concepts that also address cruise drag reduction and reductions in mass flow through the use of pulsed pneumatic blowing on a Coanda surface. Pulsed concepts demonstrate significant reductions in mass flow requirements for Circulation Control, as well as cruise drag concepts that equal or exceed conventional airfoil systems.
NASA Technical Reports Server (NTRS)
Jacobs, E.N.; Abbott, Ira H.; von Doenhoff, A.E.
1939-01-01
In order to extend the useful range of Reynolds numbers of airfoils designed to take advantage of the extensive laminar boundary layers possible in an air stream of low turbulence, tests were made of the NACA 2412-34 and 1412-34 sections in the NACA low-turbulence tunnel. Although the possible extent of the laminar boundary layer on these airfoils is not so great as for specially designed laminar-flow airfoils, it is greater than that for conventional airfoils, and is sufficiently extensive so that at Reynolds numbers above 11,000,000 the laminar region is expected to be limited by the permissible 'Reynolds number run' and not by laminar separation as is the case with conventional airfoils. Drag measurements by the wake-survey method and pressure-distribution measurements were made at several lift coefficients through a range of Reynolds numbers up to 11,400,000. The drag scale-effect curve for the NACA 1412-34 is extrapolated to a Reynolds number of 30,000,000 on the basis of theoretical calculations of the skin friction. Comparable skin-friction calculations were made for the NACA 23012. The results indicate that, for certain applications at moderate values of the Reynolds number, the NACA 1412-34 and 2412-34 airfoils offer some advantages over such conventional airfoils as the NACA 23012. The possibility of maintaining a more extensive laminar boundary layer on these airfoils should result in a small drag reduction, and the absence of pressure peaks allows higher speeds to be reached before the compressibility burble is encountered. At lower Reynold numbers, below about 10,000,000, these airfoils have higher drags than airfoils designed to operate with very extensive laminar boundary layers.
HSCT high lift system aerodynamic requirements
NASA Technical Reports Server (NTRS)
Paulson, John A.
1992-01-01
The viewgraphs and discussion of high lift system aerodynamic requirements are provided. Low speed aerodynamics has been identified as critical to the successful development of a High Speed Civil Transport (HSCT). The airplane must takeoff and land at a sufficient number of existing or projected airports to be economically viable. At the same time, community noise must be acceptable. Improvements in cruise drag, engine fuel consumption, and structural weight tend to decrease the wing size and thrust required of engines. Decreasing wing size increases the requirements for effective and efficient low speed characteristics. Current design concepts have already been compromised away from better cruise wings for low speed performance. Flap systems have been added to achieve better lift-to-drag ratios for climb and approach and for lower pitch attitudes for liftoff and touchdown. Research to achieve improvements in low speed aerodynamics needs to be focused on areas most likely to have the largest effect on the wing and engine sizing process. It would be desirable to provide enough lift to avoid sizing the airplane for field performance and to still meet the noise requirements. The airworthiness standards developed in 1971 will be the basis for performance requirements for an airplane that will not be critical to the airplane wing and engine size. The lift and drag levels that were required to meet the performance requirements of tentative airworthiness standards established in 1971 and that were important to community noise are identified. Research to improve the low speed aerodynamic characteristics of the HSCT needs to be focused in the areas of performance deficiency and where noise can be reduced. Otherwise, the wing planform, engine cycle, or other parameters for a superior cruising airplane would have to be changed.
A unified viscous theory of lift and drag of 2-D thin airfoils and 3-D thin wings
NASA Technical Reports Server (NTRS)
Yates, John E.
1991-01-01
A unified viscous theory of 2-D thin airfoils and 3-D thin wings is developed with numerical examples. The viscous theory of the load distribution is unique and tends to the classical inviscid result with Kutta condition in the high Reynolds number limit. A new theory of 2-D section induced drag is introduced with specific applications to three cases of interest: (1) constant angle of attack; (2) parabolic camber; and (3) a flapped airfoil. The first case is also extended to a profiled leading edge foil. The well-known drag due to absence of leading edge suction is derived from the viscous theory. It is independent of Reynolds number for zero thickness and varies inversely with the square root of the Reynolds number based on the leading edge radius for profiled sections. The role of turbulence in the section induced drag problem is discussed. A theory of minimum section induced drag is derived and applied. For low Reynolds number the minimum drag load tends to the constant angle of attack solution and for high Reynolds number to an approximation of the parabolic camber solution. The parabolic camber section induced drag is about 4 percent greater than the ideal minimum at high Reynolds number. Two new concepts, the viscous induced drag angle and the viscous induced separation potential are introduced. The separation potential is calculated for three 2-D cases and for a 3-D rectangular wing. The potential is calculated with input from a standard doublet lattice wing code without recourse to any boundary layer calculations. Separation is indicated in regions where it is observed experimentally. The classical induced drag is recovered in the 3-D high Reynolds number limit with an additional contribution that is Reynold number dependent. The 3-D viscous theory of minimum induced drag yields an equation for the optimal spanwise and chordwise load distribution. The design of optimal wing tip planforms and camber distributions is possible with the viscous 3-D wing theory.
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.
Trailing edge modifications for flatback airfoils.
Kahn, Daniel L. (University of California, Davis, CA); van Dam, C.P. (University of California, Davis, CA); 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.
Increasing Lift by Releasing Compressed Air on Suction Side of Airfoil
NASA Technical Reports Server (NTRS)
Seewald, F
1927-01-01
The investigation was limited chiefly to the region of high angles of attack since it is only in this region that any considerable change in the character of the flow can be expected from such artificial aids. The slot, through which compressed air was blown, was formed by two pieces of sheet steel connected by screws at intervals of about 5 cm. It was intended to regulate the width of the slot by means of these screws. Much more compressed air was required than was originally supposed, hence all the delivery pipes were much too small. This experiment, therefore, is to be regarded as only a preliminary one.
NASA Technical Reports Server (NTRS)
Eppler, R.
1979-01-01
A computer approach to the design and analysis of airfoils and some common problems concerning laminar separation bubbles at different lift coefficients are briefly discussed. Examples of application to ultralight airplanes, canards, and sailplanes with flaps are given.
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.
HSR high lift research program: Status and plans
NASA Technical Reports Server (NTRS)
Rose, Jim
1992-01-01
The viewgraphs for a report on the status and plans for High Speed Research (HSR) programs are provided. Benefits of improved high lift performance are listed. The objectives, approach, and experimental results are outlined. Future experimental tests and additional research opportunities are listed. The use of computational fluid dynamics (CFD) in high lift research is also outlined.
Status of LaRC HSCT high-lift research
NASA Technical Reports Server (NTRS)
Coe, Paul L.
1992-01-01
The viewgraphs for a status report of the NASA Langley Reseach Center High Speed Civil Transport (HSCT) High-Lift Research Program are provided. A listing of available models and previous wind tunnel studies are presented. Objectives and approach of the piloted simulation program are given. The HSCT High-Lift Research plans are listed and briefly described.
AFC-Enabled Simplified High-Lift System Integration Study
NASA Technical Reports Server (NTRS)
Hartwich, Peter M.; Dickey, Eric D.; Sclafani, Anthony J.; Camacho, Peter; Gonzales, Antonio B.; Lawson, Edward L.; Mairs, Ron Y.; Shmilovich, Arvin
2014-01-01
The primary objective of this trade study report is to explore the potential of using Active Flow Control (AFC) for achieving lighter and mechanically simpler high-lift systems for transonic commercial transport aircraft. This assessment was conducted in four steps. First, based on the Common Research Model (CRM) outer mold line (OML) definition, two high-lift concepts were developed. One concept, representative of current production-type commercial transonic transports, features leading edge slats and slotted trailing edge flaps with Fowler motion. The other CRM-based design relies on drooped leading edges and simply hinged trailing edge flaps for high-lift generation. The relative high-lift performance of these two high-lift CRM variants is established using Computational Fluid Dynamics (CFD) solutions to the Reynolds-Averaged Navier-Stokes (RANS) equations for steady flow. These CFD assessments identify the high-lift performance that needs to be recovered through AFC to have the CRM variant with the lighter and mechanically simpler high-lift system match the performance of the conventional high-lift system. Conceptual design integration studies for the AFC-enhanced high-lift systems were conducted with a NASA Environmentally Responsible Aircraft (ERA) reference configuration, the so-called ERA-0003 concept. These design trades identify AFC performance targets that need to be met to produce economically feasible ERA-0003-like concepts with lighter and mechanically simpler high-lift designs that match the performance of conventional high-lift systems. Finally, technical challenges are identified associated with the application of AFC-enabled highlift systems to modern transonic commercial transports for future technology maturation efforts.
K. Hashimoto
1985-01-01
To reduce the number of the turbine airfoils or the solidity as far as possible without increasing energy loss, a study of highly loaded turbine airfoils was conducted. These airfoils were designed for the typical velocity diagrams of the first and second stages of a jet engine low pressure turbine. With regard to the design procedures, an improved inverse method,
Tests of 16 related airfoils at high speed
NASA Technical Reports Server (NTRS)
Stack, John; Von Doenhoff, Albert E
1935-01-01
In order to provide information that might lead to the development of better propeller section, 13 related symmetrical airfoils were tested in the NACA high-speed wind tunnel for a study of the effect of thickness form on the aerodynamic characteristics. The thickness-form variables studies were the value of the maximum thickness, the position along the chord at which the maximum thickness occurs, and the value of the leading-edge radius. The tests were conducted through the low angle-of-attack range for speeds extending from 35 percent of that of sound to slightly in excess of the speed at which a compressibility burble, or breakdown of flow, occurs. The corresponding Reynolds number range is 350,000 to 750,000.
Development of the Risø wind turbine airfoils
NASA Astrophysics Data System (ADS)
Fuglsang, Peter; Bak, Christian
2004-04-01
This paper presents the wind turbine airfoil development at Risø. The design method is described together with our target characteristics for wind turbine airfoils. The use of the CFD code Ellipsys2D for prediction of final target characteristics is described together with the VELUX wind tunnel testing setup. Three airfoil families were developed; Risø-A1, Risø-P and Risø-B1. The Risø-A1 airfoil family was developed for rotors of 600 kW and larger. Wind tunnel testing and field testing showed that this airfoil family is well suited for stall and active stall control. However, sensitivity to roughness was higher than expected. Field tests of a 600 kW active stall wind turbine showed an estimated reduction in blade fatigue loading of up to 15% at the same annual energy yield and at the same time reduced blade weight and blade solidity. The Risø-P airfoils were developed to replace the Risø-A1 airfoils for use on pitch controlled wind turbines. Improved design objectives should reduce the sensitivity to roughness, but measurements are not yet available. The Risø-B1 airfoil family was developed for variable speed operation with pitch control of large megawatt sized rotors. Wind tunnel testing verified the high maximum lift for these airfoils, and the airfoils were found to be very insensitive to leading edge roughness. Performance with vortex generators and Gurney flaps in combination was found to be attractive for the blade root part. Field testing of a 1.5 MW rotor is in progress. Copyright
Gurney flap—Lift enhancement, mechanisms and applications
J. J. Wang; Y. C. Li; K.-S. Choi
2008-01-01
Since its invention by a race car driver Dan Gurney in 1960s, the Gurney flap has been used to enhance the aerodynamics performance of subsonic and supercritical airfoils, high-lift devices and delta wings. In order to take stock of recent research and development of Gurney flap, we have carried out a review of the characteristics and mechanisms of lift enhancement
NASA Technical Reports Server (NTRS)
Somers, D. M.
1981-01-01
A flapped natural laminar flow airfoil for general aviation applications, the NLF(1)-0215F, has been 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 has been achieved. The safety requirement that the maximum lift coefficient not be significantly affected with transition fixed near the leading edge has also been met. Comparisons of the theoretical and experimental results show generally good agreement.
Natural laminar flow airfoil design considerations for winglets on low-speed airplanes
NASA Technical Reports Server (NTRS)
Vandam, C. P.
1984-01-01
Winglet airfoil section characteristics which significantly influence cruise performance and handling qualities of an airplane are discussed. A good winglet design requires an airfoil section with a low cruise drag coefficient, a high maximum lift coefficient, and a gradual and steady movement of the boundary layer transition location with angle of attack. The first design requirement provides a low crossover lift coefficient of airplane drag polars with winglets off and on. The other requirements prevent nonlinear changes in airplane lateral/directional stability and control characteristics. These requirements are considered in the design of a natural laminar flow airfoil section for winglet applications and chord Reynolds number of 1 to 4 million.
Impulsive Start of a Symmetric Airfoil at High Angle of Attack
NASA Technical Reports Server (NTRS)
Katz, Joseph; Yon, Steven; Rogers, Stuart E.
1996-01-01
The fluid dynamic phenomena following the impulsive start of a NACA 0015 airfoil were studied by using a time accurate solution of the incompressible laminar Navier-Stokes equations. Angle of attack was set at 10 deg to simulate steady-state poststall conditions at a Reynolds number of 1.2 x 10(exp 4). The calculation revealed that large initial lift values can be obtained, immediately following the impulsive start, when a trapped vortex develops above the airfoil. Before the buildup of this trapped vortex and immediately after the airfoil was set into motion, the fluid is attached to the airfoil's surface and flows around the trailing edge, demonstrating the delay in the buildup of the classical Kutta condition. The transient of this effect is quite short and is followed by an attached How event that leads to the trapped vortex that has a longer duration. The just described initial phenomenon eventually transits into a fully developed separated flow pattern identifiable by an alternating, periodic vortex shedding.
Experimental and Numerical Investigations of a High Performance Co-Flow Jet Airfoil
Danah Kirk
2009-01-01
The work reflected in this thesis includes a detailed study of co-flow jet (CFJ) technologies as they are applied to a typical thin airfoil, NACA 6415, at take-off and landing speeds. Numerical analysis and experimental testing were conducted on baseline and co-flow jet airfoils of the same plan form. The CFJ mechanism employs high pressure air injected along the span
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 coefficient ranges for which the original forms had been designed before further extending the investigation.
Design of low Reynolds number airfoils. I
NASA Technical Reports Server (NTRS)
Pfenninger, W.; Vemuru, C. S.
1988-01-01
The low Reynolds number airfoils designated ASM-LRN-003 and -007 have been designed for high section L/D ratios using Drela's (1985) design-and-analysis code; close to 70-percent laminar flow is maintained on the upper surfaces, and 100-percent on the lower, at coefficients of lift of 1.0-1.3, assuming optimum laminar separation and transition control on the upper surface by means of suitable turbulators. If peak performance is critical, airfoils of this type with an undercut front lower surface and a correspondingly sharper leading edge may be resorted to.
Root region airfoil for wind turbine
Tangler, James L. (Boulder, CO); Somers, Dan M. (State College, PA)
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.
Tangler, James L. (Boulder, CO); Somers, Dan M. (State College, PA)
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.
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.
Wind Tunnel Tests of Wind Turbine Airfoils at High Reynolds Numbers
NASA Astrophysics Data System (ADS)
Llorente, E.; Gorostidi, A.; Jacobs, M.; Timmer, W. A.; Munduate, X.; Pires, O.
2014-06-01
Wind tunnel tests have been performed to measure the two-dimensional aerodynamic characteristics of two different airfoil families at high Reynolds numbers (from 3 to 12 millions) in the DNW High Pressure Wind Tunnel in Gottingen (HDG), Germany. Also, tests at a Reynolds number of 3 millions have been performed in the Low-Speed Low- Turbulence Wind Tunnel of Delft University, The Netherlands. The airfoils tested belong to two wind turbine dedicated families: the TU-Delft DU family and the ACCIONA Windpower AWA family that was designed in collaboration with CENER. Reynolds number effects on airfoil performance have been obtained in the range of 3 to 12 millions. The availability of data from two different wind tunnels has brought the opportunity to cross compare the results from the two facilities.
The effects of sound on the boundary layer of an airfoil at high angles of attack
Hutchinson, Thomas Ira
1963-01-01
THE EFFECTS OF SOUND ON THE BOUNDARY LAYER OF AN AIRFOIL AT HIGH ANGLES OF ATTACK A Thesis By THOMAS IRA HUTCHINSON S, rhr?tted to the Graduate School of the Agrtcultu, al and Mechanical College of Texas in partial f, . lfrllr:, cnt... of the requirements for the degree of MASTER OF SCIENCE January 1963 Mai"" Subject: Acrcspace Engrneering THE EFFECTS OF SOUND ON THE BOUNDARY LAYER OF AN AIRFOIL AT HIGH ANGLES OF ATTACK A Thesis THOMAS IRA HUTCHINSON Approved as to style and content by...
ADVANCED HIGH LIFT CFD ANALYSIS AND DESIGN TOOLS
R. Rudnik; S. Wallin
2004-01-01
This contribution highlights development, validation, and application of European CFD codes for high lift aerodynamics. Major contributions to the status of CFD have been achieved in EC funded European projects, such as EUROLIFT and HiAer. Concerning the flow field analysis, progress in numerical methods as well as computer resources resulted in the ability of treating complete 3D transport aircraft high
An Experimental Investigation of the Confluent Boundary Layer on a High-Lift System
NASA Technical Reports Server (NTRS)
Thomas, F. O.; Nelson, R. C.
1997-01-01
This paper describes a fundamental experimental investigation of the confluent boundary layer generated by the interaction of a leading-edge slat wake with the boundary layer on the main element of a multi-element airfoil model. The slat and airfoil model geometry are both fully two-dimensional. The research reported in this paper is performed in an attempt to investigate the flow physics of confluent boundary layers and to build an archival data base on the interaction of the slat wake and the main element wall layer. In addition, an attempt is made to clearly identify the role that slat wake / airfoil boundary layer confluence has on lift production and how this occurs. Although complete LDV flow surveys were performed for a variety of slat gap and overhang settings, in this report the focus is on two cases representing both strong and weak wake boundary layer confluence.
NASA Technical Reports Server (NTRS)
Becker, John V.
1940-01-01
Determinations of boundary-layer transition on the NACA 0012 and 2301 airfoils were made in the 8-foot high-speed wind tunnel over a range of Reynolds Numbers from 1,600,000 to 16,800,000. The results are of particular significance as compared with flight tests and tests in wind tunnels of appreciable turbulence because of the extremely low turbulence in the high-speed tunnel. A comparison of the results obtained on NACA 0012 airfoils of 2-foot and 5-foot chord at the same Reynolds Number permitted an evaluation of the effect of compressibility on transition. The local skin friction along the surface of the NACA 0012 airfoil was measured at a Reynolds Number of 10,000,000. For all the lift coefficient at which tests were made, transition occurred in the region of estimated laminar separation at the low Reynolds Numbers and approach the point of minimum static pressure as a forward limit at the high Reynolds Numbers. The effect of compressibility on transition was slight. None of the usual parameters describing the local conditions in the boundary layer near the transition point served as an index for locating the transition point. As a consequence of the lower turbulence in the 8-foot high-speed tunnel, the transition points occurred consistently farther back along the chord than those measured in the NACA full-scale tunnel. An empirical relation for estimating the location of the transition point for conventional airfoils on the basis of static-pressure distribution and Reynolds Number is presented.
Noise exposure reduction of advanced high-lift systems
NASA Technical Reports Server (NTRS)
Haffner, Stephen W.
1995-01-01
The purpose of NASA Contract NAS1-20090 Task 3 was to investigate the potential for noise reduction that would result from improving the high-lift performance of conventional subsonic transports. The study showed that an increase in lift-to-drag ratio of 15 percent would reduce certification noise levels by about 2 EPNdB on approach, 1.5 EPNdB on cutback, and zero EPNdB on sideline. In most cases, noise contour areas would be reduced by 10 to 20 percent.
Lift outs: how to acquire a high-functioning team.
Groysberg, Boris; Abrahams, Robin
2006-12-01
More and more, expanding companies are hiring high-functioning groups of people who have been working together effectively within one company and can rapidly come up to speed in a new environment. These lifted-out teams don't need to get acquainted with one another or to establish shared values, mutual accountability, or group norms; their long-standing relationships and trust help them make an impact very quickly. Of course, the process is not without risks: A failed lift out can lead to loss of money, opportunity, credibility, and even native talent. Boris Groysberg and Robin Abrahams studied more than 40 high-profile moves and interviewed team leaders in multiple industries and countries to examine the risks and opportunities that lift outs present. They concluded that, regardless of industry, nationality, or size of the team, a successful lift out unfolds over four consecutive, interdependent stages that must be meticulously managed. In the courtship stage, the hiring company and the leader of the targeted team determine whether the proposed move is, in fact, a good idea, and then define their business goals and discuss strategies. At the same time, the team leader discusses the potential move with the other members of his or her group to assess their level of interest and prepare them for the change. The second stage involves the integration of the team leader with the new company's top leadership. This part of the process ensures the team's access to senior executives-the most important factor in a lift out's success. Operational integration is the focus of the third stage. Ideally, teams will start out working with the same or similar clients, vendors, and industry standards. The fourth stage entails full cultural integration. To succeed, the lifted-out team members must be willing to re-earn credibility by proving their value and winning their new colleagues' trust. PMID:17183798
Computational design and analysis of flatback airfoil wind tunnel experiment.
Mayda, Edward A. (University of California, Davis, CA); van Dam, C.P. (University of California, Davis, CA); Chao, David D. (University of California, Davis, CA); 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.
CAA boundary conditions for airfoil noise due to high-frequency gusts
Jae Wook Kim; Alex S. H. Lau; Neil D. Sandham
2010-01-01
This paper proposes inflow and outflow boundary conditions for direct computation of airfoil noise under the influence of impinging gusts with a particular interest in the high frequency response. The proposed boundary conditions are based on an existing zonal sponge technique that has been used mostly to attenuate outgoing disturbances and absorb reflections from computational boundaries. A modified form of
Numerical design of a transonic airfoil having a Stratford pressure recovery
Rocholl, Bruce Martin
1978-01-01
and supercritical rooftop pressures. For the critical rooftop airfoil, this recovery contibutes to substantially improved off design performance such as negligible drag creep and a high drag rise Mach number. The numerical technique used employs Cartesian... Introductory Comments . Analysis at Design Mach Number Drag Rise Mach Number Determination . Determination of Low Speed Maximum Lift CHAPTER VI DESIGN AND ANALYSIS OF A SUPERCRITICAL AIRFOIL Introductory Comments . Design Phase Analysis at Design...
High School Redesign Gets Presidential Lift
ERIC Educational Resources Information Center
Adams, Caralee J.
2013-01-01
President Barack Obama applauded high school redesign efforts in his State of the Union address and encouraged districts to look to successful models for inspiration. Last week, he followed up with a request in his fiscal 2014 budget proposal for a new, $300 million competitive-grant program. Recognition is widespread that high schools need to…
Multi-objective optimization design of low-Reynolds-number airfoil in Near Space
Rong Ma; Bowen Zhong; Peiqing Liu; Wenhu Wang
2010-01-01
For the reason of practical use in propeller design for low-dynamic aircrafts in Near Space, it is very important to have high-performance low-Reynolds-number airfoils which should have high lift-drag ratio, high endurance factor and high maximum-lift-coefficient, etc. Multi-objective optimal design can effectively solve the multiple design-points problem, which is an effective means and methods for optimum aerodynamic design. The paper
Prediction of high lift: review of present CFD capability
Christopher L. Rumsey; Susan X. Ying
2002-01-01
A survey is conducted of CFD methods applied to the computation of high-lift multi-element configurations over the last 10–15 years. Both 2-D and 3-D configurations are covered. The review is organized by configuration, in an effort to glean useful insights with respect to particular successes or failings of CFD methods as a whole. In general, for both 2-D and 3-D
Unsteady Transonic Airfoil Flow Simulations using High-Order WENO Schemes
I. Klioutchnikov; J. Ballmann; H. Oliver; V. Hermes; A. Alshabu
The flow over a supercritical airfoil in the transonic regime contains a lot of complex physical phenomena and interactions.\\u000a High demands are made on the numerical model for an accurate numerical investigation of such a flow type. High-order (N ? 5) WENO (Weighted Essentially Nonoscillatory) schemes are appropriate for the numerical simulation of transonic flows,\\u000a as they have the capability
Recent developments in high temperature coatings for gas turbine airfoils
NASA Technical Reports Server (NTRS)
Goward, G. W.
1983-01-01
The importance of coatings for hot section airfoils has increased with the drive for more cost-effective use of fuel in a wide variety of gas turbine engines. Minor additions of silicon have been found to appreciably increase the oxidation resistance of plasma-sprayed NiCoCrAlY coatings on a single crystal nickel-base superalloy. Increasing the chromium content of MCrAlY coatings substantially increases the resistance to acidic (Na2SO4-SO3) hot corrosion at temperatures of about 1300 F (704 C) but gives no significant improvement beyond contemporary coatings in the range of 1600 F (871 C). Surface enrichment of MCrAlY coatings with silicon also gives large increases in resistance to acidic hot corrosion in the 1300 F region. The resistance to the thermal stress-induced spalling of zirconia-based thermal barrier coatings has been improved by lowering coating stresses with segmented structures and by controlling the substrate temperature during coating fabrication.
Two-dimensional, subsonic, flowfield analysis for arbitrary boundaries with lifting bodies
Noack, Ralph Wayne
1981-01-01
special shapes, such as the lift- ing circular cylinder and Joukowski airfoils. Rather, for an arbitrary airfoil, the constant k must be obtained as part of the solution. Finally, for an inviscid, lifting flow, the specification of in- finity boundary...
NASA Technical Reports Server (NTRS)
Lowry, John G
1941-01-01
An investigation was made in the NACA 7- by 10-foot wind tunnel to determine the effect of slot-lip location on the aerodynamic section characteristics of an NACA 23012 airfoil with a 30-percent-chord slotted flap. Tests were made with slot lips located at 90 and 100 percent of the airfoil chord and with two different flap shapes. The results are compared with a slotted flap previously developed by the National advisory Committee for Aeronautics with a slot lip located at 83 percent of the airfoil chord. The extension of the slot lip to the rear increased the section lift and pitching-moment coefficients. Comparisons made on a basis of pitching moment for a given tail length show that the Fowler type flap, lip extended to trailing edge of the airfoil, has the greatest section lift coefficient. For moderate tail lengths, 2 to 3 chord lengths, there was only a slight difference between the previously developed slotted flap and the slotted flap with slot lip extended to 90 percent of the airfoil chord. Of the three flaps tested, the Fowler flap had the lowest drag coefficient at high lift coefficients. The extension of the lower surface at the leading edge of the slot had a negligible effect on the profile drag of the airfoil-flap arrangement with the flap deflected when the lip terminated at 90 percent of the airfoil chord.
Design and validation of a high-lift low-pressure turbine blade
NASA Astrophysics Data System (ADS)
McQuilling, Mark Wayne
This dissertation is a design and validation study of the high-lift low-pressure turbine (LPT) blade designated L2F. High-lift LPTs offer the promise of reducing the blade count in modern gas turbine engines. Decreasing the blade count can reduce development and maintenance costs and the weight of the engine, but care must be taken in order to maintain turbine section performance with fewer blades. For an equivalent amount of work extracted, lower blade counts increase blade loading in the LPT section. The high-lift LPT presented herein allows 38% fewer blades with a Zweifel loading coefficient of 1.59 and maintains the same inlet and outlet blade metal angles of conventional geometries in service today while providing an improved low-Reynolds number characteristic. The computational design method utilizes the Turbine Design and Analysis System (TDAAS) developed by John Clark of the Air Force Research Laboratory. TDAAS integrates several government-funded design utilities including airfoil and grid generation capability with a Reynolds-Averaged Navier-Stokes flow solver into a single, menu-driven, Matlab-based system. Transition modeling is achieved with the recently developed model of Praisner and Clark, and this study validates the use of the model for design purposes outside of the Pratt & Whitney (P&W) design system where they were created. Turbulence modeling is achieved with the Baldwin and Lomax zero-equation model. The experimental validation consists of testing the front-loaded L2F along with a previously designed, mid-loaded blade (L1M) in a linear turbine cascade in a low-speed wind tunnel over a range of Reynolds numbers at 3.3% freestream turbulence. Hot-wire anemometry and pressure measurements elucidate these comparisons, while a shear and stress sensitive film (S3F) also helps describe the flow in areas of interest. S3F can provide all 3 components of stress on a surface in a single measurement, and these tests extend the operational envelope of the technique to low speed air environments where small dynamic pressures and curved surfaces preclude the use of more traditional global measurement methods. Results are compared between the L1M and L2F geometries along with previous data taken in the same wind tunnel at identical flow conditions for the P&W Pack B geometry.
NASA Technical Reports Server (NTRS)
Nichino, Takafumi; Hahn, Seonghyeon; Shariff, Karim
2010-01-01
This slide presentation reviews the Large Eddy Simulation of a high reynolds number Coanda flow that is separated from a round trailing edge of a ciruclation control airfoil. The objectives of the study are: (1) To investigate detailed physics (flow structures and statistics) of the fully turbulent Coanda jet applied to a CC airfoil, by using LES (2) To compare LES and RANS results to figure out how to improve the performance of existing RANS models for this type of flow.
Gurney flap—Lift enhancement, mechanisms and applications
NASA Astrophysics Data System (ADS)
Wang, J. J.; Li, Y. C.; Choi, K.-S.
2008-01-01
Since its invention by a race car driver Dan Gurney in 1960s, the Gurney flap has been used to enhance the aerodynamics performance of subsonic and supercritical airfoils, high-lift devices and delta wings. In order to take stock of recent research and development of Gurney flap, we have carried out a review of the characteristics and mechanisms of lift enhancement by the Gurney flap and its applications. Optimum design of the Gurney flap is also summarized in this paper. For the Gurney flap to be effective, it should be mounted at the trailing edge perpendicular to the chord line of airfoil or wing. The flap height must be of the order of local boundary layer thickness. For subsonic airfoils, an additional Gurney flap increases the pressure on the upstream surface of the Gurney flap, which increases the total pressure of the lower surface. At the same time, a long wake downstream of the flap containing a pair of counter-rotating vortices can delay or eliminate the flow separation near the trailing edge on the upper surface. Correspondingly, the total suction on the airfoil is increased. For supercritical airfoils, the lift enhancement of the Gurney flap mainly comes from its ability to shift the shock on the upper surface in the downstream. Applications of the Gurney flap to modern aircraft design are also discussed in this review.
Transonic characteristics of a humped airfoil
NASA Technical Reports Server (NTRS)
Tai, Tsze C.; Huson, Gregory G.; Hicks, Raymond M.; Gregorek, Gerald M.
1987-01-01
A humped airfoil concept to improve the perfomance of supercritical airfoils at off-design conditions in transonic maneuver is introduced. Theoretical aspects of the airfoil and recent results of experimental tests conducted in the high Reynolds number transonic wind tunnels at both NASA Ames Research Center and The Ohio State University are presented. Experimental evidence has shown that the humped airfoil generally has more favorable transonic characteristics in off-design high supercritical flow conditions than a regular supercritical airfoil.
Aerodynamic flow control of a high lift system with dual synthetic jet arrays
NASA Astrophysics Data System (ADS)
Alstrom, Robert Bruce
Implementing flow control systems will mitigate the vibration and aeroacoustic issues associated with weapons bays; enhance the performance of the latest generation aircraft by reducing their fuel consumption and improving their high angle-of-attack handling qualities; facilitate steep climb out profiles for military transport aircraft. Experimental research is performed on a NACA 0015 airfoil with a simple flap at angle of attack of 16o in both clean and high lift configurations. The results of the active control phase of the project will be discussed. Three different experiments were conducted; they are Amplitude Modulated Dual Location Open Loop Control, Adaptive Control with Amplitude Modulation using Direct Sensor Feedback and Adaptive Control with Amplitude Modulation using Extremum Seeking Control. All the closed loop experiments are dual location. The analysis presented uses the spatial variation of the root mean square pressure fluctuations, power spectral density estimates, Fast Fourier Transforms (FFTs), and time frequency analysis which consists of the application of the Morlet and Mexican Hat wavelets. Additionally, during the course of high speed testing in the wind tunnel, some aeroacoustic phenomena were uncovered; those results will also be presented. A cross section of the results shows that the shape of the RMS pressure distributions is sensitive to forcing frequency. The application of broadband excitation in the case adaptive control causes the flow to select a frequency to lock in to. Additionally, open loop control results in global synchronization via switching between two stable states and closed loop control inhibits the switching phenomena, but rather synchronizes the flow about multiple stable shedding frequencies.
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)
Frey, Gary A. (Poway, CA); Twardochleb, Christopher Z. (Alpine, CA)
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.
Leading-edge singularities in thin-airfoil theory
NASA Technical Reports Server (NTRS)
Jones, R. T.
1976-01-01
If the thin airfoil theory is applied to an airfoil having a rounded leading edge, a certain error will arise in the determination of the pressure distribution around the nose. It is shown that the evaluation of the drag of such a blunt nosed airfoil by the thin airfoil theory requires the addition of a leading edge force, analogous to the leading edge thrust of the lifting airfoil. The method of calculation is illustrated by application to: (1) The Joukowski airfoil in subsonic flow; and (2) the thin elliptic cone in supersonic flow. A general formula for the edge force is provided which is applicable to a variety of wing forms.
Identification of airfoil characteristics for optimum wind turbine performance / b
Miller, Leonard Scott
1983-01-01
is determined only by the airfoil's lift and drag characteristics and the effect of the local twist angle is eliminated. Therefore, the twist optimizer is an angle of attack optimi zer which isolates the effect of an airfoil's behavior on turbine performance.... This routine adjusts the blade segment angle of attack so as to yield local maximum power output. In this condition the airfoil's effect on turbine performance is isolated from any possible geometric influence due to twist . This assures that each airfoil...
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.
A critical assessment of UH-60 main rotor blade airfoil data
NASA Technical Reports Server (NTRS)
Totah, Joseph
1993-01-01
Many current comprehensive rotorcraft analyses employ lifting-line methods that require main rotor blade airfoil data, typically obtained from wind tunnel tests. In order to effectively evaluate these lifting-line methods, it is of the utmost importance to ensure that the airfoil section data are free of inaccuracies. A critical assessment of the SC1095 and SC1094R8 airfoil data used on the UH-60 main rotor blade was performed for that reason. Nine sources of wind tunnel data were examined, all of which contain SC1095 data and four of which also contain SC1094R8 data. Findings indicate that the most accurate data were generated in 1982 at the 11-Foot Wind Tunnel Facility at NASA Ames Research Center and in 1985 at the 6-inch-by-22-inch transonic wind tunnel facility at Ohio State University. It has not been determined if data from these two sources are sufficiently accurate for their use in comprehensive rotorcraft analytical models of the UH-60. It is recommended that new airfoil tables be created for both airfoils using the existing data. Additional wind tunnel experimentation is also recommended to provide high quality data for correlation with these new airfoil tables.
A critical assessment of UH-60 main rotor blade airfoil data
NASA Technical Reports Server (NTRS)
Totah, Joseph
1993-01-01
Many current comprehensive rotorcraft analyses employ lifting-line methods that require main rotor blade airfoil data, typically obtained from wind tunnel tests. In order to effectively evaluate these lifting-line methods, it is of the utmost importance to ensure that the airfoil section data are free of inaccuracies. A critical assessment of the SC1095 and SC1094R8 airfoil data used on the UH-60 main rotor blade was performed for that reason. Nine sources of wind tunnel data were examined, all of which contain SC1095 data and four of which also contain SC1094R8 data. Findings indicate that the most accurate data were generated in 1982 at the 11-Foot Wind Tunnel Facility at NASA Ames Research Center and in 1985 at the 6-inch by 22-inch transonic wind tunnel facility at Ohio State University. It has not been determined if data from these two sources are sufficiently accurate for their use in comprehensive rotorcraft analytical models of the UH-60. It is recommended that new airfoil tables be created for both airfoils using the existing data. Additional wind tunnel experimentation is also recommended to provide high quality data for correlation with these new airfoil tables.
Airfoil optimization for unsteady flows with application to high-lift noise reduction
Markus Peer Rumpfkeil
2008-01-01
The use of steady-state aerodynamic optimization methods in the computational fluid dynamic (CFD) community is fairly well established. In particular, the use of adjoint methods has proven to be very beneficial because their cost is independent of the number of design variables. The application of numerical optimization to airframe-generated noise, however, has not received as much attention, but with the
Airfoil Optimization for Unsteady Flows with Application to High-Lift Noise Reduction
Rumpfkeil, Markus Peer
wave propagation program based on the Ffowcs Williams and Hawkings (FW-H) formulation is implemented and validated. The general framework is then used to derive the adjoint equations for a novel hybrid URANS/FW-H far-field pressure fluctuations. Validation and application results for this novel hybrid URANS/FW-H
A study on high subsonic airfoil flows in relatively high Reynolds number by using OpenFOAM
NASA Astrophysics Data System (ADS)
Nakao, Shinichiro; Kashitani, Masashi; Miyaguni, Takeshi; Yamaguchi, Yutaka
2014-04-01
In the present study, numerical calculations of the flow-field around the airfoil model are performed by using the OpenFOAM in high subsonic flows. The airfoil model is NACA 64A010. The maximum thickness is 10 % of the chord length. The SonicFOAM and the RhoCentralFOAM are selected as the solver in high subsonic flows. The grid point is 158,000 and the Mach numbers are 0.277 and 0.569 respectively. The CFD data are compared with the experimental data performed by the cryogenic wind tunnel in the past. The results are as follows. The numerical results of the pressure coefficient distribution on the model surface calculated by the SonicFOAM solver showed good agreement with the experimental data measured by the cryogenic wind tunnel. And the data calculated by the SonicFOAM have the capability for the quantitative comparison of the experimental data at low angle of attack.
NASA Technical Reports Server (NTRS)
Crivellini, A.; Golubev, V.; Mankbadi, R.; Scott, J. R.; Hixon, R.; Povinelli, L.; Kiraly, L. James (Technical Monitor)
2002-01-01
The nonlinear response of symmetric and loaded airfoils to an impinging vortical gust is investigated in the parametric space of gust dimension, intensity, and frequency. The study, which was designed to investigate the validity limits for a linear analysis, is implemented by applying a nonlinear high-order prefactored compact code and comparing results with linear solutions from the GUST3D frequency-domain solver. Both the unsteady aerodynamic and acoustic gust responses are examined.
NASA Technical Reports Server (NTRS)
Edge, D. Christian; Perkins, John N.
1995-01-01
The sizing and efficiency of an aircraft is largely determined by the performance of its high-lift system. Subsonic civil transports most often use deployable multi-element airfoils to achieve the maximum-lift requirements for landing, as well as the high lift-to-drag ratios for take-off. However, these systems produce very complex flow fields which are not fully understood by the scientific community. In order to compete in today's market place, aircraft manufacturers will have to design better high-lift systems. Therefore, a more thorough understanding of the flows associated with these systems is desired. Flight and wind-tunnel experiments have been conducted on NASA Langley's B737-100 research aircraft to obtain detailed full-scale flow measurements on a multi-element high-lift system at various flight conditions. As part of this effort, computational aerodynamic tools are being used to provide preliminary flow-field information for instrumentation development, and to provide additional insight during the data analysis and interpretation process. The purpose of this paper is to demonstrate the ability and usefulness of a three-dimensional low-order potential flow solver, PMARC, by comparing computational results with data obtained from 1/8 scale wind-tunnel tests. Overall, correlation of experimental and computational data reveals that the panel method is able to predict reasonably well the pressures of the aircraft's multi-element wing at several spanwise stations. PMARC's versatility and usefulness is also demonstrated by accurately predicting inviscid three-dimensional flow features for several intricate geometrical regions.
Validation of the CQU-DTU-LN1 series of airfoils
NASA Astrophysics Data System (ADS)
Shen, W. Z.; Zhu, W. J.; Fischer, A.; Garcia, N. R.; Cheng, J. T.; Chen, J.; Madsen, J.
2014-12-01
The CQU-DTU-LN1 series of airfoils were designed with an objective of high lift and low noise emission. In the design process, the aerodynamic performance is obtained using XFOIL while noise emission is obtained with the BPM model. In this paper we present some validations of the designed CQU-DTU-LN118 airfoil by using wind tunnel measurements in the acoustic wind tunnel located at Virginia Tech and numerical computations with the inhouse Q3uic and EllipSys 2D/3D codes. To show the superiority of the new airfoils, comparisons with a NACA64618 airfoil are made. For the aerodynamic features, the designed Cl and Cl/Cd agrees well with the experiment and are in general higher than those of the NACA airfoil. For the acoustic features, the noise emission of the LN118 airfoil is compared with the acoustic measurements and that of the NACA airfoil. Comparisons show that the BPM model can predict correctly the noise changes.
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.
The ultra-low Reynolds number airfoil wake
NASA Astrophysics Data System (ADS)
Alam, Md. Mahbub; Zhou, Y.; Yang, H. X.; Guo, H.; Mi, J.
2010-01-01
Lift force and the near wake of an NACA 0012 airfoil were measured over the angle (?) of attack of 0°-90° and the chord Reynolds number ( Re c ), 5.3 × 103-5.1 × 104, with a view to understand thoroughly the near wake of the airfoil at low- to ultra-low Re c . While the lift force is measured using a load cell, the detailed flow structure is captured using laser-Doppler anemometry, particle image velocimetry, and laser-induced fluorescence flow visualization. It has been found that the stall of an airfoil, characterized by a drop in the lift force, occurs at Re c ? 1.05 × 104 but is absent at Re c = 5.3 × 103. The observation is connected to the presence of the separation bubble at high Re c but absence of the bubble at ultra-low Re c , as evidenced in our wake measurements. The near-wake characteristics are examined and discussed in detail, including the vortex formation length, wake width, spanwise vorticity, wake bubble size, wavelength of K-H vortices, Strouhal numbers, and their dependence on ? and Re c .
Prediction of laminar-turbulent transition on an airfoil at high level of free-stream turbulence
NASA Astrophysics Data System (ADS)
Chernoray, V.
2015-06-01
Prediction of laminar-turbulent transition at high level of free-stream turbulence in boundary layers of airfoil geometries with external pressure gradient changeover is in focus. The aim is a validation of a transition model for transition prediction in turbomachinery applications. Numerical simulations have been performed by using a transition model by Langtry and Menter for a number of different cases of pressure gradient, at Reynolds-number range, based on the airfoil chord, 50 000 ? Re ? 500 000, and free-stream turbulence intensities 2% and 4%. The validation of the computational results against the experimental data showed good performance of used turbulence model for all test cases.
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.
NASA Technical Reports Server (NTRS)
Weick, Fred E; Shortal, Joseph A
1933-01-01
Lift and drag tests were made on a Clark Y wing equipped with four fixed slots and a trailing-edge flap in the 5-foot vertical wind tunnel of the National Advisory Committee for Aeronautics. All possible combinations of the four slots were tested with the flap neutral and the most promising combinations were tested with the flap down 45 degrees. Considering both the maximum lift coefficient and the speed-range ratio with the flap neutral no appreciable improvement was found with the use of more than the single leading-edge slot. With the flap down 45 degrees a maximum lift coefficient of 2.60 was obtained but the particular slot combination used had a rather large minimum drag coefficient with the flap neutral. With the flap down 45 degrees the optimum combination, considering both the maximum lift coefficient and the speed-range ratio, was obtained with only the two rearmost slots in use. For this arrangement the maximum lift coefficient was 2.44.
Overview of Fundamental High-Lift Research for Transport Aircraft at NASA
NASA Technical Reports Server (NTRS)
Leavitt, L. D.; Washburn, A. E.; Wahls, R. A.
2007-01-01
NASA has had a long history in fundamental and applied high lift research. Current programs provide a focus on the validation of technologies and tools that will enable extremely short take off and landing coupled with efficient cruise performance, simple flaps with flow control for improved effectiveness, circulation control wing concepts, some exploration into new aircraft concepts, and partnership with Air Force Research Lab in mobility. Transport high-lift development testing will shift more toward mid and high Rn facilities at least until the question: "How much Rn is required" is answered. This viewgraph presentation provides an overview of High-Lift research at NASA.
H. Arbey; M. Sunyach; G. Comte-Bellot
1979-01-01
The paper describes a new experimental setup for the investigation of noise generated by airfoils. The setup involves two adjacent jets ducted into an anechoic chamber, with the airfoils under investigation placed in the median jet. This leads to improvements in the acoustic conditions of the experiment (decrease of background noise due to the jets and shift of the preferred
Aerodynamic Control of a Pitching Airfoil using Distributed Active Bleed
NASA Astrophysics Data System (ADS)
Kearney, John; Glezer, Ari
2012-11-01
Aero-effected flight control using distributed active bleed driven by pressure differences across lifting surface and regulated by integrated louver actuators is investigated in wind tunnel experiments. The interaction between unsteady bleed and the cross flows alters the apparent aerodynamic shape of the lifting surface by regulating the accumulation and shedding of vorticity concentrations, and consequently the distributions of forces and moments. The present experiments are conducted using a 2-D dynamically-pitching VR-7 airfoil model from pre- to post-stall angles of attack. The effects of leading edge bleed at high angles of attack on the formation and evolution of the dynamic stall vorticity concentrations are investigated at high reduced frequencies (k > 0.1) using PIV phase-locked to the airfoil's motion. The time-dependent bleed enables broad-range variation in lift and pitching moment with significant extension of the stall margin. In particular, bleed actuation reduces the extent of ``negative damping'' or pitching moment instability with minimal lift penalty. Supported by NTRC-VLRCOE, monitored by Dr. Mike Rutkowski.
High lift force with 275 Hz wing beat in MFI
Erik Steltz; Srinath Avadhanula; Ronald S. Fearing
2007-01-01
Abstract—The Micromechanical,Flying Insect (MFI) project aims to create a 25 mm,(wingtip to wingtip) flapping wing micro air vehicle inspired by the aerodynamics,of insect flight. A key challenge is generating,appropriate,wing,trajectories. Previous work,[1] showed,a lift of 506 µN at 160 Hz using feedforward control. In this paper, refinements to the MFI design including those in [2] increased,wing,beat frequency,to 275 Hz and,lift to
A direct-inverse method for transonic and separated flows about airfoils
NASA Technical Reports Server (NTRS)
Carlson, Leland A.
1990-01-01
A direct-inverse technique and computer program called TAMSEP that can be used for the analysis of the flow about airfoils at subsonic and low transonic freestream velocities is presented. The method is based upon a direct-inverse nonconservative full potential inviscid method, a Thwaites laminar boundary layer technique, and the Barnwell turbulent momentum integral scheme; and it is formulated using Cartesian coordinates. Since the method utilizes inverse boundary conditions in regions of separated flow, it is suitable for predicting the flow field about airfoils having trailing edge separated flow under high lift conditions. Comparisons with experimental data indicate that the method should be a useful tool for applied aerodynamic analyses.
A direct-inverse method for transonic and separated flows about airfoils
NASA Technical Reports Server (NTRS)
Carlson, K. D.
1985-01-01
A direct-inverse technique and computer program called TAMSEP that can be sued for the analysis of the flow about airfoils at subsonic and low transonic freestream velocities is presented. The method is based upon a direct-inverse nonconservative full potential inviscid method, a Thwaites laminar boundary layer technique, and the Barnwell turbulent momentum integral scheme; and it is formulated using Cartesian coordinates. Since the method utilizes inverse boundary conditions in regions of separated flow, it is suitable for predicing the flowfield about airfoils having trailing edge separated flow under high lift conditions. Comparisons with experimental data indicate that the method should be a useful tool for applied aerodynamic analyses.
Aerodynamics of S809 Airfoil at Low and Transitional Reynolds Numbers
NASA Astrophysics Data System (ADS)
Carreras, Jaime J.; Laal-Dehghani, Nader; Gorumlu, Serdar; Mehdi, Faraz; Castillo, Luciano; Aksak, Burak; Sheng, Jian
2013-11-01
The S809 is a thick airfoil extensively used in wind turbine design applications and model studies in wind tunnel. With increased interests in reducing energy production cost and understanding turbulence and turbine interactions, scaled down models (Re ~103) are often used as an alternative to full scale field experimentation (Re >106). This Reynolds number discrepancy raises the issue of scaling for the airfoil performance from laboratory studies to field scale applications. To the best of our knowledge, there are no studies existing in literature to characterize the lift- and drag-coefficients of S809 airfoil at Re less than 3 ×105 . This study is to fill the deficit in the current state of knowledge by performing high resolution force measurements. The lift and drag measurements are carried out in Texas Tech Wind Tunnel Facility using an in-house developed dual-cell force balance. The configuration eliminates the large torque and torsion often accompanied by conventional mounts. This unique design allows us to reach a measurement accuracy of 0.02N (0.1%). Comparative studies are performed on a two-dimensional airfoil with a smooth- as well as a well-engineered surface covered by micro-pillar array to simulate the surface conditions of a real life airfoil.
NASA Technical Reports Server (NTRS)
Mcdevitt, J. B.; Okuno, A. F.
1985-01-01
The supercritical flows at high subsonic speeds over a NACA 0012 airfoil were studied to acquire aerodynamic data suitable for evaluating numerical-flow codes. The measurements consisted primarily of static and dynamic pressures on the airfoil and test-channel walls. Shadowgraphs were also taken of the flow field near the airfoil. The tests were performed at free-stream Mach numbers from approximately 0.7 to 0.8, at angles of attack sufficient to include the onset of buffet, and at Reynolds numbers from 1 million to 14 million. A test action was designed specifically to obtain two-dimensional airfoil data with a minimum of wall interference effects. Boundary-layer suction panels were used to minimize sidewall interference effects. Flexible upper and lower walls allow test-channel area-ruling to nullify Mach number changes induced by the mass removal, to correct for longitudinal boundary-layer growth, and to provide contouring compatible with the streamlines of the model in free air.
1997 NASA High-Speed Research Program Aerodynamic Performance Workshop. Volume 2; High Lift
NASA Technical Reports Server (NTRS)
Baize, Daniel G. (Editor)
1999-01-01
The High-Speed Research Program and NASA Langley Research Center sponsored the NASA High-Speed Research Program Aerodynamic Performance Workshop on February 25-28, 1997. The workshop was designed to bring together NASA and industry High-Speed Civil Transport (HSCT) Aerodynamic Performance technology development participants in areas of Configuration Aerodynamics (transonic and supersonic cruise drag, prediction and minimization), High-Lift, Flight Controls, Supersonic Laminar Flow Control, and Sonic Boom Prediction. The workshop objectives were to (1) report the progress and status of HSCT aerodynamic performance technology development; (2) disseminate this technology within the appropriate technical communities; and (3) promote synergy among the scientist and engineers working HSCT aerodynamics. In particular, single- and multi-point optimized HSCT configurations, HSCT high-lift system performance predictions, and HSCT Motion Simulator results were presented along with executives summaries for all the Aerodynamic Performance technology areas.
High-precision position control of a heavy-lift manipulator in a dynamic environment
Garretson, Justin R. (Justin Richard)
2005-01-01
This thesis considers the control of a heavy-lift serial manipulator operating on the deck of a large ocean vessel. This application presents a unique challenge for high- precision control because the system must contend ...
Wind Tunnel and Numerical Analysis of Thick Blunt Trailing Edge Airfoils
NASA Astrophysics Data System (ADS)
McLennan, Anthony William
Two-dimensional aerodynamic characteristics of several thick blunt trailing edge airfoils are presented. These airfoils are not only directly applicable to the root section of wind turbine blades, where they provide the required structural strength at a fraction of the material and weight of an equivalent sharp trailing edge airfoil, but are also applicable to the root sections of UAVs having high aspect ratios, that also encounter heavy root bending forces. The Reynolds averaged Navier-Stokes code, ARC2D, was the primary numerical tool used to analyze each airfoil. The UCD-38-095, referred to as the Pareto B airfoil in this thesis, was also tested in the University of California, Davis Aeronautical Wind Tunnel. The Pareto B has an experimentally determined maximum lift coefficient of 1.64 at 14 degrees incidence, minimum drag coefficient of 0.0385, and maximum lift over drag ratio of 35.9 at a lift coefficient of 1.38, 10 degrees incidence at a Reynolds number of 666,000. Zig-zag tape at 2% and 5% of the chord was placed on the leading edge pressure and suction side of the Pareto B model in order to determine the aerodynamic performance characteristics at turbulent flow conditions. Experimental Pareto B wind tunnel data and previous FB-3500-0875 data is also presented and used to validate the ARC2D results obtained in this study. Additionally MBFLO, a detached eddy simulation Navier-Stokes code, was used to analyze the Pareto B airfoil for comparison and validation purposes.
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.
Calculation of transonic flow over supercritical airfoil sections
NASA Technical Reports Server (NTRS)
Rose, W. C.; Seginer, A.
1977-01-01
Two numerical methods for calculating the transonic flow, including viscous effects over lifting airfoil sections and experimental data are compared for turbulent flow over a supercritical airfoil. In addition, results for a NACA 64A010 airfoil at nonzero angle of attack are compared to demonstrate the applicability of the numerical methods to classical, lifting airfoils. One numerical method is a solution to the time-averaged Navier-Stokes equations throughout the entire flow field. The other is a hybrid method that combines inviscid, boundary-layer, and Navier-Stokes equations in appropriate regions of the flow field. Both methods adequately predict the surface pressures and flow field about the 64A010 airfoil at M = 0.8 and alpha = 2 deg when an appropriate turbulence model is used. The methods are not as successful for the supercritical airfoil.
Application of shock tubes to transonic airfoil testing at high Reynolds numbers
NASA Technical Reports Server (NTRS)
Cook, W. J.; Chaney, M. J.; Presley, L. L.; Chapman, G. T.
1978-01-01
Performance analysis of a gas-driven shock tube shows that transonic airfoil flows with chord Reynolds numbers of the order of 100 million can be produced, with limitations being imposed by the structural integrity of the facility or the model. A study of flow development over a simple circular arc airfoil at zero angle of attack was carried out in a shock tube at low and intermediate Reynolds numbers to assess the testing technique. Results obtained from schlieren photography and airfoil pressure measurements show that steady transonic flows similar to those produced for the same airfoil in a wind tunnel can be generated within the available testing time in a shock tube with properly contoured test section walls.
Optimization of the airfoil stroke in a high Reynolds number flow for energy harvesting
NASA Astrophysics Data System (ADS)
Guo, Xinjun; Mandre, Shreyas
2012-11-01
We investigate the heaving and pitching stroke of an airfoil for maximum energy extraction from the flow of the ambient fluid. This analysis is targeted towards optimization of oscillating airfoil or hydrofoils for wind and hydrokinetic energy conversion respectively. The goal is to study the influence of unsteady aerodynamic effects like leading edge vortex, unsteady boundary layer separation, vortex recapture, etc. We are inspired by the mechanics of insect and bird flight, which are believed to use unsteady aerodynamics for enhanced performance. Our airfoil has two degrees of freedom, heaving and pitching, and these degrees of freedom are actuated independently. We employ a variational framework for optimizing the transient stroke of the airfoil with the objective function being the time-averaged harvested power.
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.
Numerical simulations of iced airfoils and wings
NASA Astrophysics Data System (ADS)
Pan, Jianping
A numerical study was conducted to understand the effects of simulated ridge and leading-edge ice shapes on the aerodynamic performance of airfoils and wings. In the first part of this study, a range of Reynolds numbers and Mach numbers, as well as ice-shape sizes and ice-shape locations were examined for various airfoils with the Reynolds-Averaged Navier-Stokes approach. Comparisons between simulation results and experimental force data showed favorable comparison up to stall conditions. At and past stall condition, the aerodynamic forces were typically not predicted accurately for large upper-surface ice shapes. A lift-break (pseudo-stall) condition was then defined based on the lift curve slope change. The lift-break angles compared reasonably with experimental stall angles, and indicated that the critical ice-shape location tended to be near the location of minimum pressure and the location of the most adverse pressure gradient. With the aim of improving the predictive ability of the stall behavior for iced airfoils, simulations using the Detached Eddy Simulation (DES) approach were conducted in the second part of this numerical investigation. Three-dimensional DES computations were performed for a series of angles of attack around stall for the iced NACA 23012 and NLF 0414 airfoils. The simulations for both iced airfoils provided the maximum lift coefficients and stall behaviors qualitatively consistent with experiments.
NASA Astrophysics Data System (ADS)
Arko, Bryan M.
Design trends for the low-pressure turbine (LPT) section of modern gas turbine engines include increasing the loading per airfoil, which promises a decreased airfoil count resulting in reduced manufacturing and operating costs. Accurate Reynolds-Averaged Navier-Stokes predictions of separated boundary layers and transition to turbulence are needed, as the lack of an economical and reliable computational model has contributed to this high-lift concept not reaching its full potential. Presented here for what is believed to be the first time applied to low-Re computations of high-lift linear cascade simulations is the Abe-Kondoh-Nagano (AKN) linear low-Re two-equation turbulence model which utilizes the Kolmogorov velocity scale for improved predictions of separated boundary layers. A second turbulence model investigated is the Kato-Launder modified version of the AKN, denoted MPAKN, which damps turbulent production in highly strained regions of flow. Fully Laminar solutions have also been calculated in an effort to elucidate the transitional quality of the turbulence model solutions. Time accurate simulations of three modern high-lift blades at a Reynolds number of 25,000 are compared to experimental data and higher-order computations in order to judge the accuracy of the results, where it is shown that the RANS simulations with highly refined grids can produce both quantitatively and qualitatively similar separation behavior as found in experiments. In particular, the MPAKN model is shown to predict the correct boundary layer behavior for all three blades, and evidence of transition is found through inspection of the components of the Reynolds Stress Tensor, spectral analysis, and the turbulence production parameter. Unfortunately, definitively stating that transition is occurring becomes an uncertain task, as similar evidence of the transition process is found in the Laminar predictions. This reveals that boundary layer reattachment may be a result of laminar vortex shedding interaction with the blade surface, or that application of Laminar solvers with highly refined grids may be valid for investigations of transitional flow. It is recommended that this be studied in the future along with further studies of the turbulence models presented, which have proved to be worthy of additional study in the present application.
Control of unsteady separated flow associated with the dynamic pitching of airfoils
NASA Technical Reports Server (NTRS)
Ahmed, Sajeer
1991-01-01
Although studies have been done to understand the dependence of parameters for the occurrence of deep stall, studies to control the flow for sustaining lift for a longer time has been little. To sustain the lift for a longer time, an understanding of the development of the flow over the airfoil is essential. Studies at high speed are required to study how the flow behavior is dictated by the effects of compressibility. When the airfoil is pitched up in ramp motion or during the upstroke of an oscillatory cycle, the flow development on the upper surface of the airfoil and the formation of the vortex dictates the increase in lift behavior. Vortex shedding past the training edge decreases the lift. It is not clear what is the mechanism associated with the unsteady separation and vortex formation in present unsteady environment. To develop any flow control device, to suppress the vortex formation or delay separation, it is important that this mechanism be properly understood. The research activities directed toward understanding these questions are presented and the results are summarized.
Optimum lifting body shapes in hypersonic flow at high angles of attack
V. N. Golubkin; V. V. Negoda
1995-01-01
A variational procedure for the determination of lifting body configurations having a maximum lift-to-drag ratio Kmax in hypersonic flight at high angles of attack a, is proposed. It is based on an analytical solution to the problem for three-dimensional hypersonic flow over small aspect ratio wings using thin shock-layer theory. This reduces the variational problem of finding Kmax, and the
Demonstration of a Low-Lift Heat Pump for High-Power Spacecraft Thermal Control
Lawrence R. Grzyll
2006-01-01
This paper describes the development and demonstration of a prototype low-lift heat pump for high-power spacecraft thermal control The low-lift heat pump was designed to provide 25 kW of cooling at 303 K and transport this waste heat to a radiator for heat rejection. To accomplish this, a demonstration heat pump with an evaporation temperature of 298 K and a
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.
Tests of an N.A.C.A. 23012 Airfoil with a slotted deflector flap
NASA Technical Reports Server (NTRS)
House, R O
1939-01-01
Section aerodynamic characteristics of a large-chord N.A.C.A. 23012 airfoil with a slotted deflector flap were obtained in the N.A.C.A. 7- by 10-foot wind tunnel. The characteristics of an N.A.C.A. slotted flap and of a simple split flap are included for comparison. The slotted deflector flap was found to have a somewhat lower maximum lift coefficient and somewhat higher drag at high lift coefficients than the N.A.C.A. slotted flap. The high drag of the open slot with the deflector flap neutral indicates that the slot should be closed for this condition.
Airfoil wake and linear theory gust response including sub and superresonant flow conditions
NASA Technical Reports Server (NTRS)
Henderson, Gregory H.; Fleeter, Sanford
1992-01-01
The unsteady aerodynamic gust response of a high solidity stator vane row is examined in terms of the fundamental gust modeling assumptions with particular attention given to the effects near an acoustic resonance. A series of experiments was performed with gusts generated by rotors comprised of perforated plates and airfoils. It is concluded that, for both the perforated plate and airfoil wake generated gusts, the unsteady pressure responses do not agree with the linear-theory gust predictions near an acoustic resonance. The effects of the acoustic resonance phenomena are clearly evident on the airfoil surface unsteady pressure responses. The transition of the measured lift coefficients across the acoustic resonance from the subresonant regime to the superresonant regime occurs in a simple linear fashion.
NASA Technical Reports Server (NTRS)
Chawla, Kalpana
1993-01-01
Attached as appendices to this report are documents describing work performed on the simulation of a landing powered-lift delta wing, the tracking of flow features using overset grids, and the simulation of flaps on the Wright Patterson Lab's fighter-lift-and-control (FLAC) wing. Numerical simulation of a powered-lift landing includes the computation of flow about a delta wing at four fixed heights as well as a simulated landing, in which the delta wing descends toward the ground. Comparison of computed and experimental lift coefficients indicates that the simulations capture the qualitative trends in lift-loss encountered by thrust-vectoring aircraft operating in ground effect. Power spectra of temporal variations of pressure indicate computed vortex shedding frequencies close to the jet exit are in the experimentally observed frequency range; the power spectra of pressure also provide insights into the mechanisms of lift oscillations. Also, a method for using overset grids to track dynamic flow features is described and the method is validated by tracking a moving shock and vortices shed behind a circular cylinder. Finally, Chimera gridding strategies were used to develop pressure coefficient contours for the FLAC wing for a Mach no. of 0.18 and Reynolds no. of 2.5 million.
Experimental and simulated control of lift using trailing edge devices
NASA Astrophysics Data System (ADS)
Cooperman, A.; Blaylock, M.; van Dam, C. P.
2014-12-01
Two active aerodynamic load control (AALC) devices coupled with a control algorithm are shown to decrease the change in lift force experienced by an airfoil during a change in freestream velocity. Microtabs are small (1% chord) surfaces deployed perpendicular to an airfoil, while microjets are pneumatic jets with flow perpendicular to the surface of the airfoil near the trailing edge. Both devices are capable of producing a rapid change in an airfoil's lift coefficient. A control algorithm for microtabs has been tested in a wind tunnel using a modified S819 airfoil, and a microjet control algorithm has been simulated for a NACA 0012 airfoil using OVERFLOW. In both cases, the AALC devices have shown the ability to mitigate the changes in lift during a gust.
Does Small High School Reform Lift Urban Districts? Evidence from New York City
ERIC Educational Resources Information Center
Stiefel, Leanna; Schwartz, Amy Ellen; Wiswall, Matthew
2015-01-01
Research finds that small high schools deliver better outcomes than large high schools for urban students. An important outstanding question is whether this better performance is gained at the expense of losses elsewhere: Does small school reform lift the whole district? We explore New York City's small high school reform in which hundreds of new…
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 required for the airfoil-aircraft matching. Examples are presented to illustrate the flapped-airfoil design approach for a general aviation aircraft and the results are validated by comparison with results from post-design aircraft performance computations. Once the airfoil is designed to incorporate a TE flap, it is important to determine the most suitable flap angles along the wing for different flight conditions. The second part of this dissertation presents a method for determining the optimum flap angles to minimize drag based on pressures measured at select locations on the wing. Computational flow simulations using a panel method are used "in the loop" for demonstrating closed-loop control of the flaps. Examples in the paper show that the control algorithm is successful in correctly adapting the wing to achieve the target lift distributions for minimizing induced drag while adjusting the wing angle of attack for operation of the wing in the drag bucket. It is shown that the "sense-and-adapt" approach developed is capable of handling varying and unpredictable inflow conditions. Such a capability could be useful in adapting long-span flexible wings that may experience significant and unknown atmospheric inflow variations along the span. To further develop the "sense-and-adapt" approach, the method was tested experimentally in the third part of the research. The goal of the testing was to see if the same results found computationally can be obtained experimentally. The North Carolina State University subsonic wind tunnel was used for the wind tunnel tests. Results from the testing showed that the "sense-and-adapt" approach has the same performance experimentally as it did computationally. The research presented in this dissertation is a stepping stone towards further development of the concept, which includes modeling the system in the Simulink environment and flight experiments using uninhabited aerial vehicles.
Transonic Flow Past a Symmetrical Airfoil at High Angle of Attack
NASA Technical Reports Server (NTRS)
Johnson, D. A.; Bachalo, W. D.; Owen, F. K.
1981-01-01
The results of an experimental investigation of shock-induced stall and leading-edge stall on a 64A010 airfoil section are presented. Advanced nonintrusive techniques - laser velocimetry and holographic interferometry - were used in characterizing the inviscid and viscous flow regions. The measurements include Mach contours of the inviscid now regions, and mean velocity, flow direction, and Reynolds shear stress profiles in the separated regions. The experimental observations of this study are relevant to efforts to improve surface-pressure prediction methods for airfoils at or near stall.
Large-eddy simulation of flow around an airfoil on a structured mesh
NASA Technical Reports Server (NTRS)
Kaltenbach, Hans-Jakob; Choi, Haecheon
1995-01-01
The diversity of flow characteristics encountered in a flow over an airfoil near maximum lift taxes the presently available statistical turbulence models. This work describes our first attempt to apply the technique of large-eddy simulation to a flow of aeronautical interest. The challenge for this simulation comes from the high Reynolds number of the flow as well as the variety of flow regimes encountered, including a thin laminar boundary layer at the nose, transition, boundary layer growth under adverse pressure gradient, incipient separation near the trailing edge, and merging of two shear layers at the trailing edge. The flow configuration chosen is a NACA 4412 airfoil near maximum lift. The corresponding angle of attack was determined independently by Wadcock (1987) and Hastings & Williams (1984, 1987) to be close to 12 deg. The simulation matches the chord Reynolds number U(sub infinity)c/v = 1.64 x 10(exp 6) of Wadcock's experiment.
Design and experimental results for the S809 airfoil
Somers, D.M. [Airfoils, Inc., State College, PA (United States)] [Airfoils, Inc., State College, PA (United States)
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.
Shape optimization of single- and two-element airfoils on multiblock grids
NASA Technical Reports Server (NTRS)
Lacasse, James M.; Baysal, Oktay
1995-01-01
A multiblock, discrete sensitivity analysis method is used to couple a direct optimization method and a flow analysis method. The domain is divided into smaller subdomains for which the sensitivities are obtained separately. Then, an effective sensitivity equation is solved to complete the coupling of all the sensitivity information. The flow analysis is based on the thin-layer Navier-Stokes equations solved by an implicit, upwind-biased, finite-volume method. The method of feasible directions is used for the present gradient-based optimization approach. First, a transonic airfoil is optimized to investigate the behavior of the method in highly nonlinear flows as well as the effect of different blocking strategies on the procedure. A supercritical airfoil is produced from an initially symmetric airfoil with multiblocking affecting the path but not the final shape. Secondly, a two-element airfoil is shape optimized in subsonic flow to demonstrate the present method's capability of shaping aerodynamically interfering elements simultaneously. For a very low and a very high Reynolds number cases, the shape of the main airfoil and the flap are optimized to yield improved lift-to-drag ratios.
Active Control of Flow Separation Over an Airfoil
NASA Technical Reports Server (NTRS)
Ravindran, S. S.
1999-01-01
Designing an aircraft without conventional control surfaces is of interest to aerospace community. In this direction, smart actuator devices such as synthetic jets have been proposed to provide aircraft maneuverability instead of control surfaces. In this article, a numerical study is performed to investigate the effects of unsteady suction and blowing on airfoils. The unsteady suction and blowing is introduced at the leading edge of the airfoil in the form of tangential jet. Numerical solutions are obtained using Reynolds-Averaged viscous compressible Navier-Stokes equations. Unsteady suction and blowing is investigated as a means of separation control to obtain lift on airfoils. The effect of blowing coefficients on lift and drag is investigated. The numerical simulations are compared with experiments from the Tel-Aviv University (TAU). These results indicate that unsteady suction and blowing can be used as a means of separation control to generate lift on airfoils.
A lifting surface theory in rotational flow
NASA Technical Reports Server (NTRS)
Shiau, M. J.; Lan, C. E.
1983-01-01
The partial differential equation for small disturbance steady rotational flow in three dimensions is solved through an integral equation approach. The solution is obtained by using the method of weighted residuals. Specific applications are directed to wings in nonuniform subsonic parallel streams with velocity varying in vertical and spanwise directions and to airfoils in nonuniform freestream. Comparison with limited known results indicates that the present method is reasonably accurate. Numerical results for the lifting pressure of airfoil, lift, induced drag, and pitching moments of airfoil, lift, induced drag, and pitching moments of elliptic, rectangular, and delta wings in a jet, wake, or monotonic sheared stream are presented. It is shown that, in addition to the effect of local dynamic pressures, a positive velocity gradient tends to enhance the lift.
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.
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.
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.
Demonstration of a Low-Lift Heat Pump for High-Power Spacecraft Thermal Control
NASA Astrophysics Data System (ADS)
Grzyll, Lawrence R.
2006-01-01
This paper describes the development and demonstration of a prototype low-lift heat pump for high-power spacecraft thermal control The low-lift heat pump was designed to provide 25 kW of cooling at 303 K and transport this waste heat to a radiator for heat rejection. To accomplish this, a demonstration heat pump with an evaporation temperature of 298 K and a condensing temperature of 301 K was designed and built. HFC-227ea was the working fluid. This effort resulted in optimization of the centrifugal compressor impeller, diffuser, and shroud designs through extensive experimental testing. The detailed design of a magnetic bearing centrifugal compressor was completed. A prototype heat pump thermal control system was designed and fabricated which contained prototypical cold plate and condenser designs. This prototype system was extensively tested and demonstrated to measure performance parameters such as power consumption, cooling capacity, system size and mass, and other key parameters. Finally, the experimental performance was input into the theoretical trade study allowing for a comparison of the actual performance of the low-lift heat pump to a single-phase pumped loop. Inputting the experimental low-lift heat pump performance into the trade study showed that the low-lift heat pump still has lower system mass than the single-phase pumped loop for all space temperatures considered. The experimental results very closely match the theoretical results used in the trade study.
Aerodynamic Effects Caused by Icing of an Unswept NACA 65A004 Airfoil
NASA Technical Reports Server (NTRS)
Gray, Vernon H.; vonGlahn, Uwe H.
1958-01-01
The effects of ice formations on the section lift, drag, and pitching-moment coefficients of an unswept NACA 65A004 airfoil section of 6-foot chord were studied.. The magnitude of the aerodynamic penalties was primarily a function of the shape and size of the ice formation near the leading edge of the airfoil. The exact size and shape of the ice formations were determined photographically and found to be complex functions of the operating and icing conditions. In general, icing of the airfoil at angles of attack less than 40 caused large increases in section drag coefficients (as much as 350 percent in 8 minutes of heavy glaze icing), reductions in section lift coefficients (up to 13 percent), and changes in the pitching-moment coefficient from diving toward climbing moments. At angles of attack greater than 40 the aerodynamic characteristics depended mainly on the ice type. The section drag coefficients generally were reduced by the addition of rime ice (by as much as 45 percent in 8 minutes of icing). In glaze icing, however, the drag increased at these angles of attack. The section lift coefficients were variably affected by rime-ice formations; however, in glaze icing, lift increases at high angles of attack amounted to as much as 9 percent for an icing time of 8 minutes. Pitching-moment-coefficient changes in icing conditions were somewhat erratic and depended on the icing condition. Rotation of the iced airfoil to angles of attack other than that at which icing occurred caused sufficiently large changes in the pitching-moment coefficient that, in flight, rapid corrections in trim might be required in order to avoid a hazardous situation.
NASA Technical Reports Server (NTRS)
Vinci, Samuel, J.
2012-01-01
This report is the third part of a three-part final report of research performed under an NRA cooperative Agreement contract. The first part was published as NASA/CR-2012-217415. The second part was published as NASA/CR-2012-217416. The study of the very high lift low-pressure turbine airfoil L1A in the presence of unsteady wakes was performed computationally and compared against experimental results. The experiments were conducted in a low speed wind tunnel under high (4.9%) and then low (0.6%) freestream turbulence intensity for Reynolds number equal to 25,000 and 50,000. The experimental and computational data have shown that in cases without wakes, the boundary layer separated without reattachment. The CFD was done with LES and URANS utilizing the finite-volume code ANSYS Fluent (ANSYS, Inc.) under the same freestream turbulence and Reynolds number conditions as the experiment but only at a rod to blade spacing of 1. With wakes, separation was largely suppressed, particularly if the wake passing frequency was sufficiently high. This was validated in the 3D CFD efforts by comparing the experimental results for the pressure coefficients and velocity profiles, which were reasonable for all cases examined. The 2D CFD efforts failed to capture the three dimensionality effects of the wake and thus were less consistent with the experimental data. The effect of the freestream turbulence intensity levels also showed a little more consistency with the experimental data at higher intensities when compared with the low intensity cases. Additional cases with higher wake passing frequencies which were not run experimentally were simulated. The results showed that an initial 25% increase from the experimental wake passing greatly reduced the size of the separation bubble, nearly completely suppressing it.
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.
Analysis of a jet-controlled high-lift hydrofoil with a flap
S. H. Rhee; S.-E. Kim; H. Ahn; J. Oh; H. Kim
2003-01-01
A jet-controlled high-lift hydrofoil with a flap is investigated using both experimental and computational methods. Experiments were carried out in a cavitation tunnel to measure forces and moment acting on the hydrofoil, and surface pressure distribution. The measured data show the feasibility of such a device for marine applications. Computational studies have also been carried out in parallel with the
Wind tunnel wall interference in V/STOL and high lift testing: A selected, annotated bibliography
NASA Technical Reports Server (NTRS)
Tuttle, M. H.; Mineck, R. E.; Cole, K. L.
1986-01-01
This bibliography, with abstracts, consists of 260 citations of interest to persons involved in correcting aerodynamic data, from high lift or V/STOL type configurations, for the interference arising from the wind tunnel test section walls. It provides references which may be useful in correcting high lift data from wind tunnel to free air conditions. References are included which deal with the simulation of ground effect, since it could be viewed as having interference from three tunnel walls. The references could be used to design tests from the standpoint of model size and ground effect simulation, or to determine the available testing envelope with consideration of the problem of flow breakdown. The arrangement of the citations is chronological by date of publication in the case of reports or books, and by date of presentation in the case of papers. Included are some documents of historical interest in the development of high lift testing techniques and wall interference correction methods. Subject, corporate source, and author indices, by citation numbers, have been provided to assist the users. The appendix includes citations of some books and documents which may not deal directly with high lift or V/STOL wall interference, but include additional information which may be helpful.
Simplified dragonfly airfoil aerodynamics at Reynolds numbers below 8000
NASA Astrophysics Data System (ADS)
Levy, David-Elie; Seifert, Avraham
2009-07-01
Effective aerodynamics at Reynolds numbers lower than 10 000 is of great technological interest and a fundamental scientific challenge. The current study covers a Reynolds number range of 2000-8000. At these Reynolds numbers, natural insect flight could provide inspiration for technology development. Insect wings are commonly characterized by corrugated airfoils. In particular, the airfoil of the dragonfly, which is able to glide, can be used for two-dimensional aerodynamic study of fixed rigid wings. In this study, a simplified dragonfly airfoil is numerically analyzed in a steady free-stream flow. The aerodynamic performance (such as mean and fluctuating lift and drag), are first compared to a "traditional" low Reynolds number airfoil: the Eppler-E61. The numerical results demonstrate superior performances of the corrugated airfoil. A series of low-speed wind and water tunnel experiments were performed on the corrugated airfoil, to validate the numerical results. The findings indicate quantitative agreement with the mean wake velocity profiles and shedding frequencies while validating the two dimensionality of the flow. A flow physics numerical study was performed in order to understand the underlying mechanism of corrugated airfoils at these Reynolds numbers. Airfoil shapes based on the flow field characteristics of the corrugated airfoil were built and analyzed. Their performances were compared to those of the corrugated airfoil, stressing the advantages of the latter. It was found that the flow which separates from the corrugations and forms spanwise vortices intermittently reattaches to the aft-upper arc region of the airfoil. This mechanism is responsible for the relatively low intensity of the vortices in the airfoil wake, reducing the drag and increasing the flight performances of this kind of corrugated airfoil as compared to traditional low Reynolds number airfoils such as the Eppler E-61.
Tests in the variable-density wind tunnel of the NACA 23012 airfoil with plain and split flaps
NASA Technical Reports Server (NTRS)
Abbott, Ira H; Greenberg, Harry
1939-01-01
Section characteristics for use in wing design are presented for the NACA 23012 airfoil with plain and split flaps of 20 percent wing chord at a value of the effective Reynolds number of about 8,000,000. The flap deflections covered a range from 60 degrees upward to 75 degrees downward for the plain flap and from neutral to 90 degrees downward for the split flap. The split flap was aerodynamically superior to the plain flap in producing high maximum lift coefficients and in having lower profile-drag coefficients at high lift coefficients.
Navier-Stokes Analysis of a High Wing Transport High-Lift Configuration with Externally Blown Flaps
NASA Technical Reports Server (NTRS)
Slotnick, Jeffrey P.; An, Michael Y.; Mysko, Stephen J.; Yeh, David T.; Rogers, Stuart E.; Roth, Karlin; Baker, M.David; Nash, S.
2000-01-01
Insights and lessons learned from the aerodynamic analysis of the High Wing Transport (HWT) high-lift configuration are presented. Three-dimensional Navier-Stokes CFD simulations using the OVERFLOW flow solver are compared with high Reynolds test data obtained in the NASA Ames 12 Foot Pressure Wind Tunnel (PWT) facility. Computational analysis of the baseline HWT high-lift configuration with and without Externally Blown Flap (EBF) jet effects is highlighted. Several additional aerodynamic investigations, such as nacelle strake effectiveness and wake vortex studies, are presented. Technical capabilities and shortcomings of the computational method are discussed and summarized.
Tests of airfoils designed to delay the compressibility burble
NASA Technical Reports Server (NTRS)
Stack, John
1943-01-01
Fundamental investigations of compressibility phenomena for airfoils have shown that serious adverse changes of aerodynamic characteristics occur as the local speed over the surface exceeds the local speed of sound. These adverse changes have been delayed to higher free-stream speeds by development of suitable airfoil shapes. The method of deriving such airfoil shapes is described, and aerodynamic data for a wide range of Mach numbers obtained from tests of these airfoils in the Langley 24-inch high-speed tunnel are presented. These airfoils, designated the NACA 16-series, have increased critical Mach number. The same methods by which these airfoils have been developed are applicable to other airplane components.
Numerical Calculations of 3-D High-Lift Flows and Comparison with Experiment
NASA Technical Reports Server (NTRS)
Compton, William B, III
2015-01-01
Solutions were obtained with the Navier-Stokes CFD code TLNS3D to predict the flow about the NASA Trapezoidal Wing, a high-lift wing composed of three elements: the main-wing element, a deployed leading-edge slat, and a deployed trailing-edge flap. Turbulence was modeled by the Spalart-Allmaras one-equation turbulence model. One case with massive separation was repeated using Menter's two-equation SST (Menter's Shear Stress Transport) k-omega turbulence model in an attempt to improve the agreement with experiment. The investigation was conducted at a free stream Mach number of 0.2, and at angles of attack ranging from 10.004 degrees to 34.858 degrees. The Reynolds number based on the mean aerodynamic chord of the wing was 4.3 x 10 (sup 6). Compared to experiment, the numerical procedure predicted the surface pressures very well at angles of attack in the linear range of the lift. However, computed maximum lift was 5% low. Drag was mainly under predicted. The procedure correctly predicted several well-known trends and features of high-lift flows, such as off-body separation. The two turbulence models yielded significantly different solutions for the repeated case.
NASA Astrophysics Data System (ADS)
Vadillo, Jose L.
2005-07-01
Active control of flow over airfoils is currently an area of heightened interest in the aerospace community because of its potential in reducing drag, eliminating separation at high angles of attack, and modulating the aerodynamic forces and moments. We study these possibilities by performing several numerical simulations. Numerical simulations are performed by employing an Unsteady Reynolds-Averaged Navier-Stokes (URANS) equations solver in conjunction with a two-equation Shear-Stress-Transport (SST) turbulence model. In particular, the computations are performed for the following three classes of flows: (1) Subsonic flow past a 24% thick Clark-Y airfoil with a triangular bump on the upper surface with and without a synthetic jet actuator. The goal is to perform numerical simulations of this experimentally observed fluidic modification of airfoil pressure distributions leading to reduced pressure drag. The computations are compared with experiments performed at Georgia Tech. (2) Transonic flow past a NACA64A010 airfoil with a synthetic jet actuator. The goal is to control the shock/boundary layer interaction on the airfoil using a synthetic jet actuator to reduce drag as well to achieve desired modulation of aerodynamic forces and moments. (3) Subsonic flow past a commercial supercritical airfoil leveraging the presence of a Gurney flap with a synthetic jet actuator. The goal is again to improve the aerodynamic performance (increase or maintain lift and reduce drag) by using a synthetic jet actuator integrated in a bump on the pressure surface of the airfoil near the trailing edge. The computations are compared with the experiments performed at Georgia Tech. The computations as well as the experiments show the feasibility of active flow control in reducing the drag of airfoils and in achieving the desired modulation of aerodynamic forces and moments.
Computational evaluation of an airfoil with a Gurney flap
NASA Technical Reports Server (NTRS)
Jang, Cory S.; Ross, James C.; Cummings, Russell M.
1992-01-01
A 2D numerical investigation was performed to determine the effect of a Gurney flap on a NACA 4412 airfoil. A Gurney flap is a flat plate on the order of 1 to 3 percent of the airfoil chord length, oriented perpendicular to the airfoil chord line and located at the trailing edge of the airfoil. An incompressible Navier Stokes code, INS2D, was used to calculate the flow field about the airfoil. The fully turbulent results were obtained using the Baldwin-Barth one-equation turbulence model. Gurney flap sizes of 0.5 , 1, 1.25, 1.5, 2, and 3 percent of the airfoil chord were studied. Computational results were compared with experimental results where possible. The numerical solutions show that the Gurney flap increases airfoil lift coefficient with only a slight increase in drag coefficient. Use of a 1.5 percent chord Gurney flap increases the maximum lift coefficient by approximately 0.3 and decreases the angle of attack for a given lift coefficient by more than 3 deg. The numerical solutions exhibit detailed flow structures at the trialing edge and provide a possible explanation for the increased aerodynamic performance.
NASA Technical Reports Server (NTRS)
Mcghee, R. J.; Beasley, W. D.
1979-01-01
Wind tunnel tests were conducted to determine the low speed, two dimensional aerodynamic characteristics of a 13percent thick medium speed airfoil designed for general aviation applications. The results were compared with data for the 13 percent thick low speed airfoil. The tests were conducted over a Mach number range from 0.10 to 0.32, a chord Reynolds number range from 2.0 x 10 to the 6th power to 12.0 x 10 to the 6th power, and an angle of attack frange from about -8 deg to 10 deg. The objective of retaining good high-lift low speed characteristics for an airfoil designed to have good medium speed cruise performance was achieved.
M. S. Chandrasekhara
2010-01-01
A variable droop leading edge (VDLE) airfoil was successfully used to control compressible dynamic stall and its associated adverse pitching moment variations. But, the price for this success was a 10% loss of lift. A Gurney flap was then attached normally to the airfoil pressure surface at its trailing edge to recover this “lost-lift”. Gurney flaps are seldom used in
Status of CFD for LaRC's HSR high-lift program
NASA Technical Reports Server (NTRS)
Waggoner, Edgar G.; South, Jerry C., Jr.
1992-01-01
The viewgraphs for a status report for using computational fluid dynamics (CFD) for NASA Langley's Research Center High Speed Research High-Lift Program are provided. The objectives of CFD applications, the approach to developing appropriate CFD codes, and the dominant flow mechanisms are outlined. Possible CFD codes are compared; graphs comparing force coefficients, surface pressure distribution, and force are included. The plans for using CFD applications are listed.
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.
Aero-Mechanical Design Methodology for Subsonic Civil Transport High-Lift Systems
NASA Technical Reports Server (NTRS)
vanDam, C. P.; Shaw, S. G.; VanderKam, J. C.; Brodeur, R. R.; Rudolph, P. K. C.; Kinney, D.
2000-01-01
In today's highly competitive and economically driven commercial aviation market, the trend is to make aircraft systems simpler and to shorten their design cycle which reduces recurring, non-recurring and operating costs. One such system is the high-lift system. A methodology has been developed which merges aerodynamic data with kinematic analysis of the trailing-edge flap mechanism with minimum mechanism definition required. This methodology provides quick and accurate aerodynamic performance prediction for a given flap deployment mechanism early on in the high-lift system preliminary design stage. Sample analysis results for four different deployment mechanisms are presented as well as descriptions of the aerodynamic and mechanism data required for evaluation. Extensions to interactive design capabilities are also discussed.
NASA Technical Reports Server (NTRS)
Wentz, W. H., Jr.; Fiscko, K. A.
1979-01-01
Force and surface pressure distributions were measured for the 21% LS(1)-0421 modified airfoil fitted with 20% aileron, 25% slotted flap and 10% slot lip spoiler. All tests were conducted at a Reynolds number of 2.2 x 10 to the 6th power and a Mach number of 0.13. The lift, drag, pitching moments, control surface normal force and hinge moments, and surface pressure distributions are included in the results. Incremental performance of flap and aileron are discussed and compared to the GA(W)-2 airfoil. Spoiler control which shows a slight reversal tendency at high alpha, is examined.
NASA Technical Reports Server (NTRS)
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.
Aerodynamic Analysis of Trailing Edge Enlarged Wind Turbine Airfoils
NASA Astrophysics Data System (ADS)
Xu, Haoran; Shen, Wenzhong; Zhu, Weijun; Yang, Hua; Liu, Chao
2014-06-01
The aerodynamic performance of blunt trailing edge airfoils generated from the DU- 91-W2-250, DU-97-W-300 and DU-96-W-350 airfoils by enlarging the thickness of trailing edge symmetrically from the location of maximum thickness to chord to the trailing edge were analyzed by using CFD and RFOIL methods at a chord Reynolds number of 3 × 106. The goal of this study is to analyze the aerodynamic performance of blunt trailing edge airfoils with different thicknesses of trailing edge and maximum thicknesses to chord. The steady results calculated by the fully turbulent k-? SST, transitional k-? SST model and RFOIL all show that with the increase of thickness of trailing edge, the linear region of lift is extended and the maximum lift also increases, the increase rate and amount of lift become limited gradually at low angles of attack, while the drag increases dramatically. For thicker airfoils with larger maximum thickness to chord length, the increment of lift is larger than that of relatively thinner airfoils when the thickness of blunt trailing edge is increased from 5% to 10% chord length. But too large lift can cause abrupt stall which is profitless for power output. The transient characteristics of blunt trailing edge airfoils are caused by blunt body vortices at low angles of attack, and by the combined effect of separation and blunt body vortices at large angles of attack. With the increase of thickness of blunt trailing edge, the vibration amplitudes of lift and drag curves increase. The transient calculations over-predict the lift at large angles of attack and drag at all angles of attack than the steady calculations which is likely to be caused by the artificial restriction of the flow in two dimensions.
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 is hypothesized, and verified with experimental and computational data, that LEV formation always occurs at the same critical value of LESP irrespective of motion kinematics. Further, the applicability of the LESP criterion in influencing the occurrence of LEV formation is demonstrated. To model the growth and convection of leading-edge vortices, the unsteady thin-airfoil theory is augmented with discrete-vortex shedding from the leading edge. The LESP criterion is used to predict and modulate the shedding of leading-edge vorticity. Comparisons with experiments and CFD for test-cases with different airfoils, Reynolds numbers and motion kinematics, show that the method performs remarkably well in predicting force coefficients and flowfields for unsteady flows. The use of a single empirical parameter - the critical LESP value, allows the determination of onset, growth and termination of leading-edge vortex shedding. In the final part of the research, the discrete-vortex model is extended to flows where the freestream velocity is varying or small in comparison with motion velocity. With this extension, the method is made applicable to a larger set of 2D flows such as perching and hovering maneuvers, gusts, and sinusoidally varying freestream. Abstractions of perching and hovering are designed as test cases and used to validate the low-order model's performance in highly-unsteady, vortex-dominated flows. Alongside development of the low-order methodology, several features of unsteady flows are studied and analyzed with the aid of CFD and experiments. While remaining computationally inexpensive and retaining the essential flow-physics, the method is seen to be successful in prediction of both force coefficients and flow histories.
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 Control of Separation from the Slat Shoulder of a Supercritical Airfoil
NASA Technical Reports Server (NTRS)
Pack, LaTunia G.; Schaeffler, Norman W.; Yao, Chung-Sheng; Seifert, Avi
2002-01-01
Active flow control in the form of zero-mass-flux excitation was applied at the slat shoulder of a simplified high-lift airfoil to delay flow separation. The NASA Energy Efficient Transport (EET) supercritical airfoil was equipped with a 15% chord simply hinged leading edge slat and a 25% chord simply hinged trailing edge flap. The cruise configuration data was successfully reproduced, repeating previous experiments. The effects of flap and slat deflection angles on the performance of the airfoil integral parameters were quantified. Detailed flow features were measured as well, in an attempt to identify optimal actuator placement. The measurements included: steady and unsteady model and tunnel wall pressures, wake surveys, arrays of surface hot-films, flow visualization and Particle Image Velocimetry (PIV). High frequency periodic excitation was applied to delay the occurrence of slat stall and improve the maximum lift by 10 to 15%. Low frequency amplitude modulation was used to reduce the oscillatory momentum coefficient by roughly 50% with similar aerodynamic performance.
Characterization of the Effect of Wing Surface Instrumentation on UAV Airfoil Performance
NASA Technical Reports Server (NTRS)
Ratnayake, Nalin A.
2009-01-01
Recently proposed flight research at NASA Dryden Flight Research Center (DFRC) has prompted study into the aerodynamic effects of modifications made to the surfaces of laminar airfoils. The research is focused on the high-aspect ratio, laminar-flow type wings commonly found on UAVs and other aircraft with a high endurance requirement. A broad range of instrumentation possibilities, such as structural, pressure, and temperature sensing devices may require the alteration of the airfoil outer mold line as part of the installation process. This study attempts to characterize the effect of installing this additiona1 instrumentation on key airfoil performance factors, such as transition location, lift and drag curves, and stall point. In particular, the general case of an airfoil that is channeled in the spanwise direction is considered, and the impact on key performance characteristics is assessed. Particular attention is focused on exploring the limits of channel depth and low-Reynolds number on performance and stall characteristics. To quantify the effect of increased skin friction due to premature transition caused by protruding or recessed instrumentation, two simplified, conservative scenarios are used to consider two potential sources of diaturbance: A) that leading edge alterations would cause linearly expanding areas (triangles) of turbulent flow on both surfaces of the wing upstream of the natural transition point, and B) that a channel or bump on the upper surface would trip turbulent flow across the whole upper surface upstream of the natural transition point. A potentially more important consideration than the skin friction drag increment is the change in overall airfoil performance due to the installation of instrumentation along most of the wingspan. To quantify this effect, 2D CFD simulations of the flow over a representative mid-span airfoil section were conducted in order to assess the change in lift and drag curves for the airfoil in the presence of disturbed flow due to the installed instrumentation. A discussion as to the impact on high-altitude and low-speed operation of this and similar aircraft is provided.
Effect of the Gurney flap on a NACA 23012 airfoil
Neung-Soo Yoo
2000-01-01
A numerical investigation was performed to determine the effect of the Gurney flap on a NACA 23012 airfoil. A Navier-Stokes\\u000a code, RAMPANT, was used to calculate the flow field about the airfoil. Fully-turbulent results were obtained using the standardk-? two-equation turbulence model. The numerical solutions showed that the Gurney flap increased both lift and drag. These\\u000a results suggested that the
Stiffness characteristics of airfoils under pulse loading
Kevin Eugene Turner
2010-01-01
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
Two-and three-dimensional unsteady lift problems in high-speed flight
NASA Technical Reports Server (NTRS)
Lomax, Harvard; Heaslet, Max A; Fuller, Franklyn B; Sluder, Loma
1952-01-01
The problem of transient lift on two- and three-dimensional wings flying at high speeds is discussed as a boundary-value problem for the classical wave equation. Kirchoff's formula is applied so that the analysis is reduced, just as in the steady state, to an investigation of sources and doublets. The applications include the evaluation of indicial lift and pitching-moment curves for two-dimensional sinking and pitching wings flying at Mach numbers equal to 0, 0.8, 1.0, 1.2 and 2.0. Results for the sinking case are also given for a Mach number of 0.5. In addition, the indicial functions for supersonic-edged triangular wings in both forward and reverse flow are presented and compared with the two-dimensional values.
NASA Astrophysics Data System (ADS)
Ilie, Marcel
In helicopters, vortices (generated at the tip of the rotor blades) interact with the next advancing blades during certain flight and manoeuvring conditions, generating undesirable levels of acoustic noise and vibration. These Blade-Vortex Interactions (BVIs), which may cause the most disturbing acoustic noise, normally occur in descent or high-speed forward flight. Acoustic noise characterization (and potential reduction) is one the areas generating intensive research interest to the rotorcraft industry. Since experimental investigations of BVI are extremely costly, some insights into the BVI or AVI (2-D Airfoil-Vortex Interaction) can be gained using Computational Fluid Dynamics (CFD) numerical simulations. Numerical simulation of BVI or AVI has been of interest to CFD for many years. There are still difficulties concerning an accurate numerical prediction of BVI. One of the main issues is the inherent dissipation of CFD turbulence models, which severely affects the preservation of the vortex characteristics. Moreover this is not an issue only for aerodynamic and aeroacoustic analysis but also for aeroelastic investigations as well, especially when the strong (two-way) aeroelastic coupling is of interest. The present investigation concentrates mainly on AVI simulations. The simulations are performed for Mach number, Ma = 0.3, resulting in a Reynolds number, Re = 1.3 x 106, which is based on the chord, c, of the airfoil (NACA0012). Extensive literature search has indicated that the present work represents the first comprehensive investigation of AVI using the LES numerical approach, in the rotorcraft research community. The major factor affecting the aerodynamic coefficients and aeroacoustic field as a result of airfoil-vortex interaction is observed to be the unsteady pressure generated at the location of the interaction. The present numerical results show that the aerodynamic coefficients (lift, moment, and drag) and aeroacoustic field are strongly dependent on the airfoil-vortex vertical miss-distance, airfoil angle of attack, vortex characteristics, and aeroelastic response of airfoil to airfoil-vortex interaction. A decay of airfoil-vortex interactions with the increase of vertical miss-distance and angle of attack was observed. Also, a decay of airfoil-vortex interactions is observed for the case of a flexible structure when compared with the case of a rigid structure. The decay of vortex core size produces a decrease in the aerodynamic coefficients.
NASA Technical Reports Server (NTRS)
Rumsey, Christopher L.
1987-01-01
The flux splitting method of Van Leer (1982) and an implicit, upwind-biased, finite-volume scheme of Rumsey (1985) for solving the two-dimensional Reynolds-averaged Navier-Stokes equations were applied to analyze five airfoil geometries (the NACA 0012, 12-percent Joukowski, NACA 4412, NACA 65-015, and LS1-0417 sections), at high angles of attack. The effect of grid density and grid extent on the Strouhal number of the unsteady flow was computed, as well as the effect of variations in Mach number, Reynolds number, and angle of attack. It was found, that past an angle of attack of 20 deg, Strouhal numbers were independent of angle of attack (at a constant value of 0.115), agreeing well with the experimental findings of Tyler (1931).
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.
Unsteady pressure measurements on a supercritical airfoil at high Reynolds numbers
NASA Technical Reports Server (NTRS)
Hess, R. W.
1989-01-01
Steady and unsteady pressures were measured on a 14 percent supercritical airfoil at transonic Mach numbers at Reynolds numbers from 6,000,000 to 35,000,000. Instrumentation techniques were developed to measure unsteady pressures in a cryogenic tunnel at flight Reynolds numbers. Experimental steady data, corrected for wall effects show very good agreement with calculations from a full potential code with an interacted boundary layer. The steady and unsteady pressures both show a shock position that is dependent on Reynolds number. For a supercritical pressure distribution at a chord Reynolds number of 35,000,000 laminar flow was observed between the leading edge and the shock wave at 45 percent chord.
Wind-tunnel tests of a Clark Y wing with a narrow auxiliary airfoil in different positions
NASA Technical Reports Server (NTRS)
Weick, Fred E; Bamber, Millard J
1933-01-01
Aerodynamic force tests were made on a combination of a Clark Y wing and a narrow auxiliary airfoil to find the best location of the auxiliary airfoil with respect to the main wing. The auxiliary was a highly cambered airfoil of medium thickness having a chord 14.5 per cent that of the main wing. It was tested in 141 different positions ahead of, above, and behind the nose portion of the main wing, the range of the test points being extended until the best aerodynamic conditions were covered. A range of positions was found in which the combination of main wing and auxiliary gave substantially greater aerodynamic efficiency and higher maximum lift coefficients (based on total area) than the main Clark Y wing alone. In the optimum position tested, considering both the maximum lift and the speed-range ratio, the combination of main wing and auxiliary gave an increase in the maximum lift coefficient of 32 per cent together with an increase in the ratio of 21 per cent of the respective values for the main Clark Y wing alone.
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.
Flight-measured lift and drag characteristics of a large, flexible, high supersonic cruise airplane
NASA Technical Reports Server (NTRS)
Arnaiz, H. H.
1977-01-01
Flight measurements of lift, drag, and angle of attack were obtained for the XB-70 airplane, a large, flexible, high supersonic cruise airplane. This airplane had a length of over 57 meters, a takeoff gross mass of over 226,800 kilograms, and a design cruise speed of Mach 3 at an altitude of 21,340 meters. The performance measurements were made at Mach numbers from 0.72 to 3.07 and altitudes from approximately 7620 meters to 21,340 meters. The measurements were made to provide data for evaluating the techniques presently being used to design and predict the performance of aircraft in this category. Such performance characteristics as drag polars, lift-curve slopes, and maximum lift-to-drag ratios were derived from the flight data. The base drag of the airplane, changes in airplane drag with changes in engine power setting at transonic speeds, and the magnitude of the drag components of the propulsion system are also discussed.
The use of a panel code on high lift configurations of a swept forward wing
NASA Technical Reports Server (NTRS)
Scheib, J. S.; Sandlin, D. R.
1985-01-01
A study was done on high lift configurations of a generic swept forward wing using a panel code prediction method. A survey was done of existing codes available at Ames, frow which the program VSAERO was chosen. The results of VSAERO were compared with data obtained from the Ames 7- by 10-foot wind tunnel. The results of the comparison in lift were good (within 3.5%). The comparison of the pressure coefficients was also good. The pitching moment coefficients obtained by VSAERO were not in good agreement with experiment. VSAERO's ability to predict drag is questionable and cannot be counted on for accurate trends. Further studies were done on the effects of a leading edge glove, canards, leading edge sweeps and various wing twists on spanwise loading and trim lift with encouraging results. An unsuccessful attempt was made to model spanwise blowing and boundary layer control on the trailing edge flap. The potential results of VSAERO were compared with experimental data of flap deflections with boundary layer control to check the first order effects.
Alexander Ivanchin
2010-02-06
It is shown that one of the causes for tornado is Tornado Lift. At increasing vortex diameter its kinetic energy decreases to keep the moment of momentum constant. A kinetic energy gradient of such vortex is Tornado Lift. Evaluation shows that contribution of Tornado Lift in air lifting in a tornado is comparable to buoyancy according to the order of magnitude.
Ivanchin, Alexander
2010-01-01
It is shown that one of the causes for tornado is Tornado Lift. At increasing vortex diameter its kinetic energy decreases to keep the moment of momentum constant. A kinetic energy gradient of such vortex is Tornado Lift. Evaluation shows that contribution of Tornado Lift in air lifting in a tornado is comparable to buoyancy according to the order of magnitude.
Flow Prediction Around an Oscillating NACA0012 Airfoil at Re = 1,000,000
O. Frederich; U. Bunge; C. Mockett; F. Thiele
2009-01-01
\\u000a The maximum obtainable lift of a rotationally-oscillating airfoil is significantly higher than in the static or quasi-static\\u000a case. The correct prediction of dynamic stall as the basis of the dynamically increased lift is essential to quantify the\\u000a time-dependent load on the airfoil structure. This study applies unsteady RANS (URANS) and detached-eddy simulation (DES)\\u000a with various turbulence models and parameter variations
S. Spoelstra; W. G. Haije; J. W. Dijkstra
2002-01-01
This paper presents the results of a techno-economic feasibility study on two high-temperature high-lift chemical heat pumps for upgrading industrial waste heat. The study was set up in order to select the most promising heat pump concept for further development. First, a market study is performed to assess the amount of waste heat and the temperature dependence thereof. Based on
Mach number validation of a new zonal CFD method (ZAP2D) for airfoil simulations
NASA Technical Reports Server (NTRS)
Strash, Daniel J.; Summa, Michael; Yoo, Sungyul
1991-01-01
A closed-loop overlapped velocity coupling procedure has been utilized to combine a two-dimensional potential-flow panel code and a Navier-Stokes code. The fully coupled two-zone code (ZAP2D) has been used to compute the flow past a NACA 0012 airfoil at Mach numbers ranging from 0.3 to 0.84 near the two-dimensional airfoil C(lmax) point for a Reynolds number of 3 million. For these cases, the grid domain size can be reduced to 3 chord lengths with less than 3-percent loss in accuracy for freestream Mach numbers through 0.8. Earlier validation work with ZAP2D has demonstrated a reduction in the required Navier-Stokes computation time by a factor of 4 for subsonic Mach numbers. For this more challenging condition of high lift and Mach number, the saving in CPU time is reduced to a factor of 2.
Numerical study of incompressible slightly viscous flow past blunt bodies and airfoils
Cheer, A.Y.L.
1981-05-01
A grid free numerical method is used to simulate incompressible flow at high Reynolds number. The numerical method simulates the flow inside the boundary layer by vortex sheets and the flow outside this layer by vortex blobs. The algorithm produces a smooth transition between the sheets and the blobs. The accuracy of this hybrid numerical method is tested in several numerical experiments. In the first experiment, the algorithm is used to simulate slightly viscous flow past a circular cylinder. In the second experiment, the algorithm is used to simulate flow past a Joukowski airfoil at various angles of attack. In the latter case, there is no evidence of blow-up of the flow at the trailing edge of the airfoil. In both experiments, the calculated flow and its functionals (such as lift and drag coefficients) are in good agreement with both theoretical results and wind tunnel experiments.
Low-drag airfoils for transonic flow of dense gases
NASA Astrophysics Data System (ADS)
Rusak, Z.; Wang, C.-W.
Low-drag airfoils for transonic flow of BZT gases are constructed through a nonlinear small-disturbance theory on this topic. This kind of transonic flow is characterized by the high nonlinearity of the fluid thermodynamic behavior that is closely coupled with its compressible flow dynamics. Utilizing BZT gases may result in low drag exerted on airfoils operating at high transonic speeds. This advantage may be further improved by a proper design of the airfoils' shape. The modified airfoils are characterized by arcs along which the flow is sonic and that connect to a sharp tail. These airfoils may give the highest free stream Mach number for which the flow is nowhere supersonic. Using an analysis in the hodograph (velocity) plane, an analytical formula is derived to describe the sonic arc. Numerical computations using the Euler equations demonstrate that the modified airfoils have significantly lower pressure drag against standard airfoils and their critical Mach number is increased to higher values.
NASA Astrophysics Data System (ADS)
Okayasu, Shimpei; Arakawa, Yosuke; Sato, Kotaro; Shakouchi, Toshihiko; Furuya, Okitsugu
The present paper describes the influence of ground and ceiling effect on high-lift device using a circular cylinder with tangential blowing. The experiments and numerical simulations were performed under the condition of Re = 2.7× 104. In the experiments, the test, such as the time-mean surface pressure measurements on the cylinder and flow visualizations were carried out for various momentum coefficients and clearances between the cylinder and the ground (or ceiling). In the numerical simulation, a commercially available computer code, SCRYU/Tetra, was used. Qualitative agreements between the experimental and numerical results were obtained with respect to the flow fields and the lift characteristics. The experimental and numerical results show that ground effect can enhance lift-drag ratio except stalling region, and that ceiling effect can not enhance lift-drag ratio.
Turbulent intensity and Reynolds number effects on an airfoil at low Reynolds numbers
NASA Astrophysics Data System (ADS)
Wang, S.; Zhou, Y.; Alam, Md. Mahbub; Yang, H.
2014-11-01
This work investigates the aerodynamics of a NACA 0012 airfoil at the chord-based Reynolds numbers (Rec) from 5.3 × 103 to 2.0 × 104. The lift and drag coefficients, CL and CD, of the airfoil, along with the flow structure, were measured as the turbulent intensity Tu of oncoming flow varies from 0.6% to 6.0%. The analysis of the present data and those in the literature unveils a total of eight distinct flow structures around the suction side of the airfoil. Four Rec regimes, i.e., the ultra-low (<1.0 × 104), low (1.0 × 104-3.0 × 105), moderate (3.0 × 105-5.0 × 106), and high Rec (>5.0 × 106), are proposed based on their characteristics of the CL-Rec relationship and the flow structure. It has been observed that Tu has a more pronounced effect at lower Rec than at higher Rec on the shear layer separation, reattachment, transition, and formation of the separation bubble. As a result, CL, CD, CL/CD and their dependence on the airfoil angle of attack all vary with Tu. So does the critical Reynolds number Rec,cr that divides the ultra-low and low Rec regimes. It is further noted that the effect of increasing Tu bears similarity in many aspects to that of increasing Rec, albeit with differences. The concept of the effective Reynolds number Rec,eff advocated for the moderate and high Rec regimes is re-evaluated for the low and ultra-low Rec regimes. The Rec,eff treats the non-zero Tu effect as an addition of Rec and is determined based on the presently defined Rec,cr. It has been found that all the maximum lift data from both present measurements and previous reports collapse into a single curve in the low and ultra-low Rec regimes if scaled with Rec,eff.
Two-dimensional unsteady lift problems in supersonic flight
NASA Technical Reports Server (NTRS)
Heaslet, Max A; Lomax, Harvard
1949-01-01
The variation of pressure distribution is calculated for a two-dimensional supersonic airfoil either experiencing a sudden angle-of-attack change or entering a sharp-edge gust. From these pressure distributions the indicial lift functions applicable to unsteady lift problems are determined for two cases. Results are presented which permit the determination of maximum increment in lift coefficient attained by an unrestrained airfoil during its flight through a gust. As an application of these results, the minimum altitude for safe flight through a specific gust is calculated for a particular supersonic wing of given strength and wing loading.
NASA Astrophysics Data System (ADS)
Ragni, D.; Ashok, A.; van Oudheusden, B. W.; Scarano, F.
2009-07-01
The present investigation assesses a procedure to extract the aerodynamic loads and pressure distribution on an airfoil in the transonic flow regime from particle image velocimetry (PIV) measurements. The wind tunnel model is a two-dimensional NACA-0012 airfoil, and the PIV velocity data are used to evaluate pressure fields, whereas lift and drag coefficients are inferred from the evaluation of momentum contour and wake integrals. The PIV-based results are compared to those derived from conventional loads determination procedures involving surface pressure transducers and a wake rake. The method applied in this investigation is an extension to the compressible flow regime of that considered by van Oudheusden et al (2006 Non-intrusive load characterization of an airfoil using PIV Exp. Fluids 40 988-92) at low speed conditions. The application of a high-speed imaging system allows the acquisition in relatively short time of a sufficient ensemble size to compute converged velocity statistics, further translated in turbulent fluctuations included in the pressure and loads calculation, notwithstanding their verified negligible influence in the computation. Measurements are performed at varying spatial resolution to optimize the loads determination in the wake region and around the airfoil, further allowing us to assess the influence of spatial resolution in the proposed procedure. Specific interest is given to the comparisons between the PIV-based method and the conventional procedures for determining the pressure coefficient on the surface, the drag and lift coefficients at different angles of attack. Results are presented for the experiments at a free-stream Mach number M = 0.6, with the angle of attack ranging from 0° to 8°.
Simulation of self-induced unsteady motion in the near wake of a Joukowski airfoil
NASA Technical Reports Server (NTRS)
Ghia, K. N.; Osswald, G. A.; Ghia, U.
1986-01-01
The unsteady Navier-Stokes analysis is shown to be capable of analyzing the massively separated, persistently unsteady flow in the post-stall regime of a Joukowski airfoil for an angle of attack as high as 53 degrees. The analysis has provided the detailed flow structure, showing the complex vortex interaction for this configuration. The aerodynamic coefficients for lift, drag, and moment were calculated. So far only the spatial structure of the vortex interaction was computed. It is now important to potentially use the large-scale vortex interactions, an additional energy source, to improve the aerodynamic performance.
NASA Astrophysics Data System (ADS)
Lee, T.; Su, Y. Y.
2012-11-01
The small magnitude lift forces generated by both a NACA 0012 airfoil and a thin flat plate at Re = 29,000 and 54,000 were determined through the line integral of velocity, obtained with particle image velocimetry, via the application of the Kutta-Joukowsky theorem. Surface pressure measurements of the NACA0012 airfoil were also obtained to validate the lift coefficient C l. The bound circulation was found to be insensitive to the size and aspect ratio of the rectangular integration loop for pre-stall angles. The present C l data were also found to agree very well with the surface pressure-determined lift coefficient for pre-stall conditions. A large variation in C l with the loop size and aspect ratio for post-stall conditions was, however, observed. Nevertheless, the present flat-plate C l data were also found to collectively agree with the published force-balance measurements at small angles of attack, despite the large disparity exhibited among the various published data at high angles. Finally, the ensemble-averaged wake velocity profiles were also used to compute the drag coefficient and, subsequently, the lift-to-drag ratio.
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.
Experimental studies of the Eppler 61 airfoil at low Reynolds numbers
NASA Technical Reports Server (NTRS)
Burns, T. F.; Mueller, T. J.
1982-01-01
The results of an experimental study to document the effects of separation and transition on the performance of an airfoil designed for low Reynolds number operation are presented. Lift, drag and flow visualization data were obtained for the Eppler 61 airfoil section for chord Reynolds numbers from about 30,000 to over 200,000. Smoke flow visualization was employed to document the boundary layer behavior and was correlated with the Eppler airfoil design and analysis computer program. Laminar separation, transition and turbulent reattachment had significant effects on the performance of this airfoil.
Static Extended Trailing Edge for Lift Enhancement: Experimental and Computational Studies
T. Liu; Q. A. Shams
A static extended trailing edge attached to a NACA0012 airfoil section is studied for achieving lift enhancement at a small drag penalty. It is indicated that the thin extended trailing edge can enhance the lift while the zero-lift drag is not significantly increased. Experiments and calculations are conducted to compare the aerodynamic characteristics of the extended trailing edge with those
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... lift to improve their appearance. This is called cosmetic or elective surgery. The eyelid lift may be done alone or ... Codere F, Tucker N. Cosmetic blepharoplasty and browplasty. In: ... Skin. 2nd ed . Philadelphia, PA: Elsevier Mosby; 2010:chap 39.
Investigation of the Kline-Fogleman airfoil section for rotor blade applications
NASA Technical Reports Server (NTRS)
Lumsdaine, E.; Johnson, W. S.; Fletcher, L. M.; Peach, J. E.
1974-01-01
Wind tunnel tests of a wedgeshaped airfoil with sharp leading edge and a spanwise step were conducted. The airfoil was tested with variations of the following parameters: (1) Reynolds number, (2) step location, (3) step shape, (4) apex angle, and (5) with the step on either the upper or lower surface. The results are compared with a flat plate and with wedge airfoils without a step having the same aspect ratio. Water table tests were conducted for flow visualization and it was determined that the flow separates from the upper surface at low angles of attack. The wind tunnel tests show that the lift/drag ratio of the airfoil is lower than for a flat plate and the pressure data show that the airfoil derives its lift in the same manner as a flat plate.
Effect of In-Flight Ice Accretion on the Performance of a Multi-Element Airfoil
NASA Technical Reports Server (NTRS)
Khodadoust, Abdollah; Dominik, Chet; Shin, Jaiwon; Miller, Dean
1995-01-01
The effects of potential in-flight ice accretion on the aerodynamic performance of a multi-element high-lift airfoil have been investigated at moderate-to-high Reynolds numbers. The investigation was conducted in the Low Turbulence Pressure Tunnel (LTPT) at NASA Langley Research Center. Simulated ice shapes obtained from earlier testing in the Icing Research Tunnel (IRT) at NASA Lewis Research Center were used on all three elements of the multi-element configuration. Incremental performance effects due to the ice accretion are presented for both smooth and rough ice accretions. Reynolds number effects on the measured performance characteristics were also assessed. The present results confirm the importance of avoiding any ice accretions on the forward element of a lifting configuration.
A numerical study of the controlled flow tunnel for a high lift model
NASA Technical Reports Server (NTRS)
Parikh, P. C.
1984-01-01
A controlled flow tunnel employs active control of flow through the walls of the wind tunnel so that the model is in approximately free air conditions during the test. This improves the wind tunnel test environment, enhancing the validity of the experimentally obtained test data. This concept is applied to a three dimensional jet flapped wing with full span jet flap. It is shown that a special treatment is required for the high energy wake associated with this and other V/STOL models. An iterative numerical scheme is developed to describe the working of an actual controlled flow tunnel and comparisons are shown with other available results. It is shown that control need be exerted over only part of the tunnel walls to closely approximate free air flow conditions. It is concluded that such a tunnel is able to produce a nearly interference free test environment even with a high lift model in the tunnel.
Three-Dimensional High-Lift Analysis Using a Parallel Unstructured Multigrid Solver
NASA Technical Reports Server (NTRS)
Mavriplis, Dimitri J.
1998-01-01
A directional implicit unstructured agglomeration multigrid solver is ported to shared and distributed memory massively parallel machines using the explicit domain-decomposition and message-passing approach. Because the algorithm operates on local implicit lines in the unstructured mesh, special care is required in partitioning the problem for parallel computing. A weighted partitioning strategy is described which avoids breaking the implicit lines across processor boundaries, while incurring minimal additional communication overhead. Good scalability is demonstrated on a 128 processor SGI Origin 2000 machine and on a 512 processor CRAY T3E machine for reasonably fine grids. The feasibility of performing large-scale unstructured grid calculations with the parallel multigrid algorithm is demonstrated by computing the flow over a partial-span flap wing high-lift geometry on a highly resolved grid of 13.5 million points in approximately 4 hours of wall clock time on the CRAY T3E.
A finite-difference method for transonic airfoil design.
NASA Technical Reports Server (NTRS)
Steger, J. L.; Klineberg, J. M.
1972-01-01
This paper describes an inverse method for designing transonic airfoil sections or for modifying existing profiles. Mixed finite-difference procedures are applied to the equations of transonic small disturbance theory to determine the airfoil shape corresponding to a given surface pressure distribution. The equations are solved for the velocity components in the physical domain and flows with embedded shock waves can be calculated. To facilitate airfoil design, the method allows alternating between inverse and direct calculations to obtain a profile shape that satisfies given geometric constraints. Examples are shown of the application of the technique to improve the performance of several lifting airfoil sections. The extension of the method to three dimensions for designing supercritical wings is also indicated.
Prediction of unsteady airfoil flows at large angles of incidence
NASA Technical Reports Server (NTRS)
Cebeci, Tuncer; Jang, H. M.; Chen, H. H.
1992-01-01
The effect of the unsteady motion of an airfoil on its stall behavior is of considerable interest to many practical applications including the blades of helicopter rotors and of axial compressors and turbines. Experiments with oscillating airfoils, for example, have shown that the flow can remain attached for angles of attack greater than those which would cause stall to occur in a stationary system. This result appears to stem from the formation of a vortex close to the surface of the airfoil which continues to provide lift. It is also evident that the onset of dynamic stall depends strongly on the airfoil section, and as a result, great care is required in the development of a calculation method which will accurately predict this behavior.
AIAA-2004-1051 Computational Investigation of Airfoils with Miniature Trailing Edge Control
Stanford University
accurate computation is implemented. Introduction The Gurney flap is a small flap used to increase the lift. Numerous wind-tunnel tests and numerical computations have been performed on airfoils with Gurney flaps.1,35,8,9 These studies confirm that de- spite their small size, Gurney flaps can significantly increase the maximum lift
Rotating cylinder design as a lifting generator
NASA Astrophysics Data System (ADS)
Asrokin, Azharrudin; Rizal Ramly, Mohammad; Halim Ahmad, Abdul
2013-12-01
The airfoil shape of a wing has always been the design to generate lift. But few realized that a simple rotating cylinder can also create lift. However, the explanation and study of how a rotating cylinder creates lift are still complex. In remote area where it is difficult for air vehicle to access, the exploration and discovery of different configuration for design concept is rather important. Due to this reason, there is a need to think of a lift generator that can produce better lift (few fold better than conventional airfoil) at lower speed to take off in a short distance of time. This paper will explain the conditions and the design of such a wing using the rotating cylinder concept that will take off in a short time and requires little takeoff and landing strip. Spokes will be attached to the cylinder to force the surrounding air to rotate along with the cylinder. This will create a vortex that hastens the speed of the air on top of the cylinder and at the same time retarding the speed of air below the cylinder. From the results, the rougher surface cylinder produces more lift when rotating and also, higher speed rotation of the cylinder greatly changes the speed of the surrounding air, thus better lift.
Lift-off PMN-PT Thick Film for High Frequency Ultrasonic Biomicroscopy.
Zhu, Benpeng; Han, Jiangxue; Shi, Jing; Shung, K Krik; Wei, Q; Huang, Yuhong; Kosec, M; Zhou, Qifa
2010-10-01
Piezoelectric 0.65Pb(Mg(1/3)Nb(2/3))O(3)-0.35PbTiO(3) (PMN-35PT) thick film with a thickness of approximately 12 µm has been deposited on the platinum buffered Si substrate via a sol-gel composite method. The separation of the film from the substrate was achieved using a wet chemical method. The lifted-off PMN-35PT thick film exhibited good dielectric and ferroelectric properties. At 1 kHz, the dielectric constant and the dielectric loss were 3,326 and 0.037, respectively, while the remnant polarization was 30.0 µC/cm(2). A high frequency single element acoustic transducer fabricated with this film showed a bandwidth at -6 dB of 63.6% at 110 MHz. PMID:21170158
Lift-off PMN-PT Thick Film for High Frequency Ultrasonic Biomicroscopy
Zhu, Benpeng; Han, Jiangxue; Shi, Jing; Shung, K. Krik; Wei, Q.; Huang, YuHong; Kosec, M.; Zhou, Qifa
2010-01-01
Piezoelectric 0.65Pb(Mg1/3Nb2/3)O3-0.35PbTiO3 (PMN-35PT) thick film with a thickness of approximately 12 µm has been deposited on the platinum buffered Si substrate via a sol-gel composite method. The separation of the film from the substrate was achieved using a wet chemical method. The lifted-off PMN-35PT thick film exhibited good dielectric and ferroelectric properties. At 1 kHz, the dielectric constant and the dielectric loss were 3,326 and 0.037, respectively, while the remnant polarization was 30.0 µC/cm2. A high frequency single element acoustic transducer fabricated with this film showed a bandwidth at ?6 dB of 63.6% at 110 MHz. PMID:21170158
Design limits of compressible NLF airfoils
NASA Technical Reports Server (NTRS)
Viken, Jeff; Wagner, R. D.
1991-01-01
There has been considerable success in the design of practical low-speed natural laminar flow (NLF) airfoils, with significant profile drag reduction from conventional turbulent flow airfoils. The favorable lowspeed results give an incentive to explore the possibilities of high-subsonic speed NLF airfoil design. The design problem at higher freestream Mach numbers is more severe than for low-speed designs because these high-speed airplanes typically fly at higher chord Reynolds numbers than low-speed NLF airplanes. As the Mach number increases, the main priority changes from the use of sufficient acceleration to achieve NLF for low-drag, to the delay of separation in the far aft pressure recovery region. In this effort, NLF airfoils have been designed for Mach numbers ranging from 0.60 to 0.80 and chord Reynolds numbers of 30 x 10 to the 6th and 40 x 10 to the 6th.
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.
Hu, Hui
to reattach the leading edge separation bubble to the surface at higher angles of attack than a stationary/c) = angle of attack relative to free stream b = span length c = chord length cl = lift coefficient cp the leading edge of an airfoil at higher angles of attack. Many studies have been performend on the airfoil
Optical-velocimetry, wake measurements of lift and induced drag on a wing
I. Grant; G. McCutcheon; A. H. McColgan; D. Hurst
2006-01-01
Particle Tracking Velocimetry was used, in a low-speed wind-tunnel study, to obtain simultaneous cross-flow velocity measurements in three planar regions downstream of an airfoil having a NACA 0015 profile. In order to measure both the total lift and induced drag, and their distributions on the airfoil, a wake integration technique based on the control volume approach was used. The airfoil
Low-drag airfoils in transonic flow of dense gases
NASA Astrophysics Data System (ADS)
Rusak, Zvi; Wang, C.-W.
1998-11-01
Low-drag airfoils for transonic flow of BZT gases are constructed through a recent nonlinear small-disturbance theory on this topic. This kind of transonic flow is characterized by the high nonlinearity of the fluid thermodynamic behavior that is closely coupled with its compressible flow dynamics. Utilizing BZT gases may provide a low drag environment at high transonic speeds and this advantage may be further improved by a proper design of the airfoil shape. The modified airfoils are characterized by arcs along which the flow is sonic and that connect to a sharp tail. These airfoils may give the highest free stream Mach number for which the flow is nowhere supersonic. Using an analysis in the hodograph (velocity) plane, an analytical formula is derived to describe the sonic arc. Numerical computations using the Euler equations demonstrate that the modified airfoils have significantly lower pressure drag against standard airfoils and their critical Mach number is increased to higher values.
NASA Technical Reports Server (NTRS)
Glaab, Louis J.; Riley, Donald R.; Brandon, Jay M.; Person, Lee H., Jr.; Glaab, Patricia C.
1999-01-01
As part of an effort between NASA and private industry to reduce airport-community noise for high-speed civil transport (HSCT) concepts, a piloted simulation study was initiated for the purpose of predicting the noise reduction benefits that could result from improved low-speed high-lift aerodynamic performance for a typical HSCT configuration during takeoff and initial climb. Flight profile and engine information from the piloted simulation were coupled with the NASA Langley Aircraft Noise Prediction Program (ANOPP) to estimate jet engine noise and to propagate the resulting source noise to ground observer stations. A baseline aircraft configuration, which also incorporated different levels of projected improvements in low-speed high-lift aerodynamic performance, was simulated to investigate effects of increased lift and lift-to-drag ratio on takeoff noise levels. Simulated takeoff flights were performed with the pilots following a specified procedure in which either a single thrust cutback was performed at selected altitudes ranging from 400 to 2000 ft, or a multiple-cutback procedure was performed where thrust was reduced by a two-step process. Results show that improved low-speed high-lift aerodynamic performance provides at least a 4 to 6 dB reduction in effective perceived noise level at the FAA downrange flyover measurement station for either cutback procedure. However, improved low-speed high-lift aerodynamic performance reduced maximum sideline noise levels only when using the multiple-cutback procedures.
Reynolds Number Effects on a Supersonic Transport at Subsonic High-Lift Conditions (Invited)
NASA Technical Reports Server (NTRS)
Owens, L.R.; Wahls, R. A.
2001-01-01
A High Speed Civil Transport configuration was tested in the National Transonic Facility at the NASA Langley Research Center as part of NASA's High Speed Research Program. The primary purposes of the tests were to assess Reynolds number scale effects and high Reynolds number aerodynamic characteristics of a realistic, second generation supersonic transport while providing data for the assessment of computational methods. The tests included longitudinal and lateral/directional studies at transonic and low-speed, high-lift conditions across a range of Reynolds numbers from that available in conventional wind tunnels to near flight conditions. Results are presented which focus on Reynolds number and static aeroelastic sensitivities of longitudinal characteristics at Mach 0.30 for a configuration without an empennage. A fundamental change in flow-state occurred between Reynolds numbers of 30 to 40 million, which is characterized by significantly earlier inboard leading-edge separation at the high Reynolds numbers. Force and moment levels change but Reynolds number trends are consistent between the two states.
NASA Technical Reports Server (NTRS)
Johnson, W. G., Jr.
1976-01-01
An investigation was made in the 5.18 m (17 ft) test section of the Langley 300 MPH 7 by 10 foot tunnel on a rectangular, aspect ratio 6 wing which had a slotted supercritical airfoil section and externally blown flaps. The 13 percent thick wing was fitted with two high lift flap systems: single slotted and double slotted. The designations single slotted and double slotted do not include the slot which exists near the trailing edge of the basic slotted supercritical airfoil. Tests were made over an angle of attack range of -6 deg to 20 deg and a thrust-coefficient range up to 1.94 for a free-stream dynamic pressure of 526.7 Pa (11.0 lb/sq ft). The results of the investigation are presented as curves and tabulations of the chordwise pressure distributions at the midsemispan station for the wing and each flap element.
Airfoil Design and Rotorcraft Performance
WILLIAM G. BOUSMAN
2002-01-01
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
Wind tunnel evaluation of a truncated NACA 64-621 airfoil for wind turbine applications
NASA Technical Reports Server (NTRS)
Law, S. P.; Gregorek, G. M.
1987-01-01
An experimental program to measure the aerodynamic performance of a NACA 64-621 airfoil with a truncated trailing edge for wind turbine applications has been conducted in the Ohio State University Aeronautical and Astronautical Research Laboratory 6 in. by 21 in. pressurized wind tunnel. The blunted or trailing edge truncated (TET) airfoil has an advantage over similar trailing edge airfoils because it is able to streamline a larger spar structure, while also providing aerodynamic properties that are quite good. Surface pressures were measured and integrated to determine the lift, pressure drag, and moment coefficients over angles of attack ranging from -14 to +90 deg at Mach 0.2 and Reynolds numbers of 1,000,000 and 600,000. Results are compared to the NACA 0025, 0030, and 0035 thick airfoils with sharp trailing edges. Comparison shows that the 30 percent thick NACA 64-621-TET airfoil has higher maximum lift, higher lift curve slope, lower drag at higher lift coefficients, and higher chordwise force coefficient than similar thick airfoils with sharp trailing edges.
NASA Astrophysics Data System (ADS)
Fink, Martin R.
1991-08-01
Propulsive life noise is the increase in noise that occurs when airframe surfaces are placed in the propulsive system's exhaust to increase their lift force. Increased local flow velocities and turbulence levels, due to the propulsive system exhaust gases passing along the airframe lifting surfaces, cause an increase in maximum lift coefficient. The airplane's flight speed required for takeoff and landing can then be significantly reduced, allowing operation from shorter runways than those of conventional commercial airports. Unfortunately, interaction of high velocity turbulent exhaust flow with the airframe's solid surfaces generates additional noise radiation. Aeroacoustic processes that cause propulsive lift noise also are present in airframe noise and propulsive system installation noise. Research studies of propulsive lift noise led to development of improved methods of predicting noise radiation from surfaces in turbulent flows. Noise reduction and prediction methods of aircraft noise are discussed.
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.
Force control of heavy lift manipulators for high precision insertion tasks
DiCicco, Matthew A. (Matthew Adam)
2005-01-01
The inherent strength of robotic manipulators can be used to assist humans in performing heavy lifting tasks. These robots reduce manpower, reduce fatigue, and increase productivity. This thesis deals with the development ...
Low-drag airfoils for transonic flow of dense gases
Zvi Rusak; Chun-Wei Wang
2000-01-01
Low-drag airfoils for transonic flow of BZT gases are constructed through a nonlinear small-disturbance theory on this topic. This kind of transonic flow is characterized by the high nonlinearity of the fluid thermodynamic behavior that is closely coupled with its compressible flow dynamics. Utilizing BZT gases may result in low drag exerted on airfoils operating at high transonic speeds. This
Computing Aerodynamic Performance of a 2D Iced Airfoil: Blocking Topology and Grid Generation
NASA Technical Reports Server (NTRS)
Chi, X.; Zhu, B.; Shih, T. I.-P.; Slater, J. W.; Addy, H. E.; Choo, Yung K.; Lee, Chi-Ming (Technical Monitor)
2002-01-01
The ice accrued on airfoils can have enormously complicated shapes with multiple protruded horns and feathers. In this paper, several blocking topologies are proposed and evaluated on their ability to produce high-quality structured multi-block grid systems. A transition layer grid is introduced to ensure that jaggedness on the ice-surface geometry do not to propagate into the domain. This is important for grid-generation methods based on hyperbolic PDEs (Partial Differential Equations) and algebraic transfinite interpolation. A 'thick' wrap-around grid is introduced to ensure that grid lines clustered next to solid walls do not propagate as streaks of tightly packed grid lines into the interior of the domain along block boundaries. For ice shapes that are not too complicated, a method is presented for generating high-quality single-block grids. To demonstrate the usefulness of the methods developed, grids and CFD solutions were generated for two iced airfoils: the NLF0414 airfoil with and without the 623-ice shape and the B575/767 airfoil with and without the 145m-ice shape. To validate the computations, the computed lift coefficients as a function of angle of attack were compared with available experimental data. The ice shapes and the blocking topologies were prepared by NASA Glenn's SmaggIce software. The grid systems were generated by using a four-boundary method based on Hermite interpolation with controls on clustering, orthogonality next to walls, and C continuity across block boundaries. The flow was modeled by the ensemble-averaged compressible Navier-Stokes equations, closed by the shear-stress transport turbulence model in which the integration is to the wall. All solutions were generated by using the NPARC WIND code.
High-frequency microwave anti-\\/de-icing system for carbon-reinforced airfoil structures
Lambert Feher; Manfred Thumm
2001-01-01
An aircraft may be subjected to icing for a variety of meteorological reasons during the flight. Ice formation on the plane and in particular on the aerodynamically carrying structures adversely affects the flight behaviour. Conventional de-icing methods for aluminum wings are characterised by a high energy consumption during the flight and slow ice melting due to thermal diffusion of the
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.
Time resolved PIV analysis of flow over a NACA 0015 airfoil with Gurney flap
D. R. Troolin; E. K. Longmire; W. T. Lai
2006-01-01
A NACA 0015 airfoil with and without a Gurney flap was studied in a wind tunnel with Re\\u000a c = 2.0 × 105 in order to examine the evolving flow structure of the wake through time-resolved PIV and to correlate this structure with time-averaged measurements of the lift coefficient. The Gurney flap, a tab of small length (1–4% of the airfoil chord) that protrudes
Self-sustained shock oscillations on airfoils at transonic speeds
B. H. K. Lee
2001-01-01
Self-sustained shock wave oscillations on airfoils at transonic flow conditions are associated with the phenomenon of buffeting. The physical mechanisms of the periodic shock motion are not yet fully understood even though experiments performed over fifty years ago have demonstrated the presence of oscillatory shock waves on the airfoil surfaces at high subsonic speeds. The unsteady pressure fluctuations generated by
Investigations of the transonic flow around oscillating airfoils
H. Tijdeman
1977-01-01
Exploratory wind-tunnel experiments in high-subsonic and transonic flow on a conventional airfoil with oscillating flap and a supercritical airfoil oscillating in pitch are described. In the analysis of the experimental results, emphasis is placed upon the typical aspects of transonic flow, namely the interaction between the steady and unsteady flow fields, the periodical motion of the shock waves and their
Aerodynamic Characterization of a Thin, High-Performance Airfoil for Use in Ground Fluids Testing
NASA Technical Reports Server (NTRS)
Broeren, Andy P.; Lee, Sam; Clark, Catherine
2013-01-01
The FAA has worked with Transport Canada and others to develop allowance times for aircraft operating in ice-pellet precipitation. Wind-tunnel testing has been carried out to better understand the flowoff characteristics and resulting aerodynamic effects of anti-icing fluids contaminated with ice pellets using a thin, high-performance wing section at the National Research Council of Canada Propulsion and Icing Wind Tunnel. The objective of this paper is to characterize the aerodynamic behavior of this wing section in order to better understand the adverse aerodynamic effects of anti-icing fluids and ice-pellet contamination. Aerodynamic performance data, boundary-layer surveys and flow visualization were conducted at a Reynolds number of approximately 6.0 x 10(exp 6) and a Mach number of 0.12. The clean, baseline model exhibited leading-edge stall characteristics including a leading-edge laminar separation bubble and minimal or no separation on the trailing edge of the main element or flap. These results were consistent with expected 2-D aerodynamics and showed no anomalies that could adversely affect the evaluation of anti-icing fluids and ice-pellet contamination on the wing. Tests conducted with roughness and leading-edge flow disturbances helped to explain the aerodynamic impact of the anti-icing fluids and contamination. The stalling characteristics of the wing section with fluid and contamination appear to be driven at least partially by the effects of a secondary wave of fluid that forms near the leading edge as the wing is rotated in the simulated takeoff profile. These results have provided a much more complete understanding of the adverse aerodynamic effects of anti-icing fluids and ice-pellet contamination on this wing section. This is important since these results are used, in part, to develop the ice-pellet allowance times that are applicable to many different airplanes.
Aerodynamic Characterization of a Thin, High-Performance Airfoil for Use in Ground Fluids Testing
NASA Technical Reports Server (NTRS)
Broeren, Andy P.; Lee, Sam; Clark, Catherine
2013-01-01
The FAA has worked with Transport Canada and others to develop allowance times for aircraft operating in ice-pellet precipitation. Wind-tunnel testing has been carried out to better understand the flowoff characteristics and resulting aerodynamic effects of anti-icing fluids contaminated with ice pellets using a thin, high-performance wing section at the National Research Council of Canada Propulsion and Icing Wind Tunnel. The objective of this paper is to characterize the aerodynamic behavior of this wing section in order to better understand the adverse aerodynamic effects of anti-icing fluids and ice-pellet contamination. Aerodynamic performance data, boundary-layer surveys and flow visualization were conducted at a Reynolds number of approximately 6.0×10(exp 6) and a Mach number of 0.12. The clean, baseline model exhibited leading-edge stall characteristics including a leading-edge laminar separation bubble and minimal or no separation on the trailing edge of the main element or flap. These results were consistent with expected 2-D aerodynamics and showed no anomalies that could adversely affect the evaluation of anti-icing fluids and ice-pellet contamination on the wing. Tests conducted with roughness and leading-edge flow disturbances helped to explain the aerodynamic impact of the anti-icing fluids and contamination. The stalling characteristics of the wing section with fluid and contamination appear to be driven at least partially by the effects of a secondary wave of fluid that forms near the leading edge as the wing is rotated in the simulated takeoff profile. These results have provided a much more complete understanding of the adverse aerodynamic effects of anti-icing fluids and ice-pellet contamination on this wing section. This is important since these results are used, in part, to develop the ice-pellet allowance times that are applicable to many different airplanes.
Davide Lengani; Daniele Simoni; Marina Ubaldi; Pietro Zunino; Francesco Bertini
The transition and separation processes of the boundary layer developing on a flat plate under a prescribed adverse pressure\\u000a gradient typical of Ultra-High-Lift low-pressure turbine profiles have been investigated, with and without the application\\u000a of a synthetic jet (zero net mass flow rate jet). A mechanical piston has been adopted to produce an intermittent flow with\\u000a zero net mass flow
Airfoil treatments for vertical axis wind turbines
Klimas, P.C.
1985-01-01
Sandia National Laboratories (SNL) has taken three airfoil related approaches to decreasing the cost of energy of vertical axis wind turbine (VAWT) systems; airfoil sections designed specifically for VAWTs, vortex generators (VGs), and ''pumped spoiling.'' SNL's blade element airfoil section design effort has led to three promising natural laminar flow (NLF) sections. One section is presently being run on the SNL 17-m turbine. Increases in peak efficiency and more desirable dynamic stall regulation characteristics have been observed. Vane-type VGs were fitted on one DOE/Alcoa 100 kW VAWT. With approximately 12% of span having VGs, annual energy production increased by 5%. Pumped spoiling utilizes the centrifugal pumping capabilities of hollow blades. With the addition of small perforations in the surface of the blades and valves controlled by windspeed at the ends of each blade, lift spoiling jets may be generated inducing premature stall and permitting lower capacity, lower cost drivetrain components. SNL has demonstrated this concept on its 5-m turbine and has wind tunnel tested perforation geometries on one NLF section.
NASA Technical Reports Server (NTRS)
Vandam, C. P.
1993-01-01
Boundary-layer transition and relaminarization may have a critical effect on the flow development about multi-element high-lift systems of subsonic transport jets with swept wings. The purpose of the research is to study these transition phenomena in the leading-edge region of the various elements of a high-lift system. The flow phenomena studied include transition to the attachment-line flow, relaminarization, and crossflow instability, and transition. The calculations are based on pressure distributions measured in flight on the NASA Transport Systems Research Vehicle (Boeing 737-100) at a wing station where the flow approximated infinite swept wing conditions. The results indicate that significant regions of laminar flow can exist on all flap elements in flight. In future flight experiments (planned for January-February, 1994) the extent of these regions, the transition mechanisms and the effect of laminar flow on the high-lift characteristics of the multi-element system will be further explored.
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.
Unsteady measurements of a three-dimensional high-lift system
NASA Astrophysics Data System (ADS)
Moriarty, Patrick John
An experimental study of the unsteady fluid dynamics around a three-element high-lift system was performed. Four different measurement techniques were used to quantify the flowfield: hot-wire anemometry, unsteady pressure measurements, acoustic measurements, and particle image velocimetry. From these measurements, it was discovered that the unsteady flowfield around a leadingedge slat is extremely complex. In addition to unsteadiness from separating shear layers, the fluid dynamics are affected by a coupling between the acoustic and hydrodynamic disturbances. The slat trailing edge and slat gap play a primary role in contributing to unsteadiness in the slat cove. An acoustic resonance in the slat gap interacts with the vortex shedding of the slat trailing edge, and also stimulates shedding from the cusped edge on the lower surface of the slat. This acoustic excitation amplifies certain frequencies of vortex shedding and also induces an effect known as "lockin". The magnitude of the shedding is largely a function of the state of the boundary layer at the slat trailing edge. Configurations with laminar boundary layers have larger amplitude disturbances in velocity and pressure. The magnitude of the slat disturbances may also be affected by the relative thickness of the slat trailing edge. A slight nonlinear behavior of the vortex shedding frequency may be attributed to an acoustic feedback mechanism. The total sound pressure level of the slat noise scales with velocity to the fifth power. Serrating the trailing edge of the slat significantly reduces the radiated noise, as well as fluctuations of velocity and pressure in the slat cove. The discovered noise mechanisms may scale to full-size aircraft. The flowfield around the flap tip is dominated by a primary vortex rolling out from the lower surface of the flap. A secondary vortex also exists which originates from the upper surface, although, this secondary vortex is quickly assimilated with the primary vortex and is only of importance in the near field.
NASA Technical Reports Server (NTRS)
Yamauchi, G. K.; Johnson, W.
1983-01-01
The primary effects of Reynolds number on two dimensional airfoil characteristics are discussed. Results from an extensive literature search reveal the manner in which the minimum drag and maximum lift are affected by the Reynolds number. C sub d sub min and C sub l sub max are plotted versus Reynolds number for airfoils of various thickness and camber. From the trends observed in the airfoil data, universal scaling laws and easily implemented methods are developed to account for Reynolds number effects in helicopter rotor analyses.
Wei Wen Su; Bing Jun He; Hua Liang; Sam Sun
1996-01-01
A new bioreactor design that allows continuous perfusion cultivation of plant cell suspensions is described in this paper. This design incorporates an internal cell settling zone with an external-loop air-lift bioreactor. The settling zone is created by inserting a baffle plate into the upper portion of the downcomer. Using this bioreactor, Anchusa officinalis suspension culture was cultivated to a cell
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.
NASA Technical Reports Server (NTRS)
Moracz, Donald J. (Inventor); Cook, Charles R. (Inventor); Toth, Istvan J. (Inventor)
1986-01-01
An improved method of making an airfoil includes stacking plies in two groups. A separator ply is positioned between the two groups of plies. The groups of plies and the separator ply are interconnected to form an airfoil blank. The airfoil blank is shaped, by forging or other methods, to have a desired configuration. The material of the separator ply is then dissolved or otherwise removed from between the two sections of the airfoil blank to provide access to the interior of the airfoil blank. Material is removed from inner sides of the two separated sections to form core receiving cavities. After cores have been placed in the cavities, the two sections of the airfoil blank are interconnected and the shaping of the airfoil is completed. The cores are subsequently removed from the completed airfoil.
NASA Technical Reports Server (NTRS)
Moracz, Donald J. (Inventor); Cook, Charles R. (Inventor); Toth, Istvan J. (Inventor)
1984-01-01
An improved method of making an airfoil includes stacking plies in two groups. A separator ply is positioned between the two groups of plies. The groups of plies and the separator ply are interconnected to form an airfoil blank. The airfoil blank is shaped, by forging or other methods, to have a desired configuration. The material of the separator ply is then dissolved or otherwise removed from between the two sections of the airfoil blank to provide access to the interior of the airfoil blank. Material is removed from inner sides of the two separated sections to form core receiving cavities. After cores have been placed in the cavities, the two sections of the airfoil blank are interconnected and the shaping of the airfoil is completed. The cores are subsequently removed from the completed airfoil.
NASA Technical Reports Server (NTRS)
Zaguli, R. J.; Bragg, M. B.; Gregorek, G. M.
1984-01-01
Aerodynamic data from a test program in the Icing Research Tunnel are reported for a NACA 63A415 airfoil, with fowler flap, clean and with simulated ice shapes. The effect of three ice shapes on airfoil performance are presented, two of the simulated ice shapes are from earlier Icing Tunnel tests. Lift, drag, and moment coefficients are reported for the airfoil, clean and with ice, for angles of attack from approximately zero lift to maximum lift and for flap deflections of 0, 10, 20, and 30 degrees. Surface pressure distribution plots for the airfoil and flap are presented for all runs. Some preliminary oil flow visualization data are also discussed. Large drag penalties were measured in all instances. Maximum lift penalties were in general serious, and depend upon the ice shape and flap deflection.
NASA Technical Reports Server (NTRS)
Gregorek, G. M.; Kuniega, R. J.; Nyland, T. W.
1988-01-01
The aerodynamic similarity between a small (4-inch chord) wind tunnel model and a full-scale wind turbine blade (24-foot tip section with a 36-inch chord) was evaluated by comparing selected pressure distributions around the geometrically similar cross sections. The airfoils were NACA 64-621 sections, including trailing-edge ailerons with a width equal to 38 percent of the airfoil chord. The model airfoil was tested in the OSU 6- by 12-inch High Reynolds Number Wind Tunnel; the full-scale blade section was tested in the NASA Langley Research Center 30- by 60-foot Subsonic Wind Tunnel. The model airfoil contained 61 pressure taps connected by embedded tubes to pressure transducers. A belt containing 29 pressure taps was fixed to the full-scale section at midspan to obtain surface pressure data. Lift coefficients were obtained by integrating pressures, and corrections were made for the 3-D effects of blade twist and downwash in the blade tip section. The results of the two different experimental methods correlated well for angles of attack from minus 4 to 36 degrees and aileron reflections from 0 to 90 degrees.
Gregorek, G.M.; Kuniega, R.J.; Nyland, T.W.
1988-04-01
The aerodynamic similarity between a small (4-in. chord) wind tunnel model and a full-scale wind turbine blade (24-ft tip section with a 36-in. chord) was evaluated by comparing selected pressure distributions around the geometrically similar cross sections. The airfoils were NACA 64-621 sections, including trailing-edge ailerons with a width equal to 38 percent of the airfoil chord. The model airfoil was tested in the OSU 6- by 12-In. High Reynolds Number Wind Tunnel; the full-scale blade section was tested in the NASA Langley Research Center 30- by 60-Ft Subsonic Wind Tunnel. The model airfoil contained 61 pressure taps connected by embedded tubes to pressure transducers. A belt containing 29 pressure taps was fixed to the full-scale section at midspan to obtain surface pressure data. Lift coefficients were obtained by integrating pressures, and corrections were made for the three-dimensional effects of blade twist and downwash in the blade tip section. Good correlation was obtained between the results of the two different experimental methods for angles of attack from -4/degree/ to 36/degree/ and aileron deflections from 0/degree/ to 90/degree/. 4 refs., 11 figs., 1 tab.
Wind-tunnel Investigation of Two Airfoils with 25-percent-chord Gwinn and Plain Flaps
NASA Technical Reports Server (NTRS)
Ames, Milton B , Jr
1940-01-01
Aerodynamic force tests of an NACA 23018 airfoil with a Gwinn flap having a chord 25 percent of the overall chord and of an NACA 23015 airfoil with a plain flap having a 25-percent chord were conducted to determine the relative merits of the Gwinn and the plain flaps. The tests indicated that, based on speed-range ratios, the plain flap was more effective than the Gwinn flap. At small flap deflections, the plain flap had lower drag coefficients at lift-coefficient values less than 0.70. For lift coefficients greater than 0.70, however, the Gwinn flap at all downward flap deflections had the lower drag coefficients.
NASA Technical Reports Server (NTRS)
Hess, Robert W.; Seidel, David A.; Igoe, William B.; Lawing, Pierce L.
1987-01-01
Steady and unsteady pressures were measured on a 2-D supercritical airfoil in the Langley Research Center 0.3-m Transonic Cryogenic Tunnel at Reynolds numbers from 6 x 1,000,000 to 35 x 1,000,000. The airfoil was oscillated in pitch at amplitudes from plus or minus .25 degrees to plus or minus 1.0 degrees at frequencies from 5 Hz to 60 Hz. The special requirements of testing an unsteady pressure model in a pressurized cryogenic tunnel are discussed. Selected steady measured data are presented and are compared with GRUMFOIL calculations at Reynolds number of 6 x 1,000,000 and 30 x 1,000,000. Experimental unsteady results at Reynolds numbers of 6 x 1,000,000 and 30 x 1,000,000 are examined for Reynolds number effects. Measured unsteady results at two mean angles of attack at a Reynolds number of 30 x 1,000,000 are also examined.
Experimental Optimization Methods for Multi-Element Airfoils
NASA Technical Reports Server (NTRS)
Landman, Drew; Britcher, Colin P.
1996-01-01
A modern three element airfoil model with a remotely activated flap was used to investigate optimum flap testing position using an automated optimization algorithm in wind tunnel tests. Detailed results for lift coefficient versus flap vertical and horizontal position are presented for two angles of attack: 8 and 14 degrees. An on-line first order optimizer is demonstrated which automatically seeks the optimum lift as a function of flap position. Future work with off-line optimization techniques is introduced and aerodynamic hysteresis effects due to flap movement with flow on are discussed.
Airfoil Pressure Distribution Investigation in the Variable Density Wind Tunnel
NASA Technical Reports Server (NTRS)
Jacobs, Eastman N; Stack, John; Pinkerton, Robert M
1931-01-01
Report presents the results of wind tunnel tests of pressure distribution measurements over one section each of six airfoils. Pressure distribution diagrams, as well as the integrated characteristics of the airfoils, are given for both a high and a low dynamic scale or, Reynolds number VL/V, for comparison with flight and other wind-tunnel tests, respectively. It is concluded that the scale effect is very important only at angles of attack near the burble. The distribution of pressure over an airfoil having a Joukowski section is compared with the theoretically derived distribution. A further study of the distribution of pressure over all of the airfoils resulted in the development of an approximate method of predicting the pressure distribution along the chord of any normal airfoil for all attitudes within the working range if the distribution at one attitude is known.
Design and Predictions for a High-Altitude (Low-Reynolds-Number) Aerodynamic Flight Experiment
NASA Technical Reports Server (NTRS)
Greer, Donald; Hamory, Phil; Krake, Keith; Drela, Mark
1999-01-01
A sailplane being developed at NASA Dryden Flight Research Center will support a high-altitude flight experiment. The experiment will measure the performance parameters of an airfoil at high altitudes (70,000 to 100,000 ft), low Reynolds numbers (200,000 to 700,000), and high subsonic Mach numbers (0.5 and 0.65). The airfoil section lift and drag are determined from pitot and static pressure measurements. The locations of the separation bubble, Tollmien-Schlichting boundary layer instability frequencies, and vortex shedding are measured from a hot-film strip. The details of the planned flight experiment are presented. Several predictions of the airfoil performance are also presented. Mark Drela from the Massachusetts Institute of Technology designed the APEX-16 airfoil, using the MSES code. Two-dimensional Navier-Stokes analyses were performed by Mahidhar Tatineni and Xiaolin Zhong from the University of California, Los Angeles, and by the authors at NASA Dryden.
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.
Multiple piece turbine airfoil
Kimmel, Keith D (Jupiter, FL)
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.
The Monoplane as a Lifting Vortex Surface
NASA Technical Reports Server (NTRS)
Blenk, Hermann
1947-01-01
In Prandtl's airfoil theory the monoplane was replaced by a single lifting vortex line and yielded fairly practical results. However, the theory remained restricted to the straight wing. Yawed wings and those curved in flight direction could not be computed with this first approximation; for these the chordwise lift distribution must be taken into consideration. For the two-dimensional problem the transition from the lifting line to the lifting surface has been explained by Birnbaum. In the present report the transition to the three-dimensional problem is undertaken. The first fundamental problem involves the prediction of flow, profile, and drag for prescribed circulation distribution on the straight rectangular wing, the yawed wing for lateral boundaries parallel to the direction of flight, the swept-back wing, and the rectangular wing in slipping, with the necessary series developments for carrying through the calculations, the practical range of convergence of which does not comprise the wing tips or the break point of the swept-back wing. The second problem concerns the calculation of the circulation distribution with given profile for a slipping rectangular monoplane with flat profile and aspect ratio 6, and a rectangular wing with cambered profile and variable aspect ratio-the latter serving as check of the so-called conversion formulas of the airfoil theory.
Unsteady modes in the flowfield about an airfoil with a leading-edge horn-ice shape
NASA Astrophysics Data System (ADS)
Ansell, Phillip J.
An analysis of unsteady modes present in the flowfield of an airfoil with a leading-edge horn-ice shape was performed in the current study. An NACA 0012 airfoil was tested in a subsonic wind tunnel at Re = 1.8 x 106. In addition to the clean configuration, the airfoil model was also tested with a set of boundary-layer trips, a two-dimensional extrusion of a horn-ice shape casting, and an array of simulated icing configurations created using simple geometries. Time-averaged and unsteady static pressure measurements were acquired about the airfoil surface, along with unsteady wake velocity and surface hot-film array measurements. Additionally, surface and off-body flow visualization techniques were used to visualize the airfoil flowfield. A technique was also developed to determine the unsteady shear-layer reattachment location of the ice-induced laminar separation bubble downstream of the horn-ice shape using the surface hot-film array measurements. The maximum amount of unsteadiness in the iced-airfoil flowfield was observed to increase with increasing angle of attack. For a fixed angle of attack prior to stall, a change in the feature height of the simulated ice shape led to a change in the distribution of flowfield unsteadiness, but did not change the maximum levels of unsteadiness present in the flowfield. The iced-airfoil flowfield unsteadiness was primarily associated with three different frequencies. The first was represented by an increase in spectral energy across a broad-band frequency range, and was observed just upstream of shear-layer reattachment as well as downstream of shear-layer reattachment. This increase in spectral energy was caused by the regular mode of unsteadiness due to vortical motion in the separated shear layer and vortex shedding from the separation bubble. The average Strouhal number of this regular mode corresponded to StL = 0.60, and the average vortex convection velocity was observed to be 0.45Uinfinity. These values were highly consistent with those reported elsewhere in the literature. The other two frequencies were much lower and were observed as narrow-band peaks in the spectral content of the acquired measurements that were primarily present in the region covered by the ice-induced separation bubble. The first was attributed to the shear-layer flapping phenomenon and was particularly dominant in the upstream portion of the separation bubble. The Strouhal number associated with this shear-layer flapping mode corresponded to St h = 0.0185, which was consistent with those reported in studies of separation bubbles about canonical geometries. The second frequency was lower than that of shear-layer flapping and was associated with a low-frequency mode of unsteadiness that can occur prior to static stall for airfoils of thin-airfoil stall type. This low-frequency mode was characterized by a low-frequency oscillation of the airfoil circulation, and it was clearly identified in the spectral content of the iced-airfoil lift coefficient. The resulting values of Strouhal number exhibited a dependence on the airfoil angle of attack and corresponded to a range that was consistent with the Strouhal number values reported in prior studies of the low-frequency mode in the literature. Using the method for determining the unsteady shear-layer reattachment location, the average time-dependent relationship between the reattachment location and the lift coefficient was calculated. It was discovered that at the low-frequency mode, the lift coefficient leads the shear-layer reattachment location by a phase of pi/2. This phase relationship occurred due to a feedback between the airfoil circulation and the separation bubble length. This improved understanding of the low-frequency mode in the iced-airfoil flowfield was utilized in a practical example to improve the predictive qualities of a hinge-moment-based stall prediction system. This improvement in the predictive qualities was performed by identifying the intermittent signature of the low-frequency mode in the wavelet transform of the hinge moment coeffic
Euler study on porous transonic airfoils with a view toward multipoint design
NASA Technical Reports Server (NTRS)
Hartwich, Peter M.
1991-01-01
Euler solutions for steady transonic flow (Free-stream Mach 0.63-0.8, alpha = 0-2 deg) over NACA 0012 and supercritical airfoils with solid as well as porous surfaces suggest porosity as a means to realize multipoint design for transonic airfoils. The porous surfaces extend over at least 90 percent of the chord. The porosity distribution is described by a modified sine wave with several amplitudes. Either connected or separated cavities are assumed to lie underneath the upper and lower surfaces. Applied to an NACA 0012 airfoil, porosity generally increases lift, in some instances by up to 65 percent. Porous NACA 0012 airfoils in supercritical flow yield reductions of an order of magnitude in wave drag at constant lift, compared to their solid counterpart. Making the surface of a supercritical airfoil permeable also leads to sizeable reductions in wave drag at constant lift for overspeed conditions. The discussion of the computed results addresses issues such as grid sensitivity and checks for systematic errors.
Supersonic, nonlinear, attached-flow wing design for high lift with experimental validation
NASA Technical Reports Server (NTRS)
Pittman, J. L.; Miller, D. S.; Mason, W. H.
1984-01-01
Results of the experimental validation are presented for the three dimensional cambered wing which was designed to achieve attached supercritical cross flow for lifting conditions typical of supersonic maneuver. The design point was a lift coefficient of 0.4 at Mach 1.62 and 12 deg angle of attack. Results from the nonlinear full potential method are presented to show the validity of the design process along with results from linear theory codes. Longitudinal force and moment data and static pressure data were obtained in the Langley Unitary Plan Wind Tunnel at Mach numbers of 1.58, 1.62, 1.66, 1.70, and 2.00 over an angle of attack range of 0 to 14 deg at a Reynolds number of 2.0 x 10 to the 6th power per foot. Oil flow photographs of the upper surface were obtained at M = 1.62 for alpha approx. = 8, 10, 12, and 14 deg.
An exploratory investigation of airfoil sections in low Reynolds number subsonic compressible flows
NASA Astrophysics Data System (ADS)
McKee, Michael William
Recent interest in high altitude unmanned aircraft has exposed a lack of experimental data for airfoil sections in low Reynolds number, subsonic compressible flow conditions. Computational methods such as the MSES code developed my Mark Drela of MIT have been used to provide data to aircraft designers, but such codes have not been verified with experimental data. In response to this need, a proof-of-concept wind tunnel facility known as the Low Reynolds Number Transonic Test Facility has been constructed at The Ohio State University Aeronautical and Astronautical Research Laboratory. The flow conditions which may be simulated include Reynolds numbers of 1 × 105 per foot to 1.2 × 106 per foot at Mach numbers of 0.4 to 0.6. Two airfoil sections were tested at these conditions, the NACA 0012 and the PZ(1)-0215, a modern laminar airfoil design. Extensive experimental results are presented for both airfoils, including lift, drag and moment coefficients, as well as surface pressure distributions. Considerable variations in airfoil performance are observed with changes in both Reynolds number and Mach number due to the influence of laminar separation bubbles. Both short and long laminar separation bubbles are observed to increase in length with increasing Mach number, suggesting delayed boundary layer and free shear layer transition at higher Mach numbers. However, measurements of test section freestream turbulence intensity raise questions regarding these conclusions, as both frequency and magnitude of the turbulence are observed to change significantly with Mach number. While incompressible laminar boundary layer stability theory suggests that turbulence frequency content is critical to transition, its effect on the low Reynolds number compressible flow in this study is not clear. Regardless, modifications to the tunnel should be made to reduce freestream turbulence to the levels commonly seen in other low Reynolds number wind tunnel facilities. Comparisons between MSES solutions and the experimental data also suggest further improvements be made to reduce wall interference effects. A comprehensive plan for tunnel renovations is presented, including replacement of the tunnel plenum chamber and choke system, an increase in test section wall porosity, and minor improvements in tunnel instrumentation.
NASA Astrophysics Data System (ADS)
Dietz, G.; Schewe, G.; Mai, H.
2004-01-01
Recent results from flutter experiments of the supercritical airfoil NLR 7301 at flow conditions near the transonic dip are presented. The airfoil was mounted with two degrees of freedom in an adaptive solid-wall wind tunnel, and boundary-layer transition was tripped. Two limit-cycle oscillation (LCO) test cases obtained in an adaptive test-section are proposed for comparison with numerical simulations. The results link the global aerodynamic force behavior to the observed LCOs and the identified transonic dip. The time lag of the lift response to the pitching motion of the airfoil appears to be responsible for the characteristic shape of the transonic dip. The amplitude limitation of the LCOs results from a slightly nonlinear dependency of lift and moment on the amplitude of the airfoil motion. LCOs can be controlled by relatively small forces, but amplitudes strongly depend on the damping of the aeroelastic system.
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.
Wind-Tunnel Investigation of an NACA 23012 Airfoil with Various Arrangements of Slotted Flaps
NASA Technical Reports Server (NTRS)
Wenzinger, Carl J; Harris , Thomas A
1939-01-01
An investigation was made in the 7 by 10-foot wind tunnel and in the variable-density wind tunnel of the NACA 23012 airfoil with various slotted-flap arrangements. The purpose of the investigation in the 7 by 10-foot wind tunnel was to determine the airfoil section aerodynamic characteristics as affected by flap shape, slot shape, and flap location. The flap position for maximum lift; polars for arrangements favorable for take-off and climb; and complete lift, drag, and pitching-moment characteristics for selected optimum arrangements were determined. The best arrangements were tested in the variable-density tunnel at an effective Reynolds number of 8,000,000. In addition, data from both wind tunnels are included for plain, split, external-airfoil, and Fowler flaps for purposes of comparison.
On bimodal flutter behavior of a flexible airfoil
NASA Astrophysics Data System (ADS)
Drazumeric, Radovan; Gjerek, Bojan; Kosel, Franc; Marzocca, Pier
2014-02-01
The dynamic aeroelastic behavior of an elastically supported airfoil is studied in order to investigate the possibilities of increasing critical flutter speed by exploiting its chord-wise flexibility. The flexible airfoil concept is implemented using a rigid airfoil-shaped leading edge, and a flexible thin laminated composite plate conformally attached to its trailing edge. The flutter behavior is studied in terms of the number of laminate plies used in the composite plate for a given aeroelastic system configuration. The flutter behavior is predicted by using an eigenfunction expansion approach which is also used to design a laminated plate in order to attain superior flutter characteristics. Such an airfoil is characterized by two types of flutter responses, the classical airfoil flutter and the plate flutter. Analysis shows that a significant increase in the critical flutter speed can be achieved with high plunge and low pitch stiffness in the region where the aeroelastic system exhibits a bimodal flutter behavior, e.g., where the airfoil flutter and the plate flutter occur simultaneously. The predicted flutter behavior of a flexible airfoil is experimentally verified by conducting a series of systematic aeroelastic system configurations wind tunnel flutter campaigns. The experimental investigations provide, for each type of flutter, a measured flutter response, including the one with indicated bimodal behavior.
Active Flow Control on Low-Aspect Ratio, Low-Reynolds Number Airfoils
NASA Astrophysics Data System (ADS)
Munson, Matthew; Kim, Daegyoum; Dickson, William; Gharib, Morteza
2008-11-01
Insect flight observations show high-lift mechanisms that rely on leading-edge vortex stabilization. These processes are intimately coupled to the flapping motion of the insect wing. In fixed wing applications, suitable for micro-air vehicles, active flow control may be capable of providing similar influence over vortex formation and stabilization. Steady and pulsed mass injection strategies are used to explore the open-loop response of both the evolution of the flow structures and the forces experienced by the wing. Flow structures will be quantitatively visualized using Defocused Digital Particle Image Velocimetry (DDPIV) and forces measured via a six-axis balance. Insect flight typically occurs at Reynolds numbers of 10^2 to 10^4, and aspect ratios near three. For this investigation, Reynolds numbers are approximately 10^3. The airfoil models are NACA 0012 profiles with aspect ratio two.
W. Hage; R. Meyer; M. Schatz
Miniflaps at the trailing edge of airfoils (i.e., Gurney flaps) change the Kutta condition and thus produce higher lift. Unfortunately, however, the drag also increases due to the flow separation downstream of such modified trailing edges. The present work describes investigations aimed at the stabilization of the wake flow in order to achieve drag reduction and a decrease of sound
Aerothermodynamic heating and performance analysis of a high-lift aeromaneuvering AOTV concept
NASA Technical Reports Server (NTRS)
Menees, G. P.; Brown, K. G.; Wilson, J. F.; Davies, C. B.
1985-01-01
The thermal-control requirements for design-optimized aeromaneuvering performance are determined for space-based applications and low-earth orbit sorties involving large, multiple plane-inclination changes. The leading-edge heating analysis is the most advanced developed for hypersonic-rarefied flow over lifting surfaces at incidence. The effects of leading-edge bluntness, low-density viscous phenomena, and finite-rate flow-field chemistry and surface catalysis are accounted for. The predicted aerothermodynamic heating characteristics are correlated with thermal-control and flight-performance capabilities. The mission payload capability for delivery, retrieval, and combined operations is determined for round-trip sorties extending to polar orbits. Recommendations are given for future design refinements. The results help to identify technology issues required to develop prototype operational systems.
The influence of laminar separation and transition on low Reynolds number airfoil hysteresis
NASA Technical Reports Server (NTRS)
Mueller, T. J.
1984-01-01
An experimental study of the Lissaman 7769 and Miley MO6-13-128 airfoils at low chord Reynolds numbers is presented. Although both airfoils perform well near their design Reynolds number of about 600,000, they each produce a different type of hysteresis loop in the lift and drag forces when operated below chord Reynolds numbers of 300,000. The type of hysteresis loop was found to depend upon the relative location of laminar separation and transition. The influence of disturbance environment and experimental procedure on the low Reynolds number airfoil boundary layer behavior is also presented. The use of potential flow solutions to help predict how a given airfoil will behave at low Reynolds numbers is also discussed.
Fluidic Modification of the Aerodynamic Performance of an Airfoil at Low Angles of Attack
NASA Astrophysics Data System (ADS)
Desalvo, Michael; Amitay, Michael; Glezer, Ari
2002-11-01
The aerodynamic characteristics of a conventional non-symmetric airfoil are altered at low angles of attack (when the baseline flow is fully attached) by modifying its Kutta condition using a trapped vortex near the trailing edge. The trapped vortex is induced by a miniature, Gurney-like flap integrated with a synthetic jet actuator having a characteristic scale of 0.016c that is mounted on the pressure side of the airfoil immediately upstream of the trailing edge. The induced changes in the Kutta condition result in global flow modifications that affect the aerodynamic forces and moments. The aerodynamic performance can be continuously varied with the actuator momentum coefficient and the lift to pressure drag ratio is monotonically increased to a level that is higher than that of the baseline airfoil. Various actuation approaches and their effect on the airfoil performance will be discussed.
A Method for the Constrained Design of Natural Laminar Flow Airfoils
NASA Technical Reports Server (NTRS)
Green, Bradford E.; Whitesides, John L.; Campbell, Richard L.; Mineck, Raymond E.
1996-01-01
A fully automated iterative design method has been developed by which an airfoil with a substantial amount of natural laminar flow can be designed, while maintaining other aerodynamic and geometric constraints. Drag reductions have been realized using the design method over a range of Mach numbers, Reynolds numbers and airfoil thicknesses. The thrusts of the method are its ability to calculate a target N-Factor distribution that forces the flow to undergo transition at the desired location; the target-pressure-N-Factor relationship that is used to reduce the N-Factors in order to prolong transition; and its ability to design airfoils to meet lift, pitching moment, thickness and leading-edge radius constraints while also being able to meet the natural laminar flow constraint. The method uses several existing CFD codes and can design a new airfoil in only a few days using a Silicon Graphics IRIS workstation.
Viscous transonic flow over airfoils
NASA Astrophysics Data System (ADS)
Wai, J. C.; Yoshihara, H.
The TSFOIL small disturbance potential code is coupled to Green's lag entrainment method to compute transonic separated flows over airfoils. Green's equation serves as a viscous tangency condition for the TSFOIL code, eliminating an iterative procedure between the viscous and inviscid flows. The method is applied to the case of the NASA Supercritical Airfoil 12 and the 18% Circular Arc Airfoil, the latter with severe separation.
A two element laminar flow airfoil optimized for cruise. M.S. Thesis
NASA Technical Reports Server (NTRS)
Steen, Gregory Glen
1994-01-01
Numerical and experimental results are presented for a new two-element, fixed-geometry natural laminar flow airfoil optimized for cruise Reynolds numbers on the order of three million. The airfoil design consists of a primary element and an independent secondary element with a primary to secondary chord ratio of three to one. The airfoil was designed to improve the cruise lift-to-drag ratio while maintaining an appropriate landing capability when compared to conventional airfoils. The airfoil was numerically developed utilizing the NASA Langley Multi-Component Airfoil Analysis computer code running on a personal computer. Numerical results show a nearly 11.75 percent decrease in overall wing drag with no increase in stall speed at sailplane cruise conditions when compared to a wing based on an efficient single element airfoil. Section surface pressure, wake survey, transition location, and flow visualization results were obtained in the Texas A&M University Low Speed Wind Tunnel. Comparisons between the numerical and experimental data, the effects of the relative position and angle of the two elements, and Reynolds number variations from 8 x 10(exp 5) to 3 x 10(exp 6) for the optimum geometry case are presented.
An Airfoil Spanning an Open Jet
NASA Technical Reports Server (NTRS)
Stuper, J
1933-01-01
Proceeding from the fundamental problem on the mutual relation of a wing and free boundaries the distribution of the circulation is determined for an airfoil spanning an open jet of rectangular section at different aspect ratios, and then for an open jet of circular section. The solution is obtained by means of a Fourier series and computations have been performed for different values of the variables. The second part describes the experiments performed for the purpose of proving the theory. The results confirm the theory. In conclusion it defines the induced drag of a wing extending across an open jet and compares it with the drag of a monoplane having a span equal to the jet width at equal total lift.
NASA Technical Reports Server (NTRS)
Johnson, W. G., Jr.; Hill, A. S.; Eichmann, O.
1985-01-01
A wind tunnel investigation of a NASA 12-percent-thick, advanced-technology supercritical airfoil was conducted in the Langley 0.3-Meter Transonic Cryogenic Tunnel (TCT). This investigation represents another in the series of NASA/U.S. industry two-dimensional airfoil studies to be completed in the Advanced Technology Airfoil Tests program. Test temperature was varied from 220 K to 96 K at pressures ranging from 1.2 to 4.3 atm. Mach number was varied from 0.60 to 0.80. These variables provided a Reynolds number range from 4,400,000 to 40,000,000 based on a 15.24-cm (6.0-in.) airfoil chord. This investigation was designed to test a NASA advanced-technology airfoil from low to flight-equivalent Reynolds numbers, provide experience in cryogenic wind tunnel model design and testing techniques, and demonstrate the suitability of the 0.3-m TCT as an airfoil test facility. The aerodynamic results are presented as integrated force and moment coefficients and pressure distributions. Data are included which demonstrate the effects of fixed transition, Mach number, and Reynolds number on the aerodynamic characteristics. Also included are remarks on the model design, the model structural integrity, and the overall test experience.
Transonic airfoil design for helicopter rotor applications
NASA Technical Reports Server (NTRS)
Hassan, Ahmed A.; Jackson, B.
1989-01-01
Despite the fact that the flow over a rotor blade is strongly influenced by locally three-dimensional and unsteady effects, practical experience has always demonstrated that substantial improvements in the aerodynamic performance can be gained by improving the steady two-dimensional charateristics of the airfoil(s) employed. The two phenomena known to have great impact on the overall rotor performance are: (1) retreating blade stall with the associated large pressure drag, and (2) compressibility effects on the advancing blade leading to shock formation and the associated wave drag and boundary-layer separation losses. It was concluded that: optimization routines are a powerful tool for finding solutions to multiple design point problems; the optimization process must be guided by the judicious choice of geometric and aerodynamic constraints; optimization routines should be appropriately coupled to viscous, not inviscid, transonic flow solvers; hybrid design procedures in conjunction with optimization routines represent the most efficient approach for rotor airfroil design; unsteady effects resulting in the delay of lift and moment stall should be modeled using simple empirical relations; and inflight optimization of aerodynamic loads (e.g., use of variable rate blowing, flaps, etc.) can satisfy any number of requirements at design and off-design conditions.
Wind-tunnel investigation of NACA 23012, 23021, and 23030 airfoils with various sizes of split flap
NASA Technical Reports Server (NTRS)
Wenzinger, Carl J; Harris, Thomas A
1939-01-01
Report presents the results of an investigation made in the NACA 7 by 10-foot wind tunnel of large-chord NACA and 23021, and 23030 airfoils with split flaps 10, 20, 30, and 40 percent of the wing chord to determine the section aerodynamic characteristics of the airfoils as affected by airfoil thickness, flap chord, and flap deflection. The complete section aerodynamic characteristics of all the combinations tested are given in the form of graphs of lift, drag, and pitching-moment coefficients, and certain applications to aerodynamic design are discussed.
Two-fence concept for efficient trapping of vortices on airfoils
NASA Technical Reports Server (NTRS)
Rossow, Vernon J.
1992-01-01
Previous work on the use of a vortex trapped above a wing in order to produce high lift at low angles of attack is extended here. It is first postulated that the optimum way to trap a vortex is to design the airfoil section and wing so that the flow along the vortex core is minimized. It is then shown that a vertical fence both in front of and behind the separation bubble generated by the trapped vortex is an effective way to reduce the mass flow removal and its associated drag to a negligible amount. In order to show that vertical surfaces upstream and downstream of the vortex separation bubble have an opposite effect on the source requirements for vortex trapping, conformal mapping methods are used to obtain the solutions for a variety of simple two-dimensional, inviscid, incompressible flow configurations. Trapped-vortex flowfield solutions for the flow over flat plate and Clark-Y airfoils are then used to demonstrate that the heights of the fences can be tailored to make the required mass withdrawal (and therefore, the drag due to trapping) to be vanishingly small.
Computational Study of a NACA4415 airfoil using synthetic jet control
NASA Astrophysics Data System (ADS)
Lopez Mejia, Omar Dario
Synthetic jet actuators for flow control applications have been an active topic of experimental research since the 90's. Numerical simulations have become an important complement of that experimental work, providing detailed information of the dynamics of the controlled flow. This study is part of the AVOCET (Adaptive VOrticity Control Enabled flighT) project and is intended to provide computational support for the design and evaluation of closed-loop flow control with synthetic jet actuators for small scale Unmanned Aerial Vehicles (UAVs). The main objective is to analyze active flow control of a NACA4415 airfoil with tangential synthetic jets via computational modeling. A hybrid Reynolds-Averaged Navier-Stokes/Large Eddy Simulation (RANS/LES) turbulent model (called Delayed Detached-Eddy Simulation-DDES) was implemented in CDP, a kinetic energy conserving Computational Fluid Dynamics (CFD) code. CDP is a parallel unstructured grid incompressible flow solver, developed at the Center for Integrated Turbulence Simulations (CITS) at Stanford University. Two models of synthetic jet actuators have been developed and validated. The first is a detailed model in which the flow in and out of the actuator cavity is modeled. A second less costly model (RSSJ) was also developed in which the Reynolds stress produced by the actuator is modeled, based on information from the detailed model. Several static validation test cases at different angle of attack with modified NACA4415 and Dragon Eye airfoils were performed. Numerical results show the effects of the actuators on the vortical structure of the flow, as well as on the aerodynamic properties. The main effect of the actuation on the time averaged vorticity field is a bending of the separation shear layer from the actuator toward the airfoil surface, resulting in changes in the aerodynamic properties. Full actuation of the suction side actuator reduces the pitching moment and increases the lift force, while the pressure side actuator increases the pitching moment and reduces the lift force. These observations are in agreement with experimental results. The effectiveness of the actuator is measured by the change in the aerodynamic properties of the airfoil in particular the lift (Delta Cl) and moment (DeltaCm) coefficients. Computational results for the actuator effectiveness show very good agreement with the experimental values (over the range of --2° to 10°). While the actuation modifies the global pressure distribution, the most pronounced effects are near the trailing edge in which a spike in the pressure coefficient (Cp) is observed. The local reduction of Cp, for both the suction side and pressure side actuators, at xc = 0.96 (the position of the actuators) is about 0.9 with respect to the unactuated case. This local reduction of the pressure is associated with the trapped vorticity and flow acceleration close to the trailing edge. The RSSJ model is designed to capture the synthetic jet time averaged behavior so that the high actuation frequencies are eliminated. This allows the time step to be increased by a factor of 5. This ad hoc model is also tested in dynamic simulations, in which its capacity to capture the detail model average performance was demonstrated. Finally, the RSSJ model was extended to a different airfoil profile (Dragon Eye) with good results.
Optical-velocimetry, wake measurements of lift and induced drag on a wing
NASA Astrophysics Data System (ADS)
Grant, I.; McCutcheon, G.; McColgan, A. H.; Hurst, D.
2006-03-01
Particle Tracking Velocimetry was used, in a low-speed wind-tunnel study, to obtain simultaneous cross-flow velocity measurements in three planar regions downstream of an airfoil having a NACA 0015 profile. In order to measure both the total lift and induced drag, and their distributions on the airfoil, a wake integration technique based on the control volume approach was used. The airfoil model was tested in clean wing configuration and with a 3.33% Gurney flap attached to the trailing edge. The Gurney flap was found to increase the complexity of the wake and cause a systematic increase in lift and induced drag values. The changes in total lift and induced drag resulting from the fitting of a Gurney flap compared very favourably to investigations using traditional aerodynamic techniques demonstrating the validity of the method.
Viscous transonic flow over airfoils
J. C. Wai; H. Yoshihara
1981-01-01
The TSFOIL small disturbance potential code is coupled to Green's lag entrainment method to compute transonic separated flows over airfoils. Green's equation serves as a viscous tangency condition for the TSFOIL code, eliminating an iterative procedure between the viscous and inviscid flows. The method is applied to the case of the NASA Supercritical Airfoil 12 and the 18% Circular Arc
CFD aerodynamic analysis of non-conventional airfoil sections for very large rotor blades
NASA Astrophysics Data System (ADS)
Papadakis, G.; Voutsinas, S.; Sieros, G.; Chaviaropoulos, T.
2014-12-01
The aerodynamic performance of flat-back and elliptically shaped airfoils is analyzed on the basis of CFD simulations. Incompressible and low-Mach preconditioned compressible unsteady simulations have been carried out using the k-w SST and the Spalart Allmaras turbulence models. Time averaged lift and drag coefficients are compared to wind tunnel data for the FB 3500-1750 flat back airfoil while amplitudes and frequencies are also recorded. Prior to separation averaged lift is well predicted while drag is overestimated keeping however the trend in the tests. The CFD models considered, predict separation with a 5° delay which is reflected on the load results. Similar results are provided for a modified NACA0035 with a rounded (elliptically shaped) trailing edge. Finally as regards the dynamic characteristics in the load signals, there is fair agreement in terms of Str number but significant differences in terms of lift and drag amplitudes.
Surface integral analogy approaches for predicting noise from 3D high-lift low-noise wings
NASA Astrophysics Data System (ADS)
Yao, Hua-Dong; Davidson, Lars; Eriksson, Lars-Erik; Peng, Shia-Hui; Grundestam, Olof; Eliasson, Peter E.
2014-06-01
Three surface integral approaches of the acoustic analogies are studied to predict the noise from three conceptual configurations of three-dimensional high-lift low-noise wings. The approaches refer to the Kirchhoff method, the Ffowcs Williams and Hawkings (FW-H) method of the permeable integral surface and the Curle method that is known as a special case of the FW-H method. The first two approaches are used to compute the noise generated by the core flow region where the energetic structures exist. The last approach is adopted to predict the noise specially from the pressure perturbation on the wall. A new way to construct the integral surface that encloses the core region is proposed for the first two methods. Considering the local properties of the flow around the complex object-the actual wing with high-lift devices-the integral surface based on the vorticity is constructed to follow the flow structures. The surface location is discussed for the Kirchhoff method and the FW-H method because a common surface is used for them. The noise from the core flow region is studied on the basis of the dependent integral quantities, which are indicated by the Kirchhoff formulation and by the FW-H formulation. The role of each wall component on noise contribution is analyzed using the Curle formulation. Effects of the volume integral terms of Lighthill's stress tensors on the noise prediction are then evaluated by comparing the results of the Curle method with the other two methods.
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-? 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.
Transonic Lifting Flows of Pressurized Gases
NASA Astrophysics Data System (ADS)
Cramer, Mark; Morrison, Michael
1997-11-01
We consider two-dimensional, steady, inviscid transonic flows of pressurized gases over both lifting and nonlifting airfoils. To account for deviations from the ideal gas law we have incorporated a state of the art equation of state known as the Martin-Hou equation. Numerical Solutions are generated through use of a conventional finite volume scheme. Our computations reveal marked qualitative differences with the conventional transonic theory. Result of particular interest include the existence of multiple sonic points, significant decreases in the strength of compression shocks and critical mach numbers well in excess of 0.95 for NACA 0012 cross sections.
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.
Wake structure of a deformable Joukowski airfoil
Adam Ysasi; Eva Kanso; Paul K. Newton
2011-01-01
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
Wind-tunnel Tests of a Hall High-life Wing
NASA Technical Reports Server (NTRS)
Weick, Fred E; Sanders, Robert
1932-01-01
Wind-tunnel tests have been made to find the lift, drag, and center-of-pressure characteristics of a Hall high-lift wing model. The Hall wing is essentially a split-flap airfoil with an internal air passage. Air enters the passage through an opening in the lower surface somewhat back of and parallel to the leading edge, and flows out through an opening made by deflecting the rear portion of the under surface downward as a flap. For ordinary flight conditions the front opening and the rear flap can be closed, providing in effect a conventional airfoil (the Clark Y in this case). The tests were made with various flap settings and with the entrance to the passage both open and closed. The highest lift coefficient found, C(sub L) = 2.08, was obtained with the passage closed.
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.
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.
Method for forming a liquid cooled airfoil for a gas turbine
Grondahl, Clayton M. (Clifton Park, NY); Willmott, Leo C. (Ballston Spa, NY); Muth, Myron C. (Amsterdam, NY)
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.
NASA Astrophysics Data System (ADS)
Bosnyakov, S.; Kazhan, E.; Kursakov, I.; Matyash, S.; Mikhaylov, S.; Lysenkov, A.
2015-06-01
The current research concerns a half-model high-lift configuration inside the European Transonic Wind Tunnel (ETW) at landing regime. The influence of the wind-tunnel walls (both slotted and closed) is investigated and the numerical results are compared with measured data. The investigated model is a three-element landing configuration with Krueger device and flap. All calculations are performed on structured grids using EWT-TsAGI code. The computed in-tunnel results are in good agreement with uncorrected experimental data, with maximum lift predicted at the same angle of attack. The slotted wall configuration produces less wall interference than the closed wall configuration.
Design of a laminar-flow-control supercritical airfoil for a swept wing
NASA Technical Reports Server (NTRS)
Allison, D. O.; Dagenhart, J. R.
1978-01-01
An airfoil was analytically designed and analyzed for a combination of supercritical flow and laminar flow control (LFC) by boundary layer suction. A shockless inverse method was used to design an airfoil and an analysis method was used in lower surface redesign work. The laminar flow pressure distributions were computed without a boundary layer under the assumption that the laminar boundary layer would be kept thin by suction. Inviscid calculations showed that this 13.5 percent thick airfoil has shockless flows for conditions at and below the design normal Mach number of 0.73 and the design section lift coefficient of 0.60, and that the maximum local normal Mach number is 1.12 at the design point. The laminar boundary layer instabilities can be controlled with suction but the undercut leading edge of the airfoil provides a low velocity, constant pressure coefficients region which is conducive to laminar flow without suction. The airfoil was designed to be capable of lift recovery with no suction by the deflection of a small trailing edge flap.
Lift hysteresis at stall as an unsteady boundary-layer phenomenon
NASA Technical Reports Server (NTRS)
Moore, Franklin K
1956-01-01
Analysis of rotating stall of compressor blade rows requires specification of a dynamic lift curve for the airfoil section at or near stall, presumably including the effect of lift hysteresis. Consideration of the magnus lift of a rotating cylinder suggests performing an unsteady boundary-layer calculation to find the movement of the separation points of an airfoil fixed in a stream of variable incidence. The consideration of the shedding of vorticity into the wake should yield an estimate of lift increment proportional to time rate of change of angle of attack. This increment is the amplitude of the hysteresis loop. An approximate analysis is carried out according to the foregoing ideas for a 6:1 elliptic airfoil at the angle of attack for maximum lift. The assumptions of small perturbations from maximum lift are made, permitting neglect of distributed vorticity in the wake. The calculated hysteresis loop is counterclockwise. Finally, a discussion of the forms of hysteresis loops is presented; and, for small reduced frequency of oscillation, it is concluded that the concept of a viscous "time lag" is appropriate only for harmonic variations of angle of attack with time at mean conditions other than maximum lift.
Tests of NACA 0009, 0012, and 0018 Airfoils in the Full-Scale Tunnel
NASA Technical Reports Server (NTRS)
Goett, Harry J; Bullivant, W Kenneth
1939-01-01
An investigation was conducted in the NACA full-scale wind tunnel to determine the aerodynamic characteristics of the NACA 0009, 0012, and 0018 airfoils, with the ultimate purpose of providing data to be used as a basis for comparison with other wind-tunnel data, mainly in the study of scale and turbulence effects. Three symmetrical 6 by 36-foot rectangular airfoils were used. The Reynolds number range for minimum drag was form 1,800,000 to 7,000,000 and for maximum lift, from 1,700,000 to 4,500,000. The effect of rounded tips was determined for each of the airfoils. Tests were also made of the NACA 0012 airfoil equipped with a 0.20c full-span split flap hinged at 0.80c. Tuft surveys were included to show the progressive breakdown of flow near maximum lift. Momentum surveys were made in conjunction with force measurements at zero lift as an aid in converting force-test data to section coefficients.
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. PMID:24116405
NASA Technical Reports Server (NTRS)
Krenkel, A. R.
1978-01-01
The finite-step method was programmed for computing the span loading and stability derivatives of trapezoidal shaped wings in symmetric, yawed, and rotary flight. Calculations were made for a series of different wing planforms and the results compared with several available methods for estimating these derivatives in the linear angle of attack range. The agreement shown was generally good except in a few cases. An attempt was made to estimate the nonlinear variation of lift with angle of attack in the high alpha range by introducing the measured airfoil section data into the finite-step method. The numerical procedure was found to be stable only at low angles of attack.
NASA Technical Reports Server (NTRS)
Brooks, Cuyler W., Jr.; Harris, Charles D.
1987-01-01
A large chord swept supercritical laminar-flow control (LFC) airfoil was designed, constructed, and tested in the Langley 8-foot Transonic Pressure Tunnel (TPT). The LFC airfoil experiment was established to provide basic information concerning the design and compatibility of high performance supercritical airfoils with suction boundary-layer control achieved through fine slots or porous surface concepts. Shockless pressure distribution was achieved. Full chord laminar flow was achieved on upper and lower surfaces. Full chord laminar flow was maintained at subcritical speeds and over large supercritical zones. Feasibility of combined suction laminarization and supercritical airfoil technology was demonstrated.
NASA Technical Reports Server (NTRS)
Platt, Robert C; Abbott, Ira H
1937-01-01
Report presents the results of an investigation of the general aerodynamic characteristics of the NACA 23012 and 23021 airfoils, each equipped with a 0.20c external flap of NACA 23012 section. The tests were made in the NACA 7 by 10-foot and variable-density wind tunnels and covered a range of Reynolds numbers that included values corresponding to those for landing conditions of a wide range of airplanes. Besides a determination of the variation of lift and drag characteristics with position of the flap relative to the main airfoil, complete aerodynamic characteristics of the airfoil-flap combination with a flap hinge axis selected to give small hinge moments were measured in the two tunnels. Some measurements of air loads on the flap itself in the presence of the wing were made in the 7 by 10-foot wind tunnel.
An interactive boundary-layer procedure for oscillating airfoils including transition effects
NASA Technical Reports Server (NTRS)
Cebeci, Tuncer; Jang, Hong-Ming; Carr, L. W.
1989-01-01
An interactive boundary-layer method previously developed and tested for steady flows is used here in a quasi-steady manner to examine the evolution of the flow behavior around oscillating airfoils operating inlight stall conditions. The calculations encompass the airfoil and wake flows at angles of attack which lead to flow separation. The location of the onset of transition is represented by a correlation based on steady flows. The results show the large effects of the viscous layer on the variation of lift coefficient with angle of attack and reduced frequency.
Semi-empirical model for prediction of unsteady forces on an airfoil with application to flutter
NASA Technical Reports Server (NTRS)
Mahajan, Aparajit J.; Kaza, Krishna Rao V.
1992-01-01
A semi-empirical model is described for predicting unsteady aerodynamic forces on arbitrary airfoils under mildly stalled and unstalled conditions. Aerodynamic forces are modeled using second order ordinary differential equations for lift and moment with airfoil motion as the input. This model is simultaneously integrated with structural dynamics equations to determine flutter characteristics for a two degrees-of-freedom system. Results for a number of cases are presented to demonstrate the suitability of this model to predict flutter. Comparison is made to the flutter characteristics determined by a Navier-Stokes solver and also the classical incompressible potential flow theory.
Tests of Large Airfoils in the Propeller Research Tunnel, Including Two with Corrugated Surfaces
NASA Technical Reports Server (NTRS)
Wood, Donald H
1930-01-01
This report gives the results of the tests of seven 2 by 12 foot airfoils (Clark Y, smooth and corrugated, Gottingen 398, N.A.C.A. M-6, and N.A.C.A. 84). The tests were made in the propeller research tunnel of the National Advisory Committee for Aeronautics at Reynolds numbers up to 2,000,000. The Clark Y airfoil was tested with three degrees of surface smoothness. Corrugating the surface causes a flattening of the lift curve at the burble point and an increase in drag at small flying angles.
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.
Comparison of measured and calculated aircraft lift generated pressures
NASA Technical Reports Server (NTRS)
Findley, D. S.
1975-01-01
Lift generated pressures produced by large, heavy aircraft at low altitudes were investigated due to concern over their possible effects on ground objects. Aircraft lift generated pressures were calculated using elementary airfoil theory, and these values were compared with ground level measurements made during an overflight program. The predicted and the measured values were in relatively good agreement. Due to lack of experimental investigations of this phenomenon, opportunity was taken during an overflight program to use a specially instrumented test range to measure the ground pressures produced for a range of aircraft weights and distances.
Total facelift: forehead lift, midface lift, and neck lift.
Park, Dong Man
2015-03-01
Patients with thick skin mainly exhibit the aging processes of sagging, whereas patients with thin skin develop wrinkles or volume loss. Asian skin is usually thicker than that of Westerners; and thus, the sagging of skin due to aging, rather than wrinkling, is the chief problem to be addressed in Asians. Asian skin is also relatively large in area and thick, implying that the weight of tissue to be lifted is considerably heavier. These factors account for the difficulties in performing a facelift in Asians. Facelifts can be divided into forehead lift, midface lift, and lower face lift. These can be performed individually or with 2-3 procedures combined. PMID:25798381
Control of Flow Separation Using Adaptive Airfoils
NASA Technical Reports Server (NTRS)
Chandrasekhara, M. S.; Wilder, M. C.; Carr, L. W.; Davis, Sanford S. (Technical Monitor)
1996-01-01
A novel way of controlling flow separation is reported. The approach involves using an adaptive airfoil geometry that changes its leading edge shape to adjust to the instantaneous flow at high angles of attack such that the flow over it remains attached. In particular, a baseline NACA 0012 airfoil, whose leading edge curvature could be changed dynamically by 400% was tested under quasi-steady compressible flow conditions. A mechanical drive system was used to produce a rounded leading edge to reduce the strong local flow acceleration around its nose and thus reduce the strong adverse pressure gradient that follows such a rapid acceleration. Tests in steady flow showed that at M = 0.3, the flow separated at about 14 deg. angle of attack for the NACA 0012 profile but could be kept attached up to an angle of about 18 deg by changing the nose curvature. No significant hysteresis effects were observed; the flow could be made to reattach from its separated state at high angles by changing the leading edge curvature. Interestingly, the flow over a nearly semicircular nosed airfoil was separated even at low angles.
Circulation methods for lift measurement in an unsteady aerodynamic flow
NASA Astrophysics Data System (ADS)
Yuan, Jiankun; Olinger, David J.
2001-11-01
Recently an ultrasonic circulation measurement technique has been developed to determine mean lift on various aerodynamic bodies, including airfoils and delta wings^1. The ultrasonic method is based on measuring transit times of acoustic pulses along paths enclosing the aerodynamic body. In the present work, the largely unstudied extension of the ultrasonic technique to determine instantaneous lift in unsteady flows is addressed. An analytical method, based on unsteady potential flow theory, is developed which properly converts the instantaneous bound circulation measured in the ultrasonic technique to instantaneous lift. The analytical method is validated using a finite element solution of flow over an oscillating flat plate at low angle of attack and Reynolds number = 100. The bound circulation within an enclosed acoustic path is determined from the numerical simulation and converted to instantaneous lift via the developed method. Comparing these values to instantaneous lift values determined from integration of the surface pressure distribution shows that the developed method provides improved results compared to a quasi-steady Kutta-Joukowski theorem used in past investigations. The improvement in the proper phase of the instantaneous lift force is nominally larger than the improvement for rms lift amplitudes. The numerical study also shows that proper selection of the enclosing acoustic path is critical for accurate lift measurement. ^1H. Johari, & W. Durgin, Expts. in Fluids 25, 1 (1998).
Nozzle airfoil having movable nozzle ribs
Yu, Yufeng Phillip (Greenville, SC); Itzel, Gary Michael (Greenville, SC)
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)
Civil applications of high-speed rotorcraft and powered-lift aircraft configurations
NASA Technical Reports Server (NTRS)
Albers, James A.; Zuk, John
1987-01-01
Advanced subsonic vertical and short takeoff and landing (V/STOL) aircraft configurations offer new transportation options for civil applications. Described is a range of vehicles from low-disk to high-disk loading aircraft, including high-speed rotorcraft, V/STOL aircraft, and short takeoff and landing (STOL) aircraft. The status and advantages of the various configurations are described. Some of these show promise for relieving congestion in high population-density regions and providing transportation opportunities for low population-density regions.
Transonic Buffet over Symmetric Airfoils
Alexander Kuz’min; Alexey Shilkin
Supercritical airfoils typically have small curvature of the upper surface as compared with that of a conventional airfoil.\\u000a At the same time, physical phenomena associated with transonic flow over flattened surfaces have not been well understood.\\u000a For instance, the inviscid steady flow over such surfaces may exhibit an unexpected bifurcation and intricate instability\\u000a of its entire structure at certain freestream
NASA Technical Reports Server (NTRS)
Hague, D. S.; Merz, A. W.
1975-01-01
Multivariable search techniques are applied to a particular class of airfoil optimization problems. These are the maximization of lift and the minimization of disturbance pressure magnitude in an inviscid nonlinear flow field. A variety of multivariable search techniques contained in an existing nonlinear optimization code, AESOP, are applied to this design problem. These techniques include elementary single parameter perturbation methods, organized search such as steepest-descent, quadratic, and Davidon methods, randomized procedures, and a generalized search acceleration technique. Airfoil design variables are seven in number and define perturbations to the profile of an existing NACA airfoil. The relative efficiency of the techniques are compared. It is shown that elementary one parameter at a time and random techniques compare favorably with organized searches in the class of problems considered. It is also shown that significant reductions in disturbance pressure magnitude can be made while retaining reasonable lift coefficient values at low free stream Mach numbers.
NASA Technical Reports Server (NTRS)
Morris, C. E. K., Jr.; Stevens, D. D.; Tomaine, R. L.
1980-01-01
A flight investigation was conducted using a teetering-rotor AH-1G helicopter to obtain data on the aerodynamic behavior of main-rotor blades with the NLR-1T blade section. The data system recorded blade-section aerodynamic pressures at 90 percent rotor radius as well as vehicle flight state, performance, and loads. The test envelope included hover, forward flight, and collective-fixed maneuvers. Data were obtained on apparent blade-vortex interactions, negative lift on the advancing blade in high-speed flight and wake interactions in hover. In many cases, good agreement was achieved between chordwise pressure distributions predicted by airfoil theory and flight data with no apparent indications of blade-vortex interactions.
Subsonic natural-laminar-flow airfoils
NASA Technical Reports Server (NTRS)
Somers, Dan M.
1992-01-01
An account is given of the development history of natural laminar-flow (NLF) airfoil profiles under guidance of an experimentally well-verified theoretical method for the design of airfoils suited to virtually all subcritical applications. This method, the Eppler Airfoil Design and Analysis Program, contains a conformal-mapping method for airfoils having prescribed velocity-distribution characteristics, as well as a panel method for the analysis of potential flow about given airfoils and a boundary-layer method. Several of the NLF airfoils thus obtained are discussed.
Y. Mokuno; A. Chayahara; H. Yamada; N. Tsubouchi
2009-01-01
The lift-off process using ion implantation has recently been applied to produce large and thick single-crystal diamond plates by chemical vapor deposition (CVD). CVD growth conditions for undoped, as opposed to nitrogen-doped, diamond were investigated to improve the purity of plates produced by this technique. This utilized apparatus identical to that for high-rate growth with nitrogen addition under high-density plasma.
NASA Technical Reports Server (NTRS)
vanDam, C. P.; Los, S. M.; Miley, S. J.; Yip, L. P.; Banks, D. W.; Roback, V. E.; Bertelrud, A.
1995-01-01
Flight experiments on NASA Langley's B737-100 (TSRV) airplane have been conducted to document flow characteristics in order to further the understanding of high-lift flow physics, and to correlate and validate computational predictions and wind-tunnel measurements. The project is a cooperative effort involving NASA, industry, and universities. In addition to focusing on in-flight measurements, the project includes extensive application of various computational techniques, and correlation of flight data with computational results and wind-tunnel measurements. Results obtained in the most recent phase of flight experiments are analyzed and presented in this paper. In-flight measurements include surface pressure distributions, measured using flush pressure taps and pressure belts on the slats, main element, and flap elements; surface shear stresses, measured using Preston tubes; off-surface velocity distributions, measured using shear-layer rakes; aeroelastic deformations of the flap elements, measured using an optical positioning system; and boundary-layer transition phenomena, measured using hot-film anemometers and an infrared imaging system. The analysis in this paper primarily focuses on changes in the boundary-layer state that occurred on the slats, main element, and fore flap as a result of changes in flap setting and/or flight condition. Following a detailed description of the experiment, the boundary-layer state phenomenon will be discussed based on data measured during these recent flight experiments.
NASA Technical Reports Server (NTRS)
Bergrun, Norman R
1952-01-01
An empirically derived basis for predicting the area, rate, and distribution of water-drop impingement on airfoils of arbitrary section is presented. The concepts involved represent an initial step toward the development of a calculation technique which is generally applicable to the design of thermal ice-prevention equipment for airplane wing and tail surfaces. It is shown that sufficiently accurate estimates, for the purpose of heated-wing design, can be obtained by a few numerical computations once the velocity distribution over the airfoil has been determined. The calculation technique presented is based on results of extensive water-drop trajectory computations for five airfoil cases which consisted of 15-percent-thick airfoils encompassing a moderate lift-coefficient range. The differential equations pertaining to the paths of the drops were solved by a differential analyzer.
AIAA 2002-1717 Flutter Suppression for High Aspect Ratio
Stanford University
The Gurney flap is a small (typically 1% 5% chord) flap used to increase the maximum lift of an airfoil on both sin- gle element and multi-element airfoils with Gurney flaps.16 These studies confirm that despite their small size, Gurney flaps with deflections near 90 de- grees can increase maximum lift
Analysis of oscillatory motion of a light airplane at high values of lift coefficient
NASA Technical Reports Server (NTRS)
Batterson, J. G.
1983-01-01
A modified stepwise regression is applied to flight data from a light research air-plane operating at high angles at attack. The well-known phenomenon referred to as buckling or porpoising is analyzed and modeled using both power series and spline expansions of the aerodynamic force and moment coefficients associated with the longitudinal equations of motion.
Design and Predictions for High-Altitude (Low Reynolds Number) Aerodynamic Flight Experiment
NASA Technical Reports Server (NTRS)
Greer, Donald; Harmory, Phil; Krake, Keith; Drela, Mark
2000-01-01
A sailplane being developed at NASA Dryden Flight Research Center will support a high-altitude flight experiment. The experiment will measure the performance parameters or an airfoil at high altitudes (70,000 - 100,000 ft), low Reynolds numbers (2 x 10(exp 5) - 7 x 10(exp 5)), and high subsonic Mach numbers (0.5 and 0.65). The airfoil section lift and drag are determined from pilot and static pressure measurements. The locations of the separation bubble, Tollmien-Schlichting boundary-layer instability frequencies, and vortex shedding are measured from a hot-film strip. The details of the planned flight experiment are presented as well as several predictions of the airfoil performance.
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.
An airfoil flutter model suspension system to accommodate large static transonic airloads
NASA Technical Reports Server (NTRS)
Reed, W. H., III
1985-01-01
A pitch/plunge flutter model suspension system and associated two-dimensional MBB-A3 airfoil models is described. The system is designed for installation in the Langley 6-by-19-inch and 6-by-18-inch transonic blowdown wind tunnels to enable systematic study of the transonic flutter characteristics and static pressure distributions of supercritical airfoils at transonic Mach numbers. A compound spring suspension concept is introduced which simultaneously meets requirements for low plunge-mode stiffness, lightweight suspended model, and large steady lift due to angle of attack without the need for excessive static deflections of the plunge spring. The system features variable pitch and plunge frequencies, changeable airfoil rotation axes, and a self aligning control system to maintain a constant mean position of the model with changing airload.
NASA Technical Reports Server (NTRS)
Pulliam, T. H.; Nemec, M.; Holst, T.; Zingg, D. W.; Kwak, Dochan (Technical Monitor)
2002-01-01
A comparison between an Evolutionary Algorithm (EA) and an Adjoint-Gradient (AG) Method applied to a two-dimensional Navier-Stokes code for airfoil design is presented. Both approaches use a common function evaluation code, the steady-state explicit part of the code,ARC2D. The parameterization of the design space is a common B-spline approach for an airfoil surface, which together with a common griding approach, restricts the AG and EA to the same design space. Results are presented for a class of viscous transonic airfoils in which the optimization tradeoff between drag minimization as one objective and lift maximization as another, produces the multi-objective design space. Comparisons are made for efficiency, accuracy and design consistency.
Wind-tunnel investigation of NACA 23012 airfoil with various arrangements of slotted flaps
NASA Technical Reports Server (NTRS)
Wenzinger, Carl J; Harris, Thomas A
1939-01-01
An investigation was made in the NACA 7 by 10-foot wind tunnel of a large-chord NACA 23012 airfoil with several arrangements of 25.66 percent chord slotted flaps to determine the section aerodynamic characteristics as affected by slot shape, flap shape, flap location, and flap deflection. The flap position for maximum lift, the polars for arrangements considered favorable for take-off and climb, and the complete section aerodynamic characteristics for selected optimum arrangements were determined. A discussion is given of the relative merits of the various arrangement for certain selected criterions. A comparison is made of a slotted flap on the NACA 23021 airfoil with a corresponding slotted flap previously developed for the NACA 23012 airfoil.
The effects of actuation frequency on the separation control over an airfoil using a synthetic jet
NASA Astrophysics Data System (ADS)
Abe, Y.; Okada, K.; Nonomura, T.; Fujii, K.
2015-06-01
The simulation of separation control using a synthetic jet (SJ) is conducted around an NACA (National Advisory Committee for Aeronautics) 0015 airfoil by large-eddy simulation (LES) with a compact difference scheme. The synthetic jet is installed at the leading edge of the airfoil and the effects of an actuation frequency F+ (normalized by chord length and velocity of freestream) are observed. The lift-drag coefficient is recovered the most for F+ = 6. The relationship between momentum addition by turbulent mixing and large vortex structures is investigated using a phase-averaging procedure based on F+. The Reynolds shear stress is decomposed into periodic and turbulent components where the turbulent components are found to be dominant on the airfoil. The strong turbulent components appear near the large vortex structures that are observed in phase- and span-averaged flow fields.
Cadwallader, L.C.
1997-03-01
This report presents safety information about powered industrial trucks. The basic lift truck, the counterbalanced sit down rider truck, is the primary focus of the report. Lift truck engineering is briefly described, then a hazard analysis is performed on the lift truck. Case histories and accident statistics are also given. Rules and regulations about lift trucks, such as the US Occupational Safety an Health Administration laws and the Underwriter`s Laboratories standards, are discussed. Safety issues with lift trucks are reviewed, and lift truck safety and reliability are discussed. Some quantitative reliability values are given.
Transonic single-mode flutter and buffet of a low aspect ratio wing having a subsonic airfoil shape
NASA Technical Reports Server (NTRS)
Erickson, L. L.
1974-01-01
Transonic flutter and buffet results obtained from wind-tunnel tests of a low aspect ratio semispan wing model are presented. The tests were conducted to investigate potential transonic aeroelastic problems of vehicles having subsonic airfoil sections. The model employed NACA 00XX-64 airfoil sections in the streamwise direction and had a 14 deg leading edge sweep angle. Aspect ratio, and average thickness were 4.0, 0.35, and 8 percent, respectively. The model was tested at Mach numbers from 0.6 to 0.95 at angles of attack from 0 deg to 15 deg. Two zero lift flutter conditions were found that involved essentially single normal mode vibrations. With boundary layer trips on the model, flutter occurred in a narrow Mach number range centered at about Mach 0.90. The frequency and motion of this flutter were like that of the first normal mode vibration. With the trips removed flutter occurred at a slightly high Mach number but in a mode strongly resembling that of the second normal mode.
A Comparative Study Using CFD to Predict Iced Airfoil Aerodynamics
NASA Technical Reports Server (NTRS)
Chi, x.; Li, Y.; Chen, H.; Addy, H. E.; Choo, Y. K.; Shih, T. I-P.
2005-01-01
WIND, Fluent, and PowerFLOW were used to predict the lift, drag, and moment coefficients of a business-jet airfoil with a rime ice (rough and jagged, but no protruding horns) and with a glaze ice (rough and jagged end has two or more protruding horns) for angles of attack from zero to and after stall. The performance of the following turbulence models were examined by comparing predictions with available experimental data. Spalart-Allmaras (S-A), RNG k-epsilon, shear-stress transport, v(sup 2)-f, and a differential Reynolds stress model with and without non-equilibrium wall functions. For steady RANS simulations, WIND and FLUENT were found to give nearly identical results if the grid about the iced airfoil, the turbulence model, and the order of accuracy of the numerical schemes used are the same. The use of wall functions was found to be acceptable for the rime ice configuration and the flow conditions examined. For rime ice, the S-A model was found to predict accurately until near the stall angle. For glaze ice, the CFD predictions were much less satisfactory for all turbulence models and codes investigated because of the large separated region produced by the horns. For unsteady RANS, WIND and FLUENT did not provide better results. PowerFLOW, based on the Lattice Boltzmann method, gave excellent results for the lift coefficient at and near stall for the rime ice, where the flow is inherently unsteady.
NASA Astrophysics Data System (ADS)
Paskova, T.; Darakchieva, V.; Paskov, P. P.; Monemar, B.; Bukowski, M.; Suski, T.; Ashkenov, N.; Schubert, M.; Hommel, D.
2006-06-01
We have studied the effects of laser lift-off and polishing processes on the bending of free-standing HVPE grown GaN thick films. Their structural characteristics were accessed by reciprocal space mapping and lattice parameters measurements as well as by Raman scattering and photoluminescence. The in-plane strain difference between the two faces was found to have determining effect on the bending of the free-standing films. Removing the high-defect-density near-interface region either by melting caused by laser lift-off, or by polishing, or by point defects dissociation caused by high-pressure annealing was found to lead to flattening of the strain distribution along the film thickness and a significant reduction of the bending of the free-standing films.
NASA Technical Reports Server (NTRS)
Ladson, Charles L.; Hill, Acquilla S.; Johnson, William G., Jr.
1987-01-01
Tests were conducted in the 2-D test section of the Langley 0.3-meter Transonic Cryogenic Tunnel on a NACA 0012 airfoil to obtain aerodynamic data as a part of the Advanced Technology Airfoil Test (ATAT) program. The test program covered a Mach number range of 0.30 to 0.82 and a Reynolds number range of 3.0 to 45.0 x 10 to the 6th power. The stagnation pressure was varied between 1.2 and 6.0 atmospheres and the stagnation temperature was varied between 300 K and 90 K to obtain these test conditions. Tabulated pressure distributions and integrated force and moment coefficients are presented as well as plots of the surface pressure distributions. The data are presented uncorrected for wall interference effects and without analysis.
NASA Technical Reports Server (NTRS)
Lichtenstein, Jacob H
1952-01-01
Results of a low-speed wind-tunnel investigation to determine the effect of high-lift devices on the static lateral stability derivatives and the yawing derivatives of an untapered 45 degrees sweptback wing are presented. The tests were made in the curved-flow test section of the Langley stability tunnel at a Reynolds number of 1.1 X 10 to the sixth power.
Sven Eisenbach; Rainer Friedrich
2008-01-01
Incompressible flow separating from the upper surface of an airfoil at an 18° angle of attack and a Reynolds number of Re = 105, based on the freestream velocity and chord length c, is studied by the means of large-eddy simulation (LES). The numerical method is based on second-order central spatial discretization\\u000a on a Cartesian grid using an immersed boundary
Particle image velocimetry investigation of flow over unsteady airfoil with trailing-edge strip
NASA Astrophysics Data System (ADS)
Gerontakos, P.; Lee, T.
2008-04-01
The flow over a flapped NACA 0012 airfoil, oscillated slightly through the static-stall angle, was investigated by using particle image velocimetry, and was supplemented by surface pressure and dynamic-load measurements. A significant increase in the dynamic lift force and nose-down pitching moment was observed. The most pronounced flow phenomenon was the formation and detachment of an energetic leading-edge vortex compared to the no-flapped airfoil. The details of the underlying physical mechanisms responsible for the various light-stall flow processes were provided via the instantaneous velocity and vorticity fields measurements. In contrast to the Gurney flap, the inverted trailing-edge strip led to an improved negative damping while a reduced lift force. The addition of an inverted strip always led to the appearance of a Karman-type vortex shedding street immediately downstream of the strip over the entire oscillation cycle.
A Computational Study of an Oscillating VR-12 Airfoil with a Gurney Flap
NASA Technical Reports Server (NTRS)
Rhee, Myung
2004-01-01
Computations of the flow over an oscillating airfoil with a Gurney-flap are performed using a Reynolds Averaged Navier-Stokes code and compared with recent experimental data. The experimental results have been generated for different sizes of the Gurney flaps. The computations are focused mainly on a configuration. The baseline airfoil without a Gurney flap is computed and compared with the experiments in both steady and unsteady cases for the purpose of initial testing of the code performance. The are carried out with different turbulence models. Effects of the grid refinement are also examined and unsteady cases, in addition to the assessment of solver effects. The results of the comparisons of steady lift and drag computations indicate that the code is reasonably accurate in the attached flow of the steady condition but largely overpredicts the lift and underpredicts the drag in the higher angle steady flow.
Development of the Risø wind turbine airfoils
Peter Fuglsang; Christian Bak
2004-01-01
This paper presents the wind turbine airfoil development at Risø. The design method is described together with our target characteristics for wind turbine airfoils. The use of the CFD code Ellipsys2D for prediction of final target characteristics is described together with the VELUX wind tunnel testing setup. Three airfoil families were developed; Risø-A1, Risø-P and Risø-B1. The Risø-A1 airfoil family
Acoustic effects on profile drag of a laminar flow airfoil
NASA Astrophysics Data System (ADS)
Shearin, John G.; Jones, Michael G.; Baals, Robert A.
1987-09-01
A two-dimensional laminar flow airfoil (NLF-0414) was subjected to high-intensity sound (pure tones and white noise) over a frequency range of 2 to 5 kHz, while immersed in a flow of 240 ft/sec (Rn of 3 million) in a quiet flow facility. Using a wake-rake, wake dynamic pressures were determined and the deficit in momentum was used to calculate a two dimensional drag coefficient. Significant increases in drag were observed when the airfoil was subjected to the high intensity sound at critical sound frequencies. However, the increased drag was not accompanied by movement of the transition location.
Acoustic effects on profile drag of a laminar flow airfoil
NASA Technical Reports Server (NTRS)
Shearin, John G.; Jones, Michael G.; Baals, Robert A.
1987-01-01
A two-dimensional laminar flow airfoil (NLF-0414) was subjected to high-intensity sound (pure tones and white noise) over a frequency range of 2 to 5 kHz, while immersed in a flow of 240 ft/sec (Rn of 3 million) in a quiet flow facility. Using a wake-rake, wake dynamic pressures were determined and the deficit in momentum was used to calculate a two dimensional drag coefficient. Significant increases in drag were observed when the airfoil was subjected to the high intensity sound at critical sound frequencies. However, the increased drag was not accompanied by movement of the transition location.
Dynamic stall on a supercritical airfoil
W. Geissler; G. Dietz; H. Mai
2005-01-01
For the present investigations of dynamic stall a supercritical airfoil was chosen. This new airfoil designed by DLR will be used in dynamic stall control research activities (project ADASYS) planned for the near future: the leading edge portion of the airfoil will be drooped down dynamically to improve dynamic stall characteristics on the retreating side during blade motion. The optimised
Darrieus wind-turbine airfoil configurations
P. G. Migliore; J. R. Fritschen
1982-01-01
The purpose was to determine what aerodynamic performance improvement, if any, could be achieved by judiciously choosing the airfoil sections for Darrieus wind turbine blades. 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.
Knowledge Based Airfoil Aerodynamic and Aeroacoustic Design
B. Greschner; C. Yu; S. Zheng; M. Zhuang; Z. J. Wang; F. Thiele
A systematic investigation of the unsteady flows around a series of NACA airfoils is carried out. The main objective is to conduct manual design case studies on the connections between an airfoil shape characteristics and it s aerodynamic and aeroacoustic performance. The approach employs the unsteady CFD flow simulations in the near field of an airfoil and the FW-H integral
Unknown
2011-08-17
enhanced production of NO by LECs. Additionally, high concentrations of NO have been experimentally observed in the sinus region of mesenteric lymphatic vessels. The goal of this work was to develop a computational flow and mass transfer model using...
A supercritical airfoil experiment
NASA Technical Reports Server (NTRS)
Mateer, George G.; Seegmiller, H. Lee; Hand, Lawrence A.; Szodruch, Joachim
1994-01-01
The purpose of this investigation is to provide a comprehensive data base for the validation of numerical simulations. The initial results of the study (single angle of attack) were presented in ref. 1, where the effects of various parameters and the adequacies of selected turbulence models were discussed. 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 corrections 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.
Particle image velocimetry investigation of flow over unsteady airfoil with trailing-edge strip
P. Gerontakos; T. Lee
2008-01-01
The flow over a flapped NACA 0012 airfoil, oscillated slightly through the static-stall angle, was investigated by using particle\\u000a image velocimetry, and was supplemented by surface pressure and dynamic-load measurements. A significant increase in the dynamic\\u000a lift force and nose-down pitching moment was observed. The most pronounced flow phenomenon was the formation and detachment\\u000a of an energetic leading-edge vortex compared
Time resolved PIV analysis of flow over a NACA 0015 airfoil with Gurney flap
D. R. Troolin; E. K. Longmire; W. T. Lai
2006-01-01
A NACA 0015 airfoil with and without a Gurney flap was studied in a wind tunnel with Re c = 2.0 × 105 in order to examine the evolving flow structure of the wake through time-resolved PIV and to correlate this structure with time-averaged measurements of the lift coefficient. The Gurney flap, a tab of small length (1-4% of the
Martin R. Fink
1991-01-01
Propulsive life noise is the increase in noise that occurs when airframe surfaces are placed in the propulsive system's exhaust to increase their lift force. Increased local flow velocities and turbulence levels, due to the propulsive system exhaust gases passing along the airframe lifting surfaces, cause an increase in maximum lift coefficient. The airplane's flight speed required for takeoff and
The Effect of a Flap and Ailerons on the N.A.C.A. M-6 Airfoil Section
NASA Technical Reports Server (NTRS)
Higgins, George J; Jacobs, Eastman N
1928-01-01
This report contains the results obtained at the Langley Memorial Aeronautical Laboratory on an N. A. C. A. M-6 airfoil, fitted with a flap and ailerons, and tested in the variable density wind tunnel at a density of 20 atmospheres. Airfoil characteristics are given for the model up to 48 degree angle of attack with the flap set at various angles, and also with the ailerons set at similar angles. The approximate lift distribution and the center of pressure variation along the span are determined with the model at 18 degree angle of attack and with the ailerons displaced at 20 degrees. Approximate rolling moment and yawing moment coefficients are determined for the various aileron settings. A comparison of the calculated angles of zero lift and the calculated lift and moment coefficiencies with those observed is given in the appendix.
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 roughness and freestream turbulence, compared with data from the cambered vane airfoil. Stanton numbers, skin friction coefficients, aerodynamic losses, and Reynolds analogy behavior are numerically predicted for a turbine vane using the FLUENT with a k-epsilon RNG model to show the effects of Mach number, mainstream turbulence level, and surface roughness. Comparisons with wake aerodynamic loss experimental data are made. Numerically predicted skin friction coefficients and Stanton numbers are also used to deduce Reynolds analogy behavior on the vane suction and pressure sides.
NASA Technical Reports Server (NTRS)
Rashid, J. M.; Freling, M.; Friedrich, L. A.
1987-01-01
The ability of coatings to provide at least a 2X improvement in particulate erosion resistance for steel, nickel and titanium compressor airfoils was identified and demonstrated. Coating materials evaluated included plasma sprayed cobalt tungsten carbide, nickel carbide and diffusion applied chromium plus boron. Several processing parameters for plasma spray processing and diffusion coating were evaluated to identify coating systems having the most potential for providing airfoil erosion resistance. Based on laboratory results and analytical evaluations, selected coating systems were applied to gas turbine blades and evaluated for surface finish, burner rig erosion resistance and effect on high cycle fatigue strength. Based on these tests, the following coatings were recommended for engine testing: Gator-Gard plasma spray 88WC-12Co on titanium alloy airfoils, plasma spray 83WC-17Co on steel and nickel alloy airfoils, and Cr+B on nickel alloy airfoils.
NASA Technical Reports Server (NTRS)
Young, John W.
1961-01-01
A study has been made of a method for controlling the trajectory of a high-drag low-lift entry vehicle to a desired longitude and latitude on the surface of a rotating earth. By use of this control technique the vehicle can be guided to the desired point when the present position and heading of the vehicle are known and the desired longitude and latitude are specified. The present study makes use of a single reference trajectory and an estimate of the lift and side-force capabilities of the vehicle. This information is stored in a control-logic system and used with linear control equations to guide the vehicle to the desired destination. Results are presented of a number of trajectory studies which describe the operation of the control system and illustrate its ability to control the vehicle trajectory to the desired landing area.
Highly nonplanar lifting systems
NASA Technical Reports Server (NTRS)
Kroo, Ilan; McMasters, John; Smith, Stephen C.
1996-01-01
This paper deals with nonplanar wing concepts -- their advantages and possible applications in a variety of aircraft designs. A brief review and assessment of several concepts from winglets to ring wings is followed by a more detailed look at two recent ideas: exploiting nonplanar wakes to reduce induced drag, and applying a 'C-wing' design to large commercial transports. Results suggest that potential efficiency gains may be significant, while several nonaerodynamic characteristics are particularly interesting.
Time resolved PIV analysis of flow over a NACA 0015 airfoil with Gurney flap
NASA Astrophysics Data System (ADS)
Troolin, D. R.; Longmire, E. K.; Lai, W. T.
2006-08-01
A NACA 0015 airfoil with and without a Gurney flap was studied in a wind tunnel with Re c = 2.0 × 105 in order to examine the evolving flow structure of the wake through time-resolved PIV and to correlate this structure with time-averaged measurements of the lift coefficient. The Gurney flap, a tab of small length (1-4% of the airfoil chord) that protrudes perpendicular to the chord at the trailing edge, yields a significant and relatively constant lift increment through the linear range of the C L versus ? curve. Two distinct vortex shedding modes were found to exist and interact in the wake downstream of flapped airfoils. The dominant mode resembles a Kàrmàn vortex street shedding behind an asymmetric bluff body. The second mode, which was caused by the intermittent shedding of fluid recirculating in the cavity upstream of the flap, becomes more coherent with increasing angle of attack. For a 4% Gurney flap at ? = 8°, the first and second modes corresponded with Strouhal numbers based on flap height of 0.18 and 0.13. Comparison of flow around ‘filled’ and ‘open’ flap configurations suggested that the second shedding mode was responsible for a significant portion of the overall lift increment.
Measurement and Formation of Three-Dimensional Stall Cells on Two-Dimensional Airfoils
NASA Astrophysics Data System (ADS)
Sivaneri, Victor
The present work aims to identify, quantify, and understand the physics by which three-dimensional stall cells are formed on two-dimensional airfoils, using oil flow visualization, load cells, and Stereoscopic Particle Image Velocimetry (SPIV). The oil flow visualizations were conducted on two-dimensional NACA 0015 and NACA 0009 airfoils, at angles of attack ranging from 14Â° to 20Â°, Reynolds numbers ranging from 1.70x105to 4.20x105, and aspect ratios of 4, 6.67, and 13.33. Load cell data was conducted on a NACA 0015 airfoil, at angles of attack ranging from 0Â° to 20Â°, Reynolds numbers ranging from 1.70x105 to 4.20x105, and an aspect ratio of 4. SPIV experiments were conducted on a on a stalled two-dimensional NACA-0015, pitched to 18Â° angle of attack, at four Reynolds numbers ranging from 1.70x105 to 4.20x105. Oil flow visualizations were used for both qualitatively identify the stall cells and to calculate the near-surface skin friction field. The results showed that, the angle of attack, Reynolds number, aspect ratio, and the airfoil shape and thickness had a pronounced effect on the formation of stall cells. In order to understand the formation of stall cells, the effect of adding a passive disturbance, a zig-zag tape that attached to the airfoil's surface, was explored. The results showed that under some conditions, adding either a two-dimensional or a localized disturbance could alter the shape of the separation and yield the formation of stall cells. In addition, wall-mounted load cells were used to measure the lift and drag on the airfoil, in the presence or absence of stall cells. Finally, the SPIV measurements correlated the flow field over the airfoil with the oil flow visualization on the surface.
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.
Airfoil shape for a turbine nozzle
Burdgick, Steven Sebastian (Schenectady, NY); Patik, Joseph Francis (Cohoes, NY); Itzel, Gary Michael (Simpsonville, SC)
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.
NASA Astrophysics Data System (ADS)
Hinz, Denis F.; Alighanbari, Hekmat; Breitsamter, Christian
2013-02-01
The unsteady low Reynolds number aerodynamics phenomena around flapping wings are addressed in several investigations. Elsewhere, airfoils at higher Mach numbers and Reynolds numbers have been treated quite comprehensively in the literature. It is duly noted that the influence of heat transfer phenomena on the aerodynamic performance of flapping wings configurations is not well studied. The objective of the present study is to investigate the effect of heat transfer upon the aerodynamic performance of a pitching and plunging NACA0012 airfoil in the low Reynolds number flow regime with particular emphasis upon the airfoil's lift and drag coefficients. The compressible Navier-Stokes equations are solved using a finite volume method. To consider the variation of fluid properties with temperature, the values of dynamic viscosity and thermal diffusivity are evaluated with Sutherland's formula and the Eucken model, respectively. Instantaneous and mean lift and drag coefficients are calculated for several temperature differences between the airfoil surface and freestream within the range 0-100 K. Simulations are performed for a prescribed airfoil motion schedule and flow parameters. It is learnt that the aerodynamic performance in terms of the lift CL and drag CD behavior is strongly dependent upon the heat transfer rate from the airfoil to the flow field. In the plunging case, the mean value of CD tends to increase, whereas the amplitude of CL tends to decrease with increasing temperature difference. In the pitching case, on the other hand, the mean value and the amplitude of both CD and CL decrease. A spectral analysis of CD and CL in the pitching case shows that the amplitudes of both CD and CL decrease with increasing surface temperature, whereas the harmonic frequencies are not affected.
Airfoil-Shaped Fluid Flow Tool for Use in Making Differential Measurements
NASA Technical Reports Server (NTRS)
England, John Dwight (Inventor); Kelley, Anthony R. (Inventor); Cronise, Raymond J. (Inventor)
2014-01-01
A fluid flow tool includes an airfoil structure and a support arm. The airfoil structure's high-pressure side and low-pressure side are positioned in a conduit by the support arm coupled to the conduit. The high-pressure and low-pressure sides substantially face opposing walls of the conduit. At least one measurement port is formed in the airfoil structure at each of its high-pressure side and low-pressure side. A first manifold, formed in the airfoil structure and in fluid communication with each measurement port so-formed at the high-pressure side, extends through the airfoil structure and support arm to terminate and be accessible at the exterior wall of the conduit. A second manifold, formed in the airfoil structure and in fluid communication with each measurement port so-formed at the low-pressure side, extends through the airfoil structure and support arm to terminate and be accessible at the exterior wall of the conduit.
Lift-off Processes with Photoresists
Yoo, S. J. Ben
designed for lift-off attain a reproducible undercut. Such an undercut helps to prevent the resist of the resin of common negative resists maintains the undercut even at very high coating temperatures, which helps to maintain the undercut dur- ing coating. For stable lift-off processes in the µm and sub
Extraction of Subimages by Lifting Wavelet Filters
Shigeru TAKANO; Koichi NIIJIMA
2000-01-01
SUMMARY This paper proposes a method for extracting subimages from a huge reference image by learning lifting wavelet filters. Lifting wavelet filters are biorthogonal wavelet filters con- taining free parameters developed by Sweldens. Our method is to learn such free parameters using some training subimages so as to vanish their high frequency components in the y- and x- directions. The
Propeller-Design Problems of High-Speed Airplanes, Special Report
NASA Technical Reports Server (NTRS)
Dickinson, H. B.
1941-01-01
It is shown that on the basis of existing high-speed airfoil data, propeller efficiencies appreciably in excess of 40% do not appear possible at speeds above 500 miles per hour at 20,000 feet. The assumption that present propeller-blade thicknesses cannot be reduced radically, is implied. Until the reliability and applicability of the airfoil data are established, this conclusion must not be regarded as infallible. Dive tests with airplanes equipped with thrust meters and torque meters are proposed to provide an urgently needed check. The design of high-speed propellers is dictated wholly by compressibility considerations. The blade width, thickness, and pitch distribution; also the airfoil sections, the lift coefficient, the propeller diameter, and rpm must all be adjusted if reasonable efficiencies are to be maintained at airplane speeds that are now being approached. Research is urgently needed on: 1) airfoils at subsonic, sonic, and supersonic speeds; 2) propellers at high forward speeds in wind tunnels; 3)propellers in free flight at high speeds; and 4) jet propulsion and related devices. The breakdown of propeller efficiency indicated by airfoil data, should serve as an incentive for accelerated research on jet propulsion. This device may extend the attainable speed of current airplanes to the neighborhood of 550 miles per hour at 20,000 feet.
NASA Technical Reports Server (NTRS)
Aoyagi, Kiyoshi; Hickey, David H.; deSavigny, Richard A.
1961-01-01
An investigation was conducted to determine the longitudinal characteristics during low-speed flight of a large-scale VTOL airplane model with a direct lifting fan enclosed in the fuselage. The model had a shoulder-mounted unswept wing of aspect ratio 5. The effect on longitudinal characteristics of fan operation, propulsion by means of deflecting the fan efflux, trailing-edge flap deflection, and horizontal-tail height were studied.
NASA Technical Reports Server (NTRS)
Ladson, Charles L.
1959-01-01
A two-dimensional wind-tunnel investigation of the pressure distributions over several NACA 16-series airfoils with thicknesses of 4, 6, 9, and 12 percent of the chord and design lift coefficients of 0, 0.2, 1 and 0.5 has been conducted in the Langley airfoil test apparatus at transonic Mach numbers from 0.7 to 1.25. The tests ranged in Reynolds number from 2.4 x 10(exp 6) to 2.8 x l0(exp 6) and in angle of attack from -10 to 12 deg. Chordwise pressure distributions and schlieren flow photographs are presented without analysis.
NASA Technical Reports Server (NTRS)
Papadakis, Michael; Hung, Kuohsing E.; Vu, Giao T.; Yeong, Hsiung Wei; Bidwell, Colin S.; Breer, Martin D.; Bencic, Timothy J.
2002-01-01
Validation of trajectory computer codes, for icing analysis, requires experimental water droplet impingement data for a wide range of aircraft geometries as well as flow and icing conditions. This report presents improved experimental and data reduction methods for obtaining water droplet impingement data and provides a comprehensive water droplet impingement database for a range of test geometries including an MS(1)-0317 airfoil, a GLC-305 airfoil, an NACA 65(sub 2)-415 airfoil, a commercial transport tail section, a 36-inch chord natural laminar flow NLF(1)-0414 airfoil, a 48-inch NLF(1)-0414 section with a 25 percent chord simple flap, a state-of-the-art three-element high lift system, a NACA 64A008 finite span swept business jet tail, a full-scale business jet horizontal tail section, a 25 percent-scale business jet empennage, and an S-duct turboprop engine inlet. The experimental results were obtained at the NASA Glenn Icing Research Tunnel (IRT) for spray clouds with median volumetric diameter (MVD) of 11, 11.5, 21, 92, and 94 microns and for a range of angles of attack. The majority of the impingement experiments were conducted at an air speed of 175 mph corresponding to a Reynolds number of approximately 1.6 million per foot. The maximum difference of repeated tests from the average ranged from 0.24 to 12 percent for most of the experimental results presented. This represents a significant improvement in test repeatability compared to previous experimental studies. The increase in test repeatability was attributed to improvements made to the experimental and data reduction methods. Computations performed with the LEWICE-2D and LEWICE-3D computer codes for all test configurations are presented in this report. For the test cases involving median volumetric diameters of 11 and 21 microns, the correlation between the analytical and experimental impingement efficiency distributions was good. For the median volumetric diameters of 92 and 94-micron cases, however, the analysis produced higher impingement efficiencies and larger impingement limits than the experiment. It is speculated that this discrepancy is due to droplet splashing and breakup experienced by large droplets during impingement.
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.
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.
Towards Dynamic Stall Control on a Rotorcraft Airfoil
NASA Astrophysics Data System (ADS)
Alexander, Hani
1997-11-01
The unsteady stall that occurs on the retreating side of a helicopter rotor produces large torsional stresses and large losses in the aerodynamic performance. Previous results have shown that the static stall can be delayed and the hysteresis eliminated by the simple presence of a 'leading-edge wire'. The presence of the wire, of which the airfoil chord to wire diameter measures 400, has been shown to locally reduce the pressure gradient near the upper surface of the leading edge. Experimental tests of the aerodynamic performance of two wings, the Wortmann 69-H-098 and 63-137, in a pitching sinusoidal motion, have displayed the common features of the dynamic stall process. These features include the existence of a 'super-lift' or lift overshoot, a very large lift-stall and a large hysterisis associated with the cycle. Preliminary control results using a small mechanical actuator has shown a reduction in the effect of the dynamic-stall-vortex. The dynamic stall vortex is the major contributing factor to the negative effects associated with dynamic stall. Currently, steady and unsteady pressures are being measured to analyze the transient nature of the dynamic-stall-vortex motion and to verify control techniques.
Not Available
1989-01-01
This book focuses on the four major methods of artificial lift: sucker-rod pumping, gas lift, electrical submersible pumping (ESP) and hydraulic pumping. Though more than 80% of artificially lifted wells worldwide are rod-pumped, the large majority of these wells are low-volume, stripper-type producers. For this reason, sucker-rod pumping papers comprise less than 40% of the 26 SPE papers selected.
Sarychev, Michael
2007-08-01
This note describes BLS power supplies lifting techniques and provides stress calculations for lifting plate and handles bolts. BLS power supply weight is about 120 Lbs, with the center of gravity shifted toward the right front side. A lifting plate is used to attach a power supply to a crane or a hoist. Stress calculations show that safety factors for lifting plate are 12.9 (vs. 5 required) for ultimate stress and 5.7 (vs. 3 required) for yield stress. Safety factor for shackle bolt thread shear load is 37, and safety factor for bolts that attach handles is 12.8.
Self-sustained shock oscillations on airfoils at transonic speeds
NASA Astrophysics Data System (ADS)
Lee, B. H. K.
2001-02-01
Self-sustained shock wave oscillations on airfoils at transonic flow conditions are associated with the phenomenon of buffeting. The physical mechanisms of the periodic shock motion are not yet fully understood even though experiments performed over fifty years ago have demonstrated the presence of oscillatory shock waves on the airfoil surfaces at high subsonic speeds. The unsteady pressure fluctuations generated by the low-frequency large-amplitude shock motions are highly undesirable from the structural integrity and aircraft maneuverability point of view. For modern supercritical wing design with thick profiles, the shock-induced fluctuations are particularly severe and methods to reduce the shock wave amplitudes to lower values or even to delay the oscillations to higher Mach numbers or incidence angles will result in expanding the buffet boundary of the airfoil. This review begins with a recapitulation of the classical work on shock-induced bubble separation and trailing edge separation of a turbulent boundary layer. The characteristics of the unsteady pressure fluctuations are used to classify the types of shock-boundary layer interaction. The various modes of shock wave motion for different flow conditions and airfoil configurations are described. The buffet boundaries obtained using the standard trailing edge pressure divergence technique and an alternative approach of measuring the divergence of normal fluctuating forces are compared to show the equivalence. The mechanisms of self-sustained shock oscillations are discussed for symmetrical circular-arc airfoils at zero incidence and for supercritical airfoils at high incidence angles with fully separated flows. The properties of disturbances in the wake are examined from linear stability analysis of two-dimensional compressible flows. The advances in high-speed computing make predictions of buffeting flows possible. Navier-Stokes solvers and approximate boundary layer-inviscid flow interaction methods are shown to give good correlation of frequencies and other unsteady flow characteristics with experiments. Finally, passive and active methods of shock oscillation control show promising results in delaying buffet onset to higher Mach numbers or incidence angles, thus enhancing the transonic performance of airfoils.
NASA Technical Reports Server (NTRS)
Harris, Thomas A
1939-01-01
An investigation has been made in the N.A.C.A. 7- by 10-foot wind tunnel of a large-chord N.A.C.A. 23012 airfoil with several arrangements of a 40-percent-chord slotted flap to determine the section aerodynamic characteristics of the airfoil as affected by slot shape, flap location, and flap deflection. The flap positions for maximum lift, the polar for arrangements considered favorable for take-off and climb, and the complete section aerodynamic characteristics for selected optimum arrangements were determined. A discussion is given of the relative merits of the various arrangements. A comparison is made of slotted flaps of different chords on the N.A.C.A. 23012 airfoil. The best 40-percent-chord slotted flap is only slightly superior to the 25-percent-chord slotted flap from considerations of maximum lift coefficient and low drag for take-off and initial climb.
NASA Technical Reports Server (NTRS)
Polhamus, Edward C.
1996-01-01
This paper presents a survey of the effects of Reynolds number on the low- speed lift characteristics of wings encountering separated flows at their leading and side edges, with emphasis on the region near the stall. The influence of leading-edge profile and Reynolds number on the stall characteristics of two- dimensional airfoils are reviewed first to provide a basis for evaluating three- dimensional effects associated with various wing planforms. This is followed by examples of the effects of Reynolds number and geometry on the lift characteristics near the stall for a series of three-dimensional wings typical of those suitable for high-speed aircraft and missiles. Included are examples of the effects of wing geometry on the onset and spanwise progression of turbulent reseparation near the leading edge and illustrations of the degree to which simplified theoretical approaches can be useful in defining the influence of the various geometric parameters. Also illustrated is the manner in which the Reynolds number and wing geometry parameters influence whether the turbulent reseparation near the leading edge results in a sudden loss of lift, as in the two-dimensional case, or the formation of a leading-edge vortex with Rs increase in lift followed by a gentle stall as in the highly swept wing case. Particular emphasis is placed on the strong influence of 'induced camber' on the development of turbulent reseparation. R is believed that the examples selected for this report may be useful in evaluating viscous flow solutions by the new computational methods based on the Navier-Stokes equations as well as defining fruitful research areas for the high-Reynolds-number wind tunnels.
Kevin Drost; Sourabh Apte
2010-01-01
Direct numerical simulations are performed to investigate the effect of a movable leading edge on the unsteady flow at high angles of attack over a flat, thin airfoil at Reynolds number of 14700 based on the chord length. The leading edge of the airfoil is hinged at one-third chord length allowing dynamic variations in the effective angle of attack through
The Investigation of the Airfoil for the Small Wind Turbine Based on the Seagull Airfoil
Limin Qiao; Sidong Wei; Rui Gu; Xiaolin Quan; Yingjun Yang
2011-01-01
For the blades of the small wind turbine working under the conditions of Low-Reynolds, the air viscosity has relatively great influence on them. The design of the airfoil influence the performance of the small wind turbine directly. This paper based on the airfoil of the seagull, studied the aerodynamic performance of the airfoil under the Low-Reynolds and the influences of
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.
Unsteady Flow Simulation of a Controlled Airfoil
Sol Keun Jee; Omar Lopez; Jonathan Muse; Anthony Calise; Robert Moser
2008-01-01
An airfoil moving with two degrees of freedom (pitching and plunging) is simulated with a closed-loop flow control system. The simulation of the unsteady airfoil is computed using a Delayed Detached Eddy Simulation (DDES), a hybrid non-zonal RANS-LES turbulence model based on DES. The control system controls the airfoil in two modes, first through direct application of forces and torques,
NASA Technical Reports Server (NTRS)
Letko, W; Denaci, H. G.; Freed, C
1943-01-01
Hinge-moment, lift, and pressure-distribution measurements were made in the two-dimensional test section of the NACA stability tunnel on a blunt-nose balance-type aileron on an NACA 66,2-216 airfoil at speeds up to 360 miles per hour corresponding to a Mach number of 0.475. The tests were made primarily to determine the effect of speed on the action of this type of aileron. The balance-nose radii of the aileron were varied from 0 to 0.02 of the airfoil chord and the gap width was varied from 0.0005 to 0.0107 of the airfoil chord. Tests were also made with the gap sealed.
NASA Technical Reports Server (NTRS)
Loftin, Laurence K, Jr; Bursnall, William J
1950-01-01
Results are presented of an investigation made to determine the two-dimensional lift and drag characteristics of nine NACA 6-series airfoil section at Reynolds numbers of 15.0 x 10sub6, 20.0 x 10sub6, and 25.0 x 10sub6. Also presented are data from NACA Technical Report 824 for the same airfoils at Reynolds numbers of 3.0 x 10sub6, 6.0 x 10sub6, and 9.0 x 10sub6. The airfoils selected represent sections having variations in the airfoil thickness, thickness form, and camber. The characteristics of an airfoil with a split flap were determined in one instance, as was the effect of surface roughness. Qualitative explanations in terms of flow behavior are advanced for the observed types of scale effect.
NSDL National Science Digital Library
Texas Instruments
2012-08-03
"Students will work in groups to solve a real-world problem presented by the book: How Do You Lift A Lion? Using a toy lion and a lever, students will discover how much work is needed to raise the toy lion. They will use proportions to determine the force needed to lift a real lion" from TI World Math.
ERIC Educational Resources Information Center
Silva, J.; Soares, A. A.
2010-01-01
The conventional explanation of aerodynamic lift based on Bernoulli's equation is one of the most common mistakes in presentations to school students and is found in children's science books. The fallacies in this explanation together with an alternative explanation for aerofoil lift have already been presented in an excellent article by Babinsky…
Gunter, L.W.
1992-08-11
A device is described for lifting catwalk grates comprising an elongated bent member with a handle at one end and a pair of notched braces and a hook at the opposite end that act in conjunction with each other to lock onto the grate and give mechanical advantage in lifting the grate. 10 figs.
Bobo
1987-01-01
A gas lift method for lifting fluid from a borehole to the surface is described comprising the steps of: mixing at the surface a pressurized injection gas with a pressurized injection liquid without any foaming substance to form a gas-liquid pressurized non-foam mixture at a predetermined pressure less than the pressure at the bottom of the borehole, introducing the pressurized
Alon Amit; Nathan Linial; Ji?í Matoušek; Eyal Rozenman
2001-01-01
We describe here a simple probabilistic model for graphs that are lifts of a fixed base graph G, i.e., those graphs from which there is a covering man onto G. Our aim is to investigate the properties of typical graphs in this class. In particular, we show that almost every lift of G is ?(G)-connected where ?(G) is the minimal
Numerical simulation of flow induced airfoil vibrations with large amplitudes
NASA Astrophysics Data System (ADS)
Svá?ek, P.; Feistauer, M.; Horá?ek, J.
2007-04-01
The subject of this paper is the numerical simulation of the interaction of two-dimensional incompressible viscous flow and a vibrating airfoil. A solid airfoil with two degrees of freedom, which can rotate around the elastic axis and oscillate in the vertical direction, is considered. The numerical simulation consists of the finite element solution of the Navier Stokes equations, coupled with the system of ordinary differential equations describing the airfoil motion. The high Reynolds numbers considered 10 10 require the application of a suitable stabilization of the finite element discretization. The method presented in this paper is based on the laminar model and the turbulence modelling is not applied here. The time-dependent computational domain and a moving grid are taken into account with the aid of the arbitrary Lagrangian Eulerian (ALE) formulation of the Navier Stokes equations. Special attention is paid to the time discretization and the solution of the nonlinear discrete problem on each time level is performed. As a result, a sufficiently accurate and robust method is developed, which is applied to the case of flow-induced airfoil vibrations with large amplitudes after the loss of aeroelastic stability. The computational results are compared with known aerodynamical data and with results of aeroelastic calculations obtained by NASTRAN code for a linear approximation.
A frequency domain numerical method for airfoil broadband self-noise prediction
NASA Astrophysics Data System (ADS)
Zhou, Qidou; Joseph, Phillip
2007-01-01
This paper describes a numerical approach, based in the frequency domain, for predicting the broadband self-noise radiation due to an airfoil situated in a smooth mean flow. Noise is generated by the interaction between the boundary layer turbulence on the airfoil surface and the airfoil trailing edge. Thin airfoil theory is used to deduce the unsteady blade loading. In this paper, the important difference with much of the previous work dealing with trailing edge noise is that the integration of the surface sources for computation of the radiated sound field is evaluated on the actual airfoil surface rather than in the mean-chord plane. The assumption of flat plate geometry in the calculation of radiation is therefore avoided. Moreover, the solution is valid in both near and far fields and reduces to the analytic solution due to Amiet when the airfoil collapses to a flat plate with large span, and the measurement point is taken to the far field. Predictions of the airfoil broadband self-noise radiation presented here are shown to be in reasonable agreement with the predictions obtained using the Brooks approach, which are based on a comprehensive database of experimental data. Also investigated in this paper is the effect on the broadband noise prediction of relaxing the 'frozen-gust' assumption, whereby the turbulence at each frequency comprises a continuous spectrum of streamwise wavenumber components. It is shown that making the frozen gust assumption yields an under-prediction of the noise spectrum by approximately 2dB compared with that obtained when this assumption is relaxed, with the largest occurring at high frequencies. This paper concludes with a comparison of the broadband noise directivity for a flat-plat, a NACA 0012 and a NACA 0024 airfoil at non-zero angle of attack. Differences of up to 20 dB are predicted, with the largest difference occurring at a radiation angle of zero degrees relative to the airfoil mean centre line.
NASA Technical Reports Server (NTRS)
Saltzman, Edwin J.; Wang, K. Charles; Iliff, Kenneth W.
2002-01-01
This report examines subsonic flight-measured lift and drag characteristics of seven lifting-body and wing-body reentry vehicle configurations with truncated bases. The seven vehicles are the full-scale M2-F1, M2-F2, HL-10, X-24A, X-24B, and X-15 vehicles and the Space Shuttle Enterprise. Subsonic flight lift and drag data of the various vehicles are assembled under aerodynamic performance parameters and presented in several analytical and graphical formats. These formats are intended to unify the data and allow a greater understanding than individually studying the vehicles allows. Lift-curve slope data are studied with respect to aspect ratio and related to generic wind-tunnel model data and to theory for low-aspect-ratio platforms. The definition of reference area is critical for understanding and comparing the lift data. The drag components studied include minimum drag coefficient, lift-related drag, maximum lift-to drag ratio, and, where available, base pressure coefficients. The influence of forebody drag on afterbody and base drag at low lift is shown to be related to Hoerner's compilation for body, airfoil, nacelle, and canopy drag. This feature may result in a reduced need of surface smoothness for vehicles with a large ratio of base area to wetted area. These analyses are intended to provide a useful analytical framework with which to compare and evaluate new vehicle configurations of the same generic family.
An approach to constrained aerodynamic design with application to airfoils
NASA Technical Reports Server (NTRS)
Campbell, Richard L.
1992-01-01
An approach was developed for incorporating flow and geometric constraints into the Direct Iterative Surface Curvature (DISC) design method. In this approach, an initial target pressure distribution is developed using a set of control points. The chordwise locations and pressure levels of these points are initially estimated either from empirical relationships and observed characteristics of pressure distributions for a given class of airfoils or by fitting the points to an existing pressure distribution. These values are then automatically adjusted during the design process to satisfy the flow and geometric constraints. The flow constraints currently available are lift, wave drag, pitching moment, pressure gradient, and local pressure levels. The geometric constraint options include maximum thickness, local thickness, leading-edge radius, and a 'glove' constraint involving inner and outer bounding surfaces. This design method was also extended to include the successive constraint release (SCR) approach to constrained minimization.
NASA Technical Reports Server (NTRS)
Knight, Montgomery; Harris, Thomas A
1931-01-01
This experimental investigation was conducted primarily for the purpose of obtaining a method of correcting to free air conditions the results of airfoil force tests in four open wind tunnel jets of different shapes. Tests were also made to determine whether the jet boundaries had any appreciable effect on the pitching moments of a complete airplane model. Satisfactory corrections for the effect of the boundaries of the various jets were obtained for all the airfoils tested, the span of the largest being 0.75 of the jet width. The corrections for angle of attack were, in general, larger than those for drag. The boundaries had no appreciable effect on the pitching moments of either the airfoils or the complete airplane model. Increasing turbulence appeared to increase the minimum drag and maximum lift and to decrease the pitching moment.
NASA Astrophysics Data System (ADS)
Kachele, Clayton
The accurate depiction of helicopter rotor performance relies on the use of comprehensive analysis codes to efficiently predict performance, vibrations, and aeroelastic properties. Currently, most analysis relies upon airfoil tables that provide aerodynamic characteristics from a combination of wind tunnel data, empirical data, and numerical analyses. Computational Fluid Dynamics (CFD) has the potential to accurately replace the current analysis methods with the use of advanced computers; presented is a new 2-D CFD applied to rotorcraft studies. The analysis performed utilized the SC 1095 airfoil because of the large amount of previous wind tunnel testing and prior CFD analysis. In this thesis it is shown that the Flowfield Dependent Variation (FDV) method of CFD is able to produce a strong correlation of experimental and theoretical results with greater efficiency when compared to other CFD codes. The FDV method is also able to accurately predict the wind tunnel results of the lift coefficient for a rotorcraft airfoil over a range of Mach numbers.
NASA Astrophysics Data System (ADS)
Liang, Yun; Zou, Zheng-Ping; Liu, Hou-Xing; Zhang, Wei-Hao
2015-04-01
Detailed experimental investigation was carried out to investigate the interaction of unsteady wakes with boundary layer in a high-lift low-pressure turbine. Extensive measurements about boundary layer character were conducted using hot-film and hot-wire methods. In-depth analysis of the effect of wake passing frequency on boundary layer transition was carried out. The strength of separation control and profile loss variation at two wake passing frequencies were also studied. The results show that wake-induced transition can be detected in the separating shear layer, and complex vortex structures are induced by the interaction between the negative jet of wake and separation bubble. The proportions of laminar, separation and turbulence friction loss in the total loss vary with wake passing frequency, which leads to the change in the total boundary layer loss. In particular, as the wake passing frequency changes, the laminar and turbulent friction loss show opposite trends, and this indicates that the best frequency can be achieved by balancing these two types of losses. For a given high-lift profile, an optimum wake passing frequency that will lead to the minimum loss exists.
NASA Technical Reports Server (NTRS)
London, R. J.; Watts, G. A.; Sissingh, G. J.
1973-01-01
An experimental investigation to determine the dynamic characteristics of a hingeless rotor operating at moderate to high lift was conducted on a small scale, 7.5-foot diameter, four-bladed hingeless rotor model in a 7 x 10-foot wind tunnel. The primary objective of this research program was the empirical determination of the rotor steady-state and frequency responses to swashplate and body excitations. Collective pitch was set from 0 to 20 degrees, with the setting at a particular advance ratio limited by the cyclic pitch available for hub moment trim. Advance ratio varied from 0.00 to 0.36 for blades with nondimensional first-flap frequencies at 1.15, 1.28 and 1.33 times the rotor rotation frequency. Several conditions were run with the rotor operating in the transition regime. Rotor response at high lift is shown to be generally nonlinear in this region. As a secondary objective an experimental investigation of the rotor response to 4/revolution swashplate excitations at advance ratios of 0.2 to 0.85 and at a nondimensional, first-flap modal frequency of 1.34 was also conducted, using the 7 x 10-foot wind tunnel. It is shown that 4/revolution swashplate inputs are a method for substantially reducing rotor-induced, shafttransmitted vibratory forces.
NASA Astrophysics Data System (ADS)
Reddy, K. S. V.; Sharma, D. M.; Poddar, K.
While conducting 2-D experiments for flow over airfoil in small wind tunnels, models are usually mounted by spanning the tunnel. However, in large wind tunnels, it is a usual practice to test airfoil models with end plates mounted at its both ends. Thus reduces the blockage, eliminates the tip effects, enables two-dimensional configuration for the flow over airfoil, apart of being cost effective and time saving. Main objective of the present investigation is to study the effect of different sized and shaped end plates on the surface pressure distribution for a given airfoil. Experiments were carried out on a single aerofoil GA(W)-2 at National Wind Tunnel Facility in IIT-Kanpur and surface pressures were measured at constant free stream speed of 30 m/s by varying the end-plates shapes at discrete angle of attack (-10° to 30°). Present study employed three different size end plates A, B, and C. Size of the endplate A is extended 20% chord length all along the airfoil cross sectional area. And end plate B is one chord length extended on airfoil except lower surface. In addition end plate B has wedge on the upper side. End plate C is spanning the tunnel height. These selected varied shaped end-plates had a significant affect on the aerodynamic characteristics along with a considerable variations in the flow phenomena involving flow separations, tip and blockage effects. Results indicates that endplate C gives the closest 2-D configuration (Figure a). Lift curve slope increases and stall angle decreases as the end plate size increases from A to C. It appears that for small size end plates (A and B), the stalling angle delayed to higher angles because of the interaction of the outer free stream flow with the flow on the suction side of the airfoil. And this interaction provides the energy to recover pressure and sustains to higher angles at the expense of the lift. And also stalling occurs abruptly. Effect of endplates changes the moment curve slope. Negative slope observed for all the three cases (Figure b).
Downwash and Wake Behind Plain and Flapped Airfoils
NASA Technical Reports Server (NTRS)
Silverstein, Abe; Katzoff, S; Bullivant, W Kenneth
1939-01-01
Extensive experimental measurements have been made of the downwash angles and the wake characteristics behind airfoils with and without flaps and the data have been analyzed and correlated with the theory. A detailed study was made of the errors involved in applying lifting-line theory, such as the effects of a finite wing chord, the rolling-up of the trailing vortex sheet, and the wake. The downwash angles, as computed from the theoretical span load distribution by means of the Biot-Savart equation, were found to be in satisfactory agreement with the experimental results. The rolling-up of the trailing vortex sheet may be neglected, but the vertical displacement of the vortex sheet requires consideration. By the use of a theoretical treatment indicated by Prandtl, it has been possible to generalize the available experimental results so the predictions can be made of the important wake parameters in terms of the distance behind the airfoil trailing edge and the profile-drag coefficient. The method of application of the theory to design and the satisfactory agreement between predicted and experimental results when applied to an airplane are demonstrated.
A Numerical Evaluation of Icing Effects on a Natural Laminar Flow Airfoil
NASA Technical Reports Server (NTRS)
Chung, James J.; Addy, Harold E., Jr.
2000-01-01
As a part of CFD code validation efforts within the Icing Branch of NASA Glenn Research Center, computations were performed for natural laminar flow (NLF) airfoil, NLF-0414. with 6 and 22.5 minute ice accretions. Both 3-D ice castings and 2-D machine-generated ice shapes were used in wind tunnel tests to study the effects of natural ice is well as simulated ice. They were mounted in the test section of the Low Turbulence Pressure Tunnel (LTPT) at NASA Langley that the 2-dimensionality of the flow can be maintained. Aerodynamic properties predicted by computations were compared to data obtained through the experiment by the authors at the LTPT. Computations were performed only in 2-D and in the case of 3-D ice, the digitized ice shape obtained at one spanwise location was used. The comparisons were mainly concentrated on the lift characteristics over Reynolds numbers ranging from 3 to 10 million and Mach numbers ranging from 0.12 to 0.29. WIND code computations indicated that the predicted stall angles were in agreement with experiment within one or two degrees. The maximum lift values obtained by computations were in good agreement with those of the experiment for the 6 minute ice shapes and the minute 3-D ice, but were somewhat lower in the case of the 22.5 minute 2-D ice. In general, the Reynolds number variation did not cause much change in the lift values while the variation of Mach number showed more change in the lift. The Spalart-Allmaras (S-A) turbulence model was the best performing model for the airfoil with the 22.5 minute ice and the Shear Stress Turbulence (SST) turbulence model was the best for the airfoil with the 6 minute ice and also for the clean airfoil. The pressure distribution on the surface of the iced airfoil showed good agreement for the 6 minute ice. However, relatively poor agreement of the pressure distribution on the upper surface aft of the leading edge horn for the 22.5 minute ice suggests that improvements are needed in the grid or turbulence 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 ? 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.
NASA Astrophysics Data System (ADS)
Valenzuela, Fernando; Avila, Ruben
2008-11-01
The two dimensional Navier-Stokes equations for a compressible fluid have been numerically solved to investigate the effect of the viscosity and the compressibility on the instability of the laminar air flow around a NACA0012 airfoil at zero incidence angle of attack. The numerical simulations are carried out at two Re numbers (5,000 and 10,000), and at the Ma number in the range 0.2-0.9. The non-dimensional compressible Navier-Stokes equations have been solved by using a high order Spectral/hp Element Method (SEM) that uses a Discontinuous Galerkin approach. The mesh consists of 2320. The computed values of the pressure coefficient Cp, the lift coefficient CL and the drag coefficient CD, have been compared with numerical simulations and experimental data available in the literature. The Mach number countours allow to identify the presence of shock waves in the transonic flow regime. We may conclude that the discontinuous Galerkin method is capable to predict the transition stages that the laminar flow undergoes as the Ma number increases (at fixed Re number).
Large eddy simulation with periodic forcing of low-frequency flow oscillation near airfoil stall.
NASA Astrophysics Data System (ADS)
Almutairi, Jaber; Alqadi, Ibraheem; Eljack, Elteyeb
2013-11-01
The effect of periodic forcing on the airfoil performance is investigated in the present study. A large eddy simulation with a flow control technique of periodic forcing is used to remove the low-frequency flow oscillation of the NACA-0012 airfoil at a Reynolds number of 130,000 and incidence of 11.5° and thus enhancing the performance of the airfoil. The periodic forcing is introduced into the laminar boundary layer just upstream of the natural flow separation to produce perturbations in the near wall region. The amplitude of the forcing is set to be 0.3% of the freestream velocity while several different frequencies based on the detected frequency of the vortex trailing edge revealed from the simulation of the natural low-frequency flow oscillation are used. It was found that periodic forcing removes the low-frequency flow oscillation, and as a consequence improve the performance of the airfoil. It has been seen that the large fluctuation of the lift and drag coefficients are entirely eliminated and the enlargement of the turbulent boundary layer that usually occurs after separation and the associated increase of the normal velocity component are reduced sharply when the periodic forcing is added.
Two-dimensional aerodynamic characteristics of the OLS/TAAT airfoil
NASA Technical Reports Server (NTRS)
Watts, Michael E.; Cross, Jeffrey L.; Noonan, Kevin W.
1988-01-01
Two flight tests have been conducted that obtained extension pressure data on a modified AH-1G rotor system. These two tests, the Operational Loads Survey (OLS) and the Tip Aerodynamics and Acoustics Test (TAAT) used the same rotor set. In the analysis of these data bases, accurate 2-D airfoil data is invaluable, for not only does it allow comparison studies between 2- and 3-D flow, but also provides accurate tables of the airfoil characteristics for use in comprehensive rotorcraft analysis codes. To provide this 2-D data base, a model of the OLS/TAAT airfoil was tested over a Reynolds number range from 3 x 10 to the 6th to 7 x 10 to the 7th and between Mach numbers of 0.34 to 0.88 in the NASA Langley Research Center's 6- by 28-Inch Transonic Tunnel. The 2-D airfoil data is presented as chordwise pressure coefficient plots, as well as lift, drag, and pitching moment coefficient plots and tables.
NASA Astrophysics Data System (ADS)
Feng, Li-Hao; Jukes, Timothy N.; Choi, Kwing-So; Wang, Jin-Jun
2012-06-01
Flow control study of a NACA 0012 airfoil with a Gurney flap was carried out in a wind tunnel, where it was demonstrated that a dielectric-barrier-discharge (DBD) plasma actuator attached to the flap could increase the lift further, but with a small drag penalty. Time-resolved PIV measurements of the near-wake region indicated that the plasma forcing shifted the wake downwards, reducing its recirculation length. Analysis of wake vortex dynamics suggested that the plasma actuator initially amplified the lower wake shear layer by adding momentum along the downstream surface of the Gurney flap. This enhanced mutual entrainment between the upper and lower wake vortices, leading to an increase in lift on the airfoil.
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.
Control of unsteady separated flow associated with the dynamic stall of airfoils
NASA Technical Reports Server (NTRS)
Wilder, M. C.
1994-01-01
A unique active flow-control device is proposed for the control of unsteady separated flow associated with the dynamic stall of airfoils. The device is an adaptive-geometry leading-edge which will allow controlled, dynamic modification of the leading-edge profile of an airfoil while the airfoil is executing an angle-of-attack pitch-up maneuver. A carbon-fiber composite skin has been bench tested, and a wind tunnel model is under construction. A baseline parameter study of compressible dynamic stall was performed for flow over an NACA 0012 airfoil. Parameters included Mach number, pitch rate, pitch history, and boundary layer tripping. Dynamic stall data were recorded via point-diffraction interferometry and the interferograms were analyzed with in-house developed image processing software. A new high-speed phase-locked photographic image recording system was developed for real-time documentation of dynamic stall.
Lampton -1 Morphing Airfoils with Four Morphing
Valasek, John
Lampton - 1 Morphing Airfoils with Four Morphing Parameters Amanda Lampton, Adam Niksch , John Conference and Exhibit 18-21 August 2008 Honolulu, HI #12;Lampton - 2 Outline Morphing Air Vehicles Q-Learning Morphing Airfoil Results Conclusions Challenges #12;Lampton - 3 Morphing Aircraft Mutli-role platform
Rapid Design of Subcritical Airfoils Prabir Daripa
Daripa, Prabir
Rapid Design of Subcritical Airfoils Prabir Daripa Department of Mathematics Texas A&M University into Real (( = ei )) = -(), [0, 2], (7) Imag (( = 0)) = . (8) The solution of (6) subject to the above boundary data provides (). The profile Lz of the airfoil, which is the image of the unit circle = ei
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.
Drag and lift reduction of a 3D bluff-body using active vortex generators
NASA Astrophysics Data System (ADS)
Aider, Jean-Luc; Beaudoin, Jean-François; Wesfreid, José Eduardo
2010-05-01
In this study, a passive flow control experiment on a 3D bluff-body using vortex generators (VGs) is presented. The bluff-body is a modified Ahmed body (Ahmed in J Fluids Eng 105:429-434 1983) with a curved rear part, instead of a slanted one, so that the location of the flow separation is no longer forced by the geometry. The influence of a line of non-conventional trapezoïdal VGs on the aerodynamic forces (drag and lift) induced on the bluff-body is investigated. The high sensitivity to many geometric (angle between the trapezoïdal element and the wall, spanwise spacing between the VGs, longitudinal location on the curved surface) and physical (freestream velocity) parameters is clearly demonstrated. The maximum drag reduction is -12%, while the maximum global lift reduction can reach more than -60%, with a strong dependency on the freestream velocity. For some configurations, the lift on the rear axle of the model can be inverted (-104%). It is also shown that the VGs are still efficient even downstream of the natural separation line. Finally, a dynamic parameter is chosen and a new set-up with motorized vortex generators is proposed. Thanks to this active device. The optimal configurations depending on two parameters are found more easily, and a significant drag and lift reduction (up to -14% drag reduction) can be reached for different freestream velocities. These results are then analyzed through wall pressure and velocity measurements in the near-wake of the bluff-body with and without control. It appears that the largest drag and lift reduction is clearly associated to a strong increase of the size of the recirculation bubble over the rear slant. Investigation of the velocity field in a cross-section downstream the model reveals that, in the same time, the intensity of the longitudinal trailing vortices is strongly reduced, suggesting that the drag reduction is due to the breakdown of the balance between the separation bubble and the longitudinal vortices. It demonstrates that for low aspect ratio 3D bluff-bodies, like road vehicles, the flow control strategy is much different from the one used on airfoils: an early separation of the boundary layer can lead to a significant drag reduction if the circulation of the trailing vortices is reduced.
Wind tunnel results of the low-speed NLF(1)-0414F airfoil
NASA Technical Reports Server (NTRS)
Murri, Daniel G.; Mcghee, Robert J.; Jordan, Frank L., Jr.; Davis, Patrick J.; Viken, Jeffrey K.
1987-01-01
The large performance gains predicted for the Natural Laminar Flow (NLF)(1)-0414F airfoil were demonstrated in two-dimensional airfoil tests and in wind tunnel tests conducted with a full scale modified Cessna 210. The performance gains result from maintaining extensive areas of natural laminar flow, and were verified by flight tests conducted with the modified Cessna. The lift, stability, and control characteristics of the Cessna were found to be essentially unchanged when boundary layer transition was fixed near the wing leading edge. These characteristics are very desirable from a safety and certification view where premature boundary layer transition (due to insect contamination, etc.) must be considered. The leading edge modifications were found to enhance the roll damping of the Cessna at the stall, and were therefore considered effective in improving the stall/departure resistance. Also, the modifications were found to be responsible for only minor performance penalties.
Experimental Investigation of Wind-Tunnel Interference on the Downwash Behind an Airfoil
NASA Technical Reports Server (NTRS)
Silverstein, Abe; Katzoff, S
1937-01-01
The interference of the wind-tunnel boundaries on the downwash behind an airfoil has been experimentally investigated and the results have been compared with the available theoretical results for open-throat wind tunnels. As in previous studies, the simplified theoretical treatment that assumes the test section to be an infinite free jet has been shown to be satisfactory at the lifting line. The experimental results, however, show that this assumption may lead to erroneous conclusions regarding the corrections to be applied to the downwash in the region behind the airfoil where the tail surfaces are normally located. The results of a theory based on the more accurate concept of the open-jet wind tunnel as a finite length of free jet provided with a closed exit passage are in good qualitative agreement with the experimental results.
Determination of the pressure drag of airfoils by integration of surface pressures
NASA Technical Reports Server (NTRS)
Phillips, William H.
1990-01-01
A study was conducted of the causes of pressure drag of subsonic airfoils. In a previous paper by the author, the pressure drag is obtained by calculating the total drag from the momentum defect in the boundary layer at the trailing edge and subtracting the friction drag obtained from integration of surface friction along the chord. Herein, the pressure drag is obtained by integrating the streamwise components of surface pressure around the airfoil. Studies were made to verify the accuracy of the integration procedure. The values of pressure drag were much smaller than those obtained by the previous method. This lack of agreement is attributed to the difficulty of calculating boundary layer conditions in the vicinity of the trailing edge and to the extreme sensitivity of the circulation and lift to the trailing edge conditions. The results of these studies are compared with those of previous investigations.
Advanced underwater lift device
NASA Technical Reports Server (NTRS)
Flanagan, David T.; Hopkins, Robert C.
1993-01-01
Flexible underwater lift devices ('lift bags') are used in underwater operations to provide buoyancy to submerged objects. Commercially available designs are heavy, bulky, and awkward to handle, and thus are limited in size and useful lifting capacity. An underwater lift device having less than 20 percent of the bulk and less than 10 percent of the weight of commercially available models was developed. The design features a dual membrane envelope, a nearly homogeneous envelope membrane stress distribution, and a minimum surface-to-volume ratio. A proof-of-concept model of 50 kg capacity was built and tested. Originally designed to provide buoyancy to mock-ups submerged in NASA's weightlessness simulators, the device may have application to water-landed spacecraft which must deploy flotation upon impact, and where launch weight and volume penalties are significant. The device may also be useful for the automated recovery of ocean floor probes or in marine salvage applications.
ERIC Educational Resources Information Center
Weltner, Klaus
1990-01-01
Describes some experiments showing both qualitatively and quantitatively that aerodynamic lift is a reaction force. Demonstrates reaction forces caused by the acceleration of an airstream and the deflection of an airstream. Provides pictures of demonstration apparatus and mathematical expressions. (YP)
FREIGHT CONTAINER LIFTING STANDARD
POWERS DJ; SCOTT MA; MACKEY TC
2010-01-13
This standard details the correct methods of lifting and handling Series 1 freight containers following ISO-3874 and ISO-1496. The changes within RPP-40736 will allow better reading comprehension, as well as correcting editorial errors.
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.
Mechanism of Water Droplet Breakup Near the Leading Edge of an Airfoil
NASA Technical Reports Server (NTRS)
Vargas, Mario; Sor, Suthyvann; Magarino, Adelaida, Garcia
2012-01-01
This work presents results of an experimental study on droplet deformation and breakup near the leading edge of an airfoil. The experiment was conducted in the rotating rig test cell at the Instituto Nacional de Tecnica Aeroespacial (INTA) in Madrid, Spain. The airfoil model was placed at the end of the rotating arm and a monosize droplet generator produced droplets that fell from above, perpendicular to the path of the airfoil. The interaction between the droplets and the airfoil was captured with high speed imaging and allowed observation of droplet deformation and breakup as the droplet approached the airfoil near the stagnation line. Image processing software was used to measure the position of the droplet centroid, equivalent diameter, perimeter, area, and the major and minor axes of an ellipse superimposed over the deforming droplet. The horizontal and vertical displacement of each droplet against time was also measured, and the velocity, acceleration, Weber number, Bond number, Reynolds number, and the drag coefficients were calculated along the path of the droplet to the beginning of breakup. Droplet deformation is defined and studied against main parameters. The high speed imaging allowed observation of the actual mechanism of breakup and identification of the sequence of configurations from the initiation of the breakup to the disintegration of the droplet. Results and comparisons are presented for droplets of diameters in the range of 500 to 1800 microns, and airfoil velocities of 70 and 90 m/sec.
NASA Technical Reports Server (NTRS)
Prandtl , L
1920-01-01
The general basis of the theory of lifting surfaces is discussed. The problem of the flow of a fluid about a lifting surface of infinite span is examined in terms of the existence of vortexes in the current. A general theory of permanent flow is discussed. Formulas for determining the influence of aspect ratio that may be applied to all wings, whatever their plane form, are given.
BIOMECHANICAL SIMULATION OF HUMAN LIFTING
COLOBERT B. MULTON
The aim of this paper is to simulate several levels of lifting strategies from parameters depending on the subject's centre of mass movements. Usually, symmetrical lifting strategies were categorized in two major solutions (Chaffin and Andersson, 1991): the squat lift that mainly involves a knee flexion and the back lift that mainly involves hip flexions. In the literature two main
NASA Technical Reports Server (NTRS)
Quigley, H. C.; Franklin, J. A.
1977-01-01
The paper gives an overview of the technology related to lift/cruise fan VTOL aircraft, covering propulsion systems, thrust deflection, flight dynamics, controls, displays, aerodynamics, and configurations. Piloting problems are discussed, and the need for integration of power management and thrust-vector controls is pointed out. Major components for a high-bypass-ratio lift/cruise fan propulsion system for VTOL aircraft have been tested.
An airfoil pitch apparatus-modeling and control design
NASA Technical Reports Server (NTRS)
Andrews, Daniel R.
1989-01-01
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.
Drag Coefficient of Water Droplets Approaching the Leading Edge of an Airfoil
NASA Technical Reports Server (NTRS)
Vargas, Mario; Sor, Suthyvann; Magarino, Adelaida Garcia
2013-01-01
This work presents results of an experimental study on droplet deformation and breakup near the leading edge of an airfoil. The experiment was conducted in the rotating rig test cell at the Instituto Nacional de Tecnica Aeroespacial (INTA) in Madrid, Spain. An airfoil model was placed at the end of the rotating arm and a monosize droplet generator produced droplets that fell from above, perpendicular to the path of the airfoil. The interaction between the droplets and the airfoil was captured with high speed imaging and allowed observation of droplet deformation and breakup as the droplet approached the airfoil near the stagnation line. Image processing software was used to measure the position of the droplet centroid, equivalent diameter, perimeter, area, and the major and minor axes of an ellipse superimposed over the deforming droplet. The horizontal and vertical displacement of each droplet against time was also measured, and the velocity, acceleration, Weber number, Bond number, Reynolds number, and the drag coefficients were calculated along the path of the droplet to the beginning of breakup. Results are presented and discussed for drag coefficients of droplets with diameters in the range of 300 to 1800 micrometers, and airfoil velocities of 50, 70 and 90 meters/second. The effect of droplet oscillation on the drag coefficient is discussed.
Compressibility effects on dynamic stall of airfoils undergoing rapid transient pitching motion
NASA Technical Reports Server (NTRS)
Chandrasekhara, M. S.; Platzer, M. F.
1992-01-01
The research was carried out in the Compressible Dynamic Stall Facility, CDSF, at the Fluid Mechanics Laboratory (FML) of NASA Ames Research Center. The facility can produce realistic nondimensional pitch rates experienced by fighter aircraft, which on model scale could be as high as 3600/sec. Nonintrusive optical techniques were used for the measurements. The highlight of the effort was the development of a new real time interferometry method known as Point Diffraction Interferometry - PDI, for use in unsteady separated flows. This can yield instantaneous flow density information (and hence pressure distributions in isentropic flows) over the airfoil. A key finding is that the dynamic stall vortex forms just as the airfoil leading edge separation bubble opens-up. A major result is the observation and quantification of multiple shocks over the airfoil near the leading edge. A quantitative analysis of the PDI images shows that pitching airfoils produce larger suction peaks than steady airfoils at the same Mach number prior to stall. The peak suction level reached just before stall develops is the same at all unsteady rates and decreases with increase in Mach number. The suction is lost once the dynamic stall vortex or vortical structure begins to convect. Based on the knowledge gained from this preliminary analysis of the data, efforts to control dynamic stall were initiated. The focus of this work was to arrive at a dynamically changing leading edge shape that produces only 'acceptable' airfoil pressure distributions over a large angle of attack range.
NASA Technical Reports Server (NTRS)
Jorgensen, L. H.
1977-01-01
An engineering-type method is presented for computing normal-force and pitching-moment coefficients for slender bodies of circular and noncircular cross section alone and with lifting surfaces. In this method, a semi-empirical term representing viscous-separation crossflow is added to a term representing potential-theory crossflow. For many bodies of revolution, computed aerodynamic characteristics are shown to agree with measured results for investigated free-stream Mach numbers from 0.6 to 2.9. The angles of attack extend from 0 deg to 180 deg for M = 2.9 from 0 deg to 60 deg for M = 0.6 to 2.0. For several bodies of elliptic cross section, measured results are also predicted reasonably well over the investigated Mach number range from 0.6 to 2.0 and at angles of attack from 0 deg to 60 deg. As for the bodies of revolution, the predictions are best for supersonic Mach numbers. For body-wing and body-wing-tail configurations with wings of aspect ratios 3 and 4, measured normal-force coefficients and centers are predicted reasonably well at the upper test Mach number of 2.0. Vapor-screen and oil-flow pictures are shown for many body, body-wing and body-wing-tail configurations. When spearation and vortex patterns are asymmetric, undesirable side forces are measured for the models even at zero sideslip angle. Generally, the side-force coefficients decrease or vanish with the following: increase in Mach number, decrease in nose fineness ratio, change from sharp to blunt nose, and flattening of body cross section (particularly the body nose).
NASA Technical Reports Server (NTRS)
Jorgensen, L. H.
1976-01-01
An engineering-type method is presented for computing normal-force and pitching-moment coefficients for slender bodies of circular and noncircular cross section alone and with lifting surfaces. In this method, a semi-empirical term representing viscous-separation crossflow is added to a term representing potential-theory crossflow. For many bodies of revolution, computed aerodynamic characteristics are shown to agree with measured results for investigated free-stream Mach numbers from 0.6 to 2.9. For several bodies of elliptic cross section, measured results are also predicted reasonably well over the investigated Mach number range from 0.6 to 2.0 and at angles of attack from 0 to 60 deg. As for the bodies of revolution, the predictions are best for supersonic Mach numbers. For body-wing and body-wing-tail configurations with wings of aspect ratios 3 and 4, measured normal-force coefficients and centers are predicted reasonably well at the upper test Mach number of 2.0. However, with a decrease in Mach number to 0.6, the agreement for C sub N rapidly deteriorates, although the normal-force centers remain in close agreement. Vapor-screen and oil-flow pictures are shown for many body, body-wing, and body-wing-tail configurations. When separation and vortex patterns are asymmetric, undesirable side forces are measured for the models even at zero sideslip angle. Generally, the side-force coefficients decrease or vanish with the following: increase in Mach number, decrease in nose fineness ratio, change from sharp to blunt nose, and flattening of body cross section (particularly the body nose).
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.
Nonlinear aeroelastic behavior of compliant airfoils
NASA Astrophysics Data System (ADS)
Thwapiah, G.; Campanile, L. F.
2010-03-01
Since the beginning of aviation and up to the present time, airfoils have always been built as rigid structures. They are designed to fly under their divergence speed in order to avoid static aeroelastic instabilities and the resulting large deformations, which are not compatible with the typically low compliance of such airfoils. In recent years, research on airfoil morphing has generated interest in innovative ideas like the use of compliant systems, i.e. systems built to allow for large deformations without failure, in airfoil construction. Such systems can operate in the neighborhood of divergence and take advantage of large aeroelastic servo-effects. This, in turn, allows compact, advanced actuators to control the airfoil's deformation and loads, and hence complement or even replace conventional flaps. In order to analyze and design such compliant, active aeroelastic structures a nonlinear approach to static aeroelasticity is needed, which takes into account the effect of large deformations on aerodynamics and structure. Such an analytical approach is presented in this paper and applied to a compliant passive airfoil as the preliminary step in the realization of a piezoelectrically driven, active aeroelastic airfoil. Wind tunnel test results are also presented and compared with the analytic prediction. The good agreement and the observed behavior in the wind tunnel give confidence in the potential of this innovative idea.
Single Airfoil Gust Response Problem
NASA Technical Reports Server (NTRS)
Scott, James R.
2000-01-01
An unsteady aerodynamic code, called GUST3D, has been developed to solve the single, non-constant coefficient, inhomogeneous convective wave equation for flows with periodic vortical disturbance. The code uses a frequency-domain approach with second-order central differences and a Sommerfeld radiation condition in the far field. It has been extensively validated on model problems with analytical solutions. GUST3D requires as input certain mean flow quantities which are calculated separately by a potential flow solver. This solver calculates the mean flow using a Gothert's Rule approximation. On the airfoil surface, it uses the solution calculated by the potential code FL036. A figure shows the mean pressure along the airfoil surface. To calculate the unsteady pressure, GUST3D was run oil systematically refined grids to obtain a converged solution at each frequency. It was found that 24 points per wavelength was sufficient for convergence. The location of the outer grid boundary was also varied to check for sensitivity to the far-field boundary condition.
NASA Technical Reports Server (NTRS)
Sivells, James C; Spooner, Stanley H
1949-01-01
Report presents the results of an investigation conducted in the Langley 19-foot pressure tunnel to determine the maximum lift and stalling characteristics of two thin wings equipped with several types of flaps. Split, single slotted, and double slotted flaps were tested on one wing which had NACA 65-210 airfoil sections and split and double slotted flaps were tested on the other, which had NACA 64-210 airfoil sections. Both wings were zero sweep, an aspect ratio of 9, and a taper ratio of 0.4.
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.
Numerical studies of unsteady transonic flow over an oscillating airfoil
NASA Astrophysics Data System (ADS)
Chyu, W. J.; Davis, S. S.
1984-10-01
A finite-difference solution to the Navier-Stokes equations combined with a time-varying grid-generation technique was used to compute unsteady transonic flow over an oscillating airfoil. These computations were compared with experimental data (obtained at Ames Research Center) which form part of the AGARD standard configuration for aeroelastic analysis. A variety of approximations to the full Navier-Stokes equations was used to determine the effect of frequency, shock-wave motion, flow separation, and airfoil geometry on unsteady pressures and overall air loads. Good agreement is shown between experiment and theory with the limiting factor being the lack of a reliable turbulence model for high-Reynolds-number, unsteady transonic flows.
Numerical simulation of viscous transonic flows over an airfoil
NASA Astrophysics Data System (ADS)
Velichko, S. A.; Lifshitz, Yu. B.
1995-05-01
High Reynolds number viscous transonic flow is described based on an interaction of the potential outer flow with the boundary layer and wake. Following the procedure of Lighthill (1958), the solutions in these domains are matched to each other through boundary conditions. The solution to the complete problem is obtained iteratively through successive computations of the flows in the outer and inner domains. Both old and new algorithms are used for the iteration process and subsequent problem solution. Results are given for all the airfoils from the Experimental Data Base for Computer Program Assessment (AGARD-AR-138, 1979). A comparison of these results with experimental data shows the degree of agreement between these unbounded airfoil flow simulations and real transonic flow over the central part of a straight wing.
Transonic flow about a thick circular-arc airfoil
NASA Technical Reports Server (NTRS)
Mcdevitt, J. B.; Levy, L. L., Jr.; Deiwert, G. S.
1975-01-01
An experimental and theoretical study of transonic flow over a thick airfoil, prompted by a need for adequately documented experiments that could provide rigorous verification of viscous flow simulation computer codes, is reported. Special attention is given to the shock-induced separation phenomenon in the turbulent regime. Measurements presented include surface pressures, streamline and flow separation patterns, and shadowgraphs. For a limited range of free-stream Mach numbers the airfoil flow field is found to be unsteady. Dynamic pressure measurements and high-speed shadowgraph movies were taken to investigate this phenomenon. Comparisons of experimentally determined and numerically simulated steady flows using a new viscous-turbulent code are also included. The comparisons show the importance of including an accurate turbulence model. When the shock-boundary layer interaction is weak the turbulence model employed appears adequate, but when the interaction is strong, and extensive regions of separation are present, the model is inadequate and needs further development.
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.
Patrocinio, Lucas G; Patrocinio, Tomas G; Patrocinio, Jose A
2013-06-01
The subperiosteal midface-lift has benefited from significant technological advances in medicine. The endoscope now allows extensive subperiosteal undermining of facial soft tissue through minimal access incisions. Improved understanding of facial anatomy and the facial aging process now allow repositioning and remodeling of the soft tissue envelope with excellent aesthetic results. The subperiosteal midface-lift by a temporal approach is a procedure designed to rejuvenate the middle third of the face. After subperiosteal detachment, the soft tissues of the cheek, jowls, lateral canthus, and inferior orbital rim can be lifted to reestablish their youthful relationship with the underlying skeleton. It is a technique that produces satisfactory cosmetic results in most cases, causing malar augmentation, nasolabial fold improvement, and mild jowl improvement. We describe our preferred technique in detail and discuss its indications, complications, advantages, and limitations. PMID:23761124
Experimental and computational study of airfoil load alteration using oscillating fence actuator
NASA Astrophysics Data System (ADS)
Saini, Manjinder
2008-11-01
Experimental and computational studies have been conducted for characterizing the effect of a fence actuator on stationary and oscillating airfoils and assessing the authority of the fence actuator for altering the aerodynamic loading of an airfoil to suppress flutter. In particular, an oscillating fence on a NACA-23012 airfoil has been examined using Particle Image Velocimetry (PIV), time-resolved pressure measurements, and numerical simulations of a pitching airfoil. Experiments over stationary and oscillating airfoils show that the fence frequency strongly affects the evolution of the vortical structures generated by the fence and that higher actuation frequencies are more effective in producing higher suction peaks. For an oscillating airfoil, variation of the mean angle of attack strongly affects the baseline pressure distribution as expected. The disturbances, however, remain largely unaffected by the variations in mean angle of attack. It was also observed that the adverse pressure gradient at higher angles of attack caused a reduction in dissipation of the disturbances. Integrated lift and moment are used to quantify the effectiveness of the fence actuator, and the sufficiently large changes produced in these quantities show the potential of this device for altering the aerodynamic loading of a wing and suppressing flutter. Results from the computational study indicate that the fence is capable of producing significant disturbances that diminish airfoil oscillations, more so when the actuator is located close to the trailing edge. Implementation of an appropriate control scheme can be used to enhance the effects of these actuators. Pressure Sensitive Paint experiments conducted for better understanding the production and evolution of the disturbances revealed complex surface pressure behavior. The results indicate that the actuator disturbances develop three-dimensionality soon after the fence penetrates the flow because of the finite fence length effects. Furthermore, the structure was found to evolve faster for slower fence frequencies due to reduction in time scale imposed by the oscillating fence. These results also suggest that the use of a two-dimensional disturbance in the numerical model may have caused over-estimation of the effect of the actuator.
NASA Astrophysics Data System (ADS)
Roger, Michel; Schram, Christophe; Moreau, Stéphane
2014-01-01
A linear analytical model is developed for the chopping of a cylindrical vortex by a flat-plate airfoil, with or without a span-end effect. The major interest is the contribution of the tip-vortex produced by an upstream rotating blade in the rotor-rotor interaction noise mechanism of counter-rotating open rotors. Therefore the interaction is primarily addressed in an annular strip of limited spanwise extent bounding the impinged blade segment, and the unwrapped strip is described in Cartesian coordinates. The study also addresses the interaction of a propeller wake with a downstream wing or empennage. Cylindrical vortices are considered, for which the velocity field is expanded in two-dimensional gusts in the reference frame of the airfoil. For each gust the response of the airfoil is derived, first ignoring the effect of the span end, assimilating the airfoil to a rigid flat plate, with or without sweep. The corresponding unsteady lift acts as a distribution of acoustic dipoles, and the radiated sound is obtained from a radiation integral over the actual extent of the airfoil. In the case of tip-vortex interaction noise in CRORs the acoustic signature is determined for vortex trajectories passing beyond, exactly at and below the tip radius of the impinged blade segment, in a reference frame attached to the segment. In a second step the same problem is readdressed accounting for the effect of span end on the aerodynamic response of a blade tip. This is achieved through a composite two-directional Schwarzschild's technique. The modifications of the distributed unsteady lift and of the radiated sound are discussed. The chained source and radiation models provide physical insight into the mechanism of vortex chopping by a blade tip in free field. They allow assessing the acoustic benefit of clipping the rear rotor in a counter-rotating open-rotor architecture.
Flow past two in-tandem airfoils undergoing sinusoidal oscillations
NASA Astrophysics Data System (ADS)
Lee, T.
2011-12-01
The interaction of the wake, generated behind an upstream oscillating NACA 0012 airfoil, with the downstream NACA 0012 airfoil, oscillated at the same conditions but with ? = 0° and 180° different phases (relative to the upstream airfoil), was investigated by particle image velocimetry and surface pressure measurements. The results show that the axial spacing and phase difference determined the strength of the undesirable interference effects and, subsequently, the behavior of the dynamic-load loops of the downstream airfoil. The boundary-layer events on the downstream airfoil were persistently different from those observed on the baseline oscillating airfoil. The downwash induced by the upstream airfoil disrupted leading-edge vortex (LEV) formation on the downstream airfoil. The absence of LEV-induced transient effects also led to a significantly lowered aerodynamic loading and C l -hysteresis compared to the baseline airfoil. The aerodynamic performance of the ? = 180° case, however, outperformed that of the ? = 0° case.
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.
The calculation of flow over iced airfoils
NASA Technical Reports Server (NTRS)
Cebeci, Tuncer
1988-01-01
Progress toward the development of a method for predicting the flowfield of an iced airfoil is described and shown to offer the prospect of a priori calculations of the effects of ice accretion and roughness on airfoil performance. The approach is based on interaction of inviscid flow solutions obtained by a panel method and improved upon by a finite-difference boundary-layer method which, operating in an inverse mode, incorporates viscous effects including those associated with separated flows. Results are presented for smooth, rough and iced airfoils as a function of angle of attack. Those for smooth and rough airfoils confirm the accuracy of the method and its applicability to surfaces with roughness similar to that associated with insect deposition and some forms of ice. Two procedures have been developed to deal with large ice accretion and their performance is examined and shown to be appropriate to the engineering requirements.
Probabilistic aerothermal design of compressor airfoils
Garzón, Víctor E., 1972-
2003-01-01
Despite the generally accepted notion that geometric variability is undesirable in turbomachinery airfoils, little is known in detail about its impact on aerothermal compressor performance. In this work, statistical and ...
Conceptual design of a lift fan plus lift/cruise fighter aircraft
NASA Technical Reports Server (NTRS)
Hill, G. C.; Waters, M. H.
1974-01-01
Results of a design synthesis and mission analysis of a supersonic VTOL fighter aircraft are presented. Propulsive lift is provided by a single turbotip-driven lift fan and deflected thrust from a high performance turbofan cruise engine fitted with an afterburner for supersonic flight. The inlet and thrust diverter in the main engine tail-pipe are seen to be the principal design problems. V/STOL and supersonic design tradeoffs are addressed in lift fan sizing and placement, reaction and aerodynamic control sizing, fuselage volume requirements, and area ruling. Range and turn rate are used as figures of merit.-
Numerical simulations of iced airfoils and wings
Jianping Pan
2004-01-01
A numerical study was conducted to understand the effects of simulated ridge and leading-edge ice shapes on the aerodynamic performance of airfoils and wings. In the first part of this study, a range of Reynolds numbers and Mach numbers, as well as ice-shape sizes and ice-shape locations were examined for various airfoils with the Reynolds-Averaged Navier-Stokes approach. Comparisons between simulation
TAIR: A transonic airfoil analysis computer code
NASA Technical Reports Server (NTRS)
Dougherty, F. C.; Holst, T. L.; Grundy, K. L.; Thomas, S. D.
1981-01-01
The operation of the TAIR (Transonic AIRfoil) computer code, which uses a fast, fully implicit algorithm to solve the conservative full-potential equation for transonic flow fields about arbitrary airfoils, is described on two levels of sophistication: simplified operation and detailed operation. The program organization and theory are elaborated to simplify modification of TAIR for new applications. Examples with input and output are given for a wide range of cases, including incompressible, subcritical compressible, and transonic calculations.
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.
NSDL National Science Digital Library
Cislunar Aerospace
1997-01-01
This activity introduces students to the aerodynamic basics of lift and drag. The materials needed are pieces of cardboard 20 x 30 inches and other size cardboard pieces. This would be a suitable activity for small groups. Copyright 2005 International Technology Education Association
Richard Ellenbogen; Anthony Youn; Dan Yamini; Steven Svehlak
2004-01-01
BackgroundMany face lift techniques that manipulate the superficial musculoaponeurotic system (SMAS) in order to correct a perceived descent of deeper facial structures have been presented. However, such procedures can result in insufficient volume restoration, which may be correctable only by actual fat replacement.
Bruce F. Connell; Santa Ana
2002-01-01
The author describes how to use the deep layer support technique for male face lift to achieve natural-looking results and virtually undetectable scars. He discusses incisions and neck contouring in detail and provides pointers on eliminating crow's-feet. (Aesthetic Surg J 2002;22:385-396.)
NASA Technical Reports Server (NTRS)
Terrell, Kyle (inventor)
1990-01-01
A piston and cylinder assembly is disclosed which is constructed of polyvinyl chloride that uses local water pressure to perform small lifting tasks. The chamber is either pressurized to extend the piston or depressurized to retract the piston. The present invention is best utilized for raising and lowering toilet seats.
Coriolis effects enhance lift on revolving wings.
Jardin, T; David, L
2015-03-01
At high angles of attack, an aircraft wing stalls. This dreaded event is characterized by the development of a leading edge vortex on the upper surface of the wing, followed by its shedding which causes a drastic drop in the aerodynamic lift. At similar angles of attack, the leading edge vortex on an insect wing or an autorotating seed membrane remains robustly attached, ensuring high sustained lift. What are the mechanisms responsible for both leading edge vortex attachment and high lift generation on revolving wings? We review the three main hypotheses that attempt to explain this specificity and, using direct numerical simulations of the Navier-Stokes equations, we show that the latter originates in Coriolis effects. PMID:25871040
NASA Heavy Lift Rotorcraft Systems Investigation
NASA Technical Reports Server (NTRS)
Johnson, Wayne; Yamauchi, Gloria K.; Watts, Michael E.
2005-01-01
The NASA Heavy Lift Rotorcraft Systems Investigation examined in depth several rotorcraft configurations for large civil transport, designed to meet the technology goals of the NASA Vehicle Systems Program. The investigation identified the Large Civil Tiltrotor as the configuration with the best potential to meet the technology goals. The design presented was economically competitive, with the potential for substantial impact on the air transportation system. The keys to achieving a competitive aircraft were low drag airframe and low disk loading rotors; structural weight reduction, for both airframe and rotors; drive system weight reduction; improved engine efficiency; low maintenance design; and manufacturing cost comparable to fixed-wing aircraft. Risk reduction plans were developed to provide the strategic direction to support a heavy-lift rotorcraft development. The following high risk areas were identified for heavy lift rotorcraft: high torque, light weight drive system; high performance, structurally efficient rotor/wing system; low noise aircraft; and super-integrated vehicle management system.
Comparison of Lifting-line and Lifting-surface Blade Models for Rotor Wakes
NASA Astrophysics Data System (ADS)
Kini, Santosh; Godavarty, Vishwanath; Conlisk, A. T.
2001-11-01
Rotor blades are often modeled using either the lifting-line or the lifting-surface approach. Classical lifting-line theory treats the case of a high aspect ratio blade where the blade is modeled by just a bound vortex line. The tip-vortex and the trailing vortex sheet are shed from the bound vortex line. The lifting-line approach will thus be expected to be accurate for large aspect ratio wings and blades. In the lifting-surface approach the blade is modeled by a thin lifting surface to more precisely understand the aerodynamic behavior of the chordwise flow close to the blade surface and its effect on the formation of the tip-vortex and consequently the entire wake. Due to the presence of a finite chord, this model seems to be more robust as far as chordwise changes in flow patterns are concerned. The objective of the present work is to compare results from two distinct vortex-wake models. The Wake Model with Lifting-Line has a robust model for the wake. The Lifting-Surface Model with Wake Effects has a robust model for the lifting-surface with additional terms in the governing equations representing a simplified model for the wake. Bound and tip-vortex strengths, tip-vortex trajectories and velocity profiles are evaluated using these models very close to the blade. It is observed that the results obtained from these models are comparable. The minor discrepancies observed are explained in terms of the assumptions made. To establish the validity of the models the results are also compared with experimental data. Sponsored by the Rotorcraft Center of Excellence at Georgia Institute of Technology and the Army Research Office.
Measurement of the noise generation at the trailing edge of porous airfoils
NASA Astrophysics Data System (ADS)
Geyer, T.; Sarradj, E.; Fritzsche, C.
2010-02-01
Owls are commonly known for their quiet flight, enabled by three adaptions of their wings and plumage: leading edge serrations, trailing edge fringes and a soft and elastic downy upper surface of the feathers. In order to gain a better understanding of the aeroacoustic effects of the third property that is equivalent to an increased permeability of the plumage to air, an experimental survey on a set of airfoils made of different porous materials was carried out. Several airfoils with the same shape and size but made of different porous materials characterized by their flow resistivities and one non-porous reference airfoil were subject to the flow in an aeroacoustic open jet wind tunnel. The flow speed has been varied between approximately 25 and 50 m/s. The geometric angle of attack ranged from -16° to 20° in 4°-steps. The results of the aeroacoustic measurements, made with a 56-microphone array positioned out of flow, and of the measurements of lift and drag are given and discussed.
Navier-Stokes analysis of airfoils with leading edge ice accretions
NASA Technical Reports Server (NTRS)
Potapczuk, Mark G.
1993-01-01
A numerical analysis of the flowfield characteristics and the performance degradation of an airfoil with leading edge ice accretions was performed. The important fluid dynamic processes were identified and calculated. Among these were the leading edge separation bubble at low angles of attack, complete separation on the low pressure surface resulting in premature shell, drag rise due to the ice shape, and the effects of angle of attack on the separated flow field. Comparisons to experimental results were conducted to confirm these calculations. A computer code which solves the Navier-Stokes equations in two dimensions, ARC2D, was used to perform the calculations. A Modified Mixing Length turbulence model was developed to produce grids for several ice shape and airfoil combinations. Results indicate that the ability to predict overall performance characteristics, such as lift and drag, at low angles of attack is excellent. Transition location is important for accurately determining separation bubble shape. Details of the flowfield in and downstream of the separated regions requires some modifications. Calculations for the stalled airfoil indicate periodic shedding of vorticity that was generated aft of the ice accretion. Time averaged pressure values produce results which compare favorably with experimental information. A turbulence model which accounts for the history effects in the flow may be justified.
NASA Technical Reports Server (NTRS)
Capece, Vincent R.; Platzer, Max F.
2003-01-01
A major challenge in the design and development of turbomachine airfoils for gas turbine engines is high cycle fatigue failures due to flutter and aerodynamically induced forced vibrations. In order to predict the aeroelastic response of gas turbine airfoils early in the design phase, accurate unsteady aerodynamic models are required. However, accurate predictions of flutter and forced vibration stress at all operating conditions have remained elusive. The overall objectives of this research program are to develop a transition model suitable for unsteady separated flow and quantify the effects of transition on airfoil steady and unsteady aerodynamics for attached and separated flow using this model. Furthermore, the capability of current state-of-the-art unsteady aerodynamic models to predict the oscillating airfoil response of compressor airfoils over a range of realistic reduced frequencies, Mach numbers, and loading levels will be evaluated through correlation with benchmark data. This comprehensive evaluation will assess the assumptions used in unsteady aerodynamic models. The results of this evaluation can be used to direct improvement of current models and the development of future models. The transition modeling effort will also make strides in improving predictions of steady flow performance of fan and compressor blades at off-design conditions. This report summarizes the progress and results obtained in the first year of this program. These include: installation and verification of the operation of the parallel version of TURBO; the grid generation and initiation of steady flow simulations of the NASA/Pratt&Whitney airfoil at a Mach number of 0.5 and chordal incidence angles of 0 and 10 deg.; and the investigation of the prediction of laminar separation bubbles on a NACA 0012 airfoil.
Characteristics of merging shear layers and turbulent wakes of a multi-element airfoil
NASA Astrophysics Data System (ADS)
Adair, Desmond; Horne, W. Clifton
1988-02-01
Flow characteristics in the vicinity of the trailing edge of a single-slotted airfoil flap are presented and analyzed. The experimental arrangement consisted of a NACA 4412 airfoil equipped with a NACA 4415 flap whose angle of deflection was 21.8 deg. The flow remained attached over the model surfaces except in the vicinity of the flap trailing edge where a small region of boundary-layer separation extended over the aft 7 percent of flap chord. The flow was complicated by the presence of a strong, initially inviscid jet emanating from the slot between airfoil and flap, and a gradual merging of the main airfoil wake and flap suction-side boundary layer. Downstream of the flap, the airfoil and flap wakes fully merged to form an asymmetrical curved wake. The airfoil configuration was tested at an angle of attack of 8.2 deg, at a Mach number of 0.09, and a chord based Reynolds number of 1.8 x 10 to the 6th power in the Ames Research Center 7- by 10-Foot Wind Tunnel. Surface pressure measurements were made on the airfoil and flap and on the wind tunnel roof and floor. It was estimated that the wall interference increased the C sub L by 7 percent and decreased the C sub M by 4.5 percent. Velocity characteristics were quantified using hot-wire anemometry in regions of flow with preferred direction and low turbulence intensity. A 3-D laser velocimeter was used in regions of flow recirculation and relatively high turbulence intensity.
Pressure distribution over an NACA 23012 airfoil with an NACA 23012 external-airfoil flap
NASA Technical Reports Server (NTRS)
Wenzinger, Carl J
1938-01-01
Report presents the results of pressure-distribution tests of an NACA 23012 airfoil with an NACA 23012 external airfoil flap made in the 7 by 10-foot wind tunnel. The pressures were measured on the upper and lower surfaces at one chord section on both the main airfoil and on the flap for several different flap deflections and at several angles of attack. A test installation was used in which the airfoil was mounted horizontally in the wind tunnel between vertical end planes so that two-dimensional flow was approximated. The data are presented in the form of pressure-distribution diagrams and as graphs of calculated coefficients for the airfoil-and-flap combination and for the flap alone.
Endoscopically assisted biplanar forehead lift.
Ramirez, O M
1995-08-01
The standard subgaleal coronal incision used for brow lifting is limited to patients with low foreheads. The subcutaneous hairline brow lift used for patients with high foreheads has a high rate of vascular complications. However, the main advantage of the subcutaneous approach is preservation of sensation posterior to the incision line. The subperiosteal approach, on the other hand, allows a better periorbital remodeling. I have combined the subperiosteal and the subcutaneous approach to take advantage of and minimize the disadvantages of each individual approach. The advent of the endoscopic technique has allowed more accurate and controlled periorbital dissection and brow depressor muscle modification. The operation is indicated in every patient in whom the anterior hairline incision is indicated. It is a good method for decreasing the height of the forehead. The dissection is done initially in the subcutaneous plane, and about halfway on the forehead slit incisions through the galea-periosteal layer and through the temporoparietal fascia are made to continue the dissection in the deep plane. The periosteal dissection and release at the arcus marginalis is done under endoscopic control. Likewise, the brow depressor muscle modification is done under endoscopic magnification. Deep anchoring sutures fix the brow in the elevated position. Trimming and closure of the cutaneous layer are done with minimal tension. The biplanar subperiosteal-subcutaneous forehead lift has been used in 24 patients with very satisfactory results. Complications have been of a minor nature. Patients have maintained sensation posterior to the hairline incision. The height of the forehead has been decreased in every case. Frontalis muscle function has been preserved. PMID:7624404
A lifting line model for propeller noise
NASA Astrophysics Data System (ADS)
Brouwer, H. H.
1989-04-01
The possibility of calculating the sound field of a propeller from knowledge of the spanwise lift distribution, is investigated. An expression for the sound pressure is derived within a first-order approximation for blades of high aspect ratio. In this approximation the propeller is described by a system of rotating lifting lines. Numerical results for a six-bladed model propeller are compared with experimental data and the results of a theory based on a detailed description of the propeller geometry. It appears that for a high-aspect-ratio propeller, the agreement is satisfactory.
The investigation of a variable camber blade lift control for helicopter rotor systems
NASA Technical Reports Server (NTRS)
Awani, A. O.
1982-01-01
A new rotor configuration called the variable camber rotor was investigated numerically for its potential to reduce helicopter control loads and improve hover performance. This rotor differs from a conventional rotor in that it incorporates a deflectable 50% chord trailing edge flap to control rotor lift, and a non-feathering (fixed) forward portion. Lift control is achieved by linking the blade flap to a conventional swashplate mechanism; therefore, it is pilot action to the flap deflection that controls rotor lift and tip path plane tilt. This report presents the aerodynamic characteristics of the flapped and unflapped airfoils, evaluations of aerodynamics techniques to minimize flap hinge moment, comparative hover rotor performance and the physical concepts of the blade motion and rotor control. All the results presented herein are based on numerical analyses. The assessment of payoff for the total configuration in comparison with a conventional blade, having the same physical characteristics as an H-34 helicopter rotor blade was examined for hover only.
NASA Technical Reports Server (NTRS)
Mueller, T. J.
1983-01-01
Data was obtained for the NACA 663-018 airfoil in order to assess the advantages of tripping the boundary layer very near the leading edge for chord Reynolds numbers between 40,000 and 200,000. Single element trips were made from tape 2.5 mm wide and 0.15 mm thick. The unmodified tape trip in five thicknesses was placed across the span at 1.1% chord and the results of lift and drag measurements compared with the smooth airfoil case. Saw tooth geometry trips were also cut from tape and studied using five thicknesses. The saw tooth trips were placed across the span with the sharp points facing upstream at 1.1% chord and with the valley of the teeth at the 2.5% chord position. A large number of bumps and wiggles were produced in the lift coefficient versus angle of attack curves with various combinations of Reynolds number, thickness, and type of trips. The main result in the drag coefficient was an increase in C sub d min with increase in trip thickness.
Reinforcement Learning of Morphing Airfoils with Aerodynamic and Structural Effects
Valasek, John
Reinforcement Learning of Morphing Airfoils with Aerodynamic and Structural Effects Amanda Lampton Changing / Morphing for Micro Air Vehicles Learning Morphing Airfoil Results Conclusions Research Issues controllers, and mechanisms to achieve the state change. Morphing Aircraft DARPA's definition Aerospace
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.
The anchor subperiosteal forehead lift.
Ramirez, O M
1995-05-01
The coronal incision used for brow lift procedure has a high rate of localized alopecia, widening, and depression of the scar at the suture line. Other sequelae of the standard coronal brow lift incision procedure are "stretch-back" with a recurrent brow ptosis, poor brow elevation, and numbness beyond the incision line. Factors causing alopecia are tension, use of a monopolar cautery, use of key sutures with undue tension, one-layer closure, and sutures left too long. Recurrent brow ptosis may be due to anterior displacement of the posterior scalp flap, stretching of the anterior frontal skin flap, or insufficient power of the weakened frontalis muscle. Poor brow elevation may be due to unsatisfactory dissection on the glabella and orbital rims. Numbness and itching beyond the incision line are due to a low coronal incision. To avoid these problems, the following principles were followed: (1) If not contraindicated, the incision is made high on the vertex of the head, posterior to a biauricular line. (2) The pericranium is included in the frontal flap starting at the incision lines. (3) The subperiosteal dissection is continued down to the orbital rims and nasal bones. (4) The release of the periosteum at the arcus marginalis or just above allows repositioning of the brow structures. (5) The inelastic pericranium maintains the position of the elevated structures and avoids stretching of the frontal skin. (6) The integrity of the frontalis muscle is maintained completely. (7) Two large triangles of scalp resected in the posterior flaps allow fixing the position of the posterior scalp and match better the length of the anterior flap. (8) The galea periosteal rim flap allows anchoring of the frontal flap to the undersurface of the posterior scalp flap. This stabilizes the closure with minimal tension on the hair-bearing portion of the scalp. The wide surface of contact avoids depression and widening at the suture line. (9) Closure with skin staples avoids constriction of the hair follicles. (10) Hemostasis is done with a bipolar cautery. (11) No through-and-through key sutures are used. Some of these principles were introduced to the endoscopic subperiosteal forehead lift. The modifications mentioned above have been used in 92 open brow/face lift procedures with excellent aesthetic and functional results and minimal complications. PMID:7732147
NASA Astrophysics Data System (ADS)
Ragni, Daniele; Vitale, Laura; Ianiro, Andrea; Geurts, Ben; Ferreira, Carlos
2012-11-01
In the present study, an airfoil optimized for vertical axis wind turbines applications has been developed with a genetic algorithm, selecting the geometry with maximum (dcl/d ?) /cd among airfoils generated with 16 shape functions. The airfoil, operating in the curved trajectory of a vertical axis wind turbine, is usually optimized adopting conformal mappings in the straight path. Recent experimental results have shown disagreement with this approach, due to the forces determined in the curved flow path. To investigate the effects of flow rotation, an aluminum model (c=0.25m) has been manufactured from the optimized shape and further tested in the LST tunnel of the TUDelft at Reynolds number 106. Planar PIV experiments in combination with the PIV based load determination technique have been performed to simultaneously obtain velocity fields and loads. Results including velocity, pressure distributions, lift and drag are initially discussed in a steady airfoil configuration and compared with numerical results. Successively, the model has been unsteadily pitched using a magnetic linear actuator (up to 3 Hz frequency), with a free stream V? = 40 m/s corresponding to Re = 0.7 ×106. Phase locked PIV vector fields have been acquired and compared to the steadily obtained results.
E. W. M. Roosenboom
2011-01-01
The aim of the thesis is to measure the propeller slipstream properties (velocity and vorticity) and to assess the unsteady and instantaneous behavior of the propeller flow field at high disk loadings, zero thrust and thrust reverse using the image based measurement techniques. Along with its implementation of the techniques in industrial facilities particular propeller phenomena are addressed, such as
Jacono, Andrew A; Talei, Benjamin
2014-05-01
The authors' vertical neck lifting procedure is an extended deep plane facelift, which elevates the skin and SMAS-platysma complex as a composite unit. The goal is to redrape cervicomental laxity vertically onto the face rather than laterally and postauricularly. The authors consider this an extended technique because it lengthens the deep plane flap from the angle of the mandible into the neck to release the cervical retaining ligaments that limit platysmal redraping. This technique does not routinely use midline platysmal surgery because it counteracts the extent of vertical redraping. A majority of aging face patients are good candidates for this procedure in isolation, but indications for combining vertical neck lifting with submental surgery are elucidated. PMID:24745389
Tessier, P
1989-01-01
The "facial mask" is composed of all of the tissues lying on top of the skeleton: periosteum, deep adipose tissue, superficial musculo-aponeurotic tissue and skin. The periosteum is the intermediate zone between the skeleton, responsible for the shape of the face, and the more superficial tissues which complete the shapes and, most importantly, represent the mobile part of the face and consequently the site of facial expression. The secret of an effective "mask-lift" depends on complete subperiosteal dissection of the malar bones, zygomatic arches and orbital margins. This dissection can be performed via a coronal approach, but it is easier to start the subperiosteal dissection via a short vestibular incision. Subperiosteal dissection via a coronal incision is not only useful to lift the facial mask; it is also useful for remodelling the orbital margins and to obtain bone grafts from the parietal area in order to reinforce the glabella, check bones and nasogenial folds. PMID:2473674
Powered-lift aircraft technology
NASA Technical Reports Server (NTRS)
Deckert, W. H.; Franklin, J. A.
1989-01-01
Powered lift aircraft have the ability to vary the magnitude and direction of the force produced by the propulsion system so as to control the overall lift and streamwise force components of the aircraft, with the objective of enabling the aircraft to operate from minimum sized terminal sites. Power lift technology has contributed to the development of the jet lift Harrier and to the forth coming operational V-22 Tilt Rotor and the C-17 military transport. This technology will soon be expanded to include supersonic fighters with short takeoff and vertical landing capability, and will continue to be used for the development of short- and vertical-takeoff and landing transport. An overview of this field of aeronautical technology is provided for several types of powered lift aircraft. It focuses on the description of various powered lift concepts and their operational capability. Aspects of aerodynamics and flight controls pertinent to powered lift are also discussed.
Enhanced Rescue Lift Capability
NASA Technical Reports Server (NTRS)
Young, Larry A.
2007-01-01
The evolving and ever-increasing demands of emergency response and disaster relief support provided by rotorcraft dictate, among other things, the development of enhanced rescue lift capability for these platforms. This preliminary analysis is first-order in nature but provides considerable insight into some of the challenges inherent in trying to effect rescue using a unique form of robotic rescue device deployed and operated from rotary-wing aerial platforms.
Effect of chordwise forces and deformations and deformations due to steady lift on wing flutter
NASA Technical Reports Server (NTRS)
Boyd, W. N.
1977-01-01
This investigation explores the effects of chordwise forces and deformations and steady-state deformation due to lift on the static and dynamic aeroelastic stability of a uniform cantilever wing. Results of this analysis are believed to have practical applications for high-performance sailplanes and certain RPV's. The airfoil cross section is assumed to be symmetric and camber bending is neglected. Motions in vertical bending, fore-and-aft bending, and torsion are considered. A differential equation model is developed, which included the nonlinear elastic bending-torsion coupling that accompanies even moderate deflections. A linearized expansion in small time-dependent deflections is made about a steady flight condition. The stability determinant of the linearized system then contains coefficients that depend on steady displacements. Loads derived from two-dimensional incompressible aerodynamic theory are used to obtain the majority of the results, but cases using three-dimensional subsonic compressible theory are also studied. The stability analysis is carried out in terms of the dynamically uncoupled natural modes of vibration of the uniform cantilever.
NASA Technical Reports Server (NTRS)
Nagamatsu, H. T.; Dyer, R.
1984-01-01
The passive shock wave/boundary layer control for reducing the drag of 14%-thick supercritical airfoil was investigated in the 3 in. x 15.4 in. RPI Transonic Wind Tunnel with and without the top wall insert at transonic Mach numbers. Top wall insert was installed to increase the flow Mach number to 0.90 with the model mounted on the test section bottom wall. Various porous surfaces with a cavity underneath were positioned on the area of the airfoil where the shock wave occurs. The higher pressure behind the shock wave circulates flow through the cavity to the lower pressure ahead of the shock wave. The effects from this circulation prevent boundary layer separation and enthropy increase hrough the shock wave. The static pressure distributions over the airfoil, the wake impact pressure survey for determining the profile drag and the Schlieren photographs for porous surfaces are presented and compared with the results for solid surface airfoil. With a 2.8% uniform porosity the normal shock wave for the solid surface was changed to a lambda shock wave, and the wake impact pressure data indicate a drag coefficient reduction as much as 45% lower than for the solid surface airfoil at high transonic Mach numbers.
The semidiscrete Galerkin finite element modeling of compressible viscous flow past an airfoil
NASA Technical Reports Server (NTRS)
Meade, Andrew J.
1990-01-01
The primary project was the numerical simulation, by a finite element/finite difference method, of the viscous flow about an airfoil. The secondary project involved the numerical simulation of the three-dimensional separated and vortex-dominated flow about a hemispherically capped cylinder in the transonic regime. Preliminary calculations were started for the hemisphere-cylinder at 0 and 5 degree angle of attack. The solution of the flow field about airfoils and wings is required to determine the important parameters of lift, moment, and drag. Viscous effects must be accounted for if the drag is to be accurately calculated. At present there are basically two approaches to the numerical simulation of the flow field, the use of fully viscous models and the inviscid/viscous models. The fully viscous models require the solution of an approximation of the Navier-Stokes equations and therefore should simulate most of the physical mechanisms. A fast, accurate, and computationally efficient inviscid flow solver was recently developed by Hartwich. It is thought that Hartwich's program coupled to a fast, accurate, and computationally efficient boundary layer code, will make an excellent tool for airfoil design. The purpose of the primary project was to develop a compressible boundary layer code using the semidiscrete Galerkin finite element method. The numerical scheme employed used the combination of a Dorodnitsyn formulation of the boundary layer equations, with a finite difference/finite element procedure (semidiscrete Galerkin method), in the solution of the compressible two-dimensional boundary layer equations. A laminar compressible boundary layer code was developed and tested for a NACA 0012 airfoil at a Mach number of 0.5, a Reynolds number of 5000, and zero angle of attack. At present the boundary layer program solves up to, but not beyond, separation.
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.
Atlantis is lifted from its transporter in the VAB
NASA Technical Reports Server (NTRS)
2001-01-01
KENNEDY SPACE CENTER, Fla. -- This closeup shows the workers, standing on lifts, who are checking the bolts on the apparatus holding the orbiter Atlantis. The orbiter will be rotated and lifted into high bay 1 where it will be stacked with its external tank and solid rocket boosters. Space Shuttle Atlantis is scheduled to launch on mission STS-104 in early July.
NASA Technical Reports Server (NTRS)
Goradia, S. H.; Mehta, J. M.; Shrewsbury, G. S.
1977-01-01
The viscous flow phenomena associated with sharp and blunt trailing edge airfoils were investigated. Experimental measurements were obtained for a 17 percent thick, high performance GAW-1 airfoil. Experimental measurements consist of velocity and static pressure profiles which were obtained by the use of forward and reverse total pressure probes and disc type static pressure probes over the surface and in the wake of sharp and blunt trailing edge airfoils. Measurements of the upper surface boundary layer were obtained in both the attached and separated flow regions. In addition, static pressure data were acquired, and skin friction on the airfoil upper surface was measured with a specially constructed device. Comparison of the viscous flow data with data previously obtained elsewhere indicates reasonable agreement in the attached flow region. In the separated flow region, considerable differences exist between these two sets of measurements.
Discovery is lifted in the VAB
NASA Technical Reports Server (NTRS)
2001-01-01
The orbiter Discovery (seen from the front) is lifted to vertical in the transfer aisle of the Vehicle Assembly Building. It will then be lifted up and into high bay 1 for mating with its solid rocket boosters and external tank. Discovery will be launched March 8 on mission STS-102, the eighth construction flight to the International Space Station. The Shuttle will carry the Multi- Purpose Logistics Module Leonardo, the first of three pressurized modules provided by the Italian Space Agency to carry supplies and equipment to the Space Station and back to earth.
Discovery is lifted in the VAB
NASA Technical Reports Server (NTRS)
2001-01-01
In the transfer aisle of the Vehicle Assembly Building, the orbiter Discovery (seen from the back) is lifted to vertical. It will then be lifted up and into high bay 1 for mating with its solid rocket boosters and external tank. Discovery will be launched March 8 on mission STS-102, the eighth construction flight to the International Space Station. The Shuttle will carry the Multi-Purpose Logistics Module Leonardo, the first of three pressurized modules provided by the Italian Space Agency to carry supplies and equipment to the Space Station and back to earth.
Discovery is lifted in the VAB
NASA Technical Reports Server (NTRS)
2001-01-01
The orbiter Discovery is lifted by cranes in the transfer aisle of the Vehicle Assembly Building. It will next be lifted into a vertical position and into high bay 1 for mating with its solid rocket boosters and external tank. Discovery will be launched March 8 on mission STS-102, the eighth construction flight to the International Space Station. The Shuttle will carry the Multi- Purpose Logistics Module Leonardo, the first of three pressurized modules provided by the Italian Space Agency to carry supplies and equipment to the Space Station and back to Earth.
Discovery is lifted in the VAB
NASA Technical Reports Server (NTRS)
2001-01-01
The underside of orbiter Discovery is seen as it is lifted to vertical in the transfer aisle of the Vehicle Assembly Building. It will then be lifted up and into high bay 1 for mating with its solid rocket boosters and external tank. Discovery will be launched March 8 on mission STS-102, the eighth construction flight to the International Space Station. The Shuttle will carry the Multi-Purpose Logistics Module Leonardo, the first of three pressurized modules provided by the Italian Space Agency to carry supplies and equipment to the Space Station and back to earth.
Discovery is lifted in the VAB
NASA Technical Reports Server (NTRS)
2001-01-01
In the transfer aisle of the Vehicle Assembly Building the orbiter Discovery is suspended at an angle as it is lifted to a vertical position. It will next be lifted into high bay 1 for mating with its solid rocket boosters and external tank. Discovery will be launched March 8 on mission STS-102, the eighth construction flight to the International Space Station. The Shuttle will carry the Multi-Purpose Logistics Module Leonardo, the first of three pressurized modules provided by the Italian Space Agency to carry supplies and equipment to the Space Station and back to Earth.
Soccer ball lift coefficients via trajectory analysis
NASA Astrophysics Data System (ADS)
Goff, John Eric; Carré, Matt J.
2010-07-01
We performed experiments in which a soccer ball was launched from a machine while two high-speed cameras recorded portions of the trajectory. Using the trajectory data and published drag coefficients, we extracted lift coefficients for a soccer ball. We determined lift coefficients for a wide range of spin parameters, including several spin parameters that have not been obtained by today's wind tunnels. Our trajectory analysis technique is not only a valuable tool for professional sports scientists, it is also accessible to students with a background in undergraduate-level classical mechanics.
NASA Technical Reports Server (NTRS)
Silverstein, Abe; Becker, John V
1938-01-01
For the purpose of studying the transition from laminar to turbulent flow, boundary-layer measurements were made in the NACA full-scale wind tunnel on three symmetrical airfoils of NACA 0009, 0012, and 0018 sections. The effects of variations in lift coefficient, Reynolds number, and airfoil thickness on transition were investigated. Air speed in the boundary layer was measured by total-head tubes and by hot wires; a comparison of transition as indicated by the two techniques was obtained. The results indicate no unique value of Reynolds number for the transition, whether the Reynolds number is based upon the distance along the chord or upon the thickness of the boundary layer at the transition point. In general, the transition is not abrupt and occurs in a region that varies in length as a function of the test conditions.
NASA Technical Reports Server (NTRS)
Bragg, Michael B.
1993-01-01
This is the second volume of a report documenting the effect of simulated ice accretion on the aerodynamic performance of a NACA 0012 airfoil. Both an experimentally measured and a computer generated ice shape are studied. The purpose of this report is to present the results of the measurements, not an analysis of the data. Surface pressure, integrated lift and pitching moment data are presented as well as drag from a wake survey. A split hot film probe was used to document the flow-field about the airfoil with simulated ice. Data in the separation bubbles, reattached boundary layer and wake are presented. Both tabulated and graphical data are presented in the paper. The data are also available on computer disk for easy access.
NASA Technical Reports Server (NTRS)
Hackett, J. E.; Boles, R. A.; Lilley, D. E.
1976-01-01
Ground effects experiments and large/small-tunnel interference studies were carried out on a model with a 20-inch (50.8 cm) span wing. The wing, which includes a highly deflected knee-blown flap can be fitted with unflapped tips and slats. A low-mounted tailplane can be fitted to the aft fuselage. Three-component balance meaurements, made with a fixed ground equipped with a single boundary-layer blowing slot, were compared with datum, moving-ground results. Good comparisons were obtained up to model blowing momentum coefficients of approximately two, after which the particular floor blowing settings used proved insufficient to prevent floor separation in the vicinity of the model. Skin friction measurements, taken routinely along the floor centerline, proved invaluable during the analysis of results, and their use is recommended as input to determination of floor BLC setting. A careful investigation was made of pitching moments, including tail-on, close-to-ground cases, with favorable results. Drag proved the most sensitive to the change from a moving to the boundary-layer controlled ground.
NASA Technical Reports Server (NTRS)
Bergrun, N. R.
1951-01-01
An empirical method for the determination of the area, rate, and distribution of water-drop impingement on airfoils of arbitrary section is presented. The procedure represents an initial step toward the development of a method which is generally applicable in the design of thermal ice-prevention equipment for airplane wing and tail surfaces. Results given by the proposed empirical method are expected to be sufficiently accurate for the purpose of heated-wing design, and can be obtained from a few numerical computations once the velocity distribution over the airfoil has been determined. The empirical method presented for incompressible flow is based on results of extensive water-drop. trajectory computations for five airfoil cases which consisted of 15-percent-thick airfoils encompassing a moderate lift-coefficient range. The differential equations pertaining to the paths of the drops were solved by a differential analyzer. The method developed for incompressible flow is extended to the calculation of area and rate of impingement on straight wings in subsonic compressible flow to indicate the probable effects of compressibility for airfoils at low subsonic Mach numbers.
NASA Technical Reports Server (NTRS)
Wenzinger, Carl J; Bamber, Millard J
1938-01-01
A large-chord NACA 23012 airfoil was tested. The airfoil extended completely across the test section, and two-dimensional flow was approximated. The model was fitted with a full-span slotted flap having a chord 25.66 percent of the airfoil chord. The ailerons investigated extended over the entire span and each had a chord 10 percent of the airfoil chord. The types of ailerons tested were: retractable ailerons, slot-lip ailerons using the lip of the slot for ailerons, and plain ailerons on the trailing edge of the slotted flap. The data are presented in the form of curves of section lift, drag, and pitching-moment coefficients for the airfoil with flap deflected but with ailerons neutral, and of rolling-moment, yawing-moment, and hinge-moment coefficients calculated for a rectangular wing of aspect ratio 6 with a semi-span aileron and a full-span flap. For the ailerons investigated the data indicate that, from considerations of rolling and yawing moments produced and of stick forces desired, the retractable aileron is the most satisfactory means of lateral control for use with a full-span slotted flap.
Shelton, M.L.; Gregory, B.A. (Turbine Aero and Cooling Technology, General Electric Aircraft Engines, Cincinnati, OH (United States)); Doughty, R.L.; Kiss, T.; Moses, H.L. (Virginia Polytechnic Inst. and State Univ., Blacksburg, VA (United States). Mechanical Engineering Dept.)
1993-07-01
In aircraft engine design (and in other applications), small improvements in turbine efficiency may be significant. Since analytical tools for predicting transonic turbine losses are still being developed, experimental efforts are required to evaluate various designs, calibrate design methods, and validate CFD analysis tools. However, these experimental efforts must be very accurate to measure the performance differences to the levels required by the highly competitive aircraft engine market. Due to the sensitivity of transonic and supersonic flow fields, it is often difficult to obtain the desired level of accuracy. In this paper, a statistical approach is applied to the experimental evaluation of transonic turbine airfoils in the VPI and SU transonic cascade facility in order to quantify the differences between three different transonic turbine airfoils. This study determines whether the measured performance differences between the three different airfoils are statistically significant. This study also assesses the degree of confidence in the transonic cascade testing process at VPI and SU.
NASA Astrophysics Data System (ADS)
Gladden, H. J.; Proctor, M. P.
A transient technique was used to measure heat transfer coefficients on stator airfoils in a high-temperature annular cascade at real engine conditions. The transient response of thin film thermocouples on the airfoil surface to step changes in the gas stream temperature was used to determine these coefficients. In addition, gardon gages and paired thermocouples were also utilized to measure heat flux on the airfoil pressure surface at steady state conditions. The tests were conducted at exit gas stream Reynolds numbers of one-half to 1.9 million based on true chord. The results from the transient technique show good comparison with the steady-state results in both trend and magnitude. In addition, comparison is made with the STAN5 boundary layer code and shows good comparison with the trends. However, the magnitude of the experimental data is consistently higher than the analysis.
Airfoil leading-edge suction and energy conservation for compressible flow
Amiet, R.K.
1995-04-01
The leading-edge suction force produced when a flat-plate airfoil at zero angle of attack encounters a vertical gust was examined for compressible flow with a time-dependent gust. A simple derivation of the thrust force shows that the acoustic energy can be calculated using compact assumptions at low frequency, but that it must be calculated non-compactly at high frequency. For a general gust, the work done on the airfoil equals the energy taken from the fluid. For a sinusoidal gust the energy contained in the incident gust equals the sum of the energy remaining in the wake, the work done on the airfoil and the acoustic energy radiated away. Also, the relative proportions of the energy going to these three energy types depend on the gust frequency.
Approximate and exact numerical computation of supersonic flow over an airfoil
NASA Technical Reports Server (NTRS)
Lewis, T. S.; Sirovich, L.
1981-01-01
The computation of inviscid supersonic flow over a two-dimensional airfoil is considered. There are two main nonlinear approaches which lead to approximate solutions. Small-amplitude theory gives valid solutions provided the airfoil thickness is not too great and the Mach number is not too high. The second type of approximation, shock expansion theory, employs the fact that even for flows with strong shocks the effect of the down-running characteristics remains small. This leads to an analytic solution at the airfoil. It is pointed out that the approximate solution is accurate enough for many cases of interest. The numerical method furnishes a rapid correction to the solution in those cases where it is not. The characteristic-streamline coordinate system is useful both for the computation of the approximate solution and the corrections, and is also convenient for displaying and interpreting the results.
Numerical Simulation of Shock-stall Flutter of an Airfoil using the Navier-Stokes Equations
NASA Astrophysics Data System (ADS)
Isogai, K.
1993-08-01
In order to confirm qualitatively that the experimentally observed, unusual flutter phenomenon for a high-aspect-ratio (non-tailored) forward swept wing model is indeed shock-stall flutter, the aeroelastic response calculation of a two-dimensional airfoil whose vibration characteristics are similar to those of the typical section of a forward swept wing, has been performed by solving the compressible Navier-Stokes equations. By examination of the flow pattern, pressure distribution and the behavior of the unsteady aerodynamic forces during the diverging oscillation of the airfoil, it is concluded that (i) this is a shock-stall flutter, in which the large-scale shock-induced flow separation plays a dominant role and (ii) there is a mechanism of energy input into the elastic system of the airfoil, leading to nearly a single-degree-of-freedom flutter.
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.