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NASA NTRS1991

Aerodynamic characteristics of a rotorcraft airfoil designed for the tip region of a main rotor blade

A wind tunnel investigation was conducted to determine the 2-D aerodynamic characteristics of a new rotorcraft airfoil designed for application to the tip region (stations outboard of 85 pct. radius) of a helicopter main rotor blade. The new airfoil, the RC(6)-08, and a baseline airfoil, the RC(3)-08, were investigated in the Langley 6- by 28-inch transonic tunnel at Mach numbers from 0.37 to 0.90. The Reynolds number varied from 5.2 x 10(exp 6) at the lowest Mach number to 9.6 x 10(exp 6) at the highest Mach number. Some comparisons were made of the experimental data for the new airfoil and the predictions of a transonic, viscous analysis code. The results of the investigation indicate that the RC(6)-08 airfoil met the design goals of attaining higher maximum lift coefficients than the baseline airfoil while maintaining drag divergence characteristics at low lift and pitching moment characteristics nearly the same as those of the baseline airfoil. The maximum lift coefficients of the RC(6)-08 varied from 1.07 at M=0.37 to 0.94 at M=0.52 while those of the RC(3)-08 varied from 0.91 to 0.85 over the same Mach number range. At lift coefficients of -0.1 and 0, the drag divergence Mach number of both the RC(6)-08 and the RC(3)-08 was 0.86. The pitching moment coefficients of the RC(6)-08 were less negative than those of the RC(3)-08 for Mach numbers and lift coefficients typical of those that would occur on a main rotor blade tip at high forward speeds on the advancing side of the rotor disk.

Noonan, Kevin W.↗

NASA NTRS1938

Tests of N-85, N-86 and N-87 airfoil sections in the 11-inch high speed wind tunnel

Three airfoils, the N-85, the N-86, and the N-87, were tested at the request of the Bureau of Aeronautics, Navy Department, to determine the suitability of these sections for use as propeller-blade sections. Further tests of the NACA 0009-64 airfoil were also made to measure the aerodynamic effect of thickening the trailing edge in accordance with current propeller practice. The N-86 and the N-87 airfoils appear to be nearly equivalent aerodynamically and both are superior to the N-85 airfoil. Comparison of those airfoils with the previously developed NACA 2409-34 airfoils indicate that the NACA 2409-34 is superior, particularly at high speeds. Thickening the trailing edge appears to have a detrimental effect, although the effect may be small if the trailing-edge radius is less than 0.5 percent of the cord. The N-86 and the N-87 airfoils appear to be nearly equivalent.

Stack, John↗

NASA NTRS1993

Effect of an extendable slat on the stall behavior of a VR-12 airfoil

Experimental and computational tests were performed on a VR-12 airfoil to determine if the dynamic-stall behavior that normally accompanies high-angle pitch oscillations could be modified by segmenting the forward portion of the airfoil and extending it ahead of the main element. In the extended position the configuration would appear as an airfoil with a leading-edge slat, and in the retracted position it would appear as a conventional VR-12 airfoil. The calculations were obtained from a numerical code that models the vorticity transport equation for an incompressible fluid. These results were compared with test data from the water tunnel facility of the Aeroflightdynamics Directorate at Ames Research Center. Steady and unsteady flows around both airfoils were examined at angles of attack between 0 and 30 deg. The Reynolds number was fixed at 200,000 and the unsteady pitch oscillations followed a sinusoidal motion described by alpha = alpha(sub m) + 10 deg sin(omega t). The mean angle (alpha(sub m)) was varied from 10 to 20 deg and the reduced frequency from 0.05 to 0.20. The results from the experiment and the calculations show that the extended-slat VR-12 airfoil experiences a delay in both static and dynamic stall not experienced by the basic VR-12 airfoil.

Dehugues, P. Plantin↗

NASA NTRS1994

A Two Element Laminar Flow Airfoil Optimized for Cruise

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.

Steen, Gregory Glen↗

NASA NTRS1995

Control of unsteady separated flow associated with the dynamic stall of airfoils

An effort to understand and control the unsteady separated flow associated with the dynamic stall of airfoils was funded for three years through the NASA cooperative agreement program. As part of this effort a substantial data base was compiled detailing the effects various parameters have on the development of the dynamic stall flow field. Parameters studied include Mach number, pitch rate, and pitch history, as well as Reynolds number (through two different model chord lengths) and the condition of the boundary layer at the leading edge of the airfoil (through application of surface roughness). It was found for free stream Mach numbers as low as 0.4 that a region of supersonic flow forms on the leading edge of the suction surface of the airfoil at moderate angles of attack. The shocks which form in this supersonic region induce boundary-layer separation and advance the dynamic stall process. Under such conditions a supercritical airfoil profile is called for to produce a flow field having a weaker leading-edge pressure gradient and no leading-edge shocks. An airfoil having an adaptive-geometry, or dynamically deformable leading edge (DDLE), is under development as a unique active flow-control device. The DDLE, formed of carbon-fiber composite and fiberglass, can be flexed between a NACA 0012 profile and a supercritical profile in a controllable fashion while the airfoil is executing an angle-of-attack pitch-up maneuver. The dynamic stall data were recorded using point diffraction interferometry (PDI), a noninvasive measurement technique. A new high-speed cinematography system was developed for recording interferometric images. The system is capable of phase-locking with the pitching airfoil motion for real-time documentation of the development of the dynamic stall flow field. Computer-aided image analysis algorithms were developed for fast and accurate reduction of the images, improving interpretation of the results.

Wilder, M. C.↗

NASA NTRS1995

An approach to the constrained design of natural laminar flow airfoils

A design method has been developed by which an airfoil with a substantial amount of natural laminar flow can be designed, while maintaining other aerodynamic and geometric constraints. After obtaining the initial airfoil's pressure distribution at the design lift coefficient using an Euler solver coupled with an integml turbulent boundary layer method, the calculations from a laminar boundary layer solver are used by a stability analysis code to obtain estimates of the transition location (using N-Factors) for the starting airfoil. A new design method then calculates a target pressure distribution that will increase the larninar flow toward the desired amounl An airfoil design method is then iteratively used to design an airfoil that possesses that target pressure distribution. The new airfoil's boundary layer stability characteristics are determined, and this iterative process continues until an airfoil is designed that meets the laminar flow requirement and as many of the other constraints as possible.

Green, Bradford Earl↗

NASA NTRS1997

Modern Airfoil Ice Accretions

This report presents results from the first icing tests performed in the Modem Airfoils program. Two airfoils have been subjected to icing tests in the NASA Lewis Icing Research Tunnel (IRT). Both airfoils were two dimensional airfoils; one was representative of a commercial transport airfoil while the other was representative of a business jet airfoil. The icing test conditions were selected from the FAR Appendix C envelopes. Effects on aerodynamic performance are presented including the effects of varying amounts of glaze ice as well as the effects of approximately the same amounts of glaze, mixed, and rime ice. Actual ice shapes obtained in these tests are also presented for these cases. In addition, comparisons are shown between ice shapes from the tests and ice shapes predicted by the computer code, LEWICE for similar conditions. Significant results from the tests are that relatively small amounts of ice can have nearly as much effect on airfoil lift coefficient as much greater amounts of ice and that glaze ice usually has a more detrimental effect than either rime or mixed ice. LEWICE predictions of ice shapes, in general, compared reasonably well with ice shapes obtained in the IRT, although differences in details of the ice shapes were observed.

Addy, Harold E., Jr.↗

NASA NTRS1997

An Approach to the Constrained Design of Natural Laminar Flow Airfoils

A design method has been developed by which an airfoil with a substantial amount of natural laminar flow can be designed, while maintaining other aerodynamic and geometric constraints. After obtaining the initial airfoil's pressure distribution at the design lift coefficient using an Euler solver coupled with an integral turbulent boundary layer method, the calculations from a laminar boundary layer solver are used by a stability analysis code to obtain estimates of the transition location (using N-Factors) for the starting airfoil. A new design method then calculates a target pressure distribution that will increase the laminar flow toward the desired amount. An airfoil design method is then iteratively used to design an airfoil that possesses that target pressure distribution. The new airfoil's boundary layer stability characteristics are determined, and this iterative process continues until an airfoil is designed that meets the laminar flow requirement and as many of the other constraints as possible.

Green, Bradford E.↗

NASA NTRS1999

SiC/SiC Leading Edge Turbine Airfoil Tested Under Simulated Gas Turbine Conditions

Silicon-based ceramics have been proposed as component materials for use in gas turbine engine hot-sections. A high pressure burner rig was used to expose both a baseline metal airfoil and ceramic matrix composite leading edge airfoil to typical gas turbine conditions to comparatively evaluate the material response at high temperatures. To eliminate many of the concerns related to an entirely ceramic, rotating airfoil, this study has focused on equipping a stationary metal airfoil with a ceramic leading edge insert to demonstrate the feasibility and benefits of such a configuration. Here, the idea was to allow the SiC/SiC composite to be integrated as the airfoil's leading edge, operating in a "free-floating" or unrestrained manner. and provide temperature relief to the metal blade underneath. The test included cycling the airfoils between simulated idle, lift, and cruise flight conditions. In addition, the airfoils were air-cooled, uniquely instrumented, and exposed to the same internal and external conditions, which included gas temperatures in excess of 1370 C (2500 F). Results show the leading edge insert remained structurally intact after 200 simulated flight cycles with only a slightly oxidized surface. The instrumentation clearly suggested a significant reduction (approximately 600 F) in internal metal temperatures as a result of the ceramic leading edge. The object of this testing was to validate the design and analysis done by Materials Research and Design of Rosemont, PA and to determine the feasibility of this design for the intended application.

Robinson, R. Craig↗

NASA NTRS2003

An Experimental and Computational Investigation of Oscillating Airfoil Unsteady Aerodynamics at Large Mean Incidence

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.

Capece, Vincent R.↗

NASA NTRS2003

Airfoil Design and Rotorcraft Performance

The relationship between global performance of a typical helicopter and the airfoil environment, as represented by the airfoil angles of attack and Mach number, has been examined using the comprehensive analysis CAMRAD II. A general correspondence is observed between global performance parameters, such as rotor L/D, and airfoil performance parameters, such as airfoil L/D, the drag bucket boundaries, and the divergence Mach number. Effects of design parameters such as blade twist and rotor speed variation have been examined and, in most cases, improvements observed in global performance are also observed in terms of airfoil performance. The relations observed between global Performance and the airfoil environment suggests that the emphasis in airfoil design should be for good L/D, while the maximum lift coefficient performance is less important.

Bousman, William G.↗

NASA NTRS2003

Robust Airfoil Optimization in High Resolution Design Space

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.

Li, Wu↗

NASA NTRS2003

Boundary-Layer Separation Control under Low-Pressure Turbine Airfoil Conditions using Glow-Discharge Plasma Actuators

Modem low-pressure turbines, in general, utilize highly loaded airfoils in an effort to improve efficiency and to lower the number of airfoils needed. Typically, the airfoil boundary layers are turbulent and fully attached at takeoff conditions, whereas a substantial fraction of the boundary layers on the airfoils may be transitional at cruise conditions due to the change of density with altitude. The strong adverse pressure gradients on the suction side of these airfoils can lead to boundary-layer separation at the latter low Reynolds number conditions. Large separation bubbles, particularly those which fail to reattach, cause a significant degradation of engine efficiency. A component efficiency drop of the order 2% may occur between takeoff and cruise conditions for large commercial transport engines and could be as large as 7% for smaller engines at higher altitude. An efficient means of of separation elimination/reduction is, therefore, crucial to improved turbine design. Because the large change in the Reynolds number from takeoff to cruise leads to a distinct change in the airfoil flow physics, a separation control strategy intended for cruise conditions will need to be carefully constructed so as to incur minimum impact/penalty at takeoff. A complicating factor, but also a potential advantage in the quest for an efficient strategy, is the intricate interplay between separation and transition for the situation at hand. Volino gives a comprehensive discussion of several recent studies on transition and separation under low-pressure-turbine conditions, among them one in the present facility. Transition may begin before or after separation, depending on the Reynolds number and other flow conditions. If the transition occurs early in the boundary layer then separation may be reduced or completely eliminated. Transition in the shear layer of a separation bubble can lead to rapid reattachment. This suggests using control mechanisms to trigger and enhance early transition. Gad-el-Hak provides a review of various techniques for flow control in general and Volino discusses recent studies on separation control under low-pressure-turbine conditions utilizing passive as well as active devices. As pointed out by Volino, passive devices optimized for separation control at low Reynolds numbers tend to increase losses at high Reynolds numbers, Active devices have the attractive feature that they can be utilized only in operational regimes where they are needed and when turned off would not affect the flow. The focus in the present paper is an experimental Separation is induced on a flat plate installed in a closed-circuit wind tunnel by a shaped insert on the opposite wall. The flow conditions represent flow over the suction surface of a modem low-pressure-turbine airfoil ('Pak-B'). The Reynolds number, based on wetted plate length and nominal exit velocity, is varied from 50,000 to 300,000, covering cruise to takeoff conditions. Low (0.2%) and high (2.5%) Gee-stream turbulence intensities are set using passive grids. A spanwise-oriented phased-plasma-array actuator, fabricated on a printed circuit board, is surface- flush-mounted upstream of the separation point and can provide forcing in a wide frequency range. Static surface pressure measurements and hot-wire anemometry of the base and controlled flows are performed and indicate that the glow-discharge plasma actuator is an effective device for separation control. of active separation control using glow discharge plasma actuators.

Hultgren, Lennart S.↗

NASA NTRS1947

On the Theory of the Unsteady Motion of an Airfoil

The paper presents a systematical analysis of the problem of the determination of the unsteady motion about an airfoil moving in an infinite fluid that contains a system of vortices and the determination of the hydrodynamical forces acting on the airfoil. The hydrodynamical problem is reduced to the determination of the function f (xi) which transforms conformally the external region of the airfoil into the interior of a circle. The proposed methods of determining the irrotational motion of a fluid that is produced by any motion of the airfoil are especially simple and effective if the function f (xi) is rational. As an example the flow is determined for the case of an arbitrary motion of an airfoil of the Joukowsky type. The formulas obtained for the determination of the hydrodynamical forces by means of contour integration are similar to those given by S. Chaplygin. These formulas are used to determine the force acting on the airfoil in the cases where the unsteady motion is potential throughout and the circulation about the airfoil is constant and also when the fluid contains a system of vortices. A full discussion is given of the concept of virtual masses together with practical formulas for computing the virtual mass coefficients.

Sedov, L. I.↗

NASA NTRS1946

Theoretical and Experimental Data for a Number of NACA 6A-Series Airfoil Sections

The NACA 6A-series airfoil sections were designed to eliminate the trailing-edge cusp which is characteristic of the NACA 6-series sections. Theoretical data are presented for NACA 6A-series basic thickness forms having the position of minimum pressure at 30-, 40-, and 50-percent chord and with thickness ratios varying from 6 percent to 15 percent. Also presented are data for a mean line designed to maintain straight sides on the cambered sections. The experimental results of a two dimensional wind tunnel investigation of the aerodynamic characteristics of five NACA 64A-series airfoil sections and two NACA 63A-series airfoil sections are presented. An analysis of these results, which were obtained at Reynolds numbers of 3 x 10(exp 6), 6 x 10(exp 6), and 9 x 10(exp 6), indicates that the section minimum drag and maximum lift characteristics of comparable NACA 6-series and 6A-series airfoil sections are essentially the same. The quarter-chord pitching-moment coefficients and angles of zero lift of NACA 6A-series airfoil sections are slightly more negative than those of corresponding NACA 6-series airfoil sections. The position of the aerodynamic center and the lift-curve slope of smooth NACA 6-series sections. The addition of standard leading-edge roughness causes the lift-curve slope of the newer sections to decrease with increasing airfoil thickness ratio.

Loftin, Laurence K., Jr.↗

NASA NTRS2001

Feasibility of Actively Cooled Silicon Nitride Airfoil for Turbine Applications Demonstrated

Nickel-base superalloys currently limit gas turbine engine performance. Active cooling has extended the temperature range of service of nickel-base superalloys in current gas turbine engines, but the margin for further improvement appears modest. Therefore, significant advancements in materials technology are needed to raise turbine inlet temperatures above 2400 F to increase engine specific thrust and operating efficiency. Because of their low density and high-temperature strength and thermal conductivity, in situ toughened silicon nitride ceramics have received a great deal of attention for cooled structures. However, the high processing costs and low impact resistance of silicon nitride ceramics have proven to be major obstacles for widespread applications. Advanced rapid prototyping technology in combination with conventional gel casting and sintering can reduce high processing costs and may offer an affordable manufacturing approach. Researchers at the NASA Glenn Research Center, in cooperation with a local university and an aerospace company, are developing actively cooled and functionally graded ceramic structures. The objective of this program is to develop cost-effective manufacturing technology and experimental and analytical capabilities for environmentally stable, aerodynamically efficient, foreign-object-damage-resistant, in situ toughened silicon nitride turbine nozzle vanes, and to test these vanes under simulated engine conditions. Starting with computer aided design (CAD) files of an airfoil and a flat plate with internal cooling passages, the permanent and removable mold components for gel casting ceramic slips were made by stereolithography and Sanders machines, respectively. The gel-cast part was dried and sintered to final shape. Several in situ toughened silicon nitride generic airfoils with internal cooling passages have been fabricated. The uncoated and thermal barrier coated airfoils and flat plates were burner rig tested for 30 min without and with air cooling. Without cooling, the surface temperature of the flat plate reached approximately 2350 F. Starting with computer aided design (CAD) files of an airfoil and a flat plate with internal cooling passages, the permanent and removable mold components for gel casting ceramic slips were made by stereolithography and Sanders machines, respectively. The gel-cast part was dried and sintered to final shape. Several in situ toughened silicon nitride generic airfoils with internal cooling passages have been fabricated. The uncoated and thermal barrier coated airfoils and flat plates were burner rig tested for 30 min without and with air cooling. Without cooling, the surface temperature of the flat plate reached approximately 2350 F. With cooling, the surface temperature decreased to approximately 1910 F--a drop of approximately 440 F. This preliminary study demonstrates that a near-net-shape silicon nitride airfoil can be fabricated and that silicon nitride can sustain severe thermal shock and the thermal gradients induced by cooling and, thus, is a viable candidate for cooled components.

Bhatt, Ramakrishna T.↗

NASA NTRS2009

Characterization of the Effect of Wing Surface Instrumentation on UAV Airfoil Performance

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.

Ratnayake, Nalin A.↗

NASA NTRS1939

Tests of Airfoils Designed to Delay the Compressibility Burble

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.

Stack, John↗