Showing papers on "Airfoil published in 1975"

Starting vortex, separation bubbles and stall: a numerical study of laminar unsteady flow around an airfoil

Abstract: The stalling characteristics of an airfoil in a laminar viscous incompressible fluid are investigated. The governing equations in terms of the vorticity and stream function are solved using an implicit finite-difference scheme and point successive relaxation procedure. The development of the impulsively started flow, the initial generation of circulation, and the behavior of the forces at large times are studied. Following the impulsive start, the lift is at first very large and then it rapidly drops. The subsequent growth of circulation and lift is associated with the starting vortex. After incipient separation, the lift increases owing to enlargement of the separation bubble and intensification of the flow rotation in it. The extension of this bubble beyond the trailing edge causes it to rupture and brings about the stalling characteristics of the airfoil. Subsequently, new bubbles are formed near the upper surface of the airfoil and are swept away. The behavior of the lift acting on the airfoil is explained in terms of the strength and sense of these bubbles.

Low-speed aerodynamic characteristics of a 13-percent-thick airfoil section designed for general aviation applications

Abstract: Wind-tunnel tests were conducted to determine the low-speed section characteristics of a 13 percent-thick airfoil designed for general aviation applications. The results were compared with NACA 12 percent-thick sections and with the 17 percent-thick NASA airfoil. The tests were conducted ovar a Mach number range from 0.10 to 0.35. Chord Reynolds numbers varied from about 2,000,000 to 9,000,000.

Circulation Control for High Lift and Drag Generation on STOL Aircraft

Abstract: The concept of circulation control by tangential upper surface blowing over a circular trailing edge has been investigated for application to fixed wing STOL aircraft. Experimental investigations on both two- and threedimensional airfoils employing nominal blowing have demonstrated lift gains double to triple that of the conventional flapped airfoil, and associated large increases in drag (which further serve to reduce landing velocities and distances). An additional two-dimensional investigation into the basic fluid mechanics of the concept has shown that jet Mach numbers considerably above choked produced no adverse effects on the mechanism of the trailing edge Coanda flow, but instead yielded additional lift gains. These results appear quite promising where high lift generation is desired for a STOL aircraft having a nominal amount of auxiliary bleed air available, but where substitution of increased pressure ratios can produce added jet velocity to obtain the required momentum (blowing) coefficient.

Aerodynamic properties of an insect wing section and a smooth aerofoil compared.

Abstract: IN this communication I report measurements of lift and drag forces acting on corrugated aerofoils. The chordal profile of the wings in many orders of insects is corrugated1 and I have used scale models of a wing section of this type found in the hover fly Syrphus balteatus (Diptera, Syrphidae). The experimental profile is shown in Fig. 1d. The life-size chord length is 2.9 mm, and occurs 3.7 mm distal to the wing-root hinge, in a total unilateral wing length of 11.5 mm. (There is no special reason either for selecting this particular section or for choosing S. balteatus.) The model profiles were 22 times life size (chord length = 6.4 cm) and were made of stainless steel 100 µm thick. They were built to uniform chordal profile, with spans of 7.17 cm or 3.39 cm, giving aspect ratios of 1.12 and 0.53, respectively—both very low. An attempt was made to model the thick leading edge vein of the real wing by constructing the model leading edge from a stainless steel rod (diameter 1.7 mm).

Unsteady Viscous Flow on Oscillating Airfoils

Abstract: Incompressible laminar and turbulent flows over flat plates and airfoils have been investigated numerically and experimentally in unsteady flow conditions. Important differences were found between laminar and turbulent flat plate flows over a wide range of oscillation frequencies. Also, the importance of unsteady effects on laminar boundary layers was found to diminish rapidly with increasing longitudinal pressure gradients, whereas turbulent separation on airfoils was significantly affected by oscillatory motion when the incidence approached the stall angle. The calculated hysteresis in turbulent separation followed in a qualitative sense the well-known trends of dynamic stall delay and reattachment. However, the numerical analysis failed to indicate some of the important features of dynamic stall observed in the present experiment and in previous studies.

The Effect of Reynolds Number and Laminar Separation on Axial Cascade Performance

Abstract: Testing over a range of Reynolds numbers was done for three NACA 65 Profiles in cascade. The testing was carried out in the VKI C-1 Low Speed Cascade Wind Tunnel; blade chord Reynolds number was varied from 250,000 to 40,000. A semiempirical theory is developed which will predict the behavior of the shear layer across a laminar separation bubble. The method is proposed for two-dimensional incompressible flow and is applicable down to short bubble bursting. The method can be used to predict the length of the laminar bubble, the bursting Reynolds number, and the development of the shear layer through the separated region. As such it is a practical method for calculating the profile losses of axial compressor and turbine cascades in the presence of laminar separation bubbles. It can also be used to predict the abrupt leading edge stall associated with thin airfoil sections. The predictions made by the method are compared with the available experimental data. The agreement could be considered good. The method was also used to predict regions of laminar separation in converging flows through axial compressor cascades (exterior to the corner vortices) with good results. For Reynolds numbers below bursting the semiempirical theory no longer applies. For this situation the performance of an axial compressor cascade can be computed using an empirical correlation proposed by the author. Comparison of performance prediction with experiment shows satisfactory agreement. Finally, a tentative correlation, based on the NACA Diffusion Factor, is presented that allows a rapid estimation of the bursting Reynolds number of an axial compressor cascade.

On the viscous flow about the trailing edge of a rapidly oscillating plate

Abstract: The incompressible laminar flow in the neighbourhood of the trailing edge of an aerofoil undergoing sinusoidal oscillations of high frequency and low amplitude in a uniform stream is described in the limit as the Reynolds number R tends to infinity. The aerofoil is replaced by a flat plate on the assumption that leadingedge stall does not take place. It is shown that, for oscillations of non-dimensional frequency , a rational description of the flow at the trailing edge is based on a subdivision of the boundary layer above the plate into five distinct regions. Asymptotic analytic solutions are found in four of these, whilst in the fifth a linearized solution yields an estimate for the viscous correction to the circulation determined by the Kutta condition.

Coolable turbine airfoil

Abstract: An airfoil cooling system for use in a gas turbine engine having high turbine inlet temperatures is disclosed. Various construction details designed to prevent thermal deterioration are developed. The system is built around impingement, film, and convective cooling techniques which are combined to limit the temperature of the airfoil material and to reduce thermal gradients within the component.

Isolated Airfoil-Tip Vortex Interaction Noise

Abstract: An experimental investigation was conducted in the UARL Acoustic Research Tunnel to define the noise characteristics associated with the interaction of a stationary tip vortex and a downstream stationary airfoil. This model test geometry simulated, in its simplest form, the tip vortexblade interaction which occurs on single rotor helicopters during hover. For moderate to high lift test conditions, the vortex-airfoil interaction was found to cause local blade stall with an attendant increase in the blade far-field noise. These results indicated that this interaction may be an important source of helicopter broadband noise during hover. Cross-correlation measurements conducted amongst surface-mounted and far-field microphones demonstrated that the operative noise mechanism was "trailing edge noise" arising from the interaction of stall generated eddies with the airfoil trailing edge. This mechanism would be expected to be responsible for increased noise at stall conditions in other, nonrotary wing, applications.

Experimental Investigation of the High Velocity Coanda Wall Jet Applied to Bluff Trailing Edge Circulation Control Airfoils

Abstract: : A two-dimensional experimental investigation, intended to probe the mechanism for reduction in performance of circulation control elliptic airfoils in compressible flow, was conducted subsonically on a 20-percent-thick modified elliptic profile employing high Coanda wall jet velocities. The results include detailed pressure distributions (both normal and chordwise) and trailing edge shear stress measurements made with a hot film anemometer for a range of jet slot heights and jet total pressures corresponding to high subsonic, sonic, and supersonic jet velocities. Jet Mach numbers of almost 1.3 were found to have no adverse effects on the airfoil performance, and the degrading jet detachment phenomenon was never encountered. Significant differences in the jet flow field with and without an external free stream were noted, as was the deviation of the static pressure across the jet from a constant value as assumed in conventional boundary layer analysis. Airfoil lift performance was found to vary with slot height, and the detailed shear stress measurement enabled location of the jet separation point. Also discussed is the calibration and use of the hot film shear stress probe.

Unsteady Transonic Flows over an Airfoil

Abstract: Inviscid, transonic flows over the NACA 64A410 airfoil were calculated using a finite-difference analog of the Euler equations. Steady flow solutions were obtained for angles of attack of 0°, 2° and 4° at Mach 0.72. Unsteady calculations were made at the same Mach number and in the same general angle-of-attack range for a step change of angle of attack, a step change in pitching angular velocity, and for sinusoidal pitching oscillations about the midchord at reduced frequencies 0.2,1.0, and 5.0. (Reduced frequency, K^uC/U^.) A summary of the resulting surface pressures, shock locations, and the forces and moments is presented. The forces and moments obtained for the cases with harmonic oscillations were compared to the results from a linearized analysis which used, to approximate the indicial functions, the responses calculated for the step changes.

Test Techniques for High-Lift, Two-Dimensional Airfoils with Boundary Layer and Circulation Control for Application to Rotary Wing Aircraft.

Abstract: : Extensive testing experience with very high lift mono-element blown airfoils to be employed by rotary wing aircraft has necessitated the development of unconventional two-dimensional test techniques. Both experimental and analytical results are presented which should assist future investigators in conducting similar tests accurately. The primary problem of high lift two-dimensional testing, wall boundary layer separation due to severe adverse pressure gradients on the model, is discussed as are the serious errors introduced by this phenomenon. A detailed discussion is presented concerning the additional considerations which must be given to blown airfoil testing. Such additional test problems as wall and blockage corrections, Reynolds number effects, leading edge separation, and flow visualization are addressed.

Implicit approximate-factorization schemes for the low-frequency transonic equation

Abstract: Two- and three-level implicit finite-difference algorithms for the low-frequency transonic small disturbance-equation are constructed using approximate factorization techniques. The schemes are unconditionally stable for the model linear problem. For nonlinear mixed flows, the schemes maintain stability by the use of conservatively switched difference operators for which stability is maintained only if shock propagation is restricted to be less than one spatial grid point per time step. The shock-capturing properties of the schemes were studied for various shock motions that might be encountered in problems of engineering interest. Computed results for a model airfoil problem that produces a flow field similar to that about a helicopter rotor in forward flight show the development of a shock wave and its subsequent propagation upstream off the front of the airfoil.

Modern Developments in Transonic Flow

Abstract: A survey is given of transonic small disturbance theory. Basic equations, shock relations, similarity laves, lift and drag integrals are derived., The airfoil boundary value problem is formulated. Finite difference methods and computational algorithms are described. Results are compared with other calculation methods and experiments.

F100 fan stall flutter problem review and solution

Abstract: An experimental investigation was conducted to examine an airfoil durability problem in the first fan rotor of the F100 engine. This study incorporated laboratory and simulated engine flight tests, an empirical correlation of aeroelastic stability parameters from engine test data, and substantiation testing of the redesign. The results of this investigation's initial testing showed that rotor failure at high-flight Mach numbers and low altitudes was caused by torsional stall flutter instability. The results of the empirical correlation indicated that a design free of flutter required a decrease in both normalized incidence and reduced velocity. Further, the correlation indicated that the flutter was affected by inlet pressure, a heretofore undocumented phenomenon. The results of the substantiation testing confirmed that the redesign made the rotor flutter-free throughout the entire aircraft flight envelope. It was concluded that an improved stall flutter analysis was required to ensure stable fan and compressor rotor designs. It was further concluded that the effect of changes in inlet pressure level on rotor stability was, in part, the result of the accompanying changes in air density and steady-state aerodynamic loading.

Unsteady Kutta condition at high values of the reduced frequency parameter

Abstract: The unsteady Kutta condition is discussed in the light of some recent experimental measurements made near the trailing edge of a long flat plate and a 10C4 airfoil. The hierachy of disagreement from the theoretically predicted zero trailing edge loading caused by viscous instabilities is found to be acoustically correlated vortex shedding, natural vortex shedding, Tollmien-Schlichting waves, and, by implication, turbulent boundary-layer eddies. The region of significant chordwise disagreement scales with the wake perturbation wavelength of the corresponding instability. Coordinating the vorticity of the turbulent boundary layer shed from the profile airfoil with a transverse acoustic resonance produced a distinct disagreement of the Kutta condition at high reduced frequency parameters (\ = wc/U). In this case and for vortex shedding, the extrapolated loading coefficient at the trailing edge increased with the nondimensional acoustic amplitude. I. Introduction T^HE Kutta condition as applied in unsteady JL potential airfoil analyses is essentially an extension of steady theory. Kutta ! postulated that a value of circulation should be chosen in his steady potential model to avoid a velocity singularity at the sharp trailing edge of an airfoil. This condition can be established if the trailing edge is also the rear stagnation point. The resulting modeled flow pattern agrees with that observed in steady flow and also predicts the lift and its chordwise distribution well at low angles of attack. The theoretical consequences of this hypothesis are that the lift loading or chordwise vorticity jump approaches zero at the trailing edge. An alternative statement is that the surface velocities on either side of the airfoil approach a common value at the rear stagnation point. For rounded trailing edges, the position of the rear stagnation point is indeterminate , as there is no velocity singularity to be avoided and so fix its location. In this case and for the situation of real flows with viscosity, Taylor2 proposed the condition of zero net vorticity discharge to establish the steady lift value. Preston3 explained the deviation of the lift of an airfoil at low angles of incidence from the potential theory value as due to the profile alteration from the boundary-layer growth. His calculations incorporated Taylor's vorticity discharge condition. Various approximate steady lift calculation methods for the rounded trailing edge geometry have been proposed by Gostello 4 and others. These extend the upper and lower lift distributions, at a selected chordwise position, to the trailing edge to give zero loading and thereby remove the stagnation point indeterminacy. In the unsteady case there are all the previous theoretical difficulties and, in addition, the unsteady effects on the viscous boundary layer and the shed vorticity. The latter complicates the airfoil response, making it a function of the airfoil's vorticity history. However, same theoretical assumption for the Kutta condition, of no unsteady loading at the

Light aircraft lift, drag, and moment prediction: A review and analysis

Abstract: The historical development of analytical methods for predicting the lift, drag, and pitching moment of complete light aircraft configurations in cruising flight is reviewed. Theoretical methods, based in part on techniques described in the literature and in part on original work, are developed. These methods form the basis for understanding the computer programs given to: (1) compute the lift, drag, and moment of conventional airfoils, (2) extend these two-dimensional characteristics to three dimensions for moderate-to-high aspect ratio unswept wings, (3) plot complete configurations, (4) convert the fuselage geometric data to the correct input format, (5) compute the fuselage lift and drag, (6) compute the lift and moment of symmetrical airfoils to M = 1.0 by a simplified semi-empirical procedure, and (7) compute, in closed form, the pressure distribution over a prolate spheroid at alpha = 0. Comparisons of the predictions with experiment indicate excellent lift and drag agreement for conventional airfoils and wings. Limited comparisons of body-alone drag characteristics yield reasonable agreement. Also included are discussions for interference effects and techniques for summing the results above to obtain predictions for complete configurations.

Experimental Evaluation of Theories for Trailing Edge and Incidence Fluctuation Noise

Abstract: Tests were conducted to evaluate conflicting theories for trailing edge noise and for incidence fluctuation noise. A flat-plate airfoil with flush-mounted surface pressure transducers was tested in an anechoic wind tunnel at velocities of 31.5-177 m/sec (108-580 fps) and nominal 4 and 6°7o grid-generated turbulence levels. In one series of runs, the airfoil was faired into the tunnel nozzle and extended beyond the nozzle lip for studies of trailing edge noise without a leading edge and with flow on only one side. Such noise was found to vary with velocity to the fifth power and turbulence level squared, as predicted by Ffowcs Williams and Hall and by Chase. Power spectral density at high frequencies decayed approximately inversely with frequency to the 10/3 power, as predicted by Chase. The data were poorly predicted by Hayden's correlation. Additional tests were conducted with the airfoil mounted on the tunnel centerline, with flow on both sides and turbulence convected past the leading edge. Surface pressure spectra caused by incidence fluctuations were reasonably predicted by the theories of Filotas and Mugridge at high frequencies, but were more closely predicted by that of Filotas at low frequencies. Far-field spectra decayed more rapidly than was predicted by their theory within a compact source analysis when the product of Strouhal number and Mach number was larger than about 0.5. Hayden's correction for acoustic noncompactness, which may also describe a phase cancellation effect on lift force, brought Filotas' theory into agreement with far-field spectra.

Semi-buoyant aircraft

Abstract: The disclosed semi-buoyant lift-augmented aircraft, preferably of immense size, includes a fuselage of airfoil shape formed by a rigid geodesic type web framework enclosing buoyant cells pressurized to reinforce the framework. Wings are provided which impart aerodynamic lift to the aircraft, with the airfoil fuselage. The buoyancy is provided to counteract a major proportion of the great weight of the structure and to thus improve the aircraft's payload capability and range. These features maintain it airborne at reasonable speeds over a substantial altitude range.

Aerodynamic Design of High-Perf ormance Biplane Wings

Abstract: LTHOUGH biplanes were quite popular during the early days of aviation, they had virtually disappeared from service by the mid-1930's. The biplane seemed to be plagued by inherently high drag and low maximum lift coefficient. The purpose of the present study was to re-examine the reasons for the biplane's decline. It was found that for any practical biplane configuration, it is possible to design a wing system which has a much lower weight per unit area and which has essentially the same maximum useable lift coefficient as that of a comparable, well-designed, monoplane wing, in either clean or flapped configurations. Improvements in the design of fairings permit large drag reductions, relative to the earlier designs. Because of these improvements, it is possible to design a biplane whose performance is superior, for some applications, to that of a well-designed monoplane. These applications are those for which excellent low-speed maneuverability, good short-field performance, good loadcarrying ability, low cost, and rugged construction are of primary importance.! Content If a biplane wing system is to develop a high CLmax, both wings must stall nearly together. A method presented by Fuchs was used to identify those configurations for which a good stall match could be obtained. Each wing is idealized in this method as a single horseshoe vortex, with the bound part of each vortex located at the center of pressure. Satisfactory stall match was defined as a ratio of leading wing to trailing wing Coequal to unity at an assumed CLmax for the combination. Some results of this analysis are given in Fig. 1, and some nomenclature is defined in the sketch. The airfoil chord lines are parallel; results of this study showed no advantage to be gained from the use of decalage. The geometric stagger angle a for best stall match is quite insensitive to the gap/chord ratio, and it decreases as the aspect ratio increases. The best stall match always corresponds to small values of 0, the aerodynamic stagger angle. The chordwise variation in vertical velocity induced by one airfoil upon the other corresponds to a local curvature of the flow. Since within the framework of thin airfoil theory, this effect is identical to the effect of a change in the mean camber line, it will be referred to as induced camber. Data for NACA airfoils of the four- and five-digit series show increases in the maximum section lift coefficient Cfmax

Subsonic Two-Dimensional Wind Tunnel Investigations of the High Lift Capability of Circulation Control Wing Sections

Abstract: : Two series of Circulation Control Wing airfoil sections, formed by the conversion of the sharp trailing edge into a circular bluff surface with tangential upper surface blowing, were evaluated subsonically to determine their high lift characteristics as potential STOL wing sections. Parameters investigated which had noticeable effect on the blown airfoil performance include leading edge devices (type of device and degree of deflection), trailing edge configuration (radius, slot location, deflection, etc.), Reynolds number, airfoil incidence, momentum coefficient, slot height, and nozzle pressure ratio. Maximum lift coefficients roughly triple those of the flapped conventional sections were generated at incidence slightly less than the conventional stall angles and at blowing rates obtainable from bleed of state-of-the-art turbojet engines. An experimental investigation into the lift augmenting effects of pulsed unsteady blowing was conducted on a smaller radius trailing edge configuration. An additional investigation was conducted to determine the effects of spoilers or similar disturbances ahead of the jet exit. The results of the above investigations provide a data base for the prediction of the aerodynamic characteristics of aircraft employing Circulation Control trailing edges to increase their STOL capability. (Author)

Design of Low-Speed Airfoils by Numerical Optimization

Abstract: A practical procedure for the optimum design of low-speed airfoils is demonstrated. The procedure uses an optimization program based on a gradient algorithm coupled with an aerodynamic analysis program that uses a relaxation solution of the inviscid, full-potential equation. The analysis program is valid for both incompressible and compressible flow, thereby making optimum design of high-speed, shock-free airfoils possible. Results are presented for the following three constrained optimization problems at fixed angle of attack and Mach number: (1) adverse pressure-gradient minimization, (2) pitching-moment minimization; and (3) lift maximization. All three optimization problems were studied with various aerodynamic and geometric constraints.

Two-Dimensional Subsonic Wind Tunnel Evaluation of a 20-Percent-Thick Circulation Control Airfoil

Abstract: : A circulation control uncambered elliptic airfoil section with a thickness-to-chord ratio of 0.20 was tested subsonically to determine its aerodynamic characteristics. Lift coefficients up to 5 were produced at momentum coefficients, of 0.24. The initially high unblown drag coefficients, characteristic of bluff trailing edge airfoils, were greatly reduced at low values of momentum coefficient. It was therefore possible to produce equivalent lift-to-drag ratios in excess of 30 when Cl = 1.0. The ability to produce high lift coefficients essentially independent of angle of attack is indicated by the results of this investigation.

Development of a computer program to obtain ordinates for NACA 4-digit, 4-digit modified, 5-digit, and 16 series airfoils

Abstract: A computer program developed to calculate the ordinates and surface slopes of any thickness, symmetrical or cambered NACA airfoil of the 4-digit, 4-digit modified, 5-digit, and 16-series airfoil families is presented. The program produces plots of the airfoil nondimensional ordinates and a punch card output of ordinates in the input format of a readily available program for determining the pressure distributions of arbitrary airfoils in subsonic potential viscous flow.

Aerodynamic device for improving the high speed running conditions of motorcycles

Abstract: An aerodynamic device for improving the high speed running conditions of motorcycles, comprising an airfoil fitted to the motorcycle to produce a negative lifting action approximately in the vertical plane passing through the pivot axle of the front wheel.

Recent developments in dynamic stall

Abstract: Tests were performed on an airfoil in a low-turbulence subsonic wind tunnel at fixed incidence for different speeds of the wind, the airfoil being in translatory motion normal or parallel to the wind. The resulting aerodynamic forces and moment on the airfoil were measured along with local pressure and skin friction. When the incidence is small and the boundary layer attached, no unsteady effects were observed. However, for a large incidence (16 deg), in the case of boundary layer separation, strong unsteady effects are revealed when the boundary layer is reattached in periodic flow. Reattachment can be explained assuming that for the same value of free stream velocity, the gradient of the modulus of the periodic component of the relative airspeed at the upper surface of the airfoil in a direction normal to this surface is greater than the gradient of the relative air speed at the upper surface when the airfoil is not oscillating.

Helicopter rotor blade

Abstract: A helicopter main rotor has blades with an inboard airfoil having a high lift-drag ratio for efficient hovering and a "shock-free" outboard airfoil for high speed cruise. Each rotor blade has a structural section composed of metal spars and metal skins forming torque boxes accompanied by a nonstructural fairing and after body. Secured within the torque boxes and the after body are honeycomb cellular structures of a metallic and nonmetallic material, depending on the area wherein the cellular structure is secured. At a point along the blade radius, between 60% and 75%, the upper metal skin of the aft torque box changes from an outer contour surface of the blade to a flat surface parallel to and above the cord line. A fairing of fiberglass skin is secured over a nonmetallic cellular structure over the flat surface of the upper metal skin to complete the upper outer airfoil contour outboard of the 75% radius dimension. The lower outer airfoil contour is the metal skin of the aft torque box followed by a fiberglass skin joined at a trailing edge with the fiberglass upper surface.