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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 289 records · Page 16
All sourcesNASA NTRSDOE OSTI
NASA NTRS1985

Separation bubble behavior for a NACA 23009 airfoil

Results are reported of experiments with smoke-wire flow visualization applied to characterizing the conditions conducive to formation of a separation bubble on an airfoil in low Re flow. An airfoil was used which spanned the test channel of the wind tunnel designed for low turbulence and noise. A wire coated with mineral oil was stretched across the channel upstream of the airfoil and heated. The smoke it gave off was flash-illuminated and photographed with ASA 3000 film. Extending the airfoil over the entire section yielded photographs that gave an essentially two-dimensional perspective. Angles of attack from 0-10 deg were investigated at Re of 33,000, 66,000, 100,000 and 133,000. Sample photographs illustrate the usefulness of the technique and apparatus in the flow regimes examined, particularly when used in conjunction with pressure tap data from the airfoil surface.

Fisher, S. S.↗
NASA NTRS1986

The response of airfoils to periodic disturbances - The unsteady Kutta condition

Two cases of unsteady flowfields over a NACA 0012 airfoil at an angle of attack are examined. The first is the classical pitching motion about the airfoil's quarter chord. The second is the flow over a fixed airfoil immersed in the wake of the pitching airfoil. Large reduced frequencies are examined. Measurements were obtained in a water tunnel by laser Doppler velocimetry. Ensemble-averaged velocity measurements were obtained in the vicinity of the trailing edges of both the pitching and the fixed airfoil. The results indicate that the classical unsteady Kutta condition is clearly not valid. An extension of this condition earlier proposed by Giesing and Maskell is examined and some evidence is provided for its support.

Poling, D. R.↗
NASA NTRS1986

Influence of airfoil camber on convected gust interaction noise

This paper investigates the effect of airfoil steady loading on the sound generated by the interaction of an airfoil with a convected disturbance. A previous theory, which included only the incidence angle contribution to the mean loading, is extended to include camber. The theory is based on a linearization of the Euler equations about a nonuniform, 0(1) Mach number subsonic mean flow. The discussion concentrates on the case of a slightly cambered airfoil at small incidence angle, interacting with a gust whose wavelength is short compared to the airfoil chord. The small parameter representing the amount of camber and incidence, and the large parameter representing the ratio of airfoil chord to disturbance wavelength, are utilized in a singular perturbation solution to the governing equations. Acoustic power calculations reveal that the amount of sound generated increases significantly with increased loading. More importantly, it is shown that the radiated acoustic power correlates very well with the strength of the mean flow around the leading edge.

Myers, M. R.↗
NASA NTRS1987

Materials for Advanced Turbine Engines (MATE). Project 4: Erosion resistant compressor airfoil coating

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.

Rashid, J. M.↗
NASA NTRS1987

An analysis of the crossover between local and massive separation on airfoils

Massive separation on airfoils operating at high Reynolds number is an important problem to the aerodynamicist, since its onset generally determines the limiting performance of an airfoil, and it can lead to serious problems related to aircraft control as well as turbomachinery operation. The phenomenon of crossover between local separation and massive separation on realistic airfoil geometries induced by airfoil thickness is investigated for low speed (incompressible) flow. The problem is studied both for the asymptotic limit of infinite Reynolds number using triple-deck theory, and for finite Reynolds number using interacting boundary-layer theory. Numerical results are presented which follow the evolution of the flow as it develops from a mildly separated state to one dominated by the massively separated flow structure as the thickness of the airfoil geometry is systematically increased. The effect of turbulence upon the evolution of the flow is considered, and the impact is significant, with the principal effect being the suppression of the onset of separation. Finally, the effect of surface suction and injection for boundary-layer control is considered. The approach which was developed provides a valuable tool for the analysis of boundary-layer separation up to and beyond stall. Another important conclusion is that interacting boundary-layer theory provides an efficient tool for the analysis of the effect of turbulence and boundary-layer control upon separated vicsous flow.

Barnett, M.↗
NASA NTRS1987

Transonic characteristics of a humped airfoil

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.

Tai, Tsze C.↗
NASA NTRS1987

Viscous Transonic Airfoil Workshop compendium of results

Results from the Viscous Transonic Airfoil Workshop held at the AIAA 25th Aerospace Sciences Meeting at Reno, NV in January 1987, are compared with each other and with experimental data. Test cases used in this workshop include attached and separated transonic flows for three different airfoils: the NACA 0012 airfoil, the RAE 2822 airfoil, and the Jones airfoil. A total of 23 sets of numerical results from 15 different author groups are included. The numerical methods used vary widely and include: 16 Navier-Stokes methods, 2 Euler/boundary-layer methods, and 5 full-potential/boundary-layer methods. The results indicate a high degree of sophistication among the numerical methods with generally good agreement between the various computed and experimental results for attached or moderately-separated cases. The agreement for cases with larger separation is only fair and suggests additional work is required in this area.

Holst, Terry L.↗
NASA NTRS1988

Divergent-Trailing-Edge Airfoil

Wedge concept integrated into basic airfoil design. Current airfoil design, DLBA 238, developed to produce airfoil with superior geometric characteristics and equivalent cruise aerodynamic characteristics. "Trailing-edge wedge" concept improved by integration into basic airfoil design for DLBA 238. Trailing-edge thickness and maximum thickness constrained to be same as baseline DLBA 032. Base drag-penalty of wedge avoided. Airfoil upper-surface geometry held fixed.

Henne, Preston A.↗
NASA NTRS1987

Wind tunnel studies of circulation control elliptical airfoils

Effects of blown jets on the lift and drag of cambered elliptical airfoils are described. Performance changes due to a splitter plate attached to the lower surface of an elliptical airfoil near the trailing edge with and without blowing are indicated. Lift and drag characteristics of airfoils with two blown jets are compared with airfoils with single blowing jets. Airfoil designs that vary the location of a second jet relative to a fixed jet are described.

Franke, M. E.↗
NASA NTRS1988

The NASA Langley Laminar-Flow-Control (LFC) experiment on a swept, supercritical airfoil: Design overview

A large-chord, swept, supercritical, laminar-flow-control (LFC) airfoil was designed and constructed and is currently undergoing tests in the Langley 8 ft Transonic Pressure Tunnel. The experiment was directed toward evaluating the compatibility of LFC and supercritical airfoils, validating prediction techniques, and generating a data base for future transport airfoil design as part of NASA's ongoing research program to significantly reduce drag and increase aircraft efficiency. Unique features of the airfoil included a high design Mach number with shock free flow and boundary layer control by suction. Special requirements for the experiment included modifications to the wind tunnel to achieve the necessary flow quality and contouring of the test section walls to simulate free air flow about a swept model at transonic speeds. Design of the airfoil with a slotted suction surface, the suction system, and modifications to the tunnel to meet test requirements are discussed.

Harris, Charles D.↗
NASA NTRS1986

High Reynolds number tests of a Douglas DLBA 032 airfoil in the Langley 0.3-meter transonic cryogenic tunnel

A wind-tunnel investigation of a Douglas advanced-technology airfoil was conducted in the Langley 0.3-Meter Transonic Cryogenic Tunnel (0.3-m TCT). The temperature was varied from 227 K (409 R) to 100 K (180 R) at pressures ranging from about 159 kPa (1.57 atm) to about 514 kPa (5.07 atm). Mach number was varied from 0.50 to 0.78. These variables provided a Reynolds number range (based on airfoil chord) from 6.0 to 30.0 x 10 to the 6th power. This investigation was specifically designed to: (1) test a Douglas airfoil from moderately low to flight-equivalent Reynolds numbers, and (2) evaluate sidewall-boundary-layer effects on transonic airfoil performance characteristics by a systematic variation of Mach number, Reynolds number, and sidewall-boundary-layer removal. Data are included which demonstrate the effects of fixing transition, Mach number, Reynolds number, and sidewall-boundary-layer removal on the aerodynamic characteristics of the airfoil. Also included are remarks on model design and model structural integrity.

Johnson, Charles B.↗
NASA NTRS1989

Supercritical flow past a symmetrical bicircular arc airfoil

A numerical scheme is developed for computing steady supercritical flow about symmetrical airfoils, applying it to an ellipse for zero angle of attack. An algorithmic description of this new scheme is presented. Application to a symmetrical bicircular arc airfoil is also proposed. The flow field before the shock is region 1. For transonic flow, singularity can be avoided by integrating the resulting ordinary differential equations away from the body. Region 2 contains the shock which will be located by shock fitting techniques. The shock divides region 2 into supersonic and subsonic regions and there is no singularity problem in this case. The Method of Lines is used in this region and it is advantageous to integrate the resulting ordinary differential equation along the body for shock fitting. Coaxial coordinates have to be used for the bicircular arc airfoil so that boundary values on the airfoil body can be taken with one direction of the coaxial coordinates fixed. To avoid taking boundary values at + or - infinity in the coaxial co-ordinary system, approximate analytical representation of the flow field near the tips of the airfoil is proposed.

Holt, Maurice↗
NASA NTRS1989

Numerical solution of periodic vortical flows about a thin airfoil

A numerical method is developed for computing periodic, three-dimensional, vortical flows around isolated airfoils. The unsteady velocity is split into a vortical component which is a known function of the upstream flow conditions and the Lagrangian coordinates of the mean flow, and an irrotational field whose potential satisfies a nonconstant-coefficient, inhomogeneous, convective wave equation. Solutions for thin airfoils at zero degrees incidence to the mean flow are presented in this paper. Using an elliptic coordinate transformation, the computational domain is transformed into a rectangle. The Sommerfeld radiation condition is applied to the unsteady pressure on the grid line corresponding to the far field boundary. The results are compared with a Possio solver, and it is shown that for maximum accuracy the grid should depend on both the Mach number and reduced frequency. Finally, in order to assess the range of validity of the classical thin airfoil approximation, results for airfoils with zero thickness are compared with results for airfoils with small thickness.

Scott, James R.↗
NASA NTRS1989

Prop-fan airfoil icing characteristics

An icing test program was carried out during the 1986-87 and 1987-88 icing seasons at the Fluidyne Icing Tunnel. The testing consisted of evaluating the icing characteristics and aerodynamic performance characteristics of two thin, two-dimensional airfoil sections. The two airfoils were a NACA Series 16-203 and a Hamilton Standard PF1-304, 3 and 4 percent thickness-to-chord ratio airfoils, respectively. These airfoils are representative of the thin airfoils currently being evaluated for Prop-Fan propulsion systems. The tests were conducted over a wide range of icing conditions, angles-of-attack, and Mach numbers (0.3 to 0.8).

Pike, J. A.↗
NASA NTRS1989

Optimization of natural laminar flow airfoils for high section lift-to-drag ratios in the lower Reynolds number range

Relatively thin natural-laminar-flow airfoils were arranged optimally for different design lift coefficients in the wing chord Reynolds number ranges of 200,000-600,00 and 0.875 x 10 to the 6th to 2 x 10 to the 6th. The 9.5 percent thick airfoil ASM-LRN-010, the 7.9 percent thick airfoil ASM-LRN-012, the 10.4 percent thick airfoil ASM-LRN-015, and the 8.2 percent thick airfoil ASM-LRN-017 were designed for high lift-to-drag ratios using Drela's design and analysis.

Pfenninger, Werner↗
NASA NTRS1989

Numerical solution of periodic vortical flows about a thin airfoil

A numerical method is developed for computing periodic, three-dimensional, vortical flows around isolated airfoils. The unsteady velocity is split into a vortical component which is a known function of the upstream flow conditions and the Lagrangian coordinates of the mean flow, and an irrotational field whose potential satisfies a nonconstant-coefficient, inhomogeneous, convective wave equation. Solutions for thin airfoils at zero degrees incidence to the mean flow are presented in this paper. Using an elliptic coordinate transformation, the computational domain is transformed into a rectangle. The Sommerfield radiation condition is applied to the unsteady pressure on the grid line corresponding to the far field boundary. The results are compared with a Possion solver, and it is shown that for maximum accuracy the grid should depend on both the Mach number and reduced frequency. Finally, in order to assess the range of validity of the classical thin airfoil approximation, results for airfoils with zero thickness are compared with results for airfoils with small thickness.

Scott, James R.↗
NASA NTRS1989

Experience with some repeat tests on the 9 inch chord CAST-10-2/DOA 2 airfoil model in the Langley 0.3-m TCT adaptive wall test section

A co-operative testing program is in progress between the Langley Research Center (NASA) and the National Aeronautical Establishment (NAE, Canada) to validate two different techniques of airfoil testing at transonic speeds. The procedure employed is to test the same airfoil model in the NAE two-dimensional tunnel and the Langley 0.3-m Transonic Cryogenic Tunnel (0.3-m TCT). The airfoil model used in testing was CAST-10-2/DOA-2 super-critical airfoil. The Langley 0.3-m TCT has a relatively small cross section of 13 in x 13 in, giving a (h/c) ratio of 1.44 for the same 9 in chord model. The approach employed in the 0.3-m TCT aims towards eliminating the wall effects by using active walls. The top and bottom walls are flexible. By changing the wall shapes during a test in an iterative manner, the wall interference effects are reduced. The method employed to change the wall shapes is the adaptive wall technique. The current test program provided an opportunity to validate the adaptive wall technique in the 0.3-m TCT. The relatively long chord airfoil represents a severe test case to test the efficacy of the adaptive wall technique under cryogenic conditions. The program also involved removal of side wall boundary-layer thus increasing the complexity of the wall adaptation technique. This paper deals with some salient results obtained regarding repeatability of test data and possible residual interference effects.

Murthy, A. V.↗
NASA NTRS1990

Numerical computation of viscous flow about unconventional airfoil shapes

A new two-dimensional computer code was developed to analyze the viscous flow around unconventional airfoils at various Mach numbers and angles of attack. The Navier-Stokes equations are solved using an implicit, upwind, finite-volume scheme. Both laminar and turbulent flows can be computed. A new nonequilibrium turbulence closure model was developed for computing turbulent flows. This two-layer eddy viscosity model was motivated by the success of the Johnson-King model in separated flow regions. The influence of history effects are described by an ordinary differential equation developed from the turbulent kinetic energy equation. The performance of the present code was evaluated by solving the flow around three airfoils using the Reynolds time-averaged Navier-Stokes equations. Excellent results were obtained for both attached and separated flows about the NACA 0012 airfoil, the RAE 2822 airfoil, and the Integrated Technology A 153W airfoil. Based on the comparison of the numerical solutions with the available experimental data, it is concluded that the present code in conjunction with the new nonequilibrium turbulence model gives excellent results.

Ahmed, S.↗