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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 451 records · Page 25
All sourcesNASA NTRSDOE OSTI
NASA NTRS1988

Technology for pressure-instrumented thin airfoil models

A novel method of airfoil model construction was developed. This Laminated Sheet technique uses 0.8 mm thick sheets of A286 containing a network of pre-formed channels which are vacuum brazed together to form the airfoil. A 6.25 percent model of the X29A canard, which has a 5 percent thick section, was built using this technique. The model contained a total of 96 pressure orifices, 56 in three chordwise rows on the upper surface and 37 in three similar rows on the lower surface. It was tested in the NASA Langley 0.3 m Transonic Cryogenic Tunnel. Unique aerodynamic data was obtained over the full range of temperature and pressure. Part of the data was at transonic Mach numbers and flight Reynolds number. A larger two dimensional model of the NACA 64a-105 airfoil section was also fabricated. Scale up presented some problems, but a testable airfoil was fabricated.

Wigley, David A.↗
NASA NTRS1986

Transition and separation control on a low-Reynolds number airfoil

The major problem associated with the aerodynamic performance of airfoils at low Reynolds numbers is the presence of extensive laminar boundary-layer separation resulting in a large increase in presssure drag and a decrease in lift. The rapid deterioration in airfoil characteristics can be largely eliminated by artificially controlling the flow through the introduction of suitable disturbances in the boundary layer such that transition occurs ahead of the anticipated laminar separation. This paper presents the results of wind-tunnel tests conducted on a 10-cm model of LRN (1)-1007 airfoil with passive (roughness trips) and active (acoustic excitation) controls to trigger transition and suppress separation. Significant improvements in the aerodynamic characteristics of the airfoil were observed. Results of this study for a chord Reynolds number range of 40,000 to 250,000 are presented in this paper.

Mangalam, S. M.↗
NASA NTRS1988

Finite volume computation of unsteady inviscid rotational transonic flows past airfoils in rigid body motion

Unsteady inviscid transonic flow over airfoils in arbitrary rigid body motion is analyzed numerically by solving the two-dimensional unsteady Euler equations in integral form using a finite volume scheme. The solution procedure is based on an explicit Runge-Kutta time-stepping scheme wherein the spatial terms are central-differenced and a combination of second- and fourth-differences in the flow variables are used to form the numerical dissipation terms to stabilize the scheme. Unsteady calculations are started from converged steady-state solutions as initial conditions. Nonreflective boundary conditions are imposed on the far-field boundaries. Results are presented and, where possible, validated against available numerical and experimental data for airfoils subjected to a step change in angle of attack, airfoils oscillating and plunging in transonic flow, and airfoils immersed in a time-varying free stream.

Damodaran, Murali↗
NASA NTRS1988

The calculation of flow over iced airfoils

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.

Cebeci, Tuncer↗
NASA NTRS1988

A natural low frequency oscillation in the wake of an airfoil near stalling conditions

An unusually low frequency oscillation in the flow over an airfoil was explored experimentally. Wind tunnel measurements were carried out with a two dimensional airfoil model at a chord Reynolds number of 100,000. During deep stall the usual bluff-body shedding occurred at a Strouhal number. But at the onset of stall a low frequency periodic oscillation occurred, the corresponding Strouhal number being an order of magnitude lower. The phenomenon occurred in relatively unclean flow when the freestream turbulence was raised to 0.4 percent, but did not in the cleaner flow with turbulence intensity of 0.1 percent. It could also be produced by certain high frequency acoustic excitation. Details of the flow field are compared between a case of low frequency oscillation at alpha = 15 deg and a case of bluff-body shedding at alpha = 22.5 deg. The origin of the low frequency oscillation traces to the upper surface of the airfoil and is seemingly associated with the periodic formation and breakdown of a large separation bubble. The intense flow fluctuations impart significant unsteady forces to the airfoil but diminish rapidly within a distance of one chord from the trailing edge.

Zaman, K. B. M. Q.↗
NASA NTRS1988

Measurement of local convective heat transfer coefficients from a smooth and roughened NACA-0012 airfoil - Flight test data

Wind tunnels typically have higher free stream turbulence levels than are found in flight. Turbulence intensity was measured to be 0.5 percent in the NASA Lewis Icing Research Tunnel (IRT) with the cloud making sprays off and around 2 percent with cloud making equipment on. Turbulence intensity for flight conditions was found to be too low to make meaningful measurements for smooth air. This difference between free stream and wind tunnel conditions has raised questions as to the validity of results obtained in the IRT. One objective of these tests was to determine the effect of free stream turbulence on convective heat transfer for the NASA Lewis LEWICE ice growth prediction code. These tests provide in-flight heat transfer data for a NASA-0012 airfoil with a 533 cm chord. Future tests will measure heat transfer data from the same airfoil in the Lewis Icing Research Tunnel. Roughness was obtained by the attachment of small, 2 mm diameter hemispheres of uniform size to the airfoil in three different patterns. Heat transfer measurements were recorded in flight on the NASA Lewis Twin Otter Icing Research Aircraft. Measurements were taken for the smooth and roughened surfaces at various aircraft speeds and angles of attack up to four degrees. Results are presented as Frossling number versus position on the airfoil for various roughnesses and angles of attack.

Van Fossen, G. James↗
NASA NTRS1988

Experimental results for the Eppler 387 airfoil at low Reynolds numbers in the Langley low-turbulence pressure tunnel

Experimental results were obtained for an Eppler 387 airfoil in the Langley Low Turbulence Pressure Tunnel. The tests were conducted over a Mach number range from 0.03 to 0.13 and a chord Reynolds number range for 60,000 to 460,000. Lift and pitching moment data were obtained from airfoil surface pressure measurements and drag data for wake surveys. Oil flow visualization was used to determine laminar separation and turbulent reattachment locations. Comparisons of these results with data on the Eppler 387 airfoil from two other facilities as well as the Eppler airfoil code are included.

Mcghee, Robert J.↗
NASA NTRS1988

Effects of a ground vortex on the aerodynamics of an airfoil

An experimental investigation was carried out to study the aerodynamics of an airfoil with a rectangular jet exiting from its lower surface at fifty percent of the chord. The airfoil was tested with and without the influence of a ground plane. Surface static pressures were measured on the airfoil at jet to free stream velocity ratios ranging from 0 to 9. From these pressures, the variation of C sub L with velocity ratio was easily determined. The measurements indicated significant positive and negative pressure regions on the lower surface of the airfoil ahead of and after the nozzle exit respectively. The presence of a ground plane enhanced these pressure regions at low velocity ratios, but at a particular ratio for each plane location, a recirculation zone or vortex formed ahead of the jet resulting in decreased pressures and a drop in C sub L.

Krothapalli, A.↗
NASA NTRS1988

Two-dimensional aerodynamic characteristics of the OLS/TAAT airfoil

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.

Watts, Michael E.↗
NASA NTRS1988

High Lift, Low Pitching Moment Airfoils

Two families of airfoil sections which can be used for helicopter/rotorcraft rotor blades or aircraft propellers of a particular shape are prepared. An airfoil of either family is one which could be produced by the combination of a camber line and a thickness distribution or a thickness distribution which is scaled from these. An airfoil of either family has a unique and improved aerodynamic performance. The airfoils of either family are intended for use as inboard sections of a helicopter rotor blade or an aircraft propeller.

Noonan, Kevin W.↗
NASA NTRS1988

Wind tunnel-sidewall-boundary-layer effects in transonic airfoil testing-some correctable, but some not

The need to correct transonic airfoil wind tunnel test data for the influence of the tunnel sidewall boundary layers, in addition to the wall accepted corrections for the analytical investigation was carried out in order to evaluate sidewall boundary layer effects on transonic airfoil characteristics, and to validate proposed correction and the limit to their applications. This investigation involved testing of modern airfoil configurations in two different transonic airfoil test facilities, the 15 x 60 inch two-dimensional insert of the National Aeronautical Establishment (NAE) 5 foot tunnel in Ottawa, Canada, and the two-dimensional test section of the NASA Langley 0.3 m Transonic Cryogenic Tunnel (TCT). Results presented included effects of variations in sidewall-boundary layer bleed in both facilities, different sidewall boundary layer correction procedures, tunnel-to tunnel comparisons of correcte results, and flow conditions with and without separation.

Lynch, F. T.↗
NASA NTRS1989

Airfoil self-noise and prediction

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.

Brooks, Thomas F.↗
NASA NTRS1989

The structure of separated flow regions occurring near the leading edge of airfoils - including transition

Laser Doppler Velocimeter data, static pressure data, and smoke flow visualization data was obtained and analyzed to correlate with separation bubble data. The Eppler 387 airfoil was focused on at a chord Reynolds number of 100,000 and an angle of attack of 2 deg. Additional data was also obtained from the NACA 663-018 airfoil at a chord Reynolds number of 160,000 and an angle of attack of 12 deg. The structure and behavior of the transition separation bubble was documented along with the redeveloping boundary layer after reattachment over an airfoil at low Reynolds numbers. The understanding of the complex flow phenomena was examined so that analytic methods for predicting their formation and development can be improved. These analytic techniques have applications in the design and performance prediction of airfoils operating in the low Reynolds number flight regime.

Source record↗
NASA NTRS1987

Results of LFC experiment on slotted swept supercritical airfoil in Langley's 8-foot transonic pressure tunnel

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.

Brooks, Cuyler W., Jr.↗
NASA NTRS1987

Wind tunnel results of the high-speed NLF(1)-0213 airfoil

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.

Sewall, William G.↗
NASA NTRS1987

The design of an airfoil for a high-altitude, long-endurance remotely piloted vehicle

Airfoil design efforts are studied. The importance of integrating airfoil and aircraft designs was demonstrated. Realistic airfoil data was provided to aid future high altitude, long endurance aircraft preliminary design. Test cases were developed for further validation of the Eppler program. Boundary layer, not pressure distribution or shape, was designed. Substantial improvement was achieved in vehicle performance through mission specific airfoil designed utilizing the multipoint capability of the Eppler program.

Maughmer, Mark D.↗
NASA NTRS1989

Convective heat transfer measurements from a NACA 0012 airfoil in flight and in the NASA Lewis Icing Research Tunnel

Local heat transfer coefficients were measured on a smooth and roughened NACA 0012 airfoil. Heat transfer measurements on the 0.533 m chord airfoil were made both in flight on the NASA Lewis Twin Otter Icing Research Aircraft and in the NASA Lewis Icing Research Tunnel (IRT). Roughness was obtained by the attachment of uniform 2 mm diameter hemispheres to the airfoil surface in 4 distinct patterns. Flight data were taken for the smooth and roughened airfoil at various Reynolds numbers based on chord in the range 1.24 to 2.50 x 10(exp 6) and at various angles of attack up to 4 deg. During these flight tests, the free stream velocity turbulence intensity was found to be very low (less than 0.1 percent). Wind tunnel data were acquired in the Reynolds number range 1.20 to 4.25 x 10(exp 6) and at angles of attack from -4 to 8 deg. The turbulence intensity in the IRT was 0.5 to 0.7 percent with the cloud generating sprays off. A direct comparison was made between the results obtained in flight and in the IRT. The higher level of turbulence in the IRT vs. flight had little effect on the heat transfer for the lower Reynolds numbers but caused a moderate increase in heat transfer at the high Reynolds numbers. Roughness generally increased the heat transfer.

Poinsatte, Philip E.↗
NASA NTRS1990

Experimental measurements of the laminar separation bubble on an Eppler 387 airfoil at low Reynolds numbers

An experimental investigation was conducted to measure the flow velocity in the boundary layer of an Eppler 387 airfoil. In particular, the laminar separation bubble that this airfoil exhibits at low Reynolds numbers was the focus. Single component laser Doppler velocimetry data were obtained at a Reynolds number of 100,000 at an angle of attack of 2.0 degree. Static Pressure and flow visualization data for the Eppler 387 airfoil were also obtained. The difficulty in obtaining accurate experimental measurements at low Reynolds numbers is addressed. Laser Doppler velocimetry boundary layer data for the NACA 663-018 airfoil at a Reynolds number of 160,000 and angle of attack of 12 degree is also presented.

Cole, Gregory M.↗