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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 1,351 records · Page 75
All sourcesNASA NTRSDOE OSTI
NASA NTRS1986

Feasibility of generating an artificial burst in a turbulent boundary layer

Artificial bursts were generated in laminar and turbulent boundary layers. The burst-like events were produced by withdrawing near-wall fluid from two minute holes separated in the spanwise direction or by pitching a miniature delta wing that was flush-mounted to the wall. Either of these actions generated streamwise vorticity and a low-speed streak that resembled a naturally occurring one. The resulting sequence of events occurred at a given location and at controlled times, allowing detailed examination and comparison with natural, random bursts by means of flow visualization and fast-response probe measurement techniques.

Gad-El-hak, M.↗
NASA NTRS1985

An experimental investigation of apex fence flaps on delta wings

The effects on leading edge vortex formation produced by apex fences mounted on two delta wings were examined in a subsonic wind tunnel. Pressure data were collected on the upper surface of the test wings and oil flow and bubble flow visualizations were performed. The deflections studied were 74, 90 and 65 deg in one-sided and two-sided configurations. Full span and semi-span delta wing planforms were used. Relatively small, symmetrically deployed apex fences provided enhanced upper surface suction which, in some angle-of-attack situations, could lead to a pitch-up force. Asymmetrically-deployed fences showed promise for yaw control at high angles of attack, althouh the force effectiveness has yet to be quantified.

Vess, R. J.↗
NASA NTRS1985

Local heat-transfer measurements on a large scale-model turbine blade airfoil using a composite of a heater element and liquid crystals

Local heat transfer coefficients were experimentally mapped along the midchord of a five-time-size turbine blade airfoil in a static cascade operated at room temperature over a range of Reynolds numbers. The test surface consisted of a composite of commercially available materials: a mylar sheet with a layer of cholesteric liquid crystals, that change color with temperature, and a heater sheet made of a carbon-impregnated paper, that produces uniform heat flux. After the initial selection and calibration of the composite sheet, accurate, quantitative, and continuous heat transfer coefficients were mapped over the airfoil surface. The local heat transfer coefficients are presented for Reynolds numbers from 2.8 x 10 to the 5th power to 7.6 x 10 to the 5th power. Comparisons are made with analytical values of heat transfer coefficients obtained from the STANS boundary layer code. Also, a leading edge separation bubble was revealed by thermal and flow visualization.

Hippensteele, S. A.↗
NASA NTRS1984

Hovering model helicopter rotor testing

A relatively large test chamber, chosen to minimize room interference effects during helicopter hovering rotor tests, was found in initial operations to exhibit large, random variations of thrust, torque, and blade flapping. The source of these problems, unsteady recirculation of test chamber flow, was addressed by the addition of a 'flow straightener' that reduced these parameter variations to 1.5 percent of mean values. Rotor performance was then measured outside the chamber in calm, free air conditions and compared with chamber results obtained before and after modifications. Attention is given to wake flow visualization conducted in the test chamber, as well as outside.

Piziali, R. A.↗
NASA NTRS1986

Upwind Navier-Stokes solutions for separated periodic flows

The application of an upwind implicit approximate factorization Navier-Stokes algorithm to highly separated flow is described. Using both the thin-layer and complete forms of the Navier-Stokes equations, the low Reynolds number laminar flow around a circular cylinder with periodic shedding is solved. The effect of grid density, grid extent, and time step on the Strouhal number is shown. Results from both sets of equations agree within the experimental data band. Unsteady, laminar flow computations around inclined plates and separated airfoils are also described. Strouhal numbers agree to within 5 percent of experiments for inclined plates. Differences between the complete equations and the thin-layer approximation for separated periodic flows are discussed. Computations of an impulsively started circular cylinder and airfoil yield time-accurate flowfield shapes in good agreement with experimental flow visualizations. The turbulent computation of an airfoil at a high angle-of-attack is massively separated, but shows no evidence of periodicity.

Rumsey, C. L.↗
NASA NTRS1986

Unsteady pressure measurements on a biconvex airfoil in a transonic oscillating cascade

Flush-mounted dynamic pressure transducers were installed on the center airfoil of a transonic oscillating cascade to measure the unsteady aerodynamic response as nine airfols were simultaneously driven to provide 1.2 deg of pitching motion about the midchord. Initial tests were performed at an incidence and angle of 0 deg and A Mach number of 0.65 in order to obtain results in a shock-free compressible flowfield. Subsequent tests were performed at an incidence angle of 7 deg and Mach number of 0.8 in order to observe the surface pressures with an oscillating shock near the leading edge of the airfoil. Results are presented for interblade phase angles of 90 and -90 deg and at blade oscillatory frequencies of 200 and 500 Hz (semi-chord reduced frequencies up to about 0.5 at a Mach number of 0.8). Results from the zero-incidence cascade are compared with a clasical usnteady flat-plate analysis. Flow visualization results depicting the shock motion on the airfoils in the high-incidence cascade are discussed. The airfoil pressure data are tabulated.

Shaw, L. M.↗
NASA NTRS1986

Topological classification of flow separation on three-dimensional bodies

In the study of three-dimensional, separated flows, a topological approach to separation and a phenomenological approach have been considered. The topological approach is based on ideas of Poincare (1928), while the phenomenological approach distinguishes between 'open' and 'closed' separation as posed by Wang (1974). In the present paper, the topological description provided by Legendre (1956, 1965, 1982) is considered, taking into account its utilization by Tobak and Peake (1979, 1982). An attempt is made to provide a more complete framework for the description of separated as well as attached flows. Nomenclature is introduced to describe the topology of the skin-friction-line portraits on a three-dimensional body. The topological ideas are illustrated by the use of hypothetical oil-flow experiments supplemented with actual flow-visualization data when possible.

Chapman, G. T.↗
NASA NTRS1986

Natural laminar flow flight experiments on a turbine engine nacelle fairing

Flight experiments are being conducted with the objective to investigate the interactions between acoustic disturbances and laminar flow in the flight environment. In the experiments, the laminar boundary layer on the nacelles will be exposed to discrete and broadband external noises, and the effect of varying noise levels and frequencies on the stability of the laminar flow will be studied. The present paper provides an overview of the complete project and a status report on the results which have been obtained. The flight experiment is conducted with the aid of a modified research aircraft. The phase I flight tests are concerned with a quick and simple determination of natural laminar flow (NLF) behavior on an engine nacelle. Attention is given to instrumentation, the acoustic noise source, flow visualization, and the test results.

Obara, C. J.↗
NASA NTRS1986

Non-linear effects in finite amplitude wave propagation through ducts and nozzles

In this paper an extensive study of non-linear effects in finite amplitude wave propagation through ducts and nozzles is summarized. Some results from earlier studies are included to illustrate the non-linear effects on the transmission characteristics of duct and nozzle terminations. Investigaiations, both experimental and analytical, were carried out to determine the magnitudes of the effects for high intensity pulse propagation. The results derived from these investigations are presented in this paper. They include the effect of the sound intensity on the acoustic characteristics of duct and nozzle terminations, the extent of the non-linearities in the propagation of high intensity impulsive sound inside the duct and out into free field, the acoustic energy dissipation mechanism at a termination as shown by flow visualizations, and quantitative evaluations by experimental and analytical means of the influence of the intensity of a sound pulse on the dissipation of its acoustic power.

Salikuddin, M.↗
NASA NTRS1986

Experimental investigation of rotorcraft hub and shaft fairing drag reduction

A wind-tunnel test was conducted to obtain data on several rotorcraft hub and shaft fairing drag reduction configurations. Aerodynamic loads and moments were acquired for each test configuration. Limited wake pressure measurements and flow visualization (tuft) photographs were obtained for some configurations. All hub and shaft fairing configurations were tested on a 1/5-scale XH-59A model fuselage. Both coaxial and single rotor configurations were tested. All rotor assemblies were modeled with nonrotating hardware. The drag reduction methods tested included cambered elliptical hub fairings, several different shaft fairings, and strakes. Test data show that significant drag reductions can be attained with certain fairing configurations. The lowest drag values for the single rotor configurations were obtained for a cambered elliptical hub fairing with a large thickness airfoil shaft fairing. The lowest coaxial configuration drag values were obtained with cambered elliptical hub fairings and a long chord intermediate shaft fairing.

Young, L. A.↗
NASA NTRS1986

Calculation of asymmetric vortex separation on slender delta wings with a vortex-sheet model

An inviscid vortex-sheet model based on the slender wing theory is developed to examine asymmetric vortex separation at zero sideslip on delta wings. It is found that multiple asymmetric vortex configurations exist at a given angle of attack. Available data on rolling moment measurements and flow visualization are used for correlation to prove the concept.

Lan, C. E.↗
NASA NTRS1986

Flowfield survey over a 75 deg swept delta wing at an angle of attack of 20.5 deg

An experimental investigation of the flowfield over a 75 deg swept delta wing at an angle of attack of 20.5 deg has been conducted. The data include pitot pressure surveys and two types of flow visualization. Surface and flowfield visualization data were obtained at Reynolds number, Rn, ranging from 0.5 to 2.0 million in increments of 0.25 million. Detailed pitot pressure surveys were made at five longitudinal stations at Rn = 0.5, 1.0, and 1.5 million in both the primary and secondary vortices. The results indicate that Reynolds number has only a minor effect on the global structure of the flowfield in the Reynolds number range that was investigated. The boundary layer transitions from laminar to turbulent at the trailing edge of the wing at Rn = 1.0 x 10 to the 6th, and the transition moves forward to x/L = 0.4 at Rn = 2.0 x 10 to the 6th. The positions of the primary vortex cores are insensitive to Reynolds number in this range; however, the lateral position of the secondary vortex core moves outboard aft of the region where the boundary layer transitions from laminar to turbulent.

Kjelgaard, S. O.↗
NASA NTRS1987

Simulation of fluid flows during growth of organic crystals in microgravity

Several counter diffusion type crystal growth experiments were conducted in space. Improvements in crystal size and quality are attributed to reduced natural convection in the microgravity environment. One series of experiments called DMOS (Diffusive Mixing of Organic Solutions) was designed and conducted by researchers at the 3M Corporation and flown by NASA on the space shuttle. Since only limited information about the mixing process is available from the space experiments, a series of ground based experiments was conducted to further investigate the fluid dynamics within the DMOS crystal growth cell. Solutions with density differences in the range of 10 to the -7 to 10 to the -4 power g/cc were used to simulate microgravity conditions. The small density differences were obtained by mixing D2O and H2O. Methylene blue dye was used to enhance flow visualization. The extent of mixing was measured photometrically using the 662 nm absorbance peak of the dye. Results indicate that extensive mixing by natural convection can occur even under microgravity conditions. This is qualitatively consistent with results of a simple scaling analysis. Quantitave results are in close agreement with ongoing computational modeling analysis.

Roberts, Gary D.↗
NASA NTRS1987

Investigation of empennage buffeting

Theoretical methods of predicting aircraft buffeting are reviewed. For the buffeting due to leading-edge vortex breakdown, a method is developed to convert test data of mean square values of fluctuating normal force to buffeting vortex strength through an unsteady lifting-surface theory and unsteady suction analogy. The resulting buffeting vortex from the leading-edge extension of an F-18 configuration is used to generate a fluctuating flow field which produces unsteady pressure distribution on the vertical tails. The root mean square values of root bending moment on the vertical tails are calculated for a rigid configuration. Results from a flow visualization and hot films study in a water tunnel facility using a 1/48 scale model of an F-18 are included in an appendix. The results confirm that the LEX vortex is the dominant forcing function of fin buffet at high angles of attack.

Lan, C. Edward↗
NASA NTRS1987

Investigation of leading-edge flap performance on delta and double-delta wings at supersonic speeds

An investigation of the aerodynamic performance of leading-edge flaps on three clipped delta and three clipped double-delta wing planforms with aspect ratios of 1.75, 2.11, and 2.50 was conducted in the Langley Unitary Plan Wind Tunnel at Mach numbers of 1.60, 1.90, and 2.16. A primary set of fullspan leading-edge flaps with similar root and tip chords were investigated on each wing, and several alternate flap planforms were investigated on the aspect-ratio-1.75 wings. All leading-edge flap geometries were effective in reducing the drag at lifting conditions over the range of wing aspect ratios and Mach numbers tested. Application of a primary flap resulted in better flap performance with the double-delta planform than with the delta planform. The primary flap geometry generally yielded better performance than the alternate flap geometries tested. Trim drag due to flap-induced pitching moments was found to reduce the leading-edge flap performance more for the delta planform than for the double-delta planform. Flow-visualization techniques showed that leading-edge flap deflection reduces crossflow shock-induced separation effects. Finally, it was found that modified linear theory consistently predicts only the effects of leading-edge flap deflection as related to pitching moment and lift trends.

Covell, Peter F.↗
NASA NTRS1987

Investigation of dynamic ground effect

An experimental investigation of dynamic ground effect was conducted in the Univ. of Kansas wind tunnel using delta wings of 60, 70, 75 deg sweep; the XB-70 wing; and the F-104A wing. Both static and dynamic tests were made. Test data were compared to other test data, including dynamic flight test data of the XB-70 and F-104A. Limited flow visualization test were conducted. A significant dynamic effect was found for highly swept delta wings.

Chang, Ray Chung↗
NASA NTRS1987

An experimental low Reynolds number comparison of a Wortmann FX67-K170 airfoil, a NACA 0012 airfoil and a NACA 64-210 airfoil in simulated heavy rain

Wind tunnel experiments were conducted on Wortmann FX67-K170, NACA 0012, and NACA 64-210 airfoils at rain rates of 1000 mm/hr and Reynolds numbers of 310,000 to compare the aerodynamic performance degradation of the airfoils and to attempt to identify the various mechanisms which affect performance in heavy rain conditions. Lift and drag were measured in dry and wet conditions, a variety of flow visualization techniques were employed, and a computational code which predicted airfoil boundary layer behavior was used. At low angles of attack, the lift degradation in wet conditions varied significantly between the airfoils. The Wortmann section had the greatest overall lift degradation and the NACA 64-210 airfoil had the smallest. At high angles of attack, the NACA 64-210 and 0012 airfoils had improved aerodynamic performance in rain conditions due to an apparent reduction of the boundry layer separation. Performance degradation in heavy rain for all three airfoils at low angles of attack could be emulated by forced boundary layer transition near the leading edge. The secondary effect occurs at time scales consistent with top surface water runback times. The runback layer is thought to effectively alter the airfoil geometry. The severity of the performance degradation for the airfoils varied. The relative differences appeared to be related to the susceptibility of each airfoil to premature boundary layer transition.

Craig, Anthony P.↗
NASA NTRS1987

Measurement of the effect of manufacturing deviations on natural laminar flow for a single engine general aviation airplane

Renewed interest in natural laminar flow (NLF) had rekindled designer concern that manufacuring deviations may destroy the effectiveness of NLF for an operational aircraft. Experiments are summarized that attemtped to measure total drag changes associated with three different wing surface conditions on an aircraft typical of current general aviation high performance singles. The speed power technique was first used in an attempt to quantify the changes in total drag. Predicted and measured boundary layer transition locations for three different wing surface conditions were also compared, using two different forms of flow visualization. The three flight test phases included: assessment of an unpainted airframe, flight tests of the same aircraft after painstakingly filling and sanding the wings to design contours, and similar measurement after this aricraft was painted. In each flight phase, transition locations were monitored using with sublimating chemicals or pigmented oil. Two-dimensional drag coefficients were estimated using the Eppler-Somers code and measured with a wake rake in a method very similar to Jones' pitot traverse method. The net change in two-dimensional drag coefficient was approximately 20 counts between the unpainted aircraft and the hand-smoothed aircraft for typical cruise flight conditions.

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