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At least 199 records · Page 11

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Edge-Based Viscous Method for Mixed-Element Node-Centered Finite-Volume Solvers

A novel, efficient, edge-based viscous (EBV) discretization method has been recently developed, implemented in a practical, unstructured-grid, node-centered, finite-volume flow solver, and applied to viscous-kernel computations that include evaluations of meanflow viscous fluxes, turbulence-model and chemistry-model diffusion terms, and the corresponding Jacobian contributions. Initially, the EBV method had been implemented for tetrahedral grids and demonstrated multifold acceleration of all viscous-kernel computations. This paper presents an extension of the EBV method for mixed-element grids. In addition to the primal edges of a given mixed-element grid, virtual edges are introduced to connect cell nodes that are not connected by a primal edge. The EBV method uses an efficient loop over all (primal and virtual) edges and features a compact discretization stencil based on the nearest neighbors. This study verifies the EBV method and assesses its efficiency on mixed-element grids by comparing the EBV solution accuracy and iterative convergence with those of well-established solutions obtained using a cell-based viscous (CBV) discretization method. The EBV solver’s memory footprint is optimized and often smaller than the memory footprint of the CBV solver. A multifold speedup is demonstrated for all viscous-kernel computations resulting in significant reduction of the time to solutions for several benchmark mixed-element-grid computations, including simulations of a flow around NASA’s juncture-flow model and a hypersonic, chemically reacting flow around a blunt body.

Edge-based viscous method↗

NASA NTRS1973

Externally blown flap trailing edge noise reduction by slot blowing: A preliminary study

Short takeoff and landing (STOL) aircraft using externally blown flaps (EBF) for lift augmentation develop considerable jet-flap interaction noise. A proposed method to reduce the EBF trailing edge noise is to locate a slot near the trailing edge of a flap through which low velocity secondary air is blown. Limited OASPL noise data were obtained from the interaction of the jet exhaust from a 5.08 cm diameter convergent nozzle with the trailing edge of a plate, and are presented for five slot configurations located near or at the trailing edge of the plate. Also presented are some significant jet trailing edge interaction data using a mixer nozzle with one of the slot configurations.

Mckinzie, D. J., Jr.↗

NASA NTRS1974

The dyadic diffraction coefficient for a curved edge

A compact dyadic diffraction coefficient for electromagnetic waves obliquely incident on a curved edge formed by perfectly conducting curved or plane surfaces is obtained. This diffraction coefficent remains valid in the transition regions adjacent to shadow and reflection boundaries, where the diffraction coefficients of Keller's original theory fail. The method is on Keller's method of the canonical problem, which in this case is the perfectly conducting wedge illuminated by plane, cylindrical, conical, and spherical waves. When the proper ray fixed coordinate system is introduced, the dyadic diffraction coefficient for the wedge is found to be the sum of only two dyads, and it is shown that this is also true for the dyadic diffraction coefficients of higher order edges. One dyad contains the acoustic soft diffraction coefficient; the other dyad contains the acoustic hard diffraction coefficient. The expressions for the acoustic wedge diffraction coefficients contain Fresnel integrals, which ensure that the total field is continuous at shadow and reflection boundaries. The diffraction coefficients have the same form for the different types of edge illumination; only the arguments of the Fresnel integrals are different. Since diffraction is a local phenomenon, and locally the curved edge structure is wedge shaped, this result is readily extended to the curved edge.

Kouyoumjian, R. G.↗

NASA NTRS1975

Flow visualization of vortices locked by spanwise blowing over wings featuring a unique leading and trailing-edge flap system

Flow visualization studies were conducted to qualitatively determine the effects of active generation and augmentation of vortex flow over wings by blowing a discrete jet in a spanwise direction in the channel formed by extension of upper surface leading- and trailing-edge flaps. Spanwise blowing from a reflection plane over a rectangular wing was found to generate and lock a dual corotating vortex system within the channel and, at sufficient blowing rates, cause the separated flow off the upper end of the leading-edge flap to reattach to the trailing-edge flap. Test parameters included wing angle of attack, jet momentum coefficient, leading- and trailing-edge flap deflection angle, and jet location above the wing surface. Effects due to removal of the leading- and trailing-edge flap were also investigated.

Erickson, G. E.↗

NASA NTRS1975

Temporal registration of multispectral digital satellite images using their edge images

An algorithm is described which will form an edge image by detecting the edges of features in a particular spectral band of a digital satellite image. It is capable also of forming composite multispectral edge images. In addition, an edge image correlation algorithm is presented which performs rapid automatic registration of the edge images and, consequently, the grey level images.

Nack, M. L.↗

NASA NTRS1979

Improved prediction of laminar leading edge separation

Research was conducted to provide a definite criterion for the prediction of the bubble burst on airfoils typical of those used for fighter wings. The approach taken was to correlate existing airfoil bubble burst data using various parameters at the laminar separation point. The method due to Weber was modified to provide a continuous analytic solution for the velocity distribution around the airfoil leading edge. Coupling the modified Weber method with the Stratford laminar separation prediction method leads to a universal chart giving the conditions at separation as a function of stagnation location and leading edge radius. Application of the combined method to available two-dimensional airfoil data resulted in an empirical criterion presenting the limiting local velocity gradient at separation as a function of the boundary layer momentum thickness at separation for bubble burst. The correlation leads as well to the qualitative explanation of two types of laminar stall: thin airfoil and leading edge. The validity of the correlation is demonstrated by predicting the lift coefficient and angle of attack for stall on airfoils with leading edge or trailing edge flaps.

Herring, R. N.↗

NASA NTRS1979

Influence of optimized leading-edge deflection and geometric anhedral on the low-speed aerodynamic characteristics of a low-aspect-ratio highly swept arrow-wing configuration

An investigation conducted in the Langley 7 by 10 foot tunnel to determine the influence of an optimized leading-edge deflection on the low speed aerodynamic performance of a configuration with a low aspect ratio, highly swept wing. The sensitivity of the lateral stability derivative to geometric anhedral was also studied. The optimized leading edge deflection was developed by aligning the leading edge with the incoming flow along the entire span. Owing to spanwise variation of unwash, the resulting optimized leading edge was a smooth, continuously warped surface for which the deflection varied from 16 deg at the side of body to 50 deg at the wing tip. For the particular configuration studied, levels of leading-edge suction on the order of 90 percent were achieved. The results of tests conducted to determine the sensitivity of the lateral stability derivative to geometric anhedral indicate values which are in reasonable agreement with estimates provided by simple vortex-lattice theories.

Coe, P. L., Jr.↗

NASA NTRS1979

Trailing edge noise data with comparison to theory

The noise emission generated by the passage of a turbulent airstream over the trailing edge of a semiinfinite plate was measured over a large range of airstream velocity and plate geometry. The experiment was designed to validate trailing-edge noise theories. The results show that the peak of a radiation pattern moves from an upstream to a downstream direction as the velocity increases. The measured radiation pattern of the noise was in agreement with that predicted by a recent fundamental theory for leading- and trailing-edge noise. Although large changes in the character of the turbulent flow near the trailing edge effect the level and spectra of trailing-edge noise, the shape of the pattern is still accurately predicted by this theory.

Olsen, W.↗

NASA NTRS1979

High-performance wings with significant leading-edge thrust at supersonic speeds

A new class of curved-leading-edge wings with which significant levels of leading-edge thrust may be achieved at moderate supersonic speeds is suggested. A recent analysis of the factors limiting such leading-edge thrust has led to a new method for the prediction of attainable leading-edge thrust from subsonic through supersonic speeds for wings of arbitrary planform. Recent supersonic tests of a new wing shape, which largely meets design criteria given by the new prediction method, give evidence of significant levels of leading-edge thrust. The consequent unusually high levels of aerodynamic performance should renew interest in supersonic-cruise vehicle design in general and in cruise-speed selection in particular.

Robins, A. W.↗

NASA NTRS1972

Boundary-Layer Edge Conditions and Transition Reynolds Number Data for a Flight Test at Mach 20 (Reentry F)

A flight experiment, designated Reentry F, was conducted to measure heat-transfer rates for laminar, transitional, and turbulent boundary layers on a 5 deg half-angle cone 3.962 m (13 ft) long with a preflight nose radius of 2.54 mm (0.10 in.). Data were obtained over an altitude range from 36.58 to 18.29 km (120 000 to 60 000 ft) at a flight velocity of about 6.096 km/sec (20 000 ft/sec). The nominal values of the free-stream total enthalpy, sharp-cone Mach number, and the wall-to-total enthalpy ratio were 18 MJ/kg (8000 Btu/lb), 15, and 0.03, respectively. Calculated boundary-layer edge conditions that account for effects of the entropy layer and corresponding local transition Reynolds numbers are reported in the present paper. Fully developed turbulent flow occurred with essentially constant boundary-layer edge conditions near the sharp-cone values. Transition data were obtained with local edge Mach numbers ranging from about 5.55 to 15. Transition Reynolds numbers, based on local condition, were as high as 6.6 x 10(exp 7) with an edge Mach number of about 14.4 at an altitude of 24.38 km (80 000 ft). The transition could be correlated with previous flight data taken over a Mach number range from 3 to 12 in terms of parameters including the effects of local unit Reynolds number, boundary-layer wall-to-edge enthalpy ratio, and local Mach number.

Johnson, Charles B.↗

NASA NTRS1980

Experimental Study of Airfoil Trailing Edge Noise: Instrumentation, Methodology and Initial Results

The mechanisms of aerodynamic noise generation at the trailing edge of an airfoil is investigated. Instrumentation was designed, a miniature semiconductor strain-gauge pressure transducer and associated electronic amplifier circuitry were designed and tested and digital signal analysis techniques applied to gain insight into the relationship between the dynamic pressure close to the trailing edge and the sound in the acoustic far-field. Attempts are made to verify some trailing-edge noise generation characteristics as theoretically predicted by several contemporary acousticians. It is found that the noise detected in the far-field is comprised of the sum of many uncorrelated emissions radiating from the vicinity of the trailing edge. These emissions appear to be the result of acoustic energy radiation which has been converted by the trailing-edge noise mechanism from the dynamic fluid energy of independent streamwise 'strips' of the turbulent boundary layer flow.

Manley, M. B.↗

NASA NTRS1980

Low-speed aerodynamic characteristics of a highly swept arrow wing configuration with several deflected leading edge concepts

The effectiveness of leading edge concepts for minimizing or controlling leading edge flow separation was studied. Emphasis was placed on low speed performance, stability, and control characteristics of configurations with highly swept wings. Simple deflection of the leading edge, a variable camber leading edge system, and a leading edge vortex flow system were among the concepts studied. The data are presented without analysis.

Gentry, G. L., Jr.↗

NASA NTRS1980

Noise radiation from the side edges of flaps

The recently observed phenomenon of high noise radiation from the side edges of flaps in flow is investigated by way of a simple two-dimensional model problem. The model is based upon a physical picture of boundary layer vorticity being swept around the edge by spanwise flow on the flap. The model problem is developed and solved and the resulting noise radiation calculated. Further, a mathematical condition for the vortex to be captured by the potential flow and swept around the edge is derived. The results show that the sound generation depends strongly upon the strength of the vorticity and distance from the edge and that it can be more intense than the more common trailing edge noise source in agreement with the experimental observations.

Hardin, J. C.↗

NASA NTRS1980

Noise generation at the side edges of flaps

The recently observed phenomenon of high noise radiation from the side edges of flaps is investigated by way of a two-dimensional model based upon a physical picture of chordwise boundary layer vorticity being swept around the edge by spanwise flow on the flap. The trajectory and resulting noise radiation for a discrete vortex in such a flow is obtained. Further, a mathematical condition for the vortex to be captured by the flow and swept around the edge is derived. The sound generation depends strongly upon the vortex strength and distance from the edge and can be more intense than trailing edge noise in agreement with experimental observations.

Hardin, J. C.↗

NASA NTRS1980

Leading-edge deflection optimization for a highly swept arrow wing configuration

Tests were also conducted to determine the sensitivity of the lateral stability derivative C sub l sub beta to geometric anhedral. The optimized leading edge deflection was developed by aligning the leading edge with the incoming flow along the entire span. Owing to the spanwise variation of upwash, the resulting optimized leading edge was a smooth, continuously warped surface. For the particular configuration studied, levels of leading edge suction on the order of 90 percent were achieved with the smooth, continuously warped leading edge contour. The results of tests conducted to determine the sensitivity of C sub l sub beta to geometric anhedral indicate values of delta C sub l sub beta/delta T which are in reasonable agreement with estimates provided by simple vortex lattice theories.

Coe, P. L., Jr.↗

NASA NTRS1981

Contributions to the theory of side edge flap noise

High and low frequency asymptotic formulas are derived which express the acoustic frequency spectrum in terms of the wavenumber-frequency spectrum of surface pressure fluctuations on the upper surface of a part span flap, measured in-board of the side edge. Interpolations between the results are used to predict the field shape and its dependence on the subsonic forward flight speed over the whole frequency range. The radiation has the characteristics of a semibaffled dipole at small values of the Strouhal number based on the chord of the flap and the forward flight speed. The degree of Doppler amplification due to forward flight speed of the aircraft is dependent on the Strouhal number; the radiation efficiency of side-edge noise sources is governed by the value of the mean side-edge gap Strouhal number, and is larger when this number is small. Theoretical predictions extrapolated to full scale suggest that the noise generated at a single side edge can exceed that generated along the whole of the trailing edge of the flap by 3 dB.

Howe, M. S.↗

NASA NTRS1981

Icing tunnel tests of a composite porous leading edge for use with a liquid anti-ice system

The efficacy of liquid ice protection systems which distribute a glycol-water solution onto leading edge surfaces through a porous skin was demonstrated in tests conducted in the NASA Lewis icing research tunnel using a composite porous leading edge panels. The data obtained were compared with the performance of previously tested stainless steel leading edge with the same geometry. Results show: (1) anti-ice protection of a composite leading edge is possible for all the simulated conditions tested; (2) the glycol flow rates required to achieve anti-ice protection were generally much higher than those required for a stainless steel panel; (3) the low reservoir pressures of the glycol during test runs indicates that more uniform distribution of glycol, and therefore lower glycol flow rates, can probably be achieved by decreasing the porosity of the panel; and (4) significant weight savings can be achieved in fluid ice protection systems with composite porous leading edges. The resistance of composite panels to abrasion and erosion must yet be determined before they can be incorporated in production systems.

Kohlman, D. L.↗

NASA NTRS1982

On the generation of side-edge flap noise

A theory is proposed for estimating the noise generated at the side edges of part span trailing edge flaps in terms of pressure fluctuations measured just in-board of the side edge on the upper surface of the flap. Asymptotic formulas are developed in the opposite extremes of Lorentz contracted acoustic wavelength large/small compared with the chord of the flap. Interpolation between these limiting results enables the field shape and its dependence on subsonic forward flight speed to be predicted over the whole frequency range. It is shown that the mean width of the side edge gap between the flap and the undeflected portion of the airfoil has a significant influence on the intensity of the radiated sound. The results indicate that the noise generated at a single side of a full scale part span flap can exceed that produced along the whole of the trailing edge of the flap by 3 dB or more.

Howe, M. S.↗