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At least 145 records · Page 8
All sourcesNASA NTRSDOE OSTI
NASA NTRS1986

Effect of blunt trailing edge on rotor broadband noise

The production of high-frequency broadband noise by turbulent vortex shedding from rotor blades with blunt trailing edges is investigated analytically. The derivation of the governing equations, analogous to that of Kim and George (1982) for boundary-layer/trailing-edge noise, is explained, and numerical results are compared with the experimental data of Hubbard et al. (1981) and Lowson et al. (1972) in graphs. It is shown that vortex-shedding noise is a significant component of blunt-trailing-edge rotor broadband noise and that the analytical method employed gives reasonable predictions. The need for a better empirical expression for the normalized spectrum and for more measurements of surface pressure fluctuations near blunt trailing edges is indicated.

Chou, S.-T.↗
NASA NTRS1987

Navier-Stokes calculations and turbulence modeling in the trailing edge region of a circulation control airfoil

The accurate prediction of turbulent flows over curved surfaces in general and over the trailing edge region of circulation control airfoils in particular requires the coupled efforts of turbulence modelers, numerical analysts and experimentalists. The purpose of the research program in this area is described. Then, the influence on turbulence modeling of the flow characteristics over a typical circulation control wing is discussed. Next, the scope of this effort to study turbulence in the trailing edge region of a circulation control airfoil is presented. This is followed by a brief overview of the computation scheme, including the grid, governing equations, numerical method, boundary conditions and turbulence models applied to date. Then, examples of applications of two algebraic eddy viscosity models to the trailing edge region of a circulation control airfoil is presented. The results from the calculations is summarized, and conclusions drawn based on examples. Finally, the future directions of the program is outlined.

Viegas, John R.↗
NASA NTRS1988

Pressure distributions from subsonic tests of an advanced laminar-flow-control wing with leading- and trailing-edge flaps

An unswept, semispan wing model equipped with full-span leading- and trailing-edge flaps was tested in the Langley 14- by 22-Foot Subsonic Tunnel to determine the effect of high-lift components on the aerodynamics of an advanced laminar-flow-control (LFC) airfoil section. Chordwise pressure distributions near the midsemispan were measured for four configurations: cruise, trailing-edge flap only, and trailing-edge flap with a leading edge Kreuger flap of either 0.10 or 0.12 chord. Part 1 of this report presents a representative sample of the plotted pressure distribution data for each configuration tested. Part 2 (under separate cover) presents the entire set of plotted and tabulated pressure distribution data. The data are presented without analysis.

Applin, Zachary T.↗
NASA NTRS1988

Pressure distributions from subsonic tests of an advanced laminar-flow-control wing with leading- and trailing-edge flaps

An unswept, semispan wing model equipped with full-span leading- and trailing-edge flaps was tested in the Langley 14- by 22-Foot Subsonic Tunnel to determine the effect of high-lift components on the aerodynamics of an advanced laminar-flow-control (LFC) airfoil section. Chordwise pressure distributions near the midsemispan were measured for four configurations: cruise, trailing-edge flap only, and trailing-edge flap with a leading-edge Krueger flap of either 0.10 or 0.12 chord. Part 1 of this report (under separate cover) presents a representative sample of the plotted pressure distribution data for each configuration tested. Part 2 presents the entire set of plotted and tabulated pressure distribution data. The data are presented without analysis.

Applin, Zachary T.↗
NASA NTRS1989

Merging of aircraft vortex trails - Similarities to magnetic field merging

This paper discusses the phenomenological and formal similarities between the merging of aircraft vortex trails and the merging of magnetic field lines in a plasma. High-resolution photographs are shown of smoke trails from the wing tips of an airplane. These photographs show that the two vortex trails merge together downstream of the aircraft in a way similar to the merging of oppositely directed magnetic field lines in a plasma. Although there are some differences, this correspondence is apparently related to the fact that the vorticity equation in a fluid has the same mathematical form as the magnetic field equation in an MHD plasma. In both cases the merging proceeds at a rate considerably faster than would be predicted from classical estimates of the viscosity and resistivity. The enhanced merging rate in the fluid case appears to result from turbulence that increases the diffusion rate in the merging region.

Gurnett, Donald A.↗
NASA NTRS1991

The symmetric turbulent plane wake downstream of a sharp trailing edge

The analysis and modeling of the symmetric turbulent plane wake downstream of a sharp trailing edge is addressed. A compact description of the flow near the trailing edge is formulated using the results of a previous asymptotic analysis. The new description retains the two-layered structure identified in the previous work, and it clarifies the principal dynamics of the flow in the near-wake outer layer, away from the wake centerline. For zero-pressure-gradient flow, the near-wake outer layer is shown to be represented to leading order by the similarity solution that governs the outer region of the surface boundary layer. The leading perturbation in the outer layer due to the developing near-wake inner-layer flow is identified, and this is shown to be asymptotically smaller than undetermined higher-order terms associated with the surface boundary-layer flow. Results of the new near-wake analysis are used to formulate an algebraic eddy viscosity model for wake flow predictions at arbitrary distances from the trailing edge. The model is used in a numerical solution of the boundary layer equations, and computed velocity and Reynolds stress profiles are shown to compare well with experimental data.

Bogucz, E. A.↗
NASA NTRS1992

Effect of airfoil (trailing-edge) thickness on the numerical solution of panel methods based on the Dirichlet boundary condition

The practical limit of airfoil thickness ratio for which acceptable engineering results are obtainable with the Dirichlet boundary-condition-based numerical methods is investigated. This is done by studying the effect of thickness on the calculated pressure distribution near the trailing edge and by comparing the aerodynamic coefficients with available exact solutions. The first objective of this study, owing to the wide use of such computational methods, is to demonstrate the numerical symptoms that occur when the body or wing thickness approaches zero and to increase the awareness of potential users of these methods. Additionally, an effort is made to obtain the practical limits of the trailing-edge thickness where such problems will appear in the flow solution, and to propose some possible cures for very thin airfoils or those with cusped trailing edges.

Yon, Steven↗
NASA NTRS1991

Effects of leading and trailing edge flaps on the aerodynamics of airfoil/vortex interactions

A numerical procedure based on the unsteady 2D full potential equation is presently used to simulate the effects of leading-edge and trailing-edge flaps on the aerodynamics of airfoil-vortex interactions. Attention is given to unsteady flap-motion effects, which alleviate those interactions at sub- and supercritical onset flows. For subcritical interactions, the results obtained indicate that trailing-edge flaps can be used to alleviate the impulsive loads experienced by the airfoil; for supercritical interactions, a leading- rather than trailing-edge flap must be used to alleviate the interaction.

Hassan, Ahmed A.↗
NASA NTRS1993

Vortex developments over steady and accelerated airfoils incorporating a trailing edge jet

Computational and experimental studies are conducted to investigate the influence of a trailing edge jet on flow separation and subsequent vortex formation over steady and accelerated airfoils at high angles of attack. A computer code, employing the stream function-vorticity approach, is developed and utilized to conduct numerical experiments on the flow problem. To verify and economize such efforts, an experimental system is developed and incorporated into a subsonic wind tunnel where streamline and vortex flow visualization experiments are conducted. The study demonstrates the role of the trailing edge jet in controlling flow separation and subsequent vortex development for steady and accelerating flow at angles past the static stall angle of attack. The results suggest that the concept of the trailing edge jet may be utilized to control the characteristics of unsteady separated flows over lifting surfaces. This control possibility seems to be quite effective and could have a significant role in controlling unsteady separated flows.

Finaish, F.↗
NASA NTRS1950

Theoretical stability derivatives of thin sweptback wings tapered to a point with sweptback or sweptforward trailing edges for a limited range of supersonic speeds

The stability derivatives valid for a limited range of supersonic speeds are presented for a series of sweptback wings tapered to a point with sweptback or sweptforward trailing edges. These wings were derived by modifying the trailing edge of a basic triangular wing so that it coincided with lines drawn from the wing tips to the wing axis of symmetry. The stability derivatives were formulated by using the pressure distributions previously obtained for the basic triangular wing for angle of attack, constant vertical acceleration, sideslip, pitching, rolling, and yawing. Explicit expressions are given for the stability derivatives with respect to principal body axes, and conversion formulas are provided for the transformation to stability axes. The results are limited to Mach numbers for which the wing is contained within the Mach cones springing from the vertex and from the trailing edge of the center section of the wing.

Malvestuto, Frank S , Jr↗
NASA NTRS1994

Effect of leading- and trailing-edge flaps on clipped delta wings with and without wing camber at supersonic speeds

An experimental investigation of the aerodynamic characteristics of thin, moderately swept fighter wings has been conducted to evaluate the effect of camber and twist on the effectiveness of leading- and trailing-edge flaps at supersonic speeds in the Langley Unitary Plan Wind Tunnel. The study geometry consisted of a generic fuselage with camber typical of advanced fighter designs without inlets, canopy, or vertical tail. The model was tested with two wing configurations an uncambered (flat) wing and a cambered and twisted wing. Each wing had an identical clipped delta planform with an inboard leading edge swept back 65 deg and an outboard leading edge swept back 50 deg. The trailing edge was swept forward 25 deg. The leading-edge flaps were deflected 4 deg to 15 deg, and the trailing-edge flaps were deflected from -30 deg to 10 deg. Longitudinal force and moment data were obtained at Mach numbers of 1.60, 1.80, 2.00, and 2.16 for an angle-of-attack range 4 deg to 20 deg at a Reynolds number of 2.16 x 10(exp 6) per foot and for an angle-of-attack range 4 deg to 20 deg at a Reynolds number of 2.0 x 10(exp 6) per foot. Vapor screen, tuft, and oil flow visualization data are also included.

Hernandez, Gloria↗
NASA NTRS1994

Effects of leading and trailing edge flaps on the aerodynamics of airfoil/vortex interactions

A numerical procedure has been developed for predicting the two-dimensional parallel interaction between a free convecting vortex and a NACA 0012 airfoil having leading and trailing edge integral-type flaps. Special emphasis is placed on the unsteady flap motion effects which result in alleviating the interaction at subcritical and supercritical onset flows. The numerical procedure described here is based on the implicit finite-difference solutions to the unsteady two-dimensional full potential equation. Vortex-induced effects are computed using the Biot-Savart Law with allowance for a finite core radius. The vortex-induced velocities at the surface of the airfoil are incorporated into the potential flow model via the use of the velocity transpiration approach. Flap motion effects are also modeled using the transpiration approach. For subcritical interactions, our results indicate that trailing edge flaps can be used to alleviate the impulsive loads experienced by the airfoil. For supercritical interactions, our results demonstrate the necessity of using a leading edge flap, rather than a trailing edge flap, to alleviate the interaction. Results for various time-dependent flap motions and their effect on the predicted temporal sectional loads, differential pressures, and the free vortex trajectories are presented

Hassan, Ahmed A.↗
NASA NTRS2000

Massive Remnant of Evolved Cometary Dust Trail Detected in the Orbit of Halley-Type Comet 55P/Tempel-Tuttle

There is a subpopulation of Leonid meteoroid stream particles that appear to form a region of enhanced numbers density along the path of the stream. This structure has been detected in the vicinity of the parent comet, and its variation from one apparition to the next has been traced. A significant amount of known comet 55P/Tempel-Tuttle debris is in this component, called a "filament," which has dimensions exceeding by an order of magnitude that expected for a cometary dust trail. As filament particles are of a size comparable to those found in trails, the emission ages of the particles comprising the filament must be intermediate between the age of the current trail particles (which have not been observed) and the age of the background particles comprising the annual showers. The most likely explanation for this structure is planetary perturbations acting differently on the comet and large particles while at different mean anomalies relative to each other.

Jenniskens, P.↗
NASA NTRS2000

Heat Transfer in a Complex Trailing Edge Passage for a High Pressure Turbine Blade - Part 1: Experimental Measurements: Experimental Measurements - Part 1

A combined experimental and computational study has been performed to investigate the detailed heat transfer coefficient distributions within a complex blade trailing edge passage. The experimental measurements are made using a steady liquid crystal thermography technique applied to one major side of the passage. The geometry of the trailing edge passage is that of a two-pass serpentine circuit with a sharp 180-degree turning region at the tip. The upflow channel is split by interrupted ribs into two major subchannels, one of which is turbulated. This channel has an average aspect ratio of roughly 14:1. The spanwise extent of the channel geometry includes both area convergence from root to tip, as well as taper towards the trailing edge apex. The average section Reynolds numbers tested in this upflow channel range from 55,000 to 98,000. The tip section contains a turning vane near the extreme comer. The downflow channel has an aspect ratio of about 5:1, and also includes convergence and taper. Turbulators of varying sizes are included in this channel also. Both detailed heat transfer and pressure distribution measurements are presented. The pressure measurements are incorporated into a flow network model illustrating the major loss contributors.

Bunker, Ronald S.↗
NASA NTRS2002

Heat Transfer in a Complex Trailing Edge Passage for a High Pressure Turbine Blade: Simulation Results - Part 2:

A combined experimental and numerical study to investigate the heat transfer distribution in a complex blade trailing edge passage was conducted. The geometry consists of a two pass serpentine passage with taper toward the trailing edge, as well as from hub to tip. The upflow channel has an average aspect ratio of roughly 14:1, while the exit passage aspect ratio is about 5:1. The upflow channel is split in an interrupted way and is smooth on the trailing edge side of the split and turbulated on the other side. A turning vane is placed near the tip of the upflow channel. Reynolds numbers in the range of 31,000 to 61,000, based on inlet conditions, were simulated numerically. The simulation was performed using the Glenn-HT code, a full three-dimensional Navier-Stokes solver using the Wilcox k-omega turbulence model. A structured multi-block grid is used with approximately 4.5 million cells and average y+ values on the order of unity. Pressure and heat transfer distributions are presented with comparison to the experimental data. While there are some regions with discrepancies, in general the agreement is very good for both pressure and heat transfer.

Rigby, David L.↗
NASA NTRS2002

Trailing Edge Noise Prediction Based on a New Acoustic Formulation

A new analytic result in acoustics called 'Formulation 1B,' proposed by Farassat, is used to compute broadband trailing edge noise from an unsteady surface pressure distribution on a thin airfoil in the time domain. This formulation is a new solution of the Ffowcs Williams-Hawkings equation with the loading source term, and has been shown in previous research to provide time domain predictions of broadband noise that are in excellent agreement with experiment. Furthermore, this formulation lends itself readily to rotating reference frames and statistical analysis of broadband trailing edge noise. Formulation 1B is used to calculate the far field noise radiated from the trailing edge of a NACA 0012 airfoil in low Mach number flows, using both analytical and experimental data on the airfoil surface. The results are compared to analytical results and experimental measurements that are available in the literature. Good agreement between predictions and measurements is obtained.

Casper, J.↗
NASA NTRS2003

Application of Passive Porous Treatment to Slat Trailing Edge Noise

Porous trailing-edge treatment is investigated as a passive means for slat noise reduction by using time-accurate simulations based on Reynolds-averaged Navier-Stokes equations. For the model scale high-lift configuration used during previous experiments in the Low-Turbulence Pressure Tunnel at NASA Langley Research Center, application of the proposed treatment over a minute fraction of the slat surface area is shown to mitigate the noise impact of the trailing edge, with no measurable aerodynamic penalty. Assessment of the pressure fluctuations in the vicinity of the treated edge indicates a potential noise reduction in excess of 20 dB. The primary mechanism underlying this reduction is related to the reduced strength of Strouhal shedding from the finite thickness trailing edge. A secondary effect of the treatment involves an upward shift in the Strouhal-shedding frequency to a frequency band of reduced auditory sensitivity in a full-scale application.

Khorrami, Mehdi R.↗
NASA NTRS2003

Broadband Trailing Edge Noise Predictions in the Time Domain

A recently developed analytic result in acoustics, "Formulation 1B," is used to compute broadband trailing edge noise from an unsteady surface pressure distribution on a thin airfoil in the time domain. This formulation is a new solution of the Ffowcs Willliams-Hawkings equation with the loading source term, and has been shown in previous research to provide time domain predictions of broadband noise that are in excellent agreement with experimental results. Furthermore, this formulation lends itself readily to rotating reference frames and statistical analysis of broadband trailing edge noise. Formulation 1B is used to calculate the far field noise radiated from the trailing edge of a NACA 0012 airfoil in low Mach number flows, by using both analytical and experimental data on the airfoil surface. The acoustic predictions are compared with analytical results and experimental measurements that are available in the literature. Good agreement between predictions and measurements is obtained.

Casper, Jay↗