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At least 127 records · Page 7
All sourcesNASA NTRSDOE OSTI
NASA NTRS2014

Experimental Investigation of a Flexible Wing with a Variable Camber Continuous Trailing Edge Flap Design

This paper presents experimental results of a flexible wing wind tunnel model with a variable camber continuous trailing edge flap (VCCTEF) design for drag minimization, tested at the University of Washington Aeronautical Labo- ratory (UWAL). The wind tunnel test was designed to explore the relative merit of the VCCTEF concept for improved cruise efficiency through the use of low-cost aeroelastic model test techniques. The flexible wing model is a 10-scaled model of a typical transport wing and is constructed of woven fabric composites and foam core. The wing structural stiffness in bending is tailored to be half of the stiffness of a Boeing 757-era transport wing, while the torsional stiffness is about the same. This stiffness reduction results in a wing tip deflection of about 10 of the wing semi-span. The VCCTEF is a multi-segment flap design having three chordwise camber segments and five spanwise flap sections for a total of 15 individual flap elements. The three chordwise camber segments can be positioned appropriately to create a desired trailing edge camber. Elastomeric material is used to cover the gaps in between the spanwise flap sections, thereby creating a continuous trailing edge. Wind tunnel data indicate a high degree of data correlation and repeata- bility. The VCCTEF can achieve a drag reduction of up to 6.31 and an improvement in the lift-to-drag ratio (LD) of up to 4.85. The paper also presents two methods for estimating the lift coefficient of a rigid wing using a dynamic pressure correction and an aeroelastic deflection correction. Both methods provide good estimates of the rigid-wing lift coefficient.

Aerodynamics↗
NASA NTRS1976

Sound generated in the vicinity of the trailing edge of an upper surface blown flap

Far-field noise data indicated that for practical upper surface blown flap configurations, the noise radiated below the flap is dominated by the noise generated in the vicinity of the trailing edge. Sound field by turbulent mixing in the trailing edge wake is investigated experimentally and theoretically. Hot-wire measurements were made downstream of the trailing edge to determine the gross turbulence mixing characteristics of the flow. This information is used as input to a theoretical analysis of the sound field. Favorable agreement is found between predicted and measured far-field noise directivities at various frequencies and noise power spectra at various directions.

Tam, C. K. W.↗
NASA NTRS1971

Wind-tunnel investigation of effects of trailing-edge geometry on a NASA supercritical airfoil section

Wind-tunnel tests have been conducted at Mach numbers from 0.60 to 0.81 to determine the effects of trailing-edge geometry on the aerodynamic characteristics of a NASA supercritical airfoil shape. Variations in trailing-edge thicknesses from 0 to 1.5 percent of the chord and a cavity in the trailing edge were investigated with airfoils with maximum thicknesses of 10 and 11 percent of the chord.

Harris, C. D.↗
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 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 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 NTRS1977

Analysis of the separated boundary layer flow on the surface and in the wake of blunt trailing edge airfoils

The viscous flow phenomena associated with sharp and blunt trailing edge airfoils were investigated. Experimental measurements were obtained for a 17 percent thick, high performance GAW-1 airfoil. Experimental measurements consist of velocity and static pressure profiles which were obtained by the use of forward and reverse total pressure probes and disc type static pressure probes over the surface and in the wake of sharp and blunt trailing edge airfoils. Measurements of the upper surface boundary layer were obtained in both the attached and separated flow regions. In addition, static pressure data were acquired, and skin friction on the airfoil upper surface was measured with a specially constructed device. Comparison of the viscous flow data with data previously obtained elsewhere indicates reasonable agreement in the attached flow region. In the separated flow region, considerable differences exist between these two sets of measurements.

Goradia, S. H.↗
NASA NTRS1977

Experimental studies of trailing-edge conditions on an oscillating airfoil at frequency parameters of up to one

This paper describes an investigation to find the range of reduced frequencies over which the classical Kutta-Joukowsky condition is valid and the nature of deviations beyond this range. An NACA 64A010 airfoil with a 15-cm chord is supported between the side walls of a 25-cm x 35-cm low-speed wind tunnel and oscillated over reduced frequencies ranging from 0.05 to 1.2. The experimentally determined loading near the trailing edge is measured and compared with unsteady incompressible small-disturbance theory. Experimental data are obtained both with and without a boundary-layer trip. It is observed that application of the Kutta-Joukowsky condition in the theoretical analysis is valid below a reduced frequency of about 0.6 in predicting the loading in the trailing-edge region. At reduced frequencies beyond 0.8, the predicted results underestimate the measured loading near the trailing edge, and the deviations increase further with reduced frequency. Similarly, the experimental phase angle of the loading in the trailing-edge region agrees reasonably with the linear theory up to k = 0.8 but lags the predicted value beyond k = 0.8.

Satyanarayana, B.↗
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 NTRS1974

Effect of trailing-edge flap deflection on the lateral and longitudinal-stability characteristics of a supersonic transport model having a highly-swept arrow wing

A low-speed investigation has been made on a highly-swept wing model having a thickened leading edge to determine the effect of inboard trailing-edge flaps on the characteristics of the model in pitch and sideslip. The tests were made at a Mach number of 0.227 which corresponds to a Reynolds number of about 5.53 million based on the reference chord. The results showed that deflection of the trailing-edge flaps decreased the roll due to sideslip by about 20 percent at a landing lift coefficient of 0.5. The directional-stability parameter, was increased by deflection of the flaps and the range of lift coefficients over which it was positive was also increased. The longitudinal stability characteristics of the model without leading-edge devices were improved by increased flap deflection, that is, the pitch-up tendency was delayed to higher lift coefficients. The lift coefficient increment resulting from the 15 deg flap deflection compared favorably with that predicted.

Lockwood, V. E.↗
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 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 NTRS1985

A class of airfoils having finite trailing edge pressure gradients

Some new developments relevant to the design of single-element airfoils using potential flow methods are presented. In particular, the ramifications of the unbounded trailing edge pressure gradients generally present in the potential flow solution for the flow over an airfoil are examined, and the conditions necessary to obtain a class of airfoils having finite trailing edge pressure gradients developed. The incorporation of these conditions into the inverse method of Eppler for the design of low-speed airfoils is discussed, and designs generated using the modified scheme are presented for consideration. A detailed viscous analysis of one of these airfoils demonstrates a significant reduction in the strong inviscid-viscid interactions generally present near the trailing edge. These reductions offer the possibility of improved airfoil performance, as well as the possibility of improved accuracy in the methods of airfoil design and analysis.

Ormsbee, A. I.↗
NASA NTRS1947

The Effectiveness at High Speeds of a 20-Percent-Chord Plain Trailing-Edge Flap on the NACA 65-210 Airfoil

An analysis has been made of the lift control effectiveness of a 20-percent-chord plain trailing-edge flap on the NACA 65-210 airfoil section from section lift-coefficient data obtained at Mach numbers from 0.3 to 0.875. In addition, the effectiveness of the plain flap as a lift-control device has been compared with the corresponding effectiveness of both a spoiler and a dive-recovery flap on the NACA 65-210 airfoil section. The analysis indicates that the plain trailing-edge flap employed on the 10-percent-thick airfoil at Mach numbers as high as 0.875 retains at least 50-percent of its low-speed lift-control effectiveness, and is sufficiently effective in lateral control application, assuming a rigid wing, to provide adequate airplane rolling characteristics. The plain trailing-edge flap, as compared to the spoiler and the dive-recovery flap, appears to afford the most favorable characteristics as a device for controlling lift continuously throughout the range of Mach numbers from 0.3 to 0.875. At Mach numbers above those for lift divergence of the wing, either a plain flap or a dive-recovery flap may be used on a thin airplane wing to provide auxiliary wing lift when the airplane is to be controlled in flight, other than in dives, at these Mach numbers. The choice of a lift-control device for this use, however, should include the consideration of other factors such as the increments of drag and pitching moment accompanying the use of the device, and the structural and high-speed aerodynamic characteristics of the airplane which is to employ the device.

Stivers, Louis S., Jr.↗
NASA NTRS2014

Flow Phenomena in the Very Near Wake of a Flat Plate with a Circular Trailing Edge

The very near wake of a flat plate with a circular trailing edge, exhibiting pronounced shedding of wake vortices, is investigated with data from a direct numerical simulation. The separating boundary layers are turbulent and statistically identical thus resulting in a wake that is symmetric in the mean. The focus here is on the instability of the detached shear layers, the evolution of rib-vortex induced localized regions of reverse flow that detach from the main body of reverse flow in the trailing edge region and convect downstream, and phaseaveraged velocity statistics in the very near wake. The detached shear layers are found to exhibit unstable behavior intermittently, including the development of shear layer vortices as in earlier cylinder flow investigations with laminar separating boundary layers. Only a small fraction of the separated turbulent boundary layers undergo this instability, and form the initial shed vortices. Pressure spectra within the shear layers show a broadband peak at a multiple of shedding frequency. Phase-averaged intensity and shear stress distributions of the randomly fluctuating component of velocity are compared with those obtained in the near wake. The distributions of the production terms in the transport equations for the turbulent stresses are also provided.

Turbulence↗
NASA NTRSDate not supplied

Study of Trailing Edge Mini Flaps Interactions on a Wind Tunnel Scale Aspect Ratio 13.5 Common Research Model

This paper presents a computation study of an aspect ratio 13.5 wind tunnel scale Common Research Model (CRM) with trailing edge mini-flaps. A vortex-lattice model of the CRM is developed with transonic small disturbance and integral boundary-layer corrections coupled with NASTRAN equivalent beam model for the rapid CRM aeroelastic simulation. The simulations are performed with individual trailing edge mini-flap deflected and a series of groups of two adjacent trailing edge mini-flaps deflected at the same time. A surrogate model is developed to estimate the aerodynamic interactions between the flaps. To validate the surrogate model, static aeroelastic simulations are performed using FUN3D CFD solver.

Interaction↗
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.↗