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At least 343 records · Page 19
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NASA NTRSDate not supplied

Effects of Porous Gap Fillers on 30P30N Leading-Edge SlatNoise. Part I: Surface Pressure and Acoustics

The leading edge slat of a high-lift system is one of the main contributors to airframe noise during approach. In a previous experimental study, we assessed the performance of an impermeable slat gap filler as a passive flow control device to reduce the slat noise associated with a two-dimensional, three-element high-lift airfoil. The present paper, the first of two parts,represents a follow-on investigation to assess the relative efficacy of permeable gap fillers thatallow successively higher amounts of flow to pass through the gap. To evaluate the influence of this passive flow control device on the acoustics generated by the unsteady flow near the slat, experiments are conducted in an anechoic wind tunnel by mounting a gap filler to the slat element of the two-dimensional 30P30N high-lift configuration. Measurements are performed at a single geometric angle of attack (𝛼𝑘=8◦) and three different flow speeds that correspond to Reynolds numbers of𝑅𝑒𝑐=1.2𝑒6,1.5𝑒6, and1.71𝑒6, respectively. Steady surface pressure measurements are used to gauge the influence of the permeable gap filler treatments on the overall lift. The effect of each treatment on the radiated noise is analyzed via acoustic array measurements, followed by delay-and-sum beamforming to locate the slat noise sources and to provide the integrated acoustic spectra. The porous gap fillers are found to eliminate the narrowband peaks in the acoustic spectra and, also, to yield a 10 dB reduction in the broadband noise in comparison with the baseline case with no gap filler. The porous gap filler with thelowest permeability acts similar to the impermeable gap filler examined previously. However, the aerodynamics and noise reduction both degrade with increasing permeability. An accompanying abstract describes the particle image velocimetry measurements of the flowfield within the slat cove and on either side of the permeable gap f

slat noise↗
NASA NTRS1947

Investigation of the Stability and Control Characteristics of a 1/20-Scale Model of the Consolidated Vultee XB-53 Airplane with a Full-Span Leading-Edge Slat in the Langley Free-Flight Tunnel

An investigation of the low-speed; power-off stability and control characteristics of a 1/20-scale model of the Consolidated Vultee XB-53 airplane equipped with full-span leading-edge slats has been conducted in the Langley free-flight tunnel. In this investigation it was found that the-full-span leading-edge slat gave about the same maximum lift coefficient as was obtained with the outboard single slotted flap and inboard slat. The stability and control characteristics were greatly improved except near the stall where the characteristics with the full-span slat were considered unsatisfactory because of a loss of directional stability and a slight nosing-up tendency.

Bennett, Charles V.↗
NASA NTRS1975

Theoretical effect of modifications to the upper surface of two NACA airfoils using smooth polynomial additional thickness distributions which emphasize leading edge profile and which vary quadratically at the trailing edge

An investigation was conducted on a CDC 7600 digital computer to determine the effects of additional thickness distributions to the upper surface of the NACA 64-206 and 64 sub 1 - 212 airfoils. The additional thickness distribution had the form of a continuous mathematical function which disappears at both the leading edge and the trailing edge. The function behaves as a polynomial of order epsilon sub 1 at the leading edge, and a polynomial of order epsilon sub 2 at the trailing edge. Epsilon sub 2 is a constant and epsilon sub 1 is varied over a range of practical interest. The magnitude of the additional thickness, y, is a second input parameter, and the effect of varying epsilon sub 1 and y on the aerodynamic performance of the airfoil was investigated. Results were obtained at a Mach number of 0.2 with an angle-of-attack of 6 degrees on the basic airfoils, and all calculations employ the full potential flow equations for two dimensional flow. The relaxation method of Jameson was employed for solution of the potential flow equations.

Merz, A. W.↗
NASA NTRS1982

Space Shuttle Orbiter leading edge structural subsystem thermal performance

An extensive qualification test program and the STS-1 flight of the Space Shuttle Orbiter have provided the data necessary to verify the performance of the Orbiter thermal protection system. The reinforced carbon-carbon leading edge structural subsystem is used on areas of the orbiter where temperatures exceed 2300 F. The subsystem consists of the ROC nose cap and wing leading edge panels, metallic attachments, internal insulation, and interface tiles. Thermal response data from the qualification tests and the STS-1 flight, postflight inspection, and analytical predictions support the conclusion that the thermal performance of the subsystem verified the design.

Curry, D. M.↗
NASA NTRS1991

Effects of leading-edge flap oscillation on unsteady delta wing flow and rock control

The isolated and interdisciplinary problems of unsteady fluid dynamics and rigid-body dynamics and control of delta wings with and without leading-edge flap oscillation are considered. For the fluid dynamics problem, the unsteady, compressible, thin-layer Navier-Stokes (NS) equations, which are written relative to a moving frame of reference, are solved along with the unsteady, linearized, Navier-displacement (ND) equations. The NS equations are solved for the flowfield using an implicit finite-volume scheme. The ND equations are solved for the grid deformation, if the leading-edge flaps oscillate, using an ADI scheme. For the dynamics and control problem, the Euler equation of rigid-body rolling motion for a wing and its flaps are solved interactively with the fluid dynamics equations for the wing-rock motion and subsequently for its control. A four-stage Runge-Kutta scheme is used to explicitly integrate the dynamics equation.

Kandil, Osama A.↗
NASA NTRS1987

Development flight tests of JetStar LFC leading-edge flight test experiment

The overall objective of the flight tests on the JetStar aircraft was to demonstrate the effectiveness and reliability of laminar flow control under representative flight conditions. One specific objective was to obtain laminar flow on the JetStar leading-edge test articles for the design and off-design conditions. Another specific objective was to obtain operational experience on a Laminar Flow Control (LFC) leading-edge system in a simulated airline service. This included operational experience with cleaning requirements, the effect of clogging, possible foreign object damage, erosion, and the effects of ice particle and cloud encounters. Results are summarized.

Fisher, David F.↗
NASA NTRS1954

Effect of Ice Formations on Section Drag of Swept NACA 63A-009 Airfoil with Partical-span Leading-edge Slat for Various Modes of Thermal Ice Protection

Studies were made to determine the effect of ice formations on the section drag of a 6.9-foot-chord 36 degree swept NACA 63A-009 airfoil with partial-span leading-edge slat. In general, the icing of a thin swept airfoil will result in greater aerodynamic penalties than for a thick unswept airfoil. Glaze-ice formations at the leading edge of the airfoil caused large increases in section drag even at liquid-water content of 0.39 gram per cubic meter. The use of an ice-free parting strip in the stagnation region caused a negligible change in drag compared with a completely unheated airfoil. Cyclic de-icing when properly applied caused the drag to decrease almost to the bare-airfoil drag value.

PROTUBERANCES - BODIES↗
NASA NTRS1984

Engineering analysis of drooped leading-edge wings near stall

Numerical results obtained with a three-dimensional vortex panel computer program for the calculation of inviscid, incompressible (potential) flow over infinitely thin finite wings with camber are presented. The program, which is designed for wings with drooped leading edge discontinuities, is essentially a numerical representation of lifting surface theory, involving both spanwise and chordwise distributions of vorticity. The results also include two engineering approximations. First, the effect of the leading-edge discontinuities is modeled by assuming that the vortices emanating from the discontinuities aerodynamically divide the wing into three distinct sections of lower aspect ratio. Secondly, the effect of the separated flow at high angle of attack is modeled by applying rectangular vortex panels with a varying vortex strength over the portions of the wing with attached flow.

Anderson, J. D., Jr.↗
NASA NTRS1984

Leading edge flap system for aircraft control augmentation

Traditional roll control systems such as ailerons, elevons or spoilers are least effective at high angles of attack due to boundary layer separation over the wing. This invention uses independently deployed leading edge flaps on the upper surfaces of vortex stabilized wings to shift the center of lift outboard. A rolling moment is created that is used to control roll in flight at high angles of attack. The effectiveness of the rolling moment increases linearly with angle of attack. No adverse yaw effects are induced. In an alternate mode of operation, both leading edge flaps are deployed together at cruise speeds to create a very effective airbrake without appreciable modification in pitching moment. Little trim change is required.

Rao, D. M.↗
NASA NTRS1979

Summary of past experience in natural laminar flow and experimental program for resilient leading edge

The potential of natural laminar flow for significant drag reduction and improved efficiency for aircraft is assessed. Past experience with natural laminar flow as reported in published and unpublished data and personal observations of various researchers is summarized. Aspects discussed include surface contour, waviness, and smoothness requirements; noise and vibration effects on boundary layer transition, boundary layer stability criteria; flight experience with natural laminar flow and suction stabilized boundary layers; and propeller slipstream, rain, frost, ice and insect contamination effects on boundary layer transition. The resilient leading edge appears to be a very promising method to prevent leading edge insect contamination.

Carmichael, B. H.↗
NASA NTRS1990

Three-dimensional thermal-structural analysis of a swept cowl leading edge subjected to skewed shock-shock interference heating

A three-dimensional flux-based thermal analysis method has been developed and its capability is demonstrated by predicting the transient nonlinear temperature response of a swept cowl leading edge subjected to intense three-dimensional aerodynamic heating. The predicted temperature response from the transient thermal analysis is used in a linear elastic structural analysis to determine thermal stresses. Predicted thermal stresses are compared with those obtained from a two-dimensional analysis which represents conditions along the chord where maximum heating occurs. Results indicate a need for a three-dimensional analysis to predict accurately the leading edge thermal stress response.

Polesky, Sandra P.↗
NASA NTRS2016

Direct Numerical Simulation of an Airfoil with Sand Grain Roughness on the Leading Edge

As part of a computational study of acoustic radiation due to the passage of turbulent boundary layer eddies over the trailing edge of an airfoil, the Lattice-Boltzmann method is used to perform direct numerical simulations of compressible, low Mach number flow past an NACA 0012 airfoil at zero degrees angle of attack. The chord Reynolds number of approximately 0.657 million models one of the test conditions from a previous experiment by Brooks, Pope, and Marcolini at NASA Langley Research Center. A unique feature of these simulations involves direct modeling of the sand grain roughness on the leading edge, which was used in the abovementioned experiment to trip the boundary layer to fully turbulent flow. This report documents the findings of preliminary, proof-of-concept simulations based on a narrow spanwise domain and a limited time interval. The inclusion of fully-resolved leading edge roughness in this simulation leads to significantly earlier transition than that in the absence of any roughness. The simulation data is used in conjunction with both the Ffowcs Williams-Hawkings acoustic analogy and a semi-analytical model by Roger and Moreau to predict the farfield noise. The encouraging agreement between the computed noise spectrum and that measured in the experiment indicates the potential payoff from a full-fledged numerical investigation based on the current approach. Analysis of the computed data is used to identify the required improvements to the preliminary simulations described herein.

Ribeiro, Andre F. P.↗
NASA NTRS1973

The effects of leading-edge serrations on reducing flow unsteadiness about airfoils.

High frequency surface pressure measurements were obtained from wind-tunnel tests over the Reynolds number range 1.2 x 1,000,000 to 6.2 x 1,000,000 on a rectangular wing of NACA 63-009 airfoil section. A wide selection of leading-edge serrations were also added to the basic airfoil. Under a two-dimensional laminar bubble very close to the leading edge of the basic airfoil there is a large peak in rms pressure, which is interpreted as an oscillation in size and position of the bubble. The serrations divide the bubble into segments and reduce the peak rms pressures. A low Reynolds number flow visualization test on a hydrofoil in water was also conducted. A von Karman vortex street was found trailing from the rear of the foil. Its frequency is at a much lower Strouhal number than in the high Reynolds number experiment, and is related mathematically to the airfoil trailing-edge and boundary-layer thicknesses.

Schwind, R. G.↗
NASA NTRS1985

Turbine Vane External Heat Transfer. Volume 1: Analytical and Experimental Evaluation of Surface Heat Transfer Distributions with Leading Edge Showerhead Film Cooling

Progress in predictive design capabilities for external heat transfer to turbine vanes was summarized. A two dimensional linear cascade (previously used to obtain vane surface heat transfer distributions on nonfilm cooled airfoils) was used to examine the effect of leading edge shower head film cooling on downstream heat transfer. The data were used to develop and evaluate analytical models. Modifications to the two dimensional boundary layer model are described. The results were used to formulate and test an effective viscosity model capable of predicting heat transfer phenomena downstream of the leading edge film cooling array on both the suction and pressure surfaces, with and without mass injection.

Turner, E. R.↗
NASA NTRS1973

The effects of leading-edge serrations on reducing flow unsteadiness about airfoils, an experimental and analytical investigation

High frequency surface pressure measurements were obtained from wind-tunnel tests over the Reynolds number range 1.2 times one million to 6.2 times one million on a rectangular wing of NACA 63-009 airfoil section. Measurements were also obtained with a wide selection of leading-edge serrations added to the basic airfoil. Under a two-dimensional laminar bubble very close to the leading edge of the basic airfoil there is a large apatial peak in rms pressure. Frequency analysis of the pressure signals in this region show a large, high-frequency energy peak which is interpreted as an oscillation in size and position of the bubble. The serrations divide the bubble into segments and reduce the peak rms pressures. A low Reynolds number flow visualization test on a hydrofoil in water was also conducted. A von Karman vortex street was found trailing from the rear of the foil. Its frequency is at a much lower Strouhal number than in the high Reynolds number experiment, and is related to the trailing-edge and boundary-layer thicknesses.

Schwind, R. G.↗