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Results for “AIRFOILS - CLARK Y - CHARACTERISTICS”

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 37 records · Page 2
All sourcesNASA NTRSDOE OSTI
NASA NTRS1976

Two-layer thermal barrier coating for turbine airfoils - furnace and burner rig test results

A simple, two-layer plasma-sprayed thermal barrier coating system was developed which has the potential for protecting high temperature air-cooled gas turbine components. Of those coatings initially examined, the most promising system consisted of a Ni-16Cr-6Al-0.6Y (in wt%) thermal barrier coating (about 0.005 to 0.010 cm thick) and a ZrO2-12Y2O3 (in wt%) thermal barrier coating (about 0.025 to 0.064 cm thick). This thermal barrier substantially lowered the metal temperature of an air-cooled airfoil. The coating withstood 3,200 cycles (80 sec at 1,280 C surface temperature) and 275 cycles (1 hr at 1,490 C surface temperature) without cracking or spalling. No separation of the thermal barrier from the bond coating or the bond coating from the substrate was observed.

Stecura, S.↗
NASA NTRS1975

Study of a porous surface microphone sensor in an aerofoil

The porous microphone in an airfoil is described as a directional sensor which rejects flow noise. The airfoil allows the sensor to be rotated in the airflow over a wide range of yaw angles, 0 to 90 degrees, avoiding flow separation over the surface of the sensor and its associated additional flow noise. The microphone is discussed in terms of its acoustic properties, vibration sensitivity, effect of Mach number on the directivity function, and flow noise. Additional information on the acoustic calibration of the microphone, the acceleration sensitivity of the airfoil, stationary source and receiver in a moving gas, acoustic tests in airflow, and flow noise tests of the airfoil porous surface sensor is included.

Noiseux, D. U.↗
NASA NTRS1984

Application of transonic codes to aeroelastic modeling of airfoils including active controls

A study is performed using aeroelastic modeling to investigate the stability behavior of airfoils in small-disturbance transonic flow. Two conventional airfoils, NACA 64.A006 and NACA 64A010, and a supercritical airfoil, MBB A-3, are considered. Three sets of unsteady aerodynamic data are computed using three different transonic codes (LTRAN2-NLR, LTRAN2-HI, and USTS) for comparison purposes. Stability results obtained using a constant matrix, state-space, aeroelastic model are presented in a root-locus format. Use of the state-space model is demonstrated through application to flutter suppression using active controls. Aeroelastic effects due to simple, constant gain, partial feedback, control laws that utilize displacement, velocity, and acceleration sensing are studied using a variety of control gains. Calculations are also performed using linear subsonic aerodynamic theory to reveal the differences between including and not including transonic effects in the aeroelastic model. Aeroelastic stability behavior of these airfoils is physically interpreted and discussed in detail.

Batina, J. T.↗
NASA NTRS1975

Reversed cowl flap inlet thrust augmentor

An adjustable airfoil is described for varying the geometry of a jet inlet and an ejector inlet in a jet engine for providing thrust augmentation and noise reduction. The airfoil comprises essentially a plurality of segments which are extended radially outward and retracted relative to the longitudinal axis of the engine as a function of a change in the pressure differential between the upstream and downstream surfaces of the airfoil. A servo mechanism responsive to the change in the pressure differential is coupled to the airfoil to extend and retract the airfoil segments to maintain the pressure at a maximum on the downstream side of the airfoil relative to the pressure on the upstream side of the airfoil. At low speeds, such as at take-offs and landings, the airfoil is fully extended while at high speeds it is fully retracted.

Cheng, D. Y.↗
NASA NTRS1987

A truncation error injection approach to viscous-inviscid interaction

An approach to viscous-inviscid interaction which is based on truncation error injection is presented in the context of solving flow over an airfoil. A two-dimensional interpolation scheme is used to restrict the fine grid solutions to the global coarse grid. Details on the current implementation of the approach are given, and the boundary conditions being used are discussed. Inviscid results from a NACA0012 airfoil test case and the viscous results are presented.

Goble, B. D.↗
NASA NTRS2004

Category 3: Sound Generation by Interacting With a Gust

Solve the time-dependent inviscid flow equations for this geometry subject to the specified inflow/outflow mean conditions and the fluctuating inflow velocity distortion. (1) Compute the unsteady solution until periodicity in pressure is achieved by showing that at least two successive periods are identical. Periodicity must be achieved on both the airfoil surface and the inflow/outflow boundaries. (2) Once periodicity is achieved, compute the pressure frequency spectra on the reference airfoil on both the upper and lower surfaces at x=(-0.25c,0.00, +0.25c), on the inflow boundary at (x,y)={1.5c,-0.3c), (-1.5c,0.0),(-1.5c,0.3c)} and on the outflow boundary at (x,y)= {(1.5c,-0.3c),(1.5c,0.0), (1.5c,0.3c)}. Express the spectral results in dB using the standard definition 20 log(P(sub(r.m.s)/P(sub ref), where p(sub ref) == 20 microPa. (3) Extract the harmonic pressure distributions on the inflow and outflow boundaries (i.e., on x= -/+ 1.5c lines) at the fundamental frequency omega and apply a Fourier transform in y direction to identify the spatial (i.e., mode order) structure of the pressure perturbations. Express the result in dB for each mode order. Repeat the process for the frequencies 2omega and 3omega.

Envia, Edmane↗
NASA NTRS1994

Discrete sensitivity derivatives of the Navier-Stokes equations with a parallel Krylov solver

This paper solves an 'incremental' form of the sensitivity equations derived by differentiating the discretized thin-layer Navier Stokes equations with respect to certain design variables of interest. The equations are solved with a parallel, preconditioned Generalized Minimal RESidual (GMRES) solver on a distributed-memory architecture. The 'serial' sensitivity analysis code is parallelized by using the Single Program Multiple Data (SPMD) programming model, domain decomposition techniques, and message-passing tools. Sensitivity derivatives are computed for low and high Reynolds number flows over a NACA 1406 airfoil on a 32-processor Intel Hypercube, and found to be identical to those computed on a single-processor Cray Y-MP. It is estimated that the parallel sensitivity analysis code has to be run on 40-50 processors of the Intel Hypercube in order to match the single-processor processing time of a Cray Y-MP.

Ajmani, Kumud↗
NASA NTRS1947

The Oscillating Circular Airfoil on the Basis of Potential Theory

Proceeding from the thesis by W. Kinner the present report treats the problem of the circular airfoil in uniform airflow executing small oscillations, the amplitudes of which correspond to whole functions of the second degree in x and y. The pressure distribution is secured by means of Prandtl's acceleration potential. It results in a system of linear equations the coefficients of which can be calculated exactly with the aid of exponential functions and Hankel's functions. The equations necessary are derived in part I; the numerical calculation follows in part II.

Schade, T.↗
NASA NTRS1990

Transonic blade-vortex interactions noise: A parametric study

Transonic Blade-Vortex Interactions (BVI) are simulated numerically and the noise mechanisms are investigated. The 2-D high frequency transonic small disturbance equation is solved numerically (VTRAN2 code). An Alternating Direction Implicit (ADI) scheme with monotone switches is used; viscous effects are included on the boundary and the vortex is simulated by the cloud-in-cell method. The Kirchoff method is used for the extension of the numerical 2-D near field aerodynamic results to the linear acoustic 3-D far field. The viscous effect (shock/boundary layer interaction) on BVI is investigated. The different types of shock motion are identified and compared. Two important disturbances with different directivity exist in the pressure signal and are believed to be related to the fluctuating lift and drag forces. Noise directivity for different cases is shown. The maximum radiation occurs at an angle between 60 and 90 deg below the horizontal for an airfoil fixed coordinate system and depends on the details of the airfoil shape. Different airfoil shapes are studied and classified according to the BVI noise produced.

Lyrintzis, A. S.↗
NASA NTRS1990

A study of the noise mechanisms of transonic blade-vortex interactions

Transonic blade-vortex interactions (BVI) are simulated numerically and the noise mechanisms are investigated. The two-dimensional high frequency transonic small disturbance equation is solved numerically (VTRAN2 code). An ADI scheme with monotone switches is used; viscous effects are included on the boundary, and the vortex is simulated by the cloud in cell method. The Kirchhoff method is used for the extension of the numerical two-dimensional near-field aerodynamic results to the linear acoustic three dimensional far field. The viscous effects (shock/boundary layer interactions) on BVI is investigated. The different types of shock motion are identified and compared. Two important disturbances with different directivity exist in the pressure signal and are believed to be related to the fluctuating lift and drag forces. Noise directivity for different cases is shown. The maximum radiation occurs at an angle between 60 and 90 degrees below the horizontal for an airfoil-fixed coordinate system and depends on the details of the airfoil shape. Different airfoil shapes are studied and classified according to the BVI noise produced.

Lyrintzis, Anastasios S.↗
NASA NTRS1985

A simple, accurate and efficient algorithm for unsteady transonic flow

The derivation of an algorithm for computing an airfoil's response to small unsteady perturbations is described. The analysis of the far-field boundary conditions for the flow field, and the application of the time-linearization technique to develop an accurate description of the flow field are discussed. The treatment of moving shock waves is examined. The alternating direction implicit scheme of Ballhaus and Steger (1975) is utilized for the discretization of the governing equations and boundary conditions. Numerical examples demonstrating the applicability of the time-linearization method to determine the correlation between airfoil and pressure distribution are provided. The effect of wind tunnel walls on unsteady transonic testing is studied.

Fung, K.-Y.↗
NASA NTRS1981

Transonic flow studies

Major emphasis was on the design of shock free airfoils with applications to general aviation. Unsteady flow, transonic flow, and shock wave formation were examined.

Seebass, A. R.↗
NASA NTRS1983

Noise suppressor for turbo fan jet engines

A noise suppressor is disclosed for installation on the discharge or aft end of a turbo fan engine. Within the suppressor are fixed annular airfoils which are positioned to reduce the relative velocity between the high temperature fast moving jet exhaust and the low temperature slow moving air surrounding it. Within the suppressor nacelle is an exhaust jet nozzle which constrains the shape of the jet exhaust to a substantially uniform elongate shape irrespective of the power setting of the engine. Fixed ring airfoils within the suppressor nacelle therefore have the same salutary effects irrespective of the power setting at which the engine is operated.

Cheng, D. Y.↗
NASA NTRS1983

Calculations of transonic potential flows by a parameter free procedure

A parameter-free computational procedure for transonic potential flows around an airfoil is presented. The method is based on an inexact, Newton-like iterative scheme, and a preconditioned minimal residual algorithm is used to solve the linear system of equations arising at each iterate. This iterative procedure is formulated so that the difficulties associated with the approximate factorization technique, namely the requirement of iteration parameters and the boundary conditions for the intermediate variable are eliminated. Numerical experiments are given for potential flows around NACA 0012 airfoil with different Mach numbers and different angles of attack. These results are then compared with those obtained by the approximate factorization techniques.

Wong, Y. S.↗
NASA NTRS1977

Two-layer thermal barrier coating for high temperature components

A simple two-layer plasma-sprayed thermal barrier coating system was developed which has the potential for protecting high temperature air-cooled gas turbine components. Of the initially examined coatings, the most promising system is an Ni-16Cr-6Al-0.6Y (in wt%) bond coating (about 0.005 to 0.010 cm thick) and a ZrO2-12Y2O3 (in wt%) thermal barrier coating (about 0.025 to 0.064 cm thick). This thermal barrier substantially lowered the metal temperature of the air-cooled airfoil. The coating withstood 3200 cycles (80 s at 1280 C surface temperature) and 275 cycles (1 hr at 1490 C surface temperature) without cracking or spalling. No separation of the thermal barrier from the bond coating or the bond coating from the substrate was observed.

Stecura, S.↗
NASA NTRS1985

Viscous effects on transonic airfoil stability and response

An integral boundary layer model coupled to the inviscid XTRAN2L transonic small disturbance code is used to investigate viscous effects on transonic airfoil stability and response. A pulse transfer function analysis is employed for the computation of unsteady transonic airloads which yields forces for a wide range of reduced frequencies in a single flowfield computation. Nonlinear, time-marching aeroelastic solutions are obtained which exhibit the effects of viscosity on airfoil response behavior and flutter. A state-space aeroelastic model using Pade approximants to describe unsteady airloads is used to study viscosity effects on transonic airfoil stability. The state-space dynamic pressure root loci are in good general agreement with time marching damping and frequency estimates.

Berry, H. M.↗
NASA NTRS1976

EBF noise reduction through nozzle/flap positioning

Results are presented of an experimental and analytical study of the dependence of externally blown flap (EBF) noise on the relative position and shape of engine exhaust nozzle. Tests, conducted on a 1/15 scale model of a triple-slotted EBF system, indicate that a significant reduction (of up to 10 to 15 db for no forward speed case and of up to 5 to 10 db for forward speed case) is possible in the low frequency (around 63 Hz) region of the noise spectrum of the full scale device for small nozzle/flap separation distances. The overall acoustic performance, measured in PNdb, does not exhibit significant reductions. The analysis of the EBF noise is carried out for two limiting airfoil in a free jet. The analytical results also suggest that low frequency noise can be reduced by placing the nozzle close to the flow turning elements.

Kadman, Y.↗
NASA NTRS2007

Low Density, High Creep Resistant Single Crystal Super Alloy for Turbine Airfoils

A nickel-base superalloy article for use in turbines has increased creep resistance and lower density. The superalloy article includes, as measured in % by weight, 6.0-12.0% Mo, 5.5-6.5%Al, 3.0-7.0% Ta, 0-15% Co, 2.0-6.0% Cr, 1.0-4.0% Re, 0-1.5% W, 0-1.5% Ru, 0-2.0% Ti, 0-3.0% Nb, 0-0.2% Hf, 0-0.02% Y, 0.001-0.005% B, 0.01-0.04% C, and a remainder including nickel plus impurities.

MacKay, Rebecca A.↗