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At least 685 records · Page 38
All sourcesNASA NTRSDOE OSTI
NASA NTRS1984

Tests of a NACA 65(sub 1)-213 airfoil in the NASA Langley 0.3-meter transonic cryogenic tunnel

A wind-tunnel investigation was conducted to study the two dimensional aerodynamic characteristics of the NACA 65 sub 1-213 airfoil over a wide range of Reynolds numbers. Test temperature ranged from ambient to about 100K at pressures ranging from about 1.2 to 6.0 atm. Mach number was varied from 0.22 to 0.80 and Reynolds number (based on airfoil chord) from 3 million to 40 million. Data are included which demonstrate the effects of fixed transition, Mach number, and Reynolds number on the aerodynamic characteristics of the airfoil. A sample of data showing the effects of angle of attack on the pressure distribution is also given. The data are presented in an uncorrected form with no analysis.

Plentovich, E. B.↗
NASA NTRS1984

Transonic calculation of airfoil stability and response with active controls

Transonic aeroelastic stability and response analyses are performed for the MBB A-3 supercritical airfoil. Three degrees of freedom are considered: plunge, pitch, and aileron pitch. The control of airfoil stability and response in transonic flow are studied. Stability analyses are performed using a Pade aeroelastic model based on the use of LTRAN2-NLR transonic small disturbance finite difference computer code. Response analyses are performed by coupling the structural equations of motion to the unsteady aerodynamic forces of LTRAN2-NLR. The focus is on transonic time marching transient response solutions using modal identification to determine stability. Frequency and damping of these modes are directly compared in the complex s-plane with Pade model eigenvalues. Transonic stability and response characteristics of 2-D airfoils are discussed and comparisons are made. Application of the Pade aeroelastic model and time marching analyses to flutter suppression using active controls is demonstrated.

Batina, J. T.↗
NASA NTRS1984

Status of advanced airfoil tests in the Langley 0.3-meter transonic cryogenic tunnel

A joint NASA/U.S. industry program to test advanced technology airfoils in the Langley 0.3-meter Transonic Tunnel (TCT) was formulated under the Langley ACEE Project Office. The objectives include providing U.S. industry an opportunity to compare their most advanced airfoils to the latest NASA designs by means of high Reynolds number tests in the same facility. At the same time, industry would again experience in the design and construction of cryogenic test techniques. The status and details of the test program are presented. Typical aerodynamic results obtained, to date, are presented at chord Reynolds number up to 45 x 10(6) and are compared to results from other facilities and theory. Details of a joint agreement between NASA and the Deutsche Forschungs- und Versuchsantalt fur Luft- and Raumfahrt e.V. (DFVLR) for tests of two airfoils are also included. Results of these tests will be made available as soon as practical.

Ladson, C. L.↗
NASA NTRS1982

Experimental studies of flow separation of three airfoils at low speeds

Wind tunnel tests were conducted on the NASA LS(1)-0421 Mod, NACA 2412 and NASA GA(W)-2 airfoil sections at a Reynolds number of 2.2 x 10(6) and a Mach number of 0.13. Detailed measurements of flow fields associated with turbulent boundary layers of these airfoils were obtained at pre-stall, near-stall, and post-stall angles of attack. Velocity and pressure survey results over the airfoil and in the associated wake, are presented for fully attached flow conditions through the stalled flow condition. Extensive force, pressure, tuft survey, hot-film survey, local skin friction and boundary layer data are also included.

Rodgers, E. J.↗
NASA NTRS1984

Feedback in separated flows over symmetric airfoils

For a flow over an airfoil with laminar separation, a feedback cycle may exist whereby a Kelvin-Helmholtz instability wave emanating from the separation point on the airfoil surface grows along the shear layer and is diffracted as it interacts with the sharp trailing edge of the airfoil, causing acoustic radiation which, in turn, propagates upstream and regenerates the initial instability wave. The analysis is restricted to the high frequency limit. Solutions to the boundary-value problem are obtained using the slowly varying approximation and the method of matched asymptotic expansions. Resonant solutions exist for certain discrete values of the Reynolds and Strouhal numbers. The results are discussed and compared with available data.

Atassi, H. M.↗
NASA NTRS1984

Comparison of the full potential and Euler formulations for computing transonic airfoil flows

A quantitative comparison between the Euler and full potential formulations with respect to speed and accuracy is presented. The robustness of the codes used is tested by a number of transonic airfoil cases. The computed results are from four transonic airfoil computer codes. The full potential codes use fully implicit iteration algorithms. The first Euler code uses a fully implicit ADI iteration scheme. The second Euler code uses an explicit Runge Kutta time stepping algorithm which is enhanced by a multigrid convergence acceleration scheme. Quantitative comparisons are made using various plots of lift coefficient versus the average mesh spacing along the airfoil. Besides yielding an asymptotic limit to the lift coefficient, these results also demonstrate the truncation error behavior of the various codes. Quantitative conclusions regarding the full potential and Euler formulations with respect to accuracy, speed, and robustness can be presented.

Flores, J.↗
NASA NTRS1984

Calculation of steady and oscillating airfoil flow fields via the Navier Stokes equations

A Navier-Stokes calculation procedure is applied to high Reynolds number flow about steady and unsteady airfoils. The procedure solves the ensemble averaged governing equations via a linearized block implicit (LBI) technique which in general converges to a nominal steady flow within 120 time steps. The grid used in the computation is highly stretched thus resolving the turbulent boundary layers. Calculations have been compared with data for both NACA 4412 airfoil at high incidence and an NACA 0012 airfoil oscillating in dynamic stall. In both cases, good agreement is noted between measured and calculated surface pressure distributions.

Shamroth, S. J.↗
NASA NTRS1984

Investigation of sidewall boundary layer removal effects on two different chord airfoil models in the Langley 0.3-meter Transonic Cryogenic Tunnel

An investigation was carried out on two CAST 10-2 airfoil models with chords of 3 in. and 6 in. To evaluate the extent of sidewall influence on airfoil tests at transonic Mach numbers. The tests were conducted in the Langley 0.3-m Transonic Cryogenic Tunnel two-dimensional test section equipped with an upstream sidewall boundary layer removal system which reduces the boundary layer displacement thickness to about 1 percent of model halfspan from an initial 2 percent without boundary layer removal. Test results have shown the changes in the location of the shock on the upper surface of the airfoil to be about the same for both models with and without sidewall boundary layer removal. Even though large differences were noted in the high lift characteristics of the two models, the sidewall boundary layer removal had little effect on the differences. These tests also served to validate the boundary layer removal technique and the associated Mach number correction required with upstream boundary layer removal.

Murthy, A. V.↗
NASA NTRS1984

Transonic calculation of airfoil stability and response with active controls

Transonic aeroelastic stability and response analyses are performed for the MBB A-3 supercritical airfoil. Three degrees of freedom are considered: plunge, pitch, and aileron pitch. The objective of this study is to gain insight into the control of airfoil stability and response in transonic flow. Stability analyses are performed using a Padeaeroelastic model based on the use of the LTRAN2-NLR transonic small-disturbance finite-difference computer code. Response analyses are performed by coupling the structural equations of motion to the unsteady aerodynamic forces of LTRAN2-NLR. The focus of the present effort is on transonic time-marching transient response solutions using modal identification to determine stability. Frequency and damping of these modes are directly compared in the complex s-plane with Pademodel eigenvalues. Transonic stability and response characteristics of two-dimensional airfoils are discussed and comparisons are made. Application of the Padeaeroelastic model and time-marching analyses to flutter suppression using active controls is demonstrated.

Batina, J. T.↗
NASA NTRS1984

The effects of gusts on the fluctuating airloads of airfoils in transonic flow

Unsteady interactions of distributed and sharp-edged gusts with a stationary airfoil have been analyzed in two-dimensional transonic flow.A simple method of introducing such disturbances has been numerically implemented within the framework of unsteady, transonic small-disturbance theory. Representative solutions for various airfoils subjected to chordwise and transverse gusts show that the strength and unsteady motion of the shock wave on the airfoil significantly affect the flowfield development and, consequently, the dynamic airloads. Also a study was made of the reductions in the unsteady airloads that can be achieved by the proper active control motion of a trailing-edge flap, and a simple gust-alleviation strategy was developed. However, the chordwise pressure distributions associated with gusts are very different from those produced by trailing-edge flap oscillations. Consequently, the fluctuating lift and the unsteady pitching moments cannot both be eliminated simultaneously.

Mccroskey, W. J.↗
NASA NTRS1984

Numerical optimization of circulation control airfoil at high subsonic speed

A numerical procedure for optimizing the design of the circulation control airfoil for use at high subsonic speeds is presented. The procedure consists of an optimization scheme coupled with a viscous potential flow analysis for the blowing jet. The desired airfoil is defined by a combination of three baseline shapes (cambered ellipse and cambered ellipse with drooped and spiraled trailing edges). The coefficients of these shapes are used as design variables in the optimization process. Under the constraints of lift augmentation and lift-to-drag ratios, the airfoil, optimized at free-stream Mach 0.54 and alpha = -2 degrees can be characterized as a cambered ellipse with a drooped trailing edge. Experimental tests support the performance improvement predicted by numerical optimization.

Tai, T. C.↗
NASA NTRS1984

Performance of two transonic airfoil wind tunnels utilizing limited ventilation

A limited-zone ventilated wall panel was developed for a closed-wall icing tunnel which permitted correct simulation of transonic flow over model rotor airfoil sections with and without ice accretions. Candidate porous panels were tested in the Ohio State University 6- x 12-inch transonic airfoil tunnel and result in essentially interference-free flow, as evidenced by pressure distributions over a NACA 0012 airfoil for Mach numbers up to 0.75. Application to the NRC 12- x 12-inch icing tunnel showed a similar result, which allowed proper transonic flow simulation in that tunnel over its full speed range.

Lee, J. D.↗
NASA NTRS1984

A study of flow past an airfoil with a jet issuing from its lower surface

The aerodynamics of a NACA 0018 airfoil with a rectangular jet of finite aspect ratio exiting from its lower surface at 90 deg to the chord were investigated. The jet was located at 50% of the wing chord. Measurements include static pressures on the airfoil surface, total pressures in the near wake, and local velocity vectors in different planes of the wake. The effects of jet cross flow interaction on the aerodynamics of the airfoil are studied. It is indicated that at all values of momentum coefficients, the jet cross flow interaction produces a strong contra-rotating vortex structure in the near wake. The flow behind the jet forms a closed recirculation region which extends up to a chord length down stream of the trailing edge which results in the flow field to become highly three dimensional. The various aerodynamic force coefficients vary significantly along the span of the wing. The results are compared with a jet flap configuration.

Krothapalli, A.↗
NASA NTRS1984

Natural laminar flow airfoil design considerations for winglets on low-speed airplanes

Winglet airfoil section characteristics which significantly influence cruise performance and handling qualities of an airplane are discussed. A good winglet design requires an airfoil section with a low cruise drag coefficient, a high maximum lift coefficient, and a gradual and steady movement of the boundary layer transition location with angle of attack. The first design requirement provides a low crossover lift coefficient of airplane drag polars with winglets off and on. The other requirements prevent nonlinear changes in airplane lateral/directional stability and control characteristics. These requirements are considered in the design of a natural laminar flow airfoil section for winglet applications and chord Reynolds number of 1 to 4 million.

Vandam, C. P.↗
NASA NTRS1985

Transient technique for measuring heat transfer coefficients on stator airfoils in a jet engine environment

A transient technique was used to measure heat transfer coefficients on stator airfoils in a high-temperature annular cascade at real engine conditions. The transient response of thin film thermocouples on the airfoil surface to step changes in the gas stream temperature was used to determine these coefficients. In addition, gardon gages and paired thermocouples were also utilized to measure heat flux on the airfoil pressure surface at steady state conditions. The tests were conducted at exit gas stream Reynolds numbers of one-half to 1.9 million based on true chord. The results from the transient technique show good comparison with the steady-state results in both trend and magnitude. In addition, comparison is made with the STAN5 boundary layer code and shows good comparison with the trends. However, the magnitude of the experimental data is consistently higher than the analysis.

Gladden, H. J.↗
NASA NTRS1984

Aeroacoustic interaction of a distributed vortex with a lifting Joukowski airfoil

A first principles computational aeroacoustics calculation of the flow and noise fields produced by the interaction of a distributed vortex with a lifting Joukowski airfoil is accomplished at the Reynolds number of 200. The case considered is that where the circulations of the vortex and the airfoil are of opposite sign, corresponding to blade vortex interaction on the retreating side of a single helicopter rotor. The results show that the flow is unsteady, even in the absence of the incoming vortex, resulting in trailing edge noise generation. After the vortex is input, it initially experiences a quite rapid apparent diffusion rate produced by stretching in the airfoil velocity gradients. Consideration of the effects of finite vortex size and viscosity causes the noise radiation during the encounter to be much less impulsive than predicted by previous analyses.

Hardin, J. C.↗
NASA NTRS1985

Supercritical airfoil drag reduction by passive shock wave/boundary layer control in the Mach number range .75 to .90

Airfoils operating in the transonic region are subject to large increases in drag due to shock wave/boundary layer interactions. The concept of passive shock wave/boundary layer control seeks drag reduction by placing a thin cavity with a porous top surface at the airfoil chordwise position where a shock wave would normally occur. The higher pressure behind the shock wave circulates flow through the cavity to the lower pressure ahead of the shock wave. The effects from this circulation prevent boundary layer separation and reduce entropy increases through the shock wave. In this investigation this concept is studied at a freestream Mach number range of .75 and .90. The Mach number distributions over the model, the wake impact pressure surveys used to determine profile drag and schlieren photographs for 2.8 percent porosity and solid airfoil cases are presented and compared. Results indicate that the profile drag coefficient can be reduced by as much as 40 percent through the use of this passive drag control system.

Nagamatsu, H. T.↗