Chordwise Location Research Articles

Comment on "An Inverse Boundary-Layer Method for Compressible Laminar and Turbulent Boundary-Layers"

T HE inverse boundary-layer calculation is one of the classic goals of aerodynamics and the present paper is worthy of the subject. There are only a few minor points requiring clarification. The first point is the matter of turbulent separation prediction which the author feels can be adequately handled by the Cebeci/Smith boundary-layer computation down to zero skin friction; the experimental basis for this statement is said to be presented in Ref. 1.. The second point is the author's statement that the capability does not exist today for calculating partially separating flows in two dimensions. It is a fortuitous coincidence that we have recently presented a methodology for the prediction of the pressure distribution on two-dimensional airfoils with massive turbulent separation; excellent experimental agreement was achieved for the NACA 63-018 at 18° angle of attack, for the NACA 65, 2-421 at 20° and for the NASA GA(W) 1 at 21°. Our algorithm employed the Cebeci/Smith boundary-layer method, and it was found that it did not yield good results when using the zero skin-friction separation criterion as compared to the test data and to the algorithm with the Goldschmied maximum pressure-recovery criterion. It is also pointed out in Ref. 2 that meaningful separation prediction requires two parameters, i.e., chordwise location and pressure coefficient, while the evidence of Ref. 1 is concerned only with chordwise location. This can be further illustrated in Fig. 8 of the present paper: with the direct method, C/ does not go to zero (no separation) if the correct pressure distribution is employed, while with the inverse method a steeper pressure distribution is required to make C/ reach zero (separation) at the correct location. The Cebeci method does not allow to predict both the correct skinfriction and the correct pressure coefficient when separation is present. If the data of Fig. 8 are used to verify directly the separation criterion of Ref. 3, since both skinfriction coefficients and pressure distribution were measured, it is found that good agreement exists for both pressure recovery and axial separation location. Using the experimental plots of Strickland and Simpson, I locate the minimum pressure point at X=65 in. with (7=85 fps (Fig. 3-3, p. 30) and the corresponding turbulent skinfriction coefficient Cf between 0.0032 and 0.0034 (Fig. 3-30, p. 66), yielding a predicted maximum pressure-recovery coefficient Cps between 0.64 and 0.68. The minimum velocity shown in Fig. 3-3 is (7=51 fps, yielding the experimental

OTW noise correlation for several nozzle/wing geometries using a 5:1 slot nozzle with various external reflectors

The aerodynamic design requirements of the nozzle boattail for OTW configurations are different for cruise and powered lift. For cruise, a small nozzle boattail angle (<10°) is required to minimize drag; while for powered lift, a large angle (25°–30°) is needed to provide jet flow attachment to the lifting surfaces. A solution to these design requirements is to use an external flow-deflector downstream of the nozzle for powered lift and to store (retract) this deflector for cruise. For powered lift, however, the jet/deflector/wing noise transmitted to ground observers can be a problem. In the present paper, representative acoustic spectral data obtained from a model scale study of several OTW configurations with a 5:1 slot nozzle using various external deflectors are correlated in terms of deflector geometry and flow parameters. Variations in the deflector geometry include deflector size and deflector angle. In addition, geometry variations in flap setting and nozzle chordwise location are included. Three dominant noise sources are correlated: fluctuating lift noise, flap-trailing edge and deflector related noise, and jet mixing noise. Lift and thrust measurements obtained in the study for the various configurations are summarized.

スポイラの逆効きと効き遅れについて

An experimental study is performed on the transient response of a split-type spoiler in twodimensional flow. The results indicate that the time-lag and the magnitude of reverse effectiveness of a spoiler are influenced not only by the chord-wise location of the spoiler but also by the ratio of the actuation time of it to the characteristic time of the flow (e.g. chord length devided by the flow velocity). A physical explanation of the phenomena is presented and a method to predict the above mentioned quantities is proposed.

Evaluation of internal heat-transfer coefficients for impingement-cooled turbine airfoils.

Although internal impingement cooling of the leading edge of gas-turbine airfoils has been shown to be effective, previously available heat-transfer data are not generally applicable to present-day turbine designs because of the unique geometry requirements. An experimental system which allows the ready acquisition of heat-transfer data necessary for thermal design of turbine airfoils is described. A cold-flow model is developed, and the measurement of local heat-transfer coefficients on a full-size model is accomplished by considering Joulean dissipation in very thin platinum strips bonded to the model. Heattransfer results are given which show the dependence of Nusselt number on Reynolds number, geometry, and chordwise location on the inside leading-edge region of the airfoil. Dimensionless correlations are presented which allow the designer to predict heat transfer for impingement cooling in these geometries for the range of parameters tested.

Damping loss in a cascade of fluttering phased aerofoils

SummaryA study is made of the aerodynamic damping in a cascade of oscillating aerofoils in subsonic compressible flow with the system mode described by a constant interblade phasing predicted by Lane (9). The integral downwash equation is obtained as an extension of Possio's equation for an isolated aerofoil. Procedures for a practical solution of the equation have allowed the aerodynamic reactions and damping derivatives to be evaluated with a digital computer after minimisation of the critical flutter velocity with respect to interbiade phase angle. The effect of aerodynamic lag in separation conditions as a function of reduced frequency and chordwise location is compared with results without separation.

Choked Flow About Vented or Cavitating Hydrofoils

Linearized theory for two-dimensional, choked-tunnel flow is developed for a vented or cavitating hydrofoil midway between infinite parallel walls, when the upper and lower free streamlines detach at an arbitrary chordwise location and at the trailing edge, respectively. Superposition of step-profile solutions gives the arbitrary-profile solution. Numerical calculations are made for two vented hydrofoils for exhaust at 30 percent chord and tunnel height-to-chord ratio of 3.5. The agreement with experiment becomes good after correction for the pressure at the test-section, static-pressure tap due to the hydrofoil and the cavity. Various aspects of the comparison for equal cavity number of choked flow with unbounded flow are discussed.

Laminar separation of flow around symmetrical struts at zero angle of attack

This paper presents a recent rapid and accurate experimental technique for the separation of flow around a body surface and the comparison of experimental and analytical results on laminar separation point. The experimental method used is called the “dust method”. Dispersing fine dust in the separated region near the trailing edge, the separated region of any arbitrary shape of body could be determined. The experiment shows that the dimensionless chordwise location of the laminar separation point of symmetrical and two-dimensional flow is constant for geometrically similar but different size strut models and is also independent of Reynolds numbers in the region of laminar flow. The comparison of experimental and analytical data for the point of laminar separation of symmetrical and two-dimensional flow around a strut at zero angle of attack shows a good agreement.

A Method for Predicting Lift Effectiveness of Spoilers at Subsonic Speeds

As part of an investigation concerning gust alleviation controls, it was necessary to develop a means for predicting the lift effectiveness of spoilers. Methods for estimating the lift effectiveness of ordinary flap-type control devices can usually be developed by application of thin airfoil theory. However, control devices depending on flow separation for their effectiveness, such as spoilers, have not proved amenable to analyses based solely on potential flow considerations. In the case of plain spoilers, a semiempirical approach utilizing both lifting-surface theory and experimental data has provided a satisfactory basis for analyzing variations in spoiler lift effectiveness. The method for prediction which evolved proved to be applicable at subsonic Mach Numbers up to approximately 0.95 for an angle-of-attack range restricted to an upper limit of about 6°. The method takes into account the effects of the following parameters: aspect ratio, taper ratio, sweepback, airfoil thickness ratio, chord wise location and spanwise location of the control, and spanwise extent of the control. A design chart presentation has been used to show the relative effects of variations in each parameter and to provide for a rapid, straightforward prediction procedure.