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NATIONAL, ADVISORY COMMITTEE
FOR AERONAUTICS
TECHNICAL NOTE No. 1574 WIND-TUNNEL INVESTIGATION OF THE BOUNDARY LAYER ON AN , - NACA 0009 AIRFOIL HAVING 0.25- AND 0.50-AIRFOIL CHORD PLAIN SEALED FLAPS By Jack D. Brewer and Josephine F. Polhamus Langley Memorial Aeronautical Laboratory Langley Field, Va.
Washington April 1948 . w- 4 P .
NATIONAL ADvISCRY‘COMMtiELE FOR AERONAUTICS , TE;cHNICAL NOTE NO. 1574 WI- INVFSTIGATION OF THE BOUNDARY LAYER ON AN NACA 0009 AIRFOIL HAVING 0.25- AND 0.504IRFOIL CHORDPLAIN SEAIXD FLAPS By Jack D. 'Brewer and Josephine F. Polhamus -- SUMMARY An investigation was conducted to determine the boundary-lsyer characteristics of an NACA 0009 airfoil equipped with 0.25 and 0.5O-airfoil The tests were made to define as thoroughly chord plain sealed flaps.
as possible the characteristics of two of the configurations used in a comprehensive investigation of control-surface chsracteristics and also to provide additional data for comparison with previous boundary-layer analyses.
The measured velocity profiles and the boundary-layer parsmeters determined from them are presented.
INTRODUCTION .
An extensive investigation of control-surface characteristics has been conducted in the Langley b-by &foot vertical tunnel.. The greater pert of the investigation consisted of twoilimensional tests of an NACA 0009 alrfoil with various flap arrangements. Pressure-distributions of some of the models are presented in references 1 to 3. Most of the summsrized in reference 4.
force and moment measurements have been In the present investigation measurements were made of the boundary- layer characteristics of an NACA 0009 airfoil having 0.25 and 0.5C-airfoil chord plain sealed flaps. Results of force snd moment tests of the same The tests were made to define as model are reported in reference 5.
completely as pOSSibh3 the characteristics of two representative airfoil- flap configurations under the specific test conditions of the general control-surface investigation.
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A previous boundary-layer investigation (reference 6) indicated that measured boundsry-layer parameters did not agree with calculated boundary layer parameters behind the control-surface hinge line. The boundary-layer M NACA TN No. 1574 measurement8 of this investigation are intended foruse in obtaining a more accurate method for predicting the boundary layer in the regior.of the trailing edge from which prediction a correlatia of measured hinge moments and calculated boundsry-layer parameters might then be developed.
SYMBOIS C airfoil chord flap chord Cf X distance along chord distance perpendicular to surface free-stream dynamic pressure outside boundary layer qo free-stream velocity UO velocity at outer edge of boundary layer U U velocity within boundary layer 6* boundary-layer displacement thickness boundary-layer momentum thickness 9 [cf$J- +&V-j H boundary-layer Shape parameter (6*/e) H ratio of velocity at edge of boundary layer to free-stream velocity (U/U,) flap deflection 6f angle of attack for airfoil of infinite aspect ratio with =0 subscript u for uncorrected value - APPARATUS ANDMODEL The present investigation wa8 conducted in the Langley k- by &foot vertical.tunnel (described in reference 7 and modified as related in reference 2). The model, constructed of laminated mahogany to the NACA 0009 profile, had a chord of 2 feet-and completely spanned the It was equipped with plain sealed flaps having chords test section.
of 25 percent and 50 percent of the airfoil chord.
E NACA TN No* 1574 Ordinates for the airfoil are given in table I. Dimensions of the c model 8re given in figure 1.
Boundary-layer profiles were measured by means of two pressure "mice," one mounted on the upper and one on the lower surface of the model.
Each "mouse" consisted of 14 total-pressure tubes and 2 static+pressure tubes.
Each of the mice tubes WBB calibrated against a standard tube.
TEST PROCEDURE The tests were made at an average dynamic pressure of 16.2 pound8 per square foot, which, for stsndarrd atmospheric conditions, corresponds to an airspeed of approximately 79.6 miles per hour and to a test Reynolds number of approximately 1.49 x 106, based on the 2-foot chord. The turbulence factor for the Langley 4-by 6-foot vertical tunnel is 1.93.
For the present investigation the model was tested as though only one fl&p existed at a time. (See fig. 1.) The nose gap of the flap not in me was completely filled with plasticine and faired to the airfoil contour. The gap of the flap being investigated WBB sealed by a small amount of plasticine placed only at'the nose of the flap. Measurem3ntB were made for positive deflections only but, because mice were located on both the upper and lower surfaces and because the model was symmetrical, values for equivalent negative flap deflections can be obtained.
The total pressure and static pressure were measured relative to the .
total pressure in the free stream. Positions of the tubes above the surface were measured to the nearest i/128 inch.
CORRECTIONS Tunnel corrections were applied to the angle of attack by an extension of the method presented in reference 8. The equations used were as follows: For the 0.25~ flap,'- = 1.023aou + 0.0031&f =oU 'For the 0.50~ flap, = 1.023~~~~ + o.olios, =0 NACA TN No. 1574 ?
A correction for the effective center location, given in reference 9, was applied to the mice-tube heights.
Boundarplayer velocity profiles for various stations along the airfoil chord are presented in figures 2 to 8 for given flap conditions.
The velocity profiles shown are based on the velocity at the outer edge of the boundary layer U. Conversion to profiles based on the free- stream velocity can be obtained by multiplying the given velocity ratios by the factor K presented on the figures. The factor (K=e> is related to the pressure coefficient approxImatsly by the equati&: K= (l-&2 where p=p-PO Q, and p = static pressure at a pointon airfoil = static pressure in free air stream PO Some of the test points have been omitted from the figures in order to make the curves more legible.
In figures 9 to 15, the boundary-layer displacement-thiclmess parameter W/c, the momentum-thickness parameter G/c, and the shape parameter H are plotted against corrected angle of attack for various stations along the airfoil. Consistent scales could not be used throughout the figures because of the wide variations in the values of The values of H on the upper surface at the the parameters.
0.25~ station are fairly large at mostnegative angles of attack.
This condition indicates a laminar boundary layer as far back as that station at those angles.of attack. Fmer (referenoes 10 and 11) boundary layer should be indicates that values of H for a laminar substantially higher than those shown in these figures (at the higher negative angles of attack), but-as yet no explanation of-the discrepancy has been folind. The sudden break in the curve (at approximately 0' for O" flap deflection S, and at more negative angles of attack as NACA TN No. 1574 the flap is deflected) and the comparatively low values of H in the positive range of angle of attack indicate transition to turbulent flow. At the 0.46~ station and at stations further aft, transition has already occurred throughout the angle-of-attack range. These observations are substantiated by the velocity profiles.
As the angle of attack is increased, the rapid increase in the value of H near the trailing edge on the upper surface indicates an approach to separation.
The shape parameter increases to over 2.4 at the trailing edge with the 0.25~ flap deflected loo (fig. E?(c)) and becomes even larger with the O.wc flap deflection. (See fig. 15(c)-) Reference 12 predicts turbulent separation at values of H between 1.8 and 2.6 but in no case is final separation shown by the present velocity--profile results l It is possible, therefore, that the mouse tubes-near the surface, which measure an average flow, till not always indicate when separation occurs.
An indication of the variation of V/c, 8/c, and H with 8f can be obtained from-figure 16. The data are presented for various angles of attack, for the upper surface, and for only one station (x = 0.95c). Values for negative flap deflection6 are actually values for conditions on the lower surface at positive flap deflections. The discrepancies at zero angle of attack are probably caused by construction irregularities, nonuniform surface conditions, or misalinement of the air stream.
Similar plots for the other stations can be obtained from the data of figure8 9 to 15.
Figures 17 and 18 present plots of u/C against H for the two flaps tested, for various angle-&attack and f&q-deflection conditions, and for two values of y/Q. Curves from figure 9 of reference l2, obtained from a large amount of turbulent boundary layer data on various plain airfoils, are presented for coqarison with the data of the present present data for an airfoil with sealed flaps deflected up paper - 'The to 10' substantiate the conclusion of reference 12 that, for turbulent boundary layers, up is a function of H alone for a given value of y/e - CONCLUDING REMARES A boundary-layer investigation has been conducted in the Langley 4- by 6-foot vertical tunnel on an NACA 0009 airfoil having 0.25- and O.~-airfoil-chord plain sealed flaps. The purpose of the tests was to determine the characteristics of two of the configurations used in a comprehensive control-surface investigation a8 completely as possible and also to provide data for comparison with previous boundary-layer results. The data my be useful for various analyses, especially for a possible hingewent correlation. Because of the high turbulence NACA TN No. 1574 I level of the Langley & by &foot- tunnel, however, it-is suwested that only data obtained in the same tuncel at .the same Reynolds number be used in any analysis involving these results.
The measured velocfty profiles and the boundary-layer parameters determined fEcun them are presented. -- Langley Memorial Aeronautical Laborat-ory National Advisory Committee for Aeronautics Langley Field, Va., December 23, 1947 Y 3 NACA TN No. 1574 RIEFERENCES .
1. Street, William G., and Ames, Milton B., Jr.: Pressure-Distribution Investigation of an N.A.C.A. 0009 Airfoil with CqO-Percent-Chord Plain Flap and Three Tabs. NACA TN No. 734, 1939.
2. Ames, Milton B., Jr., and Sears, Richard I.: Pressure-Distribution Investigation of an N,.A.C.A. 0009 Airfoil with a 30+ercent-Chord Plain Flap and Three Tabs. NACA TN No. 759, 19h.O.
3. Ames, Milton B., Jr., and Sears, Richard I.: Pressure-Distribution Investigation of an,.N.A.C.A. 0009 Airfoil with an &Percent-Chord Plain Flap and Three Tabs. NACA TN No. 761, 1940.
Wind-Tunnel Data on the Aerodynamic Characteristics 4. Sears, Richard I.: of Airplane Control Surfaces. NACA ACR No. 3108, 1943.
Wind+unnel Investigation of an NACA 0009 Airfoil 5. Spearman, M. Leroy: with 0.25~and O.)%+U.rfoil-Chord Plain Flaps Tested Independently and in Combination. NACA TN No. 1517, 1947.
6. Mendelsohn, Robert A.: Wind-T-e1 Investigation of the Boundary Layer and Wake and Their Relation to Airfoil Characteristics - NACA 651-o~ Airfoil with a True Contour Flap and a Bsveled- Trailing-Edge Flap. NACA TN No. 1304, 1947.
7. Wenzinger, Carl J., and Harris, Thomas A.: The VerticalWind Tuulel of the National Advisory Committee for Aeronautics. NACA Rep.
No. 387, 1931.
8. Glauert, H.: Wind Tunnel Interference on Wags, Bodies and Airscrews.
R. & M. No. 1566, British A.R.C., lg33* The Behaviour of a Pitot Tube in a 9. Young, A. D., a;?d Maas, J. N.: Transverse Total-Pressure Gradient. R. & M. No. 1770, British A.R.C., 1937* 10. FaUmer, V. M.: A Further Investi tion of Solutions of the Boundary r Layer Equations. R. &M. No. 184, British A.R.C., 1937.
11. Falkner, V. M.: Stiplified Calculation of the Lminar Boundary R. & M. No. 1895, British A.R.C., 1941.
Layer.
12. von Doe,tioff, Albert E., and Tetemin, Neal: Determination of General Relations for the Behavior of Turbulent Boundary Layers.
NACA ACR No. 3Gl3, 1943.
NACA TN No. 1574 TABLE I ORDINATESFOR NACA 0009 AIRFOIL ,Etations and ordinates in percent of airfoil chord Station Ordinate 0 0 1.25 1.42 2.50 1.96 5.0 2.67 7.5 3.15 10 3.51 15 4.01 20 4.30 25 4.47 30 4.50 40 4.35 50 3*97 60 3.42 70 2.75 80 1.97 90 1.09 95 .61 100 .lO I,. E. kdius: 0.89 I 10 NACA TN No. 1574 NACA TN No. 1574 11 .
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