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CB No. L5JO0 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WA11RTIME REPORT ORIGINALLY ISSUED November 1945 as Confidential Bulletin L5JO4 EFFECTS OF REYNOLDS NUMBER AND LEADINGEDGE ROUGICiESS ON LIFT AND DRAG CHARACTERISTICS OF THE NACA 653418, a = 1.0 AIRFOIL SECTION By John H. Quinn, Jr. Langley Memorial Aeronautical Laboratory Langley Field, Va. .... ACA$ WASHINGTON NACA WARTIME REPORTS are reprints of papers originally issued to provide rapid distribution of advance research results to an authorized group requiring them for the war effort. They were pre viously held under a security status but are now unclassified. Some of these reports were not tech nically edited. All have been reproduced without change in order to expedite general distribution. L 82 DOCUMENTS DEPARTMENT Ij; !1'!' NACA CB No. L5.TOL NATIONAL ADVISORY COMMITTEE FOR AERONAJUTICS C'NFI DETIAL BULLETIN EFFECTS OF REYNOLDS IUM3IBER AND LEADINGEDGE ROUGHNESS Oil LIFT AND DRAG CHARACTERISTICS OF THE I ACA 65z11, a = 1.0 AIRFOIL SECTION By John H. Quinn, Jr. SUJ MAR Y Tests were made in the Langley twodimensional low turbulence tunnels of an !TiCA 653418, a = 1.0 airfoil section with roughness in the form of carborundtum grains applied to the leading edge. Roughness grains having average diameters of 0.0005 and 0.0007 airfoil chord were applied to the leading edge of the :.ing, and lift and drag measurements were made for a range of Reynolds num bers from 0.25 to 3.0 x 100. From a comparison of data obtained in the present tests with data obtained in tests of the smooth wing, marked reductions in maximum lift coefficient were found to be caused by the roughness throughout the test range of Reynolds number. The drag coefficient at the design .ift coefficient increased sharply and the liftcurve slope decreased rapidly at a critical Reynolds number that depended upon the size of the carborundum grains. This critical Reynolds number occurred at approximately 0.50 and 0,70 x 10 for the 0.0005 and the 0.0007chorddiameter roughness grains, respectively. With roughness, a decrease in maximum lift coefficient as great as 0.2, a decrease in liftcurve slope of 0.028, and an increase in drag coefficient at the design lift coefficient ,of C.007 were observed at a Reynolds number of 1.0 x 106. For the smooth wing at the same Reynolds number, the maximum lift coefficient was 1.19, the liftcurve slope was 0.116. and the drag coefficient was 0.0077. At Reynolds numbers greater than 1.0 x 100, the scale effect on the lift and drag charac teristics of the section with both degrees of roughness was generally in the same direction as the effect on the lift and drag characteristics of the smooth airfoil. 2 COiFIDENTIAL NACA CB :D L5J04 INTRODUCTION Several investigations have been made in the past to determine the effects of Reynolds number on the aerody namic characteristics of various airfoil sections. A recent investigation was made (reference 1) to determine the effects of both Reynolds number and stream turbulence on the lift and drag of a smooth "'.CA 6series airfoil section. The present investigation was made to determine the effects of Reynolds number on the lift and drag charac teristics of an NACA 6series section with a roughened leading edge. Tests were made, therefore, of the NACA 653418, a = 1.0 airfoil section in the Langley two dimensional lowturbulence tugnels over a range of Reynolds number from 0.25 to 5.0 x lOb. Lift and drag measure ments were made at several Reynolds numbers in this range with two degrees of roughness applied to the leading edge of the airfoil. Although the data presented herein and in reference 1 are quantitatively correct only for the NACA 655418 air foil section, the effects of Reynolds number and roughness would probably be in the same general direction and of approximately the same order of magnitude for other NACA 65series airfoil sections that do not differ greatly in thickness and camber from the !!.CA 653418. These results are also helpful in properly evaluating the merits of lowscale test data. SYMBOLS cz section lift coefficient cd section drag coefficient ci. design section lift coefficient 0C maximum section lift coefficient c airfoil chorax c airfoil chord COPFTDE:iTIAL NACA CB No. L5JO4 R Reynolds number 00 section angle of attack MODELS AND TEST METHODS The airfoil model used in the present investigation was of 6inch chord and was constructed of aluminum alloy to correspond to the ordinates of the NACA 655418, a = 1.0 airfoil section. Ordinates for this airfoil section are presented in reference 2. A photograph of the model is presented in figure 1. Roughness was simulated by applying carborundum grains of a given diameter to the leading edge of the wing with shellac. The roughness was applied to both surfaces of the airfoil as far back as 0.078c and the grains covered approximately 10 percent of the roughened area. Two degrees of roughness were obtained by use of grains having average diameters of 0.0003c (0.002 in.) and 0.0007c (0.003 to 0.005 in.). The standard roughness used in systematic airfoil investigations (reference 2) for determining the characteristics of various airfoils having transition fixed at the nose is composed of grains having average diameters of 0.3005c. This standard rough ness was thought to be considerably more severe than that caused by any manufacturing, irregularities or poor painting procedures but is not so severe as that caused by icing, mud, or damage from military combat. One grain roughness used in the present tests is larger than the standard roughness; the other is smaller. The tests were made in the Langley twodimensional lowturbulence tunnel (designated LTT) and in the Langley twodimensional lowturbulence pressure tunnel (desig nated TDT). Both tunnels are 3 feet wideand feet high and 2 were designed to test models completely spanning the jet in twodimensional flow. These tunnels are characterized by air streams having exceptionally low turbulence levels, of the order of a few hundredths of 1 percent. Lift measurements were made by an arrangement designed to integrate the pressure reactions along the floor and ceiling of the tunnel test sections and drags were measured by the wakesurvey method. CONFIDENTIAL CONFIDENTIAL NACA CB No. L5JO4 All tests were run at tunnel Mach numbers less than 0.2. Measurements were made at atmospheric pressure in the LTT, and at tunnel pressures Of 1L.7, 30, 45, 63, and 87 pounds per square inch absolute in the TDT. Corrections for the effects of tunnelwall interfer ence and airstream constriction were applied to the model as follows: c, = 0.998 c't cd = 0.999 Cd' where the primed quantities represent the values obtained in the tunnel. RESULTS AND DISCUSSION Lift characteristics for the NACA 653418, a = 1.0 airfoil section at various Reynolds numbers with two degrees of leadingedge roughness are presented in figure 2 and the drag characteristics are presented in figure 3. In figure 2(a), in which data are presented for the model having 0.0003cdiameter grains on the leading edge, a pronounced jog in the lift curve is noticeable at a lift coefficient of 0.9 at a Reynolds number of 0.50 x 106. In reference 5 such a jog was found to be associated with a region of laminar separa tion just behind the leading edge of the wing. The fact that a jog occurred in the present tests indicates that this degree of roughness did not completely eliminate laminar flow at the leading edge until a Reynolds number between 0.50 and 0.75 x 106 was attained. From figures 2(b) and 2(c), in which data are pre sented for the model having 0.0007cdiameter grains on the leading edge, no jog occurred in the lift curve at Reynolds numbers greater than 0.35 x 106. This degree of roughness therefore probably eliminated laminar flow , entirely at a Reynolds number between 0.55 and 0.50 x 106. Curves showing the variation of maximum lift coeffi cient and liftcurve slope with Reynolds number are pre sented in figure 4 and the variation of drag coefficient at the design lift coefficient with Reynolds number, in COIFPIDEITTIAL CONFIDEI'TTAL ITACA CB :o. L5J04 figure 5. Application of roughness to the leading edge of the airfoil caused values of the maximum lift coeffi cient and the liftcurve slooe that were substantially lower than the values for the smooth airfoil (fig. L.). The maximum lift coefficients and the liftcurve slopes are predominantly lower than those for the smooth wing throughout the test range of Reynolds number, and there is a critical Reynolds number at vhich the maximum lift coefficient decreases noticeably and the liftcurve slope decreases rapidly. Figure 5 shows that at this critical Reynolds number a sharp increase in the variation of the drag coefficient at the design lift coefficient with Reynolds number also occurred. The liftcurve slope decreases rapidly and the drag coefficient at the design lift coefficient increases sharply at a Reynolds number of approximately 0.70 x 106 for the 0.0005cdiameter grain roughness and of approximately 0.50 x 10b for the O.0007cdiamneter grain roughness. At a Reynolds number of 1.0 x 100, however, the differences in values of the liftcurve slope and of the drag coefficient for the two degrees of roughness disappear, and at greater Reynolds numbers, within the accuracy of the results, the values of these quantities appear to be independent of the sizes of the roughness for which data are presented. The liftcurve slopes in figure h also show that for the 0.0003cdiameter grain roughness the liftcurve slope is essentially the same as for the smooth wing up to a Reynolds number of at least 0.50 x 10b. This degree of roughness probably brought about n.o significant changes at low lift coefficients in the development of the boundary layer from that existing on the smooth wing up to a Reynolds number of 0.50 x 106. Because the maximum lift coefficient for this degree of roughness ,vas lower than that for the smooth wing throughout the entire range of test Reynolds numbers, the roughness probably did induce some change in the nature of the flow at high lift coefficients. With the 0.0007cdiameter roughness grains, the lift curve slope and maximum lift coefficient were greater than those of the smooth wing at a Reynolds number of 0.25 x 106, but at Reynolds numbers greater than 0.30 x 106 these quantities were lower than those of the smooth airfoil for both degrees of roughness. The reason for this phe nomenon is not readily evident. There is a possibility, however, that at a Reynolds number of 0.25 x o106 the roughness was not large enough to destroy the laminar CONFIDENTIAL CONFIDENTIAL NACA CB No. L5JOi flow entirely but was large enough to prevent laminar separation behind the minimum pressure point. A boundary layer velocity distribution would result, therefore, which would be different from both the smooth flow condition and the smaller roughness condition. Figure shows that at Reynolds numbers between 1.0 and 3.0 x 10O the maximum lift coefficients were less than those of the smooth wing by. approximately 0.14 and 0.20 for the 0.0003c and 0.0007cdiameter roughness grains, respectively. At a Reynolds number of 1.0 x 106, the liftcurve slope of the rough wings was 24. percent less than that of the smooth wing, but at a Reynolds number of 3.0 x 106, a decrease due to roughness of approximately 12 percent in liftcurve slope was found. A constant increment in drag coefficient at the design lift coeffi cient due to roughness of approximately 100 percent was found (fig. 5) at Reynolds numbers between 1.0 and 3.0 x 106. CONCLUSIONS A comparison of results of tests of the NACA 655148, a = 1.0 airfoil section for a range of Reynolds number from 0.23 to 3.0 x 106 with roughness grains having average diameters of 0.0003 and 0.0007 air foil chord (0.0003c and 0.0007c) with results of previous tests of the smooth wing led to the following conclusions: 1. Maximum lift coefficients of the airfoil with roughness were generally lower than those obtained on the smooth airfoil section throughout the test Reynolds number range. At a Reynolds number of 1.0 x 106 the maximum lift coefficient for the smooth wing was reduced from a value of 1.19 to 1.05 and 0.99 by the 0.0005c and 0.0007cdiameter grains, respectively. 2. There is a critical Reynolds number at which the liftcurve slope decreases rapidly and the drag coeffi cient increases sharply depending upon the size of the roughness. This critical Reynolds number was approxi mately 0.70 and 0.50 x 106 for the 0.00053c and 0.0007cdiameter grains, respectively. 3. With roughness, at a Reynolds number of 1.0 x 106, the liftcurve slope was 0.088 and the drag coefficient at the design lift coefficient was 0.0155 whereas the CONFIDENTIAL CONFIDENTIAL NACA CB No. L5Jok corresponding values for the smooth airfoil section were 0.116 and 0.0077, respectively. At Reynolds numbers greater than 1.0 x 10o the changes in liftcurve slope and Crag coefficient were nearly independent of the sizes of the roughness for the two degrees of roughness for which the effects 'were measured. 4. Large variations in the lift and drag character istics of the airfoil were found in the range of Reynolds number between 0.25 and 1.0 x 106. At Reynolds numbers greater than 1.0 x 10', the scale effect on the lift and drag characteristics of the section with both degrees of roughness was generally in the sane direction as the scale effect on the characteristics :f the smooth airfoil. Langley Memorial Aeronautical Laboratory National Advisory Committee for Aeronautics Langley Field, Va. REFERENCES 1. Quinn, John H., Jr., and Tucker, barren A.: Scale and Turbulence Effects on the Lift and Drag Character istics of the IIACA 65 341, a = 1.0 Airfoil Section. NACA ACR No. L4H11, 1J44. 2. Abbott, Ira H., von DocEnhoff, Albert E., and Stivers, Louis S., Jr.: Surimary of Airfoil Data. NACA ACR No. L5C05, 1945 3. von Doenhoff, Albert E., and Tetervin, Neal: Investi gation of the Variation of Lift Coefficient with Reynolds Number at a Moderate Angle of Attack on a LowDrag Airfoil. IACA CB, Nov. 1942. CONFIDENTIAL CONFIDENTIAL Digitized by the Internet Archive in 2011 with funding from University of Florida, George A. Smathers Libraries with support from LYRASIS anc Ihe Sloan Foundation http://www.archive.org/details/effectsofreynold001ang NACA CB No. L5J04 Fig. 1 0 bD . o* S d 00 (D O 0 04. LO "a 0) SCo m I *.4 (D Q o o 4 WW . Cl .0 ( 0 .3 E 0) Ocd I C = 40 00 l 0 a) 0 4 00 . O. NACA CB No. L5JO4 S*0 rj u~M.o OO C, 0 0 o r.,' rj \ 0 0 00 21 0O TX '*0 Ti 21 u 0 0 a a '0 t 0 43 u b o r 1' o o u cc) ld U a. S O0 S I. 0 41 0 ' a a CO a2 a a o C i . .' TI 4U0 C' 41 eJ I J .. Li '. 10 ',u8oTJjeoo D IJIT uoqo9s Fig. 2a NACA CB No. L5J04 0 d , t0 0 4 0 S a a 4d 4g oo ra 0 4 co H S 0 0 0 0 HO co E ' 0 0 > 1 . 0 CO o H0 '* 3 00 04,G 0 01 v o >& co 0 1o '4ueTOTJjeoo 4JTT uo0Toe Fig. 2b MX o w x NACA CBNo. L5J04 "C! 0 14 U'Ou0 0 00 ido '4o'J ooo0 0 f o 0 Ti a a C. 0 0 t. c00 u 0 2) o  .2 Q CJ 0 3 2  oc o g o 0 0 E0 C 02 0.25 0 1 I  0 ' 0 1a 'uaioGTjjGeo 2ITT uaGQme5 Fig. 2c NACA CB No. L5J04 R  0.25 x 106 50 .75 1.00 1.50 2.00 2.50 5.00 CONFIDENTIAL iIIE __ _I _ _... x..__^ ^ ^ _ 4     NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS I I I 40 .4 .:8 1.2 Section lift coefficient, cL (a) Model having 0.0003cdiameter grains applied to leading edge. 3DT test 867. Figure 3. Variation of section drag coefficient cd with section lift coefficient c0 for the NACA 653418 airfoil section having two degrees of leadingedge roughness. Model chord, 6 inches. .0 zA ~ .052 .0281 .02( .01 o 00 .001. CONFIDENTIAL I I I u~~~4T Fig. 3a j* NACA CB No. L5J04 .4 U .4 * Section lift coefficient, c, (b) Model having 0.0007cdiameter grains applied to leading edge. TDT test 873. Figure 3. Concluded. .036 .032 .028 .024o .020 .016 .012 .008 .00o Fig. 3b NACA CB No. L5J04 o0 I 0 000 0\ 2 .s 0o 000 o0 EO S1u 44 i z 0 I co 0o 0 to S. 0 o OA 0 k mq 029 *H OS.. o \ t 0 6 a 3 0 CC S o 0 0 a WA > r Z 0 o 0 Rg 0 43I 44 14  q.02 I C d 0 P a I> H Fig. 4 Fig. 5 NACA CB No. L5J04 _ ____ _____ 1_ L4) __ ~>1 _____ ____ I 4 Zu f ~2 u JRO z z z 0 _ti __ U o I * 00 00 0 I; a I i  I + z a < z 0 S0 o 0 A. oM4  Ca rqo 0 00 * 0p M 4 4 0 43 0 r d * 0 L4 0 m 0 0 So0 'c 0 r 43 u 0 0 L0 0 0 0 H D4 a 'i' o a 1~I 0 El 0  UNIVERSITY OF FLORIDA 3 1262 07749 275 8 U ,,VERSTy OF FLORIDA DOCU, S ,.TS DEPARTMENT ".rSTO SCENCELUBRARy '. X 1170 "1 C ~I~LSVILLE, FL 32117011 USA 