Document
NATIONAL ADVISORY COMMITTEE
FOR AERONAUTICS
REPORT No. 824
SUMMARY OF AIRFOIL DATA
By IRA H. ABBOTT, ALBERT E. VON DOENHOFF, and LOUIS S. STIVERS, Jr.
AERONAUTIC SYMBOLS
1. FUNDAMENTAL AND DERIVED UNITS Metric English Symbol Abbrevia-- Abbrevia- Unit Unit tion tion meter __________________ Length ______ foot (or mile) _________ l m ft (or rni) second _________________ Time _____ ___ second (or bour) _______ t s sec (or hr) Force ___ _____ kilogram _____ weight of 1 pound _____ F weight of) kg lb Power _______ horsepower ___________ horsepower (metric) _____ P hp
----------
miles per hOuL _______ {kilometers per hour ______ kph mph Speed _______ V meters per second _______ mps feet per second ________ fps 2. GENERAL SYMBOLS
Weight=mg JI Kin ematic viscosity
w
Standard acceleration of gravity=9.80665 m/s p Density (mass per unit volume)
u
4 2
or 32.1740 ft/sec Standard density of dry air, 0.12497 kg_ m- _s at 15° C
and 760 mm; or 0.002378 Ib-ft-4 sec
Mass=W
m
s
Specific weight of "standard" air, 1.2255 kg/m or
g
0.07651 lb/cu ft
Moment of inertia=mP. (Indicate axis of
I
radius of gyration k by proper subscript.)
Coefficient of viscosity
3. AERODYNAMIC SYMBOLS
Area
s Angle of setting of wings (relative to thrust line)
Area of wing
S~ Angle of stabilizer setting (relative to thrust
line)
Gap
G
Resultant moment
b Span o
Resultant angular velocity
c Chord
b'
A Aspect ratio, S R
Reynolds number, p Vl wherelisalineardimen-
fJ.
V True air speed
sion (e.g., for an airfoil of 1.0 ft chord, 100 mph,
standard pressure at 15° 0, the corresponding
q Dynamic pressure, ~P V'
Reynolds number is 935,400; or for an airfoil
of 1.0 m chord, 100 mps, the corresponding
L
Lift, absolute coefficient OL= q~
Reynolds number is 6,865,000)
Angle of attack
D Drag, absolute coefficient OD= q~
Angle of downwash
Angle of attack, infinite aspect ratio
Profile drag, absolute coefficient ODO=~
Angle of attack, induced
Angle of attack, absolute (measured from zero-
Induced drag, absolute coefficient OD = ~~
j qu
lift position)
Flight-path angle
'Y
Parasite drag, absolute coefficient ODP= ~S
D.
o Cross-wind force, absolute coefficient 0 = q~
REPORT No. 824
SUMMARY OF AIRFOIL DATA
By IRA H. ABBOTT, ALBERT E. VON DOENHOFF, and LOUIS S. STIVERS, Jr.
Langley Memorial Aeronautical Laboratory Langley Field, Va.
I
National Advisory Committee for Aeronautics
Headquarters, 1500 New Hampshire Avenue NW., Washington 25, D. O.
Created by act of Congress approved March 3, 1915, for the supervision and direction of the scientific study of the problems of flight (U. S. Code, title 49, sec. 241). Its membership was increased to 15 by act approved March 2, 1929. The members are appointed by the President, and serve as such without compensation.
JEROME C. HUNSAKER, Sc. D., Cambridge, Mass., Chairman LYMAN J. BRIGGS, Ph. D., Vice Chairman, Director, National AUBREY W. FITCH, Vice Admiral, United States Navy, Deputy Chief of Naval Operations (Air), Kavy Department.
Bureau of Standards.
CHARLES G. ABBOT, Sc. D., Vice Chairman, Executive Committee, WILLIAM LITTLEWOOD, M. E., Jackson Heights, Long Island, Secretary, Smithsonian Institution.
N. Y.
HENRY H. ARNOLD, General, "Gnited States Army, Commanding FRANCIS W. REICHELDERFER, Sc. D., Chief, United States General, Army Air Forces, War Department.
Weather Bureau.
WILLIAM A. M. Bl:RDEN, Assistant Secretary of Commerce for LAWRENCE B. RICHARDSON, Rear Admiral, United States Navy, Aeronautics.
Assistant Chief, Bureau of Aeronautics, Navy Department.
VANNEVAR BUSH, Sc. D., Director, Office of Scientific Research EDWARD 'VARNER, Sc. D., Civil Aeronautics Board, Washington, and Development, Washington, D. C.
D. C.
WILLIAM F. DCRAND, Ph. D., Stanford Lniversity, California.
ORVILLE WRIGHT, Sc. D., Dayton, Ohio.
OLIVER P. ECHOLS, !\iajor General, "Cnited States Army, Chief of Materiel, Maintenance, and Distribution, Army Air Forces, THEODORE P. WRIGHT, Sc. D., Administrator of Civil Aero- War Department. nautics, Department of Commerce.
GEORGE W. LEWIS, Sc. D., Director of Aeronautical Research JOHN F. VICTORY, LL. M., Secretary HENRY J. E. REID, Sc. D., Engineer-in-Charge, Langley Memorial Aeronautical Laboratory, Langley Field, Va.
SMITH J. DEFRANCE, B. S., Engineer-in-Charge, Ames Aeronautical Laboratory, Moffett Field, Calif.
EDWARD R. SHARP, 1,1,. B., Manager, Aircraft Engine Research Laboratory, Cleveland Airport, Cleveland, Ohio CARLTON KEMPER, R. S., Executive Engineer, Aircraft Engine Research Laboratory, Cleveland Airport, Cleveland, Ohio
TECHNICAL COMMITTEES
AERODYNAMICS OPERATING PROBLEMS POWER Pr,ANTS FOR AIRCRAFT MATERIALS RESEARCH COORDINATION AIRCRAFT CONSTRUCTION Coordination of Research Needs of Military and Civil Aviation Preparation of Research Programs Allocation of Problems PrevenUon of Duplication LANGLEY MEMORIAL AERONAUTICAL LABORATORY AMES AERONAUTICAL LABORATORY Langley Field, Va. l\Ioffett Field, Calif.
AIRCRAFT ENGINE RESEARCH LABORATORY, Cleveland Airport, Cleveland, Ohio Conduct, under unified control, for all agencies, of scientific research on the fundamental problems of flight OFFICE OF AERONAUTICAL INTELLIGENCE, Washington, D. C.
Collection, classification, compilation, and diss~'minatiori of scientific and technical information on aeronauticll II
CONTENTS
Page Page SUMMARY _______ .. _- - - __ - - . __ - ____ - - __ - - __ ., . _ . _ .. _ .. _____ _ EXPERIMENTAL CHARACTERISTIcs-Continued 1NTRoDucTION_ . ________________________ .. _"_C ___ .. _______ 1 Drag Chara~teristics of Smooth Airfoils-·Continued SYMBOLS ________________________ ._______________________ 1 Effects of. type of sectiOIl on drag charact.eristics .. __ _ _ 18 HIflTORICAL DE~·ELOPMENT .. ________________________ .. _ _ __ _ _ 2 Effective aspect ratio ________. _________________ .. ___ 21 DESCRIPTION OF AIRFOILS ____________________________ .. __ _ _ 3 Effect of surface irregularities on drag ____ . _____ _ __ _ _ __ _ 22 Permissible roughness ________ .. _________ __ _ _ __ __ __ _ 22 :\fethod of Combining :.\Iean Lines and Thickness Distributions ______. ______________ .__________ ______ 3 Permissible waviness ______________________ ._ _ __ _ _ __ 22 NACA Four-Digit Series Airfoils ________________ .. ______ 4 Drag with fixed transition ___ . ___ .. _________________ 24 Numbering system_ _ _____ _ _____ __ _ __ __ _ _______ _ _ _ 4 Drag with practical construction methods_ _ __ _ __ _ _ _ _ 24 Thickness distributions ________ . _ _ __ ___ __ _ __ _ __ _ _ _ 5 Effects of propeller slipstream and airplane vibration_~ 29 Mean lines ____________________________. _________ ~ 5 Lift Characteristics of Smooth Airfoils_ ________ _ _ __ __ __ _ 30 NACA Fh'e-Digit Series Airfoils ____ .... _____________.___ _ 5 Two-dimensional dat9. __________ .. ________ __ _ __ _ _ _ _ 30 Numbering system__ _ _ __ _ __ __ _ __ _ ___ __ _ __ _ __ ____ _ 5 Three-dimensional data_ _ _________________ _ _ _ __ _ _ _ 37 Thickness distributions _________________ -~_c ____ ~c-_ "5 - Lift Characteristics of Rough Airfoils. _________ _ __ _ __ __ _ 37 Mean lines ______________________ .. _ ___ _ __ __ _ __ _ _ 5 Two-dimensional data_ ___________ _______ __ __ __ _ _ _ 37 Three-dimensional data ______ . ____ . ____ .. ______ .. _ _ __ 38 N ACA I-Series Airfoils ____ ... ____ . ______________ "~_-=-=- 5 Numbering system ___________ ~________________ ___ 5 Unconservatiye Airfoils ________ .... __ . _____ ... ____________ 39- Thickness distributions _________________ . __________ 5 Pitching Moment ____________________ . ___________ .. _ __ 4() Mean lin es _______. __ _ __ _ ___ __ _ __ _ _ _ __ __ _ __ _ __ _ _ _ 5 Position of Aerodynamic Center __ . ______________ . __ .. ___ 43: High-Lift Devices_ __ __ __ ___ ___________ __ __ __ __ __ _ ___ _ 43: NACA 6-Series Airfoils________________________________ 5 N um bering system __ __ __ _ _ ____ _ _ __ _ _ ___ _ __ __ _ _ __ _ 5 Lateral-Control Devices ___________________ .. __ __ _ _ __ _ _ _ 43 Leading-Edge Air Intltkes __________ .. _____ __ __ __ _ _ __ __ _ 49 Thickness distribu tions _ _ _ _ __ __ __ _ _ __ __ __ _ __ _ _ ___ _ 6 Mean lines ___ ~ _____________________ .. _____ .. ______ 6 In terference __ .. _________________________ .. ____________ . 50 NACA 7-Series Airfoils __________ ~· ___ :.:_= _____________ ~_ 7 ApPI,ICATION TO WINO DESIGN __________ . ___ .. ______________ 51 NUmcering system_ _ _____ ___ __ ___ _____ __ ____ __ ___ 7 Application of Section Data __________________________ .. 51 Thir,kness distributions_ _ __ __ __ __ __ __ _ _ ___ __ _ _ _ __ _ 7 Selection of Root Section ___ ._ _______________ __ _ _ __ ___ _ _ 51 Selection of Tip Section ____ .. ________________ ... _ __ __ _ __ 52 THEORETICAL CONSIDERATIONS __________ . ______ .. _________ .. 8 CONCLUSIONS _________________ .. _________________ .. _ _ _ __ _ _ _ 52 Pressure Distributions _______________ c _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ 8 ApPENDIX-- METHODS OF OBTAINING DATA IN THE LANGLEY Methods of derivation of thickness distributions_ _ _ _ 8 Two-DIMENSIONAL Low-TURBULENCE TUNNELS_ __________ 54 Rapid estimation of pressure distributions ____ ~ __ _ _ __ 10 Description of Tunnels___ ___ ______________ __ __ __ __ _ __ 54 Numerical examples_ __ _ _ __ __ __ __ _ _ _ __ _ ___ _ __ __ __ _ 12 Symbols ___ .. _____________ . _ ___________ ____ _ _ __ __ _ _ _ 54 Effect of camber on pressure distribution ___ .. _____ .. _ _ 13 Measurement of Lift.. ______ . ______ ... _______ __ _ _ __ __ _ __ 55 Critical Mach Number___ __ __ __ __ __ _ _ __ __ __ __ __ ___ _ __ _ 13 Measurement of Drag .. ______________________________._ 56 Moment Coefficients___ __ __ _ ___ __ __ __ ___ ___ __ ___ _____ _ 14 Tunnel-Wall Corrections _____ .. ______ _________ __ _ __ _ _ _ _ 57 Methods of calculation ____ . _______ . ______ .. __ ._____ 14 Correction for Blocking at High Lifts ______ . ___ __ _ _ __ __ _ 59 Numerical exapl pIes _ . _ _ _ __ __ _ _ __ __ __ _ ___ __ _ _ __ __ _ 14 Comparison w.ith Experiment.. ____________________ .. __ _ _ 59 Angle of Zero Lift_ _______ __ _ __ _ ___ __ _ _ __ __ __ __ _ __ __ _ _ 14 REFERENCES _____________ .. ______________________________ 60 Methods of calculation ______________________ .... _ _ _ _ 14 TABLES ___________________ .... ___ .. ________________________ 64 Numerical examples______________________________ 14 SUPPLEMENTARY DATA: Description of Flow around Airfoils ___ ~-~~~-:c __ - ___ ~______ 15 I-Basic Thickness Forms _______ .. ____ . _. ___ . _____ . _ ___ 69 EXPERIMENTAL CHARACTERISTICS_ __ __ _ _ __ _ _ __ __ __ __ __ _ __ _ _ 16- J.I-Data for Mean Lines_____________________________ 89 Sources of Data _______ .. __________________________ . __ _ 16 III::""':Airfoil Ordinates_ ________ __ ______ __ __ __ _ __ _ _ __ _ _ 99 Drag Characteristics of Smooth Airfoils ________ .. __ _ __ __ _ 16 IV-Predicted Critical Mach Numbers ______ .__________ 113 Drag characteristics in low-drag range __ .. ______ . __ _ 16 V-Aerodynamic Characteristics of Various Airfoil Sections___ ________________ __ __________ _ _ __ _ _ _ 12~ Drag characteristics outside low-drag range_ __ __ _ ___ _ 18 III
REPORT No. 824
SUMMARY OF AIRFOIL DATA
/, By IRA H. ABBOTT, ALBERT E. VON DOENHOFF, and LOUIS S. STIVERS, JR.
SUMMARY Recent information on the aerodynamic characteristics of NACA airfoils is presented. The historical development of Recent airfoil data for both flight and 'wind~tunnel tests have NACA airfoils is briefly reviewed. New data are presented been collected and correlated insojar as possible. The flight that permit the rapid calculation of the approximate pressure data consist largely of drag measurements made' by the wake~ distributions for the older NACA four-digit and five-digit survey method. Most of~he data on airjoil section characteris~ airfoils by the same methods used for the N ACA 6-series tics were obtained in the Langley two-dimensionallow~turbulence airfoils. The general methods used to derive the basic thick- pressure tunnel. Detail data necessary for the application of ness forms for N ACA 6- and 7-series airfoils together with NAOA 6~series airfoils to wing design are presented in sup- their corresponding pressure distributions are presented.
plementary figures, together with recent datajor the NAOA 00-, Detail data necessary for the application of the airfoils to
14-, 24-, 44-, and 230-series airjoils. The general methods
wing design are presented in supplementary figures placed at used to derive the basic thickness jorms jor NAOA 6- and the end of the paper. The report includes an analysis of 7 -series airjoils and the'ir corresponding pressure distributions the lift, drag, pitching~moment, and critical-speed charac:' are presented. Data and methods are given jor rapidly obtain- teristics of the airfoils, together with a discussion of the ing the approximate pressure distributions jor N AOA four- effects of surface conditions. Available data on high-lift digit, five-digit, 6-, and 7-series airfoils.
devices are presented. Problems associated with lateral- The report includes an analysis oj the lift, drag, pitching- control devices, leading-edge air intakes, and interference moment, and critical-speed characteristics of the airfoils, to- are briefly discussed, together with aerodynamic. problems gether with a discussion of the effects of surface conditions.
of application.
Data on high-lift devices are presented. Problems associated Numbered figures are used to illustrate the text and to with lateral-control devices, leading-edge air intakes, and inter- present miscellaneous data. Supplementary figures and ference are briefly discussed. The data indicate that the effects tables are not numbered but are conveniently arranged at oj surface condition on the l~ft and drag characteristics are at the end of the report according to the numerical designation least as large as the effects of the airfoil sha,pe and must be of the airfoil section within the following headings: considered in airfoil selection and the prediction of wing charac- I-Basic Thickness Forms teristics. Airjoils permUting extensive laminar flow, such as II-Data. for Mean Lines the NAOA 6-series airfoils, have much lower drag coefficients III-Airfoil Ordinates at high speed and cru~~sing lift coefficients than earlier types of IV--Predicted Critical Mach Numbers airfoils if, and only if, the wing surfaces are suffic1~ently smooth V-Aerodynamic Charactel'is tics of Various Airfoil and fair. The NAOA 6-scries airfoils also ha,1'e favorable Sections crit?:cal-speed character'istics and do not appear to present These supplementary figures and tables present the basic 1Lnu8ual problems associated with the applicat1:on oj high-l'i:ft data for the airfoils.
and lateral-control devices.
SYMBOLS INTRODUCTION aspect ratio A considerable amount of airfoil data has been accumulated Fourier series coefficients from tests in the Langley two-dimensional low-turbulence mean-line designation, fraction of chord from lead- tunnels. Data ha,ve also been obtained from tests both in ing edge over which design load is uniform; in other wind tunnels and in flight and include the effects of derivation of thickness distributions, ba,sic length high-lift devices, surface irregularities, and interference. usually considered unity Some data are also available on the effects of ai.rfoil section wing span flap span, inboard on aileron characteristics. Although a large amount of these flap span, outboard data has been published, the scattered nature of the data drag coefficient and the limited objectives of the reports have prevented drag coefficient at zero lift adequate analysis and interpretation of the results. The lift coefficient purpose of this report is to summarize these data and to increment of maximum lift cuused by flap deflection correlate and interpret t,hem insofar as possible.
REPORT NO. 824-·NATIONAL ADVISORY COMMITTEE FOR AERONAUTiC;:, C chord XL abscissa of lower surface C aileron chord Xu absciss!'\. of upper surface a section drag coefficient Cd chordwise position of transition G},.
minimum section drag coefficient Cdml~ distance perpendicular to chord Y flap chord, inboard Cfi mean-line ordinate Ya flap chord, outboard CfO ordinate of lower surface
YL
flap-chord ratio ordinate of symmetrical thickness distrihution YI C ordinate of upper surface Yu
section aileron hinge-moment coefficient (~)
Z goc complex variable in circle plane
increment of aileron hinge-moment coefficient at z'
complex variable in near-cireIe plane constant lift a angle of attack !lCHO hinge-moment parameter !lao section aileron effectiveness parameter, ratio of .10 Cl section lift coefficient change in section angle of attack to increment of design section lift coefficient Cli aileron deflect,ion at a constant value of lift moment coefficient about aerodynamic center C ma . e.
coefficient moment coefficient about quarter-chord point C me/ angle of zero lift C section normal-force coefficient n section angle of attack
D
drag increment of section angle of attack !lH . loss of total pressure section angle of attack corresponding to design
Ho free-stream total pressure
lift coefficient h section aileron hinge moment flap 01' aileron deflection; down deflection is positi,-e exit height he flap deflection, inboard
k constant
flap deflection, outboard L lift
i\,irfoil parameter (IP-()
M Mach number
value of E at trailing edge
Mer critical Mach number
complex variable in airfoil plane
s-
OU,OL typical points on upper and lower surfaces of airfoil
O angular coordinate of z'; also, angle of which tangent
p
pressure coefficient (P-Po)
is slope of mean line go . (TiP chord) critical pressure coefficient taper ratIO Root chord resultant pressure coefficient; difference between Effective Reynolds number) local upper- and lower-surface pressure coefficients T t b I f t ( ur u ence ac or Test Reynolds number local static pressure; also, angular velocity in roll in
P
pb/2V angular coordinate of z
free-stream static pressure airfoil parameter determining radial co.)rdinato of z
Po
pb/2V helix angle of wing tip
average value of 1ft (~17r 50 .. 1ft dIP)
f}o free-stream dynamic pressure
R Reynolds number
critical Reynolds number
Rer
HISTORICAL DEVELOPMENT
s
pressure coefficient (H~o P)
The development of types of NACA airfoils now in com- mon use was started in 1929 with a systematic investigation first airfoil thickness ratio of a family of airfoils in the Langley variable-density tunnel.
second airfoil thickness ratio Airfoils of this family were designated by numbers having free-stream velocity four digits, such as the NACA 4412 airfoil. All airfoils of inlet velocity this family had the same basic thickness distribution (refer- local velocity ence 1), and the amount and type of camber was systemati- increment of local velocity cally varied to produce the family of related airfoils. This increment of local velocity caused by additional type of load distribution investigation of the NACA airfoils of the four-digit series produced airfoil sections having higher maximum lift velocity ratio corresponding to thickness i1 coefficients and lower minimum drag coefficients than those of sections developed before that time. The investigation
velocity rat.io corresponding to thickness t2
also provided information on the changes in aerodynamic distance along chord characteristics resulting from variations of geometry of the mean-line abscissa mean line and thickness ratio (reference 1).
SUMMARY OF AIRFOIJ" DATA The investigation was extended in references 2 and 3 to was obtained by empirical modification of the previously include airfoils with the same thickness distribution but used thickness distributions (reference 4). These NACA with positions of the maximum camber far forward on the 16-series sections represented the first family of the low-drag airfoil. These airfoils were designated by numbers having high-critical-speed sections.
five digits, such as the NACA 23012 airfoil. Some airfoils Successive attempts to design airfoils by approximate of this family showed favorable aerodynamic characteristics theoretical methods led to families of airfoils designated except for a large sudden loss in lift at the stall. N ACA 2- to 5-series sections (reference 11). Experience with .Although these investigations were extended to include a these sections showed that none of the approximate methods limited number of airfoils with varied thickness distribu- tried was sufficien tly accurate to show correctly the effect
tions (references 1 and 3 to 6), no extensive investigations of
of changes in profile near the leading edge. Wind-tunnel
thickness distribution were made. Comparison of experi- and flight tests of these airfoils showed that extensive laminar mental drag data at low lift coefficients with the, skin- boundary layers could be maintained at cOplparatively large friction coefficients for flat plates indicated that nearly all values of the Reynolds number if the airfoil surfaces were of the profile drag under such conditions was attributable sufficiently fair and smooth. These tests also provided to skin friction. It was therefore apparent that any pro- qualitative information on the effects of the magnitude of nounced reduction of the profile drag must be obtained by a the favorable pressure gradient, leading-edge radius, and other reduction of the skin friction through increasing the relative shape variables. The data also showed that separation of extent of the laminar boundary layer. the turbulent boundary layer over the rear of the section, Decreasing pressures in the direction of flow and low air- especially with rough surfaces, limited the extent of laminar stream turbulence were known to be favorable for laminar layer for which the airfoils should be designed. The air- flow. An attempt was accordingly made to increase the foils of these early families generally showed relatively low relative extent of laminar flow by the development of air- maximum lift coefficients and, in many cases, were designed foils having favorable pressure gradients over a greater for a greater extent of laminar flow than is practical. It was proportion of the chord than the airfoils developed in refer- learned that, although sections designed for an excessive ences 1, 2, 3, and 6. The actual attainment of extensive extent of laminar flow gave extremely low drag coefficients laminar boundary layers at large Reynolds numbers was a near the designJift coefficient when sm09th, the drag of such previously unsolved experimental problem requiring the sections became unduly large when rough, particularly at lift development of new t.est equipment with very low air- coefficients higher than the design lift. These families of stream turbulence. This work was greatly encouraged by airfoils are accordingly considered obsolete.
the experiments of Jones (reference 7), who demons~rated The NACA 6-series basic thickness forms were derived by the possibility of obtaining extensive laminar layers in flight new and improved methods described herein in the section at relatively large Reynolds numbers. Uncert.ainty with "Methods of Derivation of Thick.9.ess Distributions," in ac- regard to factors affecting separation of the turbulent cordance with design criterions established with the objective boundary layer required experiments to determine the of obtaining desirable drag, critical Mach number, and possibility of making the rather sharp pressure recoveries maximum-lift characteristics. The present-report deals largely required over the rear portion of the new type of airfoil. with the characteristics of these sections. The develop- New wind tunnels were designed specifically for testing ment of the NACA 7-series family has also been started.
airfoils under conditions closely approaching flight condi- This family of airfoils is characterized by a greater extent of tions of air-stream turbulence and Reynolds number. The laminar flow on the lower than on the upper surface. These resulting wind tunnels, the Langley two-dimensional low- slilctions permit low pitching-moment coefficients with mod- turbulence tunnel (LTT) and the Langley two-dimensional erately high design lift coefficients at the expense of some low-turbulence pressure tunnel (TDT), and the methods reduction in maximum lift and critical Mach number.
Acknowledgement is gratefully expressed for the expert
used for obtaining and correcting data are briefly described
in the appendix. In these tunnels the models completely guidance and many original contributions of Mr. Eastman span the comparatively narrow test sections; two- N. Jacobs, who initiated and supervised this work.
dimensional flow is thus provided, which obviates difficulties
previously encountered in obtaining section data from
DESCRIPTION OF AIRFOILS
tests of finite-span wings and in correcting adequately for
support interference (reference 8). METHOD OF COMBINING MEAN LINES AND THICKNESS DISTRIBUTIONS
Difficulty was encountered in attempting to design air-
The cambered airfoil sections of all N ACA families con-
foils having desired pressure distributions because of the lack
of adeql.late theory. The Theodorsen method (reference 9), sidered herein are obtained by combining a mean line and a as ordinarily used for calculating the pressure distributions thickness distribution. The, necessary geometric data and about airfoils, was not sufficiently accurate near the leading some theoretical aerodynamic data for the mean lines and
thickness distributions may be obtained from the supple-
edge for prediction of the local pressure gradients. In the
absence of a suitable theoretical method, the 9-percent- mentary figures by the methods described for each family of thick symmetrical airfoil of the N ACA 16-series (reference 10) airfoils.
REPORT NO. 824-NATIONAL ADVISORY COMMITTEE FOR AERONAU'fICS y Mean line
--- ----
---
Chord Ime
---
I I I \ ::OL(:X:L-)-:Y,;L~)-----~------------- Xv =x-Y sin 8 Yu=Yc+y, cos 8 t \ XL =x+Y, sin 8 \ YL =Yc -Yt cos 8 \, Rodius fhrou9h end of chord 1.00 '(mean-line slope ot 05 percent chord) SAMPLE CALCULATIONS FOR DERIVATION OF THE KACA 65,3-818, a=1.0 AIRFOIL 11, 11' cos 0 YI sin 0 y, cos 0 X tan 0 sin 0 Xu 1/U XL 1!L (0) (b)
I
I
0 0 0 0 0 0 0 0 0 ---------- '6:94765-' "6:3i932' .-01324 ;"(;0200 .00423 .01255 .ooon .01455 .00923 -.01055 -.005 ' 0.33696 .18422 .98288 .00706 .03765 -.02501 .05 .03831 .01264 .18744 .04294 .05029 . C5706 .25 .08093 .03580 .06996 .06979 .99756 . 00565 .08073 .24435 .11653 .25565 -.04493 .50 .08593 .04412 0 1.00000 0 .08593 .50000 .13005 .50000 -.04181 -.06979 .99756 -.00311 .04445 .75 .04456 .03580 -.06996 .75311 .08025 .74689 -.00865 1.00 0 a a 0 ---------- 1.00000 0 1. 00000 a ---------- ---------- o Thickness distribution obtained from ordinates of the N A OA 65,3--018 airfoil.
b Ordinates of the mean line, 0.8 of the ordinate for c',= 1.0.
, Slope of radius through end of chord.
FIGURE I.-Method of combining mean lines and basic thickness forms.
The process for combining a mean line and a thickness. of the leading-edge point. Because the slope at the leading distribution to obtain the desired cambered airfoil section is edge is theoretically infinite for the mean lines having a illustrated in figure 1. The leading and trailing edges are theoretically finite load at the leading edge, the slope of the defined as the forward and rearward extremities, respectively, radius through the end of the chord for such mean lines is of the mean line. The chord line is defined as the straight
usually taken as the slope of the mean line at ~=0.005. This
line connecting the leading and trailing edges. Ordinates of
c
the cambered airfoil are obtained by laying off the thickness procedure is justified by the manner in which the slope distribution perpendicular to the mean line. The abscissas,
increases to the theoretically infinite value as x/c approaches
ordinates, and slopes of the mean line are designated as Xc,
o. The slope increases slowly until very small values of x/c
Yc, and tan (J, respectively. If Xu and Yu represent, respec- are reached. Large values of the slope are thus limited to tively, the abscissa and ordinate of a typical point of the
values of x/c very close to 0 and may be neglected in practical
upper surface of the airfoil and Y t is the ordinate of the airfoil design.
symmetrical thickness distribution at chordwise position X, Tables of ordinates are included in the supplementary data the upper-surface coordinates are given by the following for all airfoils for which standard characteristics are presented.
relations: NACA FOUR-DIGIT-SERIES AIRFOILS (1) xu=X-Yt sin (J Numbering system.-The numbering system for the (2) NACA airfoils of the four-digit series (reference 1) is based on the airfoil geometry. The first integer indicates the The corresponding expressions for the lower-surface coordi- maximum value of the mean-line ordinate Yc in percent of the nates are chord. The second integer indicates the distance from the leading edge to the location of the maximum camber in (3) tenths of the chord. The last two integers indicate the airfoil thickness in percent of the chord. Thus, the NACA (4) 2415 airfoil has 2-percent camber at 0.4 of the chord from the The center for the leading-edge radius is found by drawing leading edge and is 15 percent thick.
a line through the end of the chord at the leading edge with The first two integers taken together define the mean line.
the slope equal to the slope of the mean line at that point for example, the N ACA 24 mean line. The symmetrical air- and laying off a distance from the leading edge along this line foil sections representing the thickness distribution for a equal to the leading-edge radius. This method of construc- family of airfoils are designated by zeros for the first two tion causes the cambered a.irfoils to p.roject slightly forward integers, as in the case of the N ACA 0015 airfoil.
DATA OF AIRFOIL SUMMARY thickness distributions for distributions.--The Thickness Thickness distributions.---Data for the NACA 0006,0008, same as those the the N ACA five-digit series are thickness airfoils of and 0024 0018, 0021, 0012, 0015, 0010, 0009, figures_ for airfoils of the NACA four-digit series.
supplementary in the are presented distributions for the NACA 210, 220, 230, 240, and lines.-Data be obtained Mean may thicknesses for intermediate Ordinates in the supplementary figures lines are presented 250 mean tabulated ordinates in proportion to scaling the correctly by the lines given herein for mean radius in the same form as for the The leading-edge the thickness ratio (reference 1).
line mean values for each All tabulated Values of four-digit series.
ratio.
of the thickness the square varies as or with the design the maxImum ordinate linearly with vary (vIV)2, which is equivalent to the low-speed pressure distri- for the NACA 430 mean line Thus, data data were lift coefficient.
These also presented.
of vlV are
bution, and NACA 230 for the the data by multiplying ma,y be obtained Values of (reference 9).
by Theodorsen's method obtained the NAOA 640 mean line 4:2 and for by the ratio 01 mea,n line induced by changing angle
t::.va/F
the velocity increments by the NACA 240 mean line data for multiplying the by of Pressure Distribu- Estimation "Rapid attack (see section 6: 2.
of the ratio an additional lift coefficient tions") are also presented for l-SERIES AIRFOILS NACA for v/V Values of the velocity ratio approximately unity.
approxi- may be obtained NACA I-series airfoils are des- ratios systern.-The thickness Numbering intermediate the example, for velocity increments obtained 'number-as, the by a five-digit by linear scaling of ignated mately the represents thickness first integer
v/V for the nearest The
values of 16-212 section.
tabulated NACA from the dis- the The second integer indicates designation.
series ratio; thus, the to the leading edge the chord from tenths of tance in the symmetrical section (5) position of minimum pressure for the dash indicates following first number The at zero lift.
the design lift in terms of of camber expressed the amount together be obtained two numbers
!::.Va/V may and the last
in tenths, the velocity-increment ratio coefficient Values of The com- the chord_ interpolation.
by thickness in percent of indicate the for intermediate thicknesses have minimum pressure and 62,63,64,65,66, monly used sections of this family for the NACA lines.-Data Mean are usually and the leading edge figures. the chord from the supplementary at 0.6 of in lines are presented mean the yo, the NACA 16-seI'ies sections.
mean-line ordinates referred to as the presented include The data 16-006, for the NACA and the corre- distributions.-Data eli Thickness coefficient the design lift slope dYeldx, thickness and 16-021 16-018, coefficient 16-012, 16-015, ai, the moment 16-009, attack sponding design angle of the supplemen- 10) are presented in and the velocity
P distributions (reference
pressure coefficient R , ' the resultant
c
mei4 for the data data are similar in form to characteristics figures. These aerodynamic tary theoretical
!::.v/V. The
ratio for and data values the N ACA four-digit series, All tabulated those airfoils of were obtained from thin-airfoil theory.
the same obtained in may be the maxi- linearly with vary intermediate thickness ratios for each mean line, accordingly, with mean lines for similar manner.
Ye, and data ordinate mum as commonly may be NACA 16-series airfoils Mean lines.-The of camber within the usual range amounts different the uniform-load a mean line of Data with values.
the tabulated used are cambered by scaling simply obtained the the section for under by multi- which is described be obtained (a=1.0), may thus type 22 mean line the NACA for type of If any other follows.
that by the ratio 2: 6, 62 mean line N ACA 6-series airfoils for the N ACA data plying the the airfoil stated in data for multiplying the mean line is used, this fact should be mean line by and for the NACA d.esignation.
mean line by the ratio 4:6.
the NACA 64 6-SERIES AIRFOILS NACA AIRFOILS NACA 'FIVE.DIGIT-SERIES N ACA 6-set'ies airfoils are usu- system.-The Numbering state- together with a a six-digit number by ally designated numbering system for airfoils of system.-The Numberinl~ For example, type of mean line used.
the ment showing combination of based on a series ,is the NACAlive-digit "6" is a=O.5, the 65,3-218, NACA in the designation and geometric char- theoretical aerodynamic characteristics chordwise denotes the The" 5" designation.
the series 3). The first integer indicates and acteristics (references 2 chord behind tenths of the of position of minimum pressure in the relative magnitude ,pf camber in terms of the amount at zero the basic symmetrical section the leading edge for tenths Wit coefficient; the design lift coefficient in the design the range of lift 3" following the comma gives lift. The" and third The second is thus three-halves of the first integer.
the design lift coefficient above and below tenths coefficient in the distance from the leading edge integers together indicate both surfaces.
in which favorable pressure gradients exist on maximum camber; this distance in the to the locatlon of eoefficient the design lift the dash gives following The "2" represented by the number percent of the chord is one-half airfoil thickness the two digits indicate The last in tenths.
the airfoil last two integers indicate The these integers.
shows a=0.5" designation" The in percent of the chord.
The NACA 23012 airfoil the chord.
in percent of thickness mean-line designa- When the type of mean line used.
the has a, ,aesign lift coefficient of 0.3, has its maximum thus the uniform-load that it is understood is not given, tion ratio and has a thickness of the chord, at U percent camber has been used.
line (a= 1.0) mean percen~.
of 12 REPORT NO. 824-NA'rIONAL ADVISORY COMMITTEE FOR AERONAUTICS When the mean line used is obt.ained by combining more NACA 65(318)-(1.5) (16.5), a=O.5 t.han one mean line, the design lift. coefficient used in t.he Some early experimental airfoils are designated designation is the algebraic sum of the design lift coefficients by the in- sertion of of the mean the letter lines used, "x" immediately preceding and the mea.n lines are described in the hyphen as in the designation 66,2x-115.
the statement following the number as in the following case: Thickness distributions.- Datafor available N AOA 6-series a=0.5 CII=O . .3 } thickness forms are presented in the supplementary NACA 65,3-218 ' { figures.
These data are comparable with a=l.O, Cl =-0.1 the similar data i for airfoils of the NACA four-digit series, except that Air'foils having a thickness distribution obtained ordi- by linearly nates for intermediate thicknesses may not be correctly ob- increasing or decreasing the ordinates of one of the originally tained by scaling the tabulated ordinates proportional to derived thickness distributions are designated as in the follow- the thickness ratio.
This method of changing the ordinates ing example: by a factor will, however, produce shapes satisfactorily approx- NACA 65(318)-217, a=0.5 imating members of the family if the change in thickness The significance of all of ratio the numbers except those in is small.
Values of
the v/V and 6.v./V
for intermediate parentheses is the same as before. thickness ratios The may first number be approximated as described for and the the last two numbers enclosed in NACA four-digit series.
the parentheses denote, respec- tively, the low-drag range Mean and the thickness in percent lines.-The mean lines commonly used of with the the chord of the originally derived thickness distribution.
NACA 6-series airfoils produce a uniform chordwise loading The more recent NACA 6-sories airfoils are derived as from the leading edge to the point
~=a and
a linearly de- members of thickness families having a simple relationship creasing load from between the conformal transformations for airfoils of different this point to the trailing edge.
Data for NAOA thickness ratios mean lines with values of but having minimum pressure a equal at to 0, 0.1, 0.2, the samt;\ 0.3, 0.4, 0.5, 0.6, 0.7, 0.8, 0.9, chord wise position.
These and airfoils 1.0 are presented in the are distinguished from thp earlier individually derived airfoils supplementary figures.
The ordinates were computed by writing the num- by ber indicating the following formula, which represents a simplification of the low-drag ra.nge as a. subscript; for exa.mple, the original expression for mean-line ordinates given in NACA 65 -218, a=0.5 reference 11: For NACA 6-se1'ies airfoils having a thickness ratio less than 0.12 of the chord, the subscript number indicating the low-drag range should be less than unity. Rather than usc a fmctional number, a subscript of unity was originally em- ployed for these airfoils.
Since this usa.ge is not consistent with the previous definition of a number x x indicating x~ the low- h
-c
loge c+ U- (6)
drag range, c I
the designations of a.irfoil sections having a thick- ness ratio less than 0.12 of the chord are now given without where such a number.
As an example, an N AOA 6-series airfoil having a thickness ratio of 0.10 of the chord would be designated: NAOA 65-210
1 [1
1 h=- - (1-a)210g
] (l-a)--
(1-a)2 +U I-a 2 .
e Ordinates for the basic thiclniess distributions designated by a subscript are slightly different from those for The the ideal angle of corre- attack IXI corresponding to the design sponding individually derived thickness distributions. lift coefficient is given As by before, if the ordinates of the basic thickness distribution Cit have been changed 1)Y a factor, the low-drag range and cx(==-h thick- 27l'(a+D ness ratio of the original thickness distribution are enclosed in parentheses as follows: The data are presented for a design lift coefficient Cit equal to unity. All tabulated values vary directly with NAOA 65(318)-217, a=O.5 the design lift coefficient.
Oorresponding data for similar If, howevPJ', the ordinates of a basic thickness distribution mean lines with other design lift coefficients may accordingly having a thickness ratio less be obtained simply than 0.12 of the chord have been by multiplying the tabulated values by changed by a factor, 'the number the desired design lift coefficient.
indicating the low-drag range is eliminated and only the In order original thickness to camber NAOA 6-series airfoils, ratio is mean lines are enclosed in parentheses as follows: usually used having values of a, equal to or greater than the distance from the leading edge to the location of minimum NACA 65(10)-211 pressure for the selected thickness distribution at zero lift.
If the design lift coefficient in tenths or the airfoil thickness For special purposes, load distributions other than those in percent of chord are not whole integers, the numbers corresponding to the simple mean lines may be obtained by giving these quantities are usually enclosed in parentheses as combining two or more types of mean line having positive or in the following designation: negative values of the design lift coefficient.
The geometric SUMMARY OF AIRFOIL DATA and aerodynamic characteristics of such combinations may be serial letter "B." Mean lines used for the NACA 7-series obtained by algebraic addition of the values for the compo- airfoils are obtained by combining two or more of the pre- nent mean lines.
viously described mean lines. A list of the thickness dis- NACA 7-SERIES AIRFOILS tributions and mean lines used to form these airfoils is pre-
sented in table 1. The basic thickness distribution is given
Numbering system.-The NACA 7-series airfoils are desig- a designation similar to those of the final cambered airfoils.
nated by a number of the following type (reference 12): For example, the basic thickness distribution for the NACA 747A315 NACA 747A315 and 747A415 airfoils is given the designation The first number "7" indicates the series number. The NACA 747 A015 even though minimum pressure occurs at O.4c second number "4" indicates the extent over the upper sur- on both upper and lower surfaces at zero lift. Combination face, in tenths of the chord from the leading edge, of the of this thickness distribution with the mean lines listed in region of favorable pressure gradient at the design lift coeffi- table I for the NACA 747A315 airfoil changes the pressure cient. The third number "7" indicates the extent over the distribution to the desired type as shown in figure 2.
lower surface, in tenths of the chord from the leading edge, Thickness distributions.-Data for available NACA 7- of the region of favorable pressure gradient at the design lift series thickness distributions are presented in the supple- coefficient. The significance of the last group of three num- mentary figures. These thickness distributions are indi- bers is the same as for the previous NACA 6-series airfoils. vidually derived and do not form thickness families. The The letter "A" which follows the first three numbers is a thickness ratio may, however, be changed a moderate amount-say 1 or ,2 percent-by multiplying the tabulated serial letter to distinguish different airfoils having parameters that would correspond to the same numerical designation. ordinates by a suitable factor without seriously altering their For example, a second airfoil having the same extent of characteristic features. Values of (V/V2) and of v/V for thinner or thicker thickness distributions may be approximated by favorable pressure gradient over the' upper and lower sur- the method of equation (5). If the change in thickness ratio faces, the same design lift coefficient, and the same maximum thickness as the original airfoil but having a different mean- is small, tabulated values of I1V /V may be applied'directly a with reasonable a.ccuracy.
line combination or thickness distribution would have the I.B f---
r-
-
- k"
V ,NACA 747A315
1.6
"
'(upper surface)
V
II I I I
"
l
"" "-
1.4 ,
I
~. _I 1 1 ------ ~ . NACA 747AOl5 basic rhicimess distribution :----...
~
"" /
-----
1.2
!
""
k
I
r:-:JCA I 74~A3~5
(v)'1.0
"'"
---- /' (lower surface) "-
'-.
!
"-
.8
I
"'"
.6 .4 .2
o .3 .6 .7 .8 .9 1.0
. I .4 .5 xlc FIGURE 2.-Theoretical pressure distribution for the NACA 747A315 airfoil section at the design lift coefficient and the NACA 747AOlij husir thickness dis:l'ibuUOll.
TABLE I.-ANALYSIS OF AIRFOIL DE.RIVATION Mellon-line combination 1 Airfoil Basic thickness 1 ____ -;- ___ --;- ____ , ___ _,_----;--------.-----;-----,------;----,---1 designation form a=0.6 a=0.7 a=0.8 a=0.9 a=1.0 a=O a=O.l a=0.2 a=0.3 a=0.4 a=0.5 747A315 ________ 747A015 ____________________________ "" ______________________________ _ 0.763 747A415 ________ 747A015 ____________________________________________________________ _
=O:!~ ::::::::::::: :::::: ::::::: ----ii:ioo----
.763 I The numbers in the various columns headed "Mean-line combination". indicate the magnitude orthe design lift coefficient used.
REPOR'l' NO. 824-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS THEORETICAL CONSIDERATIONS this circle in complex coordinates is PRESSURE DISTRIBUTION~ z=aefo+iq,
(7)
where A knowledge of the pressure distribution over an airfoil is desimble for structural design and for estimation of the
z complex variable in circle plane
critical Mach number and moment coefficient if tests are not available. The pressure distribution also exerts a strong ¢ angular coordinate of z or predominant influence on the boundary-layer flow and,
a basic length usually considered unity
hence, on the airfoil characteristics. It is therefore usually 1/10 constant determining radius of, circle advisable to relate the airfoil characteristics to the pressure distribution rather than directly to the airfoil geometry.
This true circle is transformed into an arbitrary, almost Methods of derivation of thickness distributions.-As circular curve by the relation mentioned in the section "Historical Development," the ~, basic symmetrical thickness distributions of the N ACA 6-
='_= e(f-fol+i{O-q,)
(8) and 7-series airfoils, together with their corresponding pres-
z
sure distributions, are derived by means of conformal trans- formations. The transformations used to relate the known the equation of the almost circular curve is flow about a circle to that about an airfoil section were iO f z' =ae + (9) developed by Theodorsen in reference .9. Figure 3 shows schematically the significance of the various phases of the where process.
z' complex variable in near-circle plane
The circle about Which the flow is originally calculated has f ae radial coordinate of z'
its center at the origin and a radius of aiD. The equation of
f} angular coordi.nate of z'
In order Jor the transformation (8) to be conformal, it is necessary that the quantity (f}-r/» (given the symbol -E) be the conjugate function of (1/I-if;0); that is, if E is represented by a Fourier series of the form Z-p/one \-------jL----"---l '" '"
e=L: An sin n</>-L: Bn cos n</>
1 1 then (if;-1/Io) is given by the relation '" '"
(1/1-1/10)="'22 An cos nr/>+::8 Bn sin nr/>
1 1
=-.'= e fll ->;.)4-I(S-tJ)
z This relationship indicates that, if the function E(r/» is given, (1/1-1/10) can be calculated as a function of r/>. Means of performing this calculation are presented in reference 13.
The transformation relating the almost circular curve to tbe airfoil shape is (10) Z-p/one f-------,fL------'L----,
where f is the complex va.'iable in the airfoil plane. The
coordinates of the airfoil x and yare the real and imaginary
parts of f, respectively. These coordinates are given hy the
relations .c= 2a cosh 1/1 cos 8 (11) y=2a sinh 1/1 sin f} (12) The velocity distribution in terms of the airfoil parameters 1/1 and € is given exactly for perfect fluid flow by the expression fO v [sin (ao+¢)+~in (aO+€TE)] e (13) f=Xri y
V= ~ (sinh21/1+sin28) [( 1- :;y +(~~)]
FIGURE 3.-Transformations used to derive airfoils nnd calculate pressure distributionR.
DATA OF AIRFOIL SUMMARY .16 where ~
/
surface of airfoil over local velocity
! /
"\
':I. ... ·dE Ir
\-
/ .08 h
r V
\dfjJ
free-stream velocity V\
v
dlf.-I
/
/ dtPJ
\ "j
attack '-- ao section angle of
"'" / ~
o
II
d¢; dE ~
i ~
1J" 0/ de/> )
0/ (i1f'L
average value of
0/0
dfjJ'dfjJ.
'\
V
/
\ /
-.08 , at trailing edge ETE value of e
/
1\ V
/
by assuming
shapes were derived ~
symmetrical J
The basic
\
-.16 e/>. These values were as.a function of
de/de/>
suitable values of V
i and are subject to pr~vious experience the basis of chosen on .24
that V
the eonditions ~
-- V
L1J"~=o
1\
.16
i \ "'/
at -cp. These conditions
to df/dcp
at e/> is equal
and de/de/>
/E shapes.
'symmetrical closed
for obtaining V II
are hecessary
\'
.08 de/>.
;; \ j
by integrating
fCe/» were obtained simply V
Values of \
V
the by obtaining the conjugate of
o/(cp) were found
Values of ----
~
V
o
the.
to make % suffieient
adding a value /
eCe/» and curve of
V-
1\
cp=1f'. This condition assures a
at
0/ equal to zero
,'alue of >I,e
/
I I sharp trailing-edge sbape. f--_ ..
C\ .. -
-.08 at any the velocity distribution as small changes in Inasmuch
/
~-- to 1 + J; proportional
appr~ximately
surface are point of the
'"
'---.
-.IB were 6 ~.'I"
df/de/>
4 5 the initially assumed values of 2
(see refrrence 14), o
radians until the tP, a process of successive approximations altered by airloil.
After the ;he XACA 643-018
!~ with", for
E'~'
of velocity distribution was obtained. 4.-Variation of airfoil parameters,p, desired type FIGURE thr basic section basic thickness form.
were obtained, the ordinates of
0/ and e
final values of (11) equations computed by were distribution thickness that could be I/; and e parameters basic airfoil t.o derive (12).
and airfoils of various obtain eonstant faetor to by a there multiplied it appeared that When these computations were made, thickness ratios, without ha.ving the aforementioned limita- value of the leading-edge radius dependent.
an optimum was recent of the more Each tions in the resulting sect.ions.
position of minimum and the thickness the airfoil upon sub- families of NACA 6-series airfoils, in which numerical If the leading-edge radius was too small, a pre- pressure.
the scripts are used in the designation, having minimum pressure pressure distribution occurred in peak in the mature and by scaling up at a given chordwise position was obtained thr leadi.ng edge as the angle of att.ack immediate v-icinity of 1/;, and E.
down the basic values of the airfoil parameters If the leading-edge radius was too large, a was increasrd.
thr
(V))
prrcent of the chord behind by (indicated peak occurred a few pressure distributions premature Theoretical t.he radius COrI'rct l(lading-edge j With the edge.
lrading ACA 65-series a.irfoils covering n range of for a, family of N pressure distribut.ion became nearly flat over the forward This figure shows .5 (a).
thickness ratios are given in figure peak the airfoil before the normal leading-edge portion of the typical increase in the magnitude of the favorable pressure param- Curves of the the higher lift coefficients.
formed at the surface, and gradient, increase in maximum velocity over e/> for the NACA against
de/de/> plotted
0/, f, dl/;/de/>, eters rear portion increase in the relative pressure recovery over the 4.
64 -018 airfoil section are given in figure Figure 5 (b) of the airfoil'with increase in thickness ratio.
an the thickness ratio of tl~at, ,,,,hen Experience has shown shows the pressure distrihution for a series of bnsic thickness by multi- originally derived basic form was increased merely and having minimum forms having a thickness ratio of 0.15 unnecessarily factor, an by a constant plying all the ordinates value of the The at various chordwise positions.
pressure critical speed of the resulting section the large decrease in the and seen to decrease coefficient is minimum pressure manner ratio in a similar Reducing the thickness occurred.
magnitude of the pressure recovery over the renr portion of caused an unnecessarily large decrease in the low-drag range.
of the movement with the rearward to increase the airfoil ACA 6-series sections was this reason, each of the earlier N For point of minimum pressure.
that it was possible later found It was individually derived.
REPOR'!' NO. 824-NATIONAL ADVISORY COMMIT'!'EE FOR AERONAU'rtCS 2.8 28 , , , , , ,
~ ~4C~ 64 -oJs
_--- MACA 65,-012
2.4 - I-- '
- c::: -------- MACA 65,-015
~
------- MACA 65 -015
====-
MACA 65,-012 _--- NACA 6~-015 -'--- NACA 65-018 NACA 64,-015
- r--
---- NACA 67,1-015 ---- MACA 65.-021
c::==::::::=- c::==::::::=-
-""\ NACA 65.-015 -~ MACA 65 -015 /6 /.6 .-:-
~, -
~ ~-- --- ~ -:::c==
--'
(VI
(V!
--
=-~ -,~
=
~ ~
f;;::-
~
~ u. \
-
" . /2 1.2
c :=:::::::=-
~
{
f ~~ ~, NACA 6tJ,-015 MACA 65.J-0I8
"
~
~ ','
~ .8 .8
"
l'-~ -~
c :::>
~
NACA 67,/-015 MACA 65 -021 .4 .4 (a) (b) .2
o .2 .4 .6 .8 1.0 o .4 .6 .8 1.0
.:ric .:rIc (a) Variation with thickness.
(b) Variation with position of minimum pressure.
FIGURE 5.-Theoretical presmre distributions for some basic symmetrical NACA 6-series airfoils at zero lift.
The pressure distribution for one of the basic symmetrical and lower surfaces of the airfoil along the chord. The term thickness distributions at various lift coefficients is shown in "load distribution" is used to signify the distribution along figure 6. At zero lift the pressure distributions over the the chord of the normal force resulting from the difference in pressure on the upper and lower surfaces.
upper and lower surfaces are the same. As the lift coefficient
The pressure distribution about any airfoil in potential
is increased, the slope of the pressure distribution over the
flow may be calculated accurately by a generalization of the
forward portion of the upper surface decreases until it becomes
methods of the previous section. Although this method is
flat at a lift coefficient of 0.22 (the end of the low-drag range).
As the lift coefficient is increased beyond this value, the :usual not unduly laborious, the computations required are too peak in the pressure distribution forms at the lead~ng edge. long to permit quick and easy calculations for large numbers Rapid estimation of pressure distributions.-In the dis- of airfoils .. The need for a simple method of quickly obtaining cussion that follows, the term "pressure distribution" is used pressme distributions with engineering accuracy has led to
the development of a method (reference 15) combining
to signify the distribution of the static pressures on the upper
features of thin- and thick-airfoil theory. This simple
method makes use of previously calculated characteristics
of a limited number of mean lines and thickness distributions
that may be combined to form large numbers of airfoils.
Thin-airfoil theory (references 16 to 18) shows that the
load distribution of a thin airfoil may be considered to consist
of: (1) a basic distribution at the ideal angle of attack and
5.0
(2) an additional distribution proportional to the angle of
attack as measured from the ideal angle of attack.
The first load distribution is a function only of the shape of
4.0
the thin airfoil, or (if the thin airfoil is considered to be a
mean line) of the mean-line geometry. Integration of this
load distribution along the chord results in a normal-force
3.0
coefficient which, at small angles of attack, is substantially
(v!
equal to a lift coefficient Cit, which is designated the ideal
2.0
or design lift coefficient. If, moreover, the camber of the
mean line is changed by multiplying the mean-line ordinates
by a constant factor, the resulting load distribution, the
ideal or design angle of attack at and the design lift coefficient
Cl may be obtaIned simply by mUltiplying the original values
i
by the same fnctor. The characteristics of a large number of
mean lines are presented in both graphical and tabular form
in the supplementary figures. The load-distribution data
are presented both in the form of the resultant pressure.
coefficient P and in the form of the corresponding velocity-
R increment ratios !.lv/V. For positive design lift coefficie~ts, FIGURE 6.-Theoretical pressure distribution for the N ACA 65.-015 airfoil at several lift
these velocity-increment ratios are positive on the upper
. coefficients.
DATA OF AIRFOIL SUMMARY (14) should,
in equation
and of ~v/V
of. v/V
The values
negative on the lower surface; the opposite is
!'Iudace and
Methods
geometry.
to the ·airfoil
correspond
of course, true for negative design lift coefficients.
of obtaining the proper values of these ratios from the values
second load distribution, which results from changing
The
in the
tabulated in the supplementary figures are presented
the" additional load
the angle of attack, is designated herein
deRig- previous section "Description of Airfoils. "
and the corresponding lift coefficient is
distribution"
AvalV has the value of zero, the resulting
When the ratio
This additional load
additional lift coefficient."
nated the"
correspond
S will
coefficient
pressure
distribution of the
about the quarter-chord
distribution contributes no moment
of the airfoil
pressure distribution
to the
of approximately
independent
is
point and, according to thin-airfoil theory,
line, and
of the mean
Cl
section at the design lift coefficient i
The addi-
the airfoil geometry except for angle of attack.
be assigned this value as a first ap-
may
the lift coefficient
tional load distribution obtained from thin-airfoil theory is
If the pressure-distribution diagram is inte-
proximation.
of limited practical application, however, because this simple
will be found to be greater
Cl
grated, however, the value of
at the leading
theory leads to infinite values of the velocity
of
by an amount dependent on the thickness ratio
than Cli
by the exact thick-airfoil
is obviated
This _difficulty
edge.
.
the basic thickness form.
that the additional load
9) which also shows
theory (reference
at some
The pressure distribution will usually be desired
distribution is neither completely independent of the aidoil
this
Cli' For
corresponding to
not
specified lift coefficient
shape nor exactly a linear function of the lift coefficient.
.ob-
must be assigned some value
~va/V
purpose the ratio
ha,s been
For this reason, the additional load distribution
a
by
multiplying the tabulated value of this ratio
tained by
by the methods of reference 9 for each of the thick-
calculated
may be
For a first approximation this factor
factor j(a).
in the supplementary figures.
ness distributions presented
assigned the value
the form of velocity-increment
in
data are presented
These
(15)
an additional lift coefficient of
AVa/V corresponding to
rat.ios
lift coeffi-
additional
For positive
approximately unity.
coeffici(1nt for which the pressure distribu-
CI is the lift
where
these. velocity-increment ratios are positive on the
c:ients,
the value of
greater accuracy is desired,
surfaces; the If
the lower tion is desired.
negat.ive on
and
upper surfaces
the
produce
and error to
by trial
may be adjusted
j(a) is true for negative additional lift coefficients.
opposite
integration
by
actual desired lift coefficient as determined
additi~n to the pressure distributions associated with
In
of the pressure-distribution diagram.
two load distributions, another pressure' distribution
thes~
Although this method of superposition of velocities has
the basic symmetrical thick-
with
exists which is associated
experience has shown
justification,
theoretical
This pres- inadequate
ness form or thickness distribution of the airfoil.
methods
by the that the results obtained are adequate for engineering use.
calculated
has been
distribution
sure
of zero the results of even the first approximations agree
In fact,
described in the previous section for the condition
at least
data and are tl,dequate for
experimental
well with
('f; )2,
in the supplementary figures as
lift and is presented A com-
preliminary consideration and selection of airfoils.
numbers to the pressure
at low Mach
parison of a first-approximation theoretical pressure distri-
which is equivalent
7.
This
VIV.
velocity ratio
as the local an experimental distribution is shown in figure
S, and bution with
coefficient
local velocity ratio is always positive and is the same for
the
upper and lower surfaces of
the
corresponding points on
form.·
thickness
~
___
c
The velocity distribution about the airfoil is thus considered
com-
independent a ~ 06
and NACA '66(215)-2/6,
to be composed of three separate
ponents as follows: 2.0
dis-
the velocity
corresponding to
(1) The distribution
angle of
at zero
thickness form
basic
tribution over the
sJ-f'a;e: Uppe~ _0.
. "
attack
/;6
r-
design load
to the
corresponding
distribution
(2) The r-o-:
V
\
distribution of the mean line
i\
~
distribution corresponding to the additional load
(3) The I-<>-
II
1.2
angle of attack
with
distribution associated
surface ~ower "-;;
At'a/V correspond-
AviV and
velocity-increment ratios
The
velocity
to the
and (3) are added ~
(2)
ing to components ~r
'"
.8
total
to obtain the
(1)
component
corresponding to
ratio
~
S
velocity at one point, from which the pressure coefficient
--Theory o Experimenf is obtained; thus, .4
(14)
.8 /.0 .2 .4 .. 6
When this formula is used, values of the ratios corresponding o
:rIc
the resulting value
and
x are added together
one value of
to
the N ACA
S is assigned to the airfoil surface
of theoretical and experimental preS3ure distributions for 7.-Comparison
of the pressure coefficient FIGURE
c, = 0.23.
a = 0.6 airfoil.
66(215}-216, X.
t the same value of
It 12 REPORT NO. 824-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS Some discrepancy naturally occurs between the results of The supplementary figures give a value of 1.182 for v/V experiment and of any theoretical method based on potential atx=0.25 for the NACA 65 -015 basic thickness form. The flow'because of the presence of the boundary layer. These desired value of v/V is obtained by applying formula (5) effects are small, however, over the range of lift coefficients as follows: v 14 for which the boundary layer is thin .and the drag coefficient V=(1.182-1) 15+1 ifllow.
=1.170 Numerical examples.-The following numerical examples are included to illustrate the method of obtaining the first- From the supplementary figures the following values of approximation pressure distributions: AVa/V are obtained at x=0.25 for the following basic thickness
Example 1: Find the pressure coefficient S at the station
forms: x=0.50 on the upper and lower surfaces of the NACA 65 -418 airfoil at a lift coefficient of 0.2.
From the description of the NACA 6-series airfoils, it is determined that this airfoil is obtained by combining the NACA 65 -018 basic thickness form with the a= 1.0 .type mean line cambered to a design lift coefficient of 0.4. The following data are obtained from the supplementary figures for this thickness form and mean line at x=0.50: By interpolation the value of AVa/V of 0.287 may be
v
assigned to the 14-percent-thick form. The desired value of V=1.235 AVa/V is then computed as follows by use of equation (15): A'v
~=0.157 a
V =(0.287) (0.6-0.2) =0.115
~=0.250
Data presented in the supplementary figures for the a=0.5 The desired. value of AVa/V is computed as follows by use of type mean lines give the value of 0.333 for Av/V at x=0.25.
equation (15): As stated in the description of the NACA 6-series airfoils, the desired value of AV/V is obtained by multiplying the
A~a=(0.157)(0.2-0.4)
tabulated value by the design lift coefficient. Thus, =-0.031
~ = (0,333) (0.2)
The desired value of AV/V is obtained by multiplying the tabulated value by the design lift coefficient as stated in the =0.067 description of the NACA 6-series airfoils. Thus, Substituting the foregoing values in equation (14) gives the
AV
values of S as follows:
V = (0.250) (0.4)
For the upper surface =0.100 S= (1.170+0.067 +0.115)2 Substituting these values in equation (14) gives the following =1.828 values of S: For the lower surface For the upper surface S= (1.235+0.100-0.031)2 S= (1.170-0.067 -0.115)2 =1.700 =0.976 Example 3: Find the pressure coefficient S at the station For the lower surface x=0.30 on the upper and lower surfaces of the NACA 2412 S= (1.235-0.100+0.031)2 airfoil at a lift coefficient of 0.5.
The description of airfoils of the NACA four-digit series =1.360 shows that the necessary data may be found from the NACA
Example 2: Find the pressure coefficient S at the station
0012 thickness form and 64 mean line in the supplementary x=0.25 on the upper and lower surfaces of the NACA figures. From these figures the following data are obtained: 65(215)-214, a=0.5 airfoil at a lift coefficient of 0.6.
At x=0.30 The airfoil designation shows that this airfoil was obtained v V=1.162 by combining a thickness form obtained by multiplying' the ordinates of the NACA 65 -015 form by the factor 14/15 At x=0.30 with the a=0.5 type mean line cambered to a design lift
A~a=0.239
coefficient of 0.2.
SUMMARY OF AIRFOIL DATA For the NACA 64 mean line at x=0.30 at the design lift coefficient is to separate the pressures on the upper and lower surfaces by an amount corresponding
b.1)
V=0.260 approximately to the design load distribution of the mean line. When the local value of the design load distribution is For the NACA 64 mean line
positive, the pressure coefficient S on the upper' surface .is
increased (decreased absolute pressure) whereas that on the lower surface is decreased. This effect is shown in figure 8 (a)
The values of b.v/V llnd eli corresponding to the airfoil
for various amounts of camber.
geometry are obtained by multiplying the foregoing values The maximum value of the pressure coefficient on the upper by the factor 2/6 as explained in the description of these surface at the design lift coefficient increases with the design airfoils; thus, lift coefficient and for a given design lift coefficient increases.
~=(0.260)(i) with decreasing values of a. The result is to cause the critical
Mach number at the design lift coefficient to decrease with =0.087 increasing camber or with the use of types of mean line con- centrating the load near' the leading edge. Figure 8 (b)
elt=(0.76)(~)
shows that the location of minimum pressure on both surfaces is not affected if a type of mean line is used having a value of =0.253 a at least as large as the value of x/e at the position of minimum pressure on the basic thickness distribution. If a
The desired value of b.va/V is obtained from equation (15)
as follows: mean line with a smaller value of a is used, the possible extent, of laminar flow along the upper surface will be reduced.
t:,.~a= (0.239) (0.5-0.253)
CRITICAL MACH NUMBER =0.059 The critical speed is defined as the free-stream speed at Substituting the proper values in equation (14) gives the which the velocity at any point along the surface of the air- values of S as follows: foil reaches the local velocity of sound. If the maximum value- For the upper surface
of the low-speed pressure coefficient S is known either experi-
mentally or from theoretical methods, the criti~al Mach,
S= (1.162+0.087+0.059)2
number may be predicted approximately by the Von Karman.
= 1.712 method (reference 19). A curve relating the critical Mach.
For the lower surface
number and the low-speed pressure coefficient S has been
calculated from the equations of reference 19 and included in ..
S= (1.162-0.087 -0.059)2 the supplementary figures. These predicted critical Mach.
= 1.032 numbers are useful for preliminary considerations in the· Effect of camber on pressure distribution.-At zero lift the
absence of test data and appear to correspond fairly well to
pressure distributions over the upper and lower surfaces of the Mach numbers a t which the local velocity of sound is.
a basic symmetrical thickness distribution are, of course, reached in the high-critical,speed range of lift coefficient~ identical. The effect of camber on the pressure distribution This criterion does not, howe~~r, appear to predict accurately, \ \
~ Up~er 15ur~oc~
r- ____ Lower surface I 1
~
-' --NACA 65 -015 NACA 65,-015 NACA 65,-015 -NACA 65,-415, (1=0.3- NACA 65,-415,0.0.5
:;,.~
1--+--I---+_+--1--.,..4-'N.A CA 65,-015 20 , , NACA 65,-415,0=0.7- 20 , ,NACA 65,-215 ,f ,/ / /;' f-rNACA 65,-415 I tf',f 'II, f-+--+-=-,....,,=f--;<=-P' ~,'"">4'NA CA 65,- 415
~
NACA 65 -415, a=0.3·
V":-I--,7,fo ,NACA 65,-6 5
~ 2
~tt;rtF
~
NACA 65,-215 1.6 V/,' ii/
"'~ I----"-
(v/ 0-~~fj 1-!-0:~ It
" " : f,'-:=::::
(il)'
, " '-..
, ~ I V,
{f/
I/V _- - tj ~0G~ ~~
~ : "
NACA 65,-415, a=0.5 1.2
~
f-i;'.': ",A ;:::
'I ~
~ ~-
~~ ~ ~
NACA 65,,-415 '~~ ~'-....
I.! ~
.8 .... ~ NACA 65,-415, a=o.7" ....
~.
.4~4--+~-+--I---+--+--I---+~ NACA 65 -615
.4 ~
NACA 65.-415 (b) (a)
o .2 .4 .6 .8 1.0 0 .2 .4 .6 .8 1.0
,rIc ,rIc (a) Amount of camber.
(b) Type of camber.
FIGURE 8.-Effect of amount and type. of cambN' on pressure distribntion at design lift.
REPORT NO. 824-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS the Mach numbers at which large changes in airfoil char- NACA 64 mean line in the supplementary figures. The acteristics occur, especially when sharp pressure peaks exist moment coefficient for this mean line is -0.157. The at the leading edge. A discussion of the characteristics of required value is then airfoil sections at supercritical Mach numbers is beyond the
c / ·( -0.157) "6
mc scop'e of this report.
For convenience, curves of predicted critical Mach num- =-0.105 ber plotted against the low-speed section lift coefficient have ANGLE OF ZERO LIFT been included in the supplementary figures for a number of airfoils. High-speed lift coefficients may be obtained by Methods of calculation.- Values of the ideal or design multiplying the low-speed lift coefficient by the factor angle of attack at corresponding to the design lift coefficient 1 .
Cit are included among the data for the various mean lines
..jl"'::'W· 'I'he critical Mach numbers have been predicted
presented in the supplementary figures. The approximate from theoretical pressure distributions. For airfoils of the values of the angle of zero lift may be obtained from the NACA four- and five-digit series and for the NACA 7-series data by using the theoretical value of the lift-curve slope airfoils, the theoretical pressure distributions were obtained for thin airfoils, 2'lr per radian. The value of alo in degrees by Theodorsen's method. For the other airfoils the theo- is then retical pressure distributions were obtained by the approxi- (16) mate method described in the preceding section.
The data in the supplementary figures show that, for any The tabulated values of aj may be scaled linearly with one type of airfoil, the maximum critical Mach number the design lift coefficient or .with the mean-line ordinates.
decreases rapidly as the thickness is increased. The effect
Although these theoretical angles of zero lift may be useful
of camber is to lower the maximum critical Mach number in preliminary design, they should not be used without and to shift the range of high critical Mach numbers iii the experimental verification for such purposes as establishing same manner as for the low drag range. For common types the washout of a wing. ./ of camber the minimum reduction in critical speed for a
Numerical exampl~s·./: The method of comPlltlng alo is
given design lift coefficient is obtained with a uniform load illustrated in the following e~amples: / type of mean line. A comparison of the data presented in Example 1: Find th~ theoretical angle of zero lift of the the supplementary figures shows that N ACA 6-series sec- NACA 65 -515, a=0.5 airfoil.
tions have consi<ierably higher maximum critical Mach This airfoil number indicates a design lift coefficient of numbers than NACA 24-, 44-, and 230-series airfoils of 0.5. Dll,ta for the NACA a=0.5 mean line indicate that corresponding thickness ratios.
ai=3.04° when Clt=1.0. The desired value of aj is then MOMENT COEFFICIENTS a;= (3.04) (0.5) Methods of calculation.-Theoretical moment coefficients =1.52° may be approximated directly from the values presented in the supplementary figures for the various mean lines. These Substituting in equation (16) gives values were obtained from thin-airfoil theory and may be scaled up or down linearly with the design lift coefficient or (57.3) (0.5) alo=1.52 with the mean-line ordinates. These theoretical values are 2'lr sufficiently accurate for preliminary considerations, but ex- =-3.0° perimental values should be used for stability and control calculations. Example 2: Find the theoretical angle of zero lift for the NACA 2415 airfoil.
Numerical examples.-The following numerical examples The description of the N ACA four-digit-series airfoils illustrate the methods of calculating the moment coefficients: Example 1: Find the theoretical moment coefficient about shows that the required values of at and Cl may be obtained i the quarter-chord point for the NACA 65 -215, a=0.5 2 by multiplying the corresponding values for the N ACA 64 airfoil.
mean line (see supplementary figures) by a factor 2/6; then The designation of the airfoil shows that the design lift coefficient of this airfoil is 0.2. From the data on the
a,=(O.74) (~)
NACA a=0.5 type mean line included inthe supplementary =0.25° figures, the value of C"'c/4 is -0.139 for a design lift coefficient of 1.0. The desired value of the moment coefficient. is
accordingly CI,;~(0.76) (~)
C / =(-0.139) (0.2) mc =0.253 =-0.028 and from equation (16) Example 2: Find the theoretical moment coefficient about (57.3) (0.253) the quarter-chord point for the NACA 4415 airfoil.
2'lr From the description of the N ACA four-digit series airfoils, the required data is found to be presented for the =-2.0° SUMMARY OF AIRFOIL DATA DESCRIPTION OF FLOW AROUND AIRFOILS lower surfaces. If the region of laminar flow is extensive, Perfect-fluid theory postulates that the flow follow the separation occur", immediately downstream from the location airfoil contour smoothly at all angles of attack with no loss of minimum pressure (reference 20) and the flow returns to of energy. Consequently, perfect-fluid theory itself gives the surface almost immediately at flight Reynolds numbers no information concerning the profile drag or the maximum as a turbulent boundary layer. This turbulent boundary lift of airfoil sections. The explanation of these phenomena layer extends to the trailing edge. If the surfaces are not is found from a consideration of the effects of viscosity, sufficiently smooth and fair, if the air stream is turbulent, which are of primary importance in a thin region near the or perhaps if the Reynolds number is sufficiently large, tran.:.
surface of the airfoil called the boundary layer.
sition from laminar to turbulent flow may occur anywhere Boundary layers in general are of two types, namely, upstream of the calculated laminar separation point.
laminar and turbulent. The flow in the laminar layer is For low and moderate lift coefficients where inappreciable smooth and free from any eddying motion. The flow in the separation occurs, the airfoil profile drag is largely caused by turbulent layer is characterized by the presence of a large skin friction and the value of the drag coefficient depends number of relatively small eddies. Because the eddies in the mainly on the relative amounts of laminar and turbulent turbulent layer produce a transfer of momentum from the flow. If the location of transition is known or assumed, the relatively fast-moving outer parts of the boundary layer to drag coefficient may be calculated with reasonable accuracy the portions closer to the surface, the distribution of average from boundary-layer theory by use of the methods of velocity is characterized by relatively higher velocities near references 23 and 24.
the surface and a greater total boundary-layer thickness in As the lift coefficient of the airfoil is increased by changing . a turbulent boundary layer than in a laminar' boundary layer the angle of attack, the resulting application of the additional developed under otherwise identical conditions. Skin fric- type of lift distribution moves the minimum-pressure point tion is therefore higher for turbulent boundary-layer flow upstream on the upper surface, and the possible extent of than for laminar flow. laminar flow is thus reduced. The resulting greater propor':' When the pressures along the airfoil surface are increasing tion of turbulent flow, together with the larger average veloc- in the direction of flow, a general deceleration takes place. At ity of flow over the surfaces, causes the drag to increase with the outer limits of the boundary layer this deceleration takes lift coefficient.
place in accordance with Bernoulli's law. Closer to the sur- In the case of many of the older types of airfoils, this face, no such simple law can be given because of the action forward movement of transition is gradual and the resulting of the viscous forces within the boundary layer. In general, variation of drag with lift coefficient occurs smoothly. The however, the relative loss of speed is somewhat greater for pressure distributions for NACA 6-series airfoils are such as particles of fluid within the boundary layer than for those at to cause transition to move forward suddenly at the end of the outer limits of the layer because the reduced kinetic the low-drag range of lift coefficients. A sharp increase in energy of the boundary-layer air limits its ability to flow drag coefficient to the value corresponding to a forward loca- against the adverse pressure gradient. If the rise in pressure tion of transition on the upper surface results. Such sudden is sufficiently great, portions of the fluid within the boundary shifts in transition give the typical drag curve for these air- layer may actually have their direction of motion reversed foils with a "sag" or "bucket" in the low-drag range. The and may start moving upstream. When this reverse occurs, same characteristic is shown to a smaller degree by some of the flow in the boundary layer is said to be "separated." the earlier airfoils such as t~e NACA 23015 when tested in Because of the increased interchange of momentum from fl, low-turbulence stream.
different parts of the layer, turbulent boundary layers are At high lift coefficients, a large part of the drag is contrib- uted by pressure or form drag resulting from separation of much more resistant to separation than are laminar layers.
the flow from the surface. The flow over the upper surface is Laminar boundary layers can only exist for a relatively shQrt characterized by a negative pressure peak near the leading distance in a region in which the pressure increases in the direction of flow. Formulas for calculating many of the edge, which causes laminar separation. The onset of tur- boundary-layereharacteristics are given in references 20 to 22. bulence causes the flow to return to the surface as a turbulent After laminar separation occurs, the flow may either boundary layer. High Reynolds numbers are favorable to .
leave the surface permanently or reattach itself in the form the development of turbulence and aid in this process. If of a turbulent boundary layer. Not much is known concern- the lift coefficient is sufficiently high or if the reestablish- ing the factors controlling this phenomenon. Laminar sep- ment of flow following laminar separation is unduly delayed aration on wings is usually not permanent at flight values of by low Reynolds numbers, the turbulent layer will separate the Reynolds number except when it occurs near the leading from the surface near the trailing edge and will cause large edge under conditions corresponding to maximum lift. The drag increases. The eventual loss in lift with increasing size of the locally separated region that is formed when the angle of attack may result either from relatively sudden laminar boundary layer separates and the flow returns to the permanent separation of the laminar boundary layer near surface decreases with increasing Reynolds number at a the leading edge or from progressive forward movement of attack.
given angle of turbulent separation. Under the latter condition, the flow The flow over aerodynamically smooth airfoils at low and the surface may be separated over a relatively large portion of moderate lift coefficients is characterized by laminar boundary
prior to maximum lift. A more extended discussion of the
layers from the leading edge back to approximately the loca- flow conditions associated with maximum lift is given in tion of the first minimum-pressure point on both upper and reference 5.
REPORT NO. 824-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS .06 EXPERIMENTAL CHARACTERISTICS o N4CA 65 - 4Z0 14211 SOURCES OF DATA ~ .04 o NACA 66(2xI5)-116 , o'NACA 65,-418U..1 rJ' .03 t> NACA 2302/ The primary source of the wind-tunnel data presented is II ....: (rough leodli7Cj .edqe) from tests in the Langley two-dimensional low-turbulence ~ '2 " NACA 63(420)-422 I. 1 I.
'c .0 . (rough /eooii7g edge) pressure timnel (TDT). The methods, used to obtain and :..:: "- "NACA 230?1 correct the data are summarized in the appendix. Design I--- a I- Is>
'" :--
".010 data obtained from tests of 2-foot-chord models in this 8'.008 , r-;;;;; tunnel are presented in the supplementary figures. I.. Turbulent I- ... -- -
- - , ----- -
- ~-
\5 .006 Some wind-tunnel data presented were' obtained in other
S
'NACA 00/2 NACA wind tunnels. In each case, the source of the data
13 .004
(l; ',Laminar is indicated and the testing techniques and corrections used II) .003 : were conventional unless otherwise indicated.
§
~
.S: .002
Most of the flight data consist of drag measurements made i"'-...
by the wake-survey method on either the airplane \\ing or it ~
, "glove" fitted over the wing as the test specimen. When- .001 ., .
.? .?f) 30ldO 6 .8 /.0 3 4 8 10 ever the measurements were obtained for a glove, this fact /Teynolds number, R is indicated in the presentation of the data. All data obtained FIGUllE 10.- Va"iation of minimum section drag coefficient with Reynolds number for several at high speeds have been reduced to cofficient form by airfoils, together with laminar and turbulent skin-friction coefficients for a flat plate.
compressible-flow methods. In the case of all such tested was a practical-construction model. It may be noted NACA flight data, precautions have 'been taken to ensure that the drag coefficient for the NACA 65 -418 airfoil at low that the results presented are not invalidated by cross Reynolds numbers is substantially higher than that of the flows of low-energy air into or out of the survey plane.
NACA 0012, whereas at high Reynolds numbers the opposite DRAG CHARACTERISTICS OF SMOOTd AIRFOILS is the case. The higher drag of the NACA 65 -418 airfoil Drag characteristics in.low~drag range.--The value of the section at low Reynolds numbers is caused by a relatively drag coefficient in the low-drag range for smooth airfoils is extensive region of laminar separation downstream of the mainly a function of the Reynolds number and the relative point of minimum pressure. This region decreases in size extent of the laminar layer and is moderately affected by the with increasing Reynolds number. These data illustrate the airfoil thickness ratio and camber. The effect on minimum inadequacy of low Reynolds number test data either to esti- drag of the position of minimum pressure which determines mate the full-scale characteristics or t~ determine the relative the possible extent of laminar flow is shown in figure 9 for merits of airfoil sections at flight Reynolds numbers (refer- some NACA 6-series airfoils. The data show a regular ences 25 and 26).
decrease in drag coefficient with rearward movement of The variation of minimum drag coefficient with camber is shown in figure 11 for a number of smooth 18-percent-thick minimum pressure.
NACA 6-series airfoils. These data show very little change
o I NAtA J3,-~/51
.016 o NACA 64r2/5 o NACA 652-2/5 {> NACA 66,-215 <:7 NACA.67,/-i?15 NACA airfoil 53-series / Cl 154 -series 0 65-serles L> {56-series
r-
'V 65,J-818 t-- i'- .'
.2 3 ~ .5 .6 .7 8 .I l Position of minimum nreSsurp. .r " FiGURE 9.-Variation of minimum drag coefficient with position of minimum pressure for some N ACA 6 series airfoils of the Same camber and thickness. R = 6 X 10'.
The variation of minimum drag coefficient with Reynolds number for several airfoils is shown in figure 10. The drag coefficient generally decreases with increasing Reynolds num .
ber up to Reynolds numbers of the order of 20 X 10 • Above this Reynolds number the drag coefficient of the NACA
I
65(420-420 airfoil remained substantially constant up to a o
.2 .4 .6 .8 Reynolds number of nearly 40X 10 • The earlier increase in Oesiqn sect/on lift coeff/cient eN drag coefficient shown by the NACA 66(2x15)-U6 airfoil FIGURE ll.-'-Variation of minimum section drag coefficient with camber for several NACA may be caused by surface irregularities because the specimen 6·series airfoil sections of I8-percent thickness ratio. R = 9 X I()<l: . SUMMARY OF AIRFOIL DATA in nlllllmum drag coefficient with increase in camber. A The data presented in the supplementary figures for the large amount of systematic data is included in figure 12 to NACA 6-series thickness forms show that the range of lift show the variation of minimum drag coefficient with thick- coefficients for low drag varies markedly with airfoil thick- ness ratio for a number of NACA airfoil sections ranging in ness. It has been possible to design airfoils of 12~percent thickness from 6 percent to 24 percent of the chord. The thickness with a total theoretical low-drag range of lift coeffi- minimum drag coefficient is seen to increase with increase in cients of 0.2. This theoretical range increases by ~pprox thickness ratio for ea,ch airfoil series. This increase, how- imately 0.2 for each 3-percent increase of airfoil thickness.
ever, is greater for the NACA four- and five-digit-series air- Figure 13 shows that the theoretical extent of the low-drag foils (fig. 12 (a)) than for the N ACA 6-scries airfoils (figs, range is approximately realized at a Reynolds number of 12 (b) to 12 (e)).
9 X 10 . Figure 13 also shows a characteristic tenclE'ncy for 016 the drag to increase to some extent toward the upper end of the low-drag range for moderately cambered airfoils, pai'- Rough ---- --- ticularly for the thicker airfoils. All data for the X ACA
.012 f-Smooth -- -
. ~
.
6 .. - 6-series airfoils show a decrease in the extent of the IO\'l-drag
-
-
J-'- .0-' range with increasing Reynolds number. Extrapolation of
-
.008 c- the rate of decrease observed at Reynolds numbers below
f-- f-*-
...- I--'
,...P-- .-- 9X 10 would indicat(> a vanishingly small low-drag range at
>-----'
:,: (4~cfl9it) _
or'
.004 flight values of the Reynolds number. Tests of a carefully <> Ll 44 I I constructed model of the XACA 65(421)-420 airfoil showed, 'V 230 (5-digtf) - (a) however, that the rate of reduction of the low-drag range I I with increasing Reynolds number decreased markedly at
°
.012 -~ Reynolds numbers above 9X 10 (fig. 14). These data indi-
.--
.'
.-0-
' -
cate that the extent of the low-drag range of this airfoil is --- -B- -' .008 reduced to about olH'-half the theoretical value at a Reynolds eli o 0 r-- 6 number of 35 X 10 • Iil ~ r> () .2 t004 r--- Ll .4 IJ' ,6 .032 'V ' .. ; '1;: (b) .~ 0 .(J
6 N)CA ~41~4/~
.028 o NACA tN. -415
~ .012
_ -b <) NACA 64,-418 \) . ' \.J LI NACA 64,-421 -I;t'
,-
-
, '(y' .024 (?
I.. .008 J qi "15 '--- '1-.'
00- c: c: lr- ~ a .I ·~,020 ~
."
t ,004
.2 <> ~ IV .4 Q)
r-- "
"
,6
'" 'V
o
(e) f P
lJ.016 I
§ 0
.§
b 1
~ ~ ,012 {;
-e Q.l
l.0
--
(;:.012 -- , .~ .0
--
V
-' , .;::
~
-s- -' (; .008 [t/ cl
i ~
~ lai1
~,008 ~ 00-
~ 1/ j
<> .2 ~ L1: ~ n ~~ r-- .004 .4 ~ ".
""CO 'V .8 .004 (d)
°
,-"'- f!.;.6 -1.2 .012 /.2 /6
-.8 -.4 ° .4 .8
-
.--' . Section Iii! coefficient, c, ~ - <;;>' )-'
.0- -
,- FrGt!RE 13.-Drag characteristics of some XACA 54·series airfoil sections of ,-urious thick- .008 nesses, cambered to a design lift coefficient of 0.4. R = 9 X 10'; TDT tests 682, 733, 735, C't I and 691.
o °
r--I .2 <> .£>.
,4
r--
,004
"
The values of the lift coefficient for which low d.rag is I i (e) obtained are determined largely by the amount of camber.
, ~ ':> 4 12 16 cO 28 Jc. The lift coefficient at the center of the low-drag range corre- J,/r[: ... ;i/ th:ckness, percen,+ or' chord
°
sponds approximately to the design lift coefficient of the (a) XACA four- and five·digit series.
mean line. The effect on the drag characteristics of various (b) XACA 63·series.
(el XACA 64-series.
amounts of camber is shown in figure 15. Section data indi- (d) X .'I.e A 55·series.
cate that the location of the low-drag range may be shifted (e) XAC A 66-series, FIGFRE 12.-Variation of minimum section drag coefficient with airfoil thickness ratio for
by even such crude camber changes as those caused by small
several NACA airfoil seNions of differel1', cal"bc~s in both smooth and rough conditions.
deflections of a plain flap. (See supplementary fig.)
6• R = 6 X 10 REPORT NO. 824-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 'c.4 .0.36
1 r 1 1 I 1 '1 \ \
I o Upper limit of low-drag range
~OXI/06 0
2.0 .032 o Lower limIt of low-drag range (5.0 (> 9.0 6 /53 1.0 .028 'V 25.0 t>. 35.0 'i' ~/.C "'~ .~ .!i u-..
.\J r---,
r-
~ .8 ....
QJ.
-- - -
C) ,
"
<t: .4 :.:::: I:: ~ .S!
\)
If
'"
,...-' , ~ if'
<N
l{j
~
~4 .008
.ij}
~ ~
. ~ .,.", J..:a -.8 .. 004 (a) (b) -/.2
, o
-1.6
o 4 8 /2 /6 20 24 28 -/.2 -.8 -,4 0 .4 .8 1.2
1.6 Reyno/ds number, R Section lift· co<'!r'ficient, C z (b) Section drag characteristics at various Reynolds numbers.
(a) Variation of upper and lower limits of low-drag range with Reynolds number.
FIGURE 14.- Variatton of low-drag range with R~ynolds number for the N ACA 65(",)-420 airfoil. TDT tests 300, 312, and 328.
, .
. OJ2 o NAc;:'A 653~ 018
Drag' characteristics outside low-drag range.-At the end
.028
--
o NACI1 85 -2 I 8
of the low~drag range the drag increases rapidly with increase
ONACA 65,-418' 6 NACA. 65 -6/8
in lift coefficient. For symmetrical and low-cambered air-
V NACA 65,3-818
foils, for which the lift coefficient at the upper end of the
low-drag range is moderate, this high rate of increase does
1/
not continue. (See fig. 15.) For highly cambered sections;
for which the lift at the upper end of the low-drag range is
1\ already high, the drag coefficient shows a continued rapid
mcrease.
1\ i\
Ii
, '\
Comparison of data for airfoils cambered with a uniform-
<flllf,
load mean line with data for airfoils cambered to carry the
"c
t\ ~ I
load farther forward shows that the uniform-load mean line
'b.
.p1f
~ ~
is favorable for obtaining low drag coefficient.s at high lift
~r\' ~ ~~y
~ I))
coefficients (fig. 16 'and reference 27).
i'-..\ ...& ~ 1\ l'h
Data for many of the airfoils given in the supplementary
.004
figures s~ow large reductions in drag with increasing Reynolds
number at high lift coefficients. This scale effect is too large
o
to be accounted for by the normal variation in skin friction
-1.6 -1.2 -.8 -.4 0 .4 .8 /.2 1.0 Section lift coeffic/ent, fl
and appears to be associated with the effect of Reynolds
FIGURE 15.-Drag characteristics of some NACA 65-seriesairfoil sections of 18 percent thick-
number on the onset of turbulent flow following laminar
ness with various amounts of camber. R = 6 X 106; TDT tests 163, 314, 802, 813, and 830.
separation near the leading edge (reference 28).
Effects of type of section on drag characteristics.-A com-
The location of the low-drag range shows some variation
parison of the drag characteristics of the NACA 23012 and of
from that predicted by simple thin-airfoil theory; This de-
three NACA 6-series airfoils is presented in figure 17_ The
parture appears to be a function of the type of mean line
drag for the NACA 6-series sections is substantially lower
used (reference 27) and the airfoil thickness. The effect of
than for the NACA 23012 section in the range of lift coeffi-
airfoil thickness is shown in figure 13, from which the center
cients corresponding to high-speed flight, and this margin
'of the low-drag range is seen to shift to higher lift coefficients
may usually be maintained through the range of lift coeffi-
with increasing airfoil thickness. This shift is partly ex-
cients useful for cruising by suitable choice of camber.
plained by the increase in lift coefficient above the design
The NACA 6-series sections show the higher maximum values
lift coefficient for the mean line obtained when the velocity
of the lift-drag ratio. At high values of the lift coefficient,
increments caused by the mean lirie are combined with thp
however, the earlier NACA sections have generally lower
velocity distribution for the basic thickness form according
drag coefficients than the NACA 6.-series airfoils.
to the approximate methods previously described.
Y1 c:1 is:: is:: :> ~ o "9 :> .... ~ "'I o .... t"' t:l :> >-3 :>
~ c:o 1.6
19 IT
/
1& I 1.2
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if -III I .8 and y/c 04.7 lsi ..
-
406, -0.060 ~ posHion c, .024- .020 .
.036 .032 .028 c.
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r--- ----...
coef'fiCient, .8 .8 TDT ~ .a~02 lift
r--- I--- r-- e--- I
I
.....
-
-
9 X 10'.
I-- 0=0.5 I .6
-fB 65r418 of
-.4 ·.6 ..... .65.:c.11,B, a Section L
f5
J:/c :ric
"
NACA i\fAt:::A nnmber CA I 653-418~ o 0" .4 .4 N1 -.8 I NACA a Reynolds "t .2 .2 -1.2 f.--"' I-- f.--"' t-- airfoils 0:5
V ~ V r--
'J 7
o =
.2 .2 -1.6 -.I -.2 -.2 -.2 -.3 -.4 -.5 _ Q ci ci E ~o .0/6 ~o
'-' ~ o \.) ~ ~
-<-..~ ~ \ ..... ~
o
~ c.012 III ".008 8'
+.:c {;
.~ .0 ..::: "- ~.004 ~
65.-418, NACA and ~ \ ~ 653-418 p..::
-r
16 NACA I,....-: the deq of V- «0, IA' 1Jt' rJ attack, I "I I
t of
I~ II anqle I VI aerodynamic characteristics
IL
-8 '1/ the of Section -16 16.-Comparison -24 FIGURE • -32 .8 .4 1.6 -.4 -.8 2.0 3.6 3.2 28 2.4 -12 -/.0 -2.0, ~ c: ",,- c:: 0 u 'l:>
.0 ;::: :g (I)
"i-.,1.2 .!!! ~ .... ""
0 .> .! .
• -./ -" -.4 -.5 -.2 >t "Q ~ "" t: 'l:> o \J ~ t: .~
..... ~ ..... ~
il:l trJ '"d o il:l >-3 Z P 00 ~ ... ~ !:3 ..... o Z >- t"' >- t:::) <l ..... w o il:l ..<j C1 o ~ ~ ..... >-3 >-3 trJ trJ ~ o il:l ~ il:l o Z >- q >-3 '"' C1 W
~ o
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I
/.6 If ,.
liP / II 1/ Iv II II VJ: I Ll I,P I:f I .
I ~ I I 1.2 I V T ~ 1/ 1/ I I..-¢ II 1/ V 1/ j1i- .8
Yi'
~ 1/ !1J -.040 -.062 -.105 -0.004 Ji(A !At
t
pasilion 1 c c.
tUIlIlCI.
'" V 1-\<7 Stand°rdrauCJhnessl--- o. .268 .265 .264 .4 0.247 I>- l- !~ _~_ -~- ~ coeffIcient, t--.
R 9.0x10 9.0 9.0 9.0 D;O.-~- ~t--. lift -- ~ \ -.4
1\ ~ ~ I T
\ Section 'I II ~ 1'0 ~ -415 f':. 10.
-.8
f"> '"
23012 64r4/5 65 66(215)-416' ~4z,-'f/~ f':. 10.
,
. NACA NACA NACA NACA
, o o c> "l O,N~C1 -1.2
o -;.6
-.I -.5 -.4 .; ci ' .032 .012 .tJ36 .028 .024 .008 .004
r.J ~-.2 'IJ Cl \) -:3 ~ ~ o
..... :\;i ::::: ..... ~
-
<ll ~
IJ",·OlO c: 8 c: ~
.~ .<.J ~
..... '\f:..016 :g V)
airfoils from tests in the Langley two·dimensional low-turbulence pressure NAC.'\.
-- some \'~ oC ~' :'t-. ,-- ~ ~ l? ru r--.
/6 ~ ",.:i L __ ,.
r ,.....
1M deq characteristics Rtf,.
;r i£.
~h W, ~o, , IffJio/ 'f', flrl - - I f/'j ~erodynamic attack, I~ 'I .JU i§'( of &/ I '-- ~ -
I'i
anqle I '--- e!
- 'f / -8 (
('l
Section 17.-Comparison of the -16 .
FIGURE -24 -32 .8 .4 1.6 1.2 3.2 -.4 -.8 3.6 28 2.4 2.0 -1.2 ~ -1.6 -2.0 ~ ~ ~ 1,)- t:: v ...
:~ ~ '" :g 0
..... <.::: -...: .I 0 -./ -:3 -:4 -:5 ~
./-:2 () l) ~
.,.:' .~ ~ ~
"Q; ......
SUMMARY OF AIRFOIL DATA Effective aspect ratio.-The combination of high drags at 0.0150 to the wing drag coefficients. The resulting drag high lift coefficients, low drags at moderate lift coefficients, coefficients have been approximated by two curves corre- and the nonregular variation of drag with lift coefficient sponding to equation (17) and matched to the drag curves shown by the NACA 6-series airfoils may lead to para- at lift coefficients of 0.2 and 1.0. These two curves corre- doxical results when the span-efficiency concept (reference 29) spond to effective aspect ratios of 9.29 for the airplane with is used for the calculation of airplane performance. In the NAOA 23018 sections and of 8.30 for the airplane with usual application of this concept, the airplane drag charac- NACA 65 -418 sections and illustrate the typical large tc!ristics are approximated by a curve of the type reduction in the effective aspect ratio obtained with such sections.
(17) It should be noted, however, that although equation (17) provides a reasonably satisfactory approximation to the This curve is usually matched to the actual drag character- drag of the airplane with NACA 23018 sections, such is not istics at a rather low and at a moderately high value of the the case for the airplane with the NAOA 65 -418 section.
lift coefficient (reference 30). The most important reason for using high aspect ratios on The application of this concept to two hypothetical air- large airplane3 is to reduce the drag at cruising lift coefficients planes with N ACA 230-. and (}5~.series sections, respectively, and to obtain high maximum values of the lift-drag ratio.
is illustrated in figure 18 (a). The wing drags of the air- For the two wings considered, the maximum value of this planes have been calculated by adding the induced drags ratio is appreciably higher for the airplane with NACA 65 -418 sections (19.8 as compared with 18.5) despite the corresponding to an aspect ratio of 10 with elliptical loading to the profile-drag coefficients of the NACA 23018 and fact that this airplane shows the lower effective aspect ratio.
65 -418 airfoils. These sections are considered representa- Figure 18 (b) shows a similar comparison with similar tive of average wing sections for a large airplane of this results for two airplanes of aspect ratio 8 and NACA 2415 aspect ratio. Ordinate scales are given in figure 18 (a) for and 65 -415 airfoils. It is accordingly concluded that the the wing drag and for the total airplane drag coefficients effective aspeCt ratio is not a satisfactory criterion for use in obtained by adding a representative constant value of airfoil selection .
./0 ./0
~ I ! .1. ,l! 1.1 I I 1.1
8M 0 ~AJA ~52-~/S IWinb; alpe~t )Otid, 8
.0 o NACA 65 -418 wing; aspecf ratio, /0 - 3 .08 t-- o NACA 24/5 wing; aspect ratio, 8 o NACA 23018 wing; aspecf,rotio, /0 r- - --- NACA 65 -415 winq; - ---- NACA 65 -418 wing; . . O{) .09 effective aspect ratio, 6.97
effective aspect ratio, 830 fI
.0 7 - - - NACA 24/5 win91 .07 - - - NACA 2.3018 wing; .~ I .1.
I effective aspect rafio, 746 effective aspect rofio, 929 I .08
III
V
.06 .06 I
J
~ .07
l'll
'-S IS
I
t- "'- .05
L
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"
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'/
If
'/ P:
II .05 8
Airplane Airplane \)) )
Y
X
~
e ~.03
Q (§L. ~1/68
0L~ -~
b = 18.5
c;, = a 0063 + a0.427 C[!\
V V'
A ma.
<lJ .04 i3
I::: \)'> CD =0.0042+ a0456 Qj c;, -00068+ a0343 c,,' .
. ~ ~ 1ft.
-K ~ .02
'- Airplane " Airplane c;, ~ a 0045 + a 0383 lfX .\..:
/(}l
V
L -<c .03 - -;/7.7
'0 l ~
'(~l =19.8
D max
./ ~
-a ma •
......--: ~
.01 .01
y
tr,J
V ,y
~ ~
--
,- .02 -
.02
-
){/
V
o
V
/
.01 .01
V
/
(a) (b)
V
V
o
o
o .z .4 .6 .8 1.0 1.2
o .8 1.0 /2 .4 .6 .2
Lift coe ffl'cient, c;, Lifl coefficienf, c;, (b) NACA 65,-415 and 2415 wings of aspect ratio 8.
(a) NACA 653-418 and 23018 wings of aspect ratio lO.
FIGURE 18.-Comparison of finite aspect'ratio drag characteristics for two types of airfoils obtained by adding the induced drag corresponding to an elliptical span loading to the section drag coefficients.
REPORT NO. 824-NATIONAL ADVISORY COMMITTEE FOR AERONATT'I'ICS EFFECT OF SURFACE IRREGULARITIES ON DRAG small sharp protuberances, in contrast to waves, tends to occur at the protuberance. Transition caused by surface Permissible roughness.-Previous work has shown large waviness appears to approach t:Qe waVe gradually as the drag increments resulting from surface roughness. (reference Reynolds number or wave size is increased. The height 31). Although a large part of these drag increments was of a small cylindrical protuberance necessary to cause transi- shown to result from forward movement of transition, sub- tion when located at 5 percent of the chord with its axis stantial drag inGrements resulted from surface roughness in normal to the surface is shown in figure 19. These data were the region of turbulent flow. It is accordingly important to maintain smooth surfaces even when extensive laminar flow cannot be expected, but the gains that may be expected from maintaining smooth surfaces are grAater for NACA 6- or 7-series airfoils when extensive laminar flows are possible.
No accurate method of specifying the surface condition necessary for extensive laminar flow at high Reynolds num-
'"
bers has been developed, although some general conclusions
"-
""- have been reached. It may be presumed that for a given i"- Reynolds number and chordwise position, the size of the t-
--
permissible roughness will vary directly with the chord of the airfoil. It is known, at one extreme, that the surfaces I 2 3 4 5 6 7 8 9 IOxIO' Wing Reynolds number, R do not have to be polished or optically smooth. Such polishing or waxing has shown no improvement in tests in FIGURE 19.-Variation with wing Reynolds number of the minimum height of a cylindrical protuberance necessary to cause premature transition. Protuberance has 0.035·inch di- the Langley two-dimensional low-turbulence tunnels when ameter with axis normal to wing surface and is located at 5 percent chord of a 9O·inch·chord applied to satisfactorily sanded surfaces. Polishing or waxing symmetrical 6-series airfoil section of 15·percent thickness and with minimum pressure at 70 percent chord.
a sllrfacethat is not Iterodynamically smooth wiH, of course, result in improvement and such finishes may be of consider- able practical value because deterioration of. the finish may obtained at rather low values of the Reynolds number and be easily Seen and pmlsibly postponed. Large models having show a large decrease in allowable height with increase in chord lengths of 5 to 8 feet tested in the Langley two- Reynolds number. This, effect of Reyn?lds number on dImensional low-turbulence tunnels are usually finished by permissible surface roughness is also evident in figure 20, sanding in the chordwise direction with N o. 3~0 carborundum in which a sharp increase in drag at a Reynolds number of paper when an aerodynamically smooth surface is desired. approximately 20 X 10 occurs for the model painted with Experience has shown the resulting finish to be satisfactory camouflage lacquer.
at flight values of the Reynolds number. Any' rougher The magnitude of the favorable gradient appears to have a surface texture should be considered as a possible source of small effect on the permissible surface roughness for laminar transition, although slightly rougher surfaces have appeared flow. Figure 21 shows that the roughness becomes more to produce satisfactory results in some cases. C important at the extremities of the low-drag range where Wind-tunnel experience in testing NACA 6-series sections the favorable pressure gradient is reduced on one surface.
and data of reference 32 show that small protuberances The effect of increasing the Reynolds number for a sllrface extending above the general surface level of an otherwise of marginal smoothness, which has an effect similar to in- satisfactory surface are more likely to cause transition than creasing the surface roughness for a given Reynolds number, small depressions. Dust particles, for example, are more is to reduce rapidly the extent of the 19w-drag range and effective than ,small scratches in producing transition if the then to increase the minimum drag coefficient (fig. 21).
material at the edges of the scratches is not forced above the The data of figure 21 were specially chosen to show this general surface level. Dust particles adheI:ing to the oil effect. In most cases, the effect of Reynolds number pre- left on airfoil surfaces by fingerprints may be expected to dominates over the effect of decreasing the magnitude of the cause transition at high Reynolds numbers. favorable pressure gradient to such an extent that the only Transition spreads from an individual disturbance with an effect is the elimination of the low-drag range (reference 34).
included angle of about 15° (references 31 and 33). A few Permissible waviness.-More difficulty is generally en- scattered specks, especially near the leading edge,' will cause countered in reducing the waviness to permissible values for the flow to be largely turbulent. This fact makes necessary the maintenance of laminar flow than in obtaining the re- an extremely thorough inspection if. low drags are to be quired surface smoothness. In addition, the specification realized. Specks sufficiently large to cause . premature of the required freedom from surface waviness is more transition on full-size wings can be felt by hand. The in- difficult than that of the required surface smoothness. The spection procedure used in the Langley two-dimensional problem is not limited merely to finding the minimum wave low-turbulence tunnels is to feel the entire surface by hand size that will cause transition under given conditions because after which the surface is thoroughly wiped with a dry cloth. the number of waves and the shape of the waves require It has been noticed that transition resulting froin individual consideration.
SUMMARY OF AIRFOIL DA'l'A .DI6 ~ ~~ \J .012 '-
'"
<lJ o " .008 [?
-f5 () -<J ~.004 :;:: \J ..9j (a) /2 16 24 28 32 36 40 44 48 56 60x10' 20 52 Reynolds number, R
.... -
c: 0.
..::: >- ~ ".008 ~ ~
-:c
o/,
Vo
!
c:004 ~ (b)
"
Ji
4 8 12 16 20 24 28 32 36 40 44 48 52x/0' a Reynolds number, R (a) Smooth condition; TDT test 328.
(b) Lacquer camouflage unimproved after painting; TDT test 461.
FIGrRE 20.-· Yariation of drag codl:cient with Reynolds numhr for a 6O-inch·chord model of the NACA 65(-121)-420 airfoil for two surface conditions .
. 016 laminar separation or even reversal of the pressure gradient.
I tRI. I I I I .1 I
Data for an airfoil section having a relatively long wave on I- 0 15.3 X 10' Smooth condition
..... -
I;:
o 149 1 I I I I I 1 I. 1 T the upper surface are given in figure 22. Marked increases
·~.012 1-0 24.8 Synthetic enamel camouflage with .\J f::, 34.6 a/I specks cut off with blade in the drag corresponding to a rapid forward movement of ~ ..... " 446 the transition point were not noticeable below a Reynolds ~ ,;;P
~
".008 number of 44 X 10 • On the other hand, transition has been ~ ~ "'-
~ 'lIS§il '/
~
~ caused at comparatively low Reynolds numbers by a series I"J
'7
is m
of small waves with a wave height of the order of a few ten- 1;:.004 .0 thousandths of an inch and a wave length of the order of -i::: \J 2 inches on the same 60-inch-chord model.
~ For the types of wave usually encountered on practical- -.4
P8 o .4 .8 1.2 1.6 2.0
5ection lift coefficient, G construction wings, the test of rocking a straightedge over the surface in a chord wise direction 'is a fairly satisfactory FIGURE 21.-Drag characteristics of N AOA 65(",)-420 airfoil for two surface conditions.
TDT tests 300 and 486.
criterion. The straightedge should rock smoothly without jarring or clicking. The straightedge test will not show the existence of waves that leave the surface convex, such as the If the wave is sufficiently large to affect the pressure wave of figure 22 and the series of small waves previously distribution in such a manner that laminar separation is mentioned. Tests of a large number of practical-construction encountered, there is little doubt that such a wave will cause models, however, ·have shown that those models which premature transition at all useful Reynolds numbers. A re- passed the straightedge test were sufficiently free of small lation between the dimension€; of a wave and' the pressure waves to permit low drags to be obtained at flight values of distribution may be found by the method of reference 35.
the Reynolds number.
The size of the wave required to reverse the favorable pres- It is not feasible to specify construction tolerances on air- sure gradient increases with the pressure gradient. Large foil ordinates with sufficient accuracy to ensure adequate negative pressure gradients would therefore appear to be freedom from waviness. If care is taken to obtain fair favorable for wavy surfaces. Experimental results have surfaces, normal tolerances may be used without causing shown this conclusion to be qualitatively correct.
serious alteration of the drag characteristics.
Little information is available on waves too small to cause ItEPORT NO. 824-NATIONAL ADVISORY COMMITTEE F'OR AERONAUTICS /Cenfer of wave
-.~c:-r =-1
Departure from fair ' -------60"--------~- airfoil surface
/
f-----' 'I'-.
-
~ 4 44 5exlD· D 8 12 16 cO 24 28 J2 J6 40 48 Winq Reynolds number. R FIGl'RE 22.-Expcrimental curve showing variation of drag corffici~nt with Hcynolds number for the N ACA 65(421)-420 airfoil section with a small amount of surface waviness.
Dra.g with fixed transition.---If the airfoil surface is suffi- result of accumulation of ice or mud or damage in military ciently rough to cause transition near the leading edge, large combat.
drag increases are to be expected. Figure 23 shows that, The variation of minimum drag coefficient with thickness although the degree of roughness has some effect, the incre- ratio for a number of NACA airfoils with standard roughness ment in minimum drag coefficient cau~ed by the smallest is shown in figure 12. These data show that the magnitudes roughness capable of producing transition is nearly as great of the minimum drag coefficients for the NACA 6-series as that caused by much larger grain roughness when the airfoils are less than the values for the NACA four- and roughness is confined to the leading edge. The degree of five-digit-series airfoils. The rate of increase of drag with roughness has a much larger effect on the drag at high lift thickness is greater for the airfoils in the rough condition coefficients. If the roughness is sufficiently large to cause than in the smooth condition.
transition at all Reynolds numbers considered, the drag of Drag with practical construction methods.- The section the airfoil with roughness only at the leading edge decreases drag coefficients of several airplane wings have been measured with increasing Reynolds number (fig. 10 and reference 36).
in flight by the wake-survey method (reference 38), and a The effect of fixing transition by means of a roughness number of practical-construction wing sections have been strip of carborundum of O.Oll-inch grain is shown in figur(~ 24. tested in the Langley two-dimensional low-turbulence The minimum drag increases progressively with forward pressure tunnel at flight values of the Reynolds number.
movement of the roughness strip. The effE'ct on the drag Flight data obtained by the NACA (referenee 38) arc sum- at high lift coefficients is not progressive; the drag increases marized in figure 26 and some data obtained by the Consoli- rapidly when the roughness is at the leading edge. Figure 25 dated Vultee Aircraft Corporation are presented in figure 27.
shows that the drag coefficients for the NACA 65(223)-422 Data obtained in the Langley two-dimensional low- and 63(420)-422 airfoils were nearly the same throughout turbulenee pressure tunnel for typical praetical-construction most of the lift range when the extent of laminar flow was sections are presented in figures 28 to 32. Figure 33 presents limited to 0.30c.
a comparison of the drag coefficients obtained in this wind All recent airfoil data obtained in the Langley two-dimen- tunnel for a model of the NACA 0012 section and in flight sional low-turbulence pressure tunnel include results with for the same model mounted on an airplane. For this case, roughened leading edge, and these data are included in the the wind-tunnel and flight data agree to within the experi- supplementary figures. Tests with roughened leading edge mental error.
were formerly made. only for a limited number of airfoil All wings for which flight data art' pres('nted in figure 2() sections, especially those having large thickness ratios were earefully finished to produce smooth surfaces. Great (reference 37). The standard roughness selected for 24-inch- care was taken to reduce surface waviness to a minimum chord models consists of O.Oll-inch carborundum grains for all the sections except the NACA 2414.5, the N-22, the applied to the airfoil surface at the leading edge over a surface Republic 8-3,13, and the NACA 27-212. Curvature-gage length of 0.08c measured from the leading edge on both sur- measurements of surface wavinE'ss for some of these airfoils are presented in referE'nce 38. Surface eonditions correspond- faces. The grains are thinly spread to cover 5 to 10 percent ing to the data of figure 27 arc descri.bed in the figure.
of this area. This standard roughness is considerably more These data show that the sections permitting extensive severe than that caused by the usual manufacturing irregu- larities or deterioration in service but is considerably less laminar flow had substantially lower drag coefficients when severe than that likely to be encountered in service as a smooth than the other sections.
>o-J t" t) >-3 Ul o ~ ~ ;.. ::>:l ....; o "J ;.. I-< ::>:l o I-< ;.. ;..
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~ r---.~_ Section "- I'\..
.004-inch-qrain .0/ ..
x~ Smooth Shelloc \ o.OO?-inch-qroin /- t-,., 10,,: roughness o x
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63(420)-422 NACA - -+=+- ~ lrY'to I- -R~ Im;;~ 2~ /6 i"'kL anp' deq 2::-~ ~ I...L.L I 63(420)/?2 /1(" drag characteristics of an ~ 0;" In '8 ~ IhfL II and l.& NACA TOT II I I attock, J I ?oxIO· 23.-Lift II I of IIIJI.11111 Airfoil: Test: R: Chord:
IL
/ FIGURE !
angie V j -8 !fI ~
If'
c Section -/6 -24 i I 7 7 5' 4 B 4 8 2 -32 .8 .4 -; /.0 -; 3.6 3.2 2.8 2.4 2.0 -/.2 -I.
-z 10 ~/.2 Cl (J ~
III s
~ :i3 ~, ~ ;:: IrJ
t"l "tI co "" li>- ~ o Z >- t-' >- t::1 ::1 m o I:tI >1 o o ~ ~ ~ t"l t=l "'i o I:tI >- t=l ::tI o Z >- q 1-3 I-< o m
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strip strip T strip x 0 coet'!7cienl; t-i..A
b
l"'" x"1= lift ~t-r---!;-=t:::::t Smoofh Roughness Roughness Roughness -;4 .~ b:::, D: N: varions chordwisc locations.
x o + o Section
C-
f"-- at x ".
1\ ....
~ I R:::- \ Q.. -;8 l1. t?<. B "- ~
r"
I" -1.2 , , o -1.6 .040 .036 .056 .052 .048 .044 .028' .024 .020r .016 .0/2 .008 .004 <J f· <lJ ~ ()) Cl
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:::: {j .§ '- ~
airfoil with O.OIl-inch-grain roughncss .32 63(420)-422 i I , J fO NACA
r--H
ofan ~~-L k>.
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:Q::"', / ~'- deg .,.el-.....- <f' ~3(420)-422 I I <lDJ Iff in.
1/ drag characteristics Ih NACA TOT /; V' and J attr1ck, II j 26x10' 1/ of Test: Airfoil: R: Chord: / 24.-Lift / , angle
_,-u.lLLLLL
II j -8 FIGURE I(j 0 II ~ V Section -16 -24 -.32 .8 .4 -;4 -.8 3.6 1.6 1.2 2.B 2.0 3.2 2.4 -/.2 -1.6 -2.0 .:; ~ (U B c:
"h.~ ~ <t.:: <:::: "" .0 .;:: ~
DATA SUMMARY OF AIRFOIL 32 X 10
\
.040 35;'215 i'\.--NAd
.036 ~
.......
$-3,/'-
~-Repu:bltC; -....
.032 , ~ 1.
~4,2-(I,4)(13,5) /NACA
r-...~
~
t-- ~
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_-~--NACA 66,2-2(14.7) , NACA 24(4.S:
fZ
".028 ~ t- I;,) S -3, 13 I>- 'Rep~b1ic
.... '
NAd 27-212 c: ·~.024 'N-22 ,
:t
o
III NACA 63(420)-422--- o l).020 + S-3,11 N-22· .-- --Republic 8' ,S-3, 13· ..
Rep0blic V /' ~ '" I . ,
ii I
I:: .016
fS-2}5
~l:?r<'NjCA
.0 ~ ,/
I
i~
..::: l) NACA 24145-
V 66,2-2(14.7)
V''NACA
V
III , , ,~---
'"'. ~
'0.012
4'l
~ V f7-2(2 ~ /'-N~CA
\ " ---
"ci: /
V
~
CA ~ 4, 2-(1. 4}(.13.5}
.008 'N1
t;7 f-J::::: k- ~ .80 .96 .48 .64 ./6 .32 (modified) c NACA 65(223)-422 Section lift coefficient, l .004 in flight on various airfoilS.
of section drag coefficients obtained 26.-Comparison FIGURE and 35--215 sections made on gloves.
of N A CA 27-212 Tests o 1.6 1.2 S 0 .4 .8 -.4 x/O -1.6 -/.2 -.8 30 0.
lift coefficient, Section O.Oll·inch-grain airfoils with of two NACA 6·series characteristics FIGURE 25.-Drag • at 0.30e. R=26XI0 roughness :---..
""
i'--
r--
wind-tunnel tests of practical-construction wing sec- The
r--
showed minimum
by the manufacturer -
tions as delivered
-
drag coefficients of the order of 0.0070 to 0.0080 in nearly all cuoped flop camouflage, to 32). o Factory finish, the airfoil section used (figs. I--- cases rpgardless of -
I -
sanded, surfaCing Surface glozed, o current be regarded as typical for good applied, camouflage, may compound Such values
o - cusped flap
produce sections to the Finishing practice.
construction camouflage, waviness reduced, () Surface seoled _ flop gop flop cusp removed, ~ smooth surfaces always produced substantial drag reductions
IJ.OI2 -
flop filled, wa<ed surfoce, ViSible woves ..... ' " None of although considerable waviness usually remained.
c: removed' cusp .~ Unless at the front spar.
fair surfaces tested had the sections .\.J c;:: Lo the front spar, at is taken to produce fair surfaces speeial care 'Q; .008 v o y--
i---"
may be expected to cause transition either \.J 0.........
the resulting wave f>--
f.<>-'
-
One it.
distance behind or a short ~
spar location 8'
at the
:.--
smooth surfaces ~ "'""" with practical-eonstruction specimen tested 004 c: Reynolds numbers up to .0 relatively low drags maintained -i:: airfoil of \.J 66(2x15)-116 (NACA 30X10 of approximately ~ about 35 forward of had no spar specimen fig. 10)~ This .4 .5 .6 .2 .3 ./
o
no spanwise stiffeners and the leading edge c,.
percent chord from Lift coe fficient, of construction resulted in This type of the effect of wing surface condition forward of the spars.
FIGURE 27.-Consolidated·Vultee flight measurements wing section.
an NACA 66(215)-1(14.5) on drag of unusually fair surfaces and is being used on some modern high-performanee airplanes.
Data for such same manufacturers.
by the the same time the mini- A comparison of the effect of airfoil section on results to 32. The pairs of models are presented in figures practical-construetion surfaces is very diffi- drag with mum construction practices are current that as long as indicate effect on quality of the surface has more because the eult at flight has relatively little effect of section type best com- used the the Probably the type of section.
the drag than .
number for military airplanes.
Reynolds values of the at parison can be obtained from pairs of models construeted REPORT NO. 824- TATIONAL ADVISORY COMMITTEE FOR AERO AUTICS <J ~.01Z..-'--r-'r-.--'-.-r-'--.-'r-.--.-.-r-~-. -'
~ r-ro's)rod es ~s ~ecJlve~ e;ceJf Ilbhtl~ sJndlJd-..l_.J..-~
.:::: 0 Both surfaces pOinted and flmshed fa rear spar ~008~t-~-r-t-t~-- r-t=1=~~=i~~~~~ , 010
cJ. = x
-+--J.
8'
- + .... + -';;:4" ii .• ~.004 r--r~ -T--+-r--b--1~1--+-+-~-~4--+-+--+~ t-i- ~C._~ {.006 ~ (10 fl r'·
:g Turbulent.·
~
'q; .004 () , ~ °4 ~~~ 8 -~- / ~ 2 -L-- 16 L-~- a ~ O -~- z. L ~ -L- Z. ~ 8 L-~- 3 ~ 2 -L-- 36 L x - / ~ 0 · c, ondilion of model ~ , 003 Reynolds number, R III ~ '0'079 As r~c~lV~d
e
.. - c - 2.65 R -"Z 6 + ,077 Painted with gray primer FIGURE 28.-Drag scale e tT ect on lOO'inch-chord practical-construction model of tbe '\l , 002 surfacer
""
NAC.'. 65(216) -3( 16.5) (approx .) airf oi l section. cl=0.2 (approx.).
x ,074 Camouflage poinled with ~ .~
'" Lorna/--
]
I siT'!'i fiT (f(tll
II
,001 .1, 41) 50 60 10 zo 100x10· .016 Reynolds number , R J , I 1 1 , I 1 1 I .
NACA 65(216)-411(approx.) as built FIGURE 29.-Variation of drag coeffici ent with Reynolds number for the NACA 23016 airfoil
° NACA 23015 (approx) as bUilt
,.1' 0 section together with lamin ar and t urbul ent skin-friction coefficients for a fi at plate .
..... ' '- ,0 12 .'!!
~
V
,\)
V /'
"- V
1 I I I
~
"
-v <J Chord In c,
/
~ \).008
° NACA 66(215)-1/6 T ,I I 1 100 015
f--
t .0/2
~o NACA 66(215)-iI6fr4wlllmodel) 84% r ,18
'"
e
G '15 ,-0 NACA 23016 l0o._f- ,19 f--- '-..
'" NACA 23016 frebudf'modell 100 ,19
'- " <iJ
I~ 2,.004 8,008 \) ..........
~ 8'
/'
ift'"
----
c: ,00 4
*'
o ,Q
-,2 o .z .4 ,6 .8 1.0
Sec ti on lift coefficient, c, t
J!
FIGURE 3O . -D rag characteristics of the NACA 65(21 6)-4 17 (approx.) and NACA 23015 4 12 16 20 24 32x10 '
o 8 28
(approx.) airfoil sections built by practical'construction methods by th e same manu- Reynolds number, R fact ur er. R= 1O.23Xl()6.
FIGURE 3 1. -8ca le effect on drag of the NACA 66(215)-116 and NACA 23016 ai rfoil sections built by practical'construction m et hods by the same manufacturer and tested as received.
,[1 16 . 016 "
"
.... ' I
,II I I I " I I I <J
, , I c: 0 Oavis wmq, onglnol candtlton .,!! ,012 I-- 1-- ..... ' 0 o Wmd tunnel NACA 65-series wing, onglnol condi/;on \) ~ ,012 ~ o Flight '-.. G <iJ <.::: 0 '-..
,008 \) <iJ
"
. 008 \)
8'
*s 8'
-0 ,004 r c: ~ ,004 ..::: c: \)
:2
\)
Jl
<1 0 8 12 16 ?O 24 28 32x10'
Jl
J2xIO ' ReynOlds number, R 12 16 20 24 28
o 4 8
Reynolds number, R FIGURE 32,-Drag scale effect for a model of the ACA 65-series airfo il section, 18,27 percent FIGUR E 33,-Comparisoo or drag coefficients measured in Hight and wind tWlJlel for t he thick, and the Davis airfoil section, 18,27 percent thick, built by practical·construction NACA 0012 airfoil section at zero lift.
methods by the same man uf act ur er. cl= 0.46 (approx.) .
SUMMARY OF AIRFOIL DATA ·0/6
10 I,clab'5\";Yo!, )'h i-i!.}aJc;' J,~)
surface and 0.1 Dc on lower surfacl3 c,=Q4D o NACA 23015' (appro x.) with de-lcer removed I_L <> NACA 65(2/6)-215, 0=0,8wllh QD75c de-ieer }c=029 /', NACA 65(216)-215, 0=0.8 with de-ieer removed I' : ~ ~< ............
.....,.
j....".- I--
-
•
32 36 40. 44 48 52xlO 0. 4 8 12 16 20 24 28 Reynolds number, R FIGURE 34.- Effect of de-ieers on the drag of two practical-construction airfoil sections with relatively smooth surfaces.
Important savings in drag may be obtained at high
represent good typical installations. The mmlmum drag
Reynolds numbers by keeping the surfaces smooth even if
coefficients for both sections with de-icers installed were of
extensive laminar flow is not realized. Drag increments resul t-
the order of 0.0070 at high Reynolds numbers.
o ing from surface roughness in turbulent flow have been shown
Effects of propeller slipstream and airplane vibration.-
to be important (reference 31). The effects of surface roughness
Very few data are available on the effect of propeller slip-
on the variation of drag with Reynolds number are shown
stream on transition or airfoil drag; the data that are avail-
in figure 29, in which the favorable scale effect usually expected
able do not show consistent results. This inconsistency may
at high Reynolds numbers was not realized_ This type of
result from variations in lift coefficient, surface condition,
scale effect may be compared with that shown for the NACA
air-stream turbulence, propeller advance-diameter ratio, and
63(420)-422 airfoil with rough leading edge but otherwise
number of blades. Tests in the Langley 8-foot high-speed
smooth surfaces (fig. 10). Drag increments obtained in
tunnel indicated transition occurring from 5 to 10 percent of
flight resulting from roughness in the turbulent boundary
the chord from the leading edge (reference 40). Drag measure-
layer with fixed transition are presented in reference 39,
ments made in the Langley 19-foot pressure tunnel (fig. 35)
The effect of the application of de-icers to the leading edge
indicated only moderate drag increments resulting from a
of two smooth airfoils is shown in figure 34. The de-icer
windmilling propeller. Although the data of figure 35 may
"boots" were installed in both cases by the manufacturer to
not be very accurate because of the difficulty of making
wake surveys in the slipstream these data seem to preclude
l
very large drag increments such as would result from move·
ment of the transition to a position close to the leading edge.
These data also seem to be confirmed by recent NACA flight
data (fig. 36), which show transition as far back as 20 percent
ct L.
......
I
"Left wing section I
I--
c:t:
:-- I
I .............. in s//Dsfreom
t=-.
1..'
-.........
<Il
i--
-9 :;---
§ 10
\R/qhf wing section
I
c: outside slipstream
~
Cl ~O ~ Airfoil sections Root, NACA 66(2xI5J-DI8 Tip, NACA 67, 1-1/.3)15 o Riqht winq sectionl--- Aspect ratio, 5.98 J Propeller tip radius- outSide S!iPstre~L Left wing section in slipstream
~ o Power on
0 Power off
IA I . I)-
o rope ler wmdmll mg L\ Propeller removed
\
""
~~
~
~
kp
!'---o
.! .5 .6
7 6 5 4 '.3 2 o .£' ..3 .4
DistanceJr~m.. f!ROdel cepfer line, ft Section lift coeffiCient, Cl FIGURE 35.-The effect of propel«lr operation on section drag coefficient of a fighter-type air- FIGURE 36.-Flight measurements of transition on an NACA 66-series wing within ana plane from tests of a model in the Langley 19-foot pressure tunnel. CL=O.lO; R=3.7Xl06. outside the Slipstream.
REPORT NO. 824-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS of the chord in the slipstream. Other unpublished NACA airfoil section. For the NACA 6-series airfoils this lift coeffi- flight data on transition on an 8-3,14.6 airfoil in the slip- cient is approximately in the center of the low-drag range.
stream indicated that laminar flow occurred as far back For airfoils having thicknesses in the range from 6 to 10 per- as 0.2c.
cent, the NAOA four- and five-digit series and the NAOA
Even less data are available on the effects of vibration on . 64-series airfoil sections have values of lift-curve slope very
transition. Tests in the Langley 8-foot high-speed tunnel
close to the value for thin airfoils (271' per radian or 0.110 per
(reference 40) showed negligible effects, but the range of degree). Variation in Reynolds number between 3XI0 and frequencies tested may not have been sufficiently wide. Some 9X 10 and variations in airfoil camber up to 4 percent chord unpublished flight data showed small but consistent rear- appear to have no systematic effect on values of lift-curve ward movements of transition outside the slipstream when slope. The airfoil thickness and the type of thickness the propellers were feathered. This effect was noticed even distribution appear to be the primary variables. For the when the propeller on the opposite side of the airplane from NAOA four- and five-digit-series airfoil sections, the lift- the survey plane was feathered and was accordingly attrib- curve 'slope decreases with increase in airfoil thickness.
uted to vibration. Recent tests in the Ames full-scale tun- For the NAOA {i-series airfoil sections, however, the lift- nel showed premature adverse scale effect on drag coefficients curve slope increases with increase in thickness and forward measured by the wake-survey method when a model-support movement of the position of minimum pressure of the basic strut vibrated. thickness form at zero lift.
Some N AOA 6-series airfoils show jogs in the lift curve
LIFT CHARACTERiSTICS OF SMOOTH AIRFOILS
at the end of the low-drag range, especially at low Reynolds
Two-dimensional data.-As explained in the section "Angle numbers. This jog becomes more pronounced with increase of Zero Lift," the angle of zero lift of an airfoil is largely of camber or thickness and with rearward movement of the determined by the camber. Thin-airfoil theory provides a position of minimum pressure. on the basic thickness form.
means for computing the angle of zero lift from the mean-line This jog decreases rapidly in severity with increasing Rey- data presented in the supplementary figures. The agree- nolds number, becomes merely a change in lift-curve slope, ment between the calculated and the experimental angle of and is practically nonexistent at a Reynolds number of zero lift depends on the type of mean line used. Comparison , 9 X 10 for most airfoils that would be considered for practical of the experimental values of the angle of zero lift obtained application. This jog may be a consideration in the selection from the supplementary figures and the theoretical values of airfoils for small low-speed airplanes. An analysis of
the flow conditions leading to this jog is presented in refer-
taken from the mean-line data shows that the agreement is
good except for the uniform-load type (a=1.0) mean line. ence 28.
The variation of maximum lift coefficient with airfoil
The angles of zero lift for this type mean line generally have
thickness ratio at a Reynolds number of 6 Xl 0 is shown in
values more positive than those predicted. The experi~
figure 39 for a number of N AOA airfoil sections. The airfoils
mental va,lues of the angles of zero lift for a number of NACA
for which data are presented in this figure have a range of
four- and five-digit and NACA 6-serie::; airfoils are presented
thickness ratio from 6 to 24 percent and cambers up to
in figure 37. The. airfoil thickness appears to have little effect
4 percent chord. From the data for the NAOA four- and
on the value of the angle of zero lift regardless of the airfoil
five-digit-series airfoil sections (fig. 39 (a»), the maximum
series. For the NACA four-digit-series airfoils, the angles of
lift coefficients for the plain airfoils appear to be the greatest
zero lift are approximately 0.93 of the value given by thin-
airfoil theory; for the NACA 230-series airfoils, this factor is for a thickness of 12 percent. In general, the rate of change
'of maximum lift coefficient with thickness ratio appears
approximately 1.08; and for the NAOA 6-series airfoils with
uniform-load type mean line, this factor is approximately to be greatest for· airfoils having a thickness less than 12 percent. The data for the NAOA 6-series airfoils (figs.
0.74.
The lift-curve slopes (fig. 38) for airfoils tested in the 39 (b) to 39 (e» also show a rapid increase in maximum lift 'Langley two-dimensional low-turbulence pressure tunnel are coefficient with increasing thickness ratio for thickness higher than those previously obtained in the tests reported ratios of less than 12 percent. For NAOA 6-series airfoil in reference 8. It is not clear whether this difference in slope sections cambered to give a design lift coefficient of not more
than 0.2, the optimum thickness ratio for maximum lift
is caused by the difference in air-stream turbulence or by
coefficient appears to be between 12 and 15 percent, except
the differences in test .methods, since the section data of
for the airfoils having the position of minimum pressure at
reference 8 were inferred from tests of models of aspect ratio 6.
60 percent chord. The optimum thickness ratio for the
The present values of the lift-curve slope were measured for
NAOA 56-series sections cambered for a design lift coeffi-
a Reynolds number of 6 X 10 and at values of the lift coeffi-
cient of not more than 0.2 appears to be 15 percent or greater.
cient approximately equal to the design lift coefficient of the
SUMMARY OF AIRFOIL DATA
I
~5~rie~
f-- eli
ODD}
I- 00 A f-- I-- ~ 0 14 (4-digit)
o .2
v 0.24 1 '1 , .4 _D 44 f--
- 1-'"
.6 'V !
'V 230 (5-digif) \ , " - - (a) (b) .. , 4 8 12 16 20 24 4 8 12 16 20 24 Airfoil thickne%, percent of chord Airfoil thickness, percent of chord (b) NACA 63-serles.
(a) NA CA four- and five-digit series.
i
I 1
f- eli c- e'i I-- f-- I- 00 - . I
I-- ,-0
f--
o .2
.2 _6 .4 I-- .4 -6 I-- \l .6 , 'V .6 () 0 ~ (d) (c) 4 8 12 16 20 24' 4 8 /2 16 20 24 Airfoif thickness, percent of chord Airfoil thiCKness; percent of chord' (d) NACA 55-series.
(c) NACA 64-series.
r eli I-- I- .2 f-'-
1-°
D .4 j---.
tr- -0 v (e)
4 8 Q m m
..clirfo!1 thickness, percenf of chord (e) NACA 66-series.
FIGURE 37.-Measured section angles of zero lift for a number of NACA airfoil sections of various thicknesses and camber. R=6XI0 • ImpORT NO. 824-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS Flab-qed s';"mb~/s indicate rough conditlori ::Smoofh - _.
I-'
... -'! l- r t-.~
-
r ':; ,'Smooth
-
'""""1 "Rough
-
- '-4i
-
-
-
- -
Series
-
0 00 e'i Rou¢"':·~ 0 14 0 0 (4-digit) .2 <> 1 'I <> /:; 44 /:; .4 (a.)
(b) 230 (5-digif) .6
"
8 10 12 14 16 1/3 20 22 8 /0 /2 14 16 18 22 24 Airfoil thickness, percent of chord Airfoil thickness, percent of chord (a) N A CA four- and five-digit series. (b) NACA 63-scries.
-1'tI°
~~.1 4 ,Smooth :,Smoofh :..l? : «!l .....
- - --
., t--
--
:#oug'h -~ t--_ I. ''Rough - I-- o ~; c,j 0 .1 0 0
.2 .z
<> <> .4 /:; .6 .4 (el (d) .. 6 17 .6 \7 8 10 12 14 /6 18 20 22 8 10 12 14 16 18 20 22 24 Airfoil thickness, percent of chord Airfoil thickness, percent of chord (c) NACA M-series.
(d) NACA 65-series.
\Smooth ..;
--
::j
-
---.-'>
"- "Rough eli 0 ~
.z
<> (e) ,4
'"
8 10 12 14 16 18 20 22 24 Airfoil fhickness, percent of chord (e) N A CA 66-series.
FIGURE 38.-Variation of lift-curve slope with airfoil thickness ratio and camber for a number of NACA airfoil sections in both the smooth and rough conditions. R=6XI06.
The available data indicate that a thickness ratio of 12 ~NACA 6-series sections increase with increasing camber percent or less is optimum for airfoils having a design lift (fig. 39 (b) to 39 (e». The addition of camber to the sym- coefficient of 0.4. metrical airfoils causes the greatest increments of maximum The maximum lift coefficient is least sensitive to variations lift coefficient for airfoil thickness ratios varying from 6 to in position of minimum pressure on the basic thickness form 12 percent. The effectiveness of camber as a means of for airfoils having thickness ratios of 6, 18, or 21 percent. increasing the maximum lift coefficient generally decreases The maximum lift coefficients corresponding to intermediate as the airfoil thickness increases beyond 12 or 15 percent.
thickness ratios increase with forward movement of the The available data indicate that the combination of a 12- position of minimum pressure, particularly for those airfoils percent-thick section and a mean line cambered for a design having design lift 'coefficients of 0.2 or less. lift coefficient of 0.4 yields the highest maximum lift The maximum lift coefficients of moderately cambered coefficient.
SUMMARY OF AIRFOIL DATA 2.B \-
%'1 [.AI""
0 -:/
~ f- <)
:c
v/
o %1';;>- / -:,j!
-/ .--
rl; .4 o 14 -e-- ,
Airfoil wITh --- Airfoil with
() ;:4 4-digit 'V .6 '(f '~ , /
/ ~ ' . split flop
spM flap 0 0
6 44r~
.2 , >:> (P , ,/, , if' -)!> .4
~23~~~
~p o 00 '<7 , .6 , tJ 144-digit (<:: ,
/
,..'
-v ::7
~-- ~ ~~ --
" 4
... 1-0-
f--.
t--.. r- h...
~ 1.6 ,......
~
-=
t::S., """'- ..:::; /,/'
c: /' ~
.0
- Plain
..:: .-- "t>-
~ -,,,,- V
- ~::-
Plain '6
/ ..0-
I oirfoil
- .
~ 1.2 ~' airfoil Vl .. ~ ........
,IJ- --":; '"II ~~~
/' ~~ .-
/ I.
L 0 .......
u --0 ~
-0- --
V--
~
.::: ....
- ... ---- Rough _.-- -:..-- Rough Smooth
(al II SiOO{h
(b)
II
.4
o 4 8 /2 /6 20 24 o 4 8 12 16 20 N
Airfoil thickness, percent of chord Airfoil thickness, percent of chord (b) NACA 63-series.
(a) NAOA fonr- and five·digit serie3.
2.8 'z.8
C" eli ~
t:::-; I-
V" ~
0 k 0 ~
l--?
i ° 0
•
",- 0 . / ,,-
<)
.2 v_
,>:>
/' ~
V if ~ / ---
_0 .2 I; .4 <-:,2.4 .... -Z.4 Airfoil with I?l' Airfoil with ,/" I:; I; .4 /'
/; 'V
.6 /) " /) , .\1) , sp/d flop h / --"':
.~ split flop
c- 00 t- 'V ·6 - , .\.)
.i3 " , / I Y
-'"
'00 <.;::: ..:: .2 ~, ¢ I '{:f ,
~
'Qy.2.0 't:J . I r, .4 ~ 20 ..,
.J))' ):Y
() () '<7
>:> - ."- , ~ .6
,~ .;,yJ ~'.:
~ /' \.)
, r, l:"' __ ><.J' .4 1=""
" r-o:~
- -
{: v <t:: .6 i-- -:::: 1.6 "" 1.6 ~ ~ ...;:::" c: I-"' ~.
~ ~ <:::::-- .e
-;::: ~ ~
k - - 0
o 7 Plain
-;? -l>.
-=~~ Plain
J/ """" ./
<lJ -'::L.£ 1.2 ~ 1.2 airfoil -0
J)- 1--- - - airfoil
/~: ,0- )o- -0
-- --J~ V ,- - - '"
V , ~ ~ - -
Ii .'
- .
~
5 ---
0-- -
.0- --
V
-§
/-
·s 8 I
.8 x . x -- -----
Rough _ .. -. --
Rouqh
~ ~
STOO(h
(c) II Smooth
(d)
II
·17
4 8 /2 16 20 24 4 8 /2 /6 20 C4 Airfoil thickness, percent of chord Airfoil thickness, percent of chord (d) NACA Go-series.
(e) N A CA 64 serIes.
Symbols with flogs correspond to Simulated split flop deflected 60'?.
2.8 eli " ~ " f-O 0
•
",- p--.
.2
0 V
~
h
1-6 .4 1:'2.4 ,,~ , , 00 : -t>
/; Airfoil with
.~ .2 l:> .'> '/ splif flap
--
<.;::: V .4 I Xf'
V
'"
~ 20 j.Y o if
I.) ~
.-
--
P-Q- ~
{: -'
----
"" 1.6 c: -6 .0 ..:: o
V f-O"
,.....--
-- Plain
'iJ,!.2 -<>
-
airfoil V _.<>.-
--
l--£ -- -- -0
-
-- -0'
,-- .c-
.§ 8 -
?
l< . I
- -- - --- Rough
~
.-
(e) II Sroih
.4
o 4 8 12 /6 20 24
Airfoil Thickness; percent of chord (e) NACA GO-series.
FIGURE 39.-Variatioll of maxiwum section lift coefficient with airfoil thickness ratio and camber for several NACA airfoil sections with and without simulated split flaps and standard roughness. R=6X106 • REPORT NO. 824-NATIONAL ADVISOIW COMMITTEE FOR AERONAUTICS The variation of maximum lift with type of mean line is approximately 0.15 to 0.20. The scale effect on the NACA shown in figure 40 for one 6-series thickness distribution. 00- and 14-series airfoils having thickness ratios less than No systematic data are available for mean lines with values 0.12c is very small.
of a less than 0.5. It should be noted,however, that airfoils The scale-effect data for the NACA 6-series airfoils (figs.
such as the N ACA 230-series sections with the maximum 41 (c) tv 41 (f» do not show an entirely systematic variation.
camber far forward show large values of maximum lift. In general, the scale effect is favorable for these airfoil Airfoil sections with maximum camber far forward and with sections. For the NACA 63- and 64-series airfoils with thickness ratios of 6 to 12 percent usually stall from the small camber, the increase iIi maximum lift coefficient with
leading edge with large sudden losses in lift. A more de-
increase in Reynolds number is generally small for thicknes.s
sirable gradual stall is obtained when the location of maxi- ratios of less than ·12 percent but is somewhat larger for the
mum camber is farther back, as for the NACA 24-,44-, and
thicker sections. The character of the scale effect for the
6-series sections with normal types of camber.
NACA 65- and 66-series airfoil sections is similar to that for
the NACA 63- and 64-series airfoils but the trends are not
2.0 so well defined. In most cases the scale effect· for NACA
6-series airfoil sections cambered for a design lift coefficient
of 0.4 or 0.6 does not vary much with airfoil thickness ratio.
The data of figure 42 show that, the maximum lift coefficient
for the NACA 63(420)-422 airfoil continues to increase with
r- ~
y--
Reynolds number, at least up to a Reynolds number of
f- I,.---
26X10 •
The values of the maximum lift coefficient presented were
obtained for steady conditions. The maximum lift coeffi-
cient may be higher when the angle of attack is increasing.
Such a condition might occur during gusts and landing
maneuvers. (See reference 41.)
Reynolds number
6 The systematic investigation of NACA 6-series airfoils
o 6.0x10 o 9.0
included tests of the airfoils with a simulated split flap de-
flected 60°. It. was believed that these tests would serve as
an indication of the effectiveness of more powerful types of
trailing-edge high-lift devices although sufficient data to verify
this assumption have not been obtained. The maximum lift
coefficients for a large number of NACA airfoil sections
obtained from tests with the simulated split flap are presented
in figure 39.
o .2 .4
.6 .8 1.0
The data for the NACA 00- and 14-series airfoils equipped
Type of camber, a
with split flap for thickness ratios from 6 to 12 percent show
FIGURE 40.-Variation of maximum lift coefficient with type of camber for some NACA
a considerable increase in maximum lift coefficient with in-
65a-418 airfoil sections from tests in the Langley two-dimensional low-turbulence pressure tunnel.
crease in thickness ratio. Corresponding data for the NACA
44-series airfoils wit,h thickness ratios from 12 to 24 percent
A comparison of the maximum lift coefficients of NACA show very little variation in maximum lift coefficient with 64-series airfoil sections cambered for a design lift coefficient thickness. For NACA 6-sel'ies airfoils equipped with split
flaps the maximum lift coefficients increase rapidly with
of 0.4 with those of the NACA 44- and 230-series sections
(fig. 39) shows that the maximum lift coefficients of the increasing thickness Over a range of thickness ratio, the range NACA 64-series airfoils are as high or higher than those of beginning at thickness ratios between 6 and 9 percent, depend-
ing upon the camber. The upper limit of this range for the
the NACA 44-series sections in all cases. The NACA 230-
series airfoil sections have maximum lift coefficients some- symmetrical NACA 64- and 65-series airfoils appears to be what higher than those of the NACA 64-series sections. greater than 21 percent and for the NACA 63- and 66-series The scale effect on the maximum lift coefficient of a large airfoils, approximately 18 percent. Between thickness ratios number of NACA airfoil sections for Reynolds numbers of 6 and 9 percent the values of maximum lift coefficient for 6 6 from 3 X 10 to 9 X 10 is shown in figure 41. The scale the symmetrical N ACA 6-series airfoils are essentially the effect for the NACA 24-, 44-, and 230~series airfoils (figs. same regardless of thickness ratio and position of minimum 41 (a) and 41 (b» having thickness ratios from 12 t024 percent pressure on the basic thickness form. The maximum lift is favorable and nearly independent of the airfoil thickness. coefficient decreases with rearward movement of minimum 6 6 IIwreasing the Reynolds number from 3 X 10 to 9 X 10 pressure for the airfoils having t,hickness ratios between 9 and
results in an increase in the maximum lift cO('fficient of
18 percent.
SUMMARY Ol!' AIRFOIL DATA 35
.0 .0 2. 2
I
;.c:: NACA 23O-series (5-diq/t)' ~ NACA 24-series (4-digit) ..<> ::::."..
>-=- R .8 .6 --;::: o J.OxIO· i"--..
~ ~
""""
t--. I::::-
~ k
I'b.
I-- .....
~ o 6'~T ~ ....... ~ <> 9.0 r-.....
f:::: :.-- t--.
Dr.
.21-- e:. 6.0 Standard I. 2 ~ r-.....
r ughness ~ r-..
.........
~ I"--
.8 .r.
~ .6 -0-.
.6 J ~ ~ ~ ~
/'
--
.........
f0- r-, NACA 14-series (4-digit) ~ .2 I .2 '-......
~ .6 ""- NACA ~4-ser/es (4-dlqit) 8 (b) ~ !o=,
IA
/. 6 r--..,;;
r-
NACA DO-series (4-digit)
~
J-o.. ~
A
i=O C,j=O.4 and 0.6
-
~ i-' '-"l (a) ~ I. 2 Symbols with flags correspond to c ·O.6 u (5 I.
2. 0 r--;
t:- "0
r--....
~ ~ '"
r-- e'i =0.2
f:::;: t--
~
~ ....
t-i>
;;:: V
.P-
b:" r-::::- ~ ~
i,.- 1--6::: el; =0.4 and 0.0 ~
~
~ ~ 6 y ;l; i'-- :::::::::
- c-
h,.. 1-0
V-
~ ........, "'1 =0.1 cl;=O.c
i'D VII
~ ~
~
c: 2 v if
V
~
~ :2
~ (J (\) v, ~ 1.6 !,-,::: ........ ::J 1-0 v .§ ?-- \'0...
CI'=O
t-o
C'i = 0 x t--.
/'-., t1'
~
~
~ 1.2 I'<> I-
IB
r--....
kf /'
b I..---' (d) ~ (e)
-- ~
....
.8
.B
.....
~ 1.6 ....
!.6 C'I =0.4 -.......::: c =o.4 and 0.8 0- 1i ?-- h:,.
r-.
V
-:::,~
V
V r- ...tc \-0
/.2 1.2 1.8 1.6 ..4' ......
R"" k
~ .......
:>--
V fO c =0.2 V)'"
e'l = 0.2 1j 1-0 t-......
~
1.2 1.2
,--
~V
>-- ~ l---
# ..-'
.8 1.6 I ....., "0 J...6-.
cll=O t-- ~ \
I~ 'F/
/.2 '" i 1.2 c11·O ..A ~I t....- f-"
V m
lL .b-
~ /.
V .....
(f) ~ ~ (e)
-
~ 4 8 12 18 20 ?4 28 32 8 Ie 16 eO C4 028 Airfoil thickness) percent of chord Airfoil thickness, percent of chord (b) NACA four- and five-digit series.
(a) N ACA four-digit series.
(d) NACA 54-series.
(c) NACA 63-serles.
(f) NACA 66-series.
(e) NACA 65-series.
FIGVI\& 41.-Variation of maximum section lift coefficient with airfoil thickness ratio at several Reynolds numbers for a number of NACA airfoil sections of different cambers.
I
pj \:':l "d o z o (¥) t.:l """ z ~ ..... o Z > t:'" > ~ ...... Ul o ~ (l o ?' .... ?' ,.., ..... H H M M "'1 o ~ > M ~ o ~ >- d ..., ,... (") Ul
i:¥
~ ~ 1 I /.f' I;> .1\, lp '1' f/ 1.2 18W l(fJ 'h ~V; -1'~ ~ .8 1,,- r:; c r1::f':::: .032 .036 .028 f-:::: 1.0 >-
r--
X,!.06 V, 0 I 255, 6.
r-- I~
10.0 14.0 coefficient, R 26.0 20.0 0..4 and .8 o <> " v a
r--- I--- lift
~ r-"~ tests ~ .6 ::::i~ ~-...; -.4 Secfion 1'1)1' I-- :x/c :--r I'D i'ilt-: .4 -.8 I'- ~ numbo.r.
I~lu.., 1\ -'-c ~,\ .2 L>.p... 'V~ Reynolds -/.2 r-- hi~h at
1/ f'--
o
"I.e .2 -.2 airfoil .004 ~o .024 .016 .012 .008 ~
c: <lJ o <.> is' II
"...
..... :~ ~ -fs .~ .,.. ~
(,1(420)-422 J:: I'-' I€l f-o NACA the 'v If (>-~ )-I...p- ~ r!: ~ ~(T !-J2~ deq , 10>,..
o ~p;;?
-
~ r-- IX
drag characteristics of J/ '- and ~ !
?
coftuck, L 42.-LiCi of
I'"
j F,GURE } angle I
I
-8
w
);i .,.cW A; Section ~ 1 -16 ..,.., , IC -'l -24 0 5 .8 .4 -32 1.6 -.4 -.8 3.6 3.2 2.8 2.4- 20 -1.2 -I. -2.0
~- ~ ., 0 t.) c: \.J
~
...:-1.2 ~ "-- <:::: " :g
SUMMARY 0]' AIRFOIL DATA
Substantial increments in maximum lift coefficient with
increase in camber are shown for the N ACA 6-series airfoils
of moderate thickness ratios (10 to 15 percent chord) with
split flaps. For the airfoils having thickness ratios of 6
I--- ~ I--
percent and for the airfoils having thickness ratios of 18 or 21 ...... l--
.--
percent, the maximum lift coefficient is affected very little by
a change in camber. For thickness ratios greater than 15
percent, the maximum lift coefficients of the N ACA 63- and ,
64-series airfoils cambered for a design lift coefficient of 0.4
I
.-
equipped with split flaps are greater than the corresponding
o 0.002 roughness
maximum lift coefficients of the NACA 44-series airfoils. o .004 roughness
o .01/ roughness
Three-dimensional data.-No recent systematic three-
LI Smooth
dimensional wing data obtained at high Reynolds numbers
are available, so that it is difficult to make any comparison
with the section data. When the maximum-lift data for
three-dimensional wings are compared with section data,
o 4 e4x/O'
8 Ie 16 cO
Reynolds number, R
account should be taken of the span load distribution over
FIGURE 43.-Effects of Reynolds number on maximum section lift coefficient CI max of the
the wing. The predicted maximum lift coefficient for the
N ACA 63(420)-422 airfoil with roughened and smooth leading edge.
wing will be somewhat lower than the maximum lift coeffi-
cients of the sections used because of the nonuniformity of
lift-coefficient data at a Reynolds number of 6 X 10 for a
the spanwise distribution of lift coefficient. The difference
large number of NACA airfoil sections with standard rough-
amounts to about 4 to 7 percent for a rectangular wing with
ness are presented in figures 39 and 41. The variation of
an aspect ratio of 6.
maxi~um lift coefficient with thickness for the NACA four··
Maximum-lift data obtained from tests of a number of
and five-digit-series airfoil sections shows the same trends
wings and airplane models in the Langley 19-foot pressure
for the airfoils with roughness as for the smooth airfoils
tunnel are presented in table II. Although section data at
except that the values are considerably reduced for all of
the Reynolds numbers necessary to permit a detailed com-
these airfoils other than the N ACA OO-series airfoils of
parisonare not available, the maximum lift coefficient for
6 percent thickness. For a given thickness ratio greater than
plain wings given in table II appears to be in general agree-
15 percent, the values of maximum lift coefficient for the
ment with values expected from section data. The data for
four- and five-digit-series airfoils are substantially the same.
the airplane'models are presented to indicate the maximum
Much less variation in maximum lift coefficient with thick-
lift coefficients obtained with varIOUS airfoils and
ness ratio is shown by the NACA 6-series airfoil sections in
eonfigurations.
the rough condition than with smooth leading edge. The
maximum lift coefficients of the 6-percent-thick airfoils are
LIFT CHARACTERISTICS OF ROUGH AIRFOILS essentially the same for both smooth and rough conditions.
Two-dimensional data.-Most recent airfoil tests, espe- The variation of maximum lift coefficient with camber, how-
cially of airfoils with the thicker sections, have included tests
ever, is about the same for the airfoils with standard rough-
with roughened leading edge (reference 37), and the available
ness as for the smooth sections. The maximum lift coeffi-
data are included in the supplementary figures. cient of airfoils with standard roughness generally decreases The effect on maximum lift coefficient of various degrees somewhat with rearward movement of the position of mini- of roughness applied to the leading edge of the NACA mum pressure except for airfoils having thickness ratios 63(420)-422 airfoil is shown in figure 23. The maximum lift greater than 18 percent, in which case some slight gain in coefficient decreases progressively with increasing roughness maximum lift coefficient results from 3, rearward movement (reference 36). For a given surface condition at the leading of the position of minimum pressure.
edge, the maximum lift coefficient increases slowly with Except for the NACA 44-series airfoils of 12 to 1.'5 percent increasing Reynolds number (fig. 43). Figure 24 shows that thickness, the present data indicate that the rough NACA roughness strips located more than 0.20e from the leading 64-series airfoil sections cambered for a design lift coefficient edge have littie effect on the maximum lift coefficient or of 0.4 have maximum lift coefficients consistently higher than
the rough airfoils of the NACA 24-, 44-, and 230-series air-
lift-curve slope. The results presented in figure 38 show
foils of comparable thickness. Standard roughness causes
that the effect of standard leading edge roughness is to de-
decrements in maximum lift coefficient of the airfoils with
crease the mt-curve slope, particularly for the thicker air-
split flaps that are substantially the same as those observed
foils having the position of minimum pressure far back.
for the plain airfoi.ls.
These data are for a Reynolds number of 6X 10 • Maximum-
REPORT NO. 824-NATIONAL ADVISORY COMMl'l"l'J£E FOR AERONAUTICS P""I I I /I¢ !t>..
1.6
"" •
I
~R, I
~
If-
~ ~
~ I , I I I I
j
\ I I I : b \ I
II \ I
I
\~
1/
-,
J
I
II
V
II
o As delivered by shop, o As delivered % shop, . o As delivered by shop, TOT test 4158
J TDT test 64 TDT test 494
I /
~ 4
o Final condition, o Final condition, o Final condition, \.J'
I
TOT test ':CO TOT test 498 TOT test 523
V I
~
J
J
o
II
II
(e) 17
(e) (b) -4
'-8 -8 o 8 16 24 -8 o 8 16 24 32
o 8 16 24 Section angle of attack, a" de-:) (b) NACA 2415. (c) NACA 23012.
(a) NACA 2412.
FIGURE H.-Lift characteristics of the NACA 23012, 2412, and 2415 airfoil sections as affected by normal model inaccuracies. R=9X10 (approx.).
The maximum lift coefficient may be lowered by failure to
maintain the true airfoil contour near the leading edge, but
no systematic data on this ·effect have been obtained. Ex-
amples of this effect that were accidentally encountered are
r- 1=
presented in figure 44, in which lift characteristics are given
I....)
C J
-;-==
S=-'"' for accurate and slightly inaccurate models. The model
1.4
inaccuracies were so small that they were not found previous
1/
lnl to the tests.
JJ r'tJ Three-dimensional data.-Tests of several airplanes in the
1.2
Langley full-scale tunnel (reference 42) show that many fac-
J.~/
~
tors besides the airfoil sections affect the maximum lift co-
'l,
V
I/; efficient of airplanes. Such factors as roughness, leakage, !.O leading-edge air intakes, armament installations, nacelles,
j
V
and fuselages make it difficult to correlate the airplane maxi-
" ....
'>-'
W
mum lift with the airfoils used, even when the flaps are
.~ .8
retracted. The various flap configurations used make such
<.J
J
~ a correlation even more difficult when the flaps are deflected.
QJ
tv
() ;},
When the flaps were retracted, both the highest and the
\.J .6
lowest maximum lift coefficients obtained in recent tests of
S
VI
--J
airplanes and complete mock-ups of conventional configura-
J
/
tions in the Langley full-scale tunnel were those obtained
W
with NACA 6-series airfoils.
Results obtained from tests of a model of an airplane in
1I
the Langley 19-foot pressure tunnel and of the airplane in
.2 [; Model (Longley 19-foot pressure tunnel) -
Airplane I I I I I I I
the Langley full-scale tunnel are presented in figure 45.
j
o Sealed condition } (Langle fi II-scale
Both tests were made at approximately the same Reynolds
o Service conddlon Lj t::nnel)
jf
number. The results show that the airplane in the service
o
condition had a maximum lift coefficient more than 0.2
1I
lower than that of the model, as wljll as a lower lift-curve
)
slope. Some improvement in the airplane lift characteristics
o 4 8 12 115 20 24
was obtained by sealing leaks. These results show that air-
Angle of atfack, a, deg
plane lift characteristics are strongly affected by details not
FIGURE 45.-The effects of surface conditions on the lilt characteristics of a fighter-type 6 reproduced on large-scale smooth models.
'airplane. R=2.8X10 • SUMMARY OF AIRFOIL DATA I. 6 I. 6 1.4 I. 4
rs-
V
Hl ..-
f-<'
~
,,;.: 1.2
)/
) t'o-.
m
"'0 1-0
II
* t1 /
1.0 1.0 ,V " 17;
I
~
V
....... ~ 17/f7 Airfoil sections
d Ii .8 8 Root, NACA 63(420)-422 -f----
.~
V
Tip, Modified NACA 65"",-517
f! I
I!~
~ "-
://
~
lV
~ .6 .......
'7
:t::; --;
/1 Airfoil sections
-.J f---- Root, Davis (c2-percent)
ilj
Tip, Dovis (9.J-percent) J, .4 .4 o Nofural transition /;
/1
o Transition fixed of .10e I
jP
0 Natural fransifion
W'
.2 .2 0 T ransi!iOI7 ,.fixed of.lDc /;
1I
j
~
o
II
-- r------- I I
i I
+
4 8 12 16 20 o 4 8 12 /6 20 24 28 An91e of otlodr, C(, d€'9 Anqle of at tack, (x, deq FIGURE 46.-The effect on the lift characteristics of fixing the transition on a model in the FIGURE 47,-The effect on the lift characteristics of fixing the transition on a model in the 6, Langley 19-foot pressure tunnel. R=2.7XlO.6. (Model with N ACA airfoil sections.)
Lang]ey 19-foot pressure tunnel. R=2.7XI0 (Model with Davis aifroil sections.)
drag coefficient at high lift coefficients. The resulting drag
Lift characteristics obtained in the IJangley 19-foot pres-
sure tunnel for two airplane models in the smooth condition coefficients may be excessive at cruising lift coefficients for and with transition fixed at the front spar are presented in heavily loaded, high-altitude airplanes. Airfoil sections that
have suitable characteristics when smooth but have excessive
figures 46 and 47. In both cases, the Ilft-curve slope was de-
creased throughout most of the lift range with fixed transi- drag coefficients when rough at lift coefficients corre-
sponding to cruising or climbing conditions are classified as
tion. The maximum lift coefficient was decreased in one
case but was increased in the other case. unconservative.
The decision as to whether a given airfoil section is conserv-
UNCONSERVATIVE AIRFOILS
ative will depend upon the power and the wing loading of
the airplane. The decision may be affected by expected
The attempt to obtain low drags, especially for long-range
airplanes, leads to high wing loadings together with relatively service and operating conditions. For example, the ability
of a multiengine airplane to fly with one or more engines in-
low span loadings. This tendency results in wings of high
operative in icing conditions or after suffering damage in
aspect ratio that require large spar depths for structural
efficiency. The largp- spar depths require the use of thick combat may be a consideration.
. As an aid in judging whether the sections are conservative, root sections.
This trend to thick root sections has been encouraged by the lift coefficient corresponding to a drag coefficient of 0.02 the relatively small increase in drag coefficient with thickness was determined from the supplementary figures for a large ratio of smooth airfoils (fig. 12). Unfortunately, airplane number of NACA airfoil sections with roughened leading wings are not usually constructed with smooth surfaces and, edges. The variation of this critical lift coefficient with air- in any case, the surfaces cannot be relied upon to stay smooth foil thickness ratio and camber is shown in figure 48. These under all service conditions. The effect of roughening the data show that, in general, the lift coefficient at which the leading edges of thick airfoils is to cause large increases in the . drag coefficient is 0.02 decreases with rearward movement of REPORT NO. 824-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS position o~ minimum pressure. The thickness ratio for Series 1.2 which this lift coefficient is a maximum usually lies between
--OO! "
-i~ (4-d/gJ/J
r--- 12 and 15 percent; variations in thickness ratio from this <>
::==
44 I I L!.
~
'" /
optimum range generally cause rather sharp decreases in the
~" ~ ,.0
'V 230 15~di9it)
~
critical lift coe'lficient. The addition of camber to the
V
~ symmetrical airfoils usually causes an increase in the critical
1/
.8 lift coefficient except for the very thick sections, in which case
"
/ ~
increasing the camber becomes relatively ineffectual and may
/
(a) .6 be actually harmful. All the data of figure 48 correspond to a Reynolds number of 6XI0 • As shown in figure 49, the drag coefficient at flight values of the Reynolds number may
""-
1.2 -A be considerably lower than the drag coefficient at a Reynolds
~
, num bel' of 6 X 10 if the roughness is confined to the leading V
/
......- edge.
r------, qj
V
I'\'
V"
r--
00 PITCHING MOMENT
/
"'" ~ o .C
/
r--
L!. .4 The variation of the quarter-chord pitching-moment coef- v .6 ~ ficient at zero angle of attack with airfoil thickness ratio and / camber is presented in figure 50 for several NACA airfoil / (b) sec.tions. The quarter-chord pitching-moment coefficients of the NACA four- and five-digit-series airfoils become less
r--
negative with increasing airfoil thickness. Almost no varia-
:-;;--
~ tion in quarter-chord pitching-moment coefficient with air-
I--
./ ~ ~
foil thickness ratio or position of minimum pressure is shown ~
:7 ~ eli
t-- by the NACA 6-series aiffoil sections. As might be expected,
" I-' V
o 0 ;I: [\, / increasing the amount of camber causes an almost uniform 0 t-- .I.
.2 negative increase in the pitching-moment coefficient.
/
t-- L!. .4 As discussed previously, the pitching moment of an airfoil 'V .6 I section is primarily a function of its camber, and thin-airfoil
(e) /
theory provides a means for estimating the pitching moment from the mean-line data presented in the supplementary / r-..
figures. A comparison of the experimental moment coeffi- cient and theoretical values for the mean lines is presented
~
y-- ""'-- CII' in figure 51. The experimental values of the moment coeffi-
v' ~
00 r--
~ cients for NACA 6-series airfoils cambered with the uniform- o .2
V
V ~
t-- .8 L!. .4 load type mean line are usually about three-quarters of the V 'V .6
V ~
theoretical values (figs. 50 and 51). Airfoils employing mean (d)
/
lines with values of a less than unity, however, have moment
.6 coefficients somewhat more negative than those indicated by theory. The use of a mean line having a value of a less than /D ?-- unity, therefore, brings about only a slight reduction in
/ ~ f'."
pitching-moment coefficient for a given design lift coefficient
~
f-.o. eli
/
t-- .8 when compared with the value obtained with a uniform-
/ ~ / load type mean line. The experimental moment coefficients
(/ .2 t-- ~ Ll .4 for the NACA 24-, 44-, and 230-series airfoils are also less
1/
.6 negative than those indicated by theory but the agreement V \, ¢ is closer than for airfoils having the uniform-load type mean (e) line.
4 8 12 16 20 24 28 32 Airfoil thickness, percenf of chora' The pitching-moment data for the airfoils equipped with simulated split flaps deflected 60° (fig. 50) indicate that the (ll) N ACA four- and five-digit spries.
(b) NACA 63-series.
value of the quarter-chord pitching·moment coefficient be- (c) NACA 64-series.
comes more negative with increasing thickness for all the (d) NACA 65-series.
(e) NACA 66-series.
airfoils tested. For the thicker NACA 6-series sections the FIGURE 4B.-Variation of the lift coefficient corresponding to a drag coefficient of 0.02 with magnitude of the moment coefficient increases with rearward thickness and camber for a number of N'ACA airfoil sections with roughened leading edges.
6• movement of the position of minimum pressure.
R=6XI0 ~ ~ ':rJ >- ...... o ...... t"
~ ~ o ~ t::l ~
~ .....
- 1.6
--1-+-
1.2 I I
, I
I x f/ I I III I ~ + I .8 /" -I'/, J /,l,Ij ~ / c[
:::0: '-'-;;/ :.i"
.4-
-
~ x I I + 6,oxI0' I I 10.0 14.0 R 0 leading edge.
coefficient, I I I +- x o 020.0 026.0 the lift
'c.-
~" ~~+
~\,. ~; ~4 I\, 0 Section
~~ ~
\ \ , \'
T ,\I
... x «
~8 roughness applied to standard -/,2 with I L , ~ I ,I I I !
'I
o
-/.6 .0401 OJ2 0201 .056, .044 .028 .024 .0/6 .012 .052 .048 .036 .008 .004 ~." \l> l>
~ a g> ~ \:)
~ i5 .;::
Jl
(modified) airfoil :72 -- 65(223)-'422 NACA (modified) an '""'~ !/e -422 /6 ~a,:; deg j~tt:: ~ 165(223 I I (;(0, ~ 258 I I in.
l/, I drag characteristics of NACA 36 tack, - TOT I I
IJ'
of and Iff of Iff I I Test: , Airfol'!: Chord: Iff 49.-Lift II onq/@ II -8 V FIGURE I/!
V Section 1',- \ 1\ -/6 ~ I" -24 I ) 1 ) I t l' t ¢ 9 :> 'i'
o
.8 .4 -32 /.8 /.2 -.8 3.6 -.4 32 2.8 2.4 2.0 -I.
-/.2 -2.
... §
~ ~ 8 =t ~
.... ~
42 REPOR'f NO. 824-NATIONAL ADVISORY COMMITTEE FOR AEHONAUTICS
I
Single flagged symbols are for 60° simulated split flap I
..
/ C - I-- Zi v Seri6'> 5 .2- V
~O} r-I
14 (4-diqit)_ f!. .4_ -..., I--!
<>
----
1-0- 'V .6 r----.
"""' --
,6
~~ I I I---
~" r--.,.
'V 230 (5-digit) -0 ~
r--
"'"""
:----,
--
;.-....
----
I--
-
(a) fb) .-:: -.';; , -.1; 4 8 12 16 20 24 28 4 b R ~ ~ ~ ZR , Airfoil thickness) p~rcent of chord Airfoil thickness) percenf of chorrl (a) N AUA four and flve·diglt serle~.
(b) NAUA f>3 selles.
.I la=O.5 me~n li~e., o ,
o
, , ~ J , C z· I-- 0'- I-- I eli r-- 0 ,....
-
0 .!
00,_ I-- - .2- <> .2 <>
f!. .4_ r---- f!. .4_
-l- I--
-
.6 >---.
'V 'V .6 r---. '---b.
~
~
t:; ~ - ..... r-- -l:l..
'-- ~ "V .-- ""- ...
r--
::-
.....: -
--;;--
r-
-tl -...
:-- (c) (d) ~ 4 8 12 16 20 24 28 4 b 12 /6 20 24 28 Air-foil thickness) percent of chord Airfoil thickness., percent of chord (c) NACA 04-series.
(d) NACA 65·selie~.
I
o
I C'i- I-- - .2- <) ; ~ -...., F=: f!.
.4 NA ........
-
~ I"'-
~ ............
- -
~ -Q "'-- .......
(e) 4 8 12 16 20 24 28 Airfoil thickness) percent of chord (e) NACA 66-series.
FIGURE OO_-Variatlon ofscction CIuarter-chord pitching-moment coefficient (measnred at an angle of attack of 0°) with airfoil thickness rntlofor se\-crnl N ACA airfoil sections of different ram bt'r R=6XI0 • SUMMARY OF AIRFOIL DATA -. I 6'
increasing thickness. For the N ACA 6-series airfoils, the
....
V
opposite appears to be the case.
~
V
.Q -.f 4 HIGH-LIFT DEVICES ~
1/
V Lift characteristics for two NACA 6-series airfoils equipped
.... -. I Z 65~-618/ ~
with plain flaps are presented in figure 53. These data
1/
6:)j-6 .s; a~O show that the maximum lift coefficient increases less rapidly
&
1/
'iii -. I with flap deflection for the more highly cambered section.
___ ·4412 ~
\ 1/
Lift characteristics of three NACA 6-series airfoils with split
65 -421, a~05\
V
-eJ -42/
flaps are presented in reference 44 and figure 54. The maxi-
~1--6~-41.s; a:05" '- .. ,,~ .6%;-4/5 .. ,,," 6 2-4/~ a=£5 .. \\
V mum-lift increments for the 12-percent-thick sections were
-6,r 4 /8
-63 4-420
only about three-fourths of that increment for the 16-percent-
V
'i
'6
thick section. The maximum lift coefficient for the thicker
6J. 4-42d a='o.s, 2I J
'66( f -,16
section with flap deflected is about the same as that obtained
i6: ..
Ir
'4
-ere
for the NACA 23012 airfoil in the now obsolete Langley
V
variable-density tunnel (reference 45) and in the Langley
~. "66(Z/5)-Z/6, a-06
1/
7- by lO-foot tunnel (reference 46).
~ -23012
Tests of a number of slotted flaps on N ACA 6-series
LL
airfoils (supplementary figures and reference 47) indicate that
l/
o -.02 -.04 -.06 -.08 -.10 -.12 -.14 -.16 the design parameters necessary to obtain high maximum Theoretical moment coefficienf for the airfoil
lifts are essentially similar to those for the the N ACA 230-
mean line about quarter-chora point
series sections (references 48 and 49). Lift data obtained
FIGURE 51.-Comparison of theoretiml and measured pitching-moment coefficients for some
for typical hinged single slotted 0.25c flaps (fig. 55 (a» on
NACA airfoils. R=6X10', the NACA 63,4-420 airfoil are presented in figure 55 (b).
POSITION OF AERODYNAMIC CENTER
A maximum lift coefficient of approximately 2.95 was ob-
tained for one of the flaps. Lift characteristics for the
The variation of chordwise position of the aerodynamic
NACA 65 -118 airfoil fitted with a double slotted flap
center corresponding to a Reynolds number of 6 X 10 for a 3
(reference 47 and fig. 56 (a» are presented in figure 56 (b).
large number of NACA airfoils is presented in figure 52.
A maximum lift coefficient of 3.28 was obtained. It may
From the data given in the supplementary figures there
be concluded that no special difficulties exist in obtaining
appears to be no systematic variation of chordwise position
high maximum lift coefficients with slotted flaps on moderately
of aerodynamic center with Reynolds number. The data
thick N ACA 6-series sections.
for the NACA 00- and 14-series airfoils, presented for thick-
Tests of airplanes in the Langley full-scale tunnel (reference
ness ratios less than 12 percent, show that the chordwise
42) have shown that expected increments of maximum lift
position of the aerodynamic center is at the quarter-chord
coefficient are obtained for split flaps (fig. 57) but not for
point and does not vary with airfoil thickness. For the
slotted flaps (fig. 58). This failure to obtain the expected
NACA 24-, 44-, and 230-series airfoils with thickness ratios
maximum-lift increments with slotted flaps may be attributed
ranging from 12 to 24 percent, the chord wise position of the
to inaccuracies of flap contour and location, roughness near
aerodynamic center is ahead of the quarter-chord point and
the flap leading edge, leakage, interference from flap sup-
moves forward with increase in thickness ratio.
ports, and deflection of flap and lip under load.
The chordwise position of the aerodynamic center is behind
the quarter-chord point for the NACA 6-series airfoils and
LATERAL-CONTROL DEVICES
moves rearward with increase in airfoil thickness, which is
in accordance with the trends indicated by perfect-fluid An adequate discussion of lateral-control devices is outside theory. There appears to be no systematic variation of the scope of this report. The following brief discussion is chordwise position of the aerodynamic center with camber or therefore limited to considerations of effects of airfoil shape position of minimum pressure on the basic thickness form for on aileron characteristics.
these airfoils. The effect of airfoil shape on aileron effectiveness may be
The data of reference 43 show important forward move- inferred from the data of figure 59 and reference 50. The ments of the aerodynamic center with increasing trailing-edge section aileron effectiveness parameter Aaol Ao is plotted
against the aileron-chord ratio Calc for a number of airfoils
angle for a given airfoil thickness. For the NACA 24-, 44-,
of different type in figure 59. Also shown in this figure
and 230-series airfoils (fig. 52) the effect of increasing
are the theoretical values of the parameter for thin airfoils.
trailing-edge angle is apparently greater than the effect of
44 REPORT NO. 824-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
,---r-,~-- .20 .26 I--'" ...-: -- l--- r--
-- S-
.---- ---- .26 t--- f'o-..
. ..,;,- ._-- . ----- -- eli =0.4 and 0.6 __ Nef=!.~ies .24 .22 Symbols wilh flo9S dorrespond fa qi = 0.1 ,-- --.
.28 .26 k
f-o-
i'- V t-- 0
tn-
"--- I--- .24 .26
--t--o---
V
r--... ,./"
l--- i-- I--- NACA 44-series ez;=0.2· .24 .22 -----,,"-- .28 1--- .26 +-- l,..-
0 f..o
....
v.
, __ V
r--
.24 .26 :;>- ~ J--.....
(b) NACA 24-series e'i =0.
.24 .22 ..
.26 .28 .......; '>--
~
~ 7>- l' l' .. - --- ~ .24 -- ~ .26 t S .
<lJ () () NACA 14-series C =0.4. and 0.6 li .~ .22 -- 1--'->- .~ .24 (5 (j ~.26 ~.28 '0 '0 () v ~
~ ~
I--
-
~ tJ .24 IJ .26 "'- "'- o o c: (a) NACA DO-series t c ;=0.2 l ~ .22 ~ .24 . iii 'iii 28 () .28
8.-
..., Q.
~
~
<lJ <lJ .'2 l- .'2
Va
I-- '-
V
-6.26 -6.26 v ~
~
,J::, (d) C" =0.4 and 0.6 c,,=O <.j .24 <.j .24 Symbols wilh flags correspond to eli = 06 .28 .28 I -0 I--
f-o-
,..-- ---' .26 .,........ .26 !r-' C =0.2 li Cli - 0.4 .24 .24 .28 .28 ../ .26 .26 V eli =0.1 C,i =0.2 .24 .24 .28 .2 8
. ..3
k-o
I--""'
V
~ ~ ~ .26 .26 r- ~ (c) (e) e<;=D Cli =0.
.2 4
8. 12 Ie 20. 24 28 o 8 12 16 20. 24 28 32
32 4 Airfoil fhiclrness, percent of chord Airfoil fhiclrness, percent of chord (a) N ACA four- and five-digit series.
(b) NACA 63-series.
(c) N ACA 64-series.
(d) NACA 65-series.
(e) NACA 56-series.
FIGURE 52.-Variation of section chordwise position of tne aerodynamic center with airfoil thickness ratio for several NACA airfoil sections of different cambers. R=6Xl()6.
SUMMARY OF AIRFOIL DATA 2.8 .P 2.4 66(215)-216- __ NACA
~ /'
Y
~
~
~2.0
rS V
WACA 65,3-618 ..... ~
~
c: .!!!
/'
~
A
~ 1.6 (j
/ V
it "" c:
.~
:2
/ /
<.J 1.2
t £2r---+---~--~---+--~----~--+---4----+--~
QJ VD-""" (\) If) \r)
§
§ (ap;'ox)I-----cf---'-l
.§ .§ 0 NACA 88(215)-216 15.0xl0 ~ -- NACA 23012 3.5 ~ .8 ~ ~ .6 f----f-----+_ (from ret: 45) (eff.) -j----f-----l o NACA 6i5,1-212 6.0 , <> NACA 65 -212 8.0 .4 .4~--+---1---~---+---+~--r---+---4----+--~
o
-20 20 40 60 80
o
o 20 40 60 100 Flop deflection, Ojj deg Flop deflection, Of; deg FWURE 53.-Maximum lift coefficients for the NACA 65,3-618 and NACA 66(215)-216 air· FWURE 54.-Maximum lift coefficients for some NACA airfoils fitted with O.20-airfoil·chord foils fitted with O.20-airfoil-chord plain flaps. R=6XlO'.
split flaps.
The data show no large consistent trends of aileron-effective- coefficient based on the wing chord. This method of analysis ness variation with airfoil section for a wide range of thick- takes into account the aileron effectiveness, the hinge ness distributions and thickness ratios. In order to evaluate moments, and the possible mechanical advantage between aileron characteristics from section data, a method of analysis the controls and the ailerons. The larger the value of Aao is necessary that will lead to results comparable to the usual for a given value of the hinge-moment parameter, the more curves of stick force against helix angle pb/2V for three- ad vantageous the combination should be for providing a dimensional data. The analysis that follows is considered large value of pb/2V for a given control force. The assump- suitable for comparing the relative merits of ailerons from tion that the aileron operates at a constant lift coefficient two-dimensional data.
as the airplane rolls is not entirely correct, however, and Two-dimensional data are presented in the form of the involves an overestimation of the effect of changing angle eq uivalent change in section angle of attack Aao required to of attack on the hinge-moment coefficient. In addition, maintain a constant section lift coefficient for various de- the span of the ailerons and other possible three-dimensional flections of the aileron from neutral. This equivalent change effects are not considered. In spite of these inaccuracies, in angle of attack is plotted against the hinge-moment param- the method provides a useful means of comparing the two- eter ACHO, which is the product of the aileron deflection from neutral and the resulting increment of hinge-moment dimensional characteristics of different ailerons.
I
11>0 >- ::l o >- t:"' ~ m g >1 o a:: a:: .... :4 tr.l tr.l "'l g >- tr.l :tl o ~ c:j ..., .... c m
:tl ~ g ..., Z !=> Qo l'-' z z >- c
~ ~ I
I
\
'" ,-
...
\
~
./ de9 -----
.---. ;/
I R=6Xl()6.
dfJ o 02
/
/
location Flap-hinge - V lift characteristics.
)1
F
deflection,
J
Maximum Flap (b)
I~
III V
~
V
a O.25-airfoil-chord hinged slotted flap.
with (b)
V
) I t I t o .8 .4 l.tS 1.2 3.
2.8 2.4 2.0 airfoil £ ~ \\) () ~
C t () () t:: \\) It) §
:~ ~ ~ .::: .~ "" .§ ~
....:' 63,4-420 NACA the ~Oo ""'200 "', , ......
~-t ......~', ./87
J'r'~~300
....."
, ~~, ......
....
'....
......
~, maximum lift coefficients for ~~_~ .......
"
--
and -- ~,-:--~ _- -=..-~-::.-;:,""',;:.J~:o::'""""'t__r,; -~--- - -
I
- configuration configuration.
,.
-~-----\;...:: $-.-.I76-----~ 1.000----------------~ -------.:::::::..,;:;: ~--.I63--1 "" Flap l -~-~ location (a) 55.-FJap Hinge FIGURE Hinge-location f (a) SUMMARY OF AIRFOIL DATA
~--=/~
Flap retracted
.1 I. I 1 ) I I I I
Lme of agreement between measured 1<----------:--.864------ /.0 and predicfed results ..
'.
V
.8
V
1<----------------,875-------------r--~ v:
/'
(a)
V
1/
~.
~ /,.--0- 3.2 .2
V
V
/
L
o .4 ,6 .8 1.2
.2 1.0
I
2.8 LlC'I' predicted FIGURE 57.-Comparison between measured values of the increments In lift coefficients due
V
t.o flap deflection and values predicted from two-dimensional data. Split flap.
~(
2.4
/
/
I . 1 1 I 1 1 1 1 I
L me of ogree/nenf between meosured
/
1.0
t--- and predictd results...... L
/
i !
V
/
.8 ,/
V
V
/
J
'/
-q .4
V
,8
V
.2
V
/
.4 1.2 .2 .4 .6 .8 1.0
o
LlC,,} predicted FIGURE 5S.-Comparison between measured values of the increments In lift coefficients due to flap dellection and values predicted from two-dimensional data. Slotted lIap.
(b)
o 20 40 60 80
Flop deflection, oJ) deg (a) Flap configuration.
(b) Maximum lift characteristics.
:FIGURE 56.-Flap configuration and maximum lift coefficients for the NACA 653-118 airfoil with a double slotted flap. R=6X10 , REPORT NO. 824-NATIONAL ADVISORY COMMIT'l'EE FOR AERONAUTICS SUPPLEMENTARY INFORMATION REGARDING TESTS OF TWO-DIMENSIONAL MODELS Air-flow characteristics
I
Symbol Basic airfoil Type of flap ---------------- Reference
TIM I R
------;--- NACA OO~~~=~~~~=~~~~~~~~~~= --;,lain __ ~==~~=~~~~~~~ ----~;_- --~~-- -=~~~~~-r~~~;- NACA 0015 _________________________________________________ do ______________________________ _ 1. 93 .10 1.4XI06 56
+
NACA 23012, __________ . ____ . ___ . ______ . ____ ..• ___ .. __ . ___ .. do ___ ... ___ ... _._, .. _ .. ____ .. __ ..
1. 60 .11 X 2.2XIO' 57,48
o NACA 66(2xI5)-OlU _ Plain, straight contour ___ ... ____ ... _ 1. 93
.10 1.4XI0' NACA 66-009._ ... ____ ... ___ "' __________ .. _____ .. ___ "'_ Plain __ ..• _._ .... _ .... _____ .... __ ..• _ 1. 93 .11 l.4XlO' 58
<>
NACA 63,4-4(17.8) (approx.) ______ . _______ . __ ._ . _____ . Internally balanced __ .... ___ ._. __ ._.
A (I) .17 2.5X106 59 \1 NACA 66(2xI5)-216, a=0.6. _________ . _____ .. ___ ... ___ .. _ .. ,.do .... __ ..... ___ .... ___ .. ______ ._ (I) .18 5.3XlO' 59 NACA 66(2xI5)-116, a=O.6 .. __________ . ____ '. __________ .. ___ do _______________________________ _ C> (I) .14 59 6.0XJO' NACA 64,2-(1.4) (13.5) _____ .. ___________ . ____ . __ ______ _ Plain_. ______ • ____ .. ______ . _____ .. __ ._ <l <I) 13.0XIO' NACA 65,2-318 (approx.) ________ ._ .... ________ ... __ .. _ Internally balanced ____________ .. __ _ !7 (1) .14 6.0XI0' 59 NACA 63(420)-521 (approx.) __ .. ______ .. _______________ .. __ do .. ___ .... _______________ . _____ _ 'q (I) 8.0X1Q6 .20 2.8XI0 NACA 66(215)-216, a=O.6 __ .. ,_ .. ______ . ___ . ___ .. ____ . _' ___ do._ .. __________ .. ______________ _ c,.. <I) to to 60
}
.48 6.8X106 NACA 66(215)-014.. ____ 00 ___ " ____ .. ____ • ___ • _____ • ,___ Plain ______ .. ________________ • ______ _ 1.93 .d .09 1.2XI0 61 NACA 66(215)-216, a=O.6 ____ ._ ._. ___ .. _. _______ . ____ . _______ do _______________________ • ______ _ <I) 6.0X1Q6
o
NACA 65,-415 _______ .. ________ . ___ . ______ . ________________ .do_. ____ • ________________ • ______ _ <I) .13 6.0X106 62
o
NACA 653-418 ___ • ____ . _______ .00 ____ • _____ • ______ • _________ do ______________________________ _ [} (I) .13 6.0XI0 62 NACA 65,-421. ___ . ___ ._. __ . ___________ .. ___ . ______ . ________ do ______________________________ _ Q (I) .13 6.0X106 62 NACA 65(112)-213 .. ________ . _____ . _____ ._. __ .. ______ ._. __ Internally balanced. _______ ._. _____ _ (I) .14 8.0XJ06
o --------
NACA 745.A317 (approx.). _____________ . ___ .. ___ ... _. ____ • ___ do ___ ..... _ ... __ .... __ .. _._ ... __ . (I) .13 6.0XlO'
o
NACA 64,3-013 (approx.)_. _______ .. ____ ... ____________ . ____ do __________ . ___________________ _ <I) .13 6.0X!06
o
NACA 64,3-1(15.5) (approx.). ____ . ____ . __ . __________________ do ______________________________ _ (1) .13 6.0XlO' l:> I Approaching 1.00.
.8 .8 ., ....
..
" .. ..
/' , ~ ..
" /'" ./ ..
.. ..
/'"
V
.. Thgoretical- ..
Theoretical
-
--- "./
v -
-
tJl ." ~.6
..
~1~·6 ..-
,
, ".. V
LV<
.; ~ .- I/) ..- <I) V
,,' <;.<
III QJ
s: /'
, t , c: ,
, - QJ 0
QJ --Experimenfal ,,- V ,
/[7
:,
~ 17
" " , :;:; , ....
-
,
(j --
, --E~perimental
IJ ,
/9-
J.P
.4 .4 ~ ~ , 'V -,.; .- , "- , QJ
QJ , iP
/
, t c: , I ,
V ~
I
t
~
::;:
c;,£o
,
./
(J () ,
""
I t I
c: !oP"x I
/
.~ :g .2
.2 .....
+ <J
/
J I V"
~
J?
I
,'V /
I
!sf
,
//
V (b)
(a)
V
o .1 .2 .3
o .1 .2 .3 .4 .4 Aileron chord ratio, calc A ileron chord ratio, calc (b) IJ range from 0° to 20°.
(a) IJ range from 0° to 10°.
FIGURE 59.-Variation of section aileron effectiveness with aileroll-chord ratio for true·airfoil·contour ailerons without exposed overhang balance on a number of airfoil sections.
Gaps sealed; c,=O.
SUMMARY OF' AIRFOIL DATA - I Air-flow characteristics Refer- Basic airfoil CI Type of aileron Refer- ence Basic airfoil CI 'l'ype of aileron --- ---- ----- ence (I) ---- ----
------------- ----------
f M R NACA 66(215)-216.a =0.6 0.20c plain 0.100 64 ------------- --- ----- .-- --- ----- --- NACA 63,4-4(17.8)(ap- .450 0.20c with 0.43c, inter- ---- NACA 0009 0 0.20c plain 1. 93 0.10 1.4XI06 63 prox.) nal balance NACA 64,2-(1.4) (I3.5) .150 0.187c plain
I (') .18 4.0X106
---- NACA~66(215)-216, a=0.6_ .100 0.20c plain (~) .33 9.0X106 64
1 I
1 Trne airfoil contoUl".
~ Approaching 1.00. Aileron deflecfion, LI6/~ deg NACA 4~4( 11.8) (opprox.j-.,-: 63,
f\
if;";'- -/2- .. _.r ... _
1\
.... ---
...:., .. .:
C
" Ef' " ---)
~ '\
, ~:-- NACA 66(215)-216, ~-.(.\ a=Q6-"" '.
'~ , ~ Stroi hi sided ;biJ, ~'"'' -2 "
V 471
"
---
p;;...- '-a,
~ li"
"
- -4
--
"12 .--- ~.:::- "':.::::~ ...
I ~ ( -' ". ' '.
.
~;""'....-- ;~~~
V
---
" -6 -4r-r-+-+-~~-r-r_+_+~I-4~~2-~~,~,~+-+-~ -_:.
True airfoil contou?)- rAileron defleelion, deq L---rr ", -8
-.0018 -.0016 ~0014 -.OOlc ,DOlO ~0008 :0006 :0004 :0002 o
.1c o, radians
u FIGURE 61.-Variation of the hinge-moment parameter ACH5 with the equivalent change in section angle of attack required to maintain a constant section !itt coefficient for deflection of true-airfoil-contour and straight-sided ailerons on the NACA 63,4-4(17.8) (approx.)
and the NAC A 66(215)-216, a=O.6 airfoil sections. Gaps sealed.
FIGURE 60.-Variation of the hinge-moment parameter ilea a with the equivalent chan~e in section angle of attack required to maintain a constant section lift coefficient for deflection of the aileron on the N ACA 0009, N ACA 64,2-(1.4)(13.5), and N ACA (\6(215)-21(\, a=O.6 airfoil sections. Gaps sealed. able. It appears, however, that the straight-sided aileron would be lE'sS advantageous than the aileron of true contour for positive dE'flections greater than 12°. In the case of For the purpose of evaluating the effect of airfoil shape on the NACA 63,4-4(17.8) (approx.) airfoil, the straight- the aileron characteristics, it is desirable to make the com- sidE'd aileron appears to have no advantage over the aileron parison with unbalanced ailerons to avoid confusion. Plots of true airfoil contour. The advantage of using straight- of the parameters for plain unbalanced flaps of true airfoil sided ailerons appears to depend markedly on the airfoil used contour on threE' airfoil sections are shown in figure 60.
but sufficient data are not available to determine the signif- The characteristics of the NACA 66(215)-216, a=0.6 section icant airfoil parameters. Figure 62 shows that in one case are essentially the same as those for the NACA 0009 airfoil the effE'ct of leading-edge roughness on the aileron character- within tIll' range of deflection for which data are available.
istics is unfavorable.
The NACA 64,2-(1.4) (13.5) airfoil shows appreciably smaller valuE'S of 6.CIIO for a given value of 6.0:0 than the other LEADING-EDGE AIR INTAKES sections prE'sented. No explanation for this difference can The problem of designing satisfactory leading-edge air be offered, although some of the difference may result from intakes is to maintain the lift, drag, and critical-speed the slightly smaller chord of the flap for this combination.
characteristics of the sections while providing low intake The effects of using straight-sided ailE'rons instE'ad of ailer- losses over a wide range of lift coefficients and intake velocity ons of true airfoil contour are shown in figure 61 for two ratios. The data of reference 65 show that desirable intake N ACA 6-series airfoils. One of the two combinations for which data are available was provided with an internal and drag characteristics can easily be maintained over a rather small range of lift coefficients for NACA 6-series air- balance whE'reas the other combination was without balance.
foils. The data of reference 65 show that the intake losses This differencE' prevents any comparison between the two increase rapidly at moderately high lift coefficients for the combinations but doE'S not affect comparison of the two shapes tested. Unpublished data taken at the Langley contours for each case. For the NACA 66(215)-216, a=O.6 Laboratory indicate that shapes such as those of reference airfoil, the straight-sided aileron has more desirable charac- 65 have low maximum lift coefficients. Recent data show teristics for tll(' range of deflections for which dat.a arE' avail- REPORT NO. 824-NATIONAL ADYlSORY COMMITTEE FOR AERONAUTICS that air-intake shapes can be provided for such airfoil sec- t:-- tions with desirable air-intake characteristics and without -16·.
- .... i-- !<§mooth
loss in maximum lift coefficient (fig. 63). Some pressure- ~~~ Roughness ot leading edqe.--> --.::.:,r :::--.
distribution data for the air intakes shown in figure 63 in- -I-
::::-
dicate that the critical speed of the section has been lowered ~ only slightly and that falling pressures in the direction of ~ fX'-8 flow were maintained for some distance from the leading .~ edge on both surfaces at lift coefficients near the design lift -4-~··
1\
Q\ coefficient for the section. Sufficient information is not ~ _ 0 available to permit such desirable configurations to be de- 'I!
signed without experimental development.
"'I J"::
V
-2 INTERFERENCE
~V
The main problem of interference at low Mach numbers)s >.5·"8
/'/
-4 considered to be that of avoiding boundary-layer separation -~ ,/
-- ,/
resulting from rapid flow expansions caused by the addition
r- I-- Aileron deflection, deg -12"
<'-,.
of induced velocities about bodies and the boundary-layer -6
----
accumulations near intersections. No recent systematic investigations of interference such as the investigation of -8 ~OOI8 ~OOI6 ~OOI4 ~OOI2 ~OOIO ~0008 ~0006 ~0004 ~OOO2 0 reference 66 have been made.
twHo, radians Some tests have been made of airfoil sections with in- FIGURE 62.-Variation of the hinge·moment parameter !1cH6 with the equivalent change in tersecting flat plates (reference 67). These configurations section angle of attack required to maintain a constant section lift coefficient for deflection may be considered to represent approximately the condition of the aileron on the NACA 64,2-(1.4)(13.5) airfoil section, smooth and with roughness at the leading edge of the airfoil. (For description of aileron, see fig. 60.)
of a wing intersection with a large flat-sided fuselage.:.i:1n 24-inch chord 1.6 1.4 # ~.
", v.
I J .p- ~
~
,!-e =0.186 ..• . V' he" 1.135
p
~ ~ 1.2 /.2
V
,
lei 'l ./
·····h ~ 1./35
V
e
t
.8 ..,
V V
J
I
IJ
\
\ tJH ,-, he =0.186
,a
qo
L
Ii /
1\ v.
(V' he =0.186
/
j
\
u Configuration R ...r/'
V
\
o Plain airfoil - 3.0" I O' t--- ~ o Ducfed model
V
I (low flow) .,. r
2.4 11
.2 <> Oucted model
/
r:I
I (high flpw) 2.4
If . tJ H, he = /.135
qOI I
;\
X J
-I.e
o
-16 -8 0 8 16 24 ~4 o .4 .8 !.2 /.0
Section onqle of attock, «" deq Sedion lift coerric~'enf, C FIGURE 63.-Lift and flow characteristics of an )<ACA 7·scries type airfoil section with leading·edge air intake.
SUMMARY OF AIRFOIL DATA
this case, the interference may be considered to result from
The usual wing theory assumes that the resultant air force
the effect on the wing of the fully developed turbulent bound- and moment on any wing section are functions of only the ary layer on the fuselage or flat plate and the accumulation of section lift coefficient (or angle of attack) and the section boundary layer in the intersection. These tests showed shape. According to this assumption, the air forces and little interference except in cases for which the boundary moments on any section are not affected by adjacent sections
layer on the airfoil alone was approaching conditions of
or other features of the wing except as such sections or
separation such as were noted with the less conservative
features affect the lift distribution and thus the local lift of
airfoils at moderately high lift coefficients.
the section under consideration. These assumptions ob-
Some scattered data on the characteristics of nacelles
viously are not valid near wing tips, near discontinuities in
mounted on airfoils permitting extensive laminar fl.ow are
deflected flaps or ailerons, near disturbing bodies, or for
presented in references 68 to 70. The data appear to in- wings with pronounced sweep or sudden changes in plan
dicate that the interference problems for conservative NACA
form, section, or twist. Under such circumstances, cross flows
6-series sections are similar to those encountered with other result in a breakdown of the concept of two-dimensional
types of airfoil. The detail shapes 'for optimum interfering
flow over the airfoil sections. In addition to th('se
bodies and fillets may, however, be different for various
cross flows, induced effects exist that are equivalent to a
sections if local excessive expansions in the flow are to be
change in camber. Such effects are particularly marked
avoided. near the wing tips for wings of normal plan form and for
Some lift and drag data for an airfoil with pusher-propeller- wings of low aspect ratio or unusual plan form. Lifting- shaft housings are presented in reference 71. These results surface theory (see, for example, reference 81) provides a indicate that protuberances near the trailing edge of wings means for· calculating wing characteristics more accurately should be carefully designed to avoid unnecessary drag than the simple lifting-line theory.
increments. Although span load distributions calculated for wings with
Another type of interference of particular importance for discontinuities such as are found with partial-span flaps high-speed airplanes results in the reduction of the critical (references 82 and 83) may be sufficiently accurate for Mach number of the combination because of the addition of structural design, such distributions are not suitable for the induced velocities associated with each body (reference predicting maximum-lift and stalling characteristics. Until 72). This effect may be kept to a minimum by the use of sufficient data are obtained to permit the prediction of the
maximum-lift and stalling characteristics of wings with
bodies with low induced velocities, by separation of inter-
discontinuities, these characteristics may best be estimated
fering bodies to the greatest possible extent, and by such
from previous results with similar wings or, in the case of
selection and arrangement of combinations that the points
of maximum induced velocity for each body do not coincide. unusual configurations, should be obtained by test.
The characteristics of intermediate wing sections must be
known for the application of wing theory, but data for such
APPLICATION TO WING DESIGN
sections are seldom available. Tests of a number of such
Detail consideration of the various factors affecting wing
intermediate sections obtained by several manufacturers for
design lies outside the scope of this report. The following
wings formed by straight-line fairing have indicated that the
discussion is therefore limited to some important aerodyna-
characteristics of such sections may be obtained with reason-
mic features that must be considered in the application of
able accuracy by interpolation of the root and tip character-
the data presented.
istics according to the thickness variation.
APPLICATION OF SECTION DATA SELECTION- OF ROOT SECTION Wing characteristics are usually predicted from airfoil- The characteristics of a wing are affected to a large extent section data by use of methods based on simple lifting-line by the root section. In the case of tapered wings formed by
straight-line fairing, the resulting nonlinear variation of sec-
theory (references 73 to 76). Application of such methods
tion along the span causes the shapes of the sections to be
to wings of conventional plan form without spanwise discon-
predominantly affected by the root section over a large part
tinuities yields results of reasonable engineering accuracy
of the wing area. The desirability of having a thick wing
(reference 77), especially with regard to such important
that provides space for housing fuel and equipment and re-
characteristics as the angle of zero lift, the lift-curve slope,
duces structural weight or permits large spans usually leads
the pitching moment, and the drag. Basically similar
to the selection of the thickest root section that is aerody-
methods not requiring the assumption of linear section lift
namically feasible. The comparatively small variation of
.characteristics (references 78 and 79) appear capable of
minimum drag coefficient with thickness ratio for smooth
yielding results of greater accuracy, especially at high lift
airfoils in the normal range of thickness ratios and the main-
coefficients. Further refinement may be made by consider-
tenance of high lift coefficient for thick sections with flaps
ation of the chordwise distribution of lift (reference 80).
deflected usually result in limitation of thickness ratio by
Wings with large amounts of sweep require special consider-
characteristics other than maximum lift and minimum drag.
ation (reference 81).
HEPORT NO. 824-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS The critical Mach number of the section is the most serious The selection of the optimum type of camber for the tip limitation of thickness ratio for high-speed airplanes. It is section presents problems for which no categorical answers can be given on the basis of existing data. The use of a type desirable to select a root section with a critical Mach number sufficiently high to avoid serious drag increases resulting from of camber that imposes heavy loads on the ailerons compli- compressibility effects at the highest level-flight speed of the cates the design of the lateral-control system and increases airplane, allowance being made for the increased velocity of its weight. The use of a type of camber that carries the lift flow over the wing resulting from interference of bodies and farther forward on the section and thus relieyes the ailerons slipstream. Available data indicate that a small margin will, however, have little effect on the maximum lift coeffi- exists between the critical Mach number and the Mach num- cient of the section unless the maximum-camber position is ber at which the drag increases sharply. As airplane speeds well forward, as for the N ACA 230-series sections. In this increase, it becomes increasingly difficult and finally impos- case a sudden loss of lift at the stall may be expected. The sible to avoid the drag increases resulting from compressibil- effects on the camber of modifications to the airfoil contour ity effects by reduction of the airfoil thickness ratio. near the trailing edge, which may be made in designing the In the cases of airplanes of such low speeds that compressi- ailerons, should not be overlooked in estimating the charac- bility considerations do not limit the thickness ratio to values teristics of the wing.
less than about 0.20, the maximum thickness ratio is limited If the root sections are at least moderately thick, it is by excessive drag coefficients at moderate and high lift usually desirable to select a tip section with a somewhat coefficients with the surfaces rough. In these cases, tho reduced thickness ratio. This reduction in thickness ratio, actual surface conditions expected for the airplane should be together with the absence of induced velocities from inter- considered in selecting the section. Consideration should fering bodies, gives a margin in critical speed that permits the camber of the tip section to be increased. This reduction in also be given to unusual conditions such as ice, mud, and thickness ratio will probably be limited by the loss in maxi- damage caused in military combat, especially in the case of milm lift coefficient resulting from too thin a section.
multiengine airplanes for which ability to fly under such A small amount of aerodynamic washout may also be conditions is desired with one or more engines inoperative.
In cases for which root sections having large thickness ratios useful as an aid in the avoidance of tip stalling. The per- are under consideration to permit the use of high aspect ratios, missible amount of washout may not be limited by the in- a realistic appraisal of the drag coefficients of such sections crease in induced drag, which is small for 1 or 2° of washout with the expected surface conditions at moderately high lift (reference 73). The limiting washout may be that which coefficients will indicate an optimum aspect ratio beyond causes the tip section to operate outside the low-drag range which corresponding increases in aspect ratio and root thick- at the high-speed lift coefficient. This limitation may be ness ratio will result in reduced performance. so severe as to require some adjustment of the camber to Inboard sections of wings on conventional airplanes are permit the use of any washout.
subject to interference effects and may be in the propeller A change in airfoil section between the root and tip may be desirable to obtain favorable stalling characteristics or slipstream. The wing surfaces are likely to be roughened by to take advantage of the greater extent of laminar flow that access doors, landing-gear retraction wells, and armament installations. Attainment of extensive laminar flows is, may be possible on the outboal"d sections. Thus, such com- therefore, less likely on the inboard wing panels than on the binations as an NACA 230-series root section with an NACA outboard panels. Unless such effects are minimized, little 44-series tip section or an N ACA 63-series root section with drag reduction is to be expected from the use of sections an N ACA 65-series tip section may be desirable.
permitting extensive laminar flow. Under these conditions, It should be noted that the tip sections may easily be so the use of sections such as the N ACA 63-series will provide heavily loaded by the use of an unfavorable plan form as to advantages if the sections are thick, because such sections are cause tip stalling with any reasonable choice of section and washout. Both high taper ratios and large amounts of more conservative than those permitting more extensive laminar flow. sweepback are unfavorable in this respect and are particu- SELECTION OF TIP SECTION larly bad when used together, because the resulting tip stall promotes longitudinal instability at the stall in addition to In order to promote desirable stalling characteristics, the the usual lateral instability.
tip section should have a high maximum lift coefficient and
a large range of angle of attack between zero and maxi- Inurn lift as compared with the root section. It is also CONCLUSIONS desirable that the tip section stall without a large sudden loss in lift. The attainment of a high maximum lift coefficient is The following conclusions may be drawn from the data often more difficult at the tip section than at the root section presented. :a.:fost of the data, particularly for the lift, drag, and pitching-moment characteristics, were obtained at tor tapered wings because of the lower Reynolds number of Reynolds numbers from 3 to 9XI0 • fhe tip section. For wings with small camber, the most 1. Airfoil sections permitting extensive laminar flow, such effective way of increasing the section maximum lift coeffi- as the NACA 6- and 7-series sections, result in substantial cient is to increase the camber. The amount of camber used reductions in drag at high-speed and cruising lift coefficients will be limited in most cases by either the critical-speed . as compared with other sections if, and only if, the wing requirements or by the requirement that the section have low drag at the high-speed lift coefficient. surfaces are fair and smooth.
SUMMARY OF AIRFOIL DATA 2. Experience with full-size wings has shown that extensive 9. The effect of leading-edge roughness is to decrease the lnminnr flows nre obtainable if the surface finish is as smooth lift-curve slope, particularly for the thicker sections having ns thnt provided by sanding in the chordwise direction with the position of minimum pressure far back.
No. 320 curborundmn paper and if the surface is free from 10. Characteristics of airfoil sections with the expected smull scattered defects and specks. Satisfactory results surface conditions must be known or estimated to provide a are usually obtained if the surface is sufficiently fair to permit satisfactory basis for the prediction of the characteristics of n straightedge to be rocked smoothly in the chordwise direc- practical-construction wings and the selection of airfoils tion without jarring or clicking.
for such wings.
3. For ,vings of moderate thickness ratios with surface 11. The N A CA 6 series airfoils provide higher critical conditions corresponding to those obtained with current Mach numbers for high-speed and cruising lift coefficients construction methods, minimum drag coefficients of the than earlier types of sections and have a. reasonable range order of 0.0080 may be expected. The values of the mini- of lift coefficients within which high critical Mach numbers mum drag coefficient for such wings depend primarily on may be obtained.
the surface condition rather than on the airfoil section. 12. The NACA 6-series sections provide lower predicted 4. Substantial reductions in drag coefficient at high critical ::\-fach numbers at moderately high lift coefficients Reynolds numbers may be obtained by smoothing the than the earlier types of sections. The limited data avail- wing surfaces, even if extensive laminar flow is not obtained. able suggest, however, that the NACA 6-series sections retain 5. The maximum lift coefficients for moderately cambered satisfactory lift characteristics up to higher Mach numbers smooth NACA 6-series airfoils with the uniform-load type than the earlier sections.
of mean line are as high as those for NACA 24- and 44-series 13. The NACA 6-series airfoils do not appear to present airfoils. The NACA 230-series airfoils have somewhat unusual problems with regard to the application of ailerons.
higher maximum lift coefficients for thickness ratios less 14. Problems associated with the avoidance of boundary- than 0.20. layer separation caused by interference are expected to be 6. The maximum lift coefficients of airfoils with flaps are similar for conservative NACA 6-series sections and other good airfoils. Detail shapes for optimum interfering bodies about the same for moderately thick NACA 6-seriessections as for the NACA 23012 section but appear to be considerably and fillets may be different for various sections if local exces- lower for thinner N ACA 6-series sections. sive expansions in the flow are to be avoided.
7. The lift-curve slopes for smooth N ACA 6-series airfoils 15. Satisfactory leading-edge air intakes may be provided are slightly higher than for N ACA 24-, 44-, and 230-series for NACA 6-series sections, but insufficient information exists airfoils and usually exceed the theoretical value for thin to allow such intakes to be designed without experimental airfoils. development.
8. Leading-edge roughness causes large reductions in maximum lift coefficient for both plain airfoils and airfoils equipped with split flaps deflected 60°. The decrement in LANGLEY ::\-IEMORIAL AERONAUTICAL LABORATORY, maximum lift coefficient resulting from standard roughness NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS, is essentially the same for the plain airfoils as for the airfoils LANGLEY FIELD, VA., March 5,1945.
equipped with the 60° split flaps.
APPENDIX
METHODS OF OBTAINING DATA IN THE LANGLEY TWO·DIMENSIONAL LOW.TURBULENCE TUNNELS By MILTON M. KLEIN DESCRIPTION OF TUNNELS
C/
section lift coefficient uncorrected for tunnel-
The Langley two-dimensional low-turbulence tunnels are wall effects
design lift coefficient
closed-throat wind tunnels having rectangular test sections
lift coefficient measured in tunnel
3 feet wide and 7% feet high and are designed to test models
completely spanning the width of the tunnel in two- moment coefficient about quarter-chord
dimensional flow. The low-turbulence level of these tunnels, point corrected for tunnel-wall effects amounting to only a few hundredths of 1 percent, is achieved moment coefficient about quarter-chord
point measured in tunnel
by the large contraction ratio in the entrance cone (approx.
F average of velocity readings of orifices on
20: 1) and by the introduction of a number of fine-
wire small-mesh turbulence-reducing screens in the widest floor and ceiling used to measure blocking
at high lifts
part of the entrance cone. The chord of models tested in
average value of F in low-lift range
these tunnels is usually about 2 feet, although the characteris-
potential function used to obtain 'I)-factor
tics at low lift coefficients of models having chords as large
total pressure in front of airfoil
as 8 feet may be determined.
The Langley two-dimensional low-turbulence tunnel oper- total pressure in wake of airfoil
ates at atmospheric pressure and has a maximum speed of coefficient of loss of total pressure III the
approximately 155 miles per hour. The Langley two-
wake (Ho~HI)
dimensional low-turbulence pressure tunnel operates at pres-
He maximum value of He
sures up to 10 atmospheres absolute and has a maximum
max
hT tunnel height
speed of approximately 300 miles per hour at atmospheric
pressure. Standard airfoil tests in this tunnel a,re made of
K=c/
Cd
2-foot-chord wooden models up to Reynolds numbers of
T
approximately 9X 10 at a pressure of 4 atmospheres absolute. L
true lift resulting from a point vortex
The lift and drag characteristics of airfoils tested in these L'
lift associated with a point vortex as
tunnels are usually measured by methods other than the use measured by integrating manometers of balances. The lift is evaluated from measurements of the m upstream limit of integration of floor and
pressure reactions on the floor and ceiling of the tunnel. The ceiling pressures
drag is obtained from measurements of static and t9tal n downstream limit of integration of floor
pressures in the wake. Moments titre usually measured by a and ceiling pressures
balance. resultant pressure coefficient; difference
,SYMBOLS
between local upper- and lower-surface
pressure coefficients
AI, A , •• An coefficients of potential function for a
symmetrical body static pressure in the wake
a fraction of chord from leading edge over free-stream dynamic pressure
which design load is uniform
static-pressure coefficient (HoqO P)
B dimensionless constant determining width
static-pressure coefficient in the wake
of wake
C chord
(Ho qo PI)
Cd drag coefficient corrected for tunnel-wall
8 distance along airfoil surface
effects
U velocity, due to row of vortices, at any
ca' drag coefficient uncorrected for tunnel-wall
point along tunnel walls
effects
v free-stream velocity
drag coefficient measured in tunnel
~V increment in free-stream velocity due to
section lift coefficient corrected for tunnel-
'blocking
wall effpcts
SUMMARY OF AIRFOIL DATA
V'
corrected indicated tunnel velocity
The factor 'T]x was obtained as follows: The image system
V"
tunnel velocity measured by static-pressure
which gives only a tangential component of velocity along the
orifices
tunnel walls is made up of an infinite vertical row of vortices
v
local velocity at any point on airfoil surface
of alternating sign as shown in figure 64. If the sign of the
w
potential function for flow past a symmetri-
vortex at the origin is assumed to be positive, the complex
cal body
potential functionj for this image system is .
distance along chord' or center line of
tunnel
ir I . h 'I1'zir I . h (Z-ihT)
j
(18)
=2'11' og sm 2hT -2'11' og sm '11' ~
y
variable of integration (B:w)
where
distance perpendicular to stream direction
Y
r strength of a single vortex
ordinate of symmetrical thickness distri-
ilt
z complex variable (x+iy)
bution
distance perpendicular to stream direction
Yw hT tunnel height
from position of Hcmax
y
slope of surface of symmetrical thickness
.+ +"+
distribution
z complex variable (x+iy)
angle of zero lift
section angle of attack corrected for tunnel-
wall effects. Upperwo// ....
m n
section angle of attack measured in tunnel
strength of a single vortex
ratio of measured lift to actual lift for any
Lower woll·· ....
type of lift distribution
.-
'IJ-factor for additional-type loading
1/a
'IJ-factor for basic mean-line loading
1/b
'IJ-factor applying to a point vortex
1/x
A component of blocking factor dependent on
shape of body
quantity used for correcting effect of body
FIGURE 64.-Image system for calculation of ~·ractor in the Langley two·dimensional
upon velocity measured by static-pressure
low·turbulence tunnels.
orifices
The velocity u, due to the row of vortices, at any point
component of blocking factor dependent on
along the tunnel walls where
size of body
potential function
hT
y=""2
stream function
is then obtained as
MEASUREMENT OF LIFT
r 'I1'X
(19)
u=2h sech hT
T
The lift carried by the airfoil induces an equal and opposite
reaction upon the floor and ceiling of the tunnel. The lift
where x is the horizontal distance from the point on the wall
may therefore be obtained by integrating the pressure dis-
to the origin. The resultant pressure coefficient P is then
R
tribution along the floor and ceiling of the tunnel, the inte-
given by
gration being accomplished with an integrating manometer.
4u
Because the pressure field theoretically extends to infinity in
PR==V
both the upstream and the downstream directions, not all the
lift is included in the length over which the integration is
2r 'I1'X
(20)
performed. It is therefore necessary to apply a correction =hTVsech hT
factor'T] that gives the ratio of the measured lift to the actual
where V is the free-stream velocity.
lift for any lift distribution. The calculation was performed
The lift manometers integrate the pressure distribution
by first finding the correction factor 'lJx applying to a point
vortex and then determining the weighted average of this along the floor and ceiling from the downstream position n factor over the chord of the model. to the upstream position m (fig. 64). For a point vortex REPORT NO. 824-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS located a distance x from the origin along the center line of The values of 'YJb and 'YJa for the Langley two-dimensional the tunnel, the limits of integration become n-x and m-x.
low-turbulence pressure tunnel are given in the following
The lift L' associated with a point vortex, as measured by
table for a model having a chord length of 2 feet, where 'YJb is the integrating manometers, is given by the 1]-factor corresponding to the basic mean-line loading (indicated by the value of a) and 'YJa is the 'YJ-factor for the additional type of loading as given by thin-airfoil theory: (21) a where qo is the free-stream dynamic pressure.
1.0 0.934; The true lift L resulting from the point vortex is given by .8 .9342 .6 .9336 .4 .9330 .2 .9325 o o .9322
L=2Q r
V
'Ia=O.92U6 The correction factor 'YJx is then In order to check the variation of 'YJa with variations in the additional type of lift distribution, the value of 'YJa was re- calculated for the class C additional lift distribution given in figure 6 of reference 74. The value of 'YJa for this case was 0.9304, as compared with 0.9296 for a thin airfoil. Because
1 in-x 7rX
= h- sech -h dx
of the small variation of 'YJa with the type of additional lift, T m-x T' the value for thin-airfoil additional lift was used for all cal- culations. The lift coefficient of the model in the tunnel which yields
uncorrected for blocking c/ is given in terms of the lift co-
efficient measured in the tunnel CIT and the design lift coeffi- (22) cient of the airfoil Cit by the following expression: (24) In the Langley two-dimensional low-turbulence tunnels, the orifices in the floor and ceiling of the tunnel used to measure the lift extend over a length of approximately 13 Because 'YJb does not differ much from 'YJa, it is not necessary feet. A plot of 1]x against x for the Langley two-dimensional that the basic loading or the design lift coefficient be known low-turbulence pressure tunnel is shown in figure 65. The with great accuracy.
1]-factor for a given lift distribution is obtained from the Because of tunnel-wall and other effects, the lift distribu- expression tion over the airfoil in the turinel does not agree exactly with the assumed lift distribution. Because of the small varia-
r 11'YJ" d (~) tions of 'YJ with lift distribution, errors caused by this effect are
J chord C
considered negligible. It can also be shown that. errors caused 'YJ=f (X)
l1d -
by neglecting the effect of airfoil thickness on the distri- chord C bution of the lift reaction along the tunnel walls are small.
1.0 MEASUREMENT OF DRAG ~r- i---...
--- The drag of an airfoil may be obtained from observations
t-- .8 of the pressures in the wake (reference 84). An approxi- mation to the drag is given by the loss in total pressure of the .6 air in the wake of the airfoil. The loss of total pressure is measured by a rake of total-pressure tubes in the wake.
When the total pressures in front of the airfoil and in the .4
wake are represented by Ho and HI, respectively, the drag
coefficient obtained from loss of total pressure Cd is T dyw
Cd = r Hc (25)
T Jwake C
o -2 ~ 0 I 2 3 4 5 6
where Disfonce downsfream from reference point in funnel, x) ft FIGURE 65.-Lift efficiency factor n. for a point vortex situated at various"positions along the
He coefficient of loss of total pressure in the wake (Ho-_HI)
center line of the tunnel. Qo SFMMARY OF AIRFOIL DATA Yw d istanc(' perpendicular to stream direetion froIll position -:-..
t---,
:-
S, of fI ;: ;---- ema t---- t---- 1./
r-I--
t-- --;--..
-t--
r--
r----
t----
r--
1.0 .9
If thr static pressure in the wak(, is rrpresented by PI!
r-- t--
r-- I---
-- I---
t-- I------
--- .9
the tl'U(, drug coeffieient uncorrected for bloeking ca' may be
--
t-- I---
r-- r--
;--..
t-- r-- t--. :--.....
---
B
shown to be (refel'l'nee 84)
r---.,.
r--- I---
~ t-- .8
t--- r-- I'--
----
t-- i"---j-.... "~ I'-....
r---- l'--.-
.7 (26) r----.-
/'--. ~ ~
tZ
K '-...
f"".-
1'--- ~
.6 I--- l-- cc'=Kc dr ~ ~
l-J-i: 1_ WaKe depth ~
f- · h' ffi' t' th k Ho-PI I 8 'W l{'re 1 IS t I' stati<,-pl'essure COl' ('len m . e wa -e --_.
""-.
"(mo,) ~ q, qo .5
"" S' _ Sfa Ie pressure
The assumption is made that the variation of total pressure f-- j..--'.J:..... --t- q,
acrcss tIl(' wake can be represented by a normal probability
.4 curve. Tht' drag cOPfficient ca' is tllPn easily obtainable from
measurements of CdT by means of a factor K, the ratio of c/
.I .3 .4 .5 .6 .7 .8 .9 J.O to Cd , which depC'uds only on 8 and the maximum value of T
lIe. If the maximum value of He is rt'prf'sentf'd by H '
emax
FH.CRE tl(t--Plot of ](us u function of l1c with 8 as U paramC'ter.
maz
tIl(' NlllH tiOll of the normal probability cune is
TIll' pou'lltial function w for a symm('trical body IS ginll b~- (28) where B is a dinH'llSiollless eonstnnt that determines the width of the wlIk(,. If a ('onvenient varinble of intpgration where ,. is tlll' fl'('l'-str('am veloeity and the coefficients AI,
Byw' 1 I . K'
A ••• H1'(' comph,x. If the tunnel ht'ight is large COIll- Y = .. - .. IS use( , t Ie ratIO IS 2 ,
c
pal'l'd to tIl(' size of tIll' body, POW('I'S of liz greater than 1
may b(, Iwgl(>ded and (29) This opl'l'Htion is ('quivaknt to rpplacing the body by a eil'elc of whieh the doublpt strength is 2'llA the term AI/z repn'- I ; Sl'nts thr disturbnnec to the fr('('-stn'am flow. The total and is ind('pend<'llt of the width of the wah. TIl(' quantity indu('pd Y('locity at the center of tIl(' body dw' to all thf'
K has b('PI1 evaluatpd for variolls values of He and 8 bv
1 . rna;c ....
imng('s is ('XP"('ss('d ill l'('f(,l'en('(' 86 as
assuming 8 to b(' constant aeross thE' wake. The drag
coeffieipnt c/ may thus be obtained from tunn('l mNl.sur('-
(30)
ments of Cd , He , and 8 A plot of K as a funetion of He
1 • T mu mn
with 8 as param('t!:'r is given in figurE' 66. A paralIc,1 t1'l'at-
where the term Al is the same as the term "4 'At2 V of
mpnt of this probll'm is given in rpfrr(,lH'P 85.
reference 86.
For eOllveniellep in tunnd cal('ulatiolls, the t'xpressioll of TUNNEL-WALL CORRECTIONS AF may be written In two-dim(,llsional flow, the tunnE'1 walls may be eOllV(,I1-
AV
(31)
V=Au
iently eOllsid(,l'('(i as having two distinct dff'cts upon thE' flow where over a mod('l in a tunnrl: (1) an inerease in the frE'e-stream velocity in titP neighborhood of thE' model bpcausr of a (32) (;onstrietion of tIl<' flow and (2) a distortion of the lift distribution from tIl!' indu(,E'd curvature of the flow.
(33) The inca-cast' in fn.>('-str('am vplocity caused by the tunnel walls (blocking pfl'Pet) is obtaill('d from consideration of an infinite vprtieal "ow of imag('s of a symmetrical body as The faetor 0' d('pends only Oil the size of the body and is given in rpf(,I'('I\(,(, 86; tll(> irriag('s rl'pn'sent the ('fl'(>d of th(' easily calculatE'd. The factor A depends on the shape of the tunnd wnlls.
body and is more diffi('ult tocttlculate. For bodi('s such as REPORT NO. 824-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
sponding points of the upper and lower surfaces, and
Rankine ovals and ellipses, simple formulas may be obtained
for calculating A. In the general case, the value of A may
{ y de{! may be replaced by an integration over the upper
be obtained from the velocity distribution over the body by .10
surface; therefore,
the expression
fo z ~~ dz=2i f y de{! (counterclockwise direction)
(34)
A=l: .fol ~ V~l+(~J dG)
or
where v is the velocity at any point on the airfoil surface and
dy ddx is the slope of the airfoil surface at any point of which
the ordinate is Yt.
Reversing the path of integration, replacing de{! by vds, replac-
y
ing ds by ~ 1 + d:~2 dx, and solving for A = lC~~l gives
ods ..... ····Polh of Inlegralien
A=.16 (I'lL ~ 11 + (dyt)J d(~)
11'.10 c V -V dx c
u, .1.'
where the integration is taken from the leading edge to the
trailing edge over the upper surface.
In addition to the error caused by blocking, an error .}xists
in the measured tunnel velocity'because of the interference
effects of the model upon the velocity indicated by the static-
pressure orifices located a few feet upstream of the model
FIGURE 67.-Sketch for derivation of A-factor.
and halfway between floor and ceiling. In order to correct
for this error, an analysis was made of the velocity distribu-
In order to obtain this expression, consider the flow past a
tion along the streamline halfway between the upper and the
symmetrical body as shown in figure 67. The potential
lower tunnel walls for Rankine ovals of various sizes and thick-
function for this flow is given by equation (28). Differen-
ness ratios. The analysis showed'that the correction could
tiating and multiplying equation (28) by z gives
be expressed, within the range of conventional-airfoil
thickness ratios, as a product of a thickness factor given
by the blocking factor A and a factor ~ which depended upon
the size of the model and the distance from the static-pressure
orifices to the midchord point of the model. The corrected
The line integral about a closed curve fa z ~~ dz will
indicated tunnel velocity V' could then be written
depend only on the term -Adz and, from the theory of
(36)
residues, is given by V'= V"(1+A~)
r dWd 'A where V" is the velocity measured by the static-pressure
Jo z dz z=-211'~ 1
orifices. In the Langley two-dimensional low-turbulence
but
tunnels, the distance from the static-pressure orifices to the
dw
midchord point of the model is approximately 5.5 feet; the
z dz dz=z dw
corresponding value of ~ for a 2-foot-chord model is approxi-
= (x+iy) (dcjJ+i dt/;)
mately 0.002.
In order to calculate the effect of the tunnel walls upon the
where e{! is the potential function and if; is the stream func-
lift distribution, a comparison is made of the lift distribution
tion. On the surface of the body dif;=O, so that
of a given airfoil in a tunnel and in free air on the basis of
thin-airfoil theory. It is assumed that the flow conditions
W
r z dd dz= { x de{!+i { y de{!
(35)
in the tunnel correspond most closely to those in free air when
Jc z Jo Jo
the additional lift in the tunnel and in free air are the same
Since the body is symmetrical, the term x de{! will have
(reference 87). On this basis the following corrections are
equal numerical values but opposite signs at corres,Ponding
derived (reference 87), in which the primed quantities refer
to the coefficients measured in the tunnel:
points of the upper and lower surfaces, and fax de{! will
vanish. The term y de{! will have equal values at corre-, cl=[1-2A(o+~) -u]cz' (37)
SUMMARY OF AIRFOIL DATA
me from unity in the high-lift range for ap.y airfoil tested in the
4' -(1+) '+ 4uc / , (38)
Il'o- Ull'o d'id ,-UIl'Io
tunnel; this variation indicates a change in blocking at high
CI Il'o
lifts. A plot of FIFo against angle of attack cxo' for a 2-foot-
chord model of the NACA 643-418 airfoil is given in figure 68.
(39)
The quantity FIFo is nearly constant for values of cxo' up to
12°; but for values of Il'o' greater than 12°, FIFo increases and
·, 4 UC m 14'
I
h f n t e oregolllg equatlons, the terms d 'Ide " UCXI ,and uc/14 the increase is partIcularly noticeable at and over the stall.
,CI CXo 0
are usually negligible for 2-foot-chord models in the Langley
1.20 two-dimensional low-turbulence tunnels.
When the effect of the tunnel walls on the pressure distri-
1.10
bution over the model is small, the wall effect on the drag is
V
~r- merely that corresponding to an increase in the tunnel E!peed.
o ,.."...
>-. :r-o-
f/.oo
The correction to the drag coefficient is therefore given by the
following relation: '
.90 ca=[1-2A(u+m ca' (40) .80
Similar considerations have been applied to the development
-
- -
16 12 8 -4 0 4 8 12 /6 20 24 of corrections for the pressure distribution in reference 87. Geometric angle of attacK, a;, deg
Equation (40) neglects the blocking due to the wake, such
FIGURE 68.-Additional blocking factor at the tunnel walls plotted against angle of attack for the NACA 643-418 airfoil.
blocking being small at low to moderate drags. The effect
of a pressure gradient in the tunnel upon loss of total pressure
A theoretical comparison was made of the blocking factor
in the wake is not easily analyzed but is estimated to be small.
Au and the velocity measured by the floor and ceiling orifices
The effect of the pressure gradient upon the drag has there-
for a series of Rankine ovals of various sizes and thickness
fore been disregarded. When the drag is measured by a
ratios. The quarter-chord point of each oval was located at
balance, the effect of the pressure gradient upon the drag is
the pivot point, the usual position of an airfoil in the tunnel.
directly additive and a correction should be applied. For
The analysis showed the relation between the blocking factor
large models, especially at high lift coefficients, the effect of
Au and the change in F to be unique for chord ler gths up
the tunnel walls is to distort the pressure distribution appre-
to 50 inches in that different bodies having the same blocking
ciably. Such distortions of the pressure distribution may
factor Au gave approximately the same value of F. For
cause large changes in the boundary flow and no adequate
chords up to 50 inches, the relationship is
corrections to any of the coefficients, 'particularly the drag, can be found.
(41)
A: =0.45 (~ -1)
CORRECTION FOR BLOCKING AT mGH LIFTS
where A VIV is the true increment in tunnel velocity due to
SO long as the flow follows the airfoil surface, the foregoing
blocking; The foregoing relation :was adopted to obtain the
relations account for the effects of the tunnel walls with suffi-
cient accuracy. When the flow leaves the surface, the block-
correction to the blocking in the range of lifts where ~o > 1.
ing increases because of the predominant effect of the wake
Considerable uncertainty exists regarding the correct
upon the free-stream velocity. Since the wake effect shows
numel1ical value of the coefficient occurring in equation (41).
up primarily in the drag, the increase in blockilig would
If a row of sources, rather than the Rankine ovals used in
logically be expressed in terms of the drag. The accurate
the present analysis, is considered to represent the effect of
measurement of drag under these conditions by means of a
the wake, the value of the coefficient in equation (41) would
rake is impractical because of spanwise movements of loW--
be approximately twice the value used. Fortunately, the
energy air. A method of correcting for increased blocking
correction amounts to only about 2 percent at maximum lift
at high angles of attack without drag measurements has
for an extreme condition with a 2-foot-chord model. Further
therefore been devised for use in the Langley two-dimensional
refinement of this correction has therefore not been attempted.
low-turbulence tunnels. '
Readings of the floor and ceiling velocities are taken a few
COMPARISON WITH EXPERIMENT
inches ahead of the quarter-chord point and averaged to
A check of the validity of the tunnel-wall corrections has
remove the effect of lift. This average F, which is a measure
been made in reference 87, which gives lift and moment
of the effective tunnel velocity, is essentially constant in the
curves for models having various ratios of chord to tunnel
low-lift range. The quantity FIFo, where Fo is the average
value of F in the low-lift range, however, shows a variation height, uncorrected and corrected for tunnel-:-wall effects.
REPORT NO. 824-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS .0
/ f\
iit" K
V
V
r
~ 2 I.
!
!
\,
I /
If
I
A
/
.4
If I
1 ~
!I
!
II
I
-.
)
!
o Airfoil B Pressure distnbution,
Pressure distribution, II
t Airfoil A
rDr test 655 I
/ TDT test 640
.8
-
er Airfoil Integratin$. manometer, Iregratinq manomr ,
I Ai~fOil ~
TDT test 6 3 and 654
'\ 1I TDr fest 618
l
(b) (a)
V
(~
.2 -8 /6 24 32 -24 -16 ~8 32
-/6 8 o o 8 16 24
Section angle of attack, «OJ deg (b) Comparison for airfoil B.
(a) Comparison for airfoil A.
FIGURE 69.-Comparison between lifts obtained from pressure·distribution measurements and lifts obtained from rea~tions on the 1100r and ceiling of the tunnel.
l Chord. in.
2.0
~ Jl""'~r";'fed tor blocking
/.6
~ J~ cor;"ecfed for' b/~ckinq
/.8 1.2 <:!
/.6
V'
>-
i"-
---'
l ~
/.4 <1 \
V
~
r
~~ &;; ~
\
/.2 .,
I
~
~ 1\
S
II
\
1.0
~
I
'"
,
1/
.8 o Balance
_J
o Integrating manometer _ .6
II
j
.4 ~.B .2 -1.2 -24 -16 -8 0 8 16 24 Section angle of attock, «., deg .6 .7 .9
o . I .2 .3 .4 .5 .8
FIGURE 70.-Comparison between lifts obtained from balance measurements and from :x/c reactions on the floor and ceiling of the tunnel.
FIGURE 71.-Comparison between corrected arid uncorrected pressure distributions for two chord sizes of a symmetrical NACA 6-series airfoil of l(i·percent thickness. ,"0=0·, SUMMARY OF AIRFOIL DATA 12. Von Doenhoff, Albert E., and Stivers, Louis S., Jr.: Aerodynamic The general agreement of the corrected curves shows that Characteristics of the NACA 747A315 and 747A415 Airfoils the method of correcting the lifts and moments is valid.
from Tests in the NACA Two-Dimensional Low-Turbulence A comparison is made in reference 87 between the theoreti- Pressure Tunnel. NACA CB No. L4I25, 1944.
cal correction factor (equation (40)) and the experimentally 13. Naiman, Irven: Numerical Evaluation by Harmonic Analysis derived corrections of reference 88. The theoretical cor- of the E-Function of the Theodorsen Arbitrary-Airfoil Potential Theory. NACA ARR No. L5H18, 1945.
rection factors were found to be in good agreement with those 14. Theodorsen, Theodore: Airfoil-Contour Modification Based on obtained experimentally.
E-Curve Method of Calculating Pressure Distribution. N ACA In order to check the validity of the 'I-factor, a comparison ARR No. L4G05, 1944.
has been made of lift values obtained from pressure dis:..
15. Allen, H. Julian: A Simplified Method for the Calculation of tributions with those obtained from the integration of the Airfoil Pressure Distribution. NACA TN No. 708, 1939.
16. Munk, Max M.: Elements of the Wing Section Theory and of floor and ceiling pressures in the tunnel. A comparison for the Wing Theory. NACA Rep. No. 191, 1924.
two airfoils given in figure 69 shows that the two methods of 17. Glauert, H.: The Elements of Aerofoil and Airscrew Theory.
measuring lift give results that are in good agreement. The Cambridge Univ. Press, 1926, pp. 87-93.
'I-factor has also been checked by comparison of the lift 18. Theodorsen, Theodore: On the Theory of Wing Sections with obtained _ from balance measurements with the integrating- Particular Reference to the Lift Distribution. N ACA Rep.
No. 383, 1931.
manometer values in figure 70.
19. Von Karman, Th.: Compressibility Effects in Aerodynamics.
Finally, a check has been made of the method of correcting Jour. Aero. ScL, vol. 8, no. 9, July 1941, pp. 337-356.
pressure distributions (reference 87) for NACA 6-series air- 20. Von Doenhoff, Albert E.: A Method of Rapidly Estimating the foils of two chord lengths at zero angle of attack in figure 71, Position of the Laminar Separation Point. NACA TN No.
in which the pressure coefficients are plotted against chord- 671, 1938.
21. Jacobs, E. N., and Von Doenhoff, A. E.: Formulas for Use in
wise position x/c. The agreement between the corrected
Boundary-Layer Calculations on Low-Drag Wings. N ACA pressure distributions for both models verifies the method of ACR, Aug. 1941. ' making the tunnel-wall corrections.
22. Von Doenhoff, Albert E., and Tetervin, Neal: Determination of General Relations for the Behavior of Turbulent Boundary Layers. NACA Rep. No. 772, 1943.
REFERENCES 23. Squire, H. B., and Young, A. D.: The Calculation of the Profile Drag of Aerofoils. R. & M. No. 1838, British A. R. C., 1938.
1. Jacobs, Eastman N., Ward, Kenneth E., and Pinkerton, Robert 24. Tetervin, Neal: A Method for the Rapid Estimation of Turbulent M.: The Characteristics of 78 Related Airfoil Sections from Boundary-Layer Thicknesses for Calculating Profile Drag.
Tests in the Variable~Density Wind Tunnel. NACA Rep.
NACA ACR No. L4Gl4, 1944.
No. 460, 1933.
25. Quinn, John H., Jr., and Tucker, Warren A.: Scale and Turbulence 2. Jacobs, Eastman N., and Pinkerton, Robert M.: Tests in the Effects on the Lift and Drag Characteristics of the Variable-Density Wind Tunnel of Related Airfoils Having the NACA 65 --418, a= 1.0 Airfoil Section. NACA ACR No. L4Hll, Maximum Camber Unusually Far Forward. NACA Rep.
1944.
No. 537, 1935._ 26. Tucker, Warren A., and Wallace, Arthur R.: Scale-Effect Tests 3. Jacobs, Eastman N., Pinkerton, Robert M., and Greenberg, in a Turbulent Tunnel of the NACA 65 -418,. a=1.0 Airfoil Harry: Tests of Related Forward-Camber Airfoils in the Section with 0.20,Airfoil-Chord Split Flap. NACA ACR No.
Variable-Density Wind Tunnel. NACA Rep. No. 610, 1937.
L4122, 1944.
4. Stack, John, and Von Doenhoff, Albert E.: Tests of 16 Related 27. Davidson, Milton, and Turner, Harold R., Jr.: Effects of Mean- Airfoils' at High Speeds. NACA Rep. No. 492, 1934.
Line Loading on the Aerodynamic Characteristics of Some Low- 5. Jacobs, Eastman N., and Sherman, Albert: Airfoil Section Drag Airfoils. NACA ACR No. 3127, 1943.
Characteristics as Affected by Variations of the Reynolds 28. Von Doenhoff, Albert E., and Tetervin, Neal: Investigation of Number. NACA Rep. No. 586, 1937. the Variation of Lift Coefficient with Reynolds Number at a Moderate Angle of Attack on a Low-Drag Airfoil. NACA 6. Pinkerton, Robert M., and Greenberg, Harry: Aerodynamic CB, Nov. 1942.
Characteristics of a Large Number of Airfoils Tested in the 29. Oswald, W. Bailey: General Formulas and Charts for the Calcula- Variable-Density Wind Tunnel. NACA Rep. No. 628, 1938.
tion of Airplane Performance. NACA Rep. No. 408, 1932.
7. Jones, B. Melvill: Flight Experiments on the Boundary Layer._ 30. Millikan, Clark B.: Aerodynamics of the Airplane. John Wiley Jour. Aero. Sci., vol. 5, no. 3, Jan. 1938, pp. 81-94.
& Sons, Inc., 1941, pp. 108-109.
8. Jacobs, Eastman N., and Abbott, Ira H.: Airfoil Section Data 31. Hood, Manley J.: The Effects of Some Common Surface Obtained in the N.A.C.A. Variable-Density Tunnel as Affected Irregularities on Wing Drag. NACA TN No. 695, 1939.
by Support Interference and Other Corrections. NACA Rep.
32. Loftin, Laurence K., Jr.: Effects of Specific Types of Surface No. 669, 1939.
Roughness on -Boundary-Layer Transition. NACA ACR No.
9. Theodorsen, Theodore: Theory of Wing Sections of Arbitrary L5J29a, 1946.
Shape. NACA Rep. No. 411, 1931.
33. Charters, Alex C., Jr.: Transition between Laminar and Turbulent 10. Stack, John: Tests of Airfoils Designed to Delay the Compress- Flow by Transverse Contamination. NACA TN No. 891,1943.
ibility Burble. NACA Rep. No. 763, 1943.
34. Braslow, Albert L.: Investigation of Effects of Various Camouflage 11. Jacobs, Eastman N.: Preliminary Report on Laminar-Flow Airfoils Paints and Painting Procedures on the Drag Characteristics of an NACA 65(421)-420, a=1.0 Airfoil Section. NACA CB No.
and Jirew Methods Adopted for Airfoil and Boundary-Layer Investigations. NACA ACR, June 1939. L4Gl7, 1944. - REPORT NO. 824-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 35. Jones, Robert T., and Cohen, Doris: A Graphical Method of 55. Jones, Robert T., and Ames, Milton B., Jr.: Wind-TUlinel Inves- Determining Pressure Distribution in Two-Dimensional Flow.
tigation of Control-Surface Characteristics. V-The Use of a NACA Rep. No. 722, 1941. Beveled Trailing Edge to Reduce the Hinge Moment of a 36. Abbott, Frank T., Jr., and Turner, Harold R., Jr.: The Effects Control Surface. NACA, ARR, March 1942.
of Roughness at High Reynolds Numbers on the Lift and Drag
56. Sears, Richard r., and Liddell, Robert B.: Wind-Tunnel Investiga-
Characteristics of Three Thick Airfoils. NACA ACR No. tion of Control-Surface Characteristics. VI-A 30-Percent- L4H21,1944.
Chord Plain Flap on the NACA 0015 Airfoil. NACA ARR, 37. Jacobs, Eastman N., Abbott, Ira H., and Davidson, Milton: June 1942.
Investigation of Extreme Leading-Edge Roughness on Thick 57. Wenzinger, Carl J., and Delano, James B.: Pressure Distribution Low-Drag Airfoils to Indicate Those Critical to Separation. over an N. A. C. A. 23012 Airfoil with a Slotted and a Plain Flap.
NACA CB, June 1942. NACA Rep. No. 633, 1938.
38. Zalovcik, John A.: Profile-Drag Coefficients of Conventional 58. Gillis, Clarence 1.., and Lockwood, Vernard E.: Wind-Tunnel and Low-Drag Airfoils as Obtained in Flight. NACA ACR Investigation of Control-Surface Characteristics. XIII-Various No. L4E31, 1944. Flap Overhangs Used with a 30-Percent-Chord Flap on an 39. Zalovcik, John A., and Wood, Clotaire: A Flight Investigation of NACA 66-009 Airfoil. NACA ACR No. 3G20, 1943.
the Effect of. Surface Roughness on Wing Profile Drag with 59. Rogallo, F. M.: Collection of Balanced-Aileron Test Data. NACA ACR No. 4All, 1944.
Transition Fixed. NACA ARR No. L4I25, 1944.
40. Hood, Manley J., and Gaydos, M. Edward: Effects of Propellers 60. Denaci, H. G., and Bird, J. D.: Wind-Tunnel Tests of Ailerons at and of Vibration on the Extent of Laminar Flow on the Various Speeds. II-Ailerons of 0.20 Airfoil Chord and True N. A. C. A. 27-212 Airfoil. NACA ACR, Oct. 1939. Contour with 0.60 Aileron-Chord Sealed Internal Balance on the 41. Silverstein, Abe, Katzoff~ S., and Hootman, James A.: Com- NACA 66,2-216 Airfoil. NACA ACR No. 3F18, 1943.
parative Flight and Full-Scale Wind-Tunnel Measurements of 61. Purser, Paul E., and Riebe, John M.: Wind-Tunnel Investigation the Maximum Lift of an~~iIi?lane. NACA Rep. No. 618, 1935. of Control-Surface Characteristics. XV-Various Contour 42. Sweberg, Harold H., and . Dirigeldein, Richard C.: Summary of Modifications of a 0.30-Airfoil-Chord Plain Flap on an Measurements' in Langley Full-Scale Tunnel of Maximum NACA 66(215)-014 Airfoil. NACA ACR No. 3L20, 1943.
Lift Coefficients and Stalling Characteristics of Airplanes. 62. Braslow, Albert L. : Wind-Tunnel Investigation of Aileron Effec- NACA Rep. No. 829, 1945. tiveness of 0.20-Airfoil-Chord Plain Ailerons of True Airfoil 43. Purser, Paul E., and Johnson, Harold S.: Effects of Trailing- Contour on N ACA 65 415, 65 -41S, and 65 -421 Airfoil Sections.
r 3 4 Edge Modifications on Pitching-Moment Characteristics of NACA CB No. L4H12, 1944.
Airfoils. NACA CB No. L4I30, 1944. 63. Sears, Richard I., and Purser, Paul E.: Wind-Tunnel Investigation 44. Fullmer, Felicien F., Jr.: Wind-Tunnel Investigation of NACA of Control-Surface Characteristics. XIV-NACA 0009 Airfoil 66(215)-216, 66,1-212, and 65 -212 Airfoils with 0.20-Airfoil- with a 20-Percent-Chord Double Plain Flap. NACA ARR Chord Split Flaps. NACA CB No. L4G10, 1944. No. 3F29, 1943.
45. Abbott, Ira H., and Greenberg, Harry: Tests in the Variable- 64. Crane, Robert M., and Holtzclaw, Ralph W.: Wind-Tunnel Inves- Density Wind Tunnel of the N. A. C. A. 23012 Airfoil with tigation of the Effects of Profile Modifications and Tabs on the Plain and Split Flaps. NACA Rep. No. 661, 1939. Characteristics of Ailerons on a Low Drag Airfoil. N ACA Rep.
46. Wenzinger, Carl J., and Harris, Thomas A.: Wind-Tunnel Investi- No. S03, 1944.
gation of N. A. C. A. 23012, 23021, and 23030 Airfoils with 65. Von Doenhoff, Albert E., and Horton, Elmer A.: Preliminary Various Sizes of Split Flap. NACA Rep. No. 668, 1939. Investigation in the NACA Low-Turbulence Tunnel of Low- 47. Bogdonoff, Seymour M.: Wind-Tunnel Investigation of a Low- Drag-Airfoil Sections Suitable for Admitting Air at the Leading Drag 'Airfoil Section with a Double Slot,ted Flap. NACA ACR Edge. NACA ACR, July 1942 No. 3120, 1943. ' 66. Jacobs, Eastman N., and Ward, Kenneth E.: Interference of Wing 4S. Wenzinger, Carl J., and Harris, Thomas A.: Wind-Tunnel Investi- and Fuselage from Tests of 209 Combinations in the N. A. C. A.
gation of an N. A. C. A. 23012 Airfoil with Various Arrangements Variable-Density Tunnel. NACA Rep. No. 540, 1935.
of Slotted Flaps. NACA Rep. No. 664, 1939. 67. Abbott, Ira H.: Interference Effects of Longitudinal Flat Plates on 49. Wenzinger, Carl J., and Harris, Thomas.A.: Wind-TunneIInvest.i- Low-Drag Airfoils. NACA CB, Nov. 1942.
gation of an N. A. C. A. 23021 Airfoil with Various Arrangements 6S. Ellis, Macon C., Jr.: Some Lift and Drag Measurements of a of Slotted Flaps. NACA Rep. No. 677, 1939.
Representative Bomber Nacelle on ,a Low-Drag Wing-II.
NACA CB, Sept. 1942.
50. Swanson, Robert S., and Crandall, Stewart M.: Analysis of A vail- 69. Ellis, Macon C., Jr.: Effects of a Typical Nacelle on the Charac- able Data on the Effectiveness of Ailerons without Exposed teristics of a Thick Low-Drag Airfoil Critically Affected by Overhang Balance. NACA ACR No. L4E01, 1944.
Leading-Edge Roughness. NACA CB No.3D27, 1943.
51. Street, William G., and Ames, Milton B., Jr.: Pressure-Distribu- 70. Allen, H. Julian, and Frick, Charles W., Jr.: Experimental Investi- tion Investigation of an ,N. A. C. A. 0009 Airfoil with a 50- gation of a New Type of Low-Drag Wing-Nacelle Combination.
Percent-Chord Plain Flap and Three Tabs. NACA TN No.
NACA ACR, July 1942.
734, 1939.
71. Abbott, Frank T., Jr.: Lift and Drag Data for 30 Pusher-Propeller 52. Ames, Milton B., Jr., and Sears, Richard 1.: Pressure-Distribution Shaft Housings on an NACA 65,3-018 Airfoil Section. NACA Investigation of an N. A. C. A. 0009 Airfoil with an SO-Percent- ACR No. 3K13, 1943.
Chord Plain Flap and Three Tabs. N ACA TN No. 761, 1940.
72. Robinson, Russell G., and Wright, Ray H.: Estimation of Critical 53. Ames, Milton B., Jr., and Sears, Richard I.: Pressure-Di!~tribution Speeds of Airfoils and Streamline Bodies. NACA ACR, March Investigation of an N. A. C. A. 0009 Airfoil with a 30-Percent- 1940.
Chord Plain Flap and Three Tabs .. NACA TN No. 759, 1940.
73. Anderson, Raymond F.: Determination of the Characteristics of 54. Sears, Richard I.: Wind-Tunnel Investigation of Control-Surface Tapered Wings. NACA Rep. No. 572, 1936.
Characteristics. I-Effect of Gap on the Aerodynamic Charac- 74. Jacobs, Eastman N., and Rhode, R. V.: Airfoil Section Charac- teristics of an N ACA 0009 Airfoil with a 30-Percent-Chord teristics as Applied to the Prediction of Air Forces and Their Plain Flap. NACA ARR, June 1941. Distribution on Wings. NACA Rep. No. 631, 1938.
SUMMARY OF AIRFOIL DA'l'A 75. Soule, H. A., and Anderson, R. F.: Design Charts Relating to the 83. Pearson, Henry A., and Anderson, Raymond F.: Calculation of Stalling of Tapered Wings. NACA Rep. No. 703, 1940. the Aerodynamic Characteristics of Tapered Wings with Partial- 76. Harmon, Sidney M.: Additional Design Charts Relating to the Span Flaps. NACA Rep. No. 665, 1939.
Stalling of Tapered Wings. NACA ARR, Jan. 1943. 84. The Cambridge University Aeronautics Laboratory: The Measure- 77. Anderson, Raymond F.: The Experimental and Calculated Char- mentof Profile Drag by the Pitot-Traverse Method. R. & M.
acteristics of 22 Tapered Wings. NACA Rep. No. 627, 1938. No. 1688, British A. R. C., 1936.
78. Tani, Itiro: A Simple Method of Calculating the Induced Velocity 85. Silverstein; A., and Katzoff, S.: A Simplified Method for Determin- of a Monoplane Wing. Rep. No. 111 (Vol. IX, 3), Aero. Res. ing Wing Profile Drag in Flight. Jour. Aero. ScL, vol. 7, no. 7, Inst., Tokyo Imperial Univ., Aug. 1934. May ]940, pp. 295-301.
79. Sherman, Albert: A Simple Method of Obtaining Span Load 86. Glauert, H.: Wind Tunnel Interference on Wings, Bodies and Distributions. NACA TN No. 732, 1939. Airscrews. R. & M. No. 1566, British A. R. C., 1933.
80. Jones, Robert T.: Correction of the Lifting-Line Theory for the 87. Allen, H. Julian, and Vincenti, Walter G.: Interference in a Two- Effect of the Chord. NACA TN No. 817, 1941. Dimensional-Flow Wind Tunnel with the Consirleration of the 81. Cohen, Doris: Theoretical Distribution of Load over a Swept- Effect of Compressibility. NACA Rep. No. 782, 1944.
Back Wing. NACA ARR, Oct. 1942. 88. Fage, A.: On the Two-Dimensional Flow past a Body of Symmet.rical 82. Pearson, H. A.: Span Load Distribution for Tapered Wings with Cross-Section Mounted in a Channel of Finite Breadth.
Partial-Span Flaps. NACA Rep. No. 585, 1937. R. & M. No: 1223, British A. R. C., 1929.
~ .." ~ >-3 00 ~ >I> ~ .....
~ ~ ? ~ o ~ ~ > Sl ..... Ul o ::tI ><i o o ~ ~ ..... ...., >-3 trl trl "'J o ::tI > trl ::tI o ~ > q >-3 ..... o Ul a for to· to· cs de- satis- satis- satis- tip entire abrupt p~rtial stall; flaps progresses stall and with wit~ toward neutral, tip with progression progresslOn characteristi stall envelops stall stall for fnll-span· flaps tips UPO; flaps detlected; no neutral flaps from root flaps span data factory; flected, extremely stall wing factory ward strong outflow resulted in severe factory. ward Stalling Abrupt With Abrupt Unsatisfactory Abrupt
I ,----
36 41 78 98 15 27 37 34 1.26 • 1. 1. 1.72 1. 1.84 1.94 1. 2.07 1. 2.03 2.06 2.04 2.11 2.15 2.40 2.51 2.43 2.49 2.52 1. 1.29 1. 2.29 3.13 3.31 3.29 1.18 1. 1.43 1.17 1.31 1.
2.44 2.49 2.36 2.49 2.54 .2.50 1 2.10 2.19 2.21 CLmr.u:
·--1 -i-:-~~-I
,,,,,:~,
R 6XIO~· I __
lX106 I I I • 2.
3.6 4.6 2.6 3.6 4.6 2.6 3.6 4.6 2.6 3.6 4.6 2.6 3.6 4.6 2.6 3.6 4.6 2.6 3.6 4.6 2. 2.8 3.3 2. 2.9 3.4 2. 2.9 3.4 2. 2.9 3.4 3.0XJ06 5. 7.4 3.3XI0' 5.6 7 8 3.3 5.2 5.8 7.2 ~ I 1- '-~-:~-X-1;;
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span
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cr. 1 10 20 30 30
__
-
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I I
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20 30 10 20 30 (percent Flap.chord
No~~
THE NACA 19-FOOT PRESSURE TUNNEL
I 1 I
_..
I I I
IN I I I o I
.j, 1 .j, 30
Jr. 60 ...
______________ _ • angle 1--1--1--1--1---1---- (deg)
1 1 I I o
I I
60 1 30 30 50
~fi 35 Flap __
I-~I--I--I-.--I--I--I------I I-o-'=='-y-'- I
I I I 1
1 1 j -J.
T None Noue
Nine None Fowler Outboard 1----1--1--1--\-·-1---1--1---1---1--- :-----., Flap MODELS TESTED
I
OF .j, 1
Split None Pliin None None Fowler Inboard -1---1 line 0 0 0 0 0 0.0 1.5 3.6 2.5 chord 66(215)-216 0.25 66(215)-(1.8)12, 66(215)-116 65(318)-{)19 65(318)-{)19 66(215)-216 66(215)-216 65(318)-{)15 65(318)-{)15 of washout, washout, washout; washout, .
NAGA NAGA NACA NAGA back 0 NACA NAGA NACA NAGA NACA 21.93 a=0.6 a=O.6 =5.82 Geometric characteristics Root: Tip: A=7.00 ~=1.00 Geometric Root: Tip: A=7.0 ~=0.5 Geometric Root: Tip: A=7.36 ~=O.25 Root: Tip; A=7.36 ~=0.25 Geometric washout, 4.0 Sweep Tip; A X=0.46 Geometric Sections: Sections: Sections: Sections; se~gf:~ACA66(215)-(1.S)(15.5), I~metric STALLING CHARACTERISTICS ~ view AND Front LIFT ·------1---------1----'---,---
-====
Contlguration view H.-MAXIMUM Plan TABLE
~ ~
~
~
c--:-------=--=:J
I v II III IV Model ___ ---I --I
~ ~ ..... t::1
w q ~ >1 o ";j :.- ~ ";j S t< ~
~ Cl 1 I I I I I I I I I I I I I I I I I ·1 to- satis- with progression stall tips data factory ward Abrupt Satisfactory No Satisfactory Satisfactory Satisfactory Sati.factory 97 95 1.33 138 1.42 1.87 1.47 1.39 1.87 2.85 1. 2.11 1.
1.34 1.39 1.91 1.92 1.32 2.25 2.77 1. 2.22 01.37 01.42 01.45 <>2.19 02.20 02.21 02.00 02.06 02.06 3.3XI0' 5.3 6.0 3.3 5.1 5.8 5.1XlO' 5.1 2.4XI0' 2.4 2.4 2.5 2.5 3.5X10' 3.6 4.1 3.1XlO' 4.1 4.9 3.1 4.1 4.8 3.6 3.5XlO' 3.5 4.0XlO' 4.8 3.1 4.0
----
-- ..
I ,1. .!-
48 38 1 38
------- --ao--- ------- ------- ---- ------- --~----
------- ------- ----.-- -------
· .-- · · ---
I .!-
,1. ,l- 60 60 70 I
60 1 .10 60 1 60 60
-- -- -- -- -- --
--
1 1 ,l- 20
2.5 35 20 -._---- -----.- ------- ------- ------- -.-._--
------- -- -- ------- --- --
--
I I I I L .!-
25 2.1 ,1. 1 ,l- 20 25 I
35 20 20 20
-- -- ---- -- --
-- ----
,1.
48 60 55 ------ ------- .------ -_.---- -------
--.6.-- . ------- ------- -- -- --
-- -- -- --
0 0 0 ,l- ,l- ,l- 55 35 60 .10 55 48 60 55 -- ---55-- :
I I
I I .\. 1
,1. ,1. ,l- ,l- Zap Split Split None None None None None None slotted Double
I I
1 ,1. ,1. ,l- ./.
Zap Plain Split None Split Split None Split None Split slotted Double slotted Extensible (15.5), 0 0 0 0 0 0 0 a=1.0 2.5 1.0 65,3-318, 2.0 65(318)-316, 63(420)-418, 66(215)-(1.8)12, ACA 67(115)-116 64(215)-418 64(215)-418 64(215)-418 65M15.
67,1-115 66,2x-415 66,2x-415 66,2x-415 N washout, NACA washout, CA ACA NACA Mod. NACA NACA NACA NACA NACA N A NACA NACA Mod. NACA .10 .
=5.82 a=l.O a=0.6 a=0.6 a=O.8 a=0.8 Tip: Tip: Geometric Root: Geometric washout, Root: Geometric washout, 1.0 Root: A=8.92 ),,=0.33 Geometric washout, 1.0 Root: A=7.77 >'=0 Geometric washout, 2.8 Root:NACA66(215)-(1.8) A >'~0.46 Tip: A=8.09 >'=0.5 Geometric washout, 0.0 Root: Tip: A=6.7 >'=0.4 Tip. A=8.92 >.=0.33 Root: N Tip: A=8.92 >'=0.33 Geometric Tip: Sections: Sections: Sections: Sections: Sections: Sections: Sections:
I~
t}::=-
-=0= 'v-o-. ~=-
~
=-=0==
==
D
.
I \
,------ v -v -,
P' ~
~ ~
b:=J l1t sr~
-
- -.~ II V
(\ V
r -
V f\ V t
V
v
-
\r- -== ---.::J ---:'.
C d
~ C- d~b
~ c:t
X VI XI IX VIl VUI XII - • Propellers win<lmi!ling.
---. ---- l:d t::l 'tJ o ~ Z 00 .,.. II>- ~ > >-3 t-< o Z > t"' ~ rJl o ~ o ~ ~ t-< ~ t::l t::l "'! > t::l o z: g; >-3 t-< rJl
~ ~ ? > c g !:O C
for stall, slotted toward satisfac- first for rapidly, --- with stall, leftwing abrupt stalls very stall extensible progression wing abrupt tory tips all conditions the Stalling characteristif!S stalling left flap, satisfactory for split flap Satisfactory Abrupt Very Extremely Satisfactory Satisfactory I
I
21 76 89 72 98~ 32 46 55 98 TUNNEL-Concluded - 1. 1.37 1.45 1. 1. 1.00 1. 1.86 1. 1.99 2.13 2.01 2.15 2.21 1. 1.42 1. 2.27 2.34 2.37 2.04 2.13 2.16 2.02 2.12 2.17 1. 1.58 1.60 2.46 2.50 2.52 1.34 1.47 1.50 2.01 2.15 2.21 1.38 2.45 2.69 1.37 2.44 2.76 2.03 CLma~ '2.29 'I.
R 5.1 3.6X1Q6 61 3.6 5.2 6.3 3.5 4.9 5.9 3.5 4.S 5.9 3.4XI06 4.8 5.6 3.4 4.8 5.6 3.0XlO' 4.1 5.0 2.9 4.0 4.9 4.9 4.9 3.6X1Q6 3.1 2.8 3.6X106 3.1 2.8 2.4XI06 3.8 5.3 2.4 3.8 5:4 2.5 3.7 5.2 2.4 3.6 5.3 2.4 3.8 5.3 ---
PRESSURE b) I
.J.
bl. 30 31 1 31 1
span ------- ----._- -._---- ------- ------- ------- ------- -.----- ------- ------- ------- ------- -- -- ..
I I I I
I I I I I I I 1 I
1 .J. 1 1
Flap bl, 65 60 1 60 50 1 61; 65
(percent ---- --
-- -- ---- --
I I I I c) .(, 25 1 25 .J.
CI. 20 ------- -------
------- -._-._- ------- ------- ------- -- ------- -- ------- ------- ------- --
----
chord
I I 1 I I I I I I I I I I I
I I
-l- .J. .J. .J. 25 1 25 1
c" 20 20 20 24 NACA 19-FOOT (percent Flap
---- ----.-- ------. ---- -- --- -- ----
.------
1 I I .J. 0 .J. 0 .J.
1 25 25
a/. 60 THE --
---r ------- ------- ------- ------- ------- ------- ------- -------
-- -- ------- -- --
angle
I
(deg) IN I 1 0 0 0 0 .J. 0 .J.
tJ" 60 45 60 Flap 55 55 35 50 50 50 38 45 -----.- j --- I I I I
I I I I I I I I I I I 1
.J. 1 .J. .J.
1 1 1
None None None Double
TESTED None sPt
slotted Double slotted Outboard Flap edge
I I
.1-
1 .J. 1 .1-
Split Split None Split Split Double slotted Double slotted slotted Slotted Slotted slotted Inboard Extensible trailing Extensible Extensible 0 0 0 0 0 3.0 0.0 2.0 1.0 1.0 (216)-215, (216)-215, straight 66(215)-016 23015.6 66,2-118 65(216)-215, straight 65(216)-215, 66(215)-216 23009 66(2x15)-1l6 line washout, washout, washout, washout, washout, line NACA NACA NACA NACA65(318)-1(18.5) NACA NACA chord NACA NACA NACA NACA NACA_66(215)-o16 NACA chord =5.34> a=0.5 a=0.8 a=0.8 a=0.5 Geometric characteristics Roots: Tip: A Geometric:washout, 0.0 Root: Tip: A=5.52 >'=0.48 Geometric Root: Tip: A=5.5 >'=0.52 Geometric Root: Tip: Geometric Tip: A=9.08 -0.10 >'=0.68 Tip: A=6.9 Geometric Root: A=9.08 >'=0.45 0.2 Root: >'=0.45 Geometric Sections: Sections: Sections: Sections: Sections: Sections: view -=Q=.
-=Q== Front AND STALLING CHARACTERISTICS OF MODELS
==--A==-
----~
~ ~
LIFT Configuration
~
lL-d
view p-~
.D:-::::>
. )::=='"
V Ol1,:J V ~
V
p.
,.... ,....
F Plan II.-MAXIMUM
c:.L C:(f
~ ~---==;
~~
~-~
I I TABLE XV XIV XVI XIII XVII Model XVIII
---
....
rJ2 q ~ ~ ~ ~ > ~ o .... t:"' t:j ~ :>
t::.l --l tip _ Occurs stall all conditions for ____ initial tips at stall except full-span flap Satisfactory Satisfactory Satisfactory Poor, UnsatisfactorY, severe Satisfactory Satisfactory I
1 1 1 I
45 86 37 51 83 56 80 ~ 1. 2.57 2. 2.6.'\ 1. 1.50 1.60 1.38 1.57 1.61 1. 1.99 2.02 1.34 1. 1.63 1.85 1.92 2.0l 1.17 1.27 1.37 2.21 2.23 2.30
1.45 2. 1.23 1.43 1. t 1.90 2.01 2.04 2.43 2.02
1--1------- 1--1-- 6 6 6 _ 5X106 4.1 2.9 5.0 5.0 3.0 5.1 1. 2.2 1.4 1.9 2.7 3.6X10 3.1 2.S 5.5X10 5.5 5.5 2.9X10 4.0 4.9 3.0 4.0 4.9 3.0 4.3 5.1 5.2 4.1XlO' 4.9 4.1 2.9XlO' 4.9 4.1 2.9X10· 4.2 4.2 2.8 __
1 1
31 .l-
1 1
65 j
65 1 50 50 50 63 1
__ , , -1
25 1
25 20 _____ ______ ______ . .
__ 1 '-
25 1 I I
,j, 1
25 1 20 20 20 IS
I
__ 1--1--1--1-_1, -1 -1 __ o .l- 25 o -1--1--1--1--1 .~ ...
::::::-: ___ 1______ 1 1-· 1_ o o .l- o 55 ,J.. 45 45 40
45 60 1 45
__ 1 1--1--1--1--1--1--11--- 1--1--1--1--1--1--1---
,J.. ,J.. j
i 1 1 1
Split None None None None None DOUble slotted slott!,d Double 1-----1--1--1--1--1--1--11---
I
.l- .l- 1
None None Split None Split Split
T None
Donble slotted slotted Fowler Double 1---1 for· 0 0 0 50 5° 1.0 1.0 2.5 2. 2.5 2. 0.0° swept 65(216)-215, straight straight straight straight 65(216)-215, 66,2-11S 66,2-11S 66,2-118 66,2-118 65(223)-221, 65(216)-215, straight 65(216)-215, line 66(215)-316, 66(2x15)-1l6 66(2x15)-1l6 66(2xI5)-116 66(2x15)-1l6 line line line washout, washout, washout, washout, washout, washout, line washout, line ACA NACA NACA NACA NACA NACA N NACA chord chord chord chord NACA NACA NACA NACA NACA NACA NACA chord =0.8 chord =9.08 a=0.5 ward3.5° a=1.0 a=0.6 a a=0.5 a=O.S Geometric Root: Root: Tip: Geometric Root: Tip: A=9.08 Geometric Root: A=6.25 Geometric Root: Tip: A=6.25 Geometric Root: Tip: A=6.25 ).=0.35 Geometric 0.375 Root: Tip: A=6.1 1<=0.47 0.375 Tip: A=12.S 1<=0.33 Geometric A ).=0.45 1.1 ),=0.45 0.20 Tip: 1<=0.35 0.375 1<=0.35 0.375 Sections: Sections: Sections: Sections: Sections: Sections: Sections:
~
±-
-w- ~
~
-==
'==:::::!- ~==
--- iE:LJ:t II p c. 0 ~ 10:- CJ
~
$
$
1L9l
.--.~.-
~ ~ ~\~
~
~
~
removed.
Fillets XX XIX XXI XXV • XXII XXIV XXIII ---I ---, ---, ---I '---I /---/ SUMMARY OF AIRFOIL DATA
SUPPLEMENTARY DATA
I-BASIC THICKNESS FORMS Page Page N AC A 0006 ___________________________________________ ~ _ 70 NACA 64 -012 _________________________________________ _ N ACA 0008 ____________________________________________ _ 70 N ACA 6~-015 _________________________________________ _ N AC A 0009 ____________________________________________ _ 70 NACA 64 -018 _______________________________ c _________ _ N ACA 0010 ____________________________________________ _ 71 N ACA 64 -02L ________________________________________ _ N ACA 0012 ____________________________________________ _ 71 N ACA 65,2-016 _________________________________________ _ NACA 0015 ____________________________________________ _ 71 N ACA 65,2-02L _______________________________________ _ N ACA 001.8- ___________________________________________ _ 72 N ACA 65,3-018_________________________________________ _ N ACA 0021 ____________________________________________ _ 72 N ACA 65-006 ________________ . ________________ .. ________ _ N ACA 0024 ____________________________________________ _ 72 N ACA 65-008 __________________________________________ _ 73 N ACA 65-009 __________________________________________ _ NACA 16-006 ____ - ______________ - - ____ - - _- - - - - _- - - _- _- -- NACA 16-009 __________________________________________ _ 73 NACA 65-010 __________________________________________ _ N ACA 16-012 ____ - _____________________________________ _ 73 NACA 65 -012 _________________________________________ _ 74 NACA 65 -015 _________________________________________ _ NACA 16-015 __________ - __ - - -" _- - - _- - - -- - - - - - - - - - - - - - - -- 83 N ACA 16-018__________________________________________ _ 74 N ACA 65 -018 _______________________________________ - __ 74 N A C A 65c02 L ________________________________________ _ N ACA 16-021- ________________________ - - _- _- - ___ - - - - _- -- N ACA 63,4-020_________________________________________ _ 75 N ACA 66,1-012-- _______________________________________ _ 75 N ACA 66,2--015--___________ - - ____ - ____ - - ___ - - __________ _ NACA 63-006 ____ - _______ - - _- - _- - - _- _- - - - - - - - - - - - - - - - - -- NACA 63-009 __________________________________________ _ 75 N ACA 66,2-01L _____________________ .__________________ _ 76 NACA 66-006 __________________________________________ _ N ACA 63-010 __________________________________________ _ 76 N ACA 66-008 __________________________________________ _ N ACA 63 -012 _____________________________ - - __ - _____ - -- N ACA 632-015 _________________________________________ _ 76 N A C A 66-009 __________________________________________ _ 77 N ACA 66-010_ _____________ ______________ __ ___________ _ N ACA 63 -018 _________________ - _- _- _- - - _- _- - - - - - - - - - - -- 86 77 N ACA 66 -012 _________________________________________ _ N ACA 63 021 ____________ - _- __ - - - _- _- - - _- _- - - - - - - ___ - -- c 77 N ACA 66 -015 _____________________________ - - __ - _- - __ - __ N ACA 64,2-015 _________________ - ________ . ___ - ____ - - ___ -_ 78 N A C A 66 -018 _________________________________________ _ N ACA 64-006 _____________ - _- ___ - __ - __ - - _- - - - __ - - -'. - - - -- 87 78 NACA 66 02L ____________________________ - ___ - _- - __ - __ NACA 64-008 _______________ - ______________ - - ___ " _- __ - -- c NA'CA 64-009 ______ ._________________________ - _____ - __ - -- 78 N A C A 67,1-01 L _______________________________________ _ NACA 64-010 .. ____________________________________ - __ - __ 79 NACA 747 AOI5 _________________________________ - _- - ____ _ REPORT NO. 824--NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 2.0 NACA 0006 BASIC THICKNESS FORM X vlV AD./V
(DIV)' I
(percent c) (percent c) 1.6 0 0 3.992 0 0 .li .938 2.015 .880 ----.-947--- 1.25 1.117 1.057 1.364 2.5 1.307 1.186 1.089 .984 5.0 1.777 1..217 1.103 .696 1.2 7.5 2.100 1. 225 1.107 .562
:--- 10 2.341
1.212 1.101 .478
-
r--
15 2.673 1.098 .378 1. 206
I--
::--- 20 1. 091 .316
- 2.869 1.190
(-r;)'
25 2.971 1.179 1.086 .272 1.162 1.078 .239 30 3.001 40 2.902 1.136 1.066 .189
~
50 2.647 1.109 1.053 .162 .8 60 2.282 1.086 1.042 .123 70 1.832 1. 057 1.028 .097 80 1.312 1. 026 1.013 .073 NACA 0006 90 .724 .980 .990 .047 95 .403 .949 .974 .032 100 .063 0 0 0 .4 L. E. radius: 0.40 percent c r---
o
NACA 0008 BASIC THlOK;);ESS FORM x 11 1.6 (vi V)' vlV d"./F (percent 0) (percent 0)
---
0 0 0 2.900 .5 .792 .890 1. 795 ---i:263--- 1.25 1.103 1. 050 1.310 2.5 1.743 1.221 1.105 .971
I' r--
1.2 5.0 2.369 1.272 1.128 .694 I ~ 2.800 1.284 1.133 .561 7.5
r-- 10 3.121 1. 277 1.130 .479
--
I--
15 3.564 1.272 1.128 .379 20 3.825 1. 259 1.122 .318 25 3.961 1. 241 1.114 .273 -, 1.106 .239 30 4.001 1.223 1.089 .188 40 3.869 1.186 .8 3.529 1.149 1. 072 .152 1.054 .121 60 3.043 1.111 2.443 1.080 1. 039 .096 NACA 0008 1.017 .071 80 1.749 1.034 .965 .968 .984 .017 .537 .939 .969 .031 .084 0 100 ... -------- -------.-.- .4 L. E. radius: 0.70 percent c V- .- I'-..
o
NACA 0009 BASIC THlCKXESS FORM 1.6 x y (t>fV)' vlV Av.IV (percent c) (percent c)
---- ----- -----
0 0 0 0 0.595 .5 .750 .866 1. 700 --T42Q--- ( 1. 25 1.083 1.041 1.283 1.2 2.5 1. 961 1.229 1.109 .963
--
~
-- 5.0 2.666 1. 299 1.140 .692
r--
7.5 3.150 1. 310 1.145 .560
:---
t-- 10 3.512 '1. 309 1.144 .479 15 4.009 1. 304 1.142 .380 20 4.303 1. 293 1.137 .318 25 4.456 1.275 1.129 .273
'\
30 4.501 1.252 1.119 .239 .8 40 4.352 1.209 1.100 .188 .151 50 3.971 1.170 1.082
\
.120 .
60 3.423 1.126 1.061 NACA 0009 70 2.748 1.087 1.043 .095 80 1. 967 1.037 1.018 .070 90 1.086 .984 .982 .046 95 .605 .933 .966 .030 .4 100 .095 0 0 0
V-
L. E. radius: 0.89 percent c
I'---
o .6
1.0 SUMMARY OF AIRFOIL DATA.
2.0 NACA 0010 BASIC THICKNESS FORM
--
x (v/V)2 v/V I:1v.jV (per~nt c) (percent c) 1.6 2.372 0 0 0 0 .844 1.618 .5 .712 ~---------- 1.061 1.0ao 1. 255 1.25 1. 578 1.112
r-- 2.5 2.178 1. 237 .955
( 1.151 .690
r-- 5.0 2.962 1.325
-- 7.5 3.500 1. 341 1.158 .559
10 3.902 1.341 1.158 .479
r---
1.158 15 4.455 1.341 .380
---
...... 1.153 .318
4.782 1. 329
('I))' 20
........
25 4.952 1. 309 1.144 .273
;\7
1.133
-- 30 1.284 .239
5·002 1.112 .188 40 4.837 1. 237
'\
.8 4.412 1.190 1.091 .150 60 3.803 1.138 1.067 .119 \ 1.046 .094 70 3.053 1.094 1.020 .069 80 2.187 1.040 .045 NACA 0010 90 1.207 .960 .980 .962 95 .672 .925 .030 100 .105 ----------- ----------- .4 L. E. radius: 1.10 percent c
V-
I'---
o
NACA 0012 BASIC THICKNESS FORM x 1.6 (vi V)' vlV I:1v.IV (perc~nt c) (percent c)
-----
1.988 0 0 0 0 .800 1.475 .5 .640 ---i:S94---
r-- 1.199
1.25 1. 010 1.005
( ---
~ 1.114 .934 2.5 2.615 1. 241 1.2 1.174 .685 5.0 3.555 1. 378
...... r---
1.184 .558 7.5 4.200 1. 402 1.188 .479 10 4.683. 1.411
r--
1.411 1.188 .381 15 5.345 ......... 1. 399 1.183 .319 20 5.737 1. 378 1.174 .273 25 5.941 6.002 1. 350 1.162 .~9
1\
1.135 .187 40 5.803 1.288 .8 1.108 .149 50 5.294 1.228 1.166 1.080 .118 60 4.563 1. 053 .092 70 3.664 1.109 NACA 0012 1.022 .068 80 2.623 1.044 .978 .044 90 1.448 .956 .952 .029 95 .807 .906 100 .126 0 0 .4
,,---
L. E. radius: 1.58 percent c c--=:-- .
r'--
o
NACA 0015 BASIC THICKNESS FORM 1.6 x u (vIV) , vlV I:1v.IV (percent c) (percent c) / r-......
----
0 0 1.600
-- 0 0
~
.546 .739 1.312 .5
I
---2:367---
r--...... .933 .966 1.112
1.25 I.~ 1. 237 1.112 .900 2.5· 3.268 ~ 1.450 1.204 .675 5.0 4.443 1.498 1. 224 .557
~ 7.5 5.250
......... 1. 520 1. 233 .479 10 5.853 1. 520 1.233 .381 15 6.682
(ilt
1. 229 .320 20 7.172 1.510
'1\ 1.484 1.218 .274
25 7.427 1. 450 1.204 .239 30 7.502 .8 1. 369 1.170 .185 40 7.254 1. 279 1.131 .146 50 6.617
\
1. 206 1.098 .115 60 5.704 NAtA 0015 4.580 1.132 1.064 .090 3.279 1.049 1.024 .065 1. 810 .945 .972 .041 .872 .934 .027 95 1. 008 .158 0 0 100 0 I
-
V
r---
L. E. radius: 2.48 percent c
-
-
-
"---
/.0 .6 .8 o .2
.:rIc
REPORT NO. 824-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 2.0 NACA 0018 BASIC THICKNESS FORM .~ y c---:-
(v/VP I v/v I flv.iF
1.6 (percent 0) (percent 0) -.....
,
---------- --- .. --
( ~
0 0 0 0 1.342 .5 .465 .682 1.178 -----------
'"
1.25 2.841 .857 .926 1.028
""-
2.5 3.922. 1. 217 1. 103 .86l ~
I
5.0 5.332 1. 507 1.228 .662
t---
1.2
-..... 7.5 6.300 1. 598 1. 264 .55fi
10 7.024 1.628 1.276 .4i9
i
15 8.018 1. 633 1. 278 .381
~
!
I"--.. 20 8.606 1. 625 1. 275 .320 25 8.912 1. 592 1. 262 .274 30 9.003 1. 556 1. 247 .238
""'" 40 8.705 1.453 1. 205 .184
1\
.8 50 7.941 1.331 1. 15·i .144 60 6.845 1. 246 1.116 .113 70 5.496 1.153 1. 074 .087
\'
80 3.935 1. 051 1. 025 .063 NACA 0018 90 2.172 .933 .966 .039 1.210 .836 .914 .025 100 .189 0 0 0 .4
,..--
i ;--- L. E. radius: 3.56 percent c
t--
:/
r--
:
t---
t-- i-'-'
l--
r---- I--------
r--
"'--
o NACA 0021 BASIC THICKNESS FORM
-
~
I
J x
/.6 (vIF)' v/V 1l.v(dV' (percent c) (percent c) I
""'"
,
"""
I
~ I
0 0 0 0 1.167 .5 .. _--------- .397 .630 1. 065 I
~
1.25 3.315 .787 .887 .946 2.5 4.576 1.182 1. 087 .818 ~ 1.2 5 6.221 1. 543 1. 242 .648 7.5 7.350 1.682 1. 297 .550 10 8.195 1. 734 1. 317 .478
~
15 9.354 1. 756 1.325 .381 20 10.040 1. 742 1. 320 .320
'"
25 10.397 1. 706 1. 306 .274 30 10.504 1.664 1.290 .2.~8
1\
40 10 . .156 1..538 1. 240 .183
""
.8 .10 9.265 1. 388 1.178 .142 60 7.986 1. 284 1. 133 .111 70 6.412 1.177 1.085 .084
\
80 4.591 1. 055 1. 027 .061 NACA 0021 90 2.534 .916 .957 .037 95 1.412 .801 .895 .023 100 .221 0 0 0 .4 ;
-
- t--
'/
'---- L. E. radius: 4.85 percent c
r--
-- e----
l--
~
!--
--
o
~ -........
NACA 0024 BASIC THICKNESS FORM
( ~i
/.6 x y (V/V)2 v/V fll'./V (percent c) (percent c)
I
~r--- ----
I
0 0 0 1. 050 J
""- ________ w __
.5 .335 .579 .964 , 1. 25 3.788 .719 .848 .870
~
2.5 5.229 1.130 1. 063 .771 , 1.2
-'
5.0 7.109 1.548 1. 244 . 6.~2 !
7.5 8.400 1. 748 1.322 . .142
~
10 9.365 1. 833 1. 354 .470 r--...
15 10.691 1.88S 1. 374 .383 20 11.475 1.871 1. 368 .321 25 11.883 1. 822 1. 350 .274 30 12.004 1.777 1. 333 .238
~
'"
.8 40 11.607 1. 631 1. 277 .181 50 10.588 1. 450 1. 204 .140 60 9.127 1.325 1.151 .109
\ 70 7.328 1.203 1. 097 .082
NACA OON 80 5.247 1. 065 1. 032 .059 \ I 90 2.896 .891 .944 .035 95 1. 613 .773 .879 .022 100 .252 0 0 0 .4
1/
-
'---- L. E. radius: 6.33 percent c
r--
---
~ ~
f-.--
~
~
.C .6 .8 1.0 o
.'T zlc SUMMARY OF AIRFOIL DATA NACA 16-006 BASIC THICKNESS FORM /. 6 x 1/ Av./F (vIF)' vlF (pCrcent c) (percent c) 0 5.471 0 0 0 1. 25 .646 1.059 1. 029 1. 376 2.5 .903 1. 085· 1. 042 .980 5.0 1. 255 1. 097 1.047 .689
/. e
7.5 1. 516 1.105 1. 051 .557 1.729 1.108 1.053 .476 r
-
-----.. 15 2.067 1.112 1. 055 .379
20 2.332 1.116 1.057 .319
(v)'
2.709 1.123 1.060 .244 2.927 1.132 1. 054 .196 50 3.000 1.137 1. 066 .160
\
60 2.917 1.141 1. 068 .130 70 2.635 1.132 1. 064 .104 80 2.099 1.104 1. 051 .077 .049
\ 90 1. 259 1. 035 1. 017
NACA 16-006 05 .707 .962 .981 .032 100 .060 0 0 0 L. E. radius: 0.176 percent c NACA 16-009 BASIC THICKNESS FORM I.
'V (vi F) , vjF Av./F (percent c) 1 (perc;nt c) 1---- --------1----- -------1------ o o 3.644 o o 1. 25 .969 1. 042 1. 021 1. 330 2.5 1. 354 1.109 1. 053 .964 1.2 5.0 1. 882 1.139 1. 067 .684
--
,--
.5M 7.5 2.274 1.152 1. 073
~ .475
10 2.593 1.158 1. 076 .378 15 . 3.101 1.168 1. 081 .319 20 3.498 1.177 1.085 .245 80 4.063 1.190 1. 091 ,197 40 4.391 1. 202 1. 096
\
.8 1.100 .160 50 4.500 1.211 .131 60 4.376 1.214 1.106
\ .103
70 3.952 1. 206 1. 099 .076 3.149 1.156 1. 075 NACA 16-009 1. 022 .047 90 1. 888 1.043
80 I
.030 95 1. 061 .939 .969 o o 100 .090 o .4 ---------------------------------- ' L. E. radius: 0.3g6-percent c r--- t...--
-
J---
---
o NACA 16-012 BASIC THICKNESS FORM u; y x (vjF) , vlF "-v.iF (percent c) (percent c) o 2.624 o o o ~ 1. 292 1.002 1.001 1. 268 I.e 1. 25
-
1.109 1. 053 .942 2.5 1. 805
/
1.173 1.083 .677 5.0 2.509
~
3.032 1.197 1. 094 .1)51 7.5
(v)" 1. 208 1.099 .473
10 3.457 1. 223 1.106 .378 15 4.135 1. 237 1.112 .319 20 4.664
\ 1. 257 1.121 .245
30 '5.417 .8 1.271 1.128 .197 40 5.855 1. 286 1.134 .161 50 6.000 1. 293 1.137 .131
\ 60 5.835
5.269 1. 275 1.129 .102 NACA 16-012 4.199 1. 203 1. 097 .075 1.051 1. 025 .045 90 2.517 .908 .953 .027 95 1. 415 o o .4 100 .120 o
----------' -----..!.----~--.--------
I- L. E. radius: 0.703 perccnt c ---'
---
-- -
o .2
.4 .6 .8 .rIc _ REPOR'!' NO. 824-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS '20 NACA 16-015 BASIC THICKNESS FORM /.6 x (vIV)' t'/V tJ.v.IV (percent c) (percent c) --' 0 0 0 0 2.041 1. 25 1. 615 .956 .978 1.209 I--- L.--- 2.5 2.257 1.105 1.051
~t-.. .916
./ 5.0 3.137 1. 200 1.095 .668 1.2 7.5 3.790 1.239 1.113 .547
/ 10 4. 322 1. 256 1.121
.471
'\
15 5.168 1. 278 1.130 .377 .- 20 5,830 1. 297 1.139 .318
(vt
6.772 30 1. 327 1.152 .245 40 7.318 1. 349 1,161 .197 7.500
\ 50 1. 364 1.168 .161
.8 7.293 60 1.374 1.172 .131 70 6.587 1. 348 1.161 .102
'\
80 5.248 1. 254 1.120 .074 90 3.147 1,053 1. 026 .043 , NACA 16-015 95 1. 768 .875 ,935 .025 100 .150 0 0 0 .4 L. E. radius: 1.100 percent c ~
r---
V- r-....
V-
~
I--. L---
o
NACA 16-018 BASIC THICKNESS FORM /.6 x (v/V) , vlV tJ.v.IV (percent c) (percent c) ,...- 0 0 0 0 1. 744
V-
1. 25 1. 938 .903 .950 1.140
~
/' 2.5 2.708
1.09,2 1. 045 .883 /,E
"""
5.0 3.764 1.217 1.103 .657 7.5 4.548 1.271 1.128 .541
\
(
5,186 10 1. 302 1.141 .468 , 15 6.202 1. 332 1.154 .376 20 6.996 1. 357 , 1.165 .318 30 8.126 1. 399 1.183 .245 40 8.782 1. 426 1.194 .198
\
50 9.000 1,447 1.203 .162 .8 8.752 60 1. 452 1. 205 .131 70 7.904 1. 421 1.192 . .102 80 6.298 1. 306 1.143 .073
\
NACA 16-018 90 3.776 1. 051 1. 025 .042 95 2.122 .837 .915 .024 lop .180 0 0 0 .4 L. E. radius: 1.584 percent c
I--
1--
-
V
r--..
V--
'-- ~
1--- .I--
-
o
NACA 16-021 BASIC THICKNESS FORM 1.6
I---r-
x (o/V) , vlV tJ.v.IV
~
!---- (percent c) (percent c)
V 0 0 0 0 1. 574
I 1\
1.25 2.261 .826 .909 1. 069 1.2 2.5 3.159 1. 062 1. 031 .828 5,0 4.391 1.221 1.105 .640
'\
7.5 5.306 1. 295 1.138 .534 10 6,050 1. 342 1.159 .463 15 7.236 1. 391 1.179 .374 8.162 20 1. 419 1.191 .317 30 9.480 1. 474 1. 214 .245
1\
10.246 .8 40 1. 506 1.227 .198 50 10.500 1. 535 1. 239 .162 lO,211 60 1. 536 1.239 .131
\ 9.221 1. 495
70 1. 223 .102 NACA 16-0EI , 80 7.348 1. 361 1.166 .072 90 4.405 1.039 1.019 .041 2.476 .801 95 .895 .023
\
100 .210 0 0 0
'" L.---
- i--
L. E. radius: 2.156 percent c
v-
r----...
V
f"--
!---
r-- I-
1---: .2 .4 .6 .8 1.0 ~/c
SUMMARY OF AIRFOIL DATA 75
2. NACA 63,4-020 BASIC THICKNESS FORM .0
I I I I
x 1/ •••• .c. ".44 (upper surface) (VfV)2 vfV Av.IV
~ (percent c) (percent c)
0 1.395 0 0 .666' 1.280
V ~ ~ .444
.5 1. 714 I.B .605 .778 1.201 .75 2.081 1. 072 .820 .906 1.25 2.638
Va
.846 1.080 1.039 ·2.5 3.606
~
~
1.130 .645 1. 277 5.0 4.947 -.;;;;: 1.176 .543 1. 383 7.5 5.964 1.207 .475 1. 456 10 6.800
V
1.245 .386 1.551 15 8.090
~
I /
1.270 .330 1. 614 20 9.006
~""
1.288 .289· 9.630' 1. 659
.2 I -;44 fower
surface) 1. 300 .257 9.955 1. 689
V- 30
~
1. 277 .219 1.630
~ 35 9.978
1.252 .192 9.765 1.567 -.......:: 1. 225 .169.
9.366 1.500 1.197 .148 1.433 50 8.819
~ 1.167 .128
/ 8.143 1. 362
.8 1.135 •• 112 7.351 1. 288 1. 213 1.101 .097 6.464 1. 066 .084 5.496 1.137
/
1. 059 1.029 .071 4.466 NAOA 63,4-020 75 .978 .989 .059 80 3.401 ;046 .896 .947 2.342 85 ,
"
.811 .901 .036 1.348
AI/ 90
.853 .023 .728 .501 O· .651 .807 I-- 0
- t----
r--
V
r--
L. E. radius: 3.16 percent c ~ ~ ~
'--
L---
o r--
NACA63-006 BASIC THICKNESS FORM x (vi V) , .IV A •• IV (per~nt c) (percent c) 0 4.483 I .0 0 2.110 .973 .986 .5 .503 1.025 1.778 1.050 .75 .609 1.039 1. 399 1. 080 1. 25 .771 1.054 .981 1.110 1.057 2.5 1. 063 .692 1.130 5 1 .. 462 surface) _.cz =.03 (upper ,,0 1.069 .562 1.142 7.5 1.766 ~-, " I 1.072 .484 1.149 10 2.010 2 I.
1.077 .384 1.159 2.386 1. 079 .321 ["" I 1.165 2.656 ·20 1.082 .279 '1.170 -:03 (lower surface) ~ 25 2.841
roo-
ff?' .245 ..
1.084 1.174 '2.954
.......... 1.082 .218
1.170 3.000 1.079 .196
r---. 1.164
40 2.971 1.073 .176 1.151 45 2.877 1.066 .158 1.137 B 50 2.723 .141 1.·057 2.517 1.118 1.047 .125 1.096 2.267 1.036 .111 1. 074 65 1.982 NACA 63-006 1.046 1. 023 .098 1.670 .085 1.020 1.010 1. 342 .073 .. 994 .997 80 1. 008 .060 .965 .982 .683 4 .047 .936. .967 .383 .032 .910 .954 .138 .886 .941 i..-- ...;;;;.: L. E. radius: 0.297 percent c NACA 63-009 BASIC T1HCKNESS FORM x vlV Av.IV (·fV)' (perc~nt c) (percent c) 1.6 3:058 0 0 c, =.08 (upper surface) 1. 889 .941 .749 .885 ",,,, ... .5 1.647 1.002 1.001 .906 , .75 , .1.339 1.051 1. 025 1.151 1.25 .961 1.130 1.063 1.582 2.5 .689 1 0 ... 1.180 1.086 2.196
~ 5
.560 .
1.2 1.205 1.098 2.655 7.5 .484
- V -..;;::: 1. 221 1.105
3.024 .386 1. 241 1.114 3.591
'I;<
~ ---.08 (lower. surface).
.324 1.255 1.120 3.997 ~ 20 .281 1.124 4.275 1.264
(vt
I .248 1.269 1.126 4.442
"'-
.220 1. 265 1.125 4:500 .196 1. 255 1.120 4.447 .8 .175 1.235 1.111 4.296 .156 1.208 1.099 4.056
"" 50
.140 1.175 1.084 3.739 1.068 .124 1.141 3.358 NACA 8:1"'009 60 1.051 .109 1.104 2.928 1.032 .095 1.065 2.458 1.012 .082 1.025 '1.966 :992 .069 .984 .4 1.471 .971 .057 .942 .990 .044 .950.
.903 .550 .• 030' .932 .868 .196 I..---: 95 .915 0 .838 ~ . L. ,E. radius: 0.631 percent c .8 /.0
o .2 .4 .6
.¥<Ic REPORT NO. 824-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS NACA 63-010 BASIC THICKNESS FORM z '1/ (VIV)2 vfV "'v.IV (peJ'cent c)
----- I-- .. --. -- I (percent c)
2. .. 6°= ~[TII ------ ------'
---- ----- ------ 0 2.775 0 0 0
,- L.) .. ~j- .--~ .. -.-----.-
.841 .917 1. 825
1 .5 .829
.989 1. 603 .75 1.004 .978 /' .c, =.10 (upper surface) 1. 037 1.018 1.316 1. 25 1. 275 1.131 1.063 .952 2.5 1. 756 , 1.092 .687 5.0 2.440 1.193 1.10() .500 2.950 1.223 7.5 1.116 .484 1 10 3.362 1.245 2 (°'>-- --~::::::::.....
1. 270 1.127 .386 15 3.994 1.13~ .325 20 4.445 1. 285
I. '(/.~-.!o 10wer surfacel--":::::::::~~· --.------- ... - .2S2
25 4.753 1. 295 1. 138 .218 4.938 1.302 1.141
~t •. /)' I I' ) ~ .220
35 5.000 1. 299 1.140 .196 40 4.938 1. 286 1.134 .175 45 4.7GG 1.2G2 1.123
\~- 1/ ~-~L- ~r' ~~
.156 50 4.496 1. 231 1.110 .139 4.140 1.193 1.092 8 r - I .123 60 3.715 1.154 1. 074 .108 3.234 1.113 1.055 .094 2.712 1. 069 1.034 1---+-----1----1--- NACA 83-010··-+---1---1----1 1. 012 .081 75 2.166 1. 025 .069 80 1. 618 .979 .989 .056 1.088 .935 .967 .945 .043
I 90 .604 .893
.4r--+~4--~_4--!~_+-~-~_+-~ .924 .030 95 .214 .853 .907 0 0 .822 L. E. radius: 0.770 percent c Or-~--_+---L·-~---L--~--~--+_--~~ NACA 63,-012 BASIC THICKNESS FORM Z '1/ (VIV)2 vfV ",v.IV
I
(percent c) (percent c) -.---.
I.B 0 2.336 0 0 0 .750 .866 1.695 .5 .985
_--------C', :./4 (upper surface)
.925 .962 1. 513 .75 1.194 1. 266 1.005 1.003 1. 25 1.519 1.129 1.063 .933 2.5 2.102
-
1. 217 1.103 .682
V- -t3::=: 5 2.925
1. 261 1.123 .559
0/ 7.5 3.542
~
1. 294 1.138 .484
---= 10 4.039
U?
--- ---
1.330 1.153 .387 15 4.799 K:', .326 5.342 1. 349 1.161
surface)-........:::~
"-.14 fower
1/ .283
1. 362 1.167 25 5.712
~
.249 1.370 1.170 30 5.930 --..::::
(vl
, .221
1. 366 1.169 35 6.000 .195
1/ 1. 348 1.161
40 5.920 .174
.......... 1. 317 1.148
45 5.704 .8 .155 1. 276 1.130 50 5.370 .137 1. 229 1.109 55 4.935 1.181 1.087 .121 60 4.420 1.131 1.063 .106 65 3.840 NACA 63,-012 1.037 .091 70 3.210 1.076 1. 011 .079 75 2.556 1. 023 .969 .984 .067 80 1. 902 .959 .055 85 1. 274 .920 .4 .933 .042 90 .707 .871 .909 .029 95 .250 .826
r.--- -r'-- .889 0
100 0 .791 I--- -------------~----------------
'--'---
L. E. radius: 1.087 percent c
-
NACA 632-015 BASIC THICKNESS FORM x 'II __ .c, =.22 (upper surfaoe) (V/V)2 vlV ",v.IV (percent c) (percent c)
I
-
._._- !
/.8 .........
f 1. 918 0 0 0 0
..-- --. 1. 513
. 5 1.204 · 600 · 775
......--
1---- .. . -'---.- 1. 379 -- . 75 1. 462 · 822 · 907 1.182 · 969 1. 25 1. 878 · 938 .903 1.105 1. 051 2.5 2.610
---------r-=:~
.674 1. 244 1.115 5 3.648
I
/ -_.-. ~ ~ 1.2 .557 1. 315 1.147 7.5 4.427 .484 1.360 1.166 10 5.055
/---.22 fower surface) ~ .388
1.415 1.190 15 6.011
i/ .330
1. 446 1.202
~ 6.693
1.467 1.211 .286
(vi 25 7.155
~ .251 1. 481 1. 217 30 7.421 .222 1.475 1. 214 35 7.5CO
~
'&. .196
I 1. 446 1. 202
40 7.386 .8 .174 1.401 1.184 ......... 45 7.099 .153 1. 345 1.160 50 6.665 1.281 1.132 .135 6.108 .118 1. 220 1.105 5.453 .102 1.155 1.075 65 4.721 1.042 .088 3.934 1.0.85 1.019 1.009 .076 3.119 .4 .976 .063 2.310 .953 .946 .051 1. 541 .894 ~ .916 .039 .852 .839
--1----1--
/-
.888 .026 .300 .789 .866 0 .750 100 0
I
r-..
'-
L. E. radius: 1.594 percent c .2 .4 .8 .8 1.0 .rIc
SUMMARY OF AIRFOIL DATA 77
2.0 NACA 633-018 BASIC THICKNESS FORM /,c, ~.32 (upper surfaco) , x y (vfV) , vfV tl.v./V (percent c) (percent c)
I
I I
I
------ ~
(
0 0 1.639 0 0 1.6 .5 1. 404 .441 .664 1. 361 -........
...--
.75 1. 713 .700 .837 1. 258 1. 25 2.217 .848 .921 1.105 C(
V
"" ~ 2.5 3.104 1. 065 1. 032 .871
5 4.362 1. 260 1.122 .663
~ 7.5 5.308 1.360 1.166 .553
/'
6.068 1. 424 1.193 .484
~ ~
"" I /
15 7.225 1. 500 1. 225 .390 1.2 20 8.048 1. 547 1. 244 .333 25 8.600 1. 579 1. 257 .289
V'"
'.32 dower surrace)
~ 30 8.913 1. 598 1. 264 .253
~
I /
35 9.000 1. 585 1. 259 .223
(~r
40 8. 845 1. 550 1. 245 .197
~
45 8.482 1. 490 1. 221 .173 50 7.942 1.411 1.188 .152
I ~
.8 55 7.256 1.330 1.153 .133 60 6.455 1. 252 1.119 .115 65 5.567 1.170 1. 082 .099
I 4.622 1. 087 1.043 .084
NACA 63.-018
" 75 3.650 1. 009 1.004 .072
80 2.691 .933 .966 .059 85 1. 787 .868 .932 .048
I
90 .985 .807 .898 .036 .4 95 .348 .753 .868 .024
f.--
r-- .712 .844 0
- 100 0
t----
V
r--
- L. E. radius: 2.120 percent c
I---
,~ l..--- t--
-
o
NACA 63,-021 BASIC THICKNESS FORM
~
/
I y x d •• /V (II/V)' v/V (percent c) (percent c)
/c, = .38 (upper surface)
~
,,"\ ~
I
1.6 1. 439 o o o o .5 .275 .524 1. 236
V 1. 583
a
.--
.75 .564 .751 1.156 1. 937 ..........
~ ''/ ./
1. 25 2.527 .725 .851 1.034 2.5 3.577 1.010 1.005 .842 V
""
'"
5 5.065 1. 260 1.122 .653 7.5 6.182 1.394 1.181 .550
~ ~
I /
1.2 10 7.080 1. 487 1. 219 .484 15 8.441 1. 592 1. 262 .392 dower surface)
II'" -.38
20 9.410 1.655 1. 286 .335
~
~
I 25 10.053 1.698 1.303 .291
30 10.412 1. 721 1. 312 .255
(v)
35 10.500 1. 709 1.307 .225
~
1.654 .198 40 10.298 1. 286
~
/ .173
45 9.854 1. 578 1. 256 ,- .8 1.479 1. 216 .150 no 9.206 1.380 1.175 .130 55 8.390
~
7.441 1. 281 1.1~2 .1l2
I
.096 NACA 63 -021 65 6.396 1.180 1.086 1.084 1. 041 .081 70 5.290 75 4.160 .994 .997 .068 .911 .954 .057 80 3.054
I
.4 .839 .916 .046 85 2.021 ---::: V- 90 1.113 .774 .880 .035
~ 95 .392 .721 .849 .023
V r--- .676 .822 o
100 o
--
j..--
I---
L. E. radiu.: 2.650 percent c
~ ~
t---.. L.---
o
NACA 64,2-015 BASIC THICKNESS FORM __ c =.20 (upper surface) z (v/l')' v/v tlv./v
x I y
, . (percent c) (percent c) , 1.0 ----- --------- -------- ------ ------- 1. 930 0 0 0 0 ~ .710 .843
~~ .5 1.216 1500
.825 .908 1. 359 .75 1.453
-
~ .962 .981 1.161 ~ 1. 25 1. ~29
~
2_ 538 1. 122 1.059 .911 2.5
V
~ ~
/ 1. 234 1.111 .678
/ 5.0 3.514 1.2 4.243 1. 288 1.135 .553 7.5
/'" 1. 323 1. 150 .477
-.20 fower surface) 10 4.838
~
1.371 1.171 .383 15 5.781
/ / ~
I 1.401 1.184 .325 20 6.464 .285 25 6.967 1. 422 1.192 ..
1. 441 1. 200 .253 30 7.307
~
1. 458 1. 207 .227 35 7.481
I
.8 1.471 .202 40 7.480 I. 213 1.432 1.197 .175 45 7.268
"
'"
I 6.850 1. 366 1.169 .156
6.311 1. 299 1.140 ; 137 NACA 64,2-015 5.670 1. 234 1.111 .122 4.944 1.16R 1. 081 .102 4.158 1.102 1.050 .086 3.338 1.019 .080 75 1·039 .4 2.506 .973 .986 .071 .g54 85 1. 698 .910 .056 .039 90 .961 .849 .921
- :--
-
V-
.027 95 .351 .791 .889
-
~ . 100 0 .739 .860 0 l---
'--
-
-
L. E. radius: 1.65 percent c .8 1.0 o .2 .4 .8 HEPORT NO. 824-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS NACA 64-006 BASIC THICKNESS FORM x (ofV) , ofV t;.o./V (pertent c) (percent c) 4.623 0 0 0 0 1.6 .5 .494 .995 .997 2.175 .75 1.058 1. 780 .596 1.029 1.25 .754 1.085 1.042 1.418 2.5 1.024 1.108 1.053 .982 5.0 1.405 1.119 1. 058 .692 7.5 1.692 1.128 1.062 .560 .' .. --- (:, ... 02 (upper surface) 10 1. 928 1.134 1.065 .483 ,,0 ..
15 2.298 1.146 1.071 .385 1.2 , 20 2.572 1.154 1. 074 .321 I ~ ...... ~ .. 25 2.772 1.160 1.077 .279 --.02 (lower surrace) 30 2.907 1.164 1. 079 .246
V
35 2,981 1.168 1.081 .220
(ill
-- 40 2.995 1.171 1.082 .198
---
45 2.919 1.160 1.077 .178 r--.
--
50 2.775 1.143 1. 069 .158 .8 55 2.575 1.124 1.060 .142 '50 2.331 1.102 1. 050 .126 65 2.050 1.079 1.039 .112 70 1.740 1.0M 1.027 .098 NACA 84-008 75 1. 412 1.028 1. 014 .085 80 1.072 1.000 1.000 .072 85 .737 .970 .985 .060 .939 90 .423 .969 .047 .4 95 .157 .908 .953 .031 100 0 .876 .936 0 L. E. radius: 0.256 percent c
o
NACA 64-008 BASIC THICKNESS FORM x (vfV) , t;.o./V v/V (percent c) (percent c) 1.6 0 0 0 0 3.M4 .5 .658 .912 .955 1. 994 ____ c, :.04 (upper surface) .. - ~- .75 .794 1. 016 1. 008 1. 686 1. 25 1. 005 1.084 1. 041 1. 367 2.5 1. 365 1.127 1.062 .969 5.0 1.875 1.152 1.073 .688 .0 , ; 7.5 ,2.259 1.167 1~ 080 .560 10 2.574 1.179 1.086 .480 ~ 15 3.069 1.195 1. 093 .385 ~ 20 3.437 1. 208 1. 099 .323 -----.04 {ower surface
~
f? 25 3.704 1. 217 1.103 .279
~ 30 3.884 1. 225 1.107 .246 35 3.979 1.230 1.109 .220
~
40 3.992 1. 235 1. 111' .198 ............
45 3.883 1. 220 1.105 .176 .8 50 3.684 1.191 1. 091 .158 55 3.411 1.163 1.078 .141 60 3.081 1.133 1. 064 .125 65 2.704 1.102 1. 050 .110 NACA 84~OO8 70 2.291 1.069 1. 034 .096 75 1.854 1.033 1. 016 .083 SO 1.404 .995 .997 .071 .4 85 .961 .957 .978 .059 ,550 .918 .958 .046 95 .206 .878 .937 .031 100 .839 .916 0
--
L. E. radius: 0.455 percent c
o
NACA 64-000 BASIC THICKNESS FORM x (v/V)' t;.v.IV v/V (percent c) (percent c) 1.6
I
3.130 ----c,=.06 (upper surroC'e) 0 0 0 .. --- , , .5 .739 .872 .934 1.905 , .75 .892 .990 .995 1. 637 ,0 1. 25 1.128 1. 075 1.037 1.340 : 2.5 1. 533 1.131 1.063 .963 /7 :::,...
~ 2.109 1.166 1.080 .686 5.0
.J
1.2 2.543 1.186 1.089 .560 7.5 ~- 2.S98 1.200 1. 095 .479
--coe (ower sur~~
3.455 1. 221 1.105 .383
V
~ 3.S68 1. 236 1.112 .323
(vY 1.246 1.116 .281
25 4.170 30, 4.373 1. 255 1.120 .248
I'
4.479 1.262 1.123 .221 ............
-- .8 4.490 1.267 1.126 .198
----+=i-- 40
4.364 1.246 1.116 .176 .158 4.1~6 1. 217 1.103 3.826 1.183 1.088 .140 NACA 84-009 3.452 1.149 1.072 .• 125 3.026 1.112 1.055 .109 .095 2.561 1.073 1.036 2.069 1.033 1.016 .082 '.4 1. 564 .992 .996" .070 1.069 .950 .975 .057 .611 .907 .952 .044 t..---" .865 .930 .030 95 .227 .822 .907 0 100 0
r--
L. E. radius: 0.579 percent c
o
.2 .6 .8 ·1.0
SUMMARY OF AIRFOIL DATA 79
1'. NACA 64-010 BASIC THICKNESS FORM ,. x II (vI V)' vlV Av.IV (percent c) (percent c) 0 0 0 2.815 .913 .5 .820 .834 1.817 /. 6 .75 .989 .962 .981 1. 586 1.25 , 1.250 1.061 1.030 1.313 surface)
,-c "'.08 (upper
z 2.5 1.701 1.130 1.063 .957 2.343 1.181 1. 087 .684 , 5 7.5 2.826 1.206 1.098 '.559 ,480 10 3:221 1.221 1.105 ~ 3.842 1.245 1.116 .386 ZO. 15
--- ~~
I.
K __ 20 4.302 1.262 1.123 .325 -..::::: 4.639 1.275 1.129 .280
~ 1.134
--.08 (lower svrface) 4.864 1.286 .246
~
V
4.980 1.295 1.138 .220 1.140 40 4.988 1.300 .199
f
45 4.843 1.279 1.131 .176
f'
4.586 1.241 1.114 .158 \ 1.096 .139 55 4.238 1.201 1.161 1.077 .124 60 3.820 65 3.345 1.120 1.058 .109
" 1.039 .095
70 2.827 1.080 NACA 64-010 1.036 1.018 .081 75 2.281 .995 .069 80 1.722 .990 .944 .972 .057 85 1.176 .949 .044
.,. 90 .671 .900
.922 .030 95 .248 .850 .805 .897 0 100 0 L.----
-
L. E. radius: 0.720 perCent c
I'--
NACA 641-012 BASIC 'rHICKNESS FORM x (vI V)' V/V t.v./V (percent c) (percent c) , {) -I /.
0 0 0 0 2.379 .5 .978 .750 .866 1. 663 surfaoe) /,c, =./Z (upper .941 , .75 1.179 .885 1. 508 1.25 1.490 1.020 1.010 1.271 2.5 2.035 1.129 1.063 .943 ~ 1.097
~ 5,0 2.810 1.204 .685
-....:
V
1.114 7.5 3.394 1.240 .569
~
~
10 3.871 1. 264 1.124 .482 /.2
--
15 4.620 1.296 1.139 .388
--- K_';12
1.149
~ 20 5.173 1.320 .328
(/ surface)
(lower 1.156
~ 25 5.576 1.338 .281
1.162 30 5.8# 1.351 .247 35 5;978 1.362 1.167 .221
r ~
1.372 1.171 ' .199 40 5.981 ~ 8 ( 1.156 45 5.798 1.335 .177 1.289 1. 136 50 ,5.480 .158 1.243 1.115 .138 55 5.056 1.195 1.093 .122 60 4.548 3.974 1.144 1.070 .103 NAOA 04 -012 1.091 1.044 70 3.350 .088 1.037 1. 018 .074 75 2.695 2.029 .981 .990 .063 1.382 .928 .963 .052 .874 .935 .045 90 .786 .825 .908 .028 95 .288
e--
,--
.775 .880 0 100 0
-
-
--..:...
L. E. radius: 1.040 percent c
o
NACA 64.-015 BASIC THICKNESS FORM x Av./17 (oJV)' vlV (percent c) (percent c)
"
1."6 / A x.c2 (upper surface) 1.939 0 0 0 0 .670 .819 1.476
~ K .5, 1.208
.762 .873 1.354 .75 1.456 ,,/ 1.188
V 1. 25 1.842 .896 .947
--
~
1.113 1.055 .916 2.5 2.028
V-- ~ ~
1.231 1.109 .'670
I- 5.0 3.004
/.2 1;284 1.133 .559 7.5 4.240
/-- 1.150 .482
10 4.842 1.323
I '~ZZ (lower surface)
~
1.172 .389 15 5.785 1.375
~
/ 1.187 .326 20 6.480 1.410 1.198 .285 25 6.985 1.434 1.454 1.206 .250
" 30 7.319
1.213 .225 35 7.482 1.470
I ~~
1.218 .202 .8 40 7.473 1.485 ...........
1.426 1.195 ' .179 45 7.224 '1.365 1.168 .158 6.810 1.140 .135
II 56 6.266 1.300
NACA 64 -015 1.110 .121 2 60 5.620 1.233 1.080 .105 65 4.895 1.167 1.049 .090 70 4.113 1.101 1.016 .078, 75 3.296 1.033 .4 .983 .065 80 2.472 .967 .902 .950 .054 85, 1.677 ~ .841 .917 .041 90 .950 ~ .785 .886 .031
- - - 95 .346
.855 0 ~ 0 .730 ~ '-- r-.-....
-
L. E. radius: 1.590 percent c /.0
o .4 .6 .8
.2
80 REPORT NO. 824-NATIONAL ADVISORY COMMI'fTEE FOR AERONAUTICS
NACA 613-018 BASIC THIOKNESS FOJlM 2.0
(perc~nt c) I (pergent c) I (v/V)' viI' tJ.v.IV
l~c,=.3e (upper surface)
---,---------------. o 0 0 0 1.646
1.6 .5 1.428 .546 .739 1.360 .75 1. 720 .705 .840 1. 269
V
---
1. 25 2. 177 . 862 .920 1. 128
/" 2. 5 3.005 1. 079 1. 039 .904
~ ~
5.0 4.186 1. 244 1.115 .669 '"'-....
--
7.5 5.076 1. 327 1.152 .558
V 10 5. 803 1. 380 1. 17.5 . 486
fr ~ ~ ./
15 6.942 1. 450 1. 204 .391 /.2 20 7.782 1. 497 1. 224 .331 25 8.391 1. 535 1. 239 .288 ----.32 ~ower surface)
v'--
~ ~o 8. 7R9 1. 562 1. 250 .255
~
M &m 1.~ I.. .m
40 8. 952 1. 600 1. 265 . 200 ~ 45 8.630 1.518 1. 232 .177
~ 50 8.114 1. 436 1.198 .154
I
.8 55 7.445 1. 354 1.164 .134 60 6.658 1.272 1.128 .117
~
65 5.782 1.190 1. 091 .102
7 70 4. 842 1. 109 1. 053 . 088
-NACA 64 -0/8 75 3.866 1. 028 1. 014 .071
m ~~ .• .m .~
85 1. 951 .879 .937 .051 .-. 90 l.m .m .~ .• - .4 95 .400 .747 .864 .027 ~ 100 0 .695 .834 0
-
/- r--
--
f--.
-
-
I--
-- I L. E. radius: 2.208 percent c
~ ~
r--
o
-
------
NACA 64.-021 BASIC THICKNESS FORM ( ,
1\
:r y (percent c) (percent c)
1~=-I_~l_'T __ '-_ tJ.".: __
K,:·c,=.44 (upper surface) I
~\
/
1.6
----
0 0
o I 0 1. 458
.5 1.646 . 462 . 680 1. 274 .75 1. 985 . 603 . 776 1. 203
V
"" ~'\ /
~
1. 25 2.517 .759 .871 1.084 2.5 3.485 1.010 1.005 .878 .0;.0 4.871 1. 248 1.117 .665
V
7.5 5.915 1. 358 1.165 . .157
/0
~ ~
1.2 10 6.769 1. 431 1.196 486 15 8.108 1. 527 1.236 .395 --·.44 (lower surface) 20 9.095 1. 593 1. 262
1/, .335
/
~
25 9.m7 1. 654 1. 281
~ .293
30 10.269 1. 681 1. 297 .259 35 10.481 1. 712
~ 1.308 .232
40 10.431 1. 709 1.307 .202
/ ~
]0.030 1.607 1. 268 .178 I .8 , 50 9.404 1. 507 1. 228 .IM 55 8.607 1.406 1.186 .134
~
60 7.678 1.307 1. 143 .116 I.
I 65
.-- 6.649 1. 209 1.099 .099 I-'-NAGA 64,-02/ 70 5.549 1.112 1.055 .084 4.416 1.020 1.010 .071 m 3.287 .932 .965 .059
L ..4 85
2.213 .851 .923 .047 ~ 1. 245 .778 .882 .036
r--
-
.449 .711 .844 .022
1----
V
1--- 100
!-. 0 .653 .808 0
I
~
J--
~
L. E. radius: 2.884 percent c ~
t--- I---
'0 ~ACA 65,2-016 BASIC THICKNESS FORM ..
, ___ c,= .20 (upper surface)
(l>jl')' I tJ.v.IV
(perc~~t c) I (perc~nt c) I
_ ..
,"' ..
! I--- ----- ----- ----- ----- -------
0 0 I 0 0 1. 950
( ,......-
. ,) 1. 202 .560 .748 1. 650
~
I--- .75 1. 423 .690 .831 1. 500
-- V 1. 25
V- 1. 796 .812 .918 1. 275
~
2.5 2.507 1. 033 1. 068 .920
V
=f /0 ~ /' 5.0 3.543 1. 217 1.103
.680 1.2 7.5 4.316 1.287 1.134 .545
v<_. __
~ 10 4.954 ·1.328 1.152 .480 --.20 fower surface) 5.958 ~1.379 1.174 .390
II ~
20 6.701 '1. 409 1.187 .- .325 25 7.252 1. 433 1.197 .285 30 7.645 1. 453 1. 205 .255
"~
8 '( 35 7.892 1. 469 1. 212 .225 40 7.995 1.484 1. 218 .200 45 7.938 1. 497 1. 224
""""- .180
50 7.672 1. 491 1. 221 .160
f
55 7.184 1. 421 1.192 .140 NAGA 65, 2-0/6 60 6.495 1.328 1.152 .125 65 5.647 1. 235 1.111 .110 70 4.713 1.147 1.071 .095 75 3.738 1.05fi 1.028 .080 80 2.759 .970 .985 .066 85 1.817 .886 .941 .050 ~ J--
t--. 90 .982 .816 .903 .040
V t--.
95 .340 .769 .877 .025 100 0 .733 .856 0 ~
~ 1--
t--- ~ L. E. radius: 1.704 percent c
o .4 .8 1.0 .C' .6
I a:>/C
SUMMARY OF AIRFOIL DATA 81
2 NACA' 65,2-02.~ BASIC THICKNESS FORM .0 y, x Ilv';v (vIV)2 vlV (percent c) (percent c)
I
",.- ~- C', =.2 (uPper surface) 1. 414, 0 0 0 0 0-_
l.---
/ '.632 1. 161
.5 1. 664 .400 .6 .707 1. 084 .75 2.040 .500
-
~ .826 .967 1. 25 2.628 .682
V
V
.971 .811 2.5 3.715 .943
~ / t\
1.110 .633 5.0 5.300 1. 232 1.173 .539 7.5 6.478 1.375 1.211 .479 10 7.433 1.467
v-- - '---.2
fower sur'fqce)
/
1. 256 .380 15 8.889 1. 577
~
1. 276 .324 I .2 20 9.917 1.628 1. 286 .281 25 10.648 1. 655 ~ 1.295 .247 30 11.142 1. 677
1/
1. 302 .220.
~ 35 11.423 1.694
1. 307 .198 40 11.499 1. 708 1.310 .178 45 11.361 1.716
~
.161 1. 712 1. 308 50 10.949
If I
~ 1. 267 .147
55 10.179 1.606 .8 1.195 .110 60 9.108 1.428
'/ 1. 274 1.129 .096
65 7.848 .093
~ 1.135 1. 065
'" 70 6.461
1.001 .080 NACA 65,2-023 75 5.015 1. 003 .893 .945 .053 80 3.618 .035 85, .803 .896 2.345 .022 .856 90 1.258 .732 .018 .682 .826 4 95 .439 V .651 .807 0 100 0
I
~
V
I---
--
L. E. radius: 2.955 percent c
~ -
~r-
c.---
NACA 65,3-Q18 BASIC THICKNESS FORM
--
x 1/ I (vIV)2 v/V 1l".IV .'
(percent c) (percent c) _._-(!,=.32 (upper surface) 8 / 1.750 0 0 0 0 I.
.806 1.387 .5 1.324 .650 ~ ~ .866 1.268 .75 1. 599 .,750 .. --0 ~ .934 1.108 1.25 2.004 .872
~
1. 010 .800 2.5 2.728 1.020 1.086 .677
c7 5.0 3.831 1.179
/
1.124 .568 7.5 4.701 1.263
~
1.149 .489
~ 10 5.424 1.320
/ /
1.2 1.180 .395 15 6.568 1.393 ~ 1.200 .334 20 7.434 1.439 ----:32 (lower surface) 1.214 .292 25 8.093 1.473
! 1/ ~ 1.226 .260
30 8.568 1.502 1. 235 .232 35 8.868 1.526' 8.900 ' 1.243 .209 40 1. 546
'{ / "~
1. 562 1. 250 .186 45 8.916 1.513 1.230 .165 50 8.593 .8 1.433 1.197 .142 55 8.045 1.348 1.161 .123 60 7.317 1. 258 1.122 .107 65 6.450
I
"-
.093 NACA 65,3-018 1.169 1.081 70 5.486 .080 1.079 1. 039 75 4.456 .066 .992 .996 80 3.300 .905 .951 .054 85 2.325
I
.818 .904 .040 90 1.324 .024 .738 .859 95 .492 .658 .811 0 100 0
- r---
~
--
--
~ L. E.'radius: 1.92 percent c I'--.
~
r---t-
-~
o
NACA 65-Q06 BASIC THICKNESS FORM x 11 I!.v.IV (v/l?)' "IV (percent c) (percent c) /6 4.815 0 0 2.110 1.022 .5 .476 1. 044 1.027 1.780 .75 .574 1.055 1.031 1.300 1.25 .717 1.063 .965 1. OSI 1.040 2.5 .956 ____ .(!,=.Ol (upper surface) .695 1.049 1.310 1.100 5.0 ,- 1.055 .560 7.5 1.589 1.112 .474 1.120 1.058 10 1.824 .--0 1. 065 .381 2.197 1.134 I """'" --- -- - ~
r=- 1.069 .322
---~O/ fower,surface) .2.482 1.143 ~ r--..
1.072 .281 2.697 1.149 ~ 1.075 .247 2.852 1.155 1.077 .220 2.952 1.159
~
1.078 .198 2.998 1.163 1.080 .178 2.983 1.166 .8 45 1.079 .160 2.900 1.165 1.070 .144 2.741 1.145 1.060 .128 2.518 1; 124 NACA65-00B 1.049 .114 2.246 1.100 1.036 .100 1.935 1. 073 1.022 .086 1.594 1.044 1.006 .074 1.233 1.013 A .900 .060 .865 .981 .972 .046 .510 .944 .950 .031 .195 .002 o .
.926 0 .858 L. E. radius: 0.240 percent c 1.0
o .2 .4 .0 .8
REPORT NO. 824-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS NACA 65-008 BASIC THICKNESS FORM x (vfV)' v/V AV./V (percent c) (percent c)
I
0 O· 0 0 3.695 .5 .627 .978 .989 2.010 1.6 .75 .756 1.010 1.005 1.61ltl 1.25 .945 1.043 1. 021 1. 340 2.5 1. 2t17 1. 08t1 1. 042 .95t1 5.0 1.745 1.125 1. 061 .689 7.5 2.118 1.145 1. 070 .560 .c, "'.04 (upper surf'ace) 10 2.432 1.158 1. 07t1 .477 ~ 15 2.931 1.178 1.085 .382 1.2 I~ 20 .3.312 1.192 1. 092 .323 '.
r.-
3.599 1.203 1.097 .281,
~
"":0+ (lower surface) 3.805 1. 210 1.100 .248
V ~
35 1.217 3.938 1.103 .221 3.998 1.222 1.105 .199 ...........
45 1.226 3.974 1.107 .178 50 L222 3.857 1.105 .160
~
55 3.688 1.193 1.092 .145 .8 3.337 L 168 L078 .128 65 2.971 1.130 1.068 .113 70 2.553 1.094 1.046 .098 NACA 65-008 75 2.096 1.055 1.027 .084 80 L617 1.014 1.007 .012 85 1.131 .971 .985 .059 90 .664 .923 .961 .044 ·4 .95 .252 .873 .934 .031 100 0 .817 .904 0 .,- r-- L. E. radius: 0.434 percent c
o
NACA 65-009 BASIC, THICKNESS FORM x 1/ (v/V)' vlV Av./V (percent c) (percent c) ,1.6 .() 0 0 0 3.270 .5 .700 .945 .1l12 1.962 :'c, =.06 (upper surf'ace) , .75 .845 .985 .992 1. 655 L25 1.058 L037 L018 1. 315 2.5 1. 421 L089 L044 .,950 O.
5.0 1. 961 1.134 1.065 .687 7.5 2.383 1.159 1.077 .560 10 2.736 L 177 1.085 .477 1.2
---
15 3.299 1.200 1.095 .382 K.
~
20 3.727 1. 216 1.103 .323 .. ·.06 (lower 8urrace)
'1/
25 4.050 L229 L 109 .280
~
30 4.282 1.238 L113 .248
If
35 4.431 1.246 1.116 .220 4.49t1 40 1.252 1.119 .198
~ 45 4.469 L258 L 122 .178
""" .8
50 4.336 1. 250 1.118 .160 4.086 55 L220 L 105 .144 60 3.743 L185 L089 '.128 65 3.328 1.145 L070 .111 NACA 65-009 70 2.856 L 103 L050 .097 75 2.342 1.059 L029 .084 80 1.805 1. 013 l.ootl .071 .4 85 1.260 .968 .981 .051l 90 .738 .912 .955 .044 95 .280 .856 .925 .. 030
....- 100 0 .797 .898 0
I-- L. E. radius: 0.552 percent c
--
o
NACA 6lHl10 BASIC THICKNESS FORM :& (vi V)' v/V Av./V (perJnt c) (percent c) 1.6
I I
0 0 0 0 2.967 ,ie, =.08 (upper surface) .5 .772 .911 .954 1.911 .75 .932 .960 .980 1.614 : O.
L25 1.169 1. 025 1.012 1.292 2.5 1.574 1.086 1.042 .932 !--
Ir-t ~
5.0 2.177 1.143 1.069 .679
I.e
7.5 2.647 1.177 L085 .558
-
~ 3.040
~ 10 1.197 1.094 .480
\08 (lower surra'ce)
(/ 15
3.666 1.224 1.106 .383
~
4.143 1.242 20 1.114 .321 1.121 25 4.503 1.257 .280
(vr
30 4.760 1.268 1.126 .248
r(
"'-
35 4.924 1.277 1.130 .222
~
40 4.996 1.284 1.133 .8 .199 45 4.968 1.290 1.136 .179 50 4.812 1.284 1.133 .160 55 4.530 1.244 I.U5 .141 NACA 65-010 60 4.146 1.202 1.096 .126 85 3.682 1.158 1.076 ; 110 70 3.156 1.112 1.055 .097 75 2.584 1.062 1.031 .082 80 1.987 1.011 L005 .070 L3g5 85 .958 .979 .058 90 .810 .903 .1l50 .045 V- 95 .306 .844 .919 .030 I'-- 100 0 .781 .884 0 L. E. radius: 0.!l87 percent c
o .Z .4 .6 .8 /.0
SUMMARY OF AIRFOIL DATA NACA 65!-012 BASIC THICKNESS FORM 2.0 .
x II (vIV) , vlV !;v.IV (percent c) (percent c)
I
I
0 0 0 0 2.444 /.6 .5 .923 .848 .921 1.776 .75 1.109 .935 .967 1.465 ----c, ... 12 (upper surface) 1. 25 1.387 1.000 1.000 1.200 ,,~-- 2.5 1.875 1.082 1.040 .931 5.0 2.606 1.162 1. 078 .702 7.5 3.172 1. 201 1. 096 .568
r-
e----
10 3.647 1.232 1.110 .480
~ ~
~ 15 4.402 1.268 1.126 .389 1.2
-- 20 4.975 1.295 1.138
.326
-
V_
~ 25 5.406 1.316 1.147 .282 --olE (lower surface)
I{/ 30 5.716 1.332 1.154 .251
~
35 5.912 1.343 1.159 .223 40 5.997 1. 350 1.162 .204
(vt
I( 45 5.949 1. 357 1.165 .188
50 5.757 1.343 1.159 .169 i'-..
.8 55 5.412 1. 295 1.138 .145 60 4.943 1.243 1.115 .127
"
"
65 4.381 1.188 1. 090 .111 1.134 1.065 70 3.743 .094 NACA 65,-012 75 3.059 1.073 1. 036 .074 2.345 80 1.010 1.005 .062 1.630 .949 .974 85 .049 90 .947 .884 .940 .038 .4 95 .356 .819 .905 .025 100 0 .748 .865 0 -I- ~-
e--
L. E. radius: 1.000 percent 'c
r----
l--- '----
o
NACA 60',-015 BASIC THICKNESS FORM x y (vi V)' vlV !;v~IV (percent c) (percent c) ___ c,=.22 (upper surface) .. ---
---- ----- ----- ------ ------
I.B 0 0 0 0 2.038.
........
.5 1.124 .654 .809 1. 729 'f .75 1. 356 .817 .004 1.300
-
~ ~
1.156 1. 25 1. 702 .939 .969 f-"' 2.5 2.324 1.063 1.031 .920 .682
~ - 5.0 3.245 1.184 1.088
V
.563 7.5 3.959 1. 241 1.114
~
~ ~ ~
.487 /.2 10 4.555 1. 281 1.132 .393 15 5.504 1.336 1.15R ~ --.22 (lower surfac~ .,334 20 6.223 ' 1.374 1.172
v-
(
6.764 1.397 1.182 .290
~
1.191 .255 30 7.152 1.418 7.396 1.199 .227 35 1.438 40' 7.498 1. 452 1. 205 .203
~
I
45 7.427 1. 464 1. 210 .184 .8 50 7.168 1. 433 1.197 .160 55 6.720 1.369 1.170 .143
/ 60 6.118 1. 297 1.139 .127
NACA 65.-015 65 5.403 1.228 1.108 .109
'" 70 4.600 1.151 ].073 .096
75 3.744 ],077 1.038 .078 ],002 .068 80 2.858 1.001 .4 85 1.977 .924 .961 .052 90 1.144 .846 .920 .038 I-- .026 95 .428 .773 .879 '
r--
r--
- 0
/ 100 0 .697 .835
~
r-- f..--
.---
r--
L. E. radius: 1.505 percent c
o
NACA 65,-{)18 BASIC THICKNESS FORM x v " ,_-- c,=.3E (upper surface) (oIV)' vlV !;v.IV (percent c) (percent c)
/
1.6
-----
----- ------
r-----"
1. 746 0 0 0 0 ~ .791 1.437
~ .5 1.337 .625
L
---
1.302 .75 1.608 .702 .8.38
.,,-
-- V
.004 1.123 1.25 2.014 .817
!~ 1.020 1.010 .858
2.5 2.751
I V ~ 1.092 .650
5.0 3.866 1.192 1.2 .542 7.5 4.733 1.275 1.129 1.153 .474 10 5. 457 1.329 ---o3E (lower surface) 1.402 1.184 .385 15 6.606
v-
~
/
.327
""'" 20 7.476 1.452 1.205
.285
(vt 25 8.129 1.488 1. 220
8.595 1. 515 1.231 .251 1.539 1.241 .225 35 8.886
'/ ~
1; 561 1.249 .203 .8 40 8.999 8.001 1.578 1.256 .182 8.568 1.526 1.235 .157 j r'.
8.008 1.440 1.200 .137 NACA 65 -018 3 7.267 1.353 1.163 .118 I ;104 6.395 1.262 1.123 70 5.426 1.170 1.082 .087 75 4.396 1.076 1.037 .074 .4 80 3.338 .985 .992 .062 r- .050 85 2.295 .896 .947
y--
r-- .039
00 1.319 .813 .002
-- i--
-- .026
95 .400 .730 .854 100 0 .657 .811 ~ ~
"-
'----
-
L. E. radius: 1.96 percent c
--
o .6 .8 LO .2 .4 ;rIc REPORT NO. 824-NATIONAL ADVISORY COMMIT'fEE FOR AERONAUTICS NACA 65.-021 BASIC THICKNESS FORM 2.0 I I I I .... c, =.41 {upper surface} x y
r-- '\
(v/V) , v/v !!.Va/V (percent c) (percent c)
{ I
I I
h
l\ 0 0 0 0 1. 531
1.8 .5 1. 522 .514 .717 1.333
I
/
.75 1.838 .607 .779 1.215
----
1.25 2.301 .740 .860 1.062 ~
~"'\ 2.5
/' f\ ./ 3.154 .960 .980 .838
5.0 4.472 1.186 1.089 .649 V 7.5 5.498 1. 293 1.137 .544 10 .
6.352 1.371 1.171 .478
/
~
15 7.700 1. 469 1.212 .388
, .I.e
8.720 1. 533 1.238 .330 ..
25 9.487
"'" 1. 580 1.257 .289
'.44 (lower surf'ace) 30 10.036 1. 621 1.273 .255
I / K\
10.375 1. 654 1.286 .229 ,/
'" 10.499 1. 680 1.296 .206
45 10,366 1. 700 1.304 .184
I /
~ 50 9.952 1. 633 1. 278 .158
"
.8 55 9.277 1.508 1.228 .139 60 8.390 1.397 L 182 .120
~~
65 7.360 1.286 1.134 .101
I
/
70 6.224 1.177 1.085 .087 65 -02/ NACA 75 5.024 1.073 1.036 .073 80 3.800 .970 .985 .058 85 2.598 .872 .934 .047
I
90 1. 484 .778 .882 .035 .546 .694 .833 .020
).--- -
I--
.616 .785 0
I--
Y
1--1-
).--- L. E. radius: 2.50 percent c
~
~
t---
I--: t--
o
NACA 66,1-012 BASIC THICKNESS FORM x y (v/V)' v/V !!'Va/V (percent c) (percent c) /.6 I 0 0 0 0 2.555 .5 .900 .854 .924 1. 780 =./2 (upper surface) .A .75 1.083 .902 .950 1. 540 1. 25 1. 343 .964 .982 1.247 , 2.5 1.803 1.069 1.034 .925 5.0 2.484 1.138 1. 067 .673
Vq
I--'" 7.5 3.019 1.175 .552 +- 1.084 J.---
~
1.2 10 3.482 1. 201 1. 096 .474 , 15 4.214 1. 237 1.112 .381
---
V 20 4.779 1.257 1.121 .319
;: '--Ii? (lower surface)
I'
25 5.218 1. 272 1.128 .280 ~ 30 5.550 1.284 1.133 .248
(vJ
35 5.786 1. 293 1.137 .220 40 5.934 1.302 1.141 .195
rl
45 5.998
1'-. 1. 309 1.144 .176
.8 50 .161 5.972 1.313 1.146 55 5.844 1.320 1.149 .144
" "'-
5.594 1.327 1.152 .130 5.165 1.297 1.139 .. 117 NACA 66,1-0/2 70 4.535 1. 221 1.105 .099 75 3.789 1.143 1. 069 .083 2.964 1. 061 1.030 .069 85 2.098 .974 .987 .053 90 1. 244 .041 .885 .941 95 .477 .028 .792 .890 l,..--- 100 0 .701 .837 0 ~
-
~
I'--
L. E. radius: 0.893 percent c
-
o
NACA 66,2-015 BASIC THICKNESS FORM I I x y :c, =.20 (upper surf'ace) (vW) , v/V AVa/V (percent c) (percent c) : /.8 0 0 0 2.085 ,.- .'\. .5 1.110 .700 .837 1. 703 .75 1. 329 .870 .933 1. 382 ~ 1. 25 1. 645 .940 .970 1.156
:---r-
---
2.5 2.229 1.048 1.024 .898 0.,
~ V
~
5.0 3.086 1.154 1.074 .656 I.R 7.5 3.757 1.210 1.100 .547
- "
---
10 4.337 1. 244 1.115 .473 ····.EO (tower surface)
(
V 5.255 1.290 1.136 .382
~
~ 20 5.964 1.323 1.150 .323
(vl
21i 6.516 1. 342 1.158 .283 30 6.933 1.359 1.166 .248 35 .222 7.230 1.374 1.172
I
.8 40 7.415 .199 1.387 1.178 45 7.495 1.182 .179 1. 397
I
50 7.460 .161 1. 407 1.186 55 7.294 1.415 1.190 .145 NACA '66,i?-0I5 60 6.961 1. 421 1.192 .131
"
65 6.405 1. 372 1.171 .122 70 5.597 1. 267 1.126 .102 75 4.652 1.162 1.078 .080
"
.4 80 3.616 1.057 1.028 .066 85 2.545 .953 .976 .050 ~
r--- 90 1.488 .848 .921 .037
V- b- f--
95 .560 .743 .862 .025 I-" 100 0 .640 .800 0
f...--
"'-----
r-
I----
-
L. E. radius: 1.384 percent c .8 1.0
o .E .4 .6
SUMMARY OF AIRFOIL DATA 85
2. NACA 66,2-018 BASIC THICKNESS FORM .0
I
i
(perc~nt c) (percXnt c) I (vIF)! I vjlT I av.IV
I I
------ ----- ------- ------ ----
_ c,=.22 (upper surfac6j 0 1.659 0 0 0
-
.5 1. 438 .590 .768 1. 317 .8 1. 209 .75 1. 730 .740 .860 1.091 1. 25 2.180 .918 .958 /0 2.5 2.938 1. 084 1.041 .867
~
- 1.103 .665
5.0 3.984 1. 217
V J--I- 1.134 .544
7.5 4.804 1. 285
~ 1.151 .469
10 5.486 1. 325
V
~ 1.172 .379
15 6.541 1. 373 2 / , 1.184 .323 20 7.342 1. 401 , 1. 422 1.192 .282 25 7.957
V
~ 22 fower surroce) 1.440 1.200 .251
If / ~ 30 8.419
~ 1. 456 1. 207 .224 35 8.741 1. 468 1. 212 .201 40 8.933 , 1. 478 1. 216 .181 45 8.998 1. 488 1. 220 .162 50 8.934 8 / 1. 497 1. 224 .146 55 8.719 60 8.316 1. 502 1. 226 .134 65 7.629 1. 442 1. 201 .102 1. 314 1.146 .089 70 6.657
I~
NACA fi6,2-0/8 75 5.523 1.185 1. 089 .078
'I
I 1. 029 .064 80 4.296 1.059 .052 85 3.027 .936 .967 .O(l 00 1. 789 .817 .004 .837 .027 95 .672 .700 .771 0 100 0 .594
1---
I---
-
V I--
t- j;-- L. E. radius: 2.30 percent c
L.---
'---t--- l---
t--
-
NACA 66-006 BASIC THICKNESS FORM x y (./V), /lva/V v/V (percent c) (percent c)
I
----- 0 4.941 0 0 1. 052 1. 026 2.500 .5 .461 2.020 1.028 .75 .554 1. 057 1. 031 1. 500 1. 25 .693 1.062 .967 1.071 1. 035 2.5 .918 .695 1. 086 1. 042 5.0 1. 257 surface) Ai'Of (upper .. 554 1.098 1. 048 7.5 1. 524 .474 " 'I 1.107 1. 052 10 1. 752 1.2 .379 1.119 1. 058 15 2.119 /0 .320 1.128 1.062 20 2.401
!'--- .278
'.'Of (lower surfoce) 2,618 1.133 1. 064 F t-....
.245 1. 138 1.067 30 2.782 "'-.
.219 2.899 1.142 1.069 1. 070 .197 1.145 40 2.971 1. 071 .178 3.000 1.148 .8 1.073 .161 2.985 1.151 1. 074 .145 2.925 1.153 1. 075 .130 2.815 1. }.i5
I
""
1. 074 .116 65 2.611 1.154 NACA 88-008 1. 057 .102 2.316 1.118 1.040 .089 75 1. 953 1. 081 1.020 .075 80 1. 543 1. 040 .996 .998 .061 85 1. 107 .4 .047 .948 .974 90 .665 .030 .890 .943 95 .262 .822 .907 0 100 0 L. E. radius: 0.223 percent c
I
o
NACA 66-008 BASIC THICKNESS FORM y x vjV av./v (v/V)' (pereent c) (percent c) Ul 3.794 0 0 0 .984 2.220 .968 .5 .610 1. 023 1.011 1. 82" .75 .735 1.388 1. 046 1. 023 1.25 .919 1. 038 .949 ,f. =.03 (upper ,surfoce) 1.078 2.5 1. 219 ,-0 1. 052 .689 1.107 5.0 1. 673 1.062 .552 2.031 1.128 7.5 .474 1. 068 2.335 1.141
~ ,1 l ""'" 10
Id? 1:076 .379
~03 (lower surface) 1.158
~ 15 2.826
1.082 .321 1.171 20 3.201
"-
1.085 .278 3.490 1.178 1. 089 .246 3.709 1.186 1. 091 .220 3.865 1.191 1. 094 .198 3.962 1.196 .8 .178 1.096 4.000 1. 201 1.098 .161 3.978 1. 205 1. 099 .145 3.896 1. 208 .130 NACA 68-008 1. 101
"'" 3,740 1.213
.115 1. 096 3.459 1. 202 .101 1.1.56 1. 075 3.062 .087 1. 050 2.574 1.103 .073 .4 1. 048 1. 024 2.027 .058 .989 .994 85 1. 447 .962 .045 .926 90 .864 .925 .029 .855 95 .338
-
.876 0 r-- .768 100 0
--
L. E. radius: 0.411 percent c 1.0 .8 I .2 .4 .6 <rIc REPORT NO. 824-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS NACA 66-009 BASIC THICKNESS FORM 2.0 x (v/V)' v/v M./V (perlent c) (percent c) 0 0 0 0 3.352 1.6 .5 .687 .930 .964 2.100 .75 .824 .999 .999 1.750 1.25 1.030 1.036 1.018 1.340 2.5 1.368 1. 079 1.039 .940 ,5.0 1.119 1.058 1.880 .686 /0, ~.05 (upper surface) 7.5 2.283 1.142 .552 1.069 10 2.626 1.159 1.077 .473 /0 15 3.178 1.178 1.085 .379 I.E ,....- I .323 20 3.601 1.190 1.091 ,
::::<: 25 8.927 1.201 .280
1,096 --:05 (lower surface) 30 4.173 1.210 1.100 .246
~
85 1.217 1.103 .220
"- 4.848
40 4.457 1.221 1.105 .197 I(
"- 45 1.108
4.499 1.228 .178 1.232 1.110 .161 50 4.475 .8 55 1.237 1.112 .145
"- 4.381
60 1.114 .130 4.204 1.240 65 1.230 1.109 .116 3.882 70 8.428 1.172 1.083 .100 NACA 56-DOB 1.055 .085 ·75 2.877 1.113 80 2.263 1.050 1.025 .071 85 .985 .992 .057 1.611 90 .915 .957 .961 :043 .4 95 .839 .916 .028 .374 0 .747 .864 0 V- ""'- L. E. radius: 0.530 percent c I
o
NACA 66-OlO BASIC THICKNESS FORM x 1/ (v/V) , Il.v./V v/V (percent c) (percent c) 1.6 0 0 0 0 3.002 .5 .759 .896 .947 2.012 /e, =.07 (upper surface) .75 .913 .972 .986 1.686 1.25 1.141 1.023 1.011 1.296 2.5 1. 516 1. 078 1.038 .931 5.0 2.087 1.125 1.061 .682 7.5 2.536 1.154 1.074 .551
,..---
/.2 10 2.917 1.174 1.084 .473
or;; - 15 3.530 1.198 1.095 .379
k
20 4.001 1.215 1.102 .322 -;07 fower surface)
~
25 4.363 1.226 1.107 .279
"-
30 4.636 1.236 1.112 .246
If 35 4.832 1.243 1.115 .220
40 4.953 1.249 1.118 .198
~
45 5.000 1.25. 1.120 .178 .8
"'
50 4.971 1.261 1.123 .161 55 4.865 1.265 1.125 .146
"" 60 4. 665 1.270 1.127 • ISO
65 4.302 1.250 1.118 .114 NACA 6{]-010 70 3.787 1.190 1.091 .099 75 3.176 1.121 1.059 .085 80 2.494 1.052 1.026 .070 85 1.773 .979 .989 .056 90 1.054 .904 .951 .043 95 .408 .821 .906 .027 -I--- 100 0 .729 .854 0
t----
r-- ~
L. E. radius: 0.662 percent c
o
NACA 661-012 BASIC THICKNESS FORM x 1/ (v/V) , v/v Il.v./V (percent c) (percent c) I /0, =.Ie (upper surface) 0 2.569 0 0 0 / .5 .906 .894 1.847 .800 , - .75 1.087 .915 .957 1.575 1.25 1.237 1.358 .980 .990 ~ 2.5 1. 808 1. 073 1. 036 .913
~~
--- 5 2.496 1.138 1. 067 .674 /,2
f- 7.5 3.037 1.177 1.085 .549
--
--
--':IE (lower 10 3.496 1. 204 1. 097 .473
~ surface)
r; 15 4.234 1.237 1.112 .380 ~
I'
20 4.801 1.259 1.122 .323 25 5.238 1.275 1.129 .280
(( 30 5.568 1.287 1.134 .246
35 5.803 1.297 1.139 .221 ~ ,8 40 5.947 1. 303 1.142 .197 45 6.000 1.311 1.145 .176
"
"-
50 5.965 1.318 1.148 .162 55 5.836 1.323 1.150- .147 NACA 66,,-012 60 5.588 1.331 1.154 .132 65 5.139 1. 302 1.141 .113 70 4.515 1. 221 1.105 .098 75 3.767 1.139 1.067 .084 80 2.944 1. 053 1.026 .069 85 2.083 .96E .984 .053 I-- 90 1.234 .879 .938 .040
-
.474 .788 .888 .031
--
100 0 .687 .829 0
'-- f..---
---
-
L. E. radius: 0.952 percent c 1.0
o .2 .6 .8
·4
SUMMARY OF AIRFOIL DATA 87
2 NACA 66z-{)15 BASIC THICKNESS FORM x (v/V) • v/V t;.v./V (percent c) (percent c)
I
I I
----- -----
2.139 0 0 0 6 .5 I.J22 .760 .872 1.652 /.
.9)6 1:431 .75 1.343 .840
_---c =.:! (upper surface)
t 1.172 1. 25 1. 675 .929 .964 ,-' 2.5 2.235 1. 055 1. 027 .895 ~ 5 3.100 1.163 1. 078 .663 1.099 .547 7.5 3.781 1. 208 1.114 .473 10 4.358 1.242
v---- -
~ 1.134 .381
~ 15 5.286 1. 288
I.
...-1---
1.148 .822 20 5.995 1.817
ELY
.280 25 6.543 1.840 1.158
14 power surface) 1.164 _.248
30 6 .. 956 1. 356
~
~ 1.170 .222 35 7.250 1.370 1.175 .200 40 7.430 1.380 1.179 .180 45 7.495 1.391
"\
1.184 .163 50 7.450 1. 401 ~
8 I
1.188 .146 55 7.283 1.411 1.192 .131 60 6.959 1.420 6.372 1.867 1.169 .118
1/
1.122 .096 70 5.576 1.260 NACA 66 -0/5 2 1.156 1. 075 .080 75 4.632 3.598 1. 053 1.026 .065 2.530 .949 .974 .051
"
.847 .920 .039 90 1. 489 95 .744 .863 .025 - .566 .639 .799 0 100 0 I-:--
-
,.--
f--
•
--
p- L. E. radius:' 1.435 percent c
--
'-- I---
--
-
o
--
NACA 66s-{)18 BASIC THICKNESS FORM x II , (V/V)2 t;.v./V viv (percent c) (percent c)
_--c,_.3 (u~Der surfacp)
,~ .. -
1.8 0 0 0 0 1..773 .806 1. 456 .5 1.323 .650 I-- .. 735 .857 1. 312 .75 1.571 ~
~ .897 1.121
1. 25 1.952 .850
- 1.002 .858
2.5 2.646 1. 005
- :...---
1.154 1. 074 .649
V 5 3.690
~
1.111 .545 0/ 7.5 4.513 1. 234
t\ ../'"
1.134 .472 1.2 10 5.210 1.285 1.162 .381 15 6.333 1. 350
----
1.180 .323 --;3 20 7.188 1. 393 fower surface} 1.423 1.193 .282 25 7.848
/ /-- ~~
1. 202 .250 30 8.346 1. 445 1. 210 .223 35 8.701 1. 464 1. 217 .201 I 40 8.918 1.481
I /
1.223 .181 45 8.998 1.496 .8 1. 509 1.228 .163 50 8.942 1.522 1. 234 .147 55 8.733 1. 534 1.238 .131 60 8.323
I 1.438 1.199 .114
65 7.580 NACA 68 -0/8 s
"
1. 302 1.141 .095 70 6.597
I
1.172 1.083 .077 75 5.451 1.045 1. 022 .061 80 4.206 .950 _ 2.934 .922 .048 .4 85
"
1. 714 .803 .896 .037 .022 ~ 95 .646 .692 .832
-
.581 .166 0 100 0
V
I--
--
I--
---
~ L~ E. radius: 1.955 percent c ~
t--
---
o
-
NACA 66.-{)21 BASIC THICKNESS FORM J '" x --····c,-.4 (upper surface)_ 11 2 t;.v./V (v/V) v/V (percent c) (percent c)
\
:;;;;0- 1.8
-----
~ 1.541 0 0 0 .761 1.314
'\~ I .. 5 1.525 .580
/' .-
.797 1. 218 .75 1.804 .635
---- .869 1.054
V 1. 25 2.240 .755
--
/" I
.952 .976 .828
\ 2.5 3.045
lJ ~ /' 1.143 1. 069 .635
5 4.269 1.2 1.116 .542 1.5 5.233 1. 246 1.148 .472 10 6.052 1.318 -----.4 fower surface) 1.185 .381 7.369 1.405
/ /' ~~
.324 8.376 1. 459 1. 208 1.224 .283
(vt 25 9.153- 1. 499
, 1. 236 .251 30 9.138 1. 528
I /1
1. 245 .224 10.154 1. 551 1.255 .202 .8 40 10.407 1.574 1.263 .183 45 10.500 1.594 1. 269 .165 10.434 1.611 I 50 1.276 .148 10.186 1.629
I 55
N4CA 88 -02/
, 1.284 .132
4 9.692 1. 648 1.228 .114 65 8.793 1.508 1.155 .093 7.610 1.335 1. 084 .073
I 6.251 1.176
.4
"
1. 015 .058 4.796 1. 031 I-- 80
,-
I -
.944 .046 ~ 85 3.324 .891
I---
.873 .034 1.924 .763
V 90
l--
""- .805 .020 95 .717 .648 .734 0 100 0 .539
V -
'-
r--- ~
L. E. radius: 2.550 percent c
-
o 1.0
.2 .8 .8 .4
--
a:/c REPORT NO. 824-NATIONAL ADVISORY COMMITTEE FOR AERONAU'l'ICS NACA 67,1-015 BASIC THICKNESS FORM
e.o
x y I:>.v./F (II/V)' v/F (percent c) (percent c) I I I ------ ------------ ------- ------- ,c, ",./2 (upper surface) , 2.M2 0 0 0 0 /.6 1.167 .650 .806 1. 560 .5 : 1.394 .970 .985 1. 370 .75 , 1. 764 1.059 1. 029 1.152 1. 25 .906 2.5 2.395 1.140 L068 .667 5.0 3.245 1. 209 1.100 ( f- , I.--- 7.5 3.900 1. 239 1.113 .548
I--
I-- 1.122 .470
~ 10 4.433 1. 259
(...:--- /' .370 15 5.283 1.285 1.134 i.e .312
V'-, 20 5.940 1.304 1.142
'- -./2 1.148 .276 power surface) 25 6.454 1. 318 1.153 .248
V/ 30 6.854 1.330
~
.221 35 7.155 1.341 1.158
(v/
1.162 .201 40 7.359 1. 351
'( 1.166 .180
45 7.475 1. 360
\~
.160 50 7.497 1.368 1.170 .8 1.173 .142 55 7.421 1.375 1.175 .124 60 7.231 1. 381 1. 388 1.178 .111 65 6.905
\
1.179 .108 70 6.402 1.390 NACA 67,/-0/5 1.149 .094 75 5.621 I. 321 4.540 1.176 1.084 .071 85 3.327 1. 018 1. 009 .060 90 2.021 .864 .930 .045 .4 95 .788 .712 .844 .025 100 .570 .755 0 I----
- r--
I--
V-
~ L. 1'~. radius: 1.523 percent c.
r--
~ I---
-
a
NACA 747A015 BASIC THICKNESS FORM I I x __ -c, =.ec (upper surface) v I:>.v./V (v/F)' v/V (percent c) (percent c) / / 1.6 ------.
----- ----- ----- ------ 2.028 0 0 0 0
~
1.199 .660 .812 1. 680 .5 .75 1. 435 .799 .894 1. 560 ,0 1. 25 1.801 .942 .971 1. 325
/" V
r-:: 2.5 2.462 1.100 1.049 .990
y
~
~ ~ 5 3.419 1. 201 1.096 .695
/.2 7.5 4.143 1. 259 1.122 .551 -...:: 10 4.743 1. 295 1.138 .465
/-
-'-:2c fower surfac~ 15 5.684 1.339 1.156 .383
(/
~
.324 20 6.384 1.369 1.170 .283 25 6.898 1. 390 1.179 .252 30 7.253 1. 409 1.187 .224 35 7.454 1.423 1.193 ~
/
.199 40 7.494 1.435 1.198 .8 .176 45 7.316 1. 391 1.179
"- 1.161 .156
50 7.003 1. 348 55 6.584 1. 306 1.143 .1~
I
" NACA U7AOl5 .1
60 6.064 1. 265 1.125 .108 65 5.449 1. 221 1.105 .093 70 4.738 1.178 1. 085 .079 75 3.921 1.115 1. 056 .4 .065 80 3.020 1. 027 1. 013 .052 85 2.086 .938 .969 .040 ~ 90 1.193 .852 .923
t.--- r-- !--
.02!l 95 .443 .774 .880
--
.018
I--+- 100 0 .703 .838
r--..:
I--
I---- I--- L. E. radius: 1.544 perceut c .8 1.0
o .2 .6
<ric SUMMARY OF AIRFOIL DATA II-DATA FOR MEAN LINES Page Page NACA mean line a=O _____ ,______________________________ 93 N ACA mean line 62_ __ ______ ____________________ ______ __ _ 90 NACA mean line 63______________________________________90 NACA mean line a=O.L_________________________________ 94 NACA mean line a=0.2 _ _______ ._________________________ 94 NACA mean line 64______________________________________ 90 NACA mean line a=0.3 _ ________________ "________________ 94 NACA mean line 65______________________________________ 91 NACA mean line 66______________________________________ 91 NACA mean line a=O.4_ _ _ ____ __ ___ _ __ _ ___ ___ _ __ __ __ _____ 95 NACA mean line 67 _____________ .. __________________ . ___ ·__ 91 . NACA mean line a=O.5__________________________________ 95 N ACA mean line 2.10_ _ ____________________________ _______ 92 NACA mean line a=0.6_ _ _ __ __ ______ __ ____ ___ _ ___ _ __ _____ 95
N ACA mean line a = 0.1- _ _ _______________________________ 96
N ACA mean line 220_ _ ___________________________________ 92 NACA mean line 230 _________________________ .. ___________ 92 NACA mean line a=O.&_ _ ________________________________ 96 N ACA mean line a=0.9_ _ _ __ __ ______ _________ ____ _ __ __ ___ 96 NACA mean line 240_____________________________________ 93 NACA mean line 250 __ .. ________________________ .. _________ 93 NACA mean line a=1.0__________________________________ 97 REPORT NO. 824-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 3.0 NACA MEAN LINE 62 2.0 el,=0.90 ",,=2.81° em ",= -0.113 --
h
x 1/, dll,/dx PR tlv/V=P.Ri4 (percent c) (percent c)
------ ------
I~
/
r-
0 0.80000 0 1.0 0 1. 25 .726 .56250 .682 .171 2.5 1. 406 .52500 1.031 .258
r---
--- 5.0 2.625
.45000 1. 314 .328
r---
7.5 3.656
-- .37500 1.503 .376
..........
lO 4.500 .30000 1. 651 .413 15 5.625 .15000 1.802 .451 20 6.000 0 1. 530 .383
~
o
25 5.977 -.00938 1.273 .318 30 5.906 -.01875 1 .. 113 .279 40 5.625 -.03750 .951 .238 NACA 62 50 5.156 -.05625 .843 .211 60 4.500 -.07500 mean line .741 .185 70 3.656 -.09375 .635 .159 80 2.625 -.11250 .525 .131 90 1.406 -.13125 .377 .094 .~ .2 .727 -.14062 .261 .065 0 -:15000 0 0 l- ~
o
NACA MEAN LINE 63 2.0 e,,=0.80 "'1=1.60° em". = -0.134 x P- 11,
.. " dll,/dx PR tlv/V=PR/4
(percent e) (percent c) r-.
V
I"'---.
1.0 0 0.40000 0
r--
1. 25 .489 .38333 .389 .097
t---
2.5 .958 .36667 .553 .138
t--
I 5.0
1.833 .33333 .788 .197 --I--.
7.5 2.625 .30000 .940 .235 lO 3.333
II .26667 1.066 .267
'~
15 4.500 .20000 1.220 .305 20 5.333 .13333 1. 259 ,315
o
25 5.833 .06667 1. 233 .3G8 6.000 0 1.160 .290 40 5.878 -.02449 .949 .237 NAt;'A 83 50 5.510 -.04898 .850 .213 mean line 4.898 -.07347 .762 .191 70 4.041 -.09796 .673 .168 80 2.939 -.12245 .560 .140 Yc 2 90 1. 592 c . -.14694 .406 .102 95 .827 -.15918 .291 .073 100 0 -.17143 0 0 .I---
o
-----
NACA MEAN LINE 64 2.
e,,=0.76 a;=0.74° em,'I=-0.157 • x y, dy,/dx PR tlv/V=PR/4 . (percent e) (percent e) I. 0 0 0 0.30000
t-- 0 0
V -
r---
1.25 .369 .29062 .257 .064
1---
2.5 .726 .28125- .391 .098 / r-...
5.0 1. 406 .26250 .546 .137 7.5 2.039 .24375 .668 .167
V
10 2.625 .22500 .748 .187
~
15 3.656 .18750 .871 .218 20 4.500 .15000 .966 .242 25 5.156 .11250 1.030 .258 30 5.625 .07500 1.040 .260 NACA 64 6.000 0 40 .999 .250 mean line 50 5.833 -.03333 .910 .228 60 5.333 ~.06667 .827 .207 70 4.500 -.10000 .750 .188 2 80 3.333 -.13333 .635 .159 90 1.833 -.16667 .466 .117 I 95 '.958 -.18333 .334 .084 100 0 -.20000 0 0
I---
r--
I-
-
o
.2 .8 .8 .4 1.0 .Jt'/c SUMMARY OF AIRFOIL DATA 3.0 NACA MEAN LINE 65 Q'i=Oo CI =0.75 Cme/I = -0.187 2.0 x 1/, dy,/dx PR Avj1T=PR/4 (percent c) (percent c) ------- -------- ------ ------ -------- 0 0 0.24000 0 0 1. 25 .296 .23400 .205 .051 1.0 2.5 .585 .22800 .294 .074 5.0 1.140 .21600 .413 .103
,r--
7.5 1. 665 .502 .20400 .126 ~
~
10 2.160 .19200 .571 .143
--
15 3.060 .16800 .679 .170 20 3.840 .14400 .760 .190
'\ 25 4.500 .12000 .824
.206
1/
30 5.040 .09600 .872 .218
o
5.760 .04800 .932 .233 50 6.000 0 .951, .238 60 5.760 -.04800 .932 .233 NACA 65 70 5.040 -.09600 .872 .218 mean line 80 3.840 -.14400 .760 .190 90 2.160 -.19200 .571 .143 95 1.140 -.21600 .413 .103
!ft . .?
100 0 -.24000 0 0 c
-
f.--- -
r--.
o
-
--
NACA MEAN LINE 66 -cI =0.76 ",=-0.74° Cm 'I' = -0.222 x 1/, dy,/dx PR AV/V=PR/4 (percent c) I (percent c) I ------- ------ --------- ------- ------- 1.0 0 0 0.20000 0 0
-
f.--- 1. 25 .247 .19583 .135 .J34
~
f.---
2.5 .490 .19167 .244 .061 ~ 5.0 .958 .18333 .334 .084 7.5 1. 406 .17500 .408 .102
---
10 1. 833 .16667 .466 .117
""
\
15 2.625 .15000 .557 .139
V
20 3.333 .13333 .635 .159
o
25 3.958 .11667 .700 .175 4.500 .10000 .750 .188 40 5.333 .06667 .827 .207 NACA 66 5.833 .03333 .910 .228 mean line 60 6.000 0 .999 .250 5.625 70 -.07500 1. 040 .260 4.500 -.15000 80 .966 .242 Yo .2 90 '2.625 -.22500 .748 .187 c 95 1. 406 -.26250 .546 .137 100 -.30000 0 0 0
t--
:-- t---.
-
a
-
NACA MEAN LINE 67 2.0 cI =0.80 a:"i=-1.60o Cm,"=-0.266 I I-"'~ y, X dy,/dx PR AV/V=PR/4
1-""""
(percent c) (percent c) L /.0 ------- ------- -------
------- ------
~
-- ~ 0 0 0.17143 0 0
~
\
.212 .137 .034 1. 25 .16837 .......
.049 2.5 .421 .16531 .195 .291 .073 5 .827 .15918
\
.356 .089
i/ 7.5 1. 217 .15306
.102 10 1. 592 .14694 .406 o .121 15 2.296 .13469 .483 2.939 .12245 .560 .140 3.520 .616 .154 NACA 67 25 .11020 4.041 .673 .168 30 .09796 mean line 4.898 .07347 .762 .191 5.510 .04898 .850 .213 5.878 .02449 .949 .237 1& .2 6.000 0 1.160 .29'0 c 5.333 -.13333 1. 259 .315 3.333 -.26667 1.066 .267 1. 833 -.33333 .788 .197 -.40000 0 0 100 0 •• >_h r-...
--
o ..? .4 .6 .8 /.0
x/c REPORT NO. 824-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 3.
I
! 1 NACA MEAN LINE 210
I e,,=0.30 ai=2.09°
em ~'I = -0. 006 I 2.
I X y, dy,/dx PR AV/V=PR/4 i (percent e) (percent e) /\ I ------- ------- ------- ------- ------- I 0 0 0.59613 0 0 1. 25 .596 .36236 1. 381 .345
\
1.0 .928 .18504 1. 565 .391 2.5 5.0 1. 114 -.00018 1. 221 .305
\ 7.5 1. 087 .781 .195
10 1. 058 .626 .156 ......... 15 .999 489 .122 20 .940 .408 .102
r--
t-- 25 .881 .348 .087 I 30 .823 .302 .075
o
40 .705 .242 .061 -.01175 50' .588 .198 .049 60 .470 .160 .040 NACA 210 70 .353 .128 .032 80 .235 .098 .025 mean line 90 .118 .065 .016 95 .059 .044 .011 Yc i 100 0 0 7" .2. 0
I
I
I I J
o
NACA MEAN LINE 220 2.0 e.,=0.30 ai=1.86° em",= -0. OlD
!
x y, ,..--" dy,/dx PR AV/V=PR/4 (percent c) (percent c) I ------- ------- -------- -------- -------- \ 1.0 0 0 0.39270 0 0 1. 25 .442 .31541 .822 .206
!""-
.251 2.5 .793 .24618 1. 003 .247 5.0 1. 257 .13192 .988
~
.225 7.5 1. 479 .04994 .900 .8t)!
10 1. 535 .00024 .200 .615 .154
- 15 1. 463
20 1. 377 .465 .116
---
o
25 1. 291 .378 .095 .326 30 1. 205 .082 40 1.033 .253 .063 NACA 220 50 .861 .205 .051 -.01722
~--
mean line 60 .689 .169 .042 70 .516 .135 .034 80 .344 .100 .025 Yo ,., 90 .172 .064 .016 (' .C 95 .086 .040 .010 100 0 0 0
o
NACA MEAN LINE 230 a,=1.65° c.,=0.30 em". = -0.014 x 1/, dy,jdx PlI AVjV=PlIj4 (percent e) (percent e)
I
1.0 0 0 0.30508 0 0 1. 25 .357 .26594 .628 .132
V ~
2.5 .666 .22929 .673 .168 ...............
6.0 1.155 .16347 .791 .198
r--
7.5 1.492 .10762 .853 .213 10 1. 701 .06174 .859 .215
o
15 1.838 -.00009 .678 .170 20 1. 767 -.02203 .519 .130 25 1. 656 .419 .105 NACA 230 30 1;546 .361 .090 1.325 .274 .069 mean line ·40 50 1. 104 .217 .054 60 .883 .177 .044 -.02208 70 .662 .144 .036 ~ •. 2 80 .442 .105 .026 90 .221 .069 .017 95 .110 .042 .011 100 0 0 0 .0
o .2 .8 1.0
.4 riC
SUMMARY OF AIRFOIL DATA 93
3.0 NACA MEAN LINE 240 c,;=0.30 Cli=1.45° Cm'I' = -0.019 2.0 x y, dy,/dx PR AVjV=PR/4 (percent c) (percent c)
------- ----- -------- ------- -------
0 0 C.25233 0 0 .094 1. 25 .301 .22877 .377 .20625 .491 .123 /.0 2.5 .572 .625 .156 5.0 1. 035 .16432 .12653 .718 .180 7.5 1. 397 .09290 .750 .188
r-----.. 10 1.671
/
.677 .169 15 1. 991 .03810 -.00010 .556 .142
I-- 20 2.079
!
- r--
-.02169 .477 .119 25 2.018 .410 .103 30 1. 890 '0 .304 .076 40 1. 620 .234 .0f,9 50 1.350 .186 .047 60 1. 080 N;1cA NO -.02700 .150 .038 70 .810 mean line .110 .028 80 .540 .071 .018 90 .270 .135 .047 .012 1& .2 100 0 0 0 e
o
NACA MEA:\, LINE 250 2.0 c,,=0.30 ai=1.26° Cm /4=-O.026 c X y, dy,/d.r PR At,/V=PR/4 (percent c) (percent c) -------- ----------- -------- -------- ------- 0.21472 0 0 0 0 /.0 .258 .19920 .281 .070 1. 25 .18416 .369 .092 2.5 .498 .922 .15502 .477 .119 5.0 .138 7.5 1.277 .12909 .552 .148
- - .10458 .592
10 1. 570 1. 982 .06162 .624 .156
/' 15
-
.153 2.199 .02674 .610
-
---
-.00007 .547 .137
o 25 2.263
2.212 -.01880 .470 .117 1. 931 .346 .087 NACA 250 1. 609 .255 .064 .049 I. 287 .197 tnean line .038 .965 .154 -.03218 .oao .644 .119 .076 .019 Yo -:> .322 C .cc .051 .013 95 .161 0 0 100 0 ,
o
NACA MEAN LINE a=O c,,=1.0 Q:i=4.56° Cm,I' = -0.083 2.0 ..............
y, x dy,/dx PR AV/V=PR/4 (percent c) (percent c) ~r-...
------ ------ -------
------ -------
'~
0 0 ----------- ---O~498--- ---O~75867--- ............. .460 1. 990 .5 .496 1.0 .641 .69212 1. 985 .75 ..............
.494 .964 .60715 1. 975 1.25 1. 641 .48892 1. 950 .488 2.5 ~, .475 2.693 .36561 1. 900 5.0 3.507 .29028 1. 850 .463 7.5 .450 4.161 .23515 1. 800 ~ .425 5.124 .15508 1.700 1.600 .400 20 5.747 .09693
'---
o
1. 500 .375 25 6.114 .05156 .350 6.277 .01482 1. 400 -.01554 1.300 .325 Q=O 35 6.273 NACA 6.130 -.04086 1.200 .300 mean line -.06201 1.100 .275 45 5.871 -.07958 1.000 .250 50 5.516 5.081 -.09395 .900 .225 -.10539 .800 .200 60 4.581 4.032 -.11406 .700 .175 -.12003 .600 .150 70 3.445 -.12329 .500 .125 75 2.836 .100 2.217 -.12371 .400 .075 -.12099 .300 85 1. 604 I-- 1.013 -.11455 .200 .050 .025 ~ .467 -.10301 .100 -.07958 0
.8 to 100 0
o .4 .8
.:rIc REPORT NO. 824-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS NACA MEAN LINE a=O.1 3.0 C =-O.OF6 mc14 y, x dll,jdx PR
I AV/T'=PRj4
(percelJt c) (perc~nt c)
I
'-- ~ k
~- -~j- -~§~')- ~=:::- -~.-~-::-
i'----
r---..
5. 0 2. 689 . 38235 1.0 7. 5 3. 551 .31067 10 4.253 .25057 15 .5. 261 . 16087 I. 7li .429
~
------ 20 5. 905 . 09981 I. 616 .404
--- 25 6. 282 . 05281 I. 515 .379
'"
30 6.449 .01498 I. 414 .354 /'--....
35 6.443 -.01617 1. 313 .328 ~
o 40 6.296 -.04210 I. 212
.303 45 6. 029 -. 06373 1.111 .278 50 5.664. -.08168 I. 010 .253 NACAa:QI li5 5.218 -.09637 .909 .227 60 4.706 -.10806 mean line .808 .202 65 4.142 -.11694 .707 .177 70 3.541 -.12307 .606 .152 75 2.916 -.12644 .505 .126 .404 80 2.281 -. 1269..~ .101 85 1. 652 -.12425 .303 .076 90 1. 045 -.11781 .202 .050 95 .482 -.10620 .101 .025 100 0 -.08258- o o -, I- ,--------.-:~------------------
~ -
o
-
NACA MEAN LINE a=0.2
CI,=l.O a,=4.17" Cm,/I=-0.094 !
~',J "'~, 'J I "", I Po I "tv-P"'I
-......;.,
-- :: ~-- --~;al-- -~~~]~i~-!I-~=~~ -~~~~~~-l\
~
I'---..
1.0 2.5 I. 530 .47592 ..............
5. 0 2. 583 . 37661 1. 667 O. 417 7.5 3.443 .31487
~
10 4. 169 . 26803 t-.....
15 5.317 . 19373 , I 20 6.117 . 12405 "'/'--....
25 6. 572 . 06345 1. 563 .391 I-...
30 6. 777 . 02030 1. 459 .365
o
35 6.789 -.01418 1. 355 .339 40 6. 646 -. 04246 1. 250 .313 45 6. 373 -. 06588 1.146 .287 NACA a=QE 50 5. 994 -. 08522 1.042 .260 mean line 55 5. 527 -. 10101 .938 .234 60 4.989 -'.11359 .834 .208 65 4.396 -.12317 .729 .182 J!s 2 70 3. 762 -. 12985 .625 .156 c .
75 3.102 -.13.363 .521 .130 80 2.431 -. 13440 .417 .104 85 l. 764 -. 13186 .313 .078 90 1. 119 -.12541 .208 .052 95 .518 -. JI361 .104 .026 I--- 100 0 -. 0~941 o o t---
......-
-
o
NACA MEAN LINE a=0.3 CI,=1.0 ai=3.84° Cm,,,= -0.106 x y, dy,jdx PR AvjT7=PR/4 (percent c) (percent c)
I
------- --------
---------
'-.....
0 0 ----------- -.--------- ----ii:~553ii-- .5 .389 ~, .75 .546 .60524 !.O -.......
1. 25 .832 .54158 2.5 1.448 .45399
~ .36344
5.0 2.458 ~ 7.5 3.293 .30780 1.538 0.385 10 4.008 .26621 15 5.172 .20246
~
20 6.052 .15068' ~
o 25 6.685 .10278
30 7.072 .04833 35 -,00205 1. 429 .367 7.175 1. 319 NACA a=Q3 40 7.074 -. O~710 .330 -,06492 .302 45 6.816 1.209 mean lioe -.08746 1.099 50 6.433 .275 5.949 -.10567 .989 55 .247 5.383 -.12014 .879 60 .220 Yc 2 c . -.13119 .769 65 4.753 .192 -.13901 .659 70 4.076 .165 -.14365 .549 75 3.368 .137 -.14500 .440 80 2.645 .110 1. 924 -.14279 .330 85 .082 -.13638 90 1. 224 .220 .055 V- I--- 95 .570 "':.12430 .110 .028 I-- ~
.......-
-.09907 100 0 0 0 .8
o .2 1.0
.4
SUMMARY OF AIRFOIL DATA 95
NACA MEAN LINE a=O.4 3.0 -------------------------_._- CI,=1.0
(perc~nt c) (perg~nt c) dy,/dx I PR f>.vjV=P~/4
2.0
---~~-- --~. 3:-- -~~O:6~~~-I-~~~~=- -~~~~~~-
. 75 . 514 . 57105 1. 25 . 784 . 51210 ~ 2. 5 1. 367 . 43106 r--......
.1. 0 2. 330 . 34764 1.0 7. 5 3. 131 . 29671 10 3.824 .25892 1. 429 0.357
~
15 4. 968 . 20185
~ 20 5. 862 . 15682
25 6. 546 . 11733 ~ 30 7. 039 . 07988 35 7.343 .04136 ~ 40 7.439 -.00721 .327 45 7. 275 -.05321 1.310 .298 50 6. 929 -. 08380 1.190 2 .268 NACA a 0A 55 6.449 -.10734 1.071 .238 60 5.864 -.12567 .952 mean line .208 65 5.199 -.13962 .833 .179 70 4.475 -.14963 .714 .149.
75 3.709 -.15589 .595 .2 .119 80 2.922 -.15837 .476 .089 85 2. 132 -. 15683 .357 .238 .060 90 1. 361 -.15062 .030 95 .636 -.13816 .119 o 100 0 -.11138 o c-
r--
-
I--.
0 ~----
-
NACA MEAN LINE a=O 5 CI,=l.O ai=3.04° Cm,/.=-0.139 .0
~~~~~I~er~~nt~J ___ dY,/~ ___
_ ~_J_:~:~~~~
----------- 0 0 ------------- ----------- 0.58195 .5 .345 . 53855 , .75 .485
~ .43360
1. 25 .735 I .0 1.295 .40815 2.5 ~ 2.205 .33070 5.0 2.970 .28365 7.5
~
3.630 .24890 0.333 4.740 .19690 1. 333 ~ 5.620 .15650 6.310 .12180
~
6.840 .09000 7.215 .05930 7.430 .02800 7.490 -.00630 NACA a=D.5 7.350 -.05305 mean line ;300 6.965 -.09765 1.200 1. 067 .267 6.405 -.12550 .233 5.725 -.14570 .933 .200 4.955 -.16015 .800 .167 4.130 -.16960 .667 ; .133 3.265 -.17435 .533 -.17415 .400 .100 2.395 .267 .067 1. 535 -.16850 .133 .033 .720 -.15561>
t--
I--- -.12660 0
t-- 0 0
r--
~ I-- NACA MEAN LINE a=0.6 - ",,=2.58° Cm,I<=-0.158 CI,=1.0 1/, X dy,/dx PR f>.v/V=PR/4 (percent c) (percent c) -------- -------
---- ------ -----
0 ----------- ----------- 0 ------------- .325 0.54825 .5 ~ .455 .50760 1.0 .75 .695 .45615 '1. 25 1. 220 .38555 2.5 2.080 .31325 5.0 2.805 .26950 7.5 ""-., 3.435 .23730
""~
4.495 .18935 5.345 .15250 1. 250 0.312
o
6.035 .12125 6.570 .09310 6.965 .06660
'"
NACA a=O.fJ 7.235 .04060 mean line 7.370 .01405 7.370 -.01435 7.220 -.04700 lie !> 6.880 -.09470 eo C .<; 6.275 -.14015 1. 094 .273 5.505 -.16595 .938 .234 4.630 -.18270 .781 .195 3.695 -.19225 .625 .156 2.720 -.19515 .469 .117 . ;112 1. 755 -.19095 .078
J...-- t---
-
r-- .825 -.17790 .156 .039
I-- 95 0 -.14550 0 0
o .2 .8 1.0
.6
--
xlc
REPOR'l' NO. 824-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS NACA MEAN LINE a=0.7 -- 3.0 , CI,=1.0 Q'i=2.09° C m ./. = -0.179 y, x dy,jdx PH !:J.vjl'=Puj4 (percent c) (percent c)
----- ------ ------ ------- -------
0 0 ------------- ----------- ----------- .5 .305 0.51620 .75 .425 .47795 1. 25 .42960 .655 2.5 .36325 . 1.160 5.0 .29545 1. 955 7.5 2.645 .25450
"-
/.0 10 3.240 .22445 15 4.245 .17995
"'-
20 5.060 .14595
~
25 5.715 .11740 1.176 0.294 30 6.240 .09200 35 6.635 .06840 I 40 6.925 .04570
~
o 45 7.095 .02315
50 7.155 0 -.02455 55 7.090 NACA a=O.l 60 6.900 -.05185 -.08475 65 6.565 mean line -.13650 70 6.030 -.18510 75 5.205 .980 .245 4.215 -.20855 .784 80 .196 Yc " C .C; 85 -.21955 .588 3.140 .147 90 2.035 -.21960 .392 .098 9.1 .965 -.20725 .196 .049 100 0 -.16985 0 - -.............
--
o
-- NACA MEAN LINE a=O.R
---
---
CI,=1.0 ai=1.5·lo C m ,/.=-0.202 EO
(perc~nt ~L (perg:nt c) -' __ ~~~ __ I ___ ~_I~~=p~
0 0 ------------- ----------- ----------- .5 .287 0.48535 .75 .404 .44925 1. 25 .616 .40359 2.5 1.077 .34104 I.p 5.0 1. 841 .27718 7.5 2.483 .23868 10 3.043 .21050 15 3.985 .16892 20 4.748 .13734 25 5.367 .11101 30 5.863 .08775 1. III 0.278
~
35 6.248 .06634
o
40 6.528 .04601
""
45 6.709 .02613 50 6.790 .00620 NACA a=0.8 55 6.770 -.01433 mean line 60 6.644 -.03611 65 6.405 -.06010 70 6.037 -.08790 Yc '" -.12311 C ."- 75 5.514 80 4.771 -.18412 -.23921 85 3.683 .833 .208 2.435 -.25583 .556 90 .139 95 1.163 -.24904 .278 .069 100 0 -.20385 0 !---
-
r---
o
-
---
NACA MEAN LINE a=O.9 Cl,=1.0 Qi=O.90° Cm,/.= -0.225 x 1/, dy,jdx PR t:J.v/'V=Pllj4 (percent c) (percent c)
I
I I
-------
--
0 0 ----------- ----------- ----0:45482-- .5 .269 .75 .379 .42064 1.25 .577 .37740 1.0 .31821 2.5 1.008 5.0 1. 720 .25786 7.5 2.316 .22153
~
.19500 10 2.835 15 3.707 .15595 20 4.410 .12644 25 4.980 .10196
\
o 30 5.435 .08047
5.787 .06084 1.053 35 0.263 6.045 .04234 NACA a=0.9 45 6.212 .02447 6.290 .00678 mean/ine 55 .
6.279 -.01111 60 6.178 -.02965 65 5.981 -.04938
+.2
70 5.681 -.07103 75 5.265 -.09583 -.12605 80 4.714 -.16727 85 3.987 -.25204 90 2.984 -.31463 .526 .132 95 1.503 ~ 0 0 ~ 100 0 -.26086
....-- ~
.8
o .4 .8 1.0
.2 SUMMARY OF AIRFOIL DATA .:J.O NACA MEAN LINE a=1.0 ai=Oo Cl,=l.O Cm",=-0.250 y, x C.O dy,/dx Pn AV/V=Pn/4 (percent c) (percent c) -------- ---,----- ------- ---- ------ 0 ----.---.---- _.--.----- .
0 --.-------- .5 . 250 0.42120 .350 .38875 .75 1. 25 .535 .34770 2.5 .930 .29155 /.0 5.0 1. 580 .23430 7.5 2.120 .19995 10 2.585 .17485 15 3.365 . 13805 20 3.980 .11030 25 4.475 .08745 .06745 30 4.860 .04925 35 5.150
o 0.250
1. 000 .03225 40 5.355 .01595 45 5.475 NACA a=I.O 50 5.515 0 -.01595 55 5.475 mean line 5.355 -.03225 5.150 -.04925 4.860 -.06745 Yo .C 4.475 -.08745 £' 3.980 -.11030 3.365 -.13805 -.17486 2.585 1. 580 -.23430 -----------
100 I 0 ---.--------- -------.---
i---
-
.6 .8 1.0
o .4
---
x/c SUMMARY OF AIRFOIL DATA III-AIRFOIL ORDINATES Page Page K A C A OOOL _- - - - - ___ - - - - __ - ________ .. __________________ 100 NACA 643-218 ______________________________________ .___ 105 KACA 0009______ ____ ____ ___ ____ ___ __________ _______ _ ___ 100 N ACA 64 -418_ _ _ ____________ ____________________ __ _____ 106 N ACA 1408 _______________________ . _________ ___________ 100 N ACA 64 -618 ____ - ____ ._ _____________ __ _______ __ ________ 106 NACA 1410______ _______________________________________ 100 N ACA 644-02L _ _ __________________________________ ____ _ 106 KACA 1412 _____________ - - ______________ .__ ______ __ __ ____ 100 NACA 64 22L _________________________ .. -----__ _ __ ___ __ 106 r NACA 2412 ____ . __ - ____ -- _____ "__ ________ ___ _______ ______ 100 N ACA 6~-42L _ ~ ________ __ __________ __ ____________ _____ 1')6 NACA 241L. _________ - __ __ _ __ _ __ _____ __ _ _____ __ ________ 100 N ACA 65, 3-018 _____ .. _ __ _______________ ___ ____ __ ___ _ ____ 106 N ACA 2418 ___________ - __________________ _______________ 100 N ACA 65,3-418, a=Q.8 _ _____________ .. ______________ _____ 106 NACA 242L __________ ~ _____________ .__ __ ____ _ __ __ _____ __ 100 NACA 65, 3-618 ____ -- _____________________ ~____________ 106 NACA 2424____ __ _ __ _____ ___ _____ ___ __ _ __ _______ _____ _ __ 100 NACA 65(216)-415, a=0.5 _____ .__________________________ 106 NACA 4412_____________________________________________ 101 N ACA 65-006 _________________________________________. __ 106 . NACA 4415 ___________ ~_________________________________ 101 N ACA 65-009 ____ - - __ - __________ ___________________ _____ 107 N ACA 4418_ _ ____ ___ ___ _______________________________ __ 101 NACA 65-206 ________ - ___ __ ___ __ _ __ __ _ __ __ _ __ _ __ __ _ _____ 107 N ACA 442L _____ ____ _ ____________________ __ ______ __ __ __ 101 N ACA 65-209___ __ ____ __ __ __ ___ __ _____ __ ____ __ __________ 107 N ACA 4424____ _______ _______________________ ________ __ _ 101 N ACA 65-210 ____ .. ___ - __ __ __ ___ __ _ __ __ __ _______ ____ ___ __ 107 N ACA 65-410___ ___ __ ___________ __ __ __ ___ _ __ __ ____ ____ __ 107 N ACA 23012 ______ - __ - - - - - - - - - ___ - - - - __ - __ c _ _ _ _ _ _ _ _ _ _ _ _ _ 101 N ACA 23015 _____________________________________ .. __ __ __ 101 N ACA 65\-012_ - _ __ _ __ ___ ___ __ ________ ___ ___ __ _____ ___ __ 107 NACA 23018____________________ ________________________ 101 N ACA 6k212_ _ _ ______________ __ ___________ __ ____ __ _ __ _ 1Q7 N ACA 2302L __ ____ _____________ ________________________ 101 NACA 65\-212, a=0.6 ____________________ --------------c- 107 NACA 23024 _______________________ ~ __ __ ____ ___ _________ 101 N ACA 65\-412_ _ _ __ __ ___ ________ __ ____ __ __ __ _ __ ____ _ __ __ 107 NACA 63,4-420 _____________________________.___________ 102 N ACA 65 -015_ _ _ _______________________________________ 107 NACA 63,4-420, a=;0.3 ____ .______________________________ 102 N ACA 65 -215 _______ - _______________ ___ __ _________ _____ 108 NACA 63(420)-422_ ___ ____ ___ ____ __ _____ _ __ _____ __ __ ____ 102 N ACA 65 415_ - _____ - ______ ,_ __ ____________________ _____ 108 r N ACA 63(420)-517 _____ .. _______________________________ __ 102 NACA 65 415, a=0.5 _____________ ~______________________ 108 r N ACA 63-006 __ "_ ____ ___ __ ________ __ ____________________ 102 NACA 65:>018 ________ .. _____________________________ ,,--- 108 NACA 63-009 ____ ~ ____ -_ _ ___ __ ___ ___ ___ _______ __________ 102 N ACA 65 -218 _________ .. _______________ ____________ _____ 108 N ACA 63-206_ ___ ____ ___________________________________ 102 N ACA 65 -418_ _ _ ______ _____ ____________________ ________ 108 N ACA 63-209 _______________________________ . ___________ c 102 NACA 65 -418, a=0.5 _____ "______________________________ 108 NACA 63-210_ ______ __ ____ __ __ ____ __ ___ ___ __ _ __ _________ 102 N ACA 65 -618_ _ _ ______ __ __ ______ ______ ____________ _____ 108 NACA 63\-012 ____________________ .. ______________ ------- 102 NACA 65 -618, a=0.5 _______________________________ ._____ 108 N ACA 63\-212_ _ _ ____ _____________ _________ __ ___________ 103 N ACA 65 -021.. _ _ __________________________________ _____ 108 N ACA 63\~412_ _____ ______ __ __ _________________________ _ 103 N ACA 65 -221.. - _____ - - - ___ - ___________ - ___________ _____ 109 N AC A 63 O15 ____________________________________ ~ _____ 103 N ACA 65 -421.. ___________ .. ______________________ c _ _ _ _ _ _ 109 r 4 N ACA 63 215_ __ _______________________________________ 103 NACA 65 421, a=0.5____________________________________ 109 r r N ACA 63 415_ ______________ __ _________________________ 103 N ACA 65(211)-114.. __ - __ - _____ - ___________ - ____________ __ __ 109 r N ACA 63 615_ __ ___ ________ __ ___________ _______________ 103 N A C A 65(12!)-·420 _______ - ___ .. ____________ - _____ .. ___ .. _ ____ _ 109 r N ACA 63 -018_ __ __ __________________________________ ___ 103 N ACA 66,1-212 _______ - __ .. __ - - - ________ - - ___________ ___ __ 109 NACA 63 -218 _______________ ~_ _________ ____ _________ ___ 103 N ACA 66(215)-016 _____________________ - ___________ _ __ __ 109 NACA 633-418 ___ .. __ ~____ _______________________________ 103 N ACA 66(215)-216_ - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - c 109 N ACA 63 -618_ __ ________ __ _____________________________ 103 NACA 66(215)-216, a=0.6_______________________________ _ 109 N ACA 63 -021.. ________________________________. _________ 104 N ACA 66(215)-416 _________ - - .. _- - __ - __ - - _- - __ - - ____ - - ___ 109 N ACA 63 22L ____________________________________ __ ___ 104 NACA 66-006___________ ________________________________ 110 r N ACA 66-009 _________________________ .. - __________ ______ 110 N ACA 63 -421.. __________ . ___ . __ . _______ ___ _____________ 104 N ACA 66-206 ________________ - ________ - - ________________ 110 NACA 64-006______ ______ ____ ________ ___ ________________ 104 N ACA 66-209 ___________ .. ___ - _________ - _- ________ ______ 110 NACA 64-0'09___________________________________________ 104 NACA ·64-108 __________________ . ___ . ____ ._ _ __ ______ __ ___ __ 104 N ACA 66-2lO_ ____________________________________ ______ 110 NACA 64-110 _________________ c__ __ _ _ _ __ _ _____ __________ 104 NACA 66\-012_ _ _ _ __ ____ ___ _____ ____________ ______ _ _____ 110 NACA 66 -212_ _ _ __ _ __ __ _ ___ ___ ___________________ ______ 110 N ACA 64-206_ ____________________ __ _________________ ___ 104 NACA 64-208 ____________________ .. __ . _ . _____ .. _ __ __ ___ ____ 104 NACA 66 015 _______________ - _________ - _- ________ ______ 110 r N ACA 66 -215 _______________ - ___________ - ___________ - __ 110 N AC A 64-209 _________ __________________________________ 104 NACA 64.,-210. ______ . ___ . _______________ ._. ______________ 105 N ACA 66 415 ___________ " _- _- _________ - _- - __ - _____ __ ___ 110 r NACA 64\-012 _____________ . ________________ .. ________ ._ _ 105 NACA 66:>018_ _ _ __ ___ __ __ ________________ _ ______ _ __ ____ 111 NACA 64\-112 __________ .______________________ _________ 105 NACA 66:>218_.______ ____________________ _______________ 111 NACA 64\-212_ _ _ __________ _____ _ ______ ___ ______________ 105 NACA 66 -418 _________ ._ _____ _____________ __ _ __ ____ _ ___ III NACA 66 02L _ _ ___ __ _ _____ ________________ ______ _ ___ __ 111 N ACA 64\-412 _______________ .. ____ .. _ .. _____ _ _____________ 105 r N ACA 66~-22L _ _ _ ________________ __ _____________ _______ 111 N AC A 64 O15 ____________ . ____ _________________________ 105 r NACA 64 -215 __________________________________ .. _______ . 105 N ACA 67,1-215 _____________________________ .. _ ___ ____ __ __ 111 NACA 64 -415 _____________________ " ____________. _____ ~__ 105 N ACA 747 A315 _______________________ .. _- __________ ~ _ __ __ 111 NACA 64:>018 __________________________ "_______________ 105 N ACA 747 M15 ______ ________________ ,,_ __ _______ ___ __ _ __ _ III
I
>l>- ...... t"' >-l >- !:d I':l "d o ~ Z ? 00 ~ z ~ o Z >- >- t:l -< ...... Ul o !:d >1 C":l o ~ ~ ...... >-l I':l I':l >:rJ o !:d ~ !:d o Z ~ ...... C":l Ul
o o
-
of
of I
0 0.05 0 0 0.10 -.646 -.126 percent percent -1. -2.491 -3.318 -3.857 -4.242 -4.733 -4.986 -5.081 -5.064 -4.803 -4.321 -2.913 -2.066 -1.141 -3.646 -7.692 -8.465 -9.450 -9.959 -9.606 -4.130 -2.280 -1.292 -3.675 -4.965 -6.614 -8.644 -7.347 -5.824 Ordinate Ordinate surface surface ~10.155 -10.124 --- --- In E.: in E.: L. L.
Lower 0 1.342 2.622 5.155 7.670 Lower 0 1.615 2.988 5.620 8.180 given 10.176 15.167 20.143 25.111 30.075 40.000 49.971 59.949 69.939 79.942 89.960 94.975 given 10.700 15.667 20.573 25.445 30.300 40.000 49.882 59.797 69.756 79.767 89.839 94.902
1412 Station 100.000 Station 100.000
--- ---- chord] through through 1.58 6.33 954 753 930 .966 airfoil chord] .126 airfoil 0 1. 2.733 3.786 4.537 5.118 5.951 6.486 6.799 6.940 6.803 6.267 5.453 4.413 3.178 0 3.892 5.449 7.552 9.052 6.352 3.502 1.
ordinates '1. ordinates 8.824 10.215 11.888 12.959 - 13.593 13.874 13.606 12.532 ------- 10.903 Ordinate Ordinat3 --- radius radius surface surface and radius: of radius:
NACA NACA and
E. E.
0 1.158 2.378 4.845 7.330 9.824 .885 14.833 19.857 24.889 29.925 40.000 50.029 70.061 80.058 90.040 95.025 0 2.012 4.380 6.820 9.300 Upper 60.051 Upper 100.000 14.333 19.427 24.555 29.700 40.000 50.U8 60.203 70.244 80.233 90.161 95.098 L. Slope L. Slope of Station Station 100.000 ---- ---- [Stations [Stations of I I I of
l
--- 0 0 0 0.05 0.10 -.901 -.512 -.105 ~2.82 -4.02 -5.51 -6.48 -7.18 -8.05 -8.52 -8.67 -8.62 -8.16 -7.31 -6.17 -4.87 -3.44 -1.88 -1.06 (-.22) percent percent -1. -2.055 -2.726 -3.157 -3.462 -3.844 -4.031 -4.091 -4.064 -3.836 -3.439 -2.914 -2.304 -1.629 Ordinate Ordinate surface surfac~
---- ---
in E.: in E.: L. L.
0 1.2& 5.0 7.5 2.5 Lower 0 1.326 2.602 5.130 7.642 10 15 30 40 50 70 80 90 95 Lower 20 25 60 given 10.146 15.139 20.120 25.093 30.063 69.949 79.951 given 100 100 40.000 49.975 59.958 89.966 94.979
1410 Statiou 100.000
2421 Station
--- -- through through 1.10 76 4.85 (.22) .832 .105 3.87 5.21 7.00 8.29 9.28 9.79 7.94 5.74 3.18 1.
airfoil chord] airfoil chord] 0 1.639 2.297 3.194 3.837 4.338 5.062 5.531 5.809 5.940 5.836 5.385 4.692 3.804 2.741 1. 10.70 11.59 12.15 12.38 12.16 11.22 ----- ordinates ordinates Ordinate --- Ordinate ------ radius radius surface surface
NACA and radius: of and radius: of
NACA
E.
E.
1.174 2.398 4.870 7.358 9.854 0 1. 2.5 5.0 7.5 Upper 14.861 19.880 24.907 29.937 40.000 50.025 60.042 70.051 80.049 90.034 95.021 10 15 20 25 30 40 50 60 70 80 90 95 Upper 100.000 100 100 L. Slope Statiou L. Slope Station ---- ---- [Stations [Stations of I I of
-I
0 0 -.87 0.05 0.10 -2.45 -5.48 -6.74 -7.09 -7.12 -6.71 -3.97 -2.80 -1.53 (-.19) percent -.659 -.378 -.084 -3.44 -4.68 -6.03 -7.18 -5.99 -5.04 percent -1.200 -2.134 -1.620 -2.458 -2.682 -2.953 -3.074 -3.101 -3.063 -2.869 -2.556 -2.153 -1.693 -1.193 Ordinate Ordinate .
surface surface --- --- in in E.: E.: ----- L.
L.
1. 2.5 5.0 7.5 0 1.311 15 Lower 2.582 5.104 7.614 Lower 0 10 20 25 30 40 50 60 70 80 90 95 given 20.096 25.074 given 10.117 15.111 30.050 40.000 49.980 59.966 69.959 79.961 89.973 10.0 100 Station 100.000 Station
. 94.984 241
---- --- --- through through
55 -
0.70 3.56 (.19) 7.17 9.34 5.08 2.81 airfoil chord] .698 .084 airfoil chord] 4.415 6.03 8.05 9.89 8.65 7.02 1.
0 1.324 1.862 2.602 4.819 4.939 -3:28 10.15 10.65 10.88 10.71 ordinates 3.138 3.558 4.171 4.574 4.869 4.502 3.931 3.193 2.305 1.271 ordinates Ordinate --- . Ordinate ---.- radius surface radius surface radius:
NACA and radius: of and
NACA
E. E.
0 1.189 2.418 4.896 7.386 9.883 0 1.25 7;5 2.5 5.0 14.889 19.904 24.926 29.950 40.000 50.020 10 15 50 70 80 90 95 Upper 60.034 70.041 80.039 90.027 95.016 Upper 20 25 30 40 60 100 100 100.000 L. L. Slope·of Station Slope Station
------ ---- ---
-------
[Stations [Stations of i of I I , I I 0 0 0 -.60 -.68 0.10 -2.67 (-.16) -1.42 -1.96 -3.15 -3.51 -4.01 -4.30 -4.46 -4.50 ~4.35 -3.97 -3.42 -2.75 -1.97 -1.09 (-.10) -2.06 -2.86 -3.84 -.4.47 -4.90 -5.42 -5.66 -5.70 -5.62 -5.25 -4.67 -3.90 -3.05 -2.15 -1.17 percent percent Or<'Jnat~· snrface Ordinate' --- surface ---- In in E.: L.
0 1.25 5.0 7.5 2.5 Lower 0 1.25 2.5 5.0 7.5 10 15 30 40 ' 50 70 80 90 95 Lower 20 25 60 given 10 15 20 25 30 40 50 60 70 90 95 given 100 100 Station 100 100 StatIOn
0009 241
--- ---
chord] through 0.89 25 34 2.48 .60 0 1. 2.67 (.10) (.16) airfoil chord] ·1.96 3.15 3.51 4.01 4.30 4.46 4.50 4.35 3.97 3.42 2.75 1.97 1.09 0 3.71 5.07 6.06 6.83 7.97 8.70 9.17 9.38 9.. 8.57 7.50 6.10 4.41 2.45 1.
airfoil -2~7i ----- ordinates ordinates Ordinate ~---- Ordinat~
---
radius surface surface and radius: of
NACA radius: NACA and
E.
E.
0 1. 2.5 5.0 7.5 0 1.25 5.0 7.5 2.5 10 15 20 25 30 40 50 60 70 80 90 95 10 15 30 40 50 70 80 90 95 Upper Upper 20 25 60 100 100 100 100 L. L. Slope Station Station
------ ---- ---
[Stations [Stations
of , of I I , I I I , I I
I i
---- 0 o 0 -.9(> -.72 -.40 -.82 -.48 0.10 -1.31 -1.78 -2.10 -2.34 -2.67 -2.87 -2.97 -3.00 -2.90 -2.65 -2.28 -1.83 -1.31 (-.06) -1.65 -2.27 -3.01 -3.46 -3.75 -4.10 -4.23 -4.22 -4.12 -3.80 -3.34 -2.76 -2.14 -1.50 (-.13) percent percent Ordinate Ordinate .
surface surface
--- ---
In in E.: -- L.
1. 5.0 0 2.5 7.5 Lower 0 1.25 2.5 5.0 7.5 10 15 30 40 50 70 80 90 95 Lower 20 25 60 given 10 15 20 25 30 40 50 70 80 90 95 given 100 100 Station 100 100 Station
0006 241
.
--- ----
chord] chord] --' through 0.40 67 1.58 ..
.95 .72 .40 (.06) (.13) airfoil 0 1.31 1.78 2.10 2.34 2.67 2.87 2.97 3.00 2.90 2.65 2.28 1.83 1.31 0 airfoil 2.99 4.13 4.96 5.63 6.61 7.26 7 7.88 7.80 7.24 6.36 5.18 3.75 2.08 1.14 -2~i5 ----- ordinates ordinates Ordinate Ordinate
--- ---
radius surface surface radius: radius: of
NACA and NACA and
E. E.
1.25 2.5 5.0 7.5 1. 5.0 0 0 2.5 7.5 10 15 20 25 30 40 50 60 70 80 90 95 10 15 25 30 40 50 70. 80 90 95 Upper Upper 20 60 100 100 100 ioo L. L. Slope Station Station
---~--- ----
---
[Stations [Stations I--' o r-'- "l >- .... ~ o .... t" t:1 ~ >- o ~ ~ >- ~ o Ul - 0.305 0 0 0.20 -1.504 -- 0 percent of -4. -6.860 -9.703 -9.482 -8.242 -6.664 _4.803 -2.673 ~.964 -3.303 -5.862 -7.647 -8.852 Ordinate -10.223 -10.454 -10.278 percent of -6.931 -7.512 -8.169 -8.416 -8.411 -7.'606 -6.698 -5.562 -4.312 -3.003 -1.655 -3.472 -4.656 -6.006 -8. ---- Ordinate --- in E.: in L.E.: L.
970 389 0 2.223 3.669 6.147 8.399 Lower surface 24.738 29.735 49.766 59.798 69.838 79.884 89.936 94.964 10.577 14.999 19.747 39.744 1. 3.464 6.225 8.847 100 Lower surface 0 25.889 30.599 40.000 49.765 59.595 69.513 79.536 89.680 94.804 11. 16.326 21.142 Station 100.000 Station ---- 4424 23024 ---~ ------ 6.33 -.---~ ------- 546 724 6.33 049 008 airfoil chord] 7.988 5.687 3.115 1. ----- 5.764 9.884 0 4.017 8.172 13. 11.690 10.
airfoil chord] 7.447 4.099 2.240 ----- ordinaws given 11. 12.528 13.237 13.535 12.928 9.651 0 3.964 5.624 7.942 Ordinate radins through ordinaws given 11.012 13.045 H.416 15.287 15.738 15.606 14.474 12.674 10.312 ---- Ordinaw radius through surface radius: of surface radius: of and NACA and NACA E.
.277 E. 1.331 3.853 6.601 9.423 30.265 50.235 60.202 70.162 80.116 90.064 95.036 15.001 20.253· 25.262 40.256 .530 Upper L. Slope 1.536 3.775 6.153 8.611 0 StatioD Upper 29.401 40.000 50.235 60.405 70.487 80.464 90.320 95.196 L. Slope 13.674 18.858 24.111 100.000 Station ------- ---- [Stations [Stations - - 30 95 0.305 (-.22) -8.76 -8. -8.83 -8.14 -7.07 -5.72 -4.13 -2.30 -1.30 0 -.74 -2.08 -3.14 -4.52 -5.55 -6.32 -7.51 -8.
0.20 percent of -5.34 -4.40 -3.35 -2.31 -1.27 (-.22) -3.48 -4.78 -5.62 -6.15 -6.7.1 -6.9g -6.92 -6.76 -6.16 -2.42 Ordinaw percent of --- E.: Ordinaw ..
---- in E.: - in L.
L. - - 0 1. 2.5 5.0 7.5 Lower surface 15 20 25 30 40 50 60 70 80 90 95 2.5 5.0 7.5 given 100 100 0 1.25 Lower surface 40 50 60 70 80 90 95 Station 10 15 20 25 30 100 100 ---- Station 23021 ---- 4.85 4.85 2.76 1.53 (.22) 8.90 7.09 5.05 6. 7.93 ----- (.22) airfoil chord] .9.13 11.19 11.80 12.05 12. 11.49 10.40 9.50 6.91 3.85 2.11 -4~87 10.03 4.45 .5.84 7.82 9.24 ----- ordinates airfoil chord] 12.04 13.88 1!.27 14.16 13.18 11.60 10.35 13.17 Ordinate --- -- ordinaws given Ordinaw radius through surface --- radius: surface of and NACA and NACA E.
21; .5 2.5 5.0 7.5 E. radius: 0 1.
. 30 50 60 70 80 90 95 10 15 20 25 40 1. 2.5 5.0 7 100 L. Slope of radius through 0 Upper 30 40 50 60 70 80 90 95 10 15 20 25 100 Station Upper 100 L. Slope Station ---- - ---- [Stations [Stations 94 09 0.305 -1. (-.19) -6.18 -7.27 -7.47 -7.37 -6.81 -5.94 -4.82 -3.48 -1.
0 -1.83 -2.71 -3.80 -4.60 -5.22 -6.86 -.93 -.55 0 0.20 -2.45 (-.19) -5.49 -5.49 -5.26 -4.70 -4.02 -3.24 -1.67 -2.11 -2.99 -4.06 -4.67 -5.06 -5.56 Ordinate percent. of ---- Ordinate ---- E.: in 0 1. 2.5 5.0 7.5 Lower surface 21\ 30 40 50 60 70 80 90 95 10 15 20 100 100 2.5 5.0 7.5 0 1.25 Lower surface 50 6a 70 80 90 95 Station 10 15 20 25 30 40 100 --- Station --- 32 3.56 3.56 (.19) 55 89 4.40 2.39 1.
9.05 7.75 6.18 4.09 5.29 6.92 8. 8.83 9.86 airfoil chord] 10.56 10.55 10.04 (.19) 10.36 8. 6.22 3.46 1.
5.00 6.75 8.06 9.11 ordinaws given in percent of airfoil chord] 12.70 11.85 10.44 10.66 11.72 12.40 12.76 radius through L. E.: ------- OrdiDaw ordinaws given --- --3~76- surface Ordinate I radius: of surface and radius: NACA E.
NACA and 25 7.5 E. 0 1.25 2.5 5.0 50 60 70 80 90 95 10 15 20 25 30 40 100 I,. Slope 1. 2.5 5.0 7.5 Upper 0 80 95 15 30 40 50 60 70 90 10 20 25 Station 100 100 L. Slope of radius through L.
Upper Station ---- ------- [Stations [Stations 0 -.90 71 0.305 (-.W) -5.96 -5.50 -4.81 -3.91 -2.83 -1.59 -3.61 -4.09 -4.84 -5.41 -5.78 -5.92 ----- 0 0.20 -1.M -2.25 -3.04 percent of 0 -.57 -.36 (-.16) Ordinaw -2.14 -1.55 -1.03 5 percent of -2.48 -3.27 ~3. -3.98 -4.18 -4.15 -3.98 -3.75 -3.25 -2.72 --- E.: -179 Ordinaw in --- E.: 5 L.
in ...
L.
0 1.25 2.5 5.0 7.5 Lower surface 50 70 80 90 95 10 15 20 25 30 40 60 100 100 1. 2.5 5.0 7.5 Lower surface 0 Statio!l 40 50 60 70 80 90 95 10 15 20 25 30 100 100 Station --- 2.48 2.48 (.16) 7.74 6.61 5.25 3.73 2.04 1.12 4. 5.89 6.90 7.64 8.52 8.92 9.08 9.05 8.59 --------- ----- airfoil chord] (.16) -3~34 9.30 7.63 5.55 3.08 1.67 4.17 5.74 6.91 7.84 9.27 3.07 ordinates given 11.25 10.53 airfoil chord] 10.25 10.92 11.25 Ordinate ordinates given surface radius through Ordinate ---- radius: surface of and radius: NACA E.
and NACA 25 0 1. 2.5 5.0 7.5 E. 70 80 90 95 15 20 25 30 40 50 60 100 100 L. Slope of radius through 5.0 7.5 Upper 0 1. 2.5 50 60 70 80 90 95 10 15 20 25 30 40 Station Upper 100 L. Slope ---- Station ---- ------ [Stations [Stations 0 -.70 0.305 -3.00 -2.16 -1.23 (-.13) -3.50 -3.97 -4.28 -4.46 -4.48 -4.17 -3.67 -1.23 -1.71 -2.26 -2.61 -2.92 0 0.20 0 -.65 -.39 -.22 -.16 Ordinate (-.13) 2 -1.40 -1.00 --- E.: -2.49 -2.74 -2.86 -2.88 -2.74 -2.50 -2.26 -1.80 ---- -1.43 -1.95 Ordinate ---- E.: L.
L.
0 1. 2.5 5.0 7.5 70 80 90 95 Lower stuface 15 20 25 30 40 50 10 .60 100 100 2.5 5.0 7.5 0 1.25 Station Lower surface 50 60 70 80 90 95 10 15 20 25 30 40 Station --_.
1.58 .92 1.58 (.13) 7.14 6.41 5.47 4.36 3.08 1.68 5.80 6.43 7.19 7.50 7.60 7.55 2.67 3.61 4.91 ----- airfoil chord] ordinaws given in percent of (.13) 9.19 8.14 6.69 4.89 2.71 1. radius through airfoil chord] 3.39 4.73 5.76 6.59 7.89 8.80 9.41 9.76 9.80 0 2.44 ------- Ordinaw ------ ordinates given in percent of surface Ordinate of --- radius: surface and radius: E.
NACA and NACA 7.5 0 2.5 5.0 E. 1.25 60 70 80 90 95 10 15 20 25 30 40 50 100 L. Slope 2.5 5.0 7.5 Upper 0 1.25 . ---- 50 60 70 80 90 95 10 15 20 25 30 40 Station- Upper 100 L. Slope of radius through Station ---- ---- [Stations [Stations
I
'"d 00 t>:) H>- z e3 .... o Z ~ ~ -< .... [IJ. o ~ (l S "" ~ .... ..., ..., t'i t'i ><j o !:d ~ g; Z >- d >-3 .... (l [IJ.
!:d t'i o !:d >-:3 Z ?
i-< o ~
I
J
704 274 982 920 935 902 0 5. 5.
0 0 0 -.503 -.609 -.771 -.683 -.383 -.138 -.985 -.707 -.250 percent of -1.057 -1.462 -1.766 -2.010 -2.386 -2.656 -2.841 -2.954 -3.000 -2.971 -2.877 -2.723 -2.517 -2.267 -I. -1.670 -1.342 -1.008 -1.194 -1.519 -2.102 -2.925 -3.542 =t~~~ -5.342 -5.712 -5.930 -6.000 - - -5.370 -4. -4.420 -3.840 -3.210 -2.556 -I. -1.
__ Ordinate
Ordinate surface ---- ---- in .5 .75 .5 .75 5.0 0 1.25 2.5 5.0 7.5 0 1.25 2.5 7.5 45 55 Lower 10 15 20 25 30 35 40 45 50 55 60 65 70 75 80 85 90 95 Lower surface ~~ 20 25 30 35 40 50 60 65 70 75 SO 85 90 95 100 100 Station Station - --- ---- I .
63-006
631-012
930 920 704 420 210 902 1.087 0.297 935 274 766 982 670 342 - .985 .707 .250 .503 .609 .771 .683 .383 .138 ].194 ].519 2.925 5.342 5.712 5. 6.000. 5. 5. 5.370 4. 3.840 3. 2.556 1. 1. 0 airfoil chord] airfoil chord] 0 2.102 3.542 g~~ 4.
2.841 2.971 2.877 2.517 I. 0 0 1.057 1. 1. 2.010 2.386 2.656 2.954 3.000 2.723 2.267 I. I. 1.008 ordinates given Ordinate Ordinate --- ---- surface surface radius: radills: and .5 .75 .5 .75
NACA
2.5 1.25 2.5 5.0 7.5 0 1.25 5.0 7.5 E. NACA o E.
20 25 30 40 4.1 55 60 65 70 75 85 90 95 ~~ 20 25 30 35 40 45 50 55 60 65 70 75 SO 85 90 95
10 15 35 50 80 100 100 Upper - Upper L.
L.
Station Station
---- ----
[Stations [Stations and ordinates given in percent of
I I I I I I I I I I
.150 .367 .102 0 0.211 0 0 0.0842 -.897 -.304 -.776 -.967 -.822 -.415 -.087 -1.301 -1.562 -1.941 -2.568 -3.386 -3.984 -4.466 -.1.178 -5.653 -5.940 -6.027 -.1.903 -5.621 -5.223 -4.738 -4.184 -3.569 -2.917 -2.239 -1.554 -1.165 -1.567 -2.121 -2.524 -2.843 -3.319 -3.648 -3.857 -3.966 -3.970 -3.867 -3.671 -3.393 -3.045 -2.644 -2.204 -1.740 -1.271 ordina~1 Ordinate
---
E.: E.: L. L.
.800 .570 .831 I. 1.634 5.489 I. 5.114 7.618 Lower surface 0 2.942 8.004 Lower surface 0 2.602 10.503 15.473 25.358 30.285 35.209 40.134 45.064 50.000 54.945 59.899 64.865 69.843 74.834 79.841 84.863 89.900 94.950 lO.ll8 15.lIO 20.098 25.083 30.067 35.049 40.032 45.015 50.000 54.987 59.976 64.968 69.964 74.962 79.964 84.970 89·979 94.990 20.422 100.000 Station 100.000 Station ---
63-210
through 0.13 2.283 939 761 121 0.770 .876 .530 63(420)-51 0 1.551 1.912 2.477 3.498 4.966 7.050 8.542 9.660 8.92.1 8.067 7.099 6.030 4.877 3.668 2.434 1.213 0 0 1.107 1.379 1. 2.753 3.372 3.877 4.665 5.240 5.647 5.910 6.030 6.009 5.861 5.599 5.235 4.786 4.264 3.684 3.061 2.414 1. 1. 0 airfoil chord] 6.104 9.633 airfoil chord] 10.416 10.887 11. 10.977 10.699 10.254 ordinat.:ls given in percent of ordinatl's given in pe.rcent of Ordinate Ordinate --- surface surface radius: radius:· and and
NACA
.669 E. .430 E.
.200 0 0 .412 .866 1.162 2.398 4.886 7.382 9.882 19.902 44.985 60.024 65.032 75.038 Upper 2.058 4.511 6.996 9.497 Upper 14.890 24.917 29.933 34.951 39.968 50.000 55.013 70.036 SO. 85.030 90.021 95.010 19.578 24.642 29.715 34.791 44.936 70.157 100.000 14.527 39.866 50.000 55.055 60.101 65.135 75.166 80.159 S5.137 90.100 95.050 L. Slope of radius L. Slope of radius through Station 100.000 Station
NACA
------- ---- ----
[Stations [Stations I I I 1 i of - 898 486 071 .083 .120 0 0 0 0.168 0 0.0842 -.696 -.833 -.675 -.317 -.033 -.445 percent -1.041 -1.878 -2.229 -2.917 -3.200 -3.379 -3.470 -3.470 -3.376 -3.201 -2.953 -2.644 -2.287 -I. -I. -1.
-1.759 -2.122 -2.660 -3.568 -4.786 -.1.691 -6.428' -7.539 -8.305 -8.797 -9.002 -8.914 -8.599 -8.113 -7.495 -6.767 -5.943 -5.049 -4.100 -3.120 -2.145 -1.226 -1. -2.505 --- Ordinate Ordinate --- --- E.: E.: in L. L.
.S13 .563 .820 Lower surface 0 I. 1.650 2.959 5.508 8.023 Lower surface. 0 1.330 2.592 5.103 7.606 10.522 15.491 20.437 25.370 30.295 35.216 40.139 45.066 50.000 54.943 59.896 64.860 69.837 74.828 79.835 84.8.18 89.897 94.949 10.106 15.099 20.088 25.075 30.060 35.044 40.029 45.014 50.000 54.988 59.978 69.967 74.966 79.968 84.973 89.981 94.991 ·64.971 100.000 Station 100.000 Stat.ion ---- ---
63-209
through through 959 147 3.82 0.681 255 765 .973 .512 63(420)-422 0 1. 2.402 3.088 4.312 6.050 7.387 8.496 9.169 7.988 6.700 5.329 3.918 2.513 1.181 0 .796 airfoil chord] airfoil chord] 12.377 12.890 12.883 12.493 11.907 11. 0 11.489 13.034 10.227 0 1. I. 2.510 3.077 3.539 4.263 4.792 5.169 5.414 5.530 5.518 5.391 5.159 4.834 4.429 3.958 3.430 2.861 2.267 1.668 1.067 ordinates given in percent of ·10.231 ordinates given Ordinate --- radius IOrdinate --- radius snrfflce of radius: radius: and and
NACA
Ie,S E. E.
0 .187 .850 .437 .680 3.98 2.041 4.492 6.977 9.478 0 1.170 2.408 4.897 7.394 9.894 Upper UppP,' surface 14.509 19.563 24.630 29.705 34.784 39.861 44.934 50.000 55.057 60.104 65.140 70.163 75.172 80. 85.142 90.103 95.051 14.901 19.912 24.925 29.940 34.956 39.971 44.986 50.000 55.012 60.022 65.029 70.033 75.034 80.032 85.027 90.019 95.009 L. Slope J,. Slope of Station Station 100.000 100.000
NACA
------ --- ------
[Stations [Stations
I , I i I I ' i I i I I I I
of 970 782 620 712 859 946 .134 .178 0 0 0 0.262 0 -.272 -.010 0.0842 percent of -.856 -.451 -.537 -.662 -.869 -.952 -.698 -.447 -.212
=~:~b perccnt
-1.502 -1.805 -2.244 -2.968 -3.914 -4.601 -5.158 -5.996 -6.570 -6.948 -7.125 -7.108 -6.934 -5.756 -5.189 -4.553 -3.856 -3.111 -2.337 -1.568 -1.144 -1.341 -1.492 -1. -1. -1. -1.982 -I. -1.900 -I. -1. -1.422 -1.196 Ordinate I -_-1 Ordinate surface surface --- in E.: E.: in L. L.
3.03 .935 .042 .797 wwer 0 1. 1.809 3.144 .5.712 1.
8.229 J.lowcr 0 2.562 5.068 7.571 10.720 15.6.53 20.542 25.396 30.192 34.992 39.855 44.757 49.692 54.656 59.647 64.661 69.695 ' 74.744 79.803 84.866 89.927 94.975 20.059 25.050 30.040 35.030 45.009 50.000 54.992 59.985 64.980 69.977 74.927 79.978 84.981 89.987 94.994 given 10.070 15.066 40.019 100.000 100.000 Station Station ___ -------- ----
63-206
through through
63A-420
3.16 0.297 241 776 356 .728 804 .551 .677 .876 .900 .454 airfoil chord] 0 1.814 2.241 5.878 9.497 5.807 4.453 3.108 1.836 0 ordinates given 2.912 4.128 7.237 8.366 8.357 7.120 airfoil chord] a=0.3 I. 2.189 2.526 3.058 3.451 3.736 3.926 4.030 4.042 3.972 3.826 3.612 3.338 3.012 2.642 2.237 I. ]. 0 10.132 11.4IO 12.296 12.781 12.848 12.594 12.089 II. JO.516 0 I.
ordinat", Ordinate Ordinate ---- radius surfncc surface and radius: radius:
NACA
197 E.
E.
.065 .260 .691 0 .458 .703
NACA
1.856 4.268 6.771 9.280 I. 2.438 4.932 7.429 9.930 Upper Upper 14.347 19.4.18 24.604 29.808 35.00S 40.145 45.243 50.308 55.344 60.353 65.339 70.305 75.256 80.197 85.13-1 90.073 95.025 . 14.934 19.941 24.950 29.960 34.970 39.981 44.991 50.000 55.008 60.015 65.020 70.023 75.023 80.022 85.019 90.013 95.006 1,. Slope of L. Slope of radius Station Station 100.000 100.000
------ --- ---
------ [Stations [Stations and of I I I : I I ' I i I - 007 471 582 196 .133 0 0.168 0
-.992 -.311 -.749 -.906 -.990 -.550 0. --
p('rcent of percent -2.458 =g~ -3.210 -4.293 -5.749 -6.732 -7.834 -8. -7.916 -7.622 -7.176 -6.613 -5.953 -5.208 -4.403 -3.MO -2.673 -1.806 -1.151 -1. -2.196 -2.655 -3.024 -3.591 -3.997 -4.275 -4.442 -4.500 -4.447 -4.296 =gg~ -3.358 -2.928 -1.966 -1.
-5.097 -7.405 Ordinate Ordinate
:~1.590 ----
in E.: in I,.
.5 .75 .785 1. 2.5 5.0 7.5 J,ower surface 0 1.070 1. 2.918 5.462 7.976 Lower surface 0 10.474 15.446 25.337 30.268 40.126 .10.000 54.948 64.873 69.852 74.844 79.850 94.953 riven 20 25 30 35 4.5 55 60 65 70 75 80 85 90 95 20.397 35.197 45.060 59.905 84.871 89.906 10 15 40 Station 100 Station . 50 -
--- ---
63-009
63,4-420
131 3.16 0.631 1.790 2.196 5.557 7.817 8.523 7.438 1. .749 .906 .990 .550 .196 airfoil chord] 0 2.827 3.954 6.793 9.424 9.492 6.253 4.990 3.684 2.379 0 airfoil chord] 11.414 11.895 11.906 11.025 10.333 0 1.151 I. 2.196 2.655 3.024 3.591 3.997 4.275 4.442 4.500 4.447 4.296 4.056 3.739 3.358 2.928 2.458 1.966 1.471 0 10.589 12.036 11.556 ordinates ordinates given Ordinate Ordinate ---
---
surface suriace radius: radius: and and
NACA
.215 .430 .887 .5 .75 E.
E.
2.082 4.538 7.024 9.526 0 1.25 2.5 5.0 7.5
NACA
14.554 19.603 24.663 29.732 34.S03 39.874 44.940 50.000 55.052 60.095 65.127 70.148 75.156 80.150 85.129 90.094 95.047 10 15 20 25 30 35 40 45 50 55 60 tiS 70 75 80 85 90 95 Upper 0 Upper 100 100 L.
L. Slope of radius through Station Station
---------
--- ---- [Stations [Stations ><: o >rj
U2 c:l ~ ~ ~ :> ..... ;0 "1 o ..... t" C; :> ..., :>
I-'- 8
, I , I I - , I .. .184 .333 .238 .571 .603 0 0
5 8 0 0 0.2527
-.716 -.193 -.960 -.297 0.1685 -1.087 -1.305 -1.646 -2.220 -3 -3:565 -4.009 -4.656 -5.095 -5.361 -5.474 -5.439 -5.243 -4.909 -4.459 -3.918 -3."311 -2.660 -1.989 -1.327 -1.211 -1.849 -2.500 -3.372 -3.998 -4.484 -5.642 -5.903 -5.906 -5.630 -5.197 '-4.6.33 -3.971 -3.241 -2.475 -1.702 Ordinate -1.458 -5.181 -5.990 Ordinat~ --- --- E.: K: .700 .975 .844 Lower surface' 0 1.509 2.802 5.340 7.853 1.
Lower surface 0 1.139 3.035 5.607 8.132 10.353 15.331 20.295 25.250 30.200 35.148 40.095 45.045 50.000 54.961 59.930 64.907 69.894 74.891 79.898 84.915 89.941 94.972 10.63.3 20.531 25.451 54.931 59.875 69.813 79.822 84.853 94.952 15.596 30.360 35.266 40.171 45.081 50.000 64.836 74.809 89.897 Station 100.000 Station 100.000 ---
632-41 633-61
1.594 878 273 541 655 293 2.120 -- airfoil chord] .931 0 1.511 1. 2.491 3.616 5.268 6.542 7.586 9.219 9.667 8. 7.534 6.330 5.073 3.800 2.531 1. 0 airfoil chord] 2. 2.964 0 1.287 1.585 4.264 5.261 6.077 7.348 8279 8.941 9.362 9.559 9.527 9.289 8.871 8.298 7.595 6.780 5.877 4.907 3.900 2.885 1.884 0 10.418 11. 11.822 12.086 12.0.16 11.767 1l.251 19.
ordinates given in percent of ordinates given in percent of Ordinate Ordinate
--- ------
surface surface radius: and and .156
NACA E. NACA .a61 .797
.300 .525 .991 9.367 0 1.965 4.393 6.868 2.198 4.660 7.147 9.647 Upper 0 Upper 14.404 19.469 24.549 29.640 34.734 39.829 44.919 50.000 55.069 60.125 65.164 70.187 75.191 SO. 85.147 90.103 95.048 14.669 19.705 24.750 29.800 34.852 39.C05 44.955 50.000 55.039 60.070 65.093 70.106 75.109 80.102 85.085 90.059 95.028 L. SlopeofradiusthroughL. 100.000 L. E. radius: Slope of radius through L.
Station Station 100.000
--- ---- -
Stations [Stations I , I I 284 553 .016 .051 .286 0 0 0 0 0.1685 -.867 -.334 -.438 percent of 0.0842 percent of -1.150 -1.388 -1.766 -2.420 -3.328 -3.999 -4.535 -5.336 -5.895 -6.259 -6.448 -6.470 -6.315 -6.004 -5.562 -5.013 -4.382 -3.691 -2.962 -2.224 -1.513 -1. -1. -1.982 -2.711 -3.711 -4.443 -5.019 -5.868 -6.448 -6.805 -6.966 -6.938 -6.702 -5.736 -5.066 -4.312 -3.506 -2.676 -1.&18 -1.096 Ordinate -6.292 Ordinate
--- ----
in in E.: L,K: L.
.601 .863 .733 0 1. 2.652 5.171 7.677 1.555 2.860 5.407 7.923 Lower surface Lower surface 0 1.013 10.177 15.166 20.148 25.125 30.100 35.074 40.048 45.023 50.000 54.981 59.965 64.953 69.947 74.945 79.949 84.957 89.970 94.986 10.423 20.355 25.301 35.177 40.114 45.054 50.000 54.954 59.917 64.890 69.875 74.872 79.881 94.968 15.398 30.240 84.901 89.931 Station 100.000 Station 100.000
--- ----
633-41
632-215
-- through 1.594 2.120 980 457 .978 .616 1.484 1.833 2.410 3.455 4.975 6.139 7.087 8.560 9.632 9.446 8.596 7.626 6.564 5.438 4.280 3.130 2.017 0 airfoil chord] airfoil chord] 0 0 1.250 1.528 1. 2.792 3.960 4.847 5.569 6.682 7.487 8.049 8.392 8.530 8. 8.194 7.768 7.203 6.524 5.751 4.906 4.014 3.105 2.213 1.368 0 10.385 10.854 11.058 10.986 10.1\72 10.148 ordinates given ordinates given Ordinate
--- Ordinate ---
• and radius: of radius through and
NACA
E. radius: NACA E.
.399 .637 .487 ---- .267 .945 0 1.120 2.348 4.829 7.323 '9.823 2.140 4:59.3 7.077 9.577 Upper surface Upper surface 0 14.834 19.852 24.875 29.900 34.926 39.952 44.977 50.000 55.019 60.035 65.047 70.053 75.055 80.051 85.043 90.030 95.014 14.602 19.645 24.699 29.760 34.823 39.886 44.946 50.000 55.046 60.083 70.125 75.128 80.119 85.099 90.069 95.032 L.
L. Slope ofradius '65.110 Slope Station Station 100.000 100.000 ---- --- - [Stations [Stations - 204 541 0 0 0 0
8 0.0842
-.852 -.300 -.467 -.032 percent of percent of -1. -1.462 -1.878 -2.610 -3.648 -4.427 -5.055 ~6.011 -6.693 -7.155 -7.421 -7.500 -7.386 -7.099 -6.665 -6.108 -5.453 -4.721 -3.934 -3.119 -2.310 -1. -1.349 -1.638 -2.105 -2.913 -4.041 -4.880 -5.547 -6.549 -7.250 -7.704 -7.940 -7.970 -7.774 -7.387 -6.839 -6.161 -5.384 -4.537 -3.650 -2.754 -1.113 .-1.894 Ordinate Ordinate
---- ---
in in E.: .5 .75 .618 .88.3 1. 2.5 5.0 7.5 Lower surface 0 1.404 2.681 5.204 7.712 Lower surface 0 10 15 20 25 40 45 50 55 60 65 70 75 80 85 90 95 10.212 15.199 20.178 25.150 30.120 35.089 40.057 45.027 50.000 54.977 59.958 64.945 69.938 74.936 , 79.941 84.951 89.966 94.984 30 35 Station 100 Station 100.000 ---- ---
632-015 633-21
1.594 501 2.120 462 on .664 5.728 6.581 9.045 5.594 .852 .300 airfoil chord] 0 1.449 1.778 2.319 3.285 4.673 7.895 8.842 9.494 9.884 9.916 9.577 8.351 7.526 6.597 4.544 3.486 2.459 1. 0 airfoil chord] 1.541 0 1.204 1. 1.878 2.610 3.648 4.427 5.055 6. 6.693 7.155 7.421 7.500 7.386 7.099 6.665 6.108 5.458 4.721 3.934 3.119 2.310 0 10.030 ordinates given ordinates given
Ordinate --- Ordinate ---
radius through L.
surface radius: of and and .5 .75 E.
NACA NACA
0 1. 5.0 7.5 .382 .617 '2·5 15 20 25 45 55 60 65 70 75 80 85 90 95 2.319 7.288 9.788 10 30 35 40 50 Upper 0 1.096 4.796 Uppe,r surface 100 14.801 19.822 24.850 29.880 44.973 50.000 55.023 60.042 65.055 70.062 80.059 85.049 95.016 L. 34.911 39.943 75.064 90.034 L. K radius: Slope Station Station 100.000 ---- ---- ~ , i ;Stations [Stations of I 1 I 716 779 265 787 .329 .074 .383 0 0 0 0 -.764 -.308 -.985 -.348 -'.871 0.1685 percent oj -1.291 -2.280 -2.685 -2.995 -3.446 -3.745 -3.919 -3.984 -3.939 -3.778 -3.514 -3.164 -2.745 -2.278 -1. -1. -1.404 -1.713 -2.217 -3.104 -4.362 -5.308 -6.068 -7.225 -8.048 -8.600 -8.913 -9.000 -8.845 -8.482 -7.942 -7.256 -6.455 -5.567 -4.622 -3.650 -2.691 -1.
-1.040 -1.
Ordinate ----- Ordinate --- , in E.: -- .5 .664 .933 .75 1.25 2.5 5.0 7.5 Lower surface 0 1. 2.743 5.273 7.782 Lower surface. 0 10.282 15.265 40.076 45.036 59.943 64.924 69.913 74.911 84.930 89.951 10 15 20 25 35 40 45 50 55 60 65 70 75 80 85 90 95 gben 20.235 25.200 30.160 35.118 50.000 54.969 79.916 94.977 30 100.000 Station 100 Station --- ---
631-412
2.120 320 719 739 1.087 .881 787 0 1. 1. 2.460 3.544 4.379 .1.063 6.138 6.929 7.499 7.872 8.059 8.062 7.894 7.576 7.125 6.562 5.899 5.153 4.344 3.492 2.618 1. 0 .985 .348 airfoil chord] , 1.071 airfoil chord] 0 1.404 1.713 2.217 3.104 4.362 5.308 6.068 7.225 8.048 8.600 8.913 9:000 8.845 8.482 7.942 7.256 6.455 5.567 4.622 3.650 2.691 1. 0 ordinates given in percent ordinaj;es Ordinate Ordinate surface radius: radius: and and .75 .336 .• .5 E.
NACA E.
NACA 633-01 8 1.25 2.5 7.5
0 1.041 2.257 4.727 7.218 9.718 0 50.000 75.089 10 15 20 25 40 45 50 55 60 65 70 7., 80 85 90 95 14.735 19.765 24.800 29.840 34.882 39.924 44.964 55.031 60.057 65.076 70.087 80.084 85.070 90.049 95.023 Upper , 5.0 30 35 Upper surface 100.000 L. Slope of radius through L. L.
Station Station [Stations [Stations 517 664 629 .066 .083 .483 .704 .651 0 0 0 0.0842 -.932 -.601 -.190 percent of -.430 0.2527 -1.661 -1.017 -1.214 -1. -2.013 -2. -3.123 -3.476 -3.972 -4.290 -4.460 -4.499 -4.407 -4.172 -3.814 -3.356 -2.823 -2.239 -1. -1.015 -1.120 -1.408 -1.912 -2.606 -3.115 -3.520 -4.124 -4.545 -4.816 -4.957 -4.970 -4.849 -4.609 -4.267 -3.~ -3.349 -2.810 ·-2.238 -1.106 Ordinate Ordinate --- I i ,0 : in E.: E.: 132 OW 795 L.
..
.583 .843 1.082 1.634 2.950 5.508 8.027 o 1.355 2.622 5.137 7.642 I'()wer surface 0 Lower surface 10.527 15.496 20.442 25.375 30.300 50.000 54.942 59.895 64.861 69.841 74.837 79.847 84.873 89.911 94.958 10.141 15. 20.118 25.100 30.080 35. 40.038 45.018 50.,000 54.984 59.971 64.962 69.957 74.955 79.958 84.965 89.975 94.988 35.222 40.143 45.068 100.000 Station 100.000 Station
I
6~2-615'
631-212
473 1.594 1.087 .566 1.622 3.963 4.554 5.470 6.137 6.901 7.030 6.799 6. 5.491 4.870 4.182 3.451 2.698 1.224 0 airfoil chord] 0 1.0.32 1.260 2.284 3.238 6.606 6.991 6.030 airfoil chord] 4.560 5.667 6.578 8.010 9.974 9.393 7.809 6.847 5.800 4.693 3.555 2.398 ' 1.947 0 1.317 1.634 2.159 3.129 9.066 9.830 8.665 0 ordinates given ordinates given in percent of 10.331 10.598 10.587 10.384 ' 1.245 Ordinate
I~ I I radius through L. '
surface surface radius: radius: of and and 492 504 558 .657 .. E.
.417 E. NACA
NACA
1.145 7.358 0 .205 .418 .866 o 2.378 4.863 9.859 2.050 14.S68 19.882, 2{.900 ~n~1 39.962 44.982 50.000 55.016 60.029 65.038 70.043 75.045 80.042 85.035 90 95.012 4. 6.973 9.473 Upper Upper 14. 19. 24.625 29.700 50.000 55.058 60.105 65.139 70.159 75.163 80.153 85.127 95.042 100.000 34.778 39.857 44.932 90.089 L. Slope Qfradius through L, Slope Station Station 100.000 [Stations [Stations
t'l >I> >-:3 '-< t< t:;) '-< ~ >-:3 >-:3 t'l ~ t'l ~ ~ '-<
::0 "d o ~ >-:3 Z ? 00 ~ ~ o Z ;.- ;.- < 111 o ~ "" l.l o ;::;: t'l "j o ;.- o ~ l.l 111
..... o ~
I I .089 0 0 0 0 0.084 -.739 -.892 -.611 -.227 -.686 -.819 -.768 -.396 -.094 percent of percent of -1.128 -1.533 -2.109 -2.543 -2.898 -3.455 -3.868 -4.170 -4.373 -4.479 -4.490 -4.364 -4.136 -3.826 -3.452 -3.026 -2.561 -2.069 -1. -1.069 -1.018 -1.344 -1.791 -2.117 -2.379 -2.781 -3.071 -3.274 -3.401 -3.449 -3.419 -3.269 -3.033 -2.731 -2.381 ~1.996 ~1.589 -1.174 Ordinate Ordinate surface
--- ---
E.: in in L.
.SO .75 .562 .820 1.328 Lower 0 1.25 2.5 5.0 7.5 Lower surface 0 2.589 5.099 7.602 45.014 50.000 54.988 59.978 64.970 69.965 74.964 79.965 94.989 given 10 15 20 25 30 35 40 45 50 55 60 65 70 75 80 85 90 95 10.101 15.095 20.085 25.073 30.059 35.044 40.029 84.970 89.979 100.000 Station 100 Station --- ---
64-009 64-209
through 0.579 0.579 2,~2 716 742 .543 .739 .892 .611 .227 .786 .959 airfoil chord] airfoil chord} 5.509 1.128 0 1.128 1.533 2.109 2.543 2.898 3.455 3.868 4.170 4.373 4.479 4.490 4.364 4.136 3.826 3.452 3.026 2.561 2.069 1.564 1.069 0 0 1. 1. 2.423 2.965 3.413 4.127 4.66.'l 5.064 5.345 5.561 5.459 .5.239 4.921 4.523 4.056 3.533 2.964 2.360 1. 0 -- ordinates ordinates given
Ordinate --- Ordinate
--- radius surface surface radius: radius; of and and
NACA NACA
.50 .75 035 E.
E.
0 1.25 2.5 5.0 7.5 .438 .680 10 15 20 25 30 35 40 45 50 55 60 65 70 75 80 85 90 95 Upper 0 1.172 2.411 4.901 7.398 9.899 Upper 14.905 50.000 55.012 60.022 65.030 70.035 75.036 95.011 100 L. 19.915 24.927 29.941 34.956 39.971 44.986 SO. 85.030 90.021 L. Slope Station Station 100.000 ---- ---- [Stations [Stations
I I i I of
.UO 0 0 0 0 0.084 -.754 -.737 -.157 -.606 -.722 -.896 -.959 -.608 -.288 -.033 -.494 -.596 -.423 percent of percent -1.740 -1.177 -1.557 -2.395 -2.949 -2.921 -2.316 -2.010 -1.673 -1.319 -1.024 -1.405 -1.692 -1.928 -2.298 -2.572 -2.772 -2.907 -2.981 -2.995 -2.919 -2.775 -2.575 -2.331 -2.050 -1.412 -1.072 -1.833 -2.055 -2.640 -2.808 -2.912 -2.788 -2.581 Ordiilate Ordinate ---- --- in E.: in L.
25 000 .50 .75 .555 .812 I. 5.088 Lower surface 0 2.5 5.0 7.5 Lower surface 0 1.320 2.579 7.590 10.091 54.989 59.980 94.990 10 15 20 25 30 35 40 45 50 55 60 65 70 75 80 85 90 95 15.085 20.076 25.065 30.052 35.039 40.026 45.012 SO. 64.973 69.969 74.968 79.969 84.973 89.981 100.000 Statiou 100 Station --- --- -
64-006 64-208
through 0.256 0.455 S06 692 740 .522 .494 .596 .754 .737 .423 .157 airfoil chord] .706 .862 airfoil chorli] 1.024 1.405 1. 1.549 2.189 5.009 5.063 4.787 4. 4.152 3.733 3.263 2.749 2.200 1.634 1.067 0 0 1. 1.928 2.298 2.572 2.772 2.907 2.981 2.995 2.919 2.775 2.575 2.331 2.050 1.412 1.072 0 0 1.110 2.681 3.089 3.741 4.232 4.598 4.856 4.978 ordinates given ordinates given Ordinate Ordinate --- --- radius surface surface radius: radius: of and and
NACA NACA
.50 .75 E. E.
0 I. 2.5 5.0 7.5 .445 .688 10 15 50 70 75 1.180 2.421 4.912 7.410 9.909 Upper 20 25 30 35 40 45 55 60 65 80 85 90 95 Upper 100 14.915 19.924 24.935 29.948 34.961 39.974 44.988 50.000 55.011 60.020 65.027 70.031 75.032 80.031 85.027 90.019 95.010 L.
L. Slope Station Station 100.000 ---- ---- i [Stations [Stations of of I 774 .524 267 775 935 951 .242 -- .094 .159 0 0 0 0 0.084 -.672 -.076 -.442 - . -.645 -.836 -.768 -.517 -.276 -.064 percent percent 0.1685 -1.461 -1. -2.289 -3.181 -4.411 -5.314 -6.029 -7.082 -7.809 -8.257 -8.464 -7.000 -6.200 -5.298 -3.344 -2.367 -1.459 -1.087 -1. -1.410 -1.624 -I. -1.877 -1. -I. -1.924 -1.824 -1.672 -1.480 -1.260 -1.020 -8.438 -8.155 -7.664 -4.335 Ordinate Ordinate surface surface ---- ---- in in E.: E.: L. L.
.541 .763 .796 1. 0 1.302 2.560 5.066 7.568 Lower 0 1.048 2.914 5.473 7.993 Lowrr given 10.494 15.465 20.415 25.351 given 10.067 15.06.'l 20.057 25.048 :lO.039 40.019 45.009 50.000 54.992 59.985 64.980 69.977 74.975 79.976 94.993 30.281 35.207 40.133 45.063 50.000 54.946 59.904 64.874 69.857 74.855 79.865 84.889 89.922 94.963 35.029 84.980 89.985 Station Station 100.000 100.000 --- ---
64-206
--
634-421
through through 2.650 0.256 412 208 719 Il5 :367 OIl 661 022 .542 .664 .859 .940 .473 airfoil chord] airfoil chord] 1.
0 I. 2.054 2.717 3.925 5.675 7.010 8.097 9.774 7.232 5.947 4.643 2.144 I. 0 1. 2. 2.444 2.970 3. 3.667 3.879 4. 4.066 4.014 3.878 3.670 3.402 3.080 2.712 2.307 1.868 1.410 0 ordinates 9.582 8.455 3.364 0 ordinates 10.993 11.837 12.352 12.044 11. 10.580 12.558 12.439 Ordinate ---- Ordinate surface radius surface radius radius: radius: and and of
NACA
NACA E. E.
.452 .902 .459 .704 .237 0 2.086 9.506 0 I. 2.440 4.934 7.432 9.933 Upper 4.527 7.007 Upper 14.535 19.585 24.649 29.719 50.000 55.054 70.143 75.145 90.078 95.037 14.937 19.943 24.952 29.961 34.971 39.981 44.991 50.000 55.008 60.015 65.020 70.023 75.025 90.015 95.007 34.793 39.867 44.937 60.096 65.126 80.135 85.111 L. Slope of 80.024 85.020 L. Slope Station 100.000 Station 100.000 --- -------- I [Stations [Stations of
I of
Ill;) -- 0 0 0 0 0.042 -.595 -.076 0.0842 -.794 -.953 -.840 -.413 -.090 percent percent -1.527 -1.861 -2.414 -3.385 -4.743 -5.75.'l -6.559 -7.765 -8.612 -9.156 -9.439 -9.469 -9.227 -8.759 -8.103 -7.295 -6.370 -5.366 -4.318 -3.264 -2.257 -1.347 -1. -1.607 -2.184 -2. -2.96.'l -3.506 -3.904 -4.191 -4.378 -4.465 -4.452 -4.295 -4.034 -3.690 -3.284 -2.830 -2.341 -1.833 -1.324 Ordinate Ordinate --- ---- .
in 1 in E.: E.: L.
L. 293 - g80 .900 .535 .6.'l3 .788 0 1.425 2.708 5.237 7.747 1. 2.550 5.055 7.557 Lower surface Lower surface 0 10.247 15.233 20.208 25.176 10.056 15.053 30.140 35.103 40.066 45.031 50.000 54.973 59.952 64.937 69.929 74.927 79.933 84.944 89.961 94.982 20.047 25.041 30.033 35.025 40.016 45.008 50.000 54.993 59.988 64.984 69.981 74. 79.981 84.984 89.988 94.994 Station 100.000 Station 100.000 --- ---
64-1
634-221
through through 2.650 0.720 369 02.'l 793 .844 .929 .406 airfoil chord] .708 airfoil chord] 0 1.627 2.001 2.628 5.375 6.601 7.593 9.111 6.262 5.054 1.629 I. 1.303 I. 2.500 3.479 4.178 4.700 5.087 5.350 5.495 5.524 5.391 5.138 1.512 ordinates given 3.757 9.485 8.512 7.426 3.849 2.693 0 0 3.037 4.786 4.356 3.860 3.313 2.729 2.120 0 ordinates given 10.204 10.946 11.383 10.309 11.529 11. 10.949 Ordinate ---- Ordinate --- radius radius surface radius: of and radius: and
NACA
075 207 NACA 000 E.
E.
.600 .465 .712 .367 0 I. 2.292 4.763 7.253 9.75.'l 0 I. 2.450 4.945 7.443 Upper 9.944 14.767 19.792 24.824 29.860 44.969 50.000 55.027 60.048 65.063 70.071 75.073 90.039 95.018 14.947 19.953 24.959 29.967 34.975 39.984 44.992 50.000 55.007 60.012 65.016 70.019 75.020 90.012 95.006 34.897 39.934 80.067 85.056 L. Slope of 80.019 85.016 L. Slope Station 100.000 Station 100.
~persurf~1 ---- ----
[Stations [Stations of I I I i I of 0 0 0 0 0.042 -.632 -.950 -.392 -.758 -.625 -.292 -.048 percent percent -1.271 -2.316 -2.733 -3.398 -3.132 -2.86.'l -1.805 -1.583 -1.937 -2.527 -3.577 -5.065 -6.182 -7.080 -8.441 -9.410 -9.854 -9.206 -8.390 -7.441 -6.396 -5.290 -4.160 -3.054 -2.021 -1.113 -1.716 -2.047 -3.039 -3.256 -3.464 -3.456 -3.335 ·-2.545 ':"2.189 -1.406 -1.006 Ordinate surface Ordinate -10.053 -10.412 ---- -10.500 -10.298 ---- E.: in in L.
2.'; 285 ,545 .75 .528 .781 .5 0 I. 2.5 5.0 7.5 0 1. 2.540 5.044 7.
Lower Lower surface 10.045 10 15 20 25 30 35 40 45 50 55 60 65 70 75 80 85 90 95 15.042 20.038 25.032 30.026 35.020 40.013 45.006 50.000 ,54.995 59.990 64.987 69.985 74.984 79.985 84.987 89.990 94.995 Station 100 Station 100.000
--- ----
I
64-1
634-021
0.455 2.650 937 113 297 .808 .364 .392 .682 .828 airfoil chord] airfoil chord] 1.058 1.457 2.032 2.832 2.777 2.302 1.802 1.
0 1.583 I. 2.527 3.577 5.065 6.182 7.080 8.441 9.410 9.854 9.206 8.390 7.441 6.396 5.290 4.160 3.054 2.021 I. 0 0 2.471 3.405 3.8.% 4.152 4.370 4.494 4.528 4.431 4.236 3.959 3.617 3.219 0 ordinates given ordinates given 10.053 10.412 10.500 10.298 Ordinate Ordinate --- --- surface surface and radius: and NACA 215 E. radius: 2.5 NACA E.
.472 .5 .75 .719 1. 2.460 4.956 7.455 9.955 0 I. 2.5 5.0 7.5 Upper 0 Upper 10 20 25 45 50 55 60 70 75 14.958 19.962 24.968 29.974 34.980 39.987 44.994 50.000 55.005 60.010 65.013 70:015 75.016 80.015 85.013 90.010 95.005 15 30 35 40 65 80 85 90 95 L. 1,. Slope of radius through Station 100 Station 100.000 --------- --- ---- [Stations [Stations o t"'
U2 c:l ;::: ;:::. ;.. ~ o >:rj ;.. .... ~ "'1 .... l:::l ;.. >-3 ;.. o CJ1
I-'- of
of I
.250 .345 0.168 0. 0. 0 0.
0..0.84 -.583 -.0.84 -.864 -.903 -.435 -.0.38 percent percent -2.166 -2.535 -1.373 -1. -2.0.65 -3.885 -4.-648 -5.282 -6.266 -7.495 -7.816 -7.949 -7.881 -7.535 -7.0.11 -6.350 -5.587 -4.752 -3.870. -2.970 -2.0.91 -1.277 -1.025 -1.262 -1.649 -2.828 -3.267 -3.576 -3.783 -3.898 -3.917 -3.839 -3.608 -3.274 -2.866 -2.406 -1.913 -1.405 -2.814 -6.984 Ordinate Ordinate E.: surface --- --- in in E.: L.
L.
.662 .931 .620 .883 Lower surface 0. 1. 2.736 5.262 7.771 Lower 0. 1.401 2.675 5.196 7.70.3 35.118 50.000 54.968 59.941 45.028 10..270. 15.255 20.228 25.195 30.158 40.0.77 45.0.37 64.922 69.910 74.906 79.911 84.924 89.945 94.973 10.203 15.192 20.172 25.147 30.119 35.0.88 40..0.58 50..000 54.976 59.957 64.943 69.935 74.932 79.936 84.946 89.962 94.981 tQQ.QOO Station Station 100.000 ---
641-412 through 643-218
through 1.0.40 2.208 305 231 473 785 .919 .716 airfoil chord) airfoil chord] 1.0.64 1. 2.393 3.430 4. 4.896 5.959 6.760 7.786 8.0.37 8.123 7.988 7.686 7.246 6.690 5.293 4.483 3.619 2.722 1.818 0. 0. 1. 1. 2.279 4.497 6.316 7.612 8.576 9.285 9.750 9.725 7.729 6.812 5.814 4.760 2.623 1. 0.
0. 1.690 7.363 6.0.33 ordipates given 3.186 5.496 9.217 8.040. 3.683 ordinates given 10.009 10.0.23 Ordinate Ordinate --- ------ radius surface surface - radius: radius: and and E.
NACA NACA E.
.338 .569 .380. .617 0. 1.045 2.264 4.738 7.229 9.730 0. 1.0.99_ 2.325 4.804 7.297 9.797 Upper Upper 14.745 19.772 24.S05 29.842 34.882 39.923 44.968 50.000 5.5.0.32 60.0.59 65.0.78 70..0.90 75.0.94 80.0.89 90.0.55 95.0.27 14.808 19.828 24.853 29.881 34.912 39.942 44.972 50.000 55.0.24 60.0.43 65.0.57 70..065 75.0.68 80.0.64 85.0.54 90.0.38 95.019 Slope of L. Slope of radius L.
Station Station 100..000 100.000 . 85.0.76
---- ----
[Stations [Stations of of - -- .0.28 - 0 6 Q.Q~4 0. 0.
-.925 -.7C8 -.269 -.400.
percent percent -1.105 -1.379 -1.846 -2.491 -2.967 -3.352 -3.945 -4.376 -4.680- -4.871 ~4._948 -4.910 -4.70.3 -4.377 -3.961 -3.477 -2.944 -2.378 -1.SOQ -1.233 -1.428 -1.720. -2.177 -3.005 -4.186 -5.0.76 -5.S03 -6.942 -7.782 _8.391 -8.789 -8.979 -8.952 -8.630. -8.114 -7.445 -6.658 -5.782 -4.842 -3.866 -2.888 -1.951 -1.10.1 Ordinate Ordinate surface ---- in E.: in L.
.582 .841 .50. .75 0 1. 2.618 5.132 7.636 0. 2.5 5.0.
Lower surface Lower 1. 7.5 given 10.135 15.128 20.114 25.0.97 30.0.79 35.0.59 40.0.39 45.0.18 50.000 04.984 59.971 64.961 69.955 74.953 79.955 84.962 89.973 94.987 given 10. 15 20. 25 30. 35 40. 45 50. 55 60 65 70. 75 95 80. 85 90 Station 100.0.0.0.
Station 100.
---
643-018
641-212
through 1.0.40.
2.208 720.
.60.4 .400.
airfoil chord] airfoil chord) 0. 1.0.25 1.245 1.593 2.218 3.123 3.815 4.386 5.291 5.968 6.470 6.815 7.008 7.0.52 6.893 6.583 6.151 5.619 5.004 4.322 3.590 2.825 2.0.54 1.303 0 0. 1.428 1. 2.177 4.186 5.0.76 5.803 6.942 8.979 4.842 2.888 1.951 3.0.05 7.782 8.391 8.789 8.952 8.630 8.114 7.445 6.658 5.782 3.866 1.101 0 ordinates ordinates Ordinate --- Ordinate radius \ surface surface radius: and and radius: E. .50 .75 NACA NACA E.
.418 .659 0. 1,25 2.5 5.0. 7.5 0 1.147 2.382 4.868 7.364 9.865 Upper Upper lO 15 20 25 30. 35 40. 45 50. 55 50 65 70. 75 80. 85 90. 95 14.872 19.886 24.903 29.921 34.941 3R 44.982 50.000 55.0.16 60.0.29 65.0.39 70..045 75.047 80..045 85.0.38 90.0.27 95.0.13 L. Slope of 100 L.
Station 10.0..000 Station
----
[Stations [Stations
of I I
I -- .0.86 .288 0.042 0. 0 0.
0 0..168 -.952 -.528 -.130 -.878 -.328 percent percent of -1.143 -1.435 -1. -2.651 -3.181 -3.612 -4.284 -4.775 -5.128 -5.358 -5.463 -5.445 -5.250 -4.928 -4.508 -4.0.12 -3.459· -2.864 -2.247 -1.631 -1.046 -1.0.91 -1.299 -1.610 -2.139 -2.857 -3.379 -3.796 -4.430. -4.882 -5.191 -5.372 -5.421 -5.330. -5.0.34 -4.60.4 -4.0.76 -3.478 -2.834 -2.167 -1.50.4 Ordinate Ordinate surface E.: --- in in E.: L.
L.
.541 .796 .70.1 .974 1.302 2.559 5.0.66 7.568 Lower 0 Lower surface 0. 1.504 2.793 5.327 7.838 10.068 15.0.64 20.0.57 25.0.49 40.019 45.009 50.000 04.992 25.244 30..0.39 35.0.29 59.985 69.977 74.976 79.978 84.981 89.986 94.993 10..338 15.319 20.286 30..197 35.147 40..0.96 45.046 50..000 04.960. 59.928 64.904 69.889 74.885 79.891 84.908 89.934 94.968 .64.9SO Station 100..0.00 Station 100.000
--- ---
641-112 642-415
through through 1.040 1.590 718 291 .446 airfoil chord] airfoil chord] .976 0 1.002 1.213 1.543 2.127 2.967 3.605 4.128 4.956 5.571 6.0.24 6.330. 6.493 6.517 6.346 6.0.32 5.604 5.084 4.489 3.836 3.143 2.427 1. 1.044 0. 1. 1.579 2.883 4.121 7.122 9.041 0. 2.0.38 5.0.75 5.864 8.0.66 8.771 9.260. 9.614 9.414 9.0.16 8.456 7.762 6.954 6.055 5.0.84 4.0.62 3.0.20. 1. 0.
ordinates given ordinates given _ radius Ordiriate . Ordinate --- --- surface surface of radius: and and radius:
NACA E. NACA
E.
.459 .704 .299 .526 .996 0 1.198 2_441 4.934 7.432 9.932 2.207 4.673 7.162 9.662 Upper Upper 0 - 14.936 19.943 24.951 29.961 34.971 39.981 44.991 50.0.00 55.008 60.015 65.020 70..023 75.0.24 SO. 85.0.19 90.014 95.007 14.681 24.756 29.S03 34.853 39.90.4 44.954 50..000. 60..0.72 65.0.96 70..111 81i.Q92 L. Slope 19.714 55.0.40. 75.115 80.10.9 90.066 95.0.32 Slope of radius Station Station L.
100.000 100.000 ---- --- [Stations [Stations of of - 0 0.
0.0.84 -.978 -.786 -.288 0 percent percent -.432 -.0.30. 0 -1.179 -1.490 -2.0.35 -2.810 -3.394 -3.871 -4.620 -5.173 -5.576 -5.844 -5.978 -5.981 -5.798 -5.480 -5.056 -4.048 -3.974 -3.350 -2.695 -2.0.29 -1.382 -1.154 -1.382 -2.338 -3.184 -4.322 -5.110 -5.682 -6.0.89 -6.346 -6.452 -6.40.2 -5.70.7 -4.549 -3.865 -3.141 -1. -3.813 -6.129 -5.171 -2.401 -1.675 -1.003 -- Ordinate Ordinate surface
--- ---
in in E.: L.
.5 .75 .601 .863 0. _ 1.25 2.5 5.0. 7.5 5.164 Lower Lower surface 0 1. 2.647 7.669 10 15 20 25 30. 35 40 45 50 55 60. 65 70 SO 85 90 given 10.169 15.160 20.143 25.122 40..048 75 95 30.099 35.0.74 45.0.23 50.000 04.980. 59.964 64.952 69.945 74.942 79.945 84.954 94.984 Station 100 Station ·89.967 100.000 ---
641-012
~642-215 through 1.0.40 1.590.
382 254 945 512 466 ------- .978 .786 .288 .662 airfoiLchord] airfoil chord] 0. 1.179 1.490 2.035 2.810 3.394 3.871 4.620 5.173 5.576 5.844 5.978 5.981 5.798 5.4SO 5.056 4.048 3.974 3.350 2.0.29 1. 0 0. 1. 1.li22 1. 2.710. 3.816 4.661 5.356 6.456 7.274 7.879 8.290. 8. 8.044 8.319 7.913 7.361 6.691 5.925 5.0.85 2.349 1.
ordinates given 2.695 4.191 3.267 0.
ordinates Ordinate --- Ordinate ------ surface surface - and radius: radius: and 000 0.55
NACA E. NACA
.75 E.
.5 .399 .637 0 1.25 2.5 5.0 7.5 0. 1.122 2.353 4.836 7.331 9.831 Upper Upper 55 11.840. 24.878 29.90.1 10 15 20 25 30 35 -40 45 50 60 65 70. 75 SO 85 90 95 19.857 34.926 39.952 44.977 SQ. 55.0.20 60..0.36 65.048 70.0.55 75.058 SO. 90..033 95.0.16 L. L. Slope of radius Station Station 100.000
--- ---- ----sir.-ll46-
[Stations [Stations of of - .068 0 0. 0.
-.767 -.916 -.926 -.503 -.154 0.084 -.950
percent percent 'j
-1.140- -1.512 -2.024 -2.400 -2.70.2 -3.168 -3.505 -3.743 -3.892 -3.950 -3.917 -3.748 -3.483 -3.143 -2.749 -2.315 -1.855 -1.386 -5.785 -6.985 -7.319 -7.482 -1.208 -1.456 -1.842 -2.528 -3.504 -4.240. -4.842 -6.480. -7.473 -7.224 -6.810. -6.266 -5.620. -4.895 -4.113 -3.296 -2.472 -1.677 Ordinate Ordinate --- --- in E.: in L.
.569 .827 .50 .75 1.337 2.599 5.110 7.613 5.0.
Lower surface 0 Lower surface O. 1. 2.5 7.5 10.113 15.106 20.0.95 25.A81 30..066 35.0.49 40.032 45.0.15 50.000 54.987 59.975 69.962 89.977 15 20 25 30.
~iven 64.967 74.960 79.962 84.968 94.988 10 35 40 45 50. 55 60. 65 70. 75 80. 85 90. 95 100.000 Station Station 100
------ --- ---
64-210
642-015
through 0..720 1.590 .867 .564 airfoil chord] airfoil chord] .950. .346 5.097 0 1.056 1.354 1.884 2.656 3.248 3.736 4.514 5.533 5.836 6.010. 6.059 5.938 5.689 5.333 4.891 4.375 3.799 3.176 2.518 1.849 1.188 0. 0. 1.208 1. 1.842 2.528 3.50.4 4.240. 4.842 5.785 6.480. 6.985 7.319 7.482 7.473 7.224 6.810 4.895 4.113 6.266 5.620. 3.296 2.472 1. 0.
ordinates ordinates given Ordinate Ordinate --- radius surface surface of and radius: and radius:
NACA .50. .75
E.
NACA E.
0 .431 .673 0 1.25 2.5 5.0. 7.5 2.401 1.163 4.890 7.387 9.887 10 15 20. 2S 30. 3" 40. 45 50. 55 60. 65 70. 75 SO Upper Upper 85 90 95 14.894 24.919 29.934 44.985 50.000 19.905 34.951 39.968 55.0.14 65.033 70.038 75.040 80.038 85.033 90.0.24 95.012 L. Slope 100 Station -60.025 100.000 Station L.
----
[Stations [Stations ~ l'J "C o ~ "'3 z o 00 .., "'" * > >-3 ..... o Z > t"' > i;;1 ~ ..... o ~ >1 (") o ~ ~ ..... >-3 >-l "'l o ~ Z '::i >-l ..... ,p
~ U1 l'J M > l'J ~ o > (")
-
I
, I
..
.185 0 0 0 0 -.864 0.168 -.208 -.476 -.574 -.717 -.956 -.865 -.510 -.195 percent of -1.523 -1.821 -8.419 -2.27Q -3.090 -4.218 -5.048 -5.718 -6.750 -, -8.0n -8.321 -8.288 ":'7.840 -7.198 -6.417 -5.535 -4.588 -3.603 -2.623 -1.692 -1.310 -1.589 -1.824 -2.197 -2.482 -2.697 -2.852 -2.952 -2.998 -2.98.1 -2.900 -2.741 -2.518 -2.246 -1.9:15 -1.594 -1.233
ordina~
Ordinate 1 --- I in E.: L.
..
.773 .5 .75 0 1.055 1.597 2.904 5.455 7.972 0 1.25 Lower surface Lower surfaee 2.5 5.0 7.5 10.472 15.447 20.401 25.343 30.277 35.206 40.135 45.064 50.000 54.945 59.901 64.869 69.850 74.846 79.855 84.878 89.913 94.958 given in percent of 15 25 40 50 75 81i 90 10 20 30 35 4.1 55 60 65 70 80 95 Station 100.000 100 Station ---
65-006
--- 0.240 2.884 594 233 airfoil chord] airfoil chord] .476 .574 .717 .956 .865 .510 .195 0 1.723 1.079 I.
2.101 2.707 3.834 5.482 6.744 7.786 9.442 9.819 8.708 7.491 6.203 4.876 3.556 2.276 0 0 1.310 1.589 1.824 2.197 2.482 2.697 2.852 2.952 2.998 2.9S:l 2.900 2.741 2.518 2.246 1.935 I. 0 ordinates given ordinate~ 10.678 11.591 12.209 12.539 12.572 12.220 11.610 10.797 Ordinate Ordinato ._--- radius through surface surface radius: of radius: and
NACA
E. .5 .75 E.
NACA
.227 .445 1.25 .903 0 2.5 5.0 7.5 0 2.096 4.545 7.028 9.528 Upper Upper 10 15 20 25 30 35 40 45 50 55 60 65 70 75 80 85 90 95 14.553 19.599 24.657 29.723 34.794 39.865 44.936 50.000 55.055 60.099 70.150 75.154 90.087 95.042 65.131 80.145 85.122 L. Slope 100 L.
Station 100.000 Station -------- ---- [Stations [Stations and of of
I \
J _
911 1;7 .. .• 0 0 0 0 0.233 -.727 -.133 0.084 -.723 -.305 -.030 -.960 percent percent -1.590 -1.909 -2.404 -3.293 -4.550 -5.486 -6.248 -9.296 -9.450 -9.360 -8.935 -8.301 -4.577 -1.848 -1.278 -7.432 -8.297 -8 -7.512 -6.607 -5.619 -3.520 -2.490 -1.539 -1.110 -1.359 -1.801 -2.411 -2.832 -3.169 -3.673 -4.022 -4.267 -4.428 -4 -4.523 -4.446 -4.251 -3.940 -3.521 -2.995 -2.409 Ordinet, surface
~rdinate ---
I E.: in
6)-41 5
19 .~~~ .6.38 .904 .• Lower surface 0 I. 2.703 5.228 7.736 Lower .7.56 40.067 45.032 1.031 1.570 10.237 15.224 20.201 25.17l 30.139 35.103 50.000 54.973 59.950 64.935 69.925 74.92.3 79.927 84.939 89.956 94.979 0 2.879 5.436 7.956 15.439 64.692 74.763 79.820 Station 100.000 Stotion 10.460 20.392 25.331 30.258 35.175 40.084 44.981 49.847 54.737 59.695 69 84.883 89.938 94.980 100.000 ---
644-221
1.498 2.884 502 761 498 533 .765 .606
airfoil chord] 65(21 a=0.5 airfoil chord]
0 1.690 5.182 6.334 1. 1.2.16 1. 1. 7.575 6.373 5.152 3.890 2.639 I. 0 2.649 2.618 3.665 7.282 8.778 9.889 9.702 8.749 7.679 6.521 5.310 4.082 2.885 0 0 2.837 4.175 5.208 6.073 7.465 8.518 9.315 9.900 9.512 8.645 ordinates given in ordinates given 11.510 11. 11.125 10.507 1().279 10.467 10.438 10.131 10.701 Ordinate Ordinat, '11.240 --- surface radius: radius: and and NACA E. E.
.362 .596 .244 .469 .930 0 1. 2.297 4.772 7.264 9.763 0 2.121 4.564 7.044 9.540 Upper Upper surface 14.776 19.799 24.829 29.8t\1 34.897 39.933 44.968 50.000 55.027 60.050 65.065 70.075 75.077 80.073 85.061 90.044 95.021 14.561 19.608 24.669 29.742 34.825 30.916 45.019 50.153 55.263 60.305 65.308 70.281 75.237 80.180 85.117 00.062 95.020 L.
L. Slope of radius through L. Slope of radius through Station Station 100.000 100.000
NACA
---- [Stations [Stations of of '
I I
.234 .461 0 0 0 0 0.253 -.995 -.298 -.449 perc~nt -5.631 -1.646 -1.985 -2.517 -3.485 -4.871 -5.915 -6.769 -8.108 -9.095 -9.807 -9.404 -8.607 -7.678 -6.649 -5.549 -4.416 -3.287 -2.213 -1.245 -1.134 -1.347 -1.641 -2.129 -2.846 -3.396 -3.843 -4.527 -5.030 -5.397 -5.645 -5.774 -5.775 -5.284 -4.760 -4.103 -3.357 -2.5f>6 -1.
Ordinate Ordinate -10.269 -10.481 -10.431 -10:030 ---- ---- F.: in .50 .75 .566 Q23 778 0 1.25 2.5 5.0 7.5 .824 .
Lower surface Lower surface - 10 15 20 25 35 40 45 50 55 60 65 70 75 80 85 90 95 giwn 0 1.113 1.659 2.970 5.533 8.060 45.085 54. 59.858 74.767 79.776 84.808 89.862 94.932 100 Station 10 15.542 20.491 25.424 30.346 35.262 40.174 50.000 64.809 6~.
Station 100.000 --- ---
644-021
65,3-618
1.92 2.884.
767 565 477 .449 airfoil chord] airfoil chord] 1. 1.985 5.549 1.245 0 1.434 1. 2.283 3.245 5.940 6.945 8. 9.806 9.537 8.398 7.135 5.771 4.336 2.868 1.435 0 0 2.517 3.485 4.871 5.915 6.769 8.108 9.095 9.807 9.404 8.607 7.678 6.649 4.416 3.287 2.213 0 4.742 ordinates given in percent ordinates 10.481 10.431 10.030 10.767 11. 11.954 12.201 12.201 11.902 11.330 10.529 10.269 Ordinate Ordinate --- --- surface surface radius: radius: and and ---- .50 224 E.
.75 NACA E.
1.25 .176 .387 .R41 0 2.5 5.0 7.5
10 Iii 35 40 45 50 55 70 75 NACA 0 2.030 4.467 6.940 9.434
Upper 20 25 30 60 65 80 8., 90 95 Upper 100 14.458 19.509 24.576 29.654 34.738 39.826 44.915 50.000 55.077 60.142 65.191 70.222 75.2.13 SO. 8.1.192 90.138 95.068 L. Slope of radius through L.
L.
100.000 Station Station - ---- ---- [Stations [Stations 465 810 174 732 .075 .456 .552 0 0.194 0 0 0
8 0.253 -.810
-.494 -.345 -.050 percent of -6.361 -3.880 -3.067 -2.251 -1.473 -1.234 -1. -I. -2.402 -3.197 -3.768 -4.220 -4.899 -5.377 -5.695 -5.866 -5.885 -5.737 -5.345 -4.805 -4.160 -3.444 -2.690 -1.922 -I. -1.184 -1.412 -I. -2.273 -3.074 -3.688 -4.193 -4.957 -5.527 -5.937 -6.217 -6.376 -6.230 -5.879 -5.339 -4.658 Ordinate Ordinate --- ---- E.: in E.: L.
695 L.
.850 0 1.141 1. 5.583 8.105 .752 Lower surface 3.018 Lower surface given 10.605 15.573 35.265 40.173 45.083 50.000 69.804 74.797 0 1.033 1.569 2.869 5.422 7.947 20.514 25.440 30.355 54.929 59.871 64.829 79.809 84.839 89.885 94.944 10.456 25.359 35.235 40.165 50.021 54.954 59.894 64.8·15 69.807 74.781 79.751 94.951 15.442 20.408 30.300 45.068 84.779 89.865 Station 100.000 Station 100.000
--- ----
643-61
through 65,3-418
1.92 2.208 534 885 937 698 344 416 .930
airfoil chord] a=0.8 airfoil chordj
5.672 9.319 9.974 9.016 7.899 6.651 5.289 3.818 2.289 0 0 1. I. 2.452 3.518 5.093 6.312 7.322 8. 9.870 8.870 7.754 6.544 5.270 2.646 1. () 0 1. 1.736 2.224 3.133 4.542 6.617 8.149 ordinates 3.963 ordinates given in percent of 10.233 10.909 II. 11.600 U.602 11.307 10.751 10.153 11.065 11. 12.065 12.163 1l.915 11.423 10.730 Ordinate --- Ordinate --- radius through radius surface surface radiu~: of and radius: and 979 E.
NACA E.
.150 .359 .248 .467 .931 .805
NACA 0 2.131 4.578 7.053 9.544
Upper 0 I. 4.417 6.895 9.395 Upper 14.558 19.592 24.641 29.700 34.765 39.8.35 44.932 49. 55.646 60.106 65.155 70.193 75.219 80.249 85.221 90.135 95.649 14.427 19.486 24.560 29.645 34.735 39.827 44.917 50.000 55.071 60.129 n.1.171 70.196 75.203 85.161 90.115 95.056 L. Slope 80.191 L. Slope of 100.000 Station Station 100.000 ---- ---- [Stations [Stations r .
.234 0 0 0 -.492 0.168 percent -.602 -.064 percent of -8.8flS -8.990 -8.916 -8.593 -8.045 -7.317 -6.450 -5.486 -4.456 -3.390 -2.325 -1.324 -1.308 -1.560 -1. -2.613 -3.536 -4.212 -4.755 -5.585 -6.182 -6.596 -6.842 -6.917 -6.809 -6.440 -5.908 -5.255 -4.515 -3.721 -2.896 -2.074 -1.293 -1.324 -1..199 -2.004 -2.728 -3.831 -4.701 -5.424 -6.568 -7.43t -8.093 -8.568 ordinate- Ordinate --- in in E.: L. 25 .737 .5 .75 0 1.014 2.848 5.391 7.905 0 1. 2.5 5.0 7.5 Lower surface 1.550 Lower surface 55 60 75 85 90 95 10.405 10 15 20 25 30 35 40 45 50 65 70 80 15.383 20.343 25.293 30.237 35.177 40.115 45.055 50.000 54.953 59.914 64.886 69.869 74.865 79.873 84.892 89.923 94.963 Station 100.000 Station
643-418
through 65,3-018
1.92 2.208 4~6 508 840 037 799 132 030 .492 airfoil chord] airfoil chord] 2.325 1.324 5.424 6.568 7.434 8.093 8.568 8.868 8.990 8.916 8.593 8.045 7.317 6.450 5. 4.456 3.390 0 0 1. I. 2.370 3.357 4.800 5.908 6.823 8.277 9.366 9.635 7.841 6.784 5.654 0 1.324 1.599 2.004 2.728 3.831 4.701 8. 4.477 3.294 2. 1. 0 ordinates given ordinates given 10.176 10.730 11. 11.093 10.820 10.320 Ordinate Ordinate ------ radius
surface I surface
radius: radius: and and .5 .75 E.
NACA E.
.26.3 0 I. 2.5 5.0 7.5 .486 .950 40 45 55 65 75 80 85 90 95
7.095 NACA 10 15 20 25 30 3" 50 60 70
0 2.152 4.609 9.595 Upper Upper 100 L.
14.(iJ7 19.657 24.707 2<J.763 34.823 39.885 44.945 50.000 55.047 60.086 65.114 70.131 75.135 80.127 85. 90.077 95.037 L. Slope of Station Station 100.000 -------- ---- [Stations [Stations ~ ~ ~ I"%j ..... ~ t"' t:::1 ~
UJ q ?d o :> 6 .... :>
..... o --:r
of .
- .226 .327 0 0 0.168
5 0 0
-.661 -.781 -.944 -.792 -.393 -.037 percent percent ~.428 -L191 -1.536 -1.791 -1.999 -2.314 -2.547 -2.710 -2.814 -2.863 -2.854 -2.773 -2.606 -2.340 -2.004 -1.621 -1.211 -6.118 -1.124 -1.356 -1.702 -2.324 -3.245 -3.959 -4.555 -5.504 -6.223 -6.764 -7.152 -7.396 -7.498 -7.427 -7.168 -6.720 -5.403 -4.600 -3.744 -2.858 -1.977. -1.144 Ordinate Ordinate surface --- surface --- E.: in in L.
.5 .75 .628 .893 0 1. 2.50 5.00 7.50 0 1.411 5.203 7.711 Lower 2.682 Lower 10 15 20 25 40 45 50 55 60 65 70 75 85 90 95 given 10.212 15.202 20.183 54.971 given 25.157 30.128 35.097 40.064 45.032 50.000 59.947 64.927 69.915 74.910 79.912 84.924 89.943 94.971 100 Station , 30 '35 100.000 Station --- chord] chord]
65-41
652-01
through 0.687 1.505 702 977 .861 airfoil .937 airfoil .428 0 1.061 1.372 1. 2.800 3.487 4.067 5.006 5.731 6.290 6.702 6.983 7.138 7.153 7.018 6.720 6.288 5.741 5.099 4.372 3.577 2.729 1.8·12 1.124 1.356 1. 2.324 7.152 0 ordinates 0 3.245 3.959 4.555 6.223 6;764 7.396 7.498 7.427 7.168 6.720 6.118 5.403 4.600 3.744 2.858 1. 0 ordinates .5,504 . 1.144 Ordinate Ordinate ------ --- radius surface radius: of and radius: and
NACA
.5 .75 E. NACA E.
ppersurface .372 .607 0 1.25 2.50 5.00 7.50 0 1.089 2.318 4.797 7.289 9.788 10 15 20 25 30 40 45 50 55 60 65 70 75 90 95 Upper U 35 80 85 14.798 19.817 24.843 29.872 34.903 39.936 44.968 50.000 55.029 60.053 65.073 70.085 75.090 80.088 85.076 90.057 100 95.029 L. Slope L.
Station Station 100.000 ---- --- [Stations [Stations ----I of of 689 963 798 267 .010 .089 .278 0 0 0.084 0 0 0.168 -.719 -.859 -.711 -.293 percent -.810 -.956 -.75L -.282 percent -1.059 -1.385 -1.859 -2.221 -2.521 -2.992 -3.346 -3.607 -3.788 -3.894 -3.925 -3.868 -3.709 -3.435 -3.075 -2.652 -2.184 -1. -1.191 ~1.160 -1.490 -1. -2.314 -2.604 -3.049 -3.378 -3.613 -3.770 -3.851 -3.855 -3.759 -3.551 -3.222 -2.801 -2.320 -1. -1.
Ordinate Ordinate surface surface --- --- in E.: in E,: L. L.
441 914 .565 .822 .653 ,920 5.102 7.608 Lower 0 1. 2.717 5.243 7.753 Lower 0 1. 2.592 given 10.106 15.101 20.091 50.000 54.986 84.962 89.972 git'en 10.254 15.243 20.219 25,189 30.154 40.077 45.038 50.000 54.965 59.936 64. 79.897 94.967 25.079 30.064 35.049 40.032 45.016 59.973 64.964 69.957 74.955 79.956 35.116 69.899 74.893 84.910 89.934 '94.986 Station Station 100.000 100,000 --- chord] --- chord] ---
65-21
651-412
through through 1.000 0.687 .819 .999 .986 .622 airfoil airfoil 1.757 2.491 4.338 4.938 5.397 5.732 5.954 6.067 6.058 5.915 5.625 5.217 4.128 3.479 2.783 2.057 1.327 0 1.010 1.236 1.588 2.234 3.227 4.010 4.672 5.741 6.562 7.193 7.658 7.971 8.139 8.139 7.963 7.602 7.085 6.440 5.686 2.9.74 1.979 0 1.273 3.069 3.555 4.712 0 4.847 3.935 0 ordinates ordinates Ordinate Ordinate --- --- radius radius surface surface radius: of and radius: and
NACA 103
E. NACA .347 ,580 E.
.435 .678 0 1.059 2.283 4.757 9.746 7.247 0 1.169 2.408 4.898 7.394 9.894 Upper 14.757 19.781 24.811 29.846 34.884 39.923 44.962 50.000 55.035 60.064 65.086 70.101 75.107 SO. 85.090 90.066 95.033 Upper 14.899 19.909 24.921 29.936 34.951 39.968 44.984 50.000 55.014 60.027 65.036 70.043 75.045 80.044 85.038 90.028 95.014· L. 100.000 L. Slope of Slope Station Station 100.000 ---- ---- [Stations [Stations of of ,036 0 0 0.110 0 0 0.084 -.852 -.595 -.191 -,648 -.772 -.948 -.587 -.221 percent percent -1.012 -1.242 -1.625 -2.185 -2.606' -2.956 -3.5iJo -3.904 -4.197 -4.401 -4.518 -4.550 -4.475 -4.283 -3.968 -3.566 -3.124 -2.640 -2.131 -1.604 -1.085 -1.233 -1.643 -1.957 -2.217 -2.625 -2.930 -3.154 -3.310 -3.401 -3.425 -3.374 -3.233 -2.991 -2.672 -2.298 -1.884 -1.009 .-1.447 Ordinate Ordinate surface surface --- --- in E.'
in E.: L.
L. 152 .601 .862 0 1.376 2.644 5.163 7.671 .559 .816 Lower Lower given 10.173 15.167 20. 25.131 30.106 35.079 40.049 45.017 49.983 59.906 64.877 69.876 74.888 79.910 84.936 89.964 94.987 given 1.323 2.583 5.092 7.595 .54.949 100.000 10.096 15.091 20.082 25.071 35.044 40.029 45.014 50.000 54. 987 59.976 64.967 69.961 74.959 79.960 84.965 89.974 Station Station 30.058 94.987 100.000 --- --- chord] chord]
65-209
651-212
through through 1.000 0.552 520 042 ,982 .521
.748 .912 .596 a=0.6 airfoil
airfoil 0 1.194 1. 2.113 3.017 3.728 4.330 5.298 6. 6.611 7.029 7.304 7.444 7.423 7.231 6.856 6.318 5.634 4.842 3.983 3.082 2.173 1.297 0 0 1.162 1.605 2.275 2.805 3.251 3.971 4.522 4.944 5.254 .1.461 5.567 5.564 5.439 5.181 4.814 4.358 3.828 3.237 2.601 1.933 1.255 0 ordinates ordinates .
Ordinate Ordinate radius radius surface surface radius: of and radius: of and NACA E.
,399 ,638
E. NACA
,441 .684 0 1.124 2.356 7.329 9.827 ,4.837 24.869 29.894 34.921 39.951 44.983 50.017 55. 65.123 70.124 75.112 80.090 85,064 0 1.177 2.417 . 4.908 7.405 9.904 Upper 14.833 19.848 60.094 90.036 95.013 Upper 100.000 L. Slope 14.909 19.918 24.929 29.942 34.956 39.971 44.986 50.000 55.013 60.024 65.033 70.039 75.041 80.040 85.035 90.026 95.013 L. Slope Station Station 100.000 ---- [Stations [Stations of of 523 797 .007 .121 O· 0.084 0 0 0.084 -.956 -.429 -.040 -.870 -.502 -.608 -.768 -.993 -.963 -.699 -.437 -.192 percent -.424 percent -4.178 -4.510 -4.743 -4.854 -4.654 -4.317 -3.872 -3.351 -2.771 -2.164 -1.548 -1.036 -1.277 -1.686 -2.287 -2.745 -3.128 -3.727 -4.882 -4.926 -1.164 -1.306 -1. -1.685 -1.802 -1.880 -1.922 -1.927 -1.888 -1. -1.646 -1.447 -1.216 Ordinate Ordinate surface surface .
---- --- E.:· in E.: in L.
L. 957 ,577 ,836 1.346 2.609 5.122 7.627 .540 .794 0 Lower Lower 54.983 59.968 64. 94.983 10.127 15.121 20.110 25.094 30.077 35.058 40.039 45.019 50.000 69.950 74.947 79.948 89.967 given 0 1.300 2.556 5.061 7.563 given 100,000 -- Station 10.064 15.061 20.055 25.047 30.038 35.029 40.019 45.010 50.000 54.991 59.984 64.978 69.974 74.972 79.973 84.976 89.982 Station' - 84.955 94.991 100.000 1 --- chord] chord]
65-206
651-212
through through 1.000 0.240 ,672 .970 .524 .642 .822 .511 airfoil airfoil 6.507 6.C:!'4 5.411 0 1.176 1. 2.058 2.919 3.593 4.162 5.073 5.770 6.300 6.687 6.942 7.068 7.044 6.860 4.715 3.954 3.140 2.302 1.463 0 1.140 1.625 2.012 2.340 2.869 3.277 3.592 3.824 3.982 4.069 4.003 3.836 3.589 3.276 2.907 2.489 2.029 1.538 1.027 0 0 ordinates ordinates '4.078 Ordinate Ordinate ------ radius radius -- surface surface of radius: radius: of and and NACA E,
E. NACA
.423 .664 .460 .706 0 1.154 2.391 4.878 7.373 9.873 0 1.200 2.444 4.939 7.437 9.936 Upper Upper 50.000 55.017 60.032 65.043 95.017· 14.879 19.890 24.906 29.923 34.942 . 39.961 44.981 70.050 75.053 80.052 85.045 90.033 L. Slope 14.939 19.945 24.953 29.962 39.981 44.990 50.000 55.009 65.022 70.026 75.028 . 80.027 90.018 95.009 L. Slope Station .85.024 Station 100,000 .60.016 . 34.971 100.000 --- [Stations [Stations of of
I
0 0 -.947 -.356 -.700 -.845 -.738 -.923 percent -,280 percent -1.875 -3.172 -3.647 -4.402 -4.975 -5.406 -5.716 -5.912 -5.997 -5.949 -5.757 -5.412 -4.943 -4.381 -3.743 -3.059 -2,345 -1.630 -1.058 -1.421 -1.961 -2.383 -2.736 -3.299 -3.727 -4.050 -4.282 -4.431 -4.496 -4.469 -4.336 -4.086 -3.743 -3.328 -2.856 -2.342 -1.805 -1.260 -1.109 -1.387 -2.606 Ordinate Ordinate surface surface --- --- in in .5 .75 .5 .75 0 1. 2.5 5.0 7.5 0 1.25 2.5 5.0 7.5 Lower Lower 10 15 20 25 30 35 40 45 50 55 60 65 70 75 80 85 90 95 given 10 15 20 25 35 40 45 50 55 60 65 70 75 80 85 90 given 30 . 95 Station Station chord] chord]
65-009
651-012
1.000 0.552 961 387 .947 .356 .700 .845 ,738 ,280 .923 airfoil airfoil le875 4.402 4.975 5,406 5.716 5.912 5.997 5.949 5.757 5.412 4.943 4.381 3.743 3.059 2.345 1.630 0 0 1.058 1.421 1. 2.383 2.736 3.299 3.727 4.050 4.282 4.431 4.496 4.469 4.336 4.086 3.743 3.328 2.856 2.342 1.805 1,260 0 1.109 1. 2.606 3.172 3.647 0 ordinates ordinates Ordinate Ordinate ------ surface sUrface radius: and radius: and NACA E, .5 .75 .5 .75 E.
NACA
0 1. 2.5 5.0 7.5 0 1.25 2.5 5.0 7.5 15 20 25 30 35 40 45 50 55 60 65 70 75 80 85 90 95 10 15 20 25 30 35 40, 4S 50 55 60 65 70 75 80 85 90 95 10 UpPer Upper 100 L.
L.
Station Station
----
(Stations [stations
~ :>- >-3 .... o Z :>- t"' I:) H
....... o 00 !:C t'l "d o l:C >-3 Z o 00 t-:> ~ :>- -< .... U2 o l:C ~ (") o ~ ~ >-3 >-3 t'l M .." o l:C :>- t'l l:C o Z :>- q >-3 .... (") U2
of I
I / 0 0 0 0
8 0.084
-.801 -.173 -.546 percent of -1.282 -1.533 -1.902 -2.560 -3.546 -4.305 -4.937 -5.930 -6.676 -7.233 -7.622 -7.856 -7.928 -7.806 -7.465 -6.913 -6.196 -5.36.5 -4.454 -3.500 -2.541 -1.621 -1.522 -1.838 -9.91\2 -9.277 -8.390 -7.360 -6.224 -5024 -3.800 -2.598 -1.484 Ordinate -2.301 -3.154 -4.472 -5.498 -6.352 -7.700 -8.720 -9.487 Ordinate -10 -10.375 -10.499 -10.366 ---- _._-- in E.: L.
.612 .875 .50 .75 1.390 Lower surface 0 2.660 5.181 7.689 Lower surface 0 1.25 2.5 5.0 7.5 10.191 15.182 20.165 25.142 30.116 35.088 40.058 45.028 50.000 54.974 59.953 64.937 69.927 74.923 79.926 84.937 89.954 94.977 10 15 20 25 30 35 40 45 50 55 60 65 70 i5 80 85 90 95 Station 100.000 100 Station ---- ---
653-21 654.021
1.96 2.50 airfoil chord] .805 .546 airfoil chord] 1.673 2.116 0 1.382 2.932 4.178 5.153 5.971 7.276 8.270 9.023 9.566 9.916 9.996 9.671 9.103 8.338 7.425 6.398 5.290 4.133 2.967 1.835 0 0 1.522 1.838 2.301 3.154 4.472 5.498 6.352 7.700 9.487 9.952 9.277 7.360 6.224 5.024 3.800 2.598 1.484 0 ordinates given in percent ordinates given 8.720 8.390 10. 10036 10.375 10499 10.366 Ordinate ---- Ordinate --- surface radius through surface and radins: of and radius:
NACA .50 .75
E. NACA E.
.388 .625 0 1. 2.5 5.0 7.5 Upper 0 1.110 2.340 4.819 7.311 9.809 10 15 20 25 30 35 40 45 50 55 60 65 70 75 80 85 90 95 Upper 14.818 19.885 24.858 29.884 34.912 39.942 44.972 50.000 55.026 60.047 65.063 70.073 75.077 80.074 85.063 90.046 95.023 L. Slope 100 Station L.
100.000 Station ._-- ---- [Stations [Stations of , of 886 999 901 568 0 0
8 0 0 0.349
-.490 8
-.846 -.391 -.056 -1.337 -1.608 -2.014 -2.751 -3.866 -4.733 -5.457 -6.606 -7.476 -8.129 -8.595 -8. -8. -8. -8. -8.008 -7.267 -6.395 -5.426 -4.396 -3.338 -2.295 -1.319 -1.055 -1.239 -1.493 -1.895 -2.469 -2.884 -3.219 -3.716 -4.071 -4.321 -4.479 -4.551 -4.540 -4.407 -4.154 -3.815 -3.404 -2.936 -2.428 -1.895 -1.361 Ordinate Ordinate ---- --- E.: L.
.50 .75 244 0 1. 2.5 5.0 7.5 .941 Lower surface 10 15 20 25 30 35. 40 45 50 55 60 65 70 75 80 85 90 95 0 1. 1.811 5.752 Lower surface 3.154 8.294 100 10.806 15.775 20.699 25.593 30.463 35.316 40.151 44.966 49.727 54.532 59.454 64.443 69.481 74.555 79.653 84.761 89.867 94.954 Station Station 100.000 --- ---
0.5
653-01
653-61
through
=
1.96 469 821 090 1.96 595 886 999 901 008 319 .920 airfoil chord] .490 0 1. 1. 2.375 3.449 5.115 6.448 7.575 9.404 9.806 8.374 6.851 5.279 3.720 2.233 0
a airfoil chord]
0 1. 1.608 2.014 2.751 4.733 5.457 2.295 ordinates given in percent 3.866 6.606 7.476 8.129 8. 8. 8. 8. 8.568 8. 7.267 6.395 5.426 4.396 3.338 1. 0 10.815 11.893 12.687 13.209 13.456 13.395 12.9i4 12.173 11.
ordinates given in percent Ordinate --- Ordinate ----- snrface radius surface and radius: and radius:
NACA .50 .75
E. E.
NACA
0 1.25 2.5 5.0 7.5 .059 .256 .689 Upper 10 15 20 25 30 35 40 45 50 55 60 70 75 80 85 90 95 65 Upper 0 1.846 4.248 6.706 9.194 100 24.407 29.537 34.684 39.849 45.034 50.273 55.468 60.546 65.557 70.519 75.445 85.239 90.133 95.046 Station L. 14.225 19.301 80.347 L. Slope of Station 100.000 ---- --_.- [Stations [Stations of , .239 .463
5 0 0.233 0 0 0.253
-.957 8
-.527 -.139 -.943 -.268 -1.132 -1.377 -1.
-2.335 -2.746 -3.081 -3.591 -3.963 -4.232 -4.411 -4.508 -4.526 -4.431 -4.226 -3.929 -3.548 -3.104 -2.609 -2.083 -1.545 -1.014 -1.146 -1.356 -1.651 -2.152 -2.880 -3.427 -3.876 -4.564 -5.072 -5,433 -5.672 -5.792 -5.784 -5.616 -5.259 -4.723 -4.053 -3.302 -2.506 -1.
Ordinate Ordinate ---- --- E.: E.: L. L.
li5 661 .755 .828 Lower surface 0 1.036 1.573 2.874 5.426 7.946 1. 1.
Lower surface 0 2.974 5.538 8.
10.451 15.432 20.389 25.329 30.257 35.175 40.084 44.981 49.848 54.738 59.693 64.686 69.706 74.747 84.863 89.923 94.973 15.545 25.426 40.174 45.085 50.000 54.923 59.859 64.811 69.781 74.770 79.780 94.932 10.569 20.494 30.348 35.262 84.811 89.862 Station 100.000 100.000 . 79.801 Station --- ---
652-41
653-61
through through 459 755 1.505 433 1.96 .715 a=0.5 0 1.233 1. 1.965 2.812 4.099 5.122 5.985 7.383 8. 9.280 9.883 8.672 7.684 6.573 5.387 airfoil chord] 9.501 4.157 2.930 1. 0 0 1.446 1.776 2.293 3.268 4.776 5.971 6.978 8.602 9.848 9.482 8.338 7.075 5.719 4.306 2.863 1. 0 airfoil chord] 10.280 10.470 10.423 10.106 ordinates given in percent 10.803 11.504 11. 972 12.210 12.186 11.877 11.293 10.479 ordinates given in percent of Ordinate --- Ordinate radius --- surface radius surface and of radius: and radius: .149
NACA
E. NACA E.
.245 .464 .927 .172 .385 .839 0 2.126 4.574 7.0.14 9.
Upper 0 2.026 4.462 6.936 9.431 Upper 14.568 19.611 24.671 29.743 34.825 39.916 45.019 50.152 55.262 60.307 65.314 70.294 75.253 90.077 95.027 80.199 85.137 L. Slope 14.455 19.506 24.574 29.652 34.738 39.826 44.915 50.000 55.077 60.141 65.189 iO.219 75.230 80.220 85.189 90.138 95.068 L. Slope of Station 100.000 Station 100.000 .---- ---- [Stations [Stations of !
683 331 .206
5 0 0.168 0 0
0.233
-.628 -.107 8 -.702 -.201
-1.008 -1.200 -1.472 -1.936 -2.599 -3.098 -3.510 percent -4.150 -4.625 -4.970 -5.205 -5.335 -5.355 -5.237 -4.962 -4.530 -3.976 -3.342 -2.654 -1.952 -1.263 -1.164 -1.378 -1. -2.197 -2.951 -3.515 -3.978 -5.213 -5.595 -5.853 -5.998 -6.026 -5.905 -5.626 -5.216 -4.696 -4.094 -3.433 -2.734 -2.024 -1.
Ordinate -4.690 Ordinate ---
I
E.: E.: L. L.
.687 .958 0 1.484 2.769 5.303 7.816 Lower surface Lower surface 0 l:g~{ 1. 2.943 5.507 8.034 10.318 15.303 20.274 25.236 30.193 35.146 40.097 45.047 50.000 54.957 59.921 64.894 69.876 74.869 79.874 84.891 89.920 94.960 25.396 40.101 44.978 49.818 10.541 15.519 20.467 30.309 35.211 .14.687 59.636 64.628 69.653 74.702 79.768 84.841 89.911 94.970 Station 100.000 Station 100.000 ---
0.5
652-41
653-41
thl"Ough through
=
1.505 440 766 770 970 897 506 930 1.96 208 480 .777 airfoil chord] 2.271 5.891 6.882 9.820 8.674 7.397 6.038 4.636 3.247 0 airfoil chord] 0 1. 1. 3.233 4.715 8.482 9.709 1.
0 1. 1. 1.900 2.680 3.863 4.794 5.578 6.842 7.809 8.550 9.093 9.455 9.639 9.617 6.542 a
ordinates given in percent of 9.374 8.910 8.260 7.462 5.532 4.447 3.320 2.175 1.058 0 10.643 11.325 11. 11. 11. 11. 10.788 ordinates given in Ordinate Ordinate --- surface surface and radius: radius: and NACA E.
.313 .542 NACA E.
.197 .411 .868 1.016 2.231 Upper 0 4.697 7.184 9.682 0 2.057 4.493 6.966 9.459 Upper 14.697 19.726 24.764 29.807 34.854 39.903 44.953 50.000 55.043 60.079 65.106 70.124 75.131 80.126 85.109 90.080 95.040 L. Slope of radius 14.481 19.533 24.604 29.691 34.789 39.899 45.022 50.182 55.313 60.364 65.372 70.347 75.298 80.232 85.159 90.089 95.030 L. Slope of radius Station 100.000 Station 100.000 ---- ---- [Stations [Stations of i I I of 449 781 743 .144
5 0 0.084 0 0
0.168
percent -.626 -.112 8 -.946 -.282
-1.070 -1.282 -1.591 -2.134 -2.925 -3.532 -4.035 -4.829 -5.426 -5.868 -6.179 ,-6.366 -6.427 -6.332 -6.065 -5.625 -5.047 -3.628 -3.848 percent -4.373 -2.061 -1.303 -1.218 -1. -1. -2.360 -3.217 -3.870 -4.410 -5.250 -5.877 -6.334 -6.648 -6.824 -6.856 -6.711 -6.362 -5.818 -5.124 -4.334 -3.480 -2.603 -1.
Ordinate
Ordina~
--- I in E.: E.: in L.
368 L.
.594 .855 .722 .997 0 1. 2.635 5.152 7.658 Lower surface Lower surface 0 1.527 2.819 5.361 7.877 15.152 10.159 ·20.137 25.118 30.096 35.073 40.048 45.024 50.000 54.979 59.961 64.947 69.938 74.935 79.937 84.945 89.960 94.980 45.057 54.949 59.906 64.874 69.854 74.846 79.853 94.954 10.381 15.364 20.329 25.284 30.232 35.175 40.116 50.000 84.873 89.908 Station 100.000 100.000
Statio~1
---
652-21
653-41
through through 1.505 418 729 1.96 018 426 2.300 0 airfoil chord] .744 0 1. 1. 2.20jJ 3.104 4.481 5.566 6.478 7.942 9.061 9.914 9.408 8.454 7.368 6.183 4.927 3.63"8 1.120 1.170 1.422 1.805 airfoil chord] 0 2.506 3.557 4.380 5.069 6.175 1. 8.123 8. 8.569 8.522 7.815 7.189 6.433 5.572 4.638 2.649 ordinates given 3.653 1. 0 10.536 Hi 244 11.140 11.091 10.774 10.1£8 '7.·658 '8.271 ordinates given Ordinate --- Ordinate .
radius surface surface and radius: of radius: and .406 .645 NACA E. E.
NACA
1.132 0 2.365 4.848 7.342 9.841 .278 .503 .973 Upper 14.848 19.863 24.882 29.904 34.927 39.952 44.976 50.000 55.021 60.039 65.053 70.062 75.065 80.063 85.055 90.040 95.020 0 2.181 4.639 7.123 9.619 Upper 50.000 55.051 60.094 75.154 95.046 100.000 L. Slope 14.636 19.671 24.716 29.768 34.825 39.884 44.943 65.126 70.146 80.147 85.127 90.092 L. Slope of radius Station station_I 100.000 --- [Stations [Stations
Ul c::1 ~ ~ ~ o "'i >- ...... ::0 "1 o ...... t-< t::I ~
>- I--'
o eo
of .114 .038 0 0 0.168 0 0 0.168 percent -.373 -- percent of -.549 -1.946 -3.602 -i,337 -1.584 -2.614 -4.370 -4.992 -5.973 -6.701 -7.235 -7.606 -7.819 -7.856 -7.676 -7.260 -6.635 -5.842 -4.940 -3.973 -2.983 -2.016 -'-1.121 -1.068 -1.261 -1.524 -1.990 -2.646 -3.147 -3.581 -4.250 -4.764 -5.156 -5.448 -5.645 -5.766 -5.807 -:;.751 -5.590 -5.282 -4.771 -4.024 -3.173 -2.253 -1.373 Ordinate Ordinate --- in E.: in E,: -
L. 5)-41
018 550 L.
.742 .697 .968 Lower surface 0 1. 1. 2.848 5.397 7.917 Lower surface 0 1.492 2.775 5.307 7.820 10.421 15.404 20.366 25.316 40.129 45.063 50.000 - 30.258 35.195 54.944 59.897 64.862 69.840 74.833 79.841 84.864 89.902 94.951 10.323 15.309 20.280 25.243 30.199 35.152 40.102 45.051 50.000 54.950 59.904 64.865 69.839 74.826 79.830 84.850 89.889 94.944 Station 100.000 Station 100.000 ------ ---- through through
65(421)-420 2.27 1.575
537 952 226
airfoil chord] 66(21
airfoil chord] 0 1. 1.864 2.374 3.358 4.866 6.066 7.060 8.665 9.885 9.060 7.861 6.563 5.200 3.813 2.441 1.150 0 1.268 1.541 1. 2.734 3.910 5.649 6.942 7.948 8.736 9.765 9.970 9.566 8.891 7.912 6.753 5.437 4.065 2.617 1. 0 ordinates given ordinates given 0 4.843 9.336 10.815 11.494 11.939 12.140 12.056 11.672 11.015 10.126 10.050 10.187 10.163 Ordinate
Ordinate ._--
surface I radius
surface ---- and radius: and of E. radius: E.
0 .258 .482 .950 .303 .532 7.083 2.225 9.677 Upper 2.152 4.603 9.579 4.596 0 1.008 4.693 7.180 Upper 19.634 14.691 19.720 24.757 55.050 70.161 NACA 24.684 29.742 34.805 39.871 44.937 50.000 5.5.056 60.103 65.138 70.160 75.167 80.159 85.136 90.098 95.049 L. Slope of radius 29.801 34.848 39.898 44.949 50.000 60.096 65.135 75.174 80.170 85.150 90.111 . 95.056 L.
Slope Station Station' 100.000 100.000
NACA
--- ---
[Stations [Stations I , !
, 230 534 138 528 112 608 0 0 0 0 0.110 0.042 -.432 percent of -.810 -.241 -1.023 -1. -1. -2.870 -3.482 -4.800 -5.410 -6.388 -1.319 -2.869 -4.702 -5.741 -6.080 -6.312 -6.462 -6.523 -6.336 -6.048 -4.866 -4.037 -2.177 -1.235 Ordina~ -2.075 -3.992 -5.865 -6.189 -6.462 -6.384 -6. -5.724 -5.174 -4. -3.807 -3.046 -2.269 -1.509 -1. -'-I. -2.127 ~3.441 -3.934 -5.290 -6.483 -5.574 -3.107 surface ~rdinatel
I
in E.: E.: L. L.
.544 .799 .893 .629 Lower 0 1.305 2.563 5.071 7.574 0 1.409 2.683 5.206 7.716 Lower surface 10.074 15.071 20.064 25.055 50.000' 59.982 64.975 10.219 15.212 20.194 25.168 30.138 35.103 40.064 45.022 49.977 54.927 59.859 64.809 69.802 74.819 79.852 84.894 89. 94.979 30.045 35.034 40.023 45.011 54.990 69.971 74.969 79.971 84.975 89.981 94.991 Sta:ion Station 100.000 100.000 --------
I
through through 1.311 1.575 073 300 642 326 501 886 937
66(215)-21 .762
airfoil chord] .557 a=0.6 airfoil chord]
0 1. 1. 1. 5.472 0 1.242 1. 1. 2.615 3.701 4.563 5.308 6.500 7.428 8.155 8.708 9.098 9.356 9.471 9.431 9.224 8.800 8.084 7.068 5.889 4.585 3.265 1. o ordinates given 2.261 3.186 3.906 4.508 6.206 6.761 7.161 7.418 7.534 7.480' 7.242 6.820 6.246 5.558 4.779 3.942 3.065 2.181 1. 0 ordinates given in percent of Ordinate Ordinate radius radius surface surface radius: radius: of and and 091 181 ppcr E. E.
.456 .701 .371 .607 0 1.195 2.437 4.929 7.426 9.926 0 1. 2.317 4.794 7.284 9.781 U Upper NACA 65(215)-114 14.929 19.936 24.945 29.955 34.966 39.977 44.989 50.000 55.010 60.018 65.025 70.029 75.031 80.029 85.025 90.019 14.788 19.806 24.832 29.862 34.897 39.936 44.978 50.023 55.073 60.141 65.191 70.198 75. 80.148 85.106 90.061 9.5.021 95.009 L. Slope of L. Slope Station Station 100.000 100.000
NACA
------- ------~ [Stations [Stations of of I I I I ' I I I I I I !
0 0 0 0.084 0.233 -.281 percent -.867 -.257 percent -1.344 -1.603 -1.965 -2.595 -3.551 -4.275 -4.869 -5.780 -6.455 -6.952 -7.293 -7.486 -7.370 -5.057 -2. -1.130 -1.344 -1.644 -2.188 -2.972 -3.580 -4.106 -4.930 -5.564 -6.054 -6.422 -6.676 -6.8.18 =~:~; -6.685 =~:~8i -4.997 =~:g;g -2.049 -1.069 -7.526 -7.010 -6.484 -5.818 -4.229 -3.360 -1.634 Ordinate Ordinate --- in E.: in E.: L. 5)-21 L.
.845 .599 .860 -- 0 1.137 1.687 3.008 1. 2.638 5.154 7.660 Lower surface 5.586 8.120 Lower surface 0 10.629 15.605 20.545 64.572 10.162 15.155 20.140 25.121 30.100 35.076 40.051 45.026 50.000 54.975 59.952 64.938 69.919 74.913 79.915 84.925 89.945 94.972 25.462 30.361 35.246 40.118 44.974 49.789 54.638 59.579 69.602 74.660 79.736 84.819 89.900 94.966 100.000 100.000 Station Station --- -------
654-421 -
through 1.575 2.50 620 991 230 484 736 496
.833 66(21 .913
a=0.5 airfoil chord] airfoil chord]
1. 6.941 0 0 1. 2.553 3.631 5.315 6.651 7.773 9.572 9.637 8.193 6.664 5.097 3.550 2.095 0 - 0 1. 1. 1.858 2.560 3.604 4.428 5.140 6.276 7.156 7.844 8.366 8. 8.980 9.092 9.060 8.875 8. 7.862 5.860 4.644 3.395 2.103 ordinates given ordinates given 10.951 12.0nO 12.765 13.258 13.470 13.362 12.890 12.056 10.942 Ordinate --- Ordiuate radius through surface radius: of and radius: and 992 000 E.
NACA E.
.155 .363 .813 .401 .640 0 1. 4.414 1.128 2.362 4.846 7.340 9.838 Upper surface 6.880 9.371. 0 Upper 14.395 19.455 24.538 29.639 M.754 45.026 50.211 50.362 65.428 75.340 95.034 14. 19.860 24.879 29.900 34.924 39.949 44.974 50.000 55.025 60.048 65.067 70.081 75.087 80.08.5 85.075 90.055 95.028 39.882 60.421 70.398 80.264 85.181 90.100 Slope of radius I,. Slope Station L.
100 Station 100.000
NACA
---- ------ ----
[Stations [Stations i I I I I .088 0 0 0 0.168 -.446 -.597 percent of percent of -I.-WI -1.676 -2.065 -2.761 -3.821 -4.633 -5.303 -6.342 -7.120 -7.691 -8.088 -8.313 -8.356 -8.176 -7.746 -7.087 -6.247 -5.299 -4.278 -3.231 -2.204 -1.248 -1.184 -1. -1. -2.378 -3.292 -4.007 -4.626 -5.605 -6.362 -6.950 -7.395 -7.706 -7.909 -7.997 -7.957 -7.780 -7.425 -6.832 -5.970 -4.966 -3.849 -2.723 -1.587 Ordinate Ordinate --- ---- in E.: in
I,. 5)-01
567 25 .753 .5 .75 0 1.032 1. 2.865 5.417 1. 2.5 5.0 7.5 Lower surface 7.938 Lower surface. 0 10.443 15.425 20.384 2.5.332 30.271 &5.204 40.135 45.066 50.000 54.911 59.892 64.855 74.824 79.833 10 15 20 25 30 35 40 45 50 55 60 65 70 75 80 85 90 95 69.832 84.857 89.896 94.949 Station 100.000 100 Station
--- ----
654-421
through 2.50 1.575 976 755 587 .597
airfoil chord] 66(21
airfoil chord] 0 1.601 1.956 2.493 7.395 7.909 7.957 7.780 7.425 6.832 5.970 4.966 2.72.3 1. 0 3.505 5.085 6.329 7.371 9.034 9.419 8.166 6.811 5.388 3.940 2.514 1.176 0 0 1.184 1.418 1. 2.378 3.292 4.007 4.626 5.605 6.362 6.950 7.706 7.997 3.849 ordinateS" given ordinates given 10.304 11. 12.158 10.531 11.271 12.433 12.640 12.556 11.467 Ordinate --- Ordinate --- radius surface surface radius: and radius: and 796 .5 NACA E. .75 E.
.247 .468 .933 0 1. 2.5 5.0 7.5 0 2.135 4.582 7.062 9.557 10 15 20 25 30 35 40 45 50 55 60 65 70 75 80 85 90 95 Upper Upper •.
14.575 19.616 24.668 29.729 34. 39.865 44.934 50.000 55.059 60.108 65.145 70.168 7.5.176 80.167 85.143 90.104 95.051 J Slope of Station L. Station 100.000
. - NACA
--- [Stations [Stations of I I I I I ' 0 966 578 --
2 0
0 0 0 0.084 0.084 -c. -.229 -.847 -.726 -.160 percent percent -1.467 -1.762 -2.188 -2.963 -4.151 =g:~~~ -7.024 =U~i -9.063 -9.344 -9.428 -9.271 -8.849 -8.182 -7.319 -6.330 -5.251 -4.128 -3.003 -1.924 -1.010 -1.233 -1.614 -2.165 -2.593 -2.963 -3.539 -3.982 -4.322 -4. -4.756 -4.863 -4.903 -4.869 -4.749 -4.523 -4.135 -3.563 -2.893 -2.167 -1.424 Ordinate Ordinate
--- ---
in E.: in E.: L. L.
.628 .892 .834 .576
1-21
0 1.410 2.684 5.209 7.720 1.343 2.605 5.116 Lower surface 0 7.621 Lower surface 10.222 15.213 20.192 25.166 30.135 35.102 40.068 45.033 50.000 54.970 59.946 64.928 69.916 74.912 79.916 94.974 10.122 15.116 20.105 25.091 30.075 35.057 40.038 45.019 50.000 54.981 59: 64.949 69.939 74.934 79.935 84.942 89.957 94.978 84.928 89.948 Station 100.000 100.000 Station ------- ---
654-221
through through 66, 0.893 2.50 902 055 447 797 .792 airfoil chord] .861 .947 airfoil chord] 0 1.567 1. 3.335 4.783 5.918 5.918 1.986 7.093 6.939 5.507 4.683 2.770 1.760 6.865 8.370 9.514 9.461 8.390 7.195 4.595 3.270 2.000 0 0 1.150 1. 2. 3.441 3.997 4.885 5.574 6.112 6.522 6.816 7.005 7.075 6.665 6.195 3.759 0 - '2.402 10.381 11.007 11.404 11. 11.461 11. 10.372 ordinates given Ordinate ------ Ordinate --- radius surface surface of radius: radius: and E.
NACA E.
.372 .608 .424 .666 0 1.090 2.314 4.791 7.280 9.778 1.157 2.395 4.884
Upper NACA 0 7.379 9.878
Upper 14.787 19.808 24.834 29.865 34.898 39.932 44.967 50.000 55.030 60.054 65.072 70.084 75.088 90.052 95.026 14.884 19.895 24.909 29.925 34.943 39.962 44.981 50.000 55. 60.036 65.051 70.061 75.066 80.065 85.058 90.043 95.022 80.084 85.072 Slope L. Slope of radius Station L.
100.000 Station 100.000 ---- ---- [Stations and ordinates given lStati~ns >-:3 ,.... ::0 M "C o ::0 >-:3 Z 9 00 t>:> ... ~ > >-:3 .... o Z > t"' > C) <! .... Ul o ::0 ><l (") o ~ ~ .... >-:3 M M "1 g > M ::0 o Z > q >-:3 H (") Ul
,..... o
of
I of
.068 0 0 0 0 -.706 -.840 -.536 0.Og4 -.092 -.451
0 percent
-1.031 -1.327 -2.ll0 -2.8.'\6 -3.082 -1.187 0.168 -I. -2.389 -:1.204 -3.467 -3.664 -3.802 -3.905 -3.R68 -1.770 -3.594 -1.272 -2.815 -2.281 -1.697 -1.099 -1.006 -1.445 -1.848 -2.454 -2.921 -3.313 -3.932 -4.397 -4.749 -5.009 -5.189 -5.287 -5.305 -5.244 -5.093 -4.816 -4.311 -3.630 -2.839 -2.003 -1.180 Ordinate surface --- E.: in E.: .686 .956 Lower surface .564 .821 0 1.481 2.759 5.289 Lower 7.801 l.:l2<J 0 2.588 5.098 7.6!)] 10.304 15.291 20.264 25.229 30.188 35.143 40.096 45.048 50.000 54.954 59.910 64.874 69.850 74.838 79.841 89.896 94.947 84.861 10.102 15.097 20.088 25.076 30.063 40.032 45.016 50.000 .54.984 59.970 64.()58 69.949 74.944 79.945 94.981 Station 100.000 35.048 84.951 89.963
~~:~~I~rdinat~
100.000 --- chord]
66-21 ----
662-415
----_._- 1.435 0.662 245 570 206 airfoil chord] .806 .980 .724 ",iI'foil 1. 1.467 1.873 0 I. 1.099 2.40\ 2.958 3.432 4.202 4.796 5.257 5.608 5.862 6.024 IL095 6.074 5.960 5.n6 5.132 4.759 4.071 3.289 2.445 1. 0 0 2.592 3.718 4.617 5.381 6.624 7.581 8.329 8.897 9.309 9.571 9.685 9.656 9.173 9.100 8.431 7.518 6.419 5.187 3.872 2.519 1.196 0 ordinall'" p:iven ill IJl'rcent ordinaws given Ordinate Ordinate --- surface surface --- and radius: and radius:
NACA
P84 E. NACA
E.
.436 .679 .314 .544 0 1.171 2.412 7.399 0 1.019 2.241 4.711 9.696 Upper 4.902 9.898 Upper 7.199 14.90:l 19.912 24.924 29.937 44. 50.000 55.016 65.042 70.051 75.056 95.019 14.709 19.736 24.771 29.812 34.857 39.904 44.952 50.000 55.046 60.090 65.126 70.1.<;0 75.162 80.1.<;9 90.104 95.053 :)4.952 :l9.968 60.030 80.055 85.049 90.037 L. Slope of radius through L. 85.139 L. SlopeofradiusthroughL.
Station Station 100.000 100.000 ----- --- [S(at.lons [Stations I I I I I 0.084 0 0.084 -.635 -.752 -.921 -.937 -.443 -.058 -.9il -.249 0 0 -1.180 -1.562 -1.857 -2.107 -2.504 -2.804 -3.031 -3.201 -3.318 -3.386 -3.404 -3.372 -3.286 -3.133 -2.852 -2.456 -1.982 -1.466 -?058 -1.269 -1. -2.045 -2.781 -3.354 -3.838· -4.6ll -5.198 -5.647 -5.983 -6.220 -6.359 -6.400 -6.347 -6.188 -5.888 -5.342 -4.603 -3.736 -2.801 -1.856 Ordinate ----
I~rdinate
E.: E.: L.
366 L.
.558 .594 .854 .814 0 1.321 2.580 .<;.088 7.591 0 1. 2.630 5.145 Lower surface Lower surface 7.651 10.092 15.088 20.079 25.069 30.056 35.043 40.029 45.014 50.000 54.986 59.973 64.962 69.954 74.950 79.950 94.982 10.152 15.146 20.132 25.114 30.094 35.071 40.048 45.024 50.000 54.977 59.955 64.937 69.925 74.919 94.974 84.956 89.965 79.921 84.930 89.948
Station 100.000 Sta~ion 100.000
----
66-209
662-215
throngh 1.435 .
0.530 135 552 .<;01 .735 .892 .690 .881 "irfoil chord] airfoil chord] I. 1.409 0 1. 2.194 2.705 3.141 3.850 4.39R 4.821 5.145 5.378 5.528 5.594 5.578 5.476 5.275 4.912 4.400 3.772 3.058 2.283 1.477 0 0 1.168 1.778 2.417 3.413 4.202 4.872 5.957 6.790 7.437 7.927 8.280 8. 8.590 8.551 8.378 8.030 7.402 6.547 5.526 4.393 3.202 2.005 0 ordinates giv('n in percNlt of ordinates given in percent. of Ordinate Ordinate --- radius through surface surface and radius: of and radius:
NACA
NACA
E. E.
.442 .686 .·103 .646 1.179 1.134 Upper 0 2.420 4.912 7.409 9.908 0 2.370 4.855 7.349 9.848 Upper 14.912 19.921 19.868 24.931 29.944 34.957 39.971 44.986 50.000 55.014 60.027 65.038 70.046 75.050 80.050 85.044 90.034 95.018 14.854 24.886 29.906 34.929 39.952 44.976 50.000 55.023 60.045 65.063 70.075 75.081 80.079 85.070 90.052 95.026 L. Slope L. Slope of radius Rtation Station 100.000 100.000 --- ~------. I
I
[Stations [Stations I I I I I I I I .054 0 0.084 0 0 -.409 -.482 -.584 -.730. -.940 -.747 -.434 -.148 -.566 percent of percent of __ -1.099 -1.234 -1.445 -1.604 -1.i23 -1.81f) -1.8R9 -1.900 -1.905 -1.882 -1.830 -1.744 -1.581 -1.344 -1.058 -1.122 -1.343 -1.675 -2.235 -3.100 =!:m -5.286 -5.99.5 -6.543 -6.956 -7.250 -7.430 -7.495 -7.450 =~:~ -6.372 -5.576 -4.632 -3.598 -2.530 -1.489 Ordinate Ordinate ---- --- in E.: in L.
2,5 .539 .793 .5 .75 Lower surface 0 1.298- 2.553 5.059 7.561 Lower surface 0 1. 2.5 5.0 7.5 10.061 15.058 20.053 25.046 30.038 35.029 40.019 45.010 50.000 54.991 59.982 64.974 69.969 74.966 79.996 94.988 10 15 20 2.<; 30 35 40 45 .<;0 55 60 65 70 75 90 84.969 89.977 ~iven 80 85 ·95 Station 100.000 Statiou 100 ---- ----
66-206
662-015
through 0.223 1.435 947 343 489 .509 .622 airfoil chord] .578 .566 airfoil chord] -.798 1.102 1.572 1.
ordinates given 0 2.268 2.791 3.196. 3.513 3.754 3.929 4.042 4.095 4.088 4.020 3.886 3.641 3.288 2.848 2.3.39 1.780 1.182 0 0 1.122 1. 1.675 2.235 3.100 3.781 4.358 5.286 5.995 6.543 6.956 7.250 7.430 7.495 7.4.<;0 7.283 6.959 6.372 5.576 4.632 2.530 1. 0 ordinates 3.598 Ordinate Ordinate
--- ---
radius surface surface and radius: of and radius:
NACA
.5 .75 E. NACA E.
.461 .707 0 1.25 2.5 5.0 7.5 1.202 Upper 0 2.447 4.941 7.439 9.939 10 15 20 25 30 35 40 45 50 55 60 65 70 75 80 85 90 95 Upper 14.942 19.947 70.031 24.954 29.962 34.971 39.981 44.990 50.000 55.009 60.018 65.026 75.034 80.034 85.031 90.023 95.012 100 L. Slope L.
Station Station 100.000
---- ----
[Stations [Stations oC I I - 0 0 0 -.687 -.824 -.961 -.374 -.853 -.716 -.157 0.084 percent of -1.030 -3.178 -4.457 -1.368 -1.880 -2.283 -2.626 -3.601 -3.927 -4.173 -4.348 -4.499 -4.475 -4.381 -4.204 -3.882 -3.428 -2.877 -2.263 -1.611 -1.014 -1.248 -1.619 -2.177 -2.611 -2.977 -3.559 -4.004 -4.342 -4.595 -4.773 -4.876 -4.905 -4.862 -4.741 -4.109 -3.543 -2.871 -2.1-17 -1.409 Ordinate Ordinate .-4.517
--- ---
-- in E.: L.
.50 .75 344 Lower sui-face 0 1. 2.5 5.0 7.5 .576 .834 Lower surface 10 20 25 30 35 40 45 50 55 60 65 70 75 80 85 90 95 0 1. 2.605 5.117 7.621 Station 100 Station 10.122 15.117 20.106 25.092 30.075 3.5.057 40.038 64.949 45.019 50.000 54.981 59.964 69.939 74.934 79.935 84.943 89.957 94.978 ---- 100.000
66-009
661-212
through 0.530 0.952 --- airfoil chord] .687 .824 .961 .374 .953 airfoil chord] .789 1.030 1.368 1.880 0 2.283 2.626 3.178 3.601 3.927 4.173 4.348 4.457 4.499 4.475 4.381 4.204 3.882 3.428 2.877 2.263 1.611 0 0 1.154 1.462 1.991 2.809 3.459 4.011 4.095 5.596 6.132 6.539 6.833 7.018 7.095 7.068 6.931 6.659 6.169 5.487 4.661 2.755 1.
ordinates given in percent ordinates given 3.739 0 Ordinate Ordinate ------ surface radius surface and radius: and radius: of .50 .75 E. N·ACA E.
0 1.25 2.5 5.0 7.5 .424 .666 10 15 20
. NACA Upper 25 30 35 40 45 50 55 60 65 70 75 80 85 90 95 0 1.156 2.395 4.883 7.379 9.878
Upper 100 14.883 19.894 24.908 29.925 34.943 39.962 44.981 50.000 55.019 60.036 65.051 70.061 75.066 80.065 85.057 90.043 95.022 L. L. Slope Station Station 100.000
------- ---
[Stations [Stations of I I
474 _J
0 0 -.461 -.554 -.693 -.918· percent -.665 -.262 -.906 0.
percent of -1. -2.119 -2.401 -2.971 -1.524 -1..752 -2.618 -2.782 -2.899 -3.000 -2.985 -2.925 -2.815 -2.611 -2.316 -1.953 -1.543 -1.107 -1.087 -1.358 -1.808 -2.496 -3.037 -3.496 -4.234 -4.801 -5.238 -5.568 -5.803 -5.947 -6.000 -5.965 -5.836 -5.588 -5.139 -4.515 -3.767 -2.944 -2.083 -1.234 Ordinate Ordinate ---
----
in in -- ----- .50 .75 .75 .5 Lower surface 0 1. 2.5 5.0 7.5 0 1.25 2.5 5.0 7.5 Lower surfilCe 10 15 20 25 30 35 40 45 50 55 60 65 70 75 80 90 95 10 15 20 25 30 35 40 45 50 55 60 65 70 75 90 95 85 80 85 Station 100 Station ---- '
66-006
661-012
- 0.223 ---- 0.952 752 543 087 358 234 .693 .918 airfoil chord] .461 .554 .665 .262 .906 .474 airfoil chord] 0 1.257 1.524 1. 2.119 2.618 2.782 2.899 2.971 3.000 2.985 2.925 2.815 2.611 2.316 1.953 1. 1.107 0 0 1. 1. 1.808 2.496 3.037 3.496 4.234 4.801 5.238 5.568 5.803 5.947 6.000 5.965 5.836 5.588 5.139 4.515 2.944 2.083 1. 0 ordinates given ordinates given 3.767 :2.401
Ordinate --- Ordinate ------
surface surface and radius: and radius: ----
NACA
.50 .75 .5 .75 E. NACA E.
1.25 0 2.5 5.0 7.5 ---- 0 1.25 2.5 5.0 i.5 10 15 15 Upper 20 25 30 35 40 45 50 55 60 65 70 75 80 85 90 95 10 20 25 30 35 40 45 50 55 60 65 70 75 80 85 90 95 Upper 100 100 L. L.
Station Station
---- ----
, [Stations [Stations ~ ~ ~ >-< ~ >-< ~ rJl d o "1 ::d "1 o t; ~ ;...
....- - -
01 of I
I I
738 793 52.5 0 0 -.717 0 0 0.274 -1. -1.804 -2.240 -3.045 -4.269 -5.283 -6.052 -7.369 -8.376 -9.1sa -9. -8. -7.610 -6.251 -4.796 -3.324 -1.924 -9.692 -.994 -.&38 -.247 -10.154 -10.407 -10.500 -10.434 -10.186 -1.160 -1.406 -1. -2.349 -2.730 -3.038 -3.501 -3.845 -4.095 -4.286 -4.411 -4.485 -4.493 -4.462 -4.381 -4.215 -3.992 -3.622 -3.053 -2.344 -1.578 Ordinate Ordinate --- --- E.: L.
.5 .75 .817 I,ower surface 0 l.25 2.5 5.0 7.5 5.513 Lower surface 0 1.102 1.648 2.959 8.028 10 15 40 55 70 20 25 30 35 45 50 60 65 75 80 85 90 95 10.524 15.479 20.402 25.302 :lO.182 35.036 39.824 44.636 49.5.13 54.526 59.546 64.607 69.727 74.8.16 79.893 84.934 89.963 94.985 Station 100 Station 100.000 ---- ----
664-021
J
through 2.550 1.544 525 738 793 592 546 airfoil chord] .717 .
airfoil chord] .639 0 1. 1.804 2.240 3.045 4.269 5.233 6.052 7.369 8.376 9.153 9. 9.692 8. 7.610 6.251 4.796 1.924 0 l.318 1.622 5.488 7.827 5.838 1.
ordinates given in percent 3.324 ordinates given in percent 0 2.106 3.01ti 4.4ll 6.390 8.897 9.687 9.516 8.753 7.859 6.878 4.78.1 3.692 2 0 - 10.154 10.407 10.500 10.434 10.186 10.497 10.499 10.216 10.121 Ordinate
--- Ordinate ----
su~ace. 1 radius
surface and radius: and radius: of .75 .5 - NACA E. E.
1.25 0 2.5 5.0 7.5 .183 .398 .852
NACA747A415
10 15 20 25 30 35 40 45 50 55 60 65 70 75 80 85 90 95 0 2.041 4.487 6.972 9.476 Upper 100 14.521 19.1'98 24.698 29.818 34.964 40.176 45.364 50.447 55.474 60.454 65.393 70.273 75.164 80.107 85.066 90.037 95.015 L.
Station L. Slope Station .. 100.000
_.~~prr
----- _ ---- {Stations [Stations / of of
I
0 0 0.168 0 0 0.232 -.676 -.011 -.405 percent -1.205 -1.412 -l. -2.256 -3.042 -3.651 -4.163 -4.977 -5.589 -6.0sa -6.399 -6.639 -6.775 -6.808 -6.736 -6.543 -6.180 -5.519 -4.651 -3.658 -2.610 -1.584 -1.031 -1.207 -1.473 -1.927 -2.518 -2.952 -3.304 -3.M3 -4.247 -4.546 -4.773 -4.926 -5.020 -5.040 -5.014 -4.930 -4.772 -4.509 -4.ll0 -.3.502 -2.743 -1.915 -1.097 Ordinate Ordinate
--- ----
E.: in E.: L. L.
.991 .771 .720 L 519 1.051 1.589 5.436 7.947 Lower surface 0 2.806 5.344 7.860 Lower surface 0 2.891 54.sa7 10.364 15.349 20.317 25.274 30.225 35.171 40.115 45.057 50.000 54.944 59.893 64.851 69.822 74.809 79.815 84.838 89.880 94.940 10.442 15.401 20.332 25.242 30.133 34.999 39.800 44.625 49.553 59.565 64.6:34 69.759 74.870 79.927 84.962 89.984 94.996 - Station 100.000 Station 100.000
------- ---- ------- ---
663-41
.- through through
747A315
1.544 1.955 U59 923 599 206 airfoil chord) airfoil chord] .481 0 1.405 1.692 2.147 3.000 4.306 5.347 6.231 7.669 8.773 9.633 9.639 8.539 7.238 5.794 4.276 2.744 1.275 0 0 1.30.> 1. 2.065 2.935 4.264 5.2l'6 6.140 7.497 8.503 9.242 9.731 9.982 9.962 9.572 8.964 8. 7.324 6.365 5.354 4.336 3.295 2.257 1.289 0 ordinates !!ivcn in percent ordinates given 10.287 10.759 11. 11.188 11.148 10. 10.464 Ordinate --- Ordinate ---- radiUS surface surface radius: and radius: and 668 E.
NACA E.
.280 .509 .229 .449 .911
'.981 NACA
Upper 0 2.194 4.656 7.140 9.636 0 2.109 4.564 7.05.1 9.558 Upper 19. 55.463 14.651 19.683 24.726 29.775 34.829 39.885 44.943 50.000 55.056 60.107 65.149 70.178 75.191 80.185 85.162 90.120 95.060 14.599 24.758 29.867 35.001 40.200 45.375 50.447 60.435 65.366 70.241 75.130 80.073 85.0~8 90.016 95.004 L.
L. Slope of radius Slope of Station Station 100.000 100.000
------ ---- ----
------- (Stations { Stations of of ' - 5Q.1 0 0 0 0.084 -.329 -.471 -- percent 0.084 - -1.268 -1.840 -2.456 -3.370 -4.085 -4.690 -5.658 -6.390 -6.952 -7.3iS -7.fi71 -7.847 -7.903 -7.839 -7.638 -7.252 -6.550 -.>.624 -4.51\5 -3.409 -2.260 -1.19fi -1.11:1 -1-320 -1.553 -2.205 -2.925 -3.473 -3.913 -4.608 -5.143 -5.558 -5.881 -6.125 -6.288 -6.380 -0.394 -6.326 -6.160 -5.875 -5.429 -4.725 -3.743 -2.65.1 -1.
Ordinate ·-1.496 Ordinate --- --- inpcrccnt in E.: K: L.
L.
.611 .872 .598 .858 1.38S Lower surface 0 2.654 5.173 7.680 Lower surface 1.372 2.639 5.1.52 7.656 Ii 10.182 15.175 20.159 25.137 30.113 35.086 40.0OR 45.029 50.000 54.972 59.946 64.925 69.911 74.905 7l'.907 84.919 89.940 94.970 10.155 15.146 20.131 25. 30.092 35.070 40.047 45.024 50.000 54.976 59.953 64.932 69.914 74.902 79.900 M.908 89.929 94.963 Station 100.000 Station 100.000 --- ---
I
663-218
67,1-215
1.955 1.52:3 460 103 f>OO .961 airfoil chord] airfoil chord] 0 1.636 2.054 4.933 1. 5.954 5.7:l5 7.348 7.3?:! 6.515 5.335 2.5:l7 L 1-368 2.828 4.002 5.724 7.004 1.982 8.742 9.317 9.731 9.989 9.828 9.394 8.610 7..>68 {;.345 5.001 3.606 2.230 0 0 1.213 1.867 2.577 3.557 4.321 4.947 7.825 8.185 8.430 8.570 8. 8.516 8.302 7.935 :l.999 0 ordinates given 10.093 10.045 Ordinate Ordinate --- --- radius through surface surface radiu" of and 128 95:1 047 NACA E.
.389 .628 .402 .642
NACA
0 1.115 2.346 4.827 7.320 1. 2.361 4.848 7.344 Upper 9.818 Upper 0 9.845 14.825 19.M1 24.863 29.887 34.914 39.942 44.971 50.000 55.028 60.054 65.075 70.089 75.095 80.093 90.060 9.>.030 . 14.854 19.869 24.887 29.908 34.930 :39. 44.976 50.000 55.024 f~l. 65.068 70.086 75.098 80.100 85.092 90.071 95.037 85.081 1,. Slope of radius through L. K radius: Rlope Rtation Station 100.000 100.000 ---- --- [Stations an(l ordinates given [Stations of 597 sal 2m -- 0 0 0 II 0.084 -.646 -.400 percent percent of -1.952 -4.513 -7.848 -8.323 -1.470 -2.128 -3.948 -4.805 -li.ti9a -7.578 -8. -9.123 -9.336 -9.091 -6.6.17 -5.355 -1.406 -1.323 -1.571 -2.li4fi -a.liYO -5.210 -1i.3a3 -7.188 -8.346 -8.701 -8.918 -8.998 -8.942 -8.733 -7.580 -Ii. -5.451 -4.206 -2.934 -1.714 -1.729 -2.854 -5. -8.7Ii5 -9.405 -9.:331 -8.621 -7.763 -3.999 -2.(150 Ordinate Ordinat<' ---- ---- in K: > HiO L.
.5 .7ii .. .628 .890 1.25 1.405 0 2 5.0 7.5 Lower surface 0 2.li77 5.200 7.709 Lower surface 15 10.212 15.20:1 59. 74.891 given in 10 20 2.> 30 35 40 45 50 55 60 fi.1) 70 75 80 85 90 9.> 20.185 25. 30.131 35.100 40.01i7 45.0:3:! 50.0110 54.968 64.913 69.897 79.894 84.908 89.933 94.966 Station Station 100.000 --- ---- ----.
663-018 664-221
1.9.55 2.550 :3:1a 701 fi97 714 763 581 f)yO 221i 595 .646 airfoil chord] airfoil chord] 0 1.323 1.571 1.952 2.li4Ii a. 4.513 5.210 Ii. 7.M8 8. 8.323 H. 5.451 4.201i 2.934 1. 0 0 1. 570 2.342 4.580 5.ti53 6.51i5 8.039 9.170 8. 7.145 5.091 2.440 1.032 0 7.188 8.346 8.918 8. 998 8.942 8.733 7.580 ordinates given 3. 9.823 3.99(; ordinat~s 10.047 10.709 11.183 11.478 11. H.lm 11.281 HI.
• 1.8ti9 Ordinate Ordinaw _._- ---- surface surfa~-e radius: and radius: of radius through and .5 .7S 095 g.
NACA E. NACA
1.25 0 2.5 .>.0 7.5 .372 .610 • 10 15 liO 20 25 30 35 40 4.> 50 5, Ii.> 70 75 80 85 90 9.> () I. 2.323 4.S00 7.291 9.788 Upper Up""r 100 55.032 75.109 9.1.034 14.797 19.815 24.MO 29.869 34.900 39.933 44.967 50.000 60.063 55.087 70.103 80.106 85.092 90.067 L. Rlopu L.
Rtation Station 100.000 ._------ ----- ------- --- Stations {Stations i SUMMARY OF AIRFOIL DATA IV-PREDICTED CRITICAL MACH NUMBERS Page Page Critical Mach number chart___ _ __ _ __ __ _ __ _ __ _ _ __ __ __ _ __ _ __ 114 Variation of critical Mach number with low-spced section lift Variation of critical Mach number with low-speed section lift coefficient-Con tin ued coefficient: For sever.al N ACA 65-series symmetrical airfoil sections of various thicknesses .. _______________________ __ _ __ _ __ _ 122 For the NACA 0006,0009, and 0012 airfoil sections_ _____ 115 For several N ACA 14-series airfoil sections of various For several N ACA 65-series airfoil sections with a thickness thicknesses ________________________________ .. __ __ _ __ 115 ratio of 0.18 and cambered for various design lift coefficients ______________________________________ ~- 123 For several N ACA 24-series airfoil sections of various thicknesses_ __ ___ _ __________ __ ______ _ __ _ __ __ _ _ ____ _ 116 For several N ACA 65-series airfoil sections of various thick- For several N ACA 44-series airfoil sections of various nesses, cambered for a design lift coefficient of 0.2__ __ _ _ 123 thicknesses __________________ . _ _ __ _ __ _ __ __ _ __ __ __ __ 116 For several N ACA 65-series airfoil sections of various thick- For several N ACA 230-series airfoil sections of various nesses, cambered for a design lift coefficient of 0.4_ _ _ __ _ 124 thicknesses________________________________________ 117 For several N ACA 65-series airfoil sections with mean line For several N ACA 63-series airfoil sections of various of the type a=0.5 and cambered for a design lift coeffi- cient of 0.4 _____________________ .. ___ '. ~ ____ __ __ _ __ _ _ 124 thicknesses, cambered for various design lift coefficients_ _ 117 For several N ACA 63-series symmetrical a~rfoil sections of For two NACA 65-series airfoil sections of different thick- various thicknesses .. ___ __ __ __ __ _ _ __ __ _ __ __ _ __ __ __ _ __ 118 nesses, cambered for a design lift coefficient of 0.6_ _ __ _ _ 125 . For several N ACA 63-series airfoil sections of various thick- For two N ACA 65-series airfoil sections with mean line of nesses, cambered for a design lift coefficient of 0.2_ _ _ _ __ 118 the type a=0.5, with different thicknesses, and c8.!llbered for a design lift coefficient of 0.6 _ _ ___________ _ __ _ __ _ _ 125 For several NACA 63-series airfoil sections of various thick- nesses, cambered for a design lift coefficient of 0.4 ____ . _ 119 For several N ACA 66-series symmetrical airfoil sections of various thicknesses _____ " __________________ _ _ __ _ _ __ _ 126 For two N ACA 63-series airfoil sections of different thick- nesses, cambered for a design lift coefficient of 0.6_ __ _ __ 119, For several N ACA 66-series airfoil sections of various thick- nesses, cambered for a design lift coefficient of 0.2_ .. __ . _ _ 126 For several N ACA 64-series symmetrical airfoil sections of For two N ACA 66-series airfoil sections of different thick- various thicknesses_ ___ __ __ _ __ __ _ __ _ __ _ __ __ __ _ __ __ _ _ 120 nesses, cambered for a design lift coefficient of 0.4 ___ .. _ _ 127 For several N ACA 64-series airfoil sections of various thick- For several N ACA 66-series airfoil sections with a thickness nesses, cambered for a design lift coefficient of O. L __ " __ 120 ratio of 0.16 and cambered for various design lift For several N ACA 64-series airfoil sections of various thick- coefficients_ _ __ ___________________________ _ __ __ _ __ _ 127 nesses, cambered for a design lift coefficient of 0.2 __ .. ___ 121 For several NACA 6-seri,cs airfoil sections with different For several NACA 64-series airfoil sections of various thick- positions of minimum pressure and various thicknesses, nesses, cambered for a design lift coefficient of 0.4_ __ __ _ 121 cambered for various design lift coefficients____________ 128 For two N ACA 7-series airfoil sections with a thickness ratio For two N ACA 64-series airfoil sections of different thick- nesses, cambered for a design lift coefficient of 0.6. __ . _ _ 122 of 0.15 and cambered for different design lift coefficients_ _ 128 REPORT NO. 824-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 1.0
l
I
1\
\ J
I
\
.9
1\
\
1\
.8
\
I\·Curve calculated from equations (8) and (62) of' ref'erence /9 \
'\
.7 I'\.
~
'"
.6 f--
b ""'-
~
'"
l'---
~
"
...........
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.........
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............
r-
r:-:.:..::
:--
~
-
--
.3 .2 ./
-
_0- - o 1.4 1.8 2.2 2.6 3.0 3.4 3.8 1.0 Low-speed pressure caefflclenf, S Critical Mach nnmber chart.
S t::::I
rfl d ~ ~ ~ o I:rj >- .... r,;J t" ~ >-
I-' I-' Cl 1.0 ~
,i'
....
I"
" ,,~ .8 ...
........
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1408 1410 1412 ......
I
,J'...,.
"- 1', c, NACA NACA NACA .6 ,
"-
,
1'-'
--
...
,',
1'\
\; ~ .4 coefficient,
-- ---
~ lift _.
'-
~ ,,,~ ...
.2 section ...
of various thicknesses.
...
f'...
I'., low-speed section lift coefficient for several N ACA 14·series airfoil sections -, ........
II' I
/, with V~,
o
Low-speed
j
A number
lV
'
....
II
, '/ , -.2 / ....
V
critical Mach
/'
of ./
/V
""V
./ ""
o ./ Variation .7 .2 -.4 !..O .8 .3 .9 '- E; ( IJ
-;'.6 ~" -6.5 ~ (; :;.::. ;S.4
.... 1.0 -.
I'--- t', .... 0009, I, "---t--.
....
.........
0006, I .
.8
__
....
"
r........ i
......
....
NACA 0006 0012 - ... ,
'"
the ,
, 1
for
'" ~
c NACA NACA NACA ... .6' ...
, I'-...
, , ,
, coefficient
'"
" , , lift
, '\ -
, i'.
-------- --- \ .4 coefficient, ,
~"" ,~
"
'
---- liTt
~ 1\
'
\ airfoil sections.
. --
1\
\
-, '-' ~~ .2
\ secfion
and '- "~ with low-speed section -,--
,\
L...._ "~ - "- ~-- ,\
1\
... number ~
'\
, o ....
Low-speed --- ,/
V
.L
t-- 1/ V
;./ ll, ._-
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II
V
I I /, /
V
Variation of critical Mach / ~
I
....
/
o
.I .7 .6' .3 .2 -.4 .9 .8 .5 1.0 I..
l" § I:: g b V
i 2l .!:: ;SA
:;:: ".,~ ,,,~ "d Z 00 t.:> ~ ~ > >-3 ,... o Z > t" ~ ~ o o ~ ~ ,... "'3 t;5 t;:J g > ga o Z :>- o "'3 ,... o Ul
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various thicknesses.
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speed
I 1/ I 1 1 I 1 I
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1-+-+-+111-+-+-+1 I I
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+-+-+-11
critical Mach number with low-speed sectionliIt coefficient for several NACA «-series airfoil sections /' of
I I
-I;:;:r,,'
-
1 I 1 I
I I
V -,4
'Il--t--t-II 5 Variation
4 01
.71 _ .~_~~ .31
.81 .21
I.O,r-'--.-r--r-'-.-,--'r-'--.-r--r-'-'--r-r-'--'--.--.-'--.--r--r--r-'--~~ l
:) ()
~ iii I:: ~
~ ~ '6 '6 S
~6 :l: ,
-
1.0
I'--
~ -......
:-::: I ~ ~ .8 series airfoil scctions ~ ..
2412 2415 24/8
2421 2424 -
~- CA
-
~
NACA NAC'<) ,NACA NACA NAC4 "
h:'-
c,
- .6
- "
~ - several NA "- - - '-....
I"'-.
- - l'- r- t--
for -.. -, ,~- - t- - -------- - ----- , ._-
i'- t--
coe/tlc/ent, .4 '- -..,
r--- b I'---
/Ift -.... -....
........ "
- ---.. --
t---... 1-.
, " various thicknesses.
r-- r--...
I'--- .2 section of
k:-
spood section lift coefficient
.-
V 100V /' :/''' r..-:: ....
:
1£ V
o
Low-speed ~ ~ , ~
P'
~ / ~ -;2 L ~ critical Mach numbcr with , of , "/ ~ ,/ ;/ 1
I
-:4 Variation
.4 .3 .! o
.9 .B .2 .7 1.0 '- \:- <u G
(
~.6 E ~ 15 G
--t.5 ~
~ >1 U1 c:j ~ :>- ~ o "'l :>- ...... ~ "'1 o ...... t"' t;j ~ ~
--:r
I /.0
-"
['\
-, -.
~-
-'"
-'I:::: r--.
-8
:----
a=O.3
:--
63-series airfoil sections
r:-
--
-.: 4-420, 4-420 c z
- NACA
-'-
_6
63(420)-517 83, 63(4eO)-422 63, --
r-- r-
-
--
NACA NACA ::: t-
--
:-
I-- coef'ficient, .4
-
-
r-- I"- lift .
--
---NACA --.-
-----------NACA --
-- --
-,
r-
t--
--
r-- t-
.2 section - t- I--
--
'-
- t--
--
speed low-speed section lift coefficient for several
-
II"-- t::-
t----
1/ ow-
o with
L
,-
D'
/
:> ,.
number r of various thicknesses, cambered for various design lift coefficients.
V
1/
/ Mach
V
.2 , • / ,/ critical
/
of /~ LV
:"
i--
~ V- ~ r-
I :I ~ 1 ? ,
o -:4
.9 .B .7 .3 .2 1.0 l. Variation
<!l § c: <:J v
1.- .Q 'S.5 ~ -- " .... &.4
~.8 /.0 ~f.,; t--...
airfoil sections ~ L-- .
.8
-
~
~
230-series 23012 23015 Z3021 23024 23018 ~ ~, I-- z ":;;:
.. _. C
NAOA .8 __ NACA NACA NACA NACA NACA I:---.. 1-- I--- .:;:: r--::::: '- - -- t---. ~ ~
k-
---------- --- -. ----- ' coefficient, . .4 -:::::: 1"-..
lift
::..::.: -
r.:-:
-:;-
1"-..
-f=:::;: -I---
~ --
....... " .2
- section
-
t'--- )::::-
of various thicknesses.
--
......
I""'::
--
i"-- 1"-- I:--- low-speed section lift coefficient for several
--;..
17' ;.::
o
with Low-speed , .)
hi , 1/', 1/ number
1 j.,'- ~ V
.. , , l~ 1/ Mach ~Z , 1:;7 )...- , / ..
, ,'" , , /V , ./ f-'
1......-- V V ( V
I . ,
.7 o
.9 .8 .3 .C -.4 1.0 l.
Variafion of critical 1.- v v ~.6 .£::.5 " ...
~ Ii ~.4
~ t"1 ~ Z ~ t:::)
I-' ..- 00 "d o ? 00 I>:) ,;:. z ..... o Z ;.- t" ;.- -<! ..... w o ~ o o ~ ~ .... ~ t"1 t'l ""-l o ~ ;.- t'l l:tl o Z ~ 1-3 .... o w 1.0
"
'.
I"-, I" ~ " f-
"r---., ,
"- ........ --
..........
"-
r' i' 1::--;'" f-
, - --
i'. , l"- "-
.8 , , ~, -218 -221 t 3 .. 63·series airfoil sections
'"
10 i' I'"
83.-215 63-206 63-20g 63-210 68.-212 (}3 63 '~ ,
""
i'
'0, l
NACA , C .~
I" 1"'-
NAC4 NACA AlACA N.l.CA NACA NACA NACA
" .6
'"
- ~ f~ 1"'-
- - "- '" coefficient of 0.2.
'~
~r- 1-- -- ----.
-- , - \ .. \~ lift - \
~\ \
, -- ---- - -
--- - '"
~ ~ L~ I\.. coeFFicient, .4 design .....
._- coefficient for several
l' 1'\ ~ 1\·
lif't
~ for a
- lift
l'~ 1\
,~ --"'"
l\ Kl
... 't-,
·
"
r- r-
secfior?
.2 cambered , .........
-
"- I r-
,
/
-- " "
low·speed section ,..
!
" - -., ~ thicknesses,
- J with
1,' ~t f.
o
--- Low-speed I
I
I~ 1/1-
, various , /
j number
I
I~ V' !/
of ...
' j
" / ·7 / r---- , -- Mach ... / I
V r7 I' , 1/
-:2 ---
/ /
/..- / /, .- --- - ,~ ---
V 1/ ... :/
/ /- , ,~ , ".., /
t--- t--- f- r----- '-----
V
~ ...
,/
- f/ / ~ r-- ~ - ~
./ .
-.4 Vuriation of critical
o
.9 .7 .2 .1 .8 .3 1.0 ...
~ -
~
~.6
I I I
j
'I 1.0 I I I I I i' I
~ U r'N l I
_
"
\ I I I +-+--+--l syrnllwt.ricul '.
i I - 1 I I
H+k .8
......
'1, I I I I
N
-021 G:l-s"I-i,'s '1- ....
"-J. , ~
I I till to.. I - I I I
'1-..... UA 63.-018 63-008 83-009 63.-012 63.-0/5 63 '
" "
I I I I I N A
'
-:::i -, c,
I I . , I I I !
"J"., .8 NACA NACA NACA NACA NACA ' NACA "..
""
I I ~ I I I sHeml
1"-
, I I I I
"
, I I I I
I I
"
'\
,-"f:--.,. '
I ~' I I I
f
-4 coef'f'icient.
'\
I ~C'-. I
I I I ' ... 11ft I I I I I
11111111
1\'"
,,.....
\' \
I I I I I of various thicknesses.
\. \
I I I
I-'--,-T-r I I I I I I !
i-'\T' .2. section
I
I I 'K I\"... 1 I
sections
~t.:I- low·speed section Iirt cO('fficient for
1 I \ I I airfoil I I I I with
{:j
........
1 I , I I
...
-I Low-speed , I I I II V numbel'
J:=1-
1 I I I I
,tlU,ftl
I I 1- I I Mach I 1 I I \ -I------+-+_+_
I
-.2 1 I I I
I I I V:P
,ktj.-~r-
U4ij-r-t~ti
_
I , I I I 1/ I I I
I
,
'V
I I I ~- Ii 1 I I
.
/ /1.
I H' I
Variation of critical -.4 31 o'!
..?I .91 .81 • ./
.7fT
1.0,
~ c: t)
JeLl;' ~. l: § ~ .....
.r:..5 ~ G.4
U1 d ~ ~ :>- ::xl ~ C "'1 :>- .... ::xl "'1 C .... t" t:I ~ :>- I-' I-' roO I I I , 1 1 ~ /.0 I I I 1 1 I
U
I I I Iii 1 1 1
ill
1 I r 1 1 I
[I II!
1 1 I 1 I 1 I 1 1 1 +1--+--+-----1-----1 .8 1 I I I I I 1 1 1 -615 I 1 I I 1 1 I -6'/8 0.0.
1 I
I 1111 I I I 1 I I I I I
63. of c.
1 I 1 I I i 1
I
.B 1 I I
1 ! I 1 I II
NACA H-I
1 I I i i I 1
[]--r
---j-----l---+--I 1 1 I 1 1 I I 1
I \
~-~
1 1 I I 1 l-----l-+ 1 ------~- ---NACA coefficient, .4 1 I 1 I 1
1=1- lift
1 1 I I 1
-,-r-1-r
I 1 I I 1
1 1 1 1 1 I 1 I I I
+
,I .2 sect,on I I I I I 1 \.
1 I I 1---+-+-+--+----- 1 I
11111111I I
wit.h low-speed section lift coefficient for two N ACA 63-s.'rics airfoil sections I I I I
I I I
Low-speed 1 I I I ]-- numher I
1 1 1 1 1 I I II different thicknesses, cambered for a design lift coefficient
++-+-+-+-+-+----1-+
-+-----t----t-t----+ of /' I 1 I 1 1 I
I I I I I rUl=W
Mach
1 ~ttn~t
v 1 1 1 I
-.E?
\ I 1 1 1
I I I I
--
\
I 1 1 1 I
\ I
1 1 1 1 1 1 1 I t-i 1 r----t--- I Variation of critical
{ill! ottl -;4
.91 .I
.8\ .11 .31
1-0,--,--,
I-tHflll±±
I '-0 ~ I
"
sections
-+-
---+--
---f--
J= I I tjj ~
"t. -~ --·--t-- .-
--+-+--
,pt
1- I '.1
'i
,I -----i
I Kt ~ , .8 '\ I
j11
63-series airfoil '~i"_ '\ -1",- .....
I 1 0.4.
", _
1\-1 f-+---
NACA of
63,-4/2-rrm - Fff4 j j
+-t++-+-+-+-
r -~-I--H--- --t---+-+----t------t
rs: c
i ftEHIEB
-
.8 -t-
-j-t--i---1-r--t-i-i
r-- I
I I I I I I I I I ,- I, 'I
1.
NACA __
-I--I---t--+-+--+--+
- _
--i
,-,-- --t-
..
coefficient
I""'"--!... r- I (or s('veral
c.!--L"\'
+--
-
__
lift
I-I I
I!
_1 ~t=-~+
-~
_ tHj
~~
-
I J ---+---t---t--+
I
.4 coefficient, ...
--_:1 coe-fficicnt
-
I--
t a design
f~~-~~fg~:$huu
-~I lift
_ lift
----
~j:::: for I~I ~--- 1-----1--1--: "- -
I I I I f---
" --...
- sl!etion
. t : I \-
t
,-----, ,2
_ :s-ection
(·allll)pI'~.1 I~r-_~- I I ' ,
I "+ 11., I
liti
--
__ -!
i r ':::::-j.", i , i I
t=~ low-Spt'pel
__ 1 t
,Lrtt:L1 I I I I I I I I I I I I I --1--'----
'J7~r:-:-- __
/ --I-~--Ltf-i- --- --
I t, ! I:
:'
--+-+-+-+--+---I--I-i-
~ with
/
-- _ --- ---!--+-\-!----I-----I--t-- :-.c..-J -,L-~
-+-tlttlJ-HJ Low-speed
e-_ __
-
"
l-(....'_J ,-
-I / IllUUhf'f various t.hick,,,",sps.
--1---1-----1---'. -1---- -
~
V 1/ of
,'II
- ,;' _ _ - -
I I I I I 1/ j-t-tjj
IVI:wh -- ,I , -- /,' /
II
,;'" t- t
- -.2
--~:=-
-
- --
II I;, 1/ 1./ V r----
critical ~' .' ---
:= ~ of
./, ,-'
I 1/ V
--
" 7,'
-
,-' --
V
I I
II I I I I I I
./ -
-" f---- ~
I-r----- Variation
"03£J11111[111 7\ -:4
Jt1---t
.al .S-V-,/ .I
.f!
1.0,,----,-----.-----.---,----,--, r.. " ~
l " ::i c: &
~ ~ -t::.5 :. '- :<.:::' (}.4 !:tI trI "d o !:tI >-:l 2: 00 "" .... :.- >-:l ...... o ~ :.- t" tl ~ ~ ......
I-' t-..:I o o 2: > ;S U1 o !:tI ><1 Q o >-:l >-:l trI trI I'%j o !:tI ~ !:tI o Z :.- d >-:l ..... Q U1 , 1.0
'"
-
............
"- -....
...
1', ~
airfoil sections
"
'"
.8 I ,
'"
I'.....
54-series h, 64-110 64,-112 64-108
"
, 1\..: 0.1.
"
of
" 6
1', c, NACA NACA NACA , ..
~ \ ,
'"
- -
K 1"-
------~- - ,
\
\ .4 coerf/clenf, \
"
r, \\
lift a design lift coefficient
\\ 1\
for
.\ \
\,
1\
\ \\\
.2
-
section
'\'
.......
-
-~
r--
speed ,
II'
o
I Low-
,f-/. II
I:
'j
II various thicknesses, cambered
I I
of '/,
V
I
/
I
v' V
/ -.2
II 17
critical Mach number with low-speed section lift coefficient for several NACA /
/
of
/ J
/ V / ~
V V
~ 3 r 1
o ~4
Variation .9 .8 .3 .2 .1 .7 1.0 I- ~~ <b ~ u
~ .;::. 6.4
~.6 ~ -!:;:.5 ~ .\.l
/.0 ....
"- """1 ,
"
1', t---..
......................
"
I' f'..
l'-- symmetrical
-,1'-- "
1 ~ .8 "- " ,
-018 -021 I"--
-0/2 I"---
64-series , "- i'.
64 64 1'-.. 1, 64-006 64-009 64, 64.-0/5 , 'r-.
.....
I"
,
't, -'I
"- 1
1 r--.. c
1', NACA NACA NACA NACA ~ NACA NACA "
"
" several NACA
I" 1"- I'....
T
--- ,
"
for " " , -
,. --
,~ ----
"
~ '" 1"-
---- ---- --- ,
'"
" ~ 1', .4 coefficient, , , \.. '\ \ I"~ 11ft
"
~ ~ 1'\
\
\
\
f-\
..}~ .
\\' 1'", 1\
\ .2 sedion ~\ ,
\
\1\
rt-
~ airfoil sections of various thicknesses.
- --
t--_ 1'\\ 1-\, , -I-- "
r-- -, I'-
o
~ ./ Low-speed II f- I-" l-
- -I--
'-I
I- V J.!J..-
, -
H [[.
j -j !/
~ 17
--.2
/
critical Mach number with low-speed section lift coefficient j-1" p) Ir of
~ II
I J J ,..,V-
-
V":" V
/ / I
V'"
V Variation
1 , Y 9 5 -.4
o
.9 .8 .7 .3 .2 ,I 1,0 I..
I..~ 'b ~
i::
'::l,.6 ~ ii· ~ ~ .I.l 6.4
m q '7 .... '7 ..., ;.. ::tI ..-: o "'1 :>- H ::0 "'1 o H t"" ~ :-- ..., ;; i-' ~ I-- , 1.0 ~ ~ S('et.iolls " -, I'--
'
1"--
i'-..
'
t-..., .8 -418 -42/ -412 l'.. t->--. I\.
3 , 4 H4-series a.irfoil
-
"
~ .64, 64.-4/5 64 64 A(~A ~ ~ 0...1.
, of
1-\
NACA NACA NACA NACA .6 C, -1"-
-
[\ r-
s(,veral N
-- --
lor (,()PtJieil'llt.
-
t-- lift.
---.---
--- ---- -
-
r-- t--
.4 eoert/cienl,
'- c(J('tliCi(,lli
""- r-.
-
lift.
-r-. flft
for a de.si/!Il
-
t- s(wliou _ t--.
t-
-
-1-- sect/on (~;unhpJ'ed t-
-
il
low-SIJ('('d
jr--
V
with ....
I
t... thi('ktl<'ssl's, ~
o
Low-speed j...,< V
I
yarions
V
A/I
of
I
L/ V
,
.L/' /
V
-.2
V 1/ V V
eriti('al :'\I,wh lIumhc'r ~, or ,/ t,.- V / ".
V /'
V V
I j Variation -=;4 .9 .8 .7 /.0 L J
~- l " ()
~ -- .v fo.4
i· E ~.5 ..:: , /.0
i' r" r:-: ~
....
.... 'r-',
"" s~cti(Jns
, ~I'-..
I"
....
....
i'-- i'--- :-- ...........
airfofl "~
"
"
" I r....
.8 ,
"
'"
-218 -221
'"
f".: lt4-s('ries ,
,,' ,,1":'
"
""
I I I t'--..
64,-212 64.-215 64 64 f':, {J4-206 M-208 64-209 64-210 ,
"
"- .; ITT l' i'-..
NACA of 0.2.
"\ " "
~I--
I C
NACA NACA NACA NACA NACA NACA NACA NACA .6 .
,
,\0-
~\ :-- ~~
- f\
s('\'('f'<tl \ '\
I I I I I I
,'\ I"
cOl'fTi<.'il'llt for
-- \:,\
\
, \
~
I I r~
lift.
\ --- - ------- ---- -.--- --- ----- , \
1\
.4 coefl/cient, ,,~~: \ .\\ t~ ~~ //1'1
~ S
~~ ~ t- e--+_.
'\~ --~ s('ction lift cocllicieut -r--.
- -
..........
,-- .. .2 section ....
I- I:::'" '- - r-- -
I--
-
r --
'- -- l- 1'-.
low-speed ft ~~r
tt- t-
I /Irt with 'fi
rF ~-L II
o
/
Low-speed I ",~
1'/ W~ if
I
J
/, various thicknesses, cumbered for a dl'sign I I.~ ;'/ 7 / I of .~ I
~ II ';'l
II
Mach Ilumber , J
/
/ / ~.
~- I -:2
/
/ h~ I~ V eritk-al "
/ 'I'
of
'7
v
, /'. / .
1./ V V-
/'.L
'/ '/ ~
V V V
9 -.4 Variation 3 o .9 .8 .7 . .2 /.0 I.
,,- q, § ~ ~
~.6 ~ ~
-<::> ",.5 .S! :~ 0·4
> >-3 Z ~ ~ .... >-<1 ~ ~ .... >-3 >-3 l:':l l:':l I'<j @ ~ l:>:l o Z ~ ~ &l
~ 'tI o ~ Z P 00 ~ ... ~ ...... o @ 8
t>:) t>:) i-'
-
/.0
-
.......
-......
i'----.
f'- , _t'---.
-
r---. t-....
, i I
-
., , ~ I ]'--..
1'-
-
f"., t--.,}... , ! i
.8 ,
-
......
-018 -02/ I r----.", ~ ...
, r" t-...
I" 65 65 65-006 6.5,-012 65.-015 , , ......
-'
"
l'-..
, , ...
, r-... c,
I"--
.0 NAC465-001} NACA NAC.4 NACA NACA ·NACA , , ,,, I,
"
, "
l'. I
, "
r" f'
, -------. --- ---- _.-.- --- 'I"-- \
" "-
.4 \ coerr,cleni; coefficient for several NACA 65-seriessymmetricai ,
--~
\
1-' " ,"-
lift
\
'\ - Ilrt \ I~ I', 1- various thicknesses.
\
\ \
-\ \ \
1\ r- .of
- ,
\
'\
, 1\ .2 section
\
..1 \
k I':-. 1\
~ \ ~ I
'
-f\-
airfoil sections
1\ t-\. t-
with low-speed section
\
-
- .....
t-
t---. o
- t'
- Low-speed
.....
~ I
II
number
-
-I- I I .
r.J.
,'11
7-
~ I I
J
Ill-
J~ II/
-.2 "7 '.L .
/ /
, , II
_
.-
,
/
, ~~.
l- ~ V
I .
~ D 1/ Variation of critical Mach
.V
o -:4
6 5 .3 .2 .1 .91 .8 .7 1.0 c..
!: qj ~
'i. ~. 15 S.4
~ r .G ;::.
1.0 -
"
~ '-----
- -~
r-..
-
- I-- .9
--
--
-
{)4;:-615 64$-618 - 0.6.
-
--
t-- c, _6 NACA NACA
-
- t-
-- -~
i"- -~
-
-------- I .4 coef':£'icient,
-
a design lift coefficient of
--
- lil"t for
I-
--
.2 sectIon 1"
V-
low-speed section lift coefficient for two N ACA 64-series airfoil sections IF- /j
V
with ,
1/
o
.-
Low-speed
/
number
'/
, of different thicknesses, cambered ,
V
V
, -.2 .
, V .- / .
.
,/
.-
Variation of critical Mach -.4
o
_9 .2 ./ .8 .7 .3 ...
1.0 t~ ~ ~.6 ~ .;:
~.5 ~ ~ ,~ ~.4
~
~ ~ !? ..... .....
m e C :; o "'1 :.. ..... ):<J "'1 o t"' t:j :.. ~
'-' tv ~ i i 1.0 ~ r- ["-". r-..... r-.... ~ , ..
" "I'-- " -- ~
"
~ ........
.~
r- 1"- r b-
- ,
f'- 1"'- r--....'
.IJ "
> '"
"'- i'
~ -221 -218 -0
'"
I" l'-.
3 4 . ' , \ ' 651-2/2 65 , 0.2.
65-Z06 oS-cOB 65.-ZI5 65 I\. I" "'-
65-210 'I' ~, ~ ]". c.
' \ .6 , ~ -- NACA l- 1'\ NACA NACA NACA NACA NACA NACA \ ' - , I . .....
r-- - I \ \ --- -~ '\
~
G.
- - .. -- \
~
r-- 1\ '\, ~
\ coeTf'icl{;mt, I .4
• - - - ----- - ',\
---- ---- \
.'\
~ - ~ lirt ~\' _\
l\-
~
r- I\"- "-
.
-
--N --,
-:\
1\
.C section
F=-:: t-- r--
1"'"- "
r
-
j
7"':::" '- 'r-- low-speed section lift coefficient for several N A C A 65-series airfoil sections
'1
II,
It
/.
/, with
: rJ-I-
o
I Low-speed ~, /-1
1//1/ 1/
I .
J
/
/
number various thicknesses, cambered for a design lift coefficient of 1- II,
V
" of , ,',
I
,
f--i
I / Mach /
/ "
, ~ I -;c .L L i.l"
lL V 1/
"
, .L ~
1/' V"
, / / /;
V 1/ 1/
/;, / 1/ /
I , 1/ lL
-:4 Variation of critical
.9 .8 .7 .21 o
.S .I /.0 1..
~ g
§ l:; ~.4
Jl {::.5 ~ ] .~
'i,.6, ! , /.0 "
"
::::::c:
"
'\ :-- 1>--", - "- ~
1-=
'\ i'-.
.B ,f=-:
"
-:.
'"
A CA 65-series airfoil sections
L::"": l'--
65,3-0/8 65,3-418 65,3-618 65,3-BIB v,
r-=-
.6
"
r--r
NACA NAC4 N.4C4 NACA -~
-
-- l'\ ~ c:--,.
-- - - -:;
1,\
.4 coer-f'tclent, -------. - -- ~ lift t- I::::: I ~ / -t:- camhefed [or vl\rious design lift r-Oefficicnts, V .2 seofion , and ~ I-- 0.18
-- /
- ',-
o[ low-speed section lift coefficient for several N - J , , - ~::- ..
/ with
l/
a ,- LOW-speed
'
h('r :
V
I
"
, num
V
"
I-
V ach
"
:vI with a thickness ratio ~ V~ V -:2 I i f--
,.-
" ./
I , of critical
-
, // /1- r- V 7 -;4 .9 .8
.3 .2 o
1.0 .f ~ Variation t::
~ § g
~.6 ~ {::.5 ~ G .'> i:: 6.4
I
>- >- H H >-l >- >- ~ t=l 'tf 0 ~ Z ~ 00 ~ ,... z >-l H 0 Z >- t"' 0 ~ U2 0 ::tI ><1 (") 0 ~ ~ >-l t=l t=l "'J 0 ::tI t=l ::tI 0 Z q I-:l H (") U2 ~ ~ I -l 1.0 ,
I I t'\. I
-l.
J, I
' ~
I 'I"':
~ j
'i]:,..I'
I I _R
'\L
0.4.
"
a=0.5 I
a=Q5 ~:~~+-+-I--l 65·series airfoil sections
+--+,+-+-+-1 +
of " .--tj" , I \ I I
I I I I I I I I 1 !',. I I I I I I I I
-4/2, NACA I I fI
~~:1:$: 654-421
-f--'--t--?
1 t-_ I
1--.\1 several
1\ _ -_~
__
for
1 I I I I I I I I I I
NACA
~:&~
--
I 1--....1 i
r-
I I
I
A
'""-I 1
-
I . 'I
om:=:
--- JI=l=t=F~i=i=~~=1=t~
----NAC'A
1 i I
-I
-+
-4- I~-_ -
__
, I I
iii
-t and cambered for a design lift coefficient : I ~~~
+
-+.
--+-1--+-1-1-1
__ I I I
:j--j !:-- I I 1 I I
+-1
jl:tl:fl=tl T+
a=0.5 : I :
II: r !
low-speed section lift coeffiCient
-l-
r~1-~+-+ I with /
--~
the type
I
+
of I /..L1
i V 1
J
+1--+-1-11-+-1 il=tI=tI~1
-:-r-p-
A-f---
I I 1---1-f-1
V
+-1
~.
"
/
, I ~~-L~~~-L~
1/ Iii
JL
I 71,(~, __ 1,..1 '. I
'1
with mean line L ___ i , 1
r.{ i 1 I
H-+-+-tt±¥i-fttttfUl
~~tJIF _~ JlL': X
critical Mach number
,'r~A_
-) I 1 I of
I 1 I L-~~
+1--+-1-11-+-1 L
~I __ 1 i 1 t t i -I' n' ~ ! ,.-,----j I !L'! i ( ,
I
~
1 I' -t---
)' ~
1 :-;V I
iii
I O~!
Variation
.gl-I+-J .3R-
1.0,....,....
~
.
c:
~ § ~
.Q 'S.5V ~ o ~ 6·4r.,~~4-+-+-+-~~
)".8 1 I , 1.0
"-
"-
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1 .', r- , '\~.
'\ ,
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11;-- : I
.8 \!
,
, p:;: 1\.' I
65·scries airfoil sections , , , : I I~'
I I
\
-+-. \,1
0.4.
-421 1\1 NACA -415 -418 z 3 4 of
-
l\i~= t ~ 65-410 65 65(47.1)-420 65 651-4/2 85 .6 c \
i,\
several \ I 1 I I I lOr NACA NACA NACA NACA NACA NACA
--1,\
-.
r.:-~:~f\- I I
--
r-
-- t- I"'::' I I
.4 coefficient,
-":"'-1-
;-- !-- -------- --- ---- -----
--
lift I 1
I t-- kL-" I
--)..-
--
'-":: Ir- t- .2 ,- , section ,
ff:::: R
A"f
I -,
1 I7 1 , I low-speed section lift coefficient
T
,rl
J -
' "
. i
I/f with / - ,
II
o , , LQw-speed
/
,
Ii: ! 1
I
/ //
-- /
, various thicknesses, cambered for a design lift coefficient ~ , of
" /
, ! , 1/ IV
"
/ /1 V 1 )/ , ~2
~ I/! / i , i
critical Mach number , , of ,
/
I \,(' /' , jp l/,
I I i 1 "d: V 1/
V
Variation
~4
.1
.9 .a .71 .3 .2
1.0 l
1..- ~ c: tJ
~ 4;..5 ~
'::e,'J.6' ~ - 6.4
U2 C1 ~ ~ ;.. ~ ~ o "1 ;.. >-< ~ "1 o >-< t"' t1 ;.. t-'l ;..
..... t..:> ~
\ /.0 -,
1,\1
!
1 U.6.
, I I I of
1\
h.1 -I~- airfoil sections
I I I--
.8
I 'I-T-----'- t--
0=0.5
--
65-seril'.3
I I , , I
-
I
-615.
-6/8,0=0.5
2 -
a design lift coefficient NACA t--+-
j
i:[ .6 65 for
--
I _ t-- two
-
for
[IT-Il I r--
NAC4 NACA
I I
cambered
-
--- and .4 coefficient coefficient,
-
r-
II lift
-
----- t--
I I I /Ift
- -
I---t-- .2
..I --
section different thicknesses,
-
low-speed section
V
,-
with
1/
with /
/
o / / --/- Low-speed number / -- / type a=O.5, / / Mach the
/ /'
-:e 1/
V
/ mean line of ,.- / /V I with /
V- Variation of critical
, .7 .I ~4 .9 .8 .3 .2 /.0,
t ~ B
~.6 ~ -<:.,'5 ~ .1.> i:: 't 6.
, I ! I ! 1
'. l l l l
/.0 'I
\
-
'\
:s;
-
.8 -r--
r I
65-series airfoil sections -6/5 -6/8 0.6.,
-
2 3
--
I I
of 65 NACA
-
-
c, .0
-
:::::: 1
T T
NACA =
NACA coefficient
, I
lift
I)
T
--
-
r
.4 coefficienf,
-
--------
I I I
- -..;-...,:
-
- Idf
-
/
.2
- ,-
II I sect,on
/.
I low-speed section lift coefficient for two
/ 1- -
/
V with
/
/
I----
o
/ --
/
Low-speed numher /
1/ I i I I
T
]
/ of diff"rent thicknessr$, cambered [or a design ~
V
-
..- V -.2 / ~/ critical Ma('h of
__
i i I I I
!
V
-1'''-
~I
/
V, I i I
/ L Yadation
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9 o
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% g
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gJ ~ ~ Z ? 00 to:) II>- ~ ::"3 ..... o Z >- t" >- tj ~ U2 ~ C o ~ ~ ~ trJ trJ ""J o ~ ~ ~ o ~ § C U2
..... ~
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"
t-.... t--, t-...
..
"
' ' ',- ..... .... .........
' ,
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: , , , airfoil sections
....
, ......
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"-
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66-scries '
" "-
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-212 -2/8 , i'
t , 0.2.
, 'K
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'\ 66.
66-210 66-206 66-209 66 66,-215 66, '\
"
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c,
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r, I'\.
NACA NACA NACA NACA NACA NACA NACA \ \ \~\ '\ - "
1\ ...6
- - \
\
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--- :"-r, -
- - \
- --- .\
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:\
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\
coefficient for several NACA "\~ --- ----- - ---- t
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a design lift coefficient of .....
\
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.1
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-
section
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I
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o
I
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r- Low-speed
I /
.
I/ L.
, !//
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/' of various thicknesses, cambered I J
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1/
/
I
/ .-
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, , .'
/ /
,
I /'
L L ,
/ -/ ./ , --
Variation of critiooi Mach number wjth low-speed sootion -.4
o
.1 .3 .2 .9 .8 .7 1.0 l..
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"
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58-006 66-009 66 66, 66.
6(h , ,
, "-
"-
c, , .6
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"
"
r'-.
NACA NACA NACA MACA NACA II(ACA ,
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11ft cO(!,Tficienf,
"-
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\ \ \ \ various thicknesses.
\
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of \
\
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1\
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r- k-- 1\
I airfoil sections
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flf
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'rf 11-
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I
, -.2
-I
tf"", I , , .
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'/ Iv' Variation of critical Mach number with low,speed section lift coefficient for several NACA 66-seriessymmetrical ) 9 ? r I -.4 .3 .1
.2 o
I.
...
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,
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66-series \
, "
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66(ei5j-016 66~215tz16 66 6~2/~-416 6
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-- t'::'
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lif'f
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-rr-
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cambered for various design lift coefficients.
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;:, _etlon
r0- I , 0.16
-
1- II low-speed section !ift coefficient for several N ACA l
--: f
with I
t=-. V
o
,
"
1-. L"O.w-$p~ed ,':
V
I.-- thickness ratio of "/ I!.
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."
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.2 ./ v Variation of critical Mach nnmber
.9 .21
.s
1.0.
~
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.. /.0
"- I~ "- sections ~
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"
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J 1
0.4.
~ of 66.-415 ~- c,
J
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I 1 t-t1
NACA AlACA
-
r-
I I
-
I
I
I I
A...6 coef"ficlenr, •••••••••
I t--
lift
-
r-
-
r--
-
.2 section
-
-
II
I
o
/
I Low-speed
I V I of dUIerent thicknesses, cambered for a design lift coefficient
V i i , i
/
V
-.2 / V 1/ / ,
lL
Variation of critical Mach number with low-speed section !ift coefficient for two NACA I , ; 1 , , ,
a ,,4
.7 .9 .8 .3 .2 1.0 "-
~ ~
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-
to
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I
I r-.... L-'-----
, 7·series airfoil sections
"
I
T
-- A.CA.
I I "-
747.4315 747.4415
-
c;, I I :-;;.
.6 -'-- two N ....
I I
NACA NAC4 for
I I
-- - t- different design lift coefficients.
I I l-
.4 coerricient, for ----- ~
I
lift t- - cambered F-- .2 section and 0.15 t-.
of
~ ---
, ,
o
/
I
Low-speed I, - I - V /
, -
/
~ - .
,
/
with a thickness ratio -:2 / , , - / - /
.::-
/ f-;;- Variation of critical Mach numher with low-speed section lift coefficient 0 8 6 d -:4 .7 .5 .4 .2 .I /. l
~
~ ~ .~ 15 .~ Q
{ :;:::
-
to
"
1~ c , , , , ' j'--; .
,
r'\
,
"
"', ~ .8 airfoil sections with " ......
:.~ ...
a=O.8 f-,. [\.. f'..
6-series ~ r 'l/ C
"
-~ various design lift coefficients.
NACA \ .6 COl' r-- ........
: ti7.I-2IS 6!J(216)--nS, 0."'0.5 65,3-418, 551-012 ~I--.
\ ['...
\ \
r-
'"
NACA NACA NACA NACA ,
~
coeff'lcient)
- , -t~ .4
\
\
- \ -- lift
~- -
-
1\
\
-- -
1--- ---- - -- ~
f.\
.2
-, '- section various thicknesses, cambered
1\
~ t-~ and
-
I- -I-
t---... It-
I o
with low-speed section lift coefficient for several , , Low-speed
V V
;'
/ ,
/V
, number po /, minimum pressure
/
'/
/ of
II
Mach , -;;?
1/ V '
]7
" -
, 1/
"
V 1/
critical
/
/
,,7 of ,-
f7 17 I -'--
~ 4 '3 ? ,
0 ,8 ,6 ." -.4
() difierent positions i.
I..
\.:' ~ ~ ~ ~ ~ ~.
-E ~~
Variation SUMMARY OF AIRFOIL DATA V-AERODYNAMIC CHARACTERISTICS OF VARIOUS AIRFOIL SECTIONS J:'age Page N ACA 0006 _____ . ____________________________________. ____ 131 NACA 64 -418 ___ . ____ . ___ . ______________ . __ .. __________ 194 NACA 0009 ____ .______________________________ c ________ •• _ 132 NACA 64 .-618 __________________ . __________________ . ___ 195 NACA 1408 ________ .. __ .. __ . ______________________________ 133 N ACA 6~-021. ________________________ .. ______ . ___ _____ 196 NACA 1410_____________________________________________ 134 N ACA 64 221. ____ . ____________________ . ___________ .... _ 197 r NACA 1412 _________ .. ________________ __________________ 135 N ACA 64 42L __________________ .. ___ . ____________ . ____ 198 r N ACA 2412_ ____ __ __ _________ ________ ___________________ 136 NACA 65,3-018..----. __________ .. _____________ . _ ___ __ ___ _ 199 N ACA 2415 _____ . _________ . ______________________________ 137 NACA 65,3-418, a o =0.8_ .. _ .. ___________________ . ___________ 200 N ACA 2418 ___ ~ ___ ______________________________________ 138 NACA 65,3-618..-- _____________________________ ~ __ ______ 201 N ACA 2421._ _____ ______ __ ___ _____________ __ __ __ __ ______ 139 NACA 65,:3-618 with 0.20e sealed plain flap_________________ 202 N ACA 2424__ _____ ____________________ __ ____ ____ __ ______ 140 NACA 65(216)-415, a=0.5________________ ________________ 203 N ACA 4412 _________________________________ c _ _ _ _ _ _ _ _ __ _ 141 NACA 65-006 __________ . __________________ .____________ 204 NACA 4415 ________________________ .. __ ____ ______________ 142 NACA 65-009 _____________ ... ___________ .. ________________ 205 NACA 4418 ______________________ . _____ - ____ .____ __ ___ __ 143 N ACA 65-206_ ________ _______________ ___________________ 206 N ACA 4421. _______________________________________ ____ _ 144 N ACA 65-209_ _____________ __ _________ _ __ . ______ _______ 207 N ACA 4424. _________ ._ __________________________________ 145 N ACA 65-210 ____________ .. _____ ___ _____ ______ __ ____ ______ 208 N ACA 23012 .. ______ ________ __________ ___________________ 146 N ACA 65-410 ____ .. __ . _____ . ______ .. _______________ _______ 209 N ACA 23015 ___________________________ - ________ " . _- - ___ 147 N ACA 65 -012 ___ .. __ '_ __________ __________________ _______ 210 NACA 23018 _________________________________________ .__ 148 NACA 65 -212 _____ . _________________ .~_ ___ _____ ________ 211 NACA 23021. _______ ._ ________ ___ __ ______ ___ __ ___ ___ ___ _ _ 149 NACA 65 -212 with 0.20e split flap (lift and moment characteristics) _ _ __ ____ __ ______________________________ 212 N ACA23024 _____________ - - ___ - ____ - ____ - ___ - _- - ____ - - - __ 150 N ACA 63,4-420--- ______________________ - ___ - _- - ____ - ___ - 151 NACA 65 -212, a=0.6 ______________________ ._____________ 213 NACA 65 -412 _____________________________ .. __________ ._ 214 N ACA 63,4-420 with 0.25e slotted flap NACA 65 -015_ _ _ _________ ______________________________ 215 (a) Configuration _________________ - - - __ - - - .. - - -- - -- --. 152 NACA 65 -215_ _ _ _ ___ ___________ ___ _____ __ ______ ________ 216 (b) Aerodynamic characteristics with hinge location 1_ _ __ 158 N A C A 65 415 ______________ . ____________________ _______ 217 (c) Aerodynamic characteristics with hinge location 2_ _ _ _ 154 r NACA 65 -415, a=0.5 ____ .____________ __________________ 218 NACA 63,4-420, a=0.3_ .. ____________________________ ~____ 155 N ACA 65 -018_ _ _ ____________________________ ____ ____ ___ 219 NACA 63(420)-422 ____________________ . -- _______ - ---- ---- 156 NACA 63(420)-517 __________________________ - _- - - _- - - - - _- 157 NACA 65:,-118 with 0.30ge double slotted flap NACA 63-006 ____________________ . _________ .____________ 158 (a) Configuration ____________ .. ______________________ - 220 NACA 63-009 _________________________________ - - _______ - 159 (b) Aerodynamic characteristics_ _ ____ __ ___ __ ___ _______ 221 NACA 63.:..206 __________ . ___________ .. ____________________ 160 N ACA 653.-218 _________________ " _________ - - ___ - __ - ____ - - 222 N ACA 65:,-418 ___ .. ___ .. ___ . _______________________ _______ 223 .NACA 63-209 ____________________ ~ _____ . ______ - -- --- ---- 161 NACA 63-210 __________________ ._ _______________________ 162 N ACA 65 -418, 0= 0.5_ ____________________________ _______ 224 N ACA 6Sa-618 ______ ._ ___________ _ _ _ ______________ _______ 225 N ACA 63 -012 ________________________ - - - _- - - - - -- - - - - .. - - 163 NACA 65a-618, a=0.5 __________________________ ._________ 226 NACA 63 -212 _______________________________ - - - __ -.--_'. 164 N ACA 65 -021. ___________________________ - ______ - ____ - _ 227 N ACA 63 -412 ______ . ____ .. ___ . ___ - ___ - - - __ - - - - - - -. - - - - - -. 165 N ACA 65 -22L __ .. _ _______ _ _ _ ____________________ _______ 228 N ACA 63 -015 ____ .. ____ c _. ____________ - ____ - - - - - _ - - - - - - - 166 N ACA 65 42 1. __________________________________ - ___ - - - 229 NACA 632-215 ___ ~ ____ ._________ ._ - ________ .. _______ .. _____ 167 r NACA 65 421, a=0.5 _____ .. ______________ c_______________ 230 NACA 63 -415 ____ .___________ . ____________ ----------.-- 168 c NACA 65(216)-114 _________________ .. __ __ _ _____ _____ _______ 231 N ACA 63 -615 ___________________ - ___ - - - __ .. - _- - - - - - - - - - - 169 NACA 63 -018_______________________ ___________________ 170 N A C A 65 (421)-420 ____________________ - ___ - - _____ - - - - __ - - - 232 a NACA 66,1-212-- _________________ ._____________________ 233' N ACA 633-218 _____ . _________ - ___ - ___ - - - - - - - - - - - - - - - - - - - 171 N ACA 63 -418 ________________________ - ______ - - - _- - - - - - - 172 NACA 66,1-212 with 0.20e split flap (lift and moment characteristics) __________________ . _____________ ' __ - ___ - _ 234 N ACA 63 -618 __________ . _'" _- ___ - ___ - - - - - - - -. - - - - - - - - - - 173 N ACA 66(215)-016 ______________________________________ . _ 235 N ACA 634 -021. __________ - __ - .. __ - - - - .. - - - -- - - - - - - - - - - - - - - 174 NACA 66(215)-216 _______________________________ - ____ -- 236 N ACA 63 221. ____ .. _________ - ___ - - _- - - - - - - - - - - - - - - - - - - - 175 r N ACA 66(215)-216 with 0.20e sealed plain flap_ ___ __ _ _ _ __ ___ 237 N ACA 634-421 _______ .. _______ - _ .. _- - _- - - -.-. - - - - - - - - - - - - - - 176 N ACA 66(125)-216 with 0.20e split flap (lift and moment NACA 64-006 ___________ --- _____ ---- --. ---------------- 177 NACA 64-'-009 ___________ .. - ________________ .. ___________ 178 characteristics) _________________________ - _- __ -- - - - __ - -- 238 NACA 66(215)-216, a=0.6_________________________________ 239 N ACA 64-108 ___ .... _ .. __ . __ - _____ . _- ___ - - - - - - - - - - - - - - - - - - - 179 NACA 64-110 ____ . ____ ._. - __________ ---------.-----.----- 180 NACA 66(215)~216, a=0.6 with 0.30e slotted and 0.10c plain, NACA 64-206 ____ .__________________ .. ______ ---' ---------- 181 flap (a) Airfoil-flap configuration _________________________ - _ 240 NACA 64-208 ______________________ . --.----------------- 182 (b) Flap configuration _____________ -_ --- -- _____ - - __ -- - 240 NACA 64-209 ____________________ - ___ -- -- - ---- - -- - -- -- -- 183 N ACA 64-210 ______ . ______ . __________________ - - - __ - - - - _- 184 (c) Aerodynamic characteristics. Slotted flap retracted___ 241 N ACA 64 -012 ____________________________ - _- - - - _- - - - - - - 185 (d) Lift and moment characteristics. Slotted flap deflected 22° _ ~ _______ .. _______ __ _________ ____ ___ 242 N AC A 64 -112 ________________________ - ___ - __ - - - _- - - - - - - 186 NACA·64 -212 ____________________________________ - - - ___ 187 (e) Lift and moment characteristics. Slotted flap deflected 27°___________________________________ 243 N ACA 64 -412 ___________________ - ___ - - - - - - - - - - - - - - - - - - - 188 N ACA 642-015 _______________________________ - - - __ - - - - - - 189 (f) Lift and moment characteristics. Slotted flap deflected 32° _____ ._____________________________ 244 NACA 64 215 ________ "_________________________________ 190 r NACA 64 -415 ___________ . ___________________ . _____ -_ -- __ 191 (g) Lift and moment characteristics. Slotted flap NACA 643-018 ___________________ . _____ ._________________ 192 deflected 87°___________________________________ 245 N ACA 66(215)-416 ______________________________ - - -_ -- - - 246 NACA 643-218_ .. ____________ ~_"_________________________ 193 REPORT NO. 824-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS Page Page NACA 66-006 _______________________ -.: _______________ - __ N ACA 66 -415 __________________ "_ ________ __ ___ _ _____ _ 247 256 N ACA 66-009_ . ___________ .. _________ - _______ - ________ - __ NACA 663--018" ________________ .. _________ ~ ______ ._______ _ NACA 66--206 ________________________________________ - __ NACA 663-218 _________________________________________ _ NACA 66-209 ___________________________________ .-- ___ - __ N ACA 663-418 _________________________________________ _ 250 259 N ACA 664-021 _______ .. ______ .. ______________________ .. ___ _ N ACA 66-210 _____________. _________________ . __ - _______ - __ 251 260 N ACA 664-221 ___________________________________ .. _____ _ N ACA 66 -012 _____________________ - - - __ - - ___ - - - _ - - - _ - -- 252 261 N ACA 67,1-215 _______________________________________ ~_ ..
NACA 66 -212 ___________ -____________________________ _ NACA 747A315 ___________ .. _________________ ._ ... _______ .
NACA 66 -015 ______________________ - - __ - - __ - - - - _- - - __ -- 254 263 NACA 66 215 ____________ . ____________________________ ._ N ACA 747 A415 _____________________________ .. _________ _ '255 r a:::
~ ~ ~ >. Ea "'l S t< t:J ~
..... ~ .....
z o en
> g o o
/.8 r-r- f- f- .
60~
I.E I T ~, I I I ' l' .L1r{ .8
l.tl I.... I r I
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It:. IP IA<! roughness I I rouijfiness '- ft11e,lered.
o· 0
J 0 . , 1 IP- c .036 .032 lf1 LO~ I , ~ I /.0.
S1b'~
I
1/ I I
pO$jtion j!:<" I , , ,
"'Df""
\/
250' .250 Standar a.c.
.110'
" 0.250
, F-' .8 coeffIcient, !.!v
1-1--
simulatid ""'10( " lift .
'" ~~ I
R
'" ~ fQ 6.0 9.0 6.0.4I-Standard
.lor;1- ~]fc 6.0 I', I
.8i.
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, ,--- ) ~ , t , ,
o
'-1.6 .2 _.;} -5 -.-{i -.2 -.1 <i 008 004 " ~ airfoil section, 24-lnch chord.
.020 . ~ ~-.2 ~
~o .024 IJ ~ E
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t·
.... ·~.018 it -6 .t::l ~
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1ft
.
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Section IE\.
' .
--, 'c4 -, '-' _
-
-- r-r- .8 .4
1.8 I.e
3.2 2.8 2.4 20 -.4 -.8
v.6 -1.2
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~ q, e IJ
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J 0 -.1 -.5 ~ J 41 3-'3
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I
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o
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:v f-f-
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If fJ
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" r roughness
IP" II I I I I
,I.l .005 , , I 0 0 \ , I. I split
~;;t, coefficient,c
.8 "-. I I ~t~ndo~d 1·1 I 1 lift
alif~ .250 Standard
Q2.50 I ".2,BO "-
~
.6 -.4
"'" I simulated
I', i'o. 1 10' .+- -f--- Section 1\ - R J!lc.
1\ '-r 1-0. I I
..
3.0x 6.0r--1-·-+1.
9.0 6.0! O.20c 6.0 1\ I'\. I I o 06,01 <> c. v 1", .4 ).
-.8 "-:--' 1\ ,,~
1\ l"
\
h\
\ 1\ \ ,2 -1.2 In 10 '- ,V-- 0 I I 1 I
4 o
.2 -.3 -.4
,2 -./ -.2 :36
ti airfoil section, 24-inch chord.
EQ 020 ~
~o .024 .016 .012 .008 .004 " ~ o <.J
QJ §
..... '(; ~ ~ ~
~ ..
J' QJ 8 ()) ~ t: ~
0(l09 ..... .~ ~ "- '15 .0 ..::: \t) NACA the rr: '- '1 \' 18 '-I~ '\. :> L1::l - I , , p ~ ~ deq \ \.
'~'~ ~ 1\ i,7!
1\ «., 1\ ~ P\7 -,j Aerodynamic characteristics of If p / ~"t;< I~ ~ '--'-- I, IV I~ f attaciT; ...r1
If; ~ I"{ If
of II '1f 6' II) 'I .?
.~ f) ~ kl angle
_W I
III -8 1// ¥ I'<:l I/.
/) VI ld Irf
} "
II f4. ~ Section
"
\ - rI-. -/8 -E!4 .8 .4 1.6 1.2 -.8 -.4 38 32 28 24 20 -1.2 -1.6 -2:.52 rS t: l> III () v t: \J q, i-~ .~
~ <l::: :g
/:::: 0\1) -.5 ..,2 ~4 -.I :t 0 ~ ~ ~ 0 u ..,:] t: 'II ~ J ..... ~ ..... '0 :<:: ....
m q ~ ~ ~ o "'l > ... ~ "'l o ... t" ~ :; ..... CI-' ~
z » 0 » .... .p.. 0 0)
1.8 I i ~ b II If b' 1.2 It: V ~ .
,.( IT 1-1: / I.< .8 19' I ~ ~ lL b"'ID i.c1~ I J r ~ .02 c .036 .032 -' W {ef!cted , 1O 11 I .4 ty'
l<\. ,.,... I
I ,.
J roughness ~ ~ I ~ " .020 .007 , 0.028 , . , - J= / .8, ~ 0 coel'ficlent, 1 I I 1 I J
."'tr""
"l' o.c. position Standard roughness Standard I lift .250 ~ f:G) 1\ f- .6 Simulateld ,$p'fif(1ap -.4 1\ Ib. ~~ O' Section }. /
to.. I'r-. I~
;ric r-.... R I 30x 8.0,,±li:.25O 9.0 6.0 6'.0 8.0 Q20c ~ IS I 1 1 1
J1J
.4- o lJ <> Ll. 'V y -.8 1\.1)., .
\\.
1\ ~\ .2 -/.2 - 1~ , 0 ! , I , 0,
o
-1.6 .2, airfoil section, 24-inch chord.
-.1 -.4 -.2 -,2, -.5 ti ~ti 018 ~ .004 ~ .024 ./lEO ~ Ci
'= (b 0-.3 ~ §
..... ~ ....
.~ ~
.~ ~ ~
~~ ~ '<::: 'f;.Of2 § ~.008 1408
:lJ· {, ~
; NAOA the IV 1<> /6 I IX '.'( I \ ' - ~1.6 ~ 10; ("'I 110.,.
deg li Ii IA ('Z1l 1--- DI ).
, OC u-
1'\ VI f-
.
Aerodynamic characteristics of 117- IJ IU I , II!
fir I,l attock, .
1)1 1.1 k-..
~ of /I I~ D;' J .
1}lI [7 '.>
I,I:J r-k>
II angle
II 0' IV'
I I -8 I , J , IU / , I~ /Q\"V In Section -16 -24 f---- f- f-- I- ~ , , , ! .
-32
o
.8 .4 -.41 -.8 1.8 1.2 3.8 3.i? Z.4 2.0 -1.2 -1.8 Z8 -2.() ~ 'II -<.!' () \J \J
.~ .1:; !i:: '" ~ "" .~ .;::: Jl
s
.I ...
..3 -.4 -;/ ~ ()
,J ~
~-;2 .!!! .V ~ ~ l
I , >- ,.., "'" >- t::I "'" m >1 0 ~ ~ "'" >-I ,.., t"l '"l 0 :.- M l:Ij 0 ~ :.- d .., .... (') m
l:Ij trl "C 0 l;:j .., ~ 0 00 1-:1 "'" z 0 Z t" :.- ..: 0 !:U 0 M !:U
I.J::>.
- ~
i I I /.6 D II A ,/' 'j:J / / n <;/ 1.2 I !'
f J
1/ V V
I 1/ V [;0 I 1 60.· II ~ I I .8 I{ iN J 'Y I 1 ~g;::P' c, .036 .(132 .(1/.'8 V w deflecfed \ 'I I /.0 / .4 ..... I 1 I I .-I
"
i \ roughness I roughness flop , \ Ij,lc , I 1 .0.09 .007, , , 0.013 1-- I, splil coefficienf, -L 0
+
.8 position I 1 1 lift :ric .J a.c. :;tondord S1andard '..: 1.250. , 0..250.
l\- l.,.:; - 1 I .6 '\ -.4 bt- simulafed I >- Section hi \
;rIc ".A
1\ I'-- R 9~++_iZ47 60.
3.~rr 6.0. 6.0. o.2o.c 6.0.
1,-\ .4 V' o [:, \?
"t;:!... o o
~ -.8 R ~
r\
\ \ ('\ 'l~\ .Z -/.2 ; ?
)~ I 1 I ,
o
.2 -/.6 -./ -.5 -,2 o <.i airfoil section, 24-inch chord.
~o .0.24 .008 .00.4 :! t: !li ~ ~
~ '"'"
c:: IlJ 8 g- c:: \,J
",,,.0.20 .~ .~ ..... .~ ~.016 -6.0./2 .....
<.9i 1410 , ,]2 -i ---I
'-- I - r- tJ
+- NACA
I the I I I of .
I I J J.
I j ~ ~ tx; -~ ~ I ITC € ~ i ~ \ I '6 I I~i i1: P deq .If ~ I lj:: J S; (K
" «OJ
I i I~ 1£ h:< I ! ~ i Aerodynamic characteristics ~
I 'si I~ , i
j/f attack) ~ / 0 'f of J ~ If ff jtf !
angle IL i _VI .t IL -8 +-+- JA V " IL I !
\ ! tI
"tl!{ Section q J q l-- "t
•
-/6
-t-1--
\
+ +
-24 r- r- r , -32 .8 .4 3.2 3.6 /.0 -.4 -.8 208 2.0 2.4- -/.2 -/.6 -2.0
C C (IJ 8 c:: ~
.~
...:-1.2 .~ ~ <0..: <::::- "" ~ ~
./ -./ -.3 -.4 -5 "t Q J-:2
~ B ~ ~
't
!.::: 'Q; ..... ~
't-.~
rJ2 c:j ~ ~ ;... ~ o I>j ;... ..... l:O ~ o >-<
t" t:::1 ;... "'l ;...
...... ~ c.TI
z » o .... 0l=Io .... N
»
/.6 -1 1- h. \- 1- 1/19 '(I ,
II/ Il r/
I I'J I.Z 'J I~ If / ; I I ' j IT r/.
I J:~ .
, .8 jJ :!l'l ;
I I
.!
~F' Cz.
.03Z .028 .1
.036 deflected i®f
/..--', I I
1.0 .4
I"
flap I I roughness
I~ I roughness I I
'~ ,Ifl
, , I -.026 \ I I ; 0 0 , , I I split position .8 coefficient,
I
-
.r/C~!I/C ~p.10.
O.C. .250 252 Standard Stondard rJ1
11 11ft
0.250 ~
~ k
i'..
.6 -.4 1'"'- IO~ ni"-- -t- -t-~ Section
1\ ro~ m I I
R \ ') 0 ~c 3.0x 6.0_ 9.0 6.0 6 \ 6.0 0.20c.simulated "t1.
1\ I I'r I I
o o <> l\ 'V 17\ .
.4 -.8 t!> l'c chord.
1\ \ - In ~ i II 1\ .2 / 24-inch -1.2 , I I I ; I +- 1 I I
V f-- ~
I 1 7r-- ) '0 1 7 I 2 <;
o .3 .4
.21 :::'.6 -.2 -. -.
<i airfoil section, 004 ~~ .016 ~O .024 .012 ~
t::: Q) o lJ ~ E: ~
i-~ .~ .lJ ~ .;:: "- <,J"fl.020 ~ Q) o '" 8' t:::
..... .~ i5 ~ ~.008
::: A AC the N Qf characteristics deq «0, AerodYll/llllic aftoclr, of angle -8 Sectit;m -16 -24 I.
2.
3. 3 2.
2.
-I. -I.
-2~"'32 A c: Q) c: ~
~ e
.... .~ .\.1 <::: :.::: ~, \I)
:::.8
o
.I -./ ~3 -.4 -5 ~ o "E
~ \J ~ 8
f,2 .;:; " ~
:t
I
0 :> t::' .... 0-1 ~ ~ .... >,l >-3 t:.! t:.! I,:j. :> t<J 0 :> c:1 .., ....
~ 1':1 '" 0 ~ >,l Z 0 00 r>:> .... z :> >,l .... Z :> t" -< U1 0 ~ 8 0 ~ ~ Z o·
I-' C/o:) ~ ·lfJ
Z » (') » I\) ~ .... I\)
- ~ /.6 b II
I
I II I ~ / )J/L ~ /.2 "/ I~ 1/ / 1/ ,rj ~ /
L
,z
/ ri
, FY
.8 V ;t:
I
k" """
I
z C I .036 .032 .028 ,.,..
I .4 1.0 roughness I
W"
t--.
-.018 I -.004 O.oOB r
r
position r--.:....
.8 D coefficient, I::l; 1/
"'"
Standard a.c.
.247
r-- ....... )... r-
I 03J9 lift
- -e --1.24.7
I\-.. 1""'1 1,\ .6.
"
-.4
~ I', 1>& r-r::-
Section ~ ~ 7 1\ I
R
aie 89 V" 3./x/0·- 5.
~ IV, !'xJ I \ o o <> 65.7 r ~ .4 -;8
1\ l\-
.2 -1.2
V- !'--
-1.6
o
.2 -.2 -.1 -2 -.(J -,4 -.s airfoil section, 24-inch chord.
.;
i '
.024 .020 ~o fJG4 ~ t:: t:J .... ~ ~
'S! ~ «: '\-. ~
~'rj ~ <lJ 8 g-.01C' "'.008
.§ -<::
'\-.~ :~.016 :;::: <;.;: is t!li
NACA the I i I I -r- I I .
~ f:J p ~ ~ :'\ \i\ \"\':3 'I :l..
\
t"'"
?r.~ '"{ .
;/' deS' :f o} IV Of ~!
/; Aerodynamic characteristics of
:r
Ii I
~ attach, !
II D / of if
I
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~ Section 1: 11- t -/6 I [ ~ -24 f- -f- -32 D .8 .4 I.e -.8 1.6 -.4 0.8 2.0 302 28 2.4 -/,2 -1.6
-e.G
~ t:: Il> e '->
'\-.~ .tJ
.il! ~ "-= ~ .§ "" ~
<t:: .I -.1 -2 -.5 ~ ~
J t <lJ c 0-:3 ~ 0
.... ....
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c:1 ~ ~ :> ~ o I-:rj >- H !:J:I "'1 H t" ~ :> ..... C/oj ~
w ~
z » o » N .- (JI
I.
/·6 7~ ~ I
1.-
1/ ' I I.< I I.e / 1/ II II 1/ r~ }, II ~v .8 V I
'"
V l c Y' .036 .0'32 .028 .4 1.0 I- roughness
r- I
-.013 -.0/3 position
r--
.8 coefficient, ol'j~Y/C t--, a.c.
.Z46 Standard .246 r-- 0.241 =t-O.OI4 lift i'-. "t ,~
-
.8 7.
10"- ~.4 r"l: I\.
x \ Section t---tt R .-z:lc 1: 30 6.0 t:P o 06.0 A 09.0 ~ I' .4 -.8 [:1;1 i'\ I I I Q I .2 -/.2
r-
''-
0 , I , t
7~ .
t
o
.2 -1.8 -.3 -.2 :"5 -.2 airfoil section, 24·inch chord.
ti <i ~ ~o .024 .0/6 .012 .004
rJ ~ (lJ o \J ~ ~
.... ~
'0
Jl.OZO t:: Q) o '-' g> t:: '-'
i-<' .~ .'-' is .Jl.008 2415
:::: :g
CA N A the I , ! I I I !L
t>-
~b
b. r-
\ ~~ 1\ r\r\ \ IT' characteristics of \ V- 1\ '<I': deg ~J IfilIY' "0) IIIl/ j'f) ;\.cr<)dyn~mic I~ Y J } attoclr,
i
-
of 1/ I ! I I
IJ' e-U
angle l VI ,l!
-8 [I' ~ II ( II Section !O ( 't~ \ -/8 -24 .8 .4 1,8 1.2 -.8.
-.4 36 3,2 2,8 2.4 2.0 -1.2 -1.6 -2:°32 ~ ~~ Q) 0 \J t:: ~ u
..... ,~ ~ :.::: Ji
..... 't: :;:: .1 3 -.1 .,Z -,4 .,5 ~
J ~ §
~ Q) B-
""-~ ~ ..... ~
..... ~ 00 ~ ~ Z ~ ~
? ao ~ ::2 o Z :> t" t:l >-I
:> -<! U2 g I-<j c o ~ ~ ::3 8 l':l >oj
l':l o !Xl :> 8 Z :> c:: ~ Q [fi
Z l> ~ N .p. .... Q)
.
-' " 1.8 ~ .£ '¢
Lo'i 1/
\'" !
/,
1 -
11 ~
I.e
Ip J h"V b" IT J IL' / ..LV tiY .8
If> ""'"
L Cz ./ .03 .oe.
.00" ~ .4 /.0 rouq/7ness -.023 -:016 - t'"'
-
-0.044
-
position
r--
coefficient, .8 O.
*f~
a.c. Standard .24C .241 ~ :4:r- 11ft
r--
0.239 A:
r-
I- I- '\ -- u.... ~I::::::, .6 '\ __ __ -.4 Section ~ ! I I I I R ;c/c \ 29xl0 8.9 5.8 ~~ I'-. \"" o· 05.8-1-
<> A
1'\ 10..
.4 <~ .
-.8 I\. l'-,.
I~ r\. r'-.
chord, \ '\. L.t--., C!..
1(1 I .2 -/.2 1"-...
1/
IJ -/.6
o
.c
.
-.2 -.1 6 d -.5 t 012 ~ airfoil section, 24-inch .020
~O _024 .004 c.; t-z ~ u_.O
~ §
~ .... .~ .1::1 ....
::::: ~-.4 t ~ U J.:'
" 'g' t.o08
.... ~ G;; :'IS .~ ~
·.~.016 NACA the of c4- characteristics deq «0, Aerodynamic attack,
a
of anqle -8 Section -16 -t?4
o
.8 .4 -,)2 1.6 1.2 -.8 3.& 3.2 28 2.4 2.0 -.4 -/.2 -1.8 -2.0 ~ ~ III
.... 8 §
.~ .~ :t: <:: :.::: ~
.1 0 -:5 -.1 ~
J ~ III 8.,.3 ~ ~
it-:2 . ~ .... ~-:4
~ ~ ~ ~ o o t< t:I
I:<j ~ .... ~ :>
- ~ CO
Z
l> o l> I\) ~ I\) -
' 1 I I , , ! 1 I I I I I , I I I
J
J I I 1.6 !;>
/ /
.'i /'
T J. 1.2
, II I~ 1/ l{ I); V I .8 s p;:~ V J V z '" l-Q~ y 1/ C .DY .032 .028 ......
I C .4 1.0 /
i<tr'r
y t---. :...-y -.011 -.018 -QO,32 position L--
I--
.8 coefficient, :i:/Cti
G.C. 111:1111
YT~ald
.2391 .241 ~ t--- t---:.,...
0.231 lift K: !-
-
I', 6 .6 -+-- v I\, -.4 Section \ R d!/C '!':if'.
,\
02.9x/0 n5.9-, 11
08.9 ~5'f
,.
I'" 1\ "- .4- lU -.8 chord.
!u [0:1'\
, 1'\ r .2
<: -1.2
i'--...
V
?
7 I 4 0
o .3
-1.6 .2.
-.4 -.2 -:2 -.5 airfoil section, 24-inch ~ 004 ~~ ~ ~o .012 t)
.024 c: Q) o t: q, S
~ .... .~ .tJ ..-::: .... \J ....
~ ".020 Q) o \J ~ t: u
.... .~ ~.OI6 ~ ·51 ~.OO8
NACA the , po- r- ~~
i
....., ;t:;: "'"
[\ r... ~
characteriStics of ~ .,....
.......
'i:J deg /
iZL
«"., l/, V . 8 Aerodynamic
AJI
-If!
ill} ~ f'j attock, ID C
or
IL I angle IJ' J -8
)W A
rl·
J (f Section , ''\.
'I: -/(J 1\ Il [ "'t -24
o
.4 .8 1.6 -.8 2.0 1.2 3.2 2.8 2.4 -.4 -3.6 -1.2 -1.6
-2£:32
~ u rS I:: " o \J I::
.... " - ~ Jl
.!! .;g .... it
I o
.
-.I -.2 -.5 It. ., ~
J t ~ \.)-.3 ~ ~
Ii...: ..... ~-:4
..... ~
I
t'i >tI t-' ~ l:C C l:C .., '2! P 00 I>:l "'" '2! >- '"'l ...... 0 '2! >- t"' >- t::f <l ...... m 0 pj Io<j ~ ...... ..., '"'l
0 0 ~ M M "j 0 pj >- M pj 0 Z >- d ..., ...... 0 m
-1=10
Z l> 0 l> N N -1=10
I I I I I I I 1.6 , \ I I I I I I I \ I I , I I I /
I I I I I \
I I I J--!-+- ..
/ ~ I
'rI' I I I' \
/,2 I , 1--
l··
+-+ --- I I I , I I I .klI'
RW' -
I , ._1-+ 1 1 I I I I I . -
I II LJ1t1l1
I I - .8 -.
I - 'ttl 1 1 I L .
/ / , , I I I I I I e I I ...
c / I
'.O , UJA
1 , I I I I I ) !.O
-+-+-+-++-+++++-1-+- .4
, I roughness I , y/c .D19 .0/4 I I , 1 , I / I I I I .
I , I I 1'1 I I poslflon- 1 , I / / I I I ...
.8 U coefTicien!, / I J:/c
, I / I I ITH-:T-II· '
a.C.
-. .C'28 231 . '!.A I I I I , 0.2181°.007 lift
/ I 1 1 r
1 1 I / / I I I I I 1 I / 1 , I I _6 -."1 / I 1 1 I' I I I I I 1 1 Section
l'. , I / / I
I , ~ R ::c/c 29xl0 5.9 9.0
5glbt:~tStandard
/ I I \ 1 I / I I I I I J;:, 1 0 0 <> " , I I I I I I I
ttff8H= L -)
.4 -;IJ I I , I I-I -- /
I , I / / / I I I
I I / I r- , I I , / , I I I ~.jnch~chor<l.
.2
TIllIDfttHffllllllllll1
/ -I.e' , I / / I I I / I I / I /
I I I III d I I I 1 I / I I
scc~ion.
1 I H-+
a
-/.6 -:2' -:/1 -.2 -.3 -.4 -.51 ti 020 £ci
.0041 -
~ltt .024/ .
~ ~ \) ~
.... .... § ~
.~ ~- '..:: ,,"'1j 2~~4'!\jrfoil the NACA of ~b .
- ~ ;>-- i"\ ~ ......
Qharac~eristics ~ !:t' deq If J/./ 0(., '11 IA ~~ ,!erodrnamic I/J - _.
- IY If( -- attack, .
If1. f-- - 1 0 0.
-- - --+-- of If ,-'I: J __ U.
I- ; V --- angle A -- -8 .1,'/ (,1 f-- r- {I I) r +- ..
_ r'- 'h .
.IL Section
~ tr- -/6
.- tr
"l.: \ ---- c"]J -24
r
rr
I
) 1 I I , ) 1 t I 1 I , ~
-:)2
o
.8 .4 -.
I.B -.8 1.2 -I. -I.
J.6 3.2 2.4 2.0 2.8 -2_ ~ tJ ~ <Il 8 .;;:
..... .~ .Ci r;:: «:: ..... ~ ..:::: .S Jl
.I
.5
'-.1 ::t ~ u -.,] .... ~ ~
l 'f-:C' ~-;4
.Il! ~ '"
><1 ...... o ...... t"' u; d ~ ~ > !:;;l o "1 > !:;;l "1 1:::1 ~ > i-' ~ i-'
z » o » ~ ~ .... I\)
1.6 lo- II.;>.
"" il J/
h~ I
:1 I/D
(I f I II ~ c;: ). 1.2 II 1/ 1.6 I", .
~ ) .A I!P/' .8 !-I-- - V .
c, deflected .0.36 .032 .028 1.0 .4
""
roughness flop I 1 r- -.051 _-.041 \1 - -0.088 N \ split \ position
t---
.8 coefficient,
""'f""
'r-r- .o.c. Standard I-- ...........
0.245 , lift +246 h 1"- l- i'...
i
.6
JIl11111111l -.4
'h ~t-..
t
l'
Section )\.
R
;rIc '-\ 8.0 o.20csirnuloted v... ~
k
03.0xIO' 06.0 09.0 L\ \18.0 .4 -.8 I I .2 -/,2 j..-
/"
o
airfoil section, 24-inch chord.
.Z -1.6 .,3 -.4 -./ -.2 -.5 -.2 ti ~
~O .024 .012 .004 l \)
~ (lJ o E: ~
..... :Q :t ~
~ 4'11~
~"d.020 c: ~ 2 15' c: ()
.~ .'-' {; ..::: ..... ~.016 .<:l c%.OOB CA N A the of I I ;l.,.
IS:'.!
...., .
~\ D,.
~ 10. ~ A fr; deq l"" '(J «0) It' Aerodynamic characteristics ~~ W I}:! IhJ L f) < II II ottoclr, JI.
tL Ii- II of
:L I{I
)
L L
~ J 1 / angle / ~ /I -8 II 14 1/ I"'tv..
~ I V f9..
L I~ I't: 5e<;tiqn I Ih- L~ 1"; [ -/6 -24 I , I ~ ~ ~ 1-
6 £3 o
.4 .4 .8 '6 .8 .4 1.6 1.2 -.4 32 2.0 2.8 2.4, -1.2 -I.
.3.6 -2!:32 ~ Q c: \I) o \) § \)
..:::. ~
.~ .\) -;:: "- ;::: .....
.~
o
_I -./ .,2 -.3 -.4 .5 ~ flo ~
<.i ~ tv ~ §
..... ~ ..... ~
I
to t>:1 '0 ~ Z ? 00 "" il>- z ~ .... o Z :> t< ~ [Ii o ~ C":l o ~ ~ ~ t'j t'j ::l to z .... C":l IJ).
~ i §
-
z ~ ~ ... UI
» ~
T'i>:> 1.6 h /y / ~ I
I
17' I 1\ II ,/ .v:l[ Id 1.2 '(!f II Iii , 'I ~~ ./ .8 l ./ deflected C .OZB .0;;6 .0:Ji!
-"' !!'" 1.0 roughness !
I I I y/c floo I / / / r--- I ~ -.040 -;040 \/ .4 \ '-Q066 , \ split
---
coefficient, .8
or/eft
a.c. position Standard .245 .241
r-- 0.24/
lift "'II
-
1'\ simulated .0 ') .
--~
+H
Section 1\ R
:J.OxIO' 6": 9.0 8.0
-.4 Q20c 6.0 ~ ,(, _xlc o o <) A " "t .4 Q h: [ .8 .2 -I.e
/- c---c.--
'- --
'J
o
.2 -.2 -./ -.5 ti DD8 E~ airfoil section, 24-lnch chord.
.0041 ~o .024 .020 ~ ~
§
~ ....
~-.4 ,] t:: Qj <:) v tJ,.DJ2 ~ ~
.... t·
.~.D/6 .V ~ <0;: 'G ~ : :]c ACA C4 ;>
,.., _1-
r', '"
'" "\
L ~ "'- ',-.. ~
'"
--- f' ....., --~ deg . - d7f\ l/, ~ h~ E «0, '!I: ,-
'"' iff
L._ ~J Aerodynamic characteristics or-the N I. - 'Y.
V / Ii i¥ attack, U -- ~ f - /J / of 'I J ~
Ii
} / I J angle ._- I J I -8 1/ - l I II / }.
II I\.
Section I !/ t\r 1:' -/6 ~ ~ - -24 ,~ - '--'--- 0 -:]2 .8 .4 -:8 .,4- 1.6 2.8 20 U 3.8 3.2 2.4 -1.6 -1.2 -eo - IS Q,- 0 (.) v 1I
........ .... ~ "'" .~ ..... ~
.~ ~ J 0 2 -./ -.3 -.5 ... .,.
J a S
~ ~-.4
.~ ~ <0,;
1t-
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i--" t
z » o » ~ ~ .... CD
/.6 , l1: I .1 ~ iii. f/.
1.2 I.hi: l.9 .
..L 1/ II ...-?- .8 ~ ~ V
1 1
VI .036 .03i? .Oi?8 .4 fO roughness t---.
-.034 -Q060 position l- ,....
.8 coefficient, ~ -..( ; a.c.
J:/C~Y/C 1
Standard .
.242 .242J-.031 I-- 0.240 lift
->-
i').. ro
\ -.....
i r
-.4 _L..L..
·.8 14 l'b .-<;l Section "- !'---.l..J R x/c R"-.
o30xl0" 06.0 <:>90 1:>6.0
.LL
"- .4 ~B L..
chord.
I' i I i I I I ! , ! ! !
.2 -1.2 I i I !
. '-.
/'
I 7 , 3 ~
o
.
.2 ~ -1.6 -.4 -.5 -./ -.2 airfoil section, 24·inch ti ~.; ~o .024 .004 ~ o '-' ~-2 <iJ E: .....~ :c::; ..::: ...: ~ ~ ~ Q) l:>-
.].020 8 e fl·012 u
~.OOB NACA the I -_.-
-
~
~
~
::tl ~ k bCi ...
['-' deq . + V ~ «"0' ~ Aerodynamic characteristics of loP ~ V.
I {/ attock, II if.
of If!
- 1// ~
IlL
~ angle ~ -8
1/
~ ..>V' F=' Section \ ,,4 -16 ~ (
r\
-24 'j C 8 4 v 6 2 8 4 ? 4 8 2 6 n I.
2.
3. 2. 2. -I.
3. -I.
-2-:: ~ ~ iJ Q) 0 \) c: u
.... "" ~
.~ .;:: ..... ..... ..... .0 ..::: -.2 -.1 -.3 -.4 ..,5 .~ ~
J ~ (lJ E
~ ~
..... :g ~
I
>.:> 11'0' z > "l ...... o Z > t"' > I;; <: ...... w. o ~ ><: l.l o ~ ~ ~ l':1 l':1 ""J o ~ > l':1 ~ o Z > q "l ...... l.l w.
~ l':1 "C o ~ "l '2: S' 00
- :t
~ ~ I\) -0.
z » (") »
'- 1.0 II r I
"
1/ p 1.2 1/ r<. VG <; I / I 60°
II :;e
1,( I--' I I .8 ...Cii?
Lo I I /
ro- deflected
I c .036 .oZ8 .032.
i.--' f- ..-<.
I 1.0 .4 roughness flap I l L::-::-:,
II
11 l~ -.010 -.009 I ,~ -0.0/9 \ split position
\
I-- I I
I I I .8 eoefTicient, ill--
<fer"·
I a.c.
.238 Standard
'-
i'... '0-0 0.229 1.236 I lift ~ Ie. k1 I simulated
- ~
I .6 ~r-- 1\ 'eli:::- ~4
-
\ " ......
Section R o.20c· ;rIc 3.0xIO· 6.0 ~ ~ I}, In o 06.0 09.0 t:. ~.ol
'"
.4 \ v -.8 -
".:.c
'--- [ <:'\ '--- .2 - -I.E
----
-
/ "-
'--- ')
o 3
.2 -1.6 airfoil section, 24-inch chord, -.2 -.1 ti -.4 -.5
i
~O \,j .004 ~ .024 ~ t.l-.
f.E :Q ~ "'-- ~
"
(J:020 ~ (J) 8 8' c: S? v
.....
:~ ~.OI6 fJ.012 . .... (g.008
32 NACA the I I
6 r"':1 ~
~ ;l== \V r"9 h, Eir,\ deg I~ d'f ,I;< «0'
W 8
Aerodynamic characteristics of /I
,p
.rl w.: / All
Iri v
/ attach, >i
Il2
/ I)! of 1/ 19 ~ / I~ / 1 angle Jl . / -8 'q / J;, VI ~ I Section ~ I 1 ,.
f'.".
&
t
-16 l ' .r:: ..
-24 .4 1.2 3.6 3.2 2.8 1.6 -.4 -.8 2.4 2.0 ~ -1.2 -1.6 IS" c:: <!l c: ~
8 -2.q,32
...... ·91 ;g.B '<..: <l:: :.:::: .~ ....
./ OlnO -./ -.2 -.3 -.4 -,5
;t "
~ J <!l () t.l ~ E;
.... ~
.~ .\:! ~ .....
'"1 >- Ul C1 ~ ~ :... :>;j o "1 >- ...... :>;j "1 o I-! t" t:::I >-l >- I-' ~ 01
z » o » ~ ~ N ~
I 1.6 - In I/" 1.2 1.(.
t1 l6ti / ,0, 60° - i&l - .8 I~ Id lQ:: i"" r<' W z C
p !9v
V .036 .028 ·defJecfed __ I i I .4 1.0 v v roughness I 1 flo;:.
y/c '\J 1"---..032 /~ '\
+
split position c-'
r-- Ie---
coefficient, .8 P-. L..- .rIc
-
Standard a.c. .239-.005 -- ~ lift 0.224 '-( -/- _.
,
-
- i\.
simulated .6 \ -L..- - -.
-.4 ~ h-- I-Z:; < Section 0....
~
R.
;c/c <:) lOx/if t\. t'-r O.20c
D60,tt~:230-1.0/6
o 09.0 66.0 \16.0 0.. .......
.4 -.8 --1-· -- chord.
I- -1-- --- '- I- - --- - .2 -f-- -1.2 -- ----- r-
V L..--
t-t- r-r- r--r---
~ r-r-
V
> > . I
0 ( I
" cf
o .2 airfoil section, 24·inch -.2 -.4 -'/.6 -.2 ti
.004 i -
.020 ~o .024 \,) Q
S QJ o c: <I> g
~ ;g -;... ~
..... "-.: t:: Q)
<J 8 \]--,.012 ~ I::: 4424
..... :~.016 'i:: "-.: \j .::? t·008 0!l
--J -j 32 NACA
-I - '- - -
t- t-
--!J -l
the !
+tJ
I iJ j --L I 1" I Jl I ..
-' /6
W~
deq roy-.
~ ~ «.)
l'I LSi'.
Aerodynamic dml'actHristics of J J A t!!~ I ~I J i- J 1[L - atiacir, .
1 1 .1 0 'I, of -r $V J -- IL j !/ 'Ii. r-t- I -.11 angle v-- -L f-- -8 / c- !
IJ Jtj
2fd
/ ~ 1 II -- Y' II I'<, Secfion / - I -,- -16 IJJJ-' I -24 II J-.1 I T r, c- I r r ~
[ nJ ! , , , ~ I
ir- I
Jt
4- o
.8 .4 I.E 1.6 -.8 2.4 2.0 3.6 3.2 2.8 -1.2 -I.
-2:'-32
(oN 8
~ ~ .§ t Jl
~n :iJ ..:::
.I o
_.3 -.5 -. -.4 It
-
I::: 8
'/-.1-.4 B ~
.... .~ ~ '>.: .... ~
I
t':I 6 :::: ..., Z rn to:) ... ~ .." .... 0 Z t:I .... -< ~ t9 t9 > .."
H:>- ::;j 9 > l:"' > < m 0 l:I:I C 0 ~ ...... .." .." t':I "".I 0 l:I:I > l:I:I 0 Z c:1 ..... C m ..... 0;,
Z » (") » I\) CAl 0 - I\)
i
1.8 ip I
I
::> II I- 1 / -,.-,- 1'1 ;:>.
(/ c; 1.2
/
.II, I .~ ~
I
l-O .8 /
I-rv
1 U l c .VJO .03Z .0'" .4 1.0 hi--"" roug7ness '-1.1 .035 .004 'V 0.035 U' 1'1' I I position
I - S
I .8 0 coefficienf,
.>fCi;Y/C
<I...
I a.c.
.241 .c47 Standard
r-- 0.241
lift
!\-
"- I- f- I- \.
.6 \ -;4 ..... _l- I \ \ Section 1\ R I 3.oxI0· 6.0_f- 8.8 6.0 oxjc I o o <:l o ~ \ .4
\
~ -.8
.~ a
iQ .2 -/.2 , ~
a
,V ! I ,
o
'-/.6 .2 "2 -.1 airfoil section, 24·inch chord.
-;2 -.J -.4 -;5 o ;::Ii .018 .olc .008 .004 ~ ~ .024
l ~ <I> o \J ~ ~
- ..... " .~ ~ .... ~
c:: ~ Cl \) ~ o (}.Oc, c::
.... .~ ~ -b .:: 23012
.0 <% , 32 NACA , the \-+- 1-.
¥ ' , I .... j
I , ". I' ~~
I I I )..,--1' ,8, 1,1, ~~ '\ IJ V,I ,
, JlI !lA I
A deq VI , II 0('.,
-
).
I§ ! Aerodynamic characteristics of , Iff
f7
attacK, /I
-
o
of II I IJ I , IA r.f angle ~ .
''rI
-8 I II!
II Section , , I , -/8 -24
D ? 8 4 0 6 2 8 4 0 4 8 c 6 n
2 z.
:J. :J. 2 -/. -I.
-2: ft ~- q,
..... ~£ ~ a ~
.<:i s:: ~ .15 Ji
0 -.5 -J -:.4-
.. ll"
J
J ~
G II> 8-.3 .... ~
1:~-:2 .cu :::
~ ~ ~
m o Is: :> ~ S t" ~ :>
P-' ~
Z l> o l> J\) W o .... en
JJ
-I1 1
ill 1.6 II,! Ie / / J
[r v
L /1 -I- I J,'/ 1.2 II V V f U
v f.1:'v
II II :?
.8 c
fa IR
.036 .03£ .oes V 1.0 p roughness
- ..-0-
-.043 -.021 .4 -o.q~ iIY Ir..
-
n""
r- pqsition
.8 coefficienf, 1-'~lLLlI_1 >/e Standord a.C .239 .243
r--
lift .l....
f-o.231
r--
.6 r--.
r"- xlCP 1---1-1- Section ' :rIc 8. 9 6.0 -.4 .......
I , 02.6 1 060 <> b "'< N-
1\ lLL
.4 ........
r"-.
1\ \ 1\ 1"1:: \ -.8
rr 1\
[ IV .2 -1.2
r- '-
,V , 0
.z o
... airfoil section, 24-inch chord.
-3 -.5 -./ _.4 ti
~o .004 i ~
.020 .024 t.J u
ai-.2 III Cl S ~
~ ~ ~
.... "-.: .....
t.J'tj ~.016 ~ t u
§.012 \) 23015
.... :~ :;:: .s:!.008 ..... ~
-
-
I- f- l- I- f- e- ~ f- f- f- '- NACA l- l- t- l- t---i f-- l- I- t-- l-
q -I- I- Jj
'l
1 1" i
-t- t-I- t-f-- r-f- t-f-- t-
T
the I - I I I I I I T I I I t ~ ~ i I I \ . !
I I I~ I I ~I- deg ii' «0) AQrodynamic characteristics of I 1M 1/ I I ) attach, I IT
a
of I I A angle II -8 7I
II II i
I II Section "I III1I _18 II -24 II l- I-
~ ~ ~ f-I- f- I- l- I- l- l- I- l- I- f-- I- I-~ I--- ~I U--.l111
1-' I- ~I- I- I- 1-1- 1-1- l- I- I- l- 1-1- l- 2
0 -3 .8 .4 1.2 3.2 1.6 -.4 3.6 28 -.8 2.4 2.0 -1.2 -1.6 -2.0 ~ ~ v fS
..... .~ '" ~ -.; .§ ~
.;g
./.;::
o
-.3 ~ -.5 -.4 ~
i!.·1 ~
.... a ~ 8
:\j ~-.2 .... ~
I ~ I-< >-l ,.., trI trI "'i 0 l:t! :>- trI l:t! 0 Z :>- q ,.., ,.., (':) m
"" ~ :>- ,.., ,... 0 Z :>- t-< :>- c; ~ ,... m 0 l:t! ~ ('") 0 ~
l:t! trI '"d 0 l:t! ,.., Z ? 00 z
t-J. ~ 00
Z » I\) CA) 0 .... en
» (1
1.6 1 , I I I I
l
~ / 1/ (" II :f I~ 1.2 -{ ~ V k:: l(' 1/ I ~I;W .8 1"'7' b:::: V c, .038 028 .032 V .4 /.0.
roughness
r-
.017 .007 0.019 position
r-- l-
coefficient, c.
.8
"'C~""
..LU~..LUJ....L D. Standard .E4/ .243 t-- I---- 0.236 lift --.... ~
'-
r- '"
.6 ~4 R.. r' '-1-1-
-
\ Seciion
Ll
R xjc "P-l L I)" oS.IXIO· 06.0 08.9,f-t-t- 66.0 \
'" ~
1\ \ .4 \ -.8 it , R. r--.
\ ~
,1: .2 -1.2
_d\
I--- ~
1/
, f) airfoil section, 24·inch chord.
o -1.6
-.2 -.3 -.4 -.5 -.f ti ,; ~O .004 & ~ ~ ~ .024 'OEO \J
f.2 ~ -..; .... ~
~ 23018 ~
Q) 8 &'.012
.... -§ .~ ..... ~.008
.~ .~.016 ~ NACA the oC , I 1 J ~ ~ ~ - I\.
't~\ ¢'I}. - ~ - .m<ml~ (.I, deg I~ t 'II.IP «" W Aerodynamic characteristics # -~ Ii I -- affac/T, !
l of ) It angle Ih Efl -8 ~ 'l IF ( I~ Section I.l' -10 , el)- -24 .8 .4 3.6 3.2 /.8 1.2 -.4 -.8 2.8 2.4 20 -1.2 -1.0 ~ -Zf!3Z
c;,~ ~ QJ 8 \J
.... " 'i..:
.\J <i:: <:::: :.::: .§
.~ I .....
-.3 -.5 -.2 -.4 !to ~
l-.I u 8 c: ~
....
.~ ~ ..... ~
~ ~ > ~ o > .... "'1 o .... r' t:::I ~
m d l:C "'1 l:C > .......
~
z » (") » N eN o N ....
I I /.5
Ip II »
~ / l:{ 1 - Ii II I .
/ 1.2 II V I.,-'/- V I It' l? V:r;.<:l
-
.8 / I;:::. pc IA l ~E( 1? (-V I c .036 .03Z .028 -1-1-
r::o
.4 1.0 .......
I .026 ~
- f<> , , I
-.008 0.072 p -1- position ~ -p-
r--
.8 0 coefficient,
>tcryjc
ac. Standard rouqhness .234 .238
-- -
I 0.223 lift
-- J'!:' ~
j- I--
-
I',.
.6 -.4 _j- _j- ih Ih ~I:' Section \ R L 8.9 5.9 :ric 3.01<10' 5.9 ~ o L:.
o ¢ ,~ I\. i'.
- .4 ~ -.8 1\ In i'..
1\ 1'\ ( ~ JU 1\ .C ....
-1.2 R" !--
7 "-
-1.6
o
.2 airfoil section, 24·inch chord.
-.2 -.1 -.J -.4 -.5 ti Ed .016 .004 ~o .020 ~ \.) <b .024 ~
1:>2 .!!1 ~ '1-.:: I ~
~ rJ
~ g 'f;.012 t: ~.008
.... .~ {; ;g v, 1<:: 'i;: 23021 , I I I .12 N A CA f-~ the D 11:--:> D-
J... '"
'-'~ /6 .~
~ to
"
- '.!
~ Jr/ 'V .~ /h ~~ cc.,deq J.~ V Aerodynamic characteristics of - I)J [ 'V 'tdi attack, ~ if I. - of J - '( tf y angle ..4!p\: J4.
-8 I/J f VI
t
1,1 Section J ,- 1\ I,l, -/6 -,~ "C '"\ - - -24
U_
0 -J2 .8 .4 1.6 -.4 -.8 2.0 1.2 3.6 2.8 2.4- 3.2 -1.2 -/.6 -2.0 ~ ~ 'II ~ tl
8 ~
.... .§ ~ <t.: ~ ~
:2
-.4 -.5 -.1 ./ :t.
/to
<;J 0 <.;)-.3 ~ §
f-.z ~ .... ~
.~ .tl 't
I
trJ t-:3 to:> z > ~ 0 > t"' > ~ ~ trJ trJ ~ Z t-:3 ... 0
~ >tl 0 ~ Z ~ 00 "'" Z ~ 1-1 Ul 0 0 0 Is: Is: ~ 1; Ul
1-4 c:.n 0
Z » 0 » I\) CN 0 I\) -1=10
lfJ ?"
'7 1.2 / / L I-' ~ d / b' In I~ .8 ,/ , iO'l/ II 1'-',1,.--' z f C y .O:J6 .032 ~ .4 /.O·Oi?8 .IV V rouqhney ylc
r-- r--
.048 Y 0.102 j.065 --+ ,
r-- position
I--
a
coefficient, .8 .rIc a.c. Standard _1-.lU~-.L11 .223 .231
r- I--- lift
1-0.212
-
"", N I--- .6 -.4 Section ::::: 1>-- I R JI/c 3.0x/0'-
-- 1\ 10. J
05.9 o 08.9 /).5.9 I'.b.
.4 \ n -.8 '\..\. ~ c \ 1,\ ( I~
.z
-1.2
-
I--
l7
['...
7 , OJ ./ -1.6 .2 . ·.2 airfoil section, 24-inch chord.
-.3 -.2 -.4 -.5 ti ~ ~
.0/6 .004 J
~c .OZ4 .020 \J ~ l2
~ .... ~
.... ~ &
~
t: Ql 8 (5'.012 c:: ~.008
~ V)
.... . .G .;:: 'Q; iJ :g
23024 ACA N the "r
r- ~ ~
~
m
/"> ~!I::: ",.,o!o.,"
i~
':1j , r-.
deg }, /,l ?I'
fA «.)
r/ Aerodynamic chamcteristiCl' I/J V; l II attach, I of
.-, II.?
~ 'I "'" .
1: onq/e '(/ -8 {'. W A 'L , } Section 1:)-'- l\,..l i -M 1:bt, ('-{ .24 .8 .4 -32 l2 1.6 -.4 -.8 32 Z.8 2.4 2.0 3.6 -lZ -1.6 -2.0
~ ~ \J §
~ ~ ~ ~
'r .!J! '.:: .1 -.3 -.5 ~ -.4 c!-.1
~ § i
..... ~-2 ~
-..;' ....
~ Is: Is: ~ ~ ~ ""l o .... t:"'
l:tI ~ :>
...... 01 ......
•
z ~ ~ W
o (J) .p. .p. N o
,~. --; 1.6 ..
:::..
.~~ [Y '::-', tri'lA, Elmlik' 1/"i-,' ,~ ,'II /.2 11/' 1/. ~ IL
~ -'
i ~
II 'ff.
Id V .8 ,Id; L ~, ..
;;;~' V z ' .OJ&' .032 ;0i?8 C '.
k ,
1,0 111
.4 roughness t--., !A- -.049 -.054 Q035 - pos/han
I--
.8
l-111 0
coeffic'-ent, c.
~ ~lr
a. Standard .264 265 f--- l- ,....
0.255 lift ~~
r- t- r-
--
r--
.e
-.4 '1:b,
-r
-
S~cfion 1\ II R 6.0,--T- ;rjc u 3.IxIO· 6.0 ~ o 09.0 o '" 1'1.
.4- -.8 'b, Iq 1\ \ ~ 1\ .2 -12 .
V
" .
, , t ) ~ t 1
-/.8
o
.2 airfoil'SIlction. 24-inch chord.
-:.2 -./ -.5 ti 020 E" .004 ~o .0.24 .
" L.3 ~
q, ~
~ f-:2 .~ ~ «.:: i- ~-.4
,] l;: <:) 1.1 &,.012 1.1.008 ....
:~.O/8 ~ ~ il .€ -;::: ~
/ 63.4-420 i 1 1 l l ...., -I J2 fi NACA b b. ~ the cr" I of ~ FI"= r" )'05: 0..
~:>: I- " .
I \ /8 .
J :;v Lh D";, I.(J.
Aerodynamic characteristics '{J U attock, lX.,deq '/I I of
~ I
, I I I I anqle -8 '(I II Secfioh \, '\ -/6 I--, i", l -24 ~ ? !1 4 0 (; 2 8 4- 0 4 8 ? 6 r), -.
I. I. -.
2.
3. 3. 2. 2.
-I. -I.
-2.':'32 fS q, ~ ~
...... .~ .1.) .... ~ ..::: .§ .:::: (,oj
.;:: .....
--;5 -./ ..
./ ;t
~ t
l <;,-;3 $. ~~4
1:" ·91 '0.;
.;g
REPORT NO. 824-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
NACA 63,4-420 with flap
~ C:> -: Or) // o ,( /// C). /, / I <\l / ~ / I
• I ;<// II
C) It ,:/// I. - /],1 / / 0) 1'-1-// / / ~. ---->I II?'/ / / //j J / / I { I I /---------,------,,"'.- 1,/( J / / tI/ ( I / / I ( I / / (I (1/ I I 1;1 \ I 1/ \/ If ,~ I 1/1 }I ", If '" I § n ] ~
i
Q.
.g
I
s
""
~
t
=
~ 'l3 'I'l <> § ill ~
:s
:§: 0; ~ t\J ~ ~ C:l -.:
~
-< J:\ -< Z
U1 d ~ ~ ~ o '%j e; ~ o >-I t:' t:! ~ >
Joooo" C'I ~ I
2 » » 0) (,.) ,J:Io ,J:Io I\) o !. pt. :r - iii"
o
~ "0.
, I 2.8
-
'/
'I>
J ~ 2.4
""
If l7
-
illl F.~ "" '/J>
-
J '/ ~/ L'!
II jP 2.0 W 11>- ,v;; IV
V yv rr
/ '>f II> I V
t< .l.-
II .v""o .03Z .oze l'1
;,V
/
l.--'m
1.6 1.2 i;J
VI'
x- J ~.k:I It" [;;7 I 30- Ij- -1 20 ~, ,)- lh lOr-- '\.40 .....
If r7 ~ ~f'II "", 1.0 1.2 If ~ coeffiCIent, --:~ cJ..
"'-:;.
'P.~ :v
.' lift :: [::::: ...,.
~ a f.
.8 ~ .8
I, ""
f'.
C"ection II •
----~ <J. 'I< 10
.4 .6 ITot'ln 'b:i b- 6'X ~!.----"-'"" R I'-- I'-- I = JJlc
I-<-: -
R Hinge It:>. ~ ~ ~
o
1.
.4
U; 10 >-.. r- ~
slotted flap.
;).., h I'-- K ~ "'- 0.25c ." ~ .2 -.4
'"
hinge location <c;.
;:1...
-
with
K
V
') I q. ~ q
o
.8 -.8 -.2 .2 ;.4 -.6 -1.0 _.2 .004 airfoil section with ~o .024 .020
~ J~ Q) 8 §
.~ t ~
'r .~ ~.
<oJ'" ~ \J ~.O/2 § ~.008'
~
.... :~.0/6 :::: IS :;:: Vi
63,4-420 NACA I (b) Aerodynamic characteristics K J:::-.
~ ~~
11 r--
I'--W lz ':of ;>-0< P-h..
deg /6
.-
~"" ly ,1\ " «0) "'\ ('I' l\ ,\,k" Iy."
lJIl \ ~ Y-Io.
It /.J1 ~ [z;( 10 ~) 1/;A ~J :J ~ ;r~ I 'I V
IM1 i¥ rt rf
aftach, 1/,4./, lIIll 'f"1 ~I/ rY l 'E of -) Vld ?
III. il V J{ -
'I
IIJI! I/~ l-;:lrtlll rY/-I II II J
II tv 17 II
(fir/ J J71J angle
V: (j ro
II II ~ If ~I';I I / J 'I
r/W I
I II;!. r!
;, IP 11, Ip III -8 ,III,
If>
'ill Ill! fI
III II )(JV Secfion I If/, lilT Wi 0 5 I rJ JSj I f 10 30 35 15 20 25 40 45 6 7XI 1t1!!1 lYU 0 0 t::. \! ~ <I po "I to.
(deg) o Ilfl fl I
ilil .
;.
-/6 , Ilf (b ~4 0 ___ .8 .4 /,2 1.6 -.4 -.8 3.6 3.2 2.8 2.0 -1.2 -1.6 -2 ,2.4 ~
rS c: 8 \::. IJ
..... .S! .0 ~ 'qi .... ~ .~ .....
'~
I
~ >-3 ..... to< t1 ..... >-<j 0 ~ ~ ..... >-3 >-3 l"l l"l "'l 0 l:tI > l"l l:tI 0 Z > q >-3 .... a w
pj t';j "d 0 ~ Z ? co "" z > 0 Z > > -< w 0 l:tI a
'""'" en ~
I
Z l> (') l> 0') W ~ ~ I\) 0 ~ ;::;: :::r :::!: D)
'tJ
1 l
-1 l
l
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I
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2.B I
17 I
IJ HJ,,+-<:II---I--...J
~
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II
I 14 I ,
II (/
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J.
- Id
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1/ j l - t....-17'
:.t> I/V ~' - -
I .) / <r H-t-t-t--+-+-I-I--l.-.LJ..J ) / I D32 ,02 .038n'-TlrT--'-~~~~ • < l6 Wlp ~ -viZ' v"--.
?o.. 17'.')'
..
, ~ ~--rv c,
'l/---;-"
'20·
" :t
~'". !L
/.0
~ '1 L2
.........
K - -.;::_
~ V
_1I
V I~ I-tl-~
- I
'
•. I) ;JJ 8 I i .
,2 '1 T
. T 8 coefficient, ..... -
~
I~ lift
~
r;.. ~~~ _ I tv~ 1()6.
F-r-~ location
.f) A ...J:....
~~
6 X Secfion
- -£ --
=
~/c
U' Hinge t-.... t-.... ~ H: R !lap.
..........
o
.4 ,2.
u.
h -I-,;f'" , slotted
\~
...,. , ,~ = =-
\ ~ l:: '"'""I- O.25C
I~ 1"-1 l-"
.2 -:4
~
['...
hinge location '{F::;: ~:'\:
"t.o
with /.
'-_ [g I
~~=r~+-kJ-+-LJ-~
o ,
/. ~~~~EO±M-ffFR=R~+ttt±[±j 7~ , f
? 4 ~ "tt -:8 ?r-r-~~~'--'~~~--~ " -: -.6 -.
-/.O ~
.,. .00·
.02. -
~. .02, airfoil section with
~ J 1: (IJ C) <J ~ ~
.i3 ... ~
.... .'1> !i:: <0.::
~ ~ 2 ~.o/ ~ ~.008
is ~
.... ·~.Ol ~ '"
63.~ I NACA Aerodynamic characteristics (e) b - -' I> ~ deq ", I:tI ~ p...", ~ I ~ J... :>-b.
Rr"- «., /0 ~ '\.1:..
~ :J< 1\ ~ i' ..
Lk I ~ ~ )" attack, II 1.# 'f-? I J I J J
if ',( II f7 of'
II ) I I I{ , )19 Ii / 'j )'
II 7
If /) ) angle .~ ~l IY' il ') '/11 1 IflR If" D Jill!
!A I§ld VI ['6
I ~ VIP
II r"'<lb
Section Illl,J I I I &>.ff
1/ Il) 'III
I !
j 1 I l -8 'J. It 'I!W!!, ~I J / (; 0 5
U ~I< rt
I I ¥PI! II 11
10 35 15 20 25 30 ct,
Id. .,/. ; V
(deg) 0 a <> L;. v I> <! 17 'lJ40 __ -16 I
"
I- , f--+- i' ? t 'J '3 'C 9 <I 0 4 8 2 8 ale
, -24
I. -. .., I.
2. 2. , -I 3. 3. -i. -2.
-2.
n
~~ e ~
.... .~ .\; <0.; 'q, "" ~ ~
5:: <!::
I!:: I!:: "<I ~ ~
Ul c: ;! ..... C1 C1
-{; I
• II
Z :t- en W ~ ~ N 0 = 0 (,.)
:t- O .
..
/.6 , I J ; r-I- .~ S- 1..2 J I .y !- ~ Ld ~ II rt
.Q
.8 V fO' I.e: ,9 Cz .Q36 .D:J! .D?8 .
.4 1.0 -- C01q/lness.
~ .008 H 00<5 .
t-- f-" ~
.8 0 coefficient,
*frHc
b...
o.c. position Stqnqa(d.
.284 .265 .
r-........ ~
R liff "- ...... ~.a;4
-
-
"
~ .6
, '0' =1= :4
Section ....... ..;;::~ f\.
I- R 9.0 30. 8.0 6.0
.qc
\\ ~ 1- o o o A ..
"- '...0
chord.
.4 ~p ~ ~ b. 1\ \.
1\ .2 -1.2 f-.
airfoil section. 2l-lnch
V i"-
') -1.6
o
0.3 .2 ~2 -./ -.4 -.J -.5 ,; =
.004 i- -
;020 .016
'10 .024 t;,) ~ \J ~
1;::-;2 .... &
.... .~ .\.1 ~ "-: ~ a ~ I;:: 'I) ~ ~.012 ~.008
.... -~ -i:: Vi
_<,;,'0 :~ ~ is-
- - - - 63.4-420.
'-- - I I 1-- OA -I- --f- A - N
"'" .- !
the
r:~ --
r--- - -
>--
'Y" deg ~R.
;..--' '/ «., / '11 [/ ~ lack.
A, f' Aerodynamic characteristics of at u; of
Is
i.l<
, I
.I 'r -} VI -8 ,/1
AI
.~
-
, , , (I' Section angle ~ f\ ; f- -/6
c"
\, 't , -24 I I ; ; ; ! : ! I !
0 -32 .8 .4 1.8 1.2 -.4 -.8 '3.6 20 28 24 3.2 -/.2 -/.6' -2.0
<S 1;::- ~ " I;:: ~
"'~ .\lj _I> <t..: "'" .~ C;; ~ 4::: 0 -:5 -.1 -.2 -.J -.4 .I :It 1'.0 ~
C) v t ~
~ 1::: q, ~ .... .'-> <0.: ....
.~ ..-::
I
'"d 0 ~ >-3 0 t.? z >- ..., .... 0 Z >- t"' >- t::l ;:S w o· ~ ~ ~ ~ .... ..., t>j t>j >- t>j >-
~ t>j Z 00 ... (") 0 ..., "'! 0 ~ ~ 0 Z ~ >-3 .... (") w
...- Ot ~
I
z » 0 » C» eN ~ I\) 0 ~ I\) I\)
- -
1.6 .
J:: JJ IJ.L rL 1, 1.2 lJL t;t
t.«
~ f;< If '] ;; 0- .8 ~.
/ c, i.'<" .036 .032 O.oZ8 roughness .4 I.
t-- -.043 .J_lI/c position -""
t--
coefficient, ;rIc .8 Standard a.c. .271 - 0.269 -0.068.
lift
r-- j.--- In r- r-
~ Q:
... -r---
~ f.---" .
\.
-.4 Secfion R 6.0XI06 I-- ~t'-, o 09.0 "8.0 ;rIc I" I'
'"\ "
chord.
.4 '\ -.8 \ "\ n .2 -1.2 t----.
r----
IV IJ 1 ,
o
.2 -.2 -.3
-.2 ·A -·.1.6
;.J
i '
.016 .004 I;J
~o .024 .020 t 8 ~
.... ..... ~
.,!! .\J <;::: ....
'
IJ' ~ e 8'.012 § ~.008
..... :l:i ..;:: .... :.:: Vj
--b
63(420)-422 airfoil section, 24·inch NACA tv' ;::J the of '<....
P.. ::> -J...
y ~.
;>- "'
::>= .....
characteristics \-"'- ~ deq ..,-':>-' ~ ~ «0' '1 Aerodynamic j '{ -
IJ) :r
attaclr, II/ ~
IL
¥ of !J (f J ~~ angle II ~ / -8 / to-..
) 1\ } Section 1\ Irt ...... -16 ......
[ -24 .8 .4 1.2 1.6 ~.4 -.8 3.01 3.2 2.8 2.4 2.0 -1.2 -1.6 -2:'~2 ~ IlJ 2 c:: I;JN
-<-' 'iJ <;::: '" ~ .Q t Jl
0 2 .I -.5 -. -.3 -.4 't.
'
l-.1 ~ IlJ I.J ~ ~
....
~ .... .... ~
-~
~ ~ ;.. >1 o "1 >- ..... ~ "1 o ..... t'" t:l >- >-l >- ....... ~
w r. 1:tJ
I U1 .....
z > o > en w N o -
~
-
/.6 Ii> ~II
.111
['L f-l f-
'['Jl
IA If
I I II J, 1.2 '1.
.bL / }: V
!.tiki'
~
-:r
z C .036 .oes .032 I J .4 1.0 I -.067 -.059 b-,
f
rtylc l _ _
~ i5:; I
a
coefficient, .8 l?::.
I
:k:/c a.e.position .264
-
t-- lift
O.264 I
i"'--N r-.l62 I--.
-
~ b:; ro
.6 ~4
-
Section O'::t---.
.[ R ;qc ~ 05.9
03.1,1t 1
b ~'\ .- .4 \ -.8 ,,:'\.
1,-\ '"
.2 -1.2 :.- airfoil section, 24-inch chord.
V
, 1
7i""- , -0
" ') 4 :J I
o
.2 -1.6 -.3 -.5 -.4 ci .;
"" ~4 rS
~ .024 .020 ~-.2 ~ \) ~ ~
"'-' ~~ ~ .... ~
t:::
cJ 8 8'.012
",-' .~ .G·OI .,:: "- fi .'2 t.OO, ~
63(420)-517 NACA the ~J i"'-i~
"
& \l-. b..
-d~
- deq hi:b '1 - !/ (0) - ~ Aerodynamic characteristics of d
r ,-
-
fA
'!f I.
attaclt, IfL ,- I 0 II ,-- of - if IF 1 - I( angle
IlL ,-
J -8 1/1 C- 1// ~
r
,-,- k Section '-\ \ -~ \ \ -16 ~ ~
1\ ,---'
-'-- .24 '-- - - 0 -32 .8 .4 1.2 1.6 -.4 -.8 3.6 3.2 2.8 2.4 2.0 _12 ~/.6 -2.0 QJ 0 v
<,J- c: <.J
f .~ .v ~ 't.:: :.::: .S? .... Jl
.;t: .I -.3 -.5 -.4 ~
-
1".1 t: QJ () U t:: ~
.... .~ .V ~-.2. "'-
-~
I
l'j '1:j ~ ? I>:) li>- ~ >- 8 .... 0 ~ >- t"' >- ~
00 l:d ~ 00 ~ .... Ul 0 l:d >-<1 0 0 ~ ~ .... 8 .., t"1 t"1 ""l 0 l:d >- t"1 l:d 0 Z >- d 8 .... 0 Ul
..... C"1
I
z en en
» 0 » ~ 0 0
1.6 1.2 I - .-\--1-1-1 .8 V~ 10 l
~I-I-+-!
~ z
.OM~-r~T- _O)2~-l--'-l---+- .Oi?8·~-l--'-l---+-
I C deflecfed I~ ~ ij;z4~rl¥->+--I-+-1--j , 1.0 .4 I --1-+-1-+-+-+-1-1 k{ -j~-+-++~-I-l -+-+-'-l--+-t-l-l--!
, rouqhness V b--th.
(lap roughness \ ",' y/c \ ~~ \ -0.048 1\ :- .8 split
+-: +==:g§}
position coefficient, iE/c Standard Standard
~~~
lift I'. ~~ ~
D;;h f-,0.258 tf H
.6' \ simulafed -.4 ~ tl -~ I I I Section .rIc IQ I,., I""i~~ R 6.0 O.ZOe 6.0-+~-r+-~~~-+~~~+-~~-I 6.0
.J.ox·i,o ~~
-I- .4 p" ~ o A '-;8 chor<!.
24·inch .2 -1.2 SOOtlQII.
~~--~--~--r-~---r--~--r--,---' H~-+~++-~J-l-+-+~-+a.c. ~-+-l-4-~l--l-' ~-+--l-l-+-+-~ f..-!-+-+++-+- ~-+-+~~+- f..-!-+-+++-+-
, o
f~-+4-+-~-+4-+-~-+4-+-~-+4-+-~-+4-+-~- J~-+-+~~~~~-+~~~~~~-+-+~~+-~~~- ~~~~~~~~~~~Iu~~~~~~~~~~~~~ I·~~-+~~~+-~~~~~~~~~I~-=#-~~~~~~
o ~/~6
.2 -;2 -.I ti -,5 OOB ~.; ~ ~O .016 .004 .024 .020 c,;
!::-.2 III Q. (,-.3 ~ ~
.... ~
' .~ .(l , ... ~-.4
r] (" 'b a 'f;.012 v.
.... .~ ~ il .;::: J; .§ 63-006l1irfoil NACA thll I I .-1 J of
-t
-< , I charncteristics ~ ..
~ bI deq ,,- r=' rjl!},.
~ :'h W «0'
~ r"c:: .i! il"
r'1 A~rodynamic Ii cJ tl lB Id II I~ attack, ~ ~ k!
1")0.
lL ...5: or ll" lA l£.
I~ .A angle ~ -8 U l)j If1 f'vJ I L ( I IfJ In
,--,-
J LL I' I Section , 1-1- -/6 -24 1.
j
r t 1
, , [ l
'"I ,,[
o
.8 -;72 .4 1.8 3.6 2.8 20 1.2 -.4 -.8' 3.2 2.4 -/.2 -1.6 -2.0 ~ <S
..... .~ <2 'Q; ~ :.:::: .;::: ~
.§ -:5 -.4 ./ -./ ~
J Q.) ~ §
r-.2 .~ :::: ·~-.3 .... ~
.~
~ ~ ~ t< 1::l ~
~ :!l ~ -.
~
I
z ~ o ~ CU o o (0
'0) : I I I : I 1.6 I I I I I I I I I I I ' I I I I 1.2 1 1 1 I I I
I I I I I I I
I I I
I I
I I I I 60· I I I
I I
.8 1 I I I I I -r-.--,-,~~~~~~~ I I IIIII111111 I z I I I C .036,.., .0321 .0<'81 deflected
I I
I I t 1.0 .4 I I 1 1 rouqhness flap
I A I
I I .1,.
'V
0~n
I I 1
, I !
split ~,
I I I
I ~tflOn .8 coefficient, I I 1 I 1 I I Standard
I I I [o.c. 1:~f11;Z~ou9hnes5. I
lift
J 01;8
I I
I I
I
I
simulated
·1 .6
-.4 I I I!
..1..1 ITT -t-i-t- $r?ction I I I :efc I I 1 1 1 1 OZOc
olf",{O;- ~~:&-t 'VtJ.O [76.0
I I
I I .4 -.8 I I t I I I I
I I I I I I I I
.2.
-I.Z I I I I I I I I .
o
°rct
Or+-r~~~~~~~~O+~~~~~~~~~~4-~ airfoil section, 24-inch chord.
.2, '-1.6
_)
-.3 till -.2 -.4 -5'
i' '
~ ~ ~
.0241 l::
.~ ~ ~
..... ,g, 1
6{HlO9 AOA N the of I I I I I ~teristi(lS I ,
--
i\ I'"' ro
deg ~ 1,.
t"\ J'
fl ""
«01 v \'" Aerodyrupnie I~ ~ Ih
z
~ I~ ..t If.
IP ~ attoclr, Ii P ~ II '£, Y of C .
II ~ ~ 0 "'....., III angle
III rj 1)1
II
VI 1,,( -8
~ A ,/.(. i"'< I,
'I
Section
, 1 1
-18 .
, -24 -- .8 .4 -v 1.2 3.2 1.6 ... 1.2 3.6 2.8 -.4 2.4 2.0 ..
-1.6 ~2.q,2 q, t::
<S 8
1:' .~ ~ <0..: ;::: "" .2 t ~
-
.I .... <l" -.5 -.4 '
,J":I t:: ~ ~ ~""'-'
..... .91 .g 'q;-.2 ..... ~
!:':I t"1 ~ ....... "C 0 !:':I 8 Z ? 00 "" ~ >- 8 .... 0 Z >- t-< >- t:I <! I--< l:Jj ><i ~ 1-1 trl trl ~ 0 U1 0 C":l 0 ~ I-< 8 trl "j 0 l:Jj >- l:Jj 0 Z >- c:j 1-1 I-< C":l U1 I"
Z l> 0 l> en ~ N 0 0)
1.6' L- -I- -- I 1.2 60·
n I I 1 1
.8 17 'do- I L I Ir(IJ;1 ~ deflecfed l r-t;: f0 1 I c .036 :.032 .028 ~ I I 1 , I. .4 1.0 roughness I .. I flap ~ roughness I ,I -
we
I I Cr'~ I 0.005 -.0/1 \ 0 :r- I 1 1 \ split \ \ \~ position r--, .8 0 coefficient,
>feU
~fr.
II
a.e. Standard Standard .254 0.251 lift ~ ~i'>!l -1250
1\
.: r.s;, simulated .6 -.4 Section R 6.0 8.0 J.OI(/06 (S'0---t- 0.20e 8.0 :9.0 .x/c a () 'V. /7 o c; .4 -.8 .2 - -:1.2 - ; I""""" 0 , , , -/.8
o
.2 -2 airfoil section, 24-inch chord.
-./ --3 -.5 q ~o .02 .0/6 .004 ~o .02· ~ III
f-.2 §
.~ .", :t: ..... .~ ~--.4
~ :8
t;)"d g-.012 ~.008 ..... .ts i::
.~ .<:i <i:: ~ {5 VI
63-206
J J J J
J --I
--I 32 I ~ > 1 '-'- thc NACA of 1 1 .
I
M
~f"t, deg
'*' lV;: Ii
R"- .A
~~ «0'
1 ~ v ~ ~ ;;x;; ~ Aerodynamic characteristics ~ I~ ~ 1\ ff ~ It:l 1,\ I Ii attack, f. ~ I~ lat .J J' of ft anqle I l± J.f1 :r:f )..
-8 j I
, J
,~ In r<: L.i J J Section ~6 I -24 .
.-
•
I r ) ?
o· B 6 4 0 0 C .0 .4 .6
,+ .8 ,8 -32
2. I.
3. 2. 2.
-I. -I.
-2.
Cb ~ ~ .c:: IJ 4-.~ Jb .~ .;::: ~ .l:i <i:: .....
.~ -.5.
-./ •J ~
J Cb a-.3 t:: E
~-.2 .~ ot- ~-.4
:::: .
.il!
I-' O':l I-' t::I ~ "'j :;- .... !:>:I "'j o ..... t" :;.
c:j ~ ~ :;. !:>:I "1. o Ul • N U) » 0 » en w O· z /.6 !
j t'9 IP I ~ I/j 1.2 II V 1.1 P /.'P' 60f '« ~ .8 I I ;:;-'~ V I I ~ t;r- z C I I V defleoted .028 .036 .03t .1~ I I .4 , 1.0 .1'0 I I ~ roughneSs , roughness flap I 'VIc I It' -.032 , !
I"'" -0.018 --r split , I position ('1 p- ~ 0 , coefficIent, '62 ~ ;rIc Standard a .c.
iStandard r lift I :I.... T~~3j_.028 q.260 '1 TTl !!o.t::...
I simulated -.4 \ .6 I 1 I I " 1\ SectIon ...
I I .0 R o.20c :rIc 03.0xI0· 060 09.0 [76.0 /::;6.0 \19.0 - lfijo -.8 .4 .- -1.2 .2 I ~ 1 airfoil section, 24-inch chord.
7 I 4 -1.6 r , o ]V , '0 -.5 -.4 ti .2 i -.2 e: (,j \) 8 t ~ .0041 f-.z 'i:: .....
.024 .020 ~o 63-209 8 ~.012 § c,;" c: t.008 ~ ·~.016 ~ 'Q; ~ ......
NACA the of I i H "1 b- H -; I ---j a=r 1-1 - f--l H I ~rol deq k..,H JI 0:-1- «01 I- 8 ~iQ _Q Id Aerodynamic characteristics .1 l}i - J J1 V IF attach, If' of II I I-<:~ II I // angle A I Iff -8 H ~ ~ .1 119' I~ VI I.J,· Section -/6 I -24 1--1 I- I- ~ l- 1-1.- l- I- 1-1.- 1-'- I- I- ~ '- I- .4 .8 -.8 -/.6 -2.~2 1.2 -.4 -1.2 1.6 3.2 2.8 2.4 20 3.6 8 c: u CoI- .... .:::: Jj f .~ .U ~ 'Q; ,..
r~ -.5 .....
-.4 ;t ~-.3 ,l-.1 I:: III B .... ~ -t.:' .~ ~ -...-.2
I r:JJ
t-3 1-1 0 Z :.- t" :.- t::I < 1-1 0 );j >< (1 0 ~ ~ .... t-3 t-3 t;j t;j b:j 0 );j :.- t;j );j 0 Z ~ .... (1
);j t;j "d 0 );j t-3 Z ? 00 ~ .. z :.- Ul
~ I:,J ~ I
z » 0 » en
/.6
p- p- I"'"
'J
r7
]J::f /.2
II DIP
V P 1,./ r6 I I I -I 60° P' F.,2[7
I I
/ .8 II I?
J ,/ t .(1:::: II" .02 c~ .()38 .032 hntfSS deflecfed [L; .4 / 1.0 I LA rou flap rouqhness '1/ \ \ -.033 -.03 , -0.031 split .8 0 ~y/c coefficient, position ~ ~ c.
:c/c 11ft Stanford o. Standard .281 .26/ .....
p... >l::-,~ I I I--.. I I .6' \ _ -:4 simulated I I Secfion
i~ I
:x:lc I'D; R ~ 6.0 3.0x10 9.0 Q20c 6.0 6.; I'M IQI'-..
.4 o <> '" 'V /7,6.0 -:8 o p..Q: c \{ b- ~lh 1\ 1:0- .2 -1.2 I"-- ,[...-- , I 1 ! ,
o 3 -I.b
airfoil section, 24·inch chord.
.2 -.2 -./ -.4
ti .. -.5
~ .0/6 .004
~o .020 .008 ~ ~-.2 ~ 8_. ~ ~
.024 ~ .. J,j ~ ~ i.. ;f, ~
<:.J'tj ~ ~ ~.0/2 \)
..... ~ <;:: {, .~ i::: ~ :~ 63-210 NACA the - /8 >0 ~ Cl ~ ~ ~ deq 'V
, I , &
i{ I~ ~. «OJ /;'!f' .1'1 " I? ~ Aerodynamic characteristics of
"
p, 1/ If ", ottack, J!
III of '/ij III III v~ /J J angle Ii jj VI -8 1/ ~~ 15 1/ IIi f LJ< h.. " I'" Section -15 -24 " ) , -:12 .8 .4 -.8 ?.4 /.6 -.4 28 20 :).$ 3.2 -1.2 -1.6 -20 \J {,)~ q, a t-.. t:
f!.2 .~ ~ «...: ~ ~ ~
.I 3 0 :2 -.5 -.4 'It
III 8 I:: r·
l-.I ...:- t .,... ~
.() ~- "-: C/o:) ..- 0) r;J o .... t" t::j ~ :.- ~ ~ ~ o I:2j e:; Ul cj I » en w - 0 ... N z » 0 1.6 I IP rs:>!
IY I / 1.2 I~ I):: / I'" [7 ':;f 7 ./y rr .8 / I ~ J.
I I L--' 1..-\ I Cr I.
deflected F-' .036 1 .032 I .4 I II I /.O·OZ8 /( I"",~ rouqhness rOU9hness / flap I ion ylc II .(!; ~'P " ...,030 T I , .II -0.017 split j..,019 -+ " I 1 1 >- -I 0 coef"ficient, I .8 :,...., ;cjc J 265 Standard Standard a.c.posi .265 lift 1.265 ~ ~ simulated t-....
-.4 .6 l!- "'~ Section I i'- f' f\. R 6.0 O.20e 6.0 I I ;cIt: ~ I~ <>9.0 'l6J.[ OJ.O"Pi 06.0 ~ /7 \ "" ::::-- 1\ \ -.8 .4 '" N 1,\ ['\ \ ~ [\ U" -/.2 .2 ,\ ry ~ r\ airfoil section, 24·inch chord.
:,...- ~i- -1.6 () t o -.1 -.5 ti ci .2 -.2 tf () \.1_.3 ~ g ~-.4 ~~.2 :" 'i.:: 1; ....
.004 .016 .020 ~o .024 631-012 8 [5'.012 c,i\) i3 ~ ~.008 ~. :iJ ..:: Q; I I I NACA \-- the I I il I ~ 1\ I I c~ :c~ deg \l LQ I" r" I I - I VC., I I l 1'{ 8 ~ :u P II.
Aerodynamic characteristics of 1/ PI ILl tv ~ lJ IJ> - ~ IL attaclT, - .Ii Il IF of -r IF III ~ - IP r....
III !
II III angle· IAr; 1,.1 -8 I, .
f'< '// 0- 'VI \ ") \ IJ-' <:.~I u:x Section , LJ[1d \ L ~ -16 - -24 .4 .8 -.4 -.8 -1.2 -1.6 -2~~Z 1.2 1.6 3.2 2.8 2.4 2.0 3.6 Q) 8 ~ ~ .~ ~ -;.:: 't: :.::: .§ .1' I -.5 -.4 . t ~ ~-.3 <IJ () V ~ J-.1 ~ .~ ~-.2 ....
..... .\.1
I
l:O l':1 >tI 0 l:O t-3 Z ? ao I>:) li>- z :> t-3 I-< 0 Z :> t" :> t:;; -< I-< U1 0 l:O ~ ("l 0 ~ ~ I-< >'l >'l l':1 M I:tj 0 l:O :> M l:O 0 Z :> q t-3 I-< ("l Ul I-' Ct.l JoI:>.
I
Z » ~ en w - I\) .... I\)
1.6 In II / ;) J- ~ 1.2 II f } :/ I !I ~ J:: Le .8 .A:~:r ~ It'. 1 .1 ' .~ !~ I "..
CI ~ I ,O.JZ .036 .OZ8 rr <r D..~ deflected I 1.0 .4 rn;: I I ~ roughness I roughness flap , , y/c y;; I , -.034 -.029 ,
--
I coefficient, split j 0 .8 position ~ I ~r--£I ~ lift
r-- I
zlc a.c.
Standard Standard .263 .263 I'z:>.' tQ Pel I,
-
" '(> 1 1
.6 -.4
'\ "" la:-n simulated
1 1 1 1 Section
1\ f::: 1 1
\ ;rIc ~ 1\ I R 1~ 6.0 9.0 o.20c 6.0 6.0'1 ...7.~rr'-f.2541o.040 6.0.!.
I .4
o o o l> '17
"
-.8 , .2 -1.2
-
~ '--
r- ~ I I
? ! , ,I
o
.2 -1.0 -.I -.J -.5 airfoil section, 24·inch chord.
-.2 -,4 -.2 <i ~O .024 .008 ,004 r."E" ~ \J ~
..,. ~
~ ~ ...... ~
~
c: 8 t:J, \l § (J
","'.020 ,~ ,I..)
..... ~.Olo +:,.0/2 :;::: i!!
631-212 NACA the of C4 '-- 'J
file-
~ ~ ~ 17 ~ ~~ iI>l< ~ deq ~ /' /I' ~ '\ I II. «., I IV J/ !I I Aerodynamic characteristics t1i"L VI !iS Ii ottae/{, W 16 j !// ~ aT ., .
II ~ A '-<; III ~ I J angle VI -8 II I') IP '} III ~I f';l...
"'Ii I ,I I{{ Section I I'i: lj 111 ( -/6 -24
-
.8 .4 -32 /.6 -.8 3.6 J.Z -.4 2.8 2.4 2.0 -1.2 ."1.2 -/.6 • -20
,:; ~ 'll 8 c: (J
.~ ~
-\. ..;: <!: :.:::: :2
.I -./ -.3 -.4 -.5 ~
J":2 ~
.~ 8 ~
...... .t ~ 'Q; .....
~ ~ >- ~ "':I .... "':I o .... t" ~ ~
Ul c::l o >- l:d t:i I-' ~ 01 I 0l=Io I\)
z l=i » en Col) - -
r~ ..r1 1.6 III ] JJ' II r'
f7 V::s
L(17
n 1.2
V vv
J / !I l(
>-
iT' ~ 60' V Ji~ .8 f} IT' ~ ~ I ) z ....
I C .036 .OJ/?
.oes ..t'>'- ff1' deflected 1.0 .4 ~ roughness rouqhness ) I flap , y/c ~ \ -.080 -.073 \ I--- split coefficient, .8 0 position '"fQD95 r'Y I .)
-
r-.., r-c R'il
lift :C;c-l- o.c.
Standard .271 Standard .270
I\. r--: 0.271 I
-
Ii'. /'t:).
!Q, .6 -.4 r--.. simulafed I 5ection
"
Q ~ ;,:Ic ['-, /": R .
3.0xI0' 6.0 6.0 0.20c 6.0 I\.
.4 ).. [76.0 b.
.1'- o o 09.0 "V "- ~I'\ -.8 \ 1,\ ,0 ( I<n· .2 -/.2 I-- ,V > 0 , I 1 ,
o
.. -1.6
-. -.3
-.2 '-.4 -.5 oj
i
~o .024 .012 .004
\,j \J ~ ~
~
- t~-.2 :Q :::: .... ~
e3'.020 ~ ~ \J &' \J
:t.016 {j
.... :\j .§ ;::: ~.008
NA.CA 63.-412 airfoil section, 24·inch chord.
the / , ~
~ 1\ I
;lB" ·Il.
~- !RM.
1\ I~ deq Il'l >V .Il ~ F ~ «.> j, I~ Aerodynamic characteristics of & ~ ~ I III attack, f/ 8 .I.
IJ- of
"
j !J If:f t Ii ,L I anq/e J1.
1// I.-, , .... -8 I II U ~:.... ~ I 'j _I !~ Section -16 - I- I -24 , - I- , ; , , ) , .4 .8 -32 /.6 3.2 -.8 3.6 28 2.0 2.4 ~/.2 -I.
-/.
-2
Q ~ (II 8 S \J
.... .... Jl
~ '<.: ~ :;::.4
o
.1 -.3 -.4 -./ -.5 It
l-.2 & 8 ~ S
-..:- ......
:t:; .;:: 't ~
I
t-:3 ..... ..... ~ ~ ..... t-:3 t-:3 t;j t;j o.j t;j ~ t;j "d 0 ~ Z ? 00 t.:> "'" z :> 0 Z :> t" :> ~ w 0 ~ ><l Q 0 0 ~ :> ~ Z :> q t-:3 .... Q w ~ 0) 0) I
Z )10 0 )10 en w I\) 0 ... (II
I I 1.6 V JP I ~ '/ I 1.2 Ii¥ J It> V I '1" II f 60· /.y V .8 / ~ V /.,; z C V .i.
.QJ2 .036 .028
,l:> ... ""
~ deflected 1.0 .4 ~ I I ~
... roughness
roughness flap
, I
ylc , , -.027 -.034
...-
I I coefficient, split 0 .8 --I 1°.
r-- ~ lift
I I I .rIc ac.posiTtOn Standard Standard .272 ......:) t-- I--- .6 Z\ -.4 simulated ;;l. I"Irl
-
Section l\.
;rIc "< iZ\. i"'- R 6.0 9.~±-++i.271 6.0 0.20c 6.0 (J.O
JO~fT-M:270
1\ i"1 ~ .
.4 o 'V l7 o /:;
\. " f' o -.8
In IQ
"-
N't o.l'.
"
!\ .'tl.
\ .2 ,\ -1.2 h' 1\ I--- ~ I-" ..
\
['-
.V- 0 , ,
o
.2 -1.6 airfoil section, 24-inch chord.
-.I -.2 -.3 -;5 -.2 -.4 <i ~~ .004 ~o .024 .012 .008
I;) ~ Q) CJ (J ~ ~
"" :Q ::: .... ":<
c}l.020 S Ib 8 g- § \)
...:'
~.OI6 il ~ J!
63.-{)15 ACA N the of I ::> .p> "\: b-
"
~ ~~ "'jt:J :lj..
~ I)" l:k: deq ~ r' A.-:r' «., II ~ ~\,. VI v-v 'L P'\ ; Aerodynamic characteristics U ~ If d II attack, J.
~ ~ III I,f' of III I~ \'1 j f) tl V v~ !J l..
II IF III 'I M '1/
!J -8
1/ I\.. If ~ lIT' '(f ~ .J.
Section anqle ir' tJ h.
h~ -16
n ""
-24 , ,
o
.8 .4 -32
1.6 -.8
-.4 3.6 28 3.2 ?4 20 -1.2 -/.6 -2.0 ~ \) IS I\) 8 t:: ;.:-1.2 ~
;g ..... ~ ~
o
.I -./ -.3 -.4 -.5 :t ~
/-.2 CJ \.) ~ t
~
.... .§ ~ 'Q; i-.
~ ts: ts:
a o !oj ~ ~ o 1-1 t"' ~ :>
...... ~ 'I I
z » 0 » 0) w N I\) ... U1
/.6 I I I ! I
-~ IJ
b 1I
lL
ri'} r L> II II.
J L P lr1 II I /[;:; IL ':f .8 L 'rf.~ .f""' .
V
....:!!. c(
V !P' 1 .032 .036 .028 ....
1. I .cJ
deflected 1.0 .4 ",.J , I I I I ~ rouqhness flap roughness , y/c LL ., -.024 -.020 ~ r-- .
!
coefficient, sfllf 0 .B
-+ ~Q052
\I~ '--
r- 1---, lift
rIc a.c.pos/lion _1.
Standard Standard ,,~ .266 .267 I ~:J 0.269 .
r-- ~ ;0- .q -.4
"- :::::::~ I simulated
I Section ~ I'Q I N; \ :rIc r--:: R 6.0 9.0 6.0 6.0 1 J.OxIO· Q.20c I\, h h .4
, [76.0
o l> -.8
o r-... I'.
1\ 10 chord.
K. R.. ~r-...
I\, "\ 24-inch L- .2 -1.2 j...---" I-.
'-.. '-L- ,V , ! , ~ o airfoil section, -1.6 .2 -.3 .:...4 -;5
-.2 '-./ -:2
<i .004 ~o .024 .0/2
l ~ ~ u ~ §
~ ~
'jo: :::: ~
1.:,,·020 C 8 ~ § u
""fo.:" .~ ~.016 {) 1:: c!!l.008 632""215 ,]2 ACA the N , of -- , L _L- ,~ "<:l~ ~ ....
~ :>
~ Q ~ /6 0<:1 1"' FFl ~~ ~~ deq 1)1 IE;5 ~ - {)Co,
,- Ix:. rrI IB ;r
1)1 10 ~ Aerodynamic characteristics - g , ~~ ~
1/ Id !!r'
attacif, ):\.
I#' L'7 .11 I~ ~ .af Y'
,II 'J ! ~
f)
"""'"
.v /L ~
anqle
"'"
III J '11 -8 I Ill' J '<.
IlL 1r1 I": I
"
rt1! ~I 1..6 Section ~ I~ 'JDI~ I'r. -16 ~
\
'10
rq
-24 8 0 • -32
1.6 -.4 -.8
3.2 2.4 2.0 3.6 2.8 -/2 -1.6 -20 ~ ~ Cb (J
8. §
~ ~ J; -..:-1.2 <t: :,::.4 .I -./ -.5 -3 ~4 It
Q) 8
l'-.2 ~ ~ !::
1-..- o\) ~ 1-.. ~
I
t..:> >-l t"' t:l 1-1 !:Q >< ~ ~ 1-1 >-l M M "'l :.- trI :.- c:l >-l 1-1 !:Q trI "d 0 !:Q >-l Z s:> 00 >l"- z ;.. 1-1 0 Z :.- :.- <1 IJl 0 (l 0 >-l 0 !:Q !:Q 0 Z (l ffl t-<
&3
•
z » ~ 0) W III .pa. - U1
1.6 I IA I P' jIj.V 1.2
!'r' II vcr i..R
I V
.•
} .
1& ~ I ,A 60° V ;8 V'- ~ l' '::' I z C l.r. I .0.36 028 .032 L&~ rv I, def/ecfe.d 1.0.
.4 IQ' .1 jO~qh~e~s roughness flap
\
ylc -.036 r--I-. f', coefficient,
o
) .8 splif position -I- i:!r: I"- 1 liff
-- f'r
pa
;rIc ac.' J.
.264-0043 5,tondard .264+.-.039 .262 I ' TS(a;,dfrp'
-
1'\ ~ I .6 -.4 t--..
'" I I I I
simulated Section ~ 1\ I I I J 11 ;rIc .'t ~i!:-; R 3.0xIO· 6.0 ~ 6.0 020c .4
u Vt?
o 06.0 09.0 ;6 'V -.8 I\.. ~ I\: , r I':: I~ ~ ___ 21-inch chord.
.2 -1.2 L- ~ section,
r--- ~
?
o uio'Coil
~/.6 .2 -.1 -.3 -.2 -.4 -.5 -.2 ti ~.; .024 .016 .012 ~O .004 ~ 411\
~ 8 v ~ 8
~ .... ~
~ k ..
.", III ~ c: u
~~.020 S 8 ~.008
.... :iJ :t {) .0 .;::: 6.1, N ACA the oC
,- !-
:>
~- ~
:b&
cb~ruct()ristics '-" -~ l-¥ 3ti rv ~ I~ ~ ~ deg I 0-
17 ~ ~
J wC!
b. 0(0' /
rl
,r
Aerodynamic If Id IV' Ii
'"
VI IJP attaclr, A Ii f
o
A <vi< af VI Iii 'fr A J ~ fZ.
VI t
II rl IV-
/, I angle if "<>~ \l, ., -8 to
II
/} "''\ '1: VI S; ,~ II b: 1"- I'-' Section
"
f-1LVI \.
-16 --
c
'-- ,- -24 - -32 .8 /.6 -.8 -.4 3.6 3.2 2.8 2.4 2.0 -1.2 -/.6 ~1.2 -2.0
~~ ~ v c: v Ql
.... .~ .~ ~ ~.4
~ I/) .1 -.I -.5 -;3 -.4 :;t ~ ~1!.~-.2 v ~ t:
.~ ~
..... .iJ <i:: ~ .....
Ul d ~ ~ >- ~ o "'l >- ...... P:l "'l o ...... t"' t1 >- ..., >-
..... 0) ~
I
z » (") » en eN I\) en .... UI
1,6 Ip l1" In T ~ 1/ , .I I 'I- ';1 1.2
fS! IF}
1'1'/ II If}: l Ii' , 60", W I I .8 , (U c IL:>.Ili ~ ~~ .036 .032 .028 IQ: deflected 1.0 .4 f-.
2==tE i~ur~e~s rOUghness flap
'-
vic r:tl \ \ -.040
, ~
I--
t::l; -0.037 coef"f'icient, splif 0 .8 --1.
=+-.043 position h ~
"-
:--- I'n ;-..... fii'O lift :rIc a.c. ~fa;'~O~d \ Standard .266 I , 1.266
\
- "0.
Ib: I::" .6 -.4
t-.... r-: simulated
T T Section IQ I-l-I- T ;xlc I'-t-r: R 6.0T 3.0)~H-266 6.0 90 6.0, 020c
M n-I-n
.4
P',6P
o o <) l> '7 -.8 .
('\ 1\ .2 -1.2 ~
IV 1 ! 1 l
o
-1.6 .2 -.I -.3 -.4 -.5_ -.2 ~
020 f -
~o .024 .012 .008 .004
""' ~.-.2 q" () \J ~ ~
.... :~ :t ..... ~
....
~ q" 8 O u
" t)- c:
...:' :iJ :t.OI6 -t .Cl ..::: Ji
63.-615 airfoil section, 24·inch chord.
NACA the I I , I ~ 7'; '1 ,.$)'M ~:'O~ ~ ~I'l deg ~ Id ~ .I, «0' ,1:1 ~ 1>1 1\ ~fi 'iL Aerodynamic characteristics of A~ W if' A /lil f/ 1/ affacfr, .
~, ~ 1/ j, '.J/ of' ._- 1/ IV!'
-' I/~ ~ I; J/ .n ~I p.., A angle >k 1/ :/, -8 II t "'Iv'
A
J.
1/ "'"
D~ .1 'Q ~ III K Section I:r-' I~ :h ( -24 , ~ ~ , , r 7 t ? , , 7 .8 ~32 -.
1.6 -.
3.2 2.0 3.6 28 2.4 -I.
-I. -2._
,,~ 8 ~
y/.2 .Ql .13 ~ 1i <t:: "" ~.4 ~
o
.I -./ -.3 -.4 -.5 ~ ~ J-.2 c: () u c: ~
..... Jl.l .\J s:: 'Q3 ~
.....
I
I-t "'J l;d t;:J '"d 0 l;d >-3 Z ? 00 ~ li"- z > >-3 I-t 0 Z > t:"' > t:1 -< 1I1 0 l;d >1 0 0 ~ ~ I-t >-1 >-3 t;:J t"! 0 l;d > t;:J l;d 0 ~ > q >-1 .... 0 1I1
- " 0
I
z »
0 » en w Co) Q)
0 ....
1.6 I b M / II 1.2 ~ ,I /- I I I cC 60· ~
V l
/ .8 AY ,P L l c V IJ.
I .036 .O!JZ .QZ8 ~v p deflected I /.0 .4 lL!' PO I I I I
N =
roughness fla~ roughnessl I , yjc 005 I
II
-:0/9 -:02u " I-- L-- \ I split 0 coefficient, .8 position
0-+- =e'
I I tondard lift
r-- ~
r :rjc o.c.
Standard .27/ .27/ r-c r I-- ~ Ff277
-
.6 l.>.
simulated -.4
-
-
Section 0 '0-1.--..
t-t-
-
;zJc v 0...
I'Y\ _ R 3.0x/(J6 6.0 90 6.0 D.20c 8m 6.
.\ ~~ 1\ J .4
o o <> L; ~
. -.8 ~ \ ~0 \ 1\ ,,~ ."
1\ oK.)
1-1- ( .L I'" .2 -1.2 t',\ !--"- t--
V i'-
')
o
-1.6 .2 -.I -3 -.4 ~5 -.2 -.2 <i ~ci
-
~O .024 .018 .012 .008 .004
" ~ v
~ ~ § ~
~ 'f... ......
.....
c"d·020 IIJ B 8' r- V
:t -b
,,- .~ .\) ~C> .;::: tJ)
63,-018 airfoil section, 24-inch enord, - .32 NACA the - ,- 'U ~ fQ.
~~
PC
wt;'«
1'1 ~ ::>r rc: I~
In n ; 1(') 3fl ;>,., de9 'Y I~ :?-' h~
IV or"
'fI - If Ij Aerodynamic characteristics of /J Y Ii !If !~ , ottoc~ /) & II!
A ~ of
IJf Ir
~ 'V III If ~j I~ A ot7Cjle II
M VI'ii'
'i -8 ,jJ '(j ~ tL ~ /I ~ ,}J 'a I~
-
Section -tJ, qa· 4- -16 b IUC ..
' -24 -32 .8 .4 1.6 -.4 -.8
3.6 3.2 28 20
24 -/.2 -1.6 -20 ($' III ~ I;:
t~1.2 .:Ii .V 't... ...... .,,::: ...... :g ~
.;::: .!
-.I -:3 -.4 -.5 -.2 ~ ~
J ~ B ~ f
<;::
..... . .V 't ...... ~
Ul Cl ~ ~ ~ >- .... l;d ""1 o .... to< ~
I-l ~ I-l I (X)
z » 0 » en w Co) I\) -
/.6 .: I Q Ififl V I 1.2 II V f.
rrl ___
t :
1/ 1/
... 60· .
II .....
1.8 "-'
I I
.8 V ~
I
.
l..l
l ~ I.r, c .036 .032 .028 ./ deflected 1.0 .4 i.-'" I I I I roughness roughness
'~ flap
, ylc
~047 -.042
"
[---l- -0.050 coefficient, "'plit 0 .8 position
t---
I-- t-. lift D.c. ' J:lc
Standard Ston~ard
.271 ~ 0.273
- -
l'- .6 -.4
- K. re{Q simulated
..., <:: Section l\
R=Fi
;rjc \ "'I:§.
R 6.0AflS.272 o.20c 6.
29x10' 9.0 6.0 6.0 I\: [1) ~ .4 "\ 171 o o <> A " 1\ -.8 ~ IQ ~ .2 -/.2
--
f'-_
V
') airfoil section. 24-inch chord.
o
.2 -/.6 -./ -.3 -.2 -.4 -:5 <i ~O .024 .016 .012 .004
l o ()
~-.2 ~ o
;.:- :\3 <i:: \ '<
....
(1.020 c: Q) 8 8- c: l}
.~ .\)
~ :E il ..g ~.008
63s-218 ACA 1 I 1 1 I I , 1 - N the of ~ ~ i'-' ~ ~ li"l\ /6 ~ ~ '/ ":1:"1 o~ :A '>6: J/ deg .L1 lR I~ , if' o ;~ 19 'Ii oe II 1J;Ji Aerodynamic characteristics c',f !
If) "! ~
~ II I~ !L k;ls; attock., II ~ ~ ......
!Jl !J , 'I .lill of I} (f ""'-
Ii
ty rv.
VI j I) angle I ~I VI Ib ~ J -8 (l,
III y
Ii ..< III ]\.., n II l.d Section 'P 1"'< I~ -/6 I ( I ~ -24 - '-- - .8 .4 -32 1.6 -.8 3.6 -.4 3.2 2.8 2.0 2.4 -/.2 -/.6 -2.0
~ S Q) 8 c: u
~
~1.2 ;g .... it- "" :g
.1 • -.4 -.1 -3 -.5 ~ ft
J-.2 Q) 0 \) ~ ~
S
.... :~ .... ~
:t
I
!:O trl "C 0 !:O 8 Z ? 00 ~ ... z :> 8 .... 0 Z :> t"' :> t:I -<j .... rJ2 0 ~ 0 ~ ~ .... 8 8 trl trl >:g trl Z :> c::j rJ2
~ -l ~ a 0 !:O :> !:O 0 8 .... a
I
Z w Co) ~ .... Q)
» 0 » en
1.6 c- -- - n FP .....
II" '1--.- I II ':l 1.2 II! ~ I
IL ,jf
/ j~ p 60' ~f' -
L
f-?
./ ~
""'"
l c LA .038 .O$Z '.ORB
.,....~ '"
deflected 1.0 ..
.4 I I roughness t::::"" flop rouqhness y/c \1.
-.051 , <.:-.: ,
r-- ~ -0.052
split \ coefficient, .8 +-.057 tn i'9 lift
r--- - r- ~ a.c.positlon
;x/c..../...
Standard Standard .272 ~ 0272 1.27/
r-- I--- I
.6 -.4 -....:~ simulated
1'--' "" I
Section /(JO I--- 1\ x a:/c Cl'--:
:\ l'Q
B \.
3.0 6.0 a20c 6.0 60 ":::t:
'\ '"
1'\
! -
.4 f7 o o 09.0 66.0 'V -.8
I Ib I\, ~
\ , "(
Lq IQ'\
-
.
.2 -/.2 I--"" i--
/'
'--
IJ
o
.2 -1.6 -./ -3 airfoil section, 24·inch chord.
-.2 -.2 -.4 -.5 ti ~ci .~4 .016 .012 ~o .004 \j ~ Q) () \.J ~
§ ~
...... ~ .... ......
.... Q:; 3 8' § \J
","3.020 ...... is
:g 'f: ..::: JiS.008
633-418 \- NACA .
t-r- 1---'- the t- 'r-- of
-+- 24
I- I I _.
)c ?:1 ~ ~ ;)n !£
'-ru
rco -.2 ~ f-f~+- deg 12 10 >t:. ;r: I t- f- ~ IF I- «0, .-
~ 1£
Aerodynamic characteristics
"'"
IlL ,. III
II/ V r'r'
W
a/tock, ~ d II/ p'''J..
.
- -
r
III If of' .
~ -. - .11 'IL l£ ( ,(
Vi ~~
onqle II> rJ \v'-;;l II ) -8 II
v VI
~ ~
I rL
$
n
1.)0',..
Section ~v 1.:0 ~v j: -/6 1-1d: I I I , -24 ! I I , .8 -32 1.6 -.4 -.8 .3.6 3.2 2.8 2.4 2.0 -1.6 -2.0 \.J -1.2 \J ~ \.J ~
...:-1.2 .~ .\J '!... ..... ~ Ji
'i:: , :;::.4
./ -.1 -.3 -.4 -.5 "t ,/-.2 ~ ~ B t f:: -t.-- .....
. ;):! 'Q) ~
d ~ ~ >< o "'J .... o t'
U2 !:lj > !:lj "'J .... ~
I--' C/o:)
"
.
I
z » 0 » 0') CAl W 0') .... (X)
1.8 ~ In J
, , , ~ln
If
JL 1/. I be:.
1.2
l' ./?/
I ~ J -
ki, ~
/1 l'itil'"
60° ip Y.!
.8 If, >-<)-" '-( c, J).'-"" .03(; .032 .028 deflected 1.0 .4 I I I
r-- !£ roughness
rouqhness flap \
, -'-'-
y/c \ , -.016
r-
-0.012 coefficient, splif 0 Zl .8 -I- i-.013 position --':';~Ia.
I--- Y' i'1:::r lift
;rIc
a.c.
Standar.d standard .267 "'- 1.266 I I .......
I---:- I'-~ I" 1-10: .0 simulated -.4 :A.:: :1.
i'\ I Section I\.
1"-.. l'< i ;rIc b i I Or-,.. J.Oxf2T-t-0267 6.0 D.20c 6.0 6.0 el',. 9.0 8.0 . R .4 o o () ';7 -.8 I\.
: 0 ;Q. 1D.b.., .c -1.2
'- I !
I , , i I
IV , 0 I 7 , I
o .2, -1.6 airfoil section, 24·inch chord.
-.I - 3 -.4 -.5 -.2 -.2 ti ,,.0 .004 ~O .024 .016 .012 .008
" () () R
S ~ o
-...:- ;g \ .,... ~
-
~"'.020 c: <0 2 c:.- o c: ()
.'l! .~ ~ ~ .g
.... . c%
63,..618 NACA the ('J' '"'(:;p :J.., )n:) U r :::0 !:f2P'"' deq ~
a: :YrU k-r
\ ao, ,\ .~ J;.
~ 14'1 1.
Aerodynamic characteristics of II V
IJI
Iff :? I~ ~
attacK, ,1:
IW IJ!
ff ~ of '-'--Iv.
).i f)
r
~I !J fl
IP I'Z
angle IfI Y M It I -8 rj ~ III I" , Ii
VI r'!' lP
} ~~ I'! '0 P
(f ,'"
Section jq
Iv ({
-/6 -24 I , , 1 ,
o
.8 -32 1.6 -.8 3.6 32 -.4 2.8 2.4 2.0 -1.2 -1.6 -2.0
,,~ ~ I\) 8 ()
~
-...:-/.2 ~ <;... ~ :.::: &.4
o
./ -./ -.5 -,.3 -.4 ~ Q
l-.2 .~ " ~
...... ;g 'Q;
~
."10..
I
'tI 8 t.:> 2! >- >-3 H 0 2! >- t:"' >- t:I <: .... 0 l:C ~ C 0 ~ ~ H >-'J 8 trI trI I%J 0 l:C >- trI l:C 0 2! >- d >-'J .... C Ul
l:C trI 0 l:C 2! ~ 00 "'" Ul
JoI::>.
,..... "-l
I w .....
Z » 0 » en .JIo 0 N
1.6 /.2 J.
-
V
I 1I !:Y ~ .8 II IJ If1 I)~ z C 1/ .038 .032 .028 If<' deflected 1.0 .4 I
"
......
roughness " )... ~IX
, -.001
r-- t-l
~ ~ -0.006 split flop coefficient, .8 position H .....
t---
I-- 1-1: lift
.1:/c!/le 0.0.
Standard .2 IY 0.276
--
r-- 101, .6 -.4
I!.\ "'" simulated
"\
- Section
L--- I\,
~l8.273
J!lc Ib R 6.0 o.2Oc 30xl0 9.0 6.0
1\ '" I·\:~
.4 .
o o ¢ c.
-.8 ~
~gtfE_~~n<~O~d~~u~h~e~sllll
) '-' ~L\ chord.
1\ __ 1"1\ 1< 24-inch .c.: -1.2
r-
~
V
~ 1 ! 7 ,
:J~ "
o
.2 ''4 -/.6 ' -.2. -.2 -.3 -.4 -.5 airfoil section, d ~o .016 .008 ~ .012 .0
.02 l o (J ~
S III S
:i:! -+- <
...... ~
c: <II o u 8' c: ()
.~ .IJ {,
~ :£ :g c95
.~'tJ.020 63<-021 ACA N the ot :fi ::n ~ - ~ 7"'11\
:n
~ ~ ""'9; L.. L..
yi c""'"r~~ I ~ de<] .d
""' '14~
I( I"ll ~ «., IA IA IV II IV Aerodynamic characterIstIcs 1/ J" Q H ~).( on attack, .4 T lb 'if f'Z !J of : '1/ IP -...; ~ J.
'II rJ 16 /J
~I If( 'l"" "'"'"
/, angle '/ /'"iN ~ II .J -8
"
If r; III I\;( lrl I ~ <;) Section
-
eLlA I'i:)d -16 I -24 .
.4 -32 1.6 -,4 "'.8 3.6 3.2 2.8 2.4 2.0 ~/.2 -/.2 -1.6 -Z,O c; a; III 8. c::
.... :~ :::: ;::: :.:::: :g ~
.1 -.I -.3 -.4 -.5 }
J-.2 ~ ~ !:: ~
:~ ·u
-+-~ ~ -+- ~
~ ~ ~ o "':l > 1-1 >
i;l S t" ~ 1-1, ~ ~ Ol I
z » 0 » en w "" N N .-
1.6
-
) ·19 'Y I 1.2 1/1/' I~ W ~ I 60· 1/ 1 I riJ; .8
:1
~ I';;.
~ p- ' .036 .032 .028 l...-' :Lo1Q'
:>- deflected
1.0 .4 )- I , , ,
•
L..l flap rouqhness rouqhness , , vic ,
, -.017
~ -.033
r--
coefficient, .8 spiff
--+
In
r-- ~
~Lt.. lift :rIc I I o.c. position Standard Standard .270 .26'9 f- I 0.274=(004
-- ---
~ .6 -;4
"" I I simulated
6 r I
Section ~ 1\ I 1 \ \ ;rIc \ ~ ""'iIi~ R 6.0 o.20c 60 .J.O>r10 6.0 6.0 1\ 1'1 ~ 1'-" .4
'\ o o 09.0 6 'V
1\ I'\.hl\ -.8 ~ t\ 1'\:1\.
\; I?
.2 -1.2 "
-
/'
, '''-.... , 0 , r 7, ,
o
.2 -US airroil section, 24-inch chord.
-.I -.3, -.4 -,2 -.5 ti .; ~O .024 .004 ~ Cl u
~-.2 ~ ~
-..: ~ ~
- 'Q; "i-..
1.>"0.020 c: Q) 8 0- tl c: u
.~ .!2 ::::.016 ~.OI2 :g <!!l.008
.....
63.-221 , ' 32
~ r- ~
t- l- l- t- NACA
~ ~ ~
H- ~~ ~~ 1---1- ~~ ~ -- :::;:: H f-i:: i+.j 8=' ~ H- r-
t-+--
TT
the
I
j- --+-
1 1!--1-
of
rr~r- H-f~
,,'-'- rv-+
] ttl--!-
- =li
': f-+-...,....-+--'- ~ '
; I
'- 'q
--tt+'-+- ~f-rTl " ~ ~ lil ~ I-
I- :::rn=l- >- ~ I- ,I m
~
~
I ~-FfFttJ ::
i
K-1-tH-l--t 1""'1.-- H-rT
Ft
FFI-
I- ~~I-I;~
II
I-+- f-+- '+ !--I- hi-
~ I 'I
++ deg
I I I I 1 a o,
"" 6(>' ~~~~-i± a
-+-f-+- f-+- ! I ,.. "c- I ,
ff!-- 8
I1"~
iii
~
- ~I..ftt
,,' I "
Aerodynamic characteristics /::
"
rtf I I I I I I
I H
T
/X U I,
t-" +-~'"
attacA,
i
j"
~H- -I-
III " H-+-I ~I- "'_
I I UI.I
I
of' IjH- I-
1--1 ~-p y 14--~1 I I
II.
~ l1L~
lJh
..
-,
'Tt
'"
..tt::::::
~
angle
'v
11 ~
n-rr-+- +-
I"T
+=
-8
=r=-ff
~W
-+- 71
+--+-+-t+ "
~
,,' ' +-~ ,",..
lP
i
LL 1 1 1 ,,+-'~~i--LH-L
I-It Section
tt±±t
--I-<
I- ~ ,-++rttt+l ;=1$§Hf H ~ffi ~ l---. I-+-+- I '0[" I
t+-,,~,
.I
..
. -16 I I
~~LL'
j
tl~:='.~
T
~~IIII
-
I +- '
+- T
T
1 1 I I I
II
-24
I I I "",'- " I I : I
TFFW~U-JlQ-\lt'lrH-+-1--
, I
~ rt--+-Jt-+--L-1f1= 1- +-
~ ~ -.--,-t---H~
lr-r-t-+-...!-----L
H U u W t:::~~~ >- ~
I- LL W LL'
':::' 1 ' "~ _
. ,,- ': llitt
.Et ~I-'
,lL 8 Ott 'LL-~
-32 : Ui -.
3.6 2.0 -.
3.2 2.8 2.4 -I.
-I. -2.
t: o ",- III 8.
.S! .\J <;:: ~
~1.2 , ;:::: ::::: :.g.4
.
./ o
-./ -.3 -.4 -.5 -.2 If..
~ eO
~ ~ III 8 ~ ~
..... :Q ::::. ~
........
I
t:;) .... w I:1J >-<: 0 ~ ~ .... 1-3 1-3 b<j b<j "i 0 l:d > b<j I:1J 0 Z > d 1-3 .... 0 W
l':j "d 0 I:1J 1-3 Z ~ 00 K .., z > 1-3 .... 0 Z > t" > <: 0 0
I:1J
,...... " ~
I
z » (') » en w ",. ,J:Io I\) -
1.6 I I , , "l<J P
W '-
/,2 ~ W I II ~p 9'0 I II h-
J J .8
k!:Y ~ .
rr
z C ,036 .00Z I .OZ8
'" IdeTf~cted I I
1.0 .4 I I I I ro.uglmess roughnersj
t-- (101' I
\~
y/c I -.030 -.02 \ ",,'0
r-- ,.-
',I I spiff coef'ficient, p'0sifion II .8
""" I I
c.
tandard
,..-- liff
I""- h. J1IC I
0.
d., Standard .27. .275 0.27~+.025 I j
-
~ _= tion I- ~ j .6 simulajer -.4 1\ I I I I ~ Sec
\ ,,~
- I I I
J J
;rIc I~ I R 6.0 9.0 6.0 0.20c 6.~ 6.0 3.0xI0· .\ l'bIQ !'... I .4 , !7
\ o o <> 'V
-.8 if, In ~ .2 -1.2 :- /' \.......
') 0 o airfoil section, 24-inch chord.
-1.6 .E -.I -.3 -.5 -.2 -,2 ti ~~ ~
,024 .012 .004 '" \)
~ Q) o c: <II E: o ~-.4 .~ ..... :Q ~ .....
~ o \J
,].020 c: Q) g- c: ()
.~ .\) ~.016 {, ~ ~.008 ....
63,-421 NACA the I I I , I \ ~ baS characteristics of yu l,<!(C
ru
deg rvv ~,JUQa J:l'P 'f' «0) !f IJ i1" S Aerodynamic 11ft ~ IIf 1(.0( ~- d, N 1/ aftock, 1/ 'V-kl.,
la V
'"'
oT
1/ '1/ Ii "-'t rv"
~~ IJ f}
rL '---
e Ii ~""1: y III fl I'r J -~'--- angle rt I:::» ~ II W ~ l -8 II. ~I 10 ~~ /, II/ V ~ ~ 1';( Ip J Wi ~I Section Io.L 1.Io[ f1Jt -16 L -24 .8 .4 -32 1.6 3.6 28 -.4 -.8 3.2 2.4 20 ~1.2 -1,2 '-1,0 -20 (\) \)
~ t 8 Q)
~ <,..: ;;:: ]
.... :.::: II) .I 0 -./ -.3 -,4 -.5 !t ~ c: \J c:
.l-,2 o ~
..... .~ ~ 't ..... ~
~ ~ ;..- ~ o I:rj .... ~ I:rj 8 t< ~
gj >- I> I-l ~ ~
I
z ~ 0 0)
» 0 » 0) 0
1.6 /,2 - 60"- .8 j-j-j- .
I;t., I
1m
J c, yl'" I I .036-r-r- .032 ~ I deflected I /.0.028 I .4
I y
II .,.. I 1'1
I ~ I I rouql:mess roughness f"lap
'~'
\ IAl I I \ -.0/4 \ J~fT I I -0.068 coefficient, .8 spIff ~ylc position I-LI--<' PI/"-,' I
I
lift .""<.1 rt: I zlc a.c.
Sfandard b. .258+,014 .256 Standard
"" I
0.259
.'"'
1\ .6 \ -.4 I~ I I simulated \ Section hl'h p.;tg.
14 I I! IIIII
\ L zlc b. I R 3.0x/O· o.20c 6.0 I V' .4 '76.0 o 06.0 <:'>9.0 L:;.60) -.8 chord.
24·inch .2 -/'2 OIection, 0 ! 1 ! I I I
o
-1.6
.z airfoil
-. -3 -.5
-.4 -.2 ,:~ .016 ~O .024 .004 ~-.2 ~ \J ~ E: (.1
....... ~ ...... ~
~ B' € u
<"".020 8
. .v :::: {3.012 :t:; ~.008 .......
64-006 NACA the deg ~ !'f'I~ ~ ~ "0' iltI1.
'i; .J I r~ Aerodynamic characterisUcs of d ~ \ ~lz, A. 1\
w 1'(
attack, }-' ~ 'd ~
°
of .d I' 11' .1 9 ~ III ~ II a/79le J r"(~ P In.
-8 CJ.
hi ~ :d: !t II;; ~ [h I,.,!'i' Sectio/7 -16 -24 > j I 1- ~ \' 7 ! I 1 , -32 .'
.8 .4 1.6 3.6 2.8 2.0 -I. -2- 3.2 2.4 -/.
-1.2
IS C:" ~ 8 ~ ~
.... .U ~ V) .S! .;: ..... 'l::: ~ .I ~3 -.5 ~.1 -.4 It ~ t:: ~
l-.z 8 ~
..... .~ .~ :::: ....
I
~ ~ t>:l ~ ;.- l"J ~ ;.- 1-3 ..... C':l w ~ "d ~ 1-3 Z ~ 00 I>:) li>- z ;.- 1-3 1-1 0 Z ;.- t"' ;.- I:! -< 1-1 W 0 ~ ~ C':l 0 1-1 1-3 1-3 l"J ""J 0 0 Z d l"J 0 I--' 'I 00 I
Z » 0 » en oI=ao 0 0 (0
I I ~ /.2 ,_
r
/ ~ / T.
-1--1- I _, I J 60°r-i- )- - _ .8 /~ :- I I I r If I I .of1: c( ./ V .032 .036 .OZ8 J ~v I I..L deflected I I 1.0 A ,
ro- I
, ~IV r.-' V roughness flap roughness I , y/c I \ , -.027 0.015 I coefficient, SrIii' 0
i-·
-l .8 posit'-on I c.
I,ft I oXic a Standard Standard .262 ""ttb I'v..
0.P57 tI;l .6 SIi'nulote1 -.4 '0 ~ Section \
H=~:260
\ a-/c 1\ IS 6:: 1- I R 6.0 9.0 020c 30x10¢ 6.0 6.0 I .4 17 o o L>. V'6.0~ Qrd.
o J...'\. ""'i'- -.8 b h I'\.
~\ h '\I'\.
Z4-inchrC .2 -/.2 l....- ~ ! ! , ,
o
-/.6 airfoil section, .2 -.I -.2 -.J -.4 -:5 Q ~O .024 .016' .004
l ~ III 2 ~ t
"
~ ¥-.2 ..::: "'- .,.. .Q
~.020 QJ 8 ~ ~
t
..... .\J ~ il· .~ .;:: ~.008
64-009 JZ ACA the N ·~f - /6
" •
~~ fl I~ ~ I deq I I I I I (}(o.
i'~ I~ !E
1<
Aerodynamic characteristics , !.~, r;; 1 ~~ IJ ~ ot/acA, J; JJ.
1/ IT of' 7 ~ ~ VI fI 1// t,c, Ii onqle ~""
"- III
I i -8 II J "'u rq K~ bh- 1// I I I ....
Il, S~ction -/6 '.
-24 r r , -32 .8 -.4 -.8 3.6 3.2 1.6 2.8 2.4 2.0 -/.2 -1.6 -20 Q)
~ ~ 8 §
~
~/.2 :i3 ~ "-.: it: '-':: ~.4
./ .
-.5 -./ -.3 -4 It
\J (IJ 8
,l-:2 t ~ t
'f ~ .... ~
c:j ~ ~ :> ~ ':;J ..... l:d '=l o .... to< t:I
w ::c o :> :> ,..;; :>
...... ~ CO
•
Q)
z » ~ .p. - o
"0) I , 1 i I i I ! i ! ! , , , , i i 1 1 1 i I I I /.6 , I 1.2 I
r7
I (J ~ \--1---<> I 60° I I ] / .8 f /' ID I / c, I .0~6 .032 .028 p'.
I deflected 1.0 I .4 / s:o~ k I 1.'1 F I roughness roughness}
'tv'
flop y/c r
, .029
.014 , - '--- I f.029 coefficient, -I- 0 .8 split position
(, lJ(
I .+- ~ -I- I lift xlc a.c.
.256 .255 Standard T\... 1 0.259 15tandord
I
.6 -.4 ~ "i~1--..
1\
simulated. T
J
\
" Section
~I_, .ric ~~ I R 30x1O' ~\ 6.0 0.20e 6.0 6.0 I .4 06.0 09.0 'V 17 {O, -.8 I;;
1-10
.
1--[- .- .2 -/.2 - 1-1-- l.-- I- --
> t , I r , , 1
o
.2 -1.6 -.2 -.3 -.5 airfoil section, 24-inch chord.
-.2 -.4
"
i .024 .016 .012 ~o .004
tJ ~ IV () \.J ~ ()
i-!' .... ~ ~ .....
..,,,.020 c:: IV 8 8' c:: ~
.~ ~ {) .~ .... ~.008
,,-' 64-108 I I J I- NACA the of I 1 ")...
de<] 11 t~ ""\lh ~
~~ 1-l
".>
j;j~ ~ Ii Y\ 1\ Aerodynamic characteristics ~ .I .J. ~ !l.
attacir., if, 1«) I r< of
J
J l/ J an91e II 10 \I -8 lit ~ ~,y r \.
'I \ kz Section -16 -24
~ I I r
! ; 1 r > , , o .4 -32 .8 1.6 -.4 -.8 2.0 3.6 3.2 2.8 2.4 ,1.2 -1.2 -1.6 -2.0
I..l- t IV 8 c: IJ
..... "- <t :.:::: :g ~
~
o
.1 -.1 -.2 -:.3 -.4 -.5
... c;-
J 8 ~ ~
£~ ~.
:0 .;: 'Qj ....
I
I:'J "C t-:3 "" >l>- z :> >-3 I-< 0 Z :> t" t:l 0 l:ti ><: '-l 0 ~ ~ I-< >-3 >-3 I:'J t"l "'1 0 l:ti :> I:'J l:ti 0 Z :> q >-3 ..... '-l W l:ti 0 l:ti Z ? 00 :> <1 .1-< W
..... 00 0
I
Z ~ 0 ~ en ~ - - 0
1.6
&
v< 7
U
1.2 L I) / K::: ~ I
J
6qo
~ I
J
C .8 I
J
II 1.rI·.,f I I .
~ .q:~ '<J c, I I .036 .032 .OZB !:{ ~ I deflected 1.0.
I .4 / I 1 I y"(f / I rouqhness roughness T/op , y/c
, -.0.18 -.0.22
, .1 coefficient, 0.
.8 split postfion ---l-,. +,0./3 ~ 1 lift .r/c o.c.
standard Standard b;:~ I 1'-'1-,.,
0..260. 1260 126/ J
1Zl. ~ ~ .6 ....
-.4 ,\ I simu~ated I- Section
IL'. ~ I 1 1 1 1
;r/c ~ R t
3.Dx1D6 9.0. 0.2Dc 6.0. 6·f
b I'" I .4
o 06.0. <> 7
66.o...:r -.8 c .\ I\'~ ( IY .2 -1.2 i-
~ -
? !
0.
-1.6 .2 -,/ -.2 -3 -.5 airfoil section, 24-inch chord.
-.2 -.4
"
.0.24 ~ .0./6 .0./2 .0.0.4
~D ~ ~ V ~
l S ~
~ .....
~ ....
rJ·D2D 8 8' ~ v
.\:! is
.... .~ :t :;:: ~.D08
64-110 N ACA the
-I- -t
I
EE , 1
rl
~
~
+
"_ 'J.='
dm: rr- H =
1:cK: rr-
r ~I
f8
..1 J r-r--r--i t-r;::
-ll-
' l';; l/ ~r tx,,,deg
I 1 fu-:-
.Jl
~
~
'++=++4
Aerodynamic characteristics of
'-I-
r +,
~, r----r-r: ,...I..,,/!,'-t--r-r"
r-r- r+--+--+-,--,-r 1
J J 1 ++ attoclr,
"
-A I ;- !-- J t;::-l----+---t-TI j TI' 0.
of
ill!:. ---, ~
fi
--.---.---.--, ~
~
-'i ~ ---r-I
'f{:
-l-t---r- -t -I- -1-+
-!-+--+++iiTIQ.
-+-+
I angle
~
r-r-
~~+++~+ill~~ t i 1
-8
-IJ~~~lit1=r1=~ -
H= , , t±
fL
-n ------, rv 11
, r
-t..!.-I-
\
-
-~#
~I
SectIon
ttH-lr-r ++
-16
---r-
J-I -I
-l-t-t---,-,-r- 4-t-t---,-,!=
]
++ 1
Till J: -----,-l----+---t-TI -+-+---t-TI 1 I
~-I
-
r -24
n
r , - ,-----,
" r-
rr
~
J j
W
I
I
~;,,±;
-----, ~ -,-, r-r-r ,~ r
h tl1r:++-++ rj [ D rj-- t:t ttL rr
,[ ,r JEt+- .C'1-t..!.-I--I--I-
-32 .8 1.6 -.8 -.4 3.6 2.0.
3.2 2.8 2.4 -/.2 ~1.2 -1.6 -2.0.
I,)~ c: ~ \.J (:: ...
.~ ~ "- ;::: "" :13.4 ~
0.
.1 ~3 -./ -.4 -.5 "t ~
/-.2 c: 9i 8 c: ~
... .\,1 ::: ... ~
(fJ r:j ~ ~ ~ ~ ~ ~ ~ "'1 o .... t"' t:j ~
...... 00 ......
I
z » o » en ~ I\) o en
I /.6 /,2 :~ f 1111 I?" I I .8 '.J;:~ P'-' , V / 1/ V I::::: i I c, IP lro iLl I -t- .035 .032 .028 , I--'-' 1 .~ deflected 1.0 I .4 ~~ I I , I I ~ F" I I roughness roughness!
flop \ y/c ,.- ..... / I .01 , -020 , jt I -0020 I coefficient, (7 .8 split -I- position / I I lift 1\ I .rIc G.c.
Standard Sfondard .254=+ .253 _\
/\ t\ I
~ ~ J-tx:: .6 -.4 1<1
IW: simulated I I
Section
In b. I I
oXic R
3.0'~.255 o.20c 6.f
6.0 I .4 o V 060 09.0 .6 '760 -.8 .2 -/.2 -- ,= 0 r , I 1 I
o
.2 -1.6 -- ,,3 -.5 airfoil section, 24-inch chord.
-.2 -.4 -2
"
~ ~o .024 .016 .012 .004
l ~ \j ~
~
..... ;g ~ .....
~
f5 c:: v
",,,,.020 ~ 8
:~ ~ {i :g ~.008
.....
64-206 ACA N the i I I ~ 10 ~ IS ~ '''''P '--t."'" R~ 1'1:'
- deq
'tK: oCl rtf 'Y "0, <v )i2 Aerodynamic characteristics of J
I' "
Ph \ J W I~ ~ attacir, j, j
r ;g
Ii;!
---- J ~ ~ 'V'~ of P 1# f l ~ ~ 11/ f -- /I angle ~ P I ~ -8 ~ ~ 1SJ:"<7 I K ItF 1/ ~Ib.f .Q:~ Section t- - -/6 -24 -32 .8 1.6 -.4 -.8 3.6 3.2 2.8 2.4 2.0 -1.2 ~1.2 -1.6 -2.0 ",-
~ ~ 3 c:: ~
.....
~ "- ~ :2.4 ~
.I -, -:./ -.4 -.5 } ~
J-.2 .... c: c: ~
.~ ;g "- ..... ~
I
M "d 1-3 ()O I>:) ... ~ >- ..., 0-; 0 Z >- t" >- t:::I <: >-< U1 0 l:d ~ 0 0 ~ ~ >-< 1-3 1-3 l".J l".J "'J 0 l:d >- l".J l:d 0 Z >- q 1-3 ..... 0 U1
!:d 0 ~ Z 9
I-' 00 ~ I
Z » 0 » en 0l=Io N 0 (X)
1.8 < /.2 k' 1-1- /bV I I 60" .1/ I I I .8 ~ 1/ v:: I :v- ....... c, / .032 .OJ6 .028 )1';..< deflected , 1.0 .4 I /.
W I rouqhness rouqhness
'v' .","
, y/c , -.005 -.007 j\ \ coefficient, -l- _!
s.olit flop 0 .8 position "'" ~ ~~ I I lif!
x!c a.C. 257 .257 Standard Standard _~ -".. kro --r-.
!:> \ \ .6 "- -.4 simulated "'-......
_ Section
"'~ I'h
;cfc R r 9.0 6.0 a20c
JO'm=f.?56·rOO5 6.0 6.0 6.0
<">i"- " .4
o o l> 'V V
o
-.8 24·inch chord.
.2 -/,2 section, l-,-::' 7·
o
.2 -/.6 -j -.2 -.2 -.3 -.4 -.5 ti
-
~o .004
.024 l
~ <b () \) ~ ~
..... ~ 2081litfoil ~ ..... ] "-.
\J (\) g- ()
",...,.020 (;
..... ,~ t.0/6 -6.012 ~ ~.008
64, 32 ACA tile N of characlc·l'istics I I v '
,<
!
deq (Vi-\.
0. A ~~ ~ (Xoo ~ tll, l.ci 'f" ~ Aerodynamic ~a If!
l-d If ~ attack, I~ ~ ~ of , / ry .Q
II 'If ~
anqle ;c II tP ' -8 / ~ !lr)., 1/ "q 'VI I¥ 10.
Sedion -16 -24 .8 -32 .4 -.8 3.6 1.6 -.4 3.2 2.0 2.8 2.4 -1.2 -1.6 -2.0 t:
&3 QJ 8 () (\)
1.3' V')
...::1.2 ;g "-. ~ <::::: :g
.!
-.1 -.3 -.4 -.5 ~ ~
':-.2 t5 8 ~ !2
..... .~ «:: 'Qj ~
.....
..
r:Jl q ts: ~ o I:z;l ~ ~ 1-1 t< t;:j ~
>
..... 00 ~
•
z 0 CD
» 0 » en -1=10 N
1.6 ! !
I ') } / II l{L ;> 1.2 II V P' ;y 'f I ~ I/ / 60° l.I / .8 ty Iyd~ 1/ z .,l- C .036 .032 .028 V ~ deflected I 1.0 I .4 If I ~~ roughness
flap ~Ju~h~els
, ,~ ~ -.029 -.041 -QOI5
1\
coefficient, splif 0 Aosifion .8 ~ r:<: I I -.l lift [tI... lei
*$""
a.c.
St~nJa)d
Standard .259 Q- 1.26/ I, ~ ~b \ .6 simulated -.4 ti ~ ?-.. f"-. I I Section I i I ;rjc 1\ (
n-lJ.-r
6.0 3.0x/06_rJ258 9.0 6.0 D.2oc 6.0 !Qh .4 o o " V 6.0
<> 6
I -.8 , .2 -1.2 r- I) ~,--- ~ ) I
o
.2 -/.6 airfoil section, 24-inch chord.
-.
-.2 -.3 -.4 -.5 -.2 ci
.. -
~o .024 .012 .004 (,) QJ \.)
~ o ~ ~ ..... '0 ~ ..... ~ ~
~"!j.020 c: <U o o g- t:: \)
.~ Je
.... .!2 :;:.0/6 :g ~.008
64-209 NACA the of ~ ~ ; I ~ I ~ de9 jill ~ \)7 .,j I I I<!> «0) I~ II I~ !R Aerodynamic characteristics ~7~ ~ fV J ~ attack, ( I~ ~ J J or ~ J koLn J A j ~ angle Iv~ A h -8 "'i}~ ~ III ~I I.
1.II[IJ,.
·IU Section -16 -24 r ~ , r , , ~ oj 4 ~
cJ 8 4 o
.4 -32 1.6 -.
3.6 3.2 2.8 2.4 2.0 -.4 ~/.2 -1.2 -1.6 -2.
~ t:: \.)
rS Cb
..... ~ :g It)
;t i8· :.::: .~ ,I -.3 -.4 -./ -.5 It.
~ \)
l-'2 t:: " o t:: ~
.~ .\,1 ~
.... :::: .....
I
t; H [f). ~ >1 ~ ~ H >-3 >-3 1':1 ~ ~ >- 1':1 ~ 0 2: >- c:l >-3 H 0 rJ).
~ 1':1 "d 0 ~ >'3 2: ? 00 I>:> .,. 2: >- >'3 H 0 Z >- t"' >- -< 0 0 0 1':1 0
joj::..
I
z » 0 » 0) I\)
1.6 b / ~ I.{ b I 1.2 ,J !IJ ~ I I I I 60· II FJV iI1V .8 )
p I/f-<' I
l/ h- c, V ~ .U-.Jf.J .03Z .ozo ~~ deflected tr~ l !.O .4 I I--< I roughness rouqhness flop , y/c -.016 -.01/ , 1\ +- coeffiCient, split .8 _ positIon M: ~th 11ft I'-" rg oXic a.c.
Standard Standard .259 .258 "'\ I-n ated
- 0.259£016
F:::-t... .x 1\ 1 I \ .6 -.4 p: R 1 L I I simu Section I ~ T T .ric t-..
R
3.0xIO· 6.0,l Q20c 6.0 6.f
r\ t\. l"t I .4 V 06.0 'V o 090 L> -.8
. i I
'-r\ .2 -/'2 !-
,,-
o -/.6
.z
-,/ -.2 -.3 -.5 airfoil section, 24-inch chord.
-.4 -.2 .; s· .024 .0/6 ~ ~o .012 .004
<J £ \) ~
.... ~
..... ~ ~ .f: ~ \) (,)'tl.020 ~ Q.> 8 g> c: ~
..... :~ :::: is JIl.ooa
64-2lO NACA I I i 1 , the ! , I I ) I~ I \ ! I }., ~ ~ 111 be, fi deq Id !o£ ~ «0' 10 ~A 1. , I... II; !:f I~ Acrodynamic charactcristics of -'-- )
1..< '"
1..-< M / 10'" I~ t>'" attack, I~ ~ I,..
/, of IV ~ If /, 11/ J J.
I II J angle 1# -8 IA VI lIP\;> I'¥"'& i
dl
o ..110 III Section -16 -24 .8 .4 -JZ -.8 3.2 1.6 -.4 3.6 2.8 2.4 2.0 ~1.2 -1.2 -1.6 -20
<J~ ~ ~ u c: \)
..... ~ "- .... "- :::: f. ~
.I -./ -.3 -,5 -.2 -.4 .". ;r ~
J
~ 8 t t:
..... :\3 -i:: 'Qj ..... ~
~ w q :s: :s:" ;.. l:O o "i ;.. >-< l:O "1 o >-< t"' t1 ~ ;..
!-' 00 <:.II I
Z l> o l> en .J::ao - o -.. I\)
! I 1.6
--
;> ;:J »
1 r I
I I i\:Y ,~
/ P /./
- ~ /t5 1.2 Y W ~ / ~ I 60 / V- I I , .8 I
V v
:;: I"'" c, V L,c In r< .036 .032 .028
V ~ deflecled
1.0 .4
I rtf
I I ~ ~ roughness roughness
\: ~~r
ylc
, .017
\ 0.029 -.OOC coeffiCient, split flap
a
t
.8 position f..-- lift I-- ~c ac. Standard .259 Standard ~"- 1'i 1.262 1
-
-
J
.6 ...
- -A f.n.'r k-,L 1 1 simulated '( Section _ "LI.. ........
;rIc ~ 1'( In. h R D.20c 8.0
3.0iT-1:256 6.0 9.0 8.0
" .......
1\ i'. I"' \ .4
)., o o o 66.0" \1
'Q.. J.... "'- --:8
,c, '\ ......
'"\ \
I"
.2 c
-1.2 \ "\
r-- i?l
~
I"-- 7
o
-1.6 .2 -.1 -.4 -.2 --:5 -.2 airfoil section, 24·inch chord.
" <i
020 ~ .024 .016 .0/2 .004
~O C ~ B ~ ~
~ ..... :~ :;:: .....
.....
\:: III 8 g- c: II
"' ·92 .\j i::
-1--.""- :t: i3 .0 ~.008
G4r·Ql2 .32 ACA N the I h f-f- ~Q.I /6 ~-> VB,.
~ Ir;.~ I'
rv
deg I I 1 ~ K.
£)(O!
!-t:' I !J ~ Aerodynamic characteristics of ~ Wf"
"
'1 l£i ~ attock, !!
~ of
I n
~ ~ J c- 11 "S1.
angle .
~J III il -8 I ~ L1l h - \ , Section K.
fuld \ -/0 ~ I
I -24
r-
;- ,- r-
r- .8 .4 -32 1.6 -.8 3.0 3.2 2.8 2.4 2.0 -.4 -/.2 -I.B -20
~ ~ Q) 8 ~
'" .U 'i::: '<-: ~ :.::::: ~
~/.c :;:: ./ -./ -.4 -;5 -;.3 ~
-
l-.2 ~ \) " 1::
~
..... ~ 1
"'Qi ~ ...... >- t" E; ...... U2 0 J:lj ><l ~ 0 ~ ~ ...... ..., >-3 M t:j "1 0 J:lj >-- M J:lj 0 Z >- c::: ..., ...... ~ U2
1:il "d 0 ~ Z ? 00 .,., "'" * 0 Z -<
~
""""
I
» en ~ .... .... N
Z » 0 -
/.6
'-1
Ln II [.1j .,., 10
J ~ 1(' II ~ J.
I .L 'I 1.2 C ,1: ¢. / ~ / ~ ~ U I / 60 11 I .B P- £. In/" Pol I I l c c, j:;J:/l I~ .0.16 .03Z .QZ8 / ~ I~ deflected 1.0 .4 11,1 I 7 I: h.~ I roughness ~oygtn~'F flop '~ , ylc I -
, -.039
, -0.013 ,I coefficient,
o
.8 split -l- +-.01 position I c
r-- ~ lift
;a?dfrf' .rIc _~ o.
Standard 5
l2!
r-... 1.265 1.267 ! I 1
[U. I I .6 -.4 rC~ 10. I I I I simulated Section \. r. !'-J.
I I I I ~I ~/c I"- 1\ N
f
R \ 3.0x/O· 020c 6.0 6 P... 6u fT .4 V l
" o 060 09.0 ,:, '::7
" -.8
"'r !1 :--...
\ ~ 1<': Ki I\' .2 <1\ -1.2 1'1' f0- ..........
.--
0 ! ,
°
.z -1.6 airfoil section, 24-inch chord.
-. -.3 -.2 -.4 -.5 -,2.
EO " .016 .0/2 ~o .024 .004
\,) ~ Cl \) ~ ~
..... ~ ..... ~ Q;
1]:020 ~ ~ 8 0- t \J
:~ :::: ~
-<-:' :2 J1j.OOB
64i-112 N ACA the '- FJ ~ "i7 Lc ~~ I
~ ~ II II r: 10 !;7 1"-
\ ia if \ deq
" :" ~
1I \ u o, 1I1 III i!
~ J I ~ Aerodynamie characteristics of d.
I I~ It II £( '(!
attaclr, .& ~ .,.
o
f) of
"
~
"
(f N~ I~ ~ /1 (i{ Iv.
I angle III . -8
IJ 4""
if 'v.
i
VI ~
,v
1'1 I ,I M Section '(1[1' ;h l q -/6 -24 ! , , -32 .8 1.6 -.4 -.8 3.6 3.2 28 2.4 20 -1.2 -/.6 ~1.2 -20 \.>--- I..> \.) ~
ij3 ~ S
.... 'l:: '::: =<::.4 If) :<::i ~ ....
.I -./ ~3 -.4 -.5 :t.
~ <::i
l-.z Cl \.) t ~
...... .~ . .;::: 'Q:j ..... Q ~
~ ~ ~ > ~ ~ > ...... ;J o ...... t< tl ~
I-' 00 "
I -1=10
z » o » en - N .... N
- 1.6
I(-
1, .1)1 IA:
lr-. IT V
J II l;Z' I 1.2 -( / 1<1 II I ~ II V I 60· V I f .8 ;:.
p I
11-1---
V I.;: c, loOk? ~ .0036 .032 Jt( I, deflected 1.0.
.4 I :HtE; I roughness flap
'~
y/c IK: I \ -.024 -.024 \ 1-- .1 coefficient, .8 -l- ~OOI3 split position I ~ r-... lift t-- .x:jc a.G.
Standard .262 .262 f'..t I::.rll !
0.262 1.1.
1"'-,.
~ .6 \ -.4
r-r k
1_ simulated
~;a,:d;,r;:1'~=()~~~1n~S~
Section
1\ k'< I I I
T I I
7~
J
'-< .xjc I'~ 1\ R 3.0X/06 6.0 020c 6.0 6.0 l<:>. b.. p.., .4
o o 09.0 /). "V V'rf
-.8 r--, IQ ~ ~ \ 1\ .2 -1.2 I--
r-
-
')
o
-1.6 airfoil section, 24-inch chord.
.2 -./ -.5 -.2 -.2 -.4 ti ci .024 .012 .004 ~O .016
r.,.,fE ~ Q] Cl 1.>,.3 ~ 8
...... ~
-!-!' ~ "-
~ \J
;:.020 1:;. III B g-
il ..:: ~.008
r ::: 641-212 NACA the I I !
R.. b-
\ "f- ~ \ I 1/ I Cl '\C
&:l II M
~'h ~ \ \ z::.
~;>. I~ 1\ deq '\
""
~ il"() 'off /l~ ~ 0(0' "tl: Lv- J J Iv Aerodynamic characteristics of L y
r
/} 11!'
l attock, I) "'- I~
o
I .4 of Iii 'I .1 p V-<;;<..
II) r; /) .
P
~I ~ I. on91e II 19 -8 II ~ ilfr-..
~ II II ~ I ;'a I tJ:: I ''l, ILJ K Section r .
InJ -16 .
-24 I -32 .8 .4 -.8 1.6 -.4 3.6 3.2 2.8 2.4 2.0 -1.2 ,1.2 -1.6 -2.0 I.J \J (J- ~ t:::
...... .~ ~ "" ;g ~
"-- <::: .I .
-.5 -.1 ~3 - It.
'
./-.2 ~ Q] 8 c: Qj ~
..... .i;! ~ ...... ~
I
I-! ~ I-! 0-3 0-3 t"l t"l "%J 0 pj >- t"l pj 0 Z >- q 0-3 I-! (":) Ul
t"l I'd pj 0-3 Z ? 00 b:> ll>- z >- 0-3 I-! 0 Z >- t-< >- I::' < Ul 0 ~ (":) 0 ~
pj 0 t-' 00 00 I
,J::Io ,J::Io ... N
Z l> 0 l> en
r- /.6 'I .-- 1;1, I IV II ).. Ipl/ fd I /,2 I~ II Ii-' I V ~ c Irl 60· / ~fN I .8 ~ Bi
I
.036 .0.32 028 ~ V I.J: 1.0.
c, ILY
, I
deflected \ I .4 \ a- pPr-'" \ , ,
- ~
roughness rouqhness flap .8 ylc -.046 -.034 ::::x "'(:~
-
-0053 coefficient, --I- split
posdion =t
......
-
~t- lift '0 I'., oXic a.c.
.6 .266 .266 Standard .267
'\. r--- 15tand~rd
~ U. ~ .zoIc -.4 I: f'., r----, I I I simulated I I Section
P- ~ I I
~ .4 '\.
R "'~
< 3.0xI0· 6.0 020c 6.0
(\. ~6.f
06.0 090· 6. 'V o -.8 , .A .2
---
-/,2
V r--
')
o
.2 -./ airfoil section, 24,inch chord.
-.3 -.4 -,5 -.2 -.2 ti i ~ ~o .024 .016 .004 tJ ~ \)
~ c: ~ ()
.... ~ .... ~
-
QJ 8 g' c: \J
",,,.020 ~
.... :t i3.012 :g ~.008
'·G G4,-412 - I- - NACA --- ._- _.- ~ -- -- 24 W .- \ r?-?
-- .J..
,..- d\ ~ ~b -- -- 16
"-
-p >.... .~ .- ...
I , I , , I~ ~ ~ - deq b
Ln t:. iF
- ._'-- if' if{ lXo, ~ jl "b _.
(V' Aerodynamic characterislics of the I~ II ~ 1-.
.
V I~ ottocfr, F 1.,( I ~
I y of
I
I k?
II Iv- // anqle If I~ '{ ) b "'I'Q -8 I If ' I ~V
, ,"J
I
I'L' R ~
[ Section f II -16 (I-' -24 .8 .4 -32 -.8 /.6 -.4 ZO ..i6 3.2 28 24 -1.2 -1.6 d.2 -2.0 ~ \.J c:: ~' c: ~
.S! .~ :::: ~ :g V)
·.. ;t.
.I -./ ~3 -.4 -.5 :!t
/-.2 \) c:: ~
~ Q)
~~ ..... ~
:g ;:: .()
q ~ ~ ~ >-1 o .." :> ..... !:d .." o ..... t" tl :> >-'3 :>
Ul i---< 00 ~ I
z -1=10
» (') » en I\) 0 .... U1
1.6
b f5>
')
Ld
hi-, '-
/0
IT 'I"i / .
1.2 1.6 ~,.
f I I j' ,.;::: 1/ I I ~ool ;- f?
Ii ) .8 L d ?-' .Jif:V I l c I/" I .03Z .036 .028 deflrc~e1
I};2,V ""
1.0 .4 .,- l..< ~ I I I
~ 'I roughness roUrhM~.s1
(Iof I , y/c I -.006 ,-.007 , -0.007
r-- Ie-"
I split coefficient,
position o
.8
f
I I-- ~ a.c. .r/C I lift Standard Standard ,267 1- 0.268 I .268 I I I
r-- "-
- ~
I .6 I simulated -.4 ~ n- I.r, 6 I I Section I'" r-..;~ I
.xlc '"
f\ I R
\ ~g
3.0x10 6.0 9.0 6.01. o.20c
'" ~~
\ .4 o D <> ; t:; '\.
R " -.8
\ ~ \ <~ \ ~ 1\ r""' \ .2 IQ l\ I\. f1-I -12 ~
..... r-- 1\
'-
,~ , 0 I !
, 7 , ¢
o
airfoil section, 24-inch chord.
.2 -1.6 -.3 -.4 -.5 -,2 -.2 .; .ci ~o .024 .016 .012 .008 .004
"" QJ Q () c: 'll
'\:? 12
-...:' ,~ 'i-- ~
:t.
~ \.l ",,".020 "-~ c: (J ~
:Q :;:: ~
is <J5
64.-{)is NACA the r<.
1'-' r'Q \..
~ ~ ~ lo;Jt ~ ""I !\ deg .l.'f \.
~ ilL p.~ v «0' I/' I~ l!!.
I ~ I Aerodynamic characteristics of " VI 1 ~ g IL ~ attack, ill 1,,( 11', of' IIJ I)" K7,.
~ Ii P V !.
""<;it & If.I /, angle III V M f) /I. -8 \-z (I 'I J 1I' f Yo rf II
, Section
Wi
:;.. 'h. -16
-24 i ! , .8 .4 • -;Je 1.6 -.4 -.8 3.6 3.2 e.8 2.4 e.0 -12 -1.6
-ao
~ Q) c:: (J Q) ::5
~/.2 .§ .~ ~ <::: '.::: <0
:g
.I -.1 -" -.4 -.5 ~ 'll'..J o l-.z c: I.l ~ E:
'I--~ .~ ~
~ ~
I
t"l Z t..:> '" >-< >-< ~ ~ >-< tr1 tr1 >-<
eo 0 ::0 "d 0 ::0 '" ? 00 li>- z .... 0 Z .... t'" .... Sl U2 0 ::0 >-1 ("} 0 ..., ...., tr1 >rj 0 ::0 .... ::0 0 Z .... G ,..., ("} U2 ......
- I
Z » 0 » en ~ N I\) .... U1
I 1.6' M
IL
:r 10
r I
II I II P I 1.2 I I fT VII' I I I f{ l):'i7' IN / II 60°, hi I~ to I .8 ed rrt p- V l/ / r& c, A .032 .036 .028 deflec
v
1.0 .4 1 roughness ~ flap
\1
["'-' -.038 -.014 \.
-0.045 !--- split coefficient, position .8 ....
"11'11111111 lift
- ~ A: *r""
ac. J-fonJ.a'!'dlrc\u~h),e1sl
Standard .266 .265 ~ I T
-
" Ii
.6 simulated -,4 i\. ..., :"tJ, I ,II Section 1\ i'-. t'1: I I I oc/c \ r-...
R
~g
\ "'( 3.0,14f-:-(U6'7 6.0 9.0 6.0 '0.20c .;:.
1\ lilJ]]
\ .4 \ o o <> 6. ; -.8 1* IQ.
I\.
\ '\1'-' t; \ ~ ~ .2 -1.2
r- I-'-
I
/- "'-
')
o
.z -/.6 airfoil section, 24-inch chord.
-./ -.2 -.3 ~.4 -.5 -.2 ti ~.
020 " .012 .008 .004
~O .024 " (J
~ \I) o ~ ~
~ <t..: ~
" ..... .....
<J. ~ \I) o o g- ~ o
.~ {, ..... .0 '2:..016' ~ ~ 64:-215 NACA the I I I I , i I I . , I I I j ~
J '1 ~
~ r\ 4"
~ :Id P" \ ~ r'--'J ~
v
~ ~ :F deg },"fJ \ ~ r> l. ...
~J Ib. 41 "0, --
~Pl\ "'-
I[ IIA Aerodynamit' characteristics of I) d 'I h ~ i"'" altac/<;) D.
VI ii' V' -- f) V" VI IJJ of !J VI It<:l I.
/1 d IF ';;o.kz.
L-_ J // angle '/ If Ih ,- !J -8 W' III H fl: VI ~ .A 'n: :/ ~ ~ L- Section ) /O- -16 1\ L _i- -24 --', .8 .4 -32 1.6 -.8 3.6 -.4 .3.2 28 2.4 2.0 ~1.2 ~ -1.2 -1.6
... ~ \I) (J -2.0
..... :Q ~ <!::: "" ~ ~
./ -.1 -.3 -.5 -.4 ~ " 1-.2 ~ \I) o o t:: ~
... ~ <t..: ~
....
d ~ ~ > ~ o "':1 > I-< P::l "':1 o I-< t"' tj > :;
(fJ - <:0 -
I
z » (') » en ,JlI. J\) ,JlI. .... U1
1.6 Ir' ~ IL
r.L
I?
V II 1//;: I /,2 J. ;.
// , I 1/ - Id lh: / /
W
r 6Q"
" 1/ 1/1 I I 1 .8 It! W / 1/ 1/ c, -"'~ .03Z .0.16 .OZO deflected 1.0 .4 A: rouqhne~s flap roughness II:LI ~ \ ylc <r: ~051 -.040 \ -""
==- b
coefficient,
flit o
position --l- ~O.O70 .8
I"'- s III
C lift
r-- r- r< :tIc I
a.
Standard Standard .265 .264. ated i" t-,.,['< I 0.264 I ~ ~
>-- I~ I
- ~
1. 1.
.6 -.4 Si'l"1.
1\ 1"'- f' I I / / /
\ section /1::<. / / I L 1\ ~ x/O .rIc I
1,\ I"
R
3 6.0 g{J 6.0! ggc;c
"'
['.;1'. IT/
.4 [76:0 o \l -.8 o <) .6.
1\ r-,: chotd.
\ l\ r\.
.2 -,/.2 ~
V r--
0 r ) I
o
-1.6 airfoil section, 24-inch .Z -. -:3 -:5 -:Z -.4 -.2
E~
~o .024 .016 .004 I; ~ ~
v ~ ~
...,." :Q :::: ..... ~
~"!J.020 ~ q, e 8' ~ u
...... . .\J {s.OI2 :;::: :::: ~.008 64z-415 ACA N the of - I- --I f0- '- 1--1 :J- Cl
e-
~ :L \
"':t r""
Iv r'.
b.
n r'-' ~ 1'1' fz.
ill, 1\ /Y deg ;.I'{;Y" 1;;1 1\ "0,
'':1-' rr
I}"
(J ~ Aerodynamic characteristics '!J ~ Id : It ~ attack, 1:J :1 l~ 0 /I of ~ i ~ "'h- U angle A I l It -8 ~ II J Jl I y .lI.
III
Section
"
"
-16 \.
e -24 " ,.....
r i i r f) ~ o 4 y -32 '::J 4 tJ A .8 -.4 1.6 -.8 2.8 -1.2 -1.6 .l6 3.2 Z.4 2.0 -2.
..,- QJ 8 8 hl
~~ .;:: ..... .... \I)
"'f."1.2 .... :.::: :;::: .!
-.5 -./ -.3 -.4 ;:t
8 c: ~
/-.2
..... ~ 4i £
~~ ....
I
l:l:I t;j "d >-3 Z 00 I>:> "'" z >- >-3 >-
~ ~ t.:) 0 l:l:I ? .... 0 Z t" >- t:=' <: .... Ul 0 l:l:I >-<1 (l 0 ~ ~ .... >-3 >-3 M t;j "'1 0 l:l:I >- t;j l:l:I 0 Z >- Cl >-3 ....
(l Ul I ~ ~
Z ~ 0 en (.,) 0 ... (X)
, /.6 V ~
11 lL2
/ /,2 / / P J / I/V If 60· ..dV 7" E .8 v: d ~ ~.
c, I~ ~ .03' .tJ3Z .028 IL' ~ deflected 1.0 .
.4 ~ :
"
.y rouqhness nap rOU9hness \1 y/c .054 , -.044 --- ..- .
coet'ficient,
split o
.8 position
-I- :::f~053
Ie
r-- ~
lift ....
ac.
Standord .268 standard .266 "i: I
r-- ~ ~
.6 -.4 :D t6a simulated
-
---.
SecTion ~
+-+-tJ
ric f\... ~ IY.
R 3.0xI0' 6.0 9.~ 6.0 0.20c 6.0 1\ .4 /I P
L\ o <> .6 '76.0
o -.8
'"
I \ ~ '\: .\ I'\.; .2 '0' -1.2 , j..-- t-- Kq .
I J
,V
~ /1'-:-. I , ?O 1 ) '/
o
-1.6 -. -. airfoil section, 2{·inch chord.
~3
-.2 -.4 ~5 <i ~o .024 .016 .012 .008 .004 I'
"
5 QJ B ~ ~
' ~ ,,' & ..... .....
r..'tl.020 c:: <lj' q, 3 8" t \J
:\j .;:: '-
.... {) :g .3i
64,-018 N ACA the
:, ~
[IT ~ ~
1\
I I A I', "F b
~ ~ ~IIIIIII
I "~ If1 I 1,0.- d"'q 1\ 11; lIMr'l'¥-,
l? WSM&
.~
I~ Lm «"
rJ' II
Aerodynamic characteristics of
II I 17
~
,I
II ~ I h: ottack, 7) LV I
D
of
w I 1/ rv'
j
Iff I II
1.
II m.,
r7I "f I
anqle
7J
1/1 ~ -8
fh 14l
f\t if/
Ii
:l
III I", IlJ[ Section
I
-16
~ f1q-'
ILQ
Ih ~
~IIIU IQ
-24 1 t f i I !
4 0 -'1~ .8
1.6 -.4i
-.8 3.6 3.2 28 Z.4 2.0 -1.2 -/.6 -2.0
~N \:: \j I!J 8 \J
,,~1.2 r-
.~ ..... .... ..... :.::: .<5 .;::. ~
.~ .1 -.1 -.2 -.3 -,4 -.5 ~
J c:: 0 \J ~
,,~ ~ 'Qj ~
.S! ....
~ ~ :> ~ o '"l :> >-< ::0 '"l o ..... t"' t::; ~ :> \;0 C/.:I
U1 q -
I
z c» ~ I\) (X)
» 0 » W ....
L6 I i I I c
/
I IV k' I 1.2 II V~ I ~V I~ ''1 ~ I 60° ..fL tf/ .8 /" I.(J
Ire
1/ S¥ p Cl ~ .036 .032 .0<'8 defleded 1.0 .4 I-'" , I I I I roughness rouqhness flap , y/c ,
, -.021 -.053
r-- ~
coefficient, split ~ .8 position c.
lift - - .......
;r/c a. Standard Standard .Z69 .271 '\:.h I .
0.267
"'"
r-- ~
.6 -.4 '\ ¢ -...;;;~ simulated I I
- Section
) 'C ~ I ~ \ ;rIc '\ " .......
R \ 6.0 3.0x/O· 6.0 0.20c 6.0 6.0 1\ lv..' b..
\ .4 17
o o 09.0 L'. \l
."-h -.8 Ii 1\ q ';;'\ ~ \\
.\ '"
- .2 -1.2 I.l, 1\ ~
- --
i'-
0 t 7 , I ~ ,
,V ) t
o airfoil section, 24-inch chord.
.~~ -/.6
.2 -
-.4 -.2
"
,004 ~o .012 .008 ~
.024 l ~-.2 () ~ §
.Q) -+-~ :~ ;:::: -+- "<.
-
III 8 (J
<.J",.020 s.:: 15'
.~ .~ ~.016 t ~ ~
..... 64,..218 .]2 -- NACA - - the of :J ,v
Itt
~ l,~~ Ib ~
r- ~ "'if
::0 R:J IYI be po 16·
\ r
0- lo Qc ...,~ ;a ~ ;I" deq '( /; PI 1\ D.
(xo, ~ 11 ~ J Aerodynamic characteristics
w ~ V
iJ II u 1<7.
n II "'- .1J aftock., '& 7J IY 11 IJ "'19- of I~ ~I I, ~ (/ > I 'I _.
'"
IP .
t
angle Ii I.f ~I - -8 ~ (f' 1/1 'If I.
" II
III !f I I\. l{j ld Section
"
rl I -16 b \ I\, I"\n, \ q -24 , I I I I I I I I ! I I I I I I !
i I I , I -.]2 .8 1.6 -.4 '-.8 3.6 3.2 28 2.4 2.0 -1.2 -1.6 -20
~ III 8 c: u
~ "-= <t: ~ ;g.4 ~
-+-~/.2 .§
.J -./ -.3 -.4 -.5 ~ ~
s.:: 8
l-.z ~ s:
..... .~ <i::: Q3 ..... ~
I
to;l .... z ;..- >'l Z t' ;..- t::j -<j ..... U2 ~ o-<i ~ ~ ..... >-l >-l M M '"=1 ~ ;..- M ~ 0 ~ ;.- r. >-l to< " (jl
l:d '"d 0 ~ Z ? 00 ." ..... 0 ;..- 0 (l 0 C
I-' ~
I
Z :J> 0 :J> en ~ (,.) ~ .... (X)
1.6 / ).
I J /
VI
j 1.2
r
rr
/ I I.
V.
II E ~ 117 rJ I I
V iP
.8 /' A
...-
I t c ~ 1 .u~o .032 Ideflecfed60" 1.0. .4 L'I I I ......", flap roughness L \~ ...
, -.049 --.050 ,
-- n . -0.05/
split 0 coefft,-ient, .8 position V'~ r-- I-- r.
lift
'/c$
Q.C Stondardrouqhness Standard '\. S::~ 0.267 127/ 1.273 IL_LIL f..-- R. b...
r--
.6 -.4 1\ ~ simulated Section i\ '", 1 1 I \ \ .rIc \ N 0 1 0 1 I R
, 3.0xIO· 9. 6.0 D.20c
6.0 6.0 0- I .4 o 06. <) "" 17 'J' ~ -.8 I \.'a-,
AU '"\"
1\ 1\ .2 -1.2 ~ I--
/' '-
o
2 airfoil section, 24-inch chord.
-.I -/.6 -.2 -:3 -.5 ti ~ ~ .016
~o .024 .012 .004 J
~-.2 Qj () (J ~ ~
~-.4 .... ;g '" .,..
ey,
<.:".020 c:: 8 C) c:: u
.~ ." ::: -b -;::: ~.008
...:- .0 A fl'h-418 C A N the P ;{H u \'i\.
--~ :;> f.'> JI Y> '\ I i vc.~ ..p deg .
b tyl :t '-nf' I 12: If) ~t1 «., ~I {, Aerodynamic charaPleristic, of lOt II I IQ l// 1.1 attock, li' ~ f-., ,.( il of II y ~ , I)
"
~/ J. t.
\ at7Cjle ill 'I 11 !J !
-8
."
II
n
1/ t::::t ~ IP t f III I\, If-' ~rL ~ Section J V 1\ I:h- \. -16 -24 -32 /.6 -.8 3.6 -.4 3.Z 2.8 Z.4 2.0 -1.6 -I.Z -I.Z -20 ~- ~.8 v I::
~
..... .~ .~ ~ ..... ~ :g.4
.I -./ -2 -.3 -.4 -.5 ~ ~
,J ~ a ~ ~
~ ..... .~ ..... .;::: 1; ....
U2 q Is: ~ ~ o to;J ~ ~ o ..... t"' ~
J--l ~ ~ I
z ~ ~ Co)
en en - Q)
I I 1.6 ~ j /'P I/- lo.
Ij, 1.2
lr; 'V
[ \ L IL t>'; il ~ 60·
-1-1
.8 V -l.
P c, II .036 .03Z .028 Ii':!
deflected I 1.0 .4 I'" I I I I:::::" roughness ~ flop roughness1 I \ y/c \ .005 I -.019 0.014
t-- 1', ;'\
I split coefficient,
o
.8 'N~ position
'Y\ 1
I
:-- Y'I'Y lift
r- rc
.z/c I a.c.
Standard .215 Standard .213
""
~ 0 i"'" I .....
~ t-- tJ0.289 .6 -.4
"
f:'-- I"" I simulated I
-
- Section Ll "'" f:::..
I Z/C
-'"
l'.. I" r:t::: g
I '1"1 R \
3'~YT 6.0 9.0 6.0.! o.20c
It': I .
.4
.\ 'I~ o o <> C> ~ ~
_\ -.8 In.
, - .s: , In. ilj ~ .2 ·/.2 ~ V
V 1'---
f)
o
.2 -1.6 -.2 -.1 c3 -.4 -:5 ti airfoil section, 24·inch chord.
020 .016 .004 ~o .024 .0/2 '(J08
l ~ \j ~ ~
~
... - l-.2 .,., ..... ~
,.. c: (lJ e 8' § I.J
.!!!
.... .V :t {) .;:::: .J!
643-618 N AOA the ~ ,'t
:L .1
b't I" >0, -6 ~ PY--t< ?o .0: \oN ~ (':t I --'--- II I-Q ra:
~
5.~ deg .(/''t; Z.- ~ - Io<ll '\ I l j if' PI'.{ et'o, t:!,
"
~ lP~ V kx '--- Ii' IR It, I'r ,-- Aerodynamic characteris.tics of II A n.
attock) II Ij3 .....,
JJ 'Z "'" of
II Ii
ill ilf
~ A III ~ IfiJ ~ f) onqle V[ '( rt7 /..
-8 /I J if Ic7 II JJI 1/ A IL Ii VI 19- 14 Section III III f:7 'q. ~la -16 -24 -32 .8 1.0 -.4 -.8 3.2 ·3.6 2.8 2.4 2.0 -1.2 -1.6 -2.0
'" ~
<tJ 8 C-
"",1.2 ;g '>- ..... :2.4 M
~ .1 -./ -.3 -.4 -.5 ~
-
/-.2 () v t ~
..... .~ .\J .;:: 'Qj ~
.....
I
I'd ~ :>- 0-3 ...... t"' :>- t1 H !;d >-<: ~ ~ H >-3 0-3 t<"l t<"l "'l 0 !;d :>- t<"l :>- 0-3 ......
!;d t<"l 0 !;d 0-3 Z ? 00 li>- z 0 Z :>- -<l 'l1 0 (") 0 !;d 0 Z c:j (") Cf2
I-' 1:0 Q:I I
Z l> l> en ~ .j:Io 0 I\) -
1.6 J -'l J..
If ~ 1.2 """I:~ ,1, / II ....
r+- ? r I I /.
I ~ SO' II I U I /~ .8 II I/' I l~ V; c,
V I)V ~
.036 .032 .028 ,......
deflected I~ -Lr, 1.0 .
.4
v
I
. I
~ , I roughness roughness flop , ylc -.044 -.077
" aOl6
-- r::---
. coefflcienf, sflif ~ 0 .8
---
r-- II II lift
:ric Standard a.c. position Standard ,-..
.270 .273 .274 I'-. I I
r-- I-- I I
.6 "Yl simulated -.4
- I
-
\ Section ~ -'-' I I I .rIc \ ~ I R \
6.0. D.20c ~~
.3.0xI0· 6.0 'S;J.
I .4
o 090 ~
o Ll.
-.8 '1\.\ 1\1\ " "\ \ \ .2 6 -1.2
> \
I-- r--
V i'--- ')
o
airfoil sect.ion, 24·inch chord.
.2 -1.6 -./ -.2 -.3 -,2 -.4 -.5_ ti
-
.024 .012 .004
~O .l ~ ~ () ~ §
.... ~ ~
-....
(,)".020 ~ <lJ 8 g- §
{) :,::
~.016 ~.008
--..-
64,-021 NACA the I ~ ~ '\~ 24
? ~~
:)OCrk> -"::' ~
r R.. rc
--
r--
"
',ll~ ,
¥liP :ru --r-' Ik;I
't:l ~ ~ IZ.- deq I \ ~ JI \ ,
I ~ ~
o oc ;'<) 'f ~~ ~ "( -I !l Aerodynamic characteristics of If If:!
, Ii l1/ ~ l~ ~ atfack, 1"1 II -vI'""'- I, E of 'If 'V IJ ~I Ii{ 'I //1 I ~ y ~ ~I ~ , lit:' I I, -8 U r\.., -. ~ .. lr:i II V.
IJf ., .
Sectionongle .n1 -/6
- 1M
['q.
\ \Q -24 .8 -32 I.S -.8 -.4 3.2 2.8 2.0 3.6 2.4 -1.2 -1.6 -20
IS" I\) 8 ~ ()
.~ .~ 0:,..: ~
-..:-1.2 ..-::: .... ~ :;::.4
.1 -./ -.2 --.3 -.4 --.5 ~
-
J o () ~ t:
.~ .() .;:: "'- ~
.... 'Q3
~ ~ ~ > l:>j ..-. o "1 > ...... l:>j "1 o ...... t< t::I ~ >
{fJ I-' ~ --t I
~ N
Z l> o l> <» .j:o. N ....
1.6 ! I I ...
. I t:> " /.2 II I i I I 1M [.0 I f:,.
II IJi I 60· I,d'! ¥i: ~ I J .8
1-0 lit
~ r;, i 1 I I V h- c{ .036 .032 .028 I deflected 1.0 .4 I I ~ roughness roughness flop J.
\11 lion ~c ~ -.003 Q -.008
t- k"'"""
split coefficient, .8 ~ 0 posi , I -.l c.
r-- I---
lift ;rIc o. Sfandard Sfondard .273 .27/ 0.275 'b .
r--- f.--
..
.6.
_ simulated -.4
I" l:\ m
I \ Section
-
~ :z:/c I'..f'I..
R \ D.20c 3.0.xI0· 6.1 9.0 6.0 6.0 6.0 IQ ht\
-
.4 \ V o t;. 'V
1\ ; o o .8
C ~ I\.lq
1\
1<: I.E .2 J..--
r--
V 1'---
I I , , 0
;1·
o
.2 -1.6 airfoil scction, 24-inch chord.
-.I -.2 -.3 -.5 -.2 -.4 o
"
~O .024 .012 .008 .004 ~
~ ~ \.l ~ ~
...:' :g '<
~ -,..
\3
r:;.020 ~ 8 is
.\b t,
-..!' 'iE.016 ~ ~
64,-221 NACA the , L 24
t- ;J :f
\ j.:!
::n
~ ~ '\
..... "'i7
~ )cI J;d.
\ I \ 1\ h{) po I.rl deq ~ ~ I" ;t; Ii I~ f<\ Irf ~~ «0, ;: J/f .: .1 !j 'r Aerodynamic characteristics of .j ~ fJ f :/ W If offael<.,
I.r II'
~ J1 of ~.
II !I 1:5' "A~ ~ 11/ f /} 'Ij III -8 1/ If ~
Ai'-
JJ
tL -.., -f
I III 'If '..-' lRJ Section angle /o! "'-
\
. U[ -16
jDCr-' -24 .
-32 .8 .4 I.B 3.6 3.2 -.4 -.8 2.8 2.4 2.0 01.2 -1.2 -1.6 -2.0 ..
~~ ~ \b 8 c: ()
.!!? .'-> <;::: "-- ~ ~ .'2 ...... ~
"i .I -.3 - 4 -./ -.5 :t \.l
./-.2 s:: o ~
~
·t·O .g ~ ..... ~
I
tzj ~ .... w ~ ~ ~ .... ...., ...., tzj >- t'l !:d 0 Z >- Cl ...., ..... w
!:d "d 0 !:d >"l 2: ? 00 ~ "'" z >- >"l .... 0 Z >- t"' <: 0 ('1 0 t'l "'l g ('1
.... (0
I ~ ~ N ....
z » 0 » en .,..
1.6 I i ! I !
1 I
1)"
1.2 12~
I7J rl
(
Ll l' T hllCl ~ 60· I .8 1,/ K ~:>' t c t-< !:Q .036 .03Z .028 I I
L...... t-< IQ deflected I I
/.0 .4 R:s I I ~ h- roughness roughness f-- flap \1 y/c '>- '>: , -.035. -.047 , I-
r- -0.017
coefficient, split 0 position .8 1'-- iN" I'" I~ lift
r- I--
a.c. iE/c Sfandard Standard .276 .
~ I" 0.277
-
~ I"" ~ t- .6 -.4 simulated )y, !\.
'I Secfion \.
~ r\ t-t-
_~.278
iE/C ~ ~ R 6.0 3.0xI0· 9.0 6.0 0.20c 6.0 6.0 K..
"- .4- V' o "V
o <> 6 -.8
IQ f')J; chord.
",In 24-inch .2 -/.2 c t--
1/"'"
0 , I~ I I :> ¢ ~
o 3
-1.6 .2 airfoil section, -.
-.2 -.2 -.5 ~ ~ .02 .0/6
~O .012 .004 l
~ 'll o °-. ~ ~
.... '0 ~-.4 ..::: <0..: ....
~.020 c: <Il 8 8' c: ()
.!,1
'i-.~ .9! :::: {, .Q ..... ~.008
64,-421 C A N A the
:.r,. ;r r-'
~::J .r.~ tPR ~ ~ ~ c>-<~ ~ ~ ~ ~ P1 ~ ;;. <l .,p 1 deq I ~ 5{ d"-~ ~ :t;ly o, ~ oe II ~ - I~ ~ ;f',.",- !'W1'7 Aerodynamic characteristics of ilJ[Jl IU {; /) ~ Y J!
attack, /J ~I I~
<r 0
J1 {f of
1& tr
J.
/, P -
fj ""Ii'-.
'aiM b..
I/J ~I /,!?
.d angle I-' tt II, '(/ J II -8 III ~ ~~ II III 1,,\11 Ie!
Ii
'"
~ -A.f Section I~ Y" -16 .J: '\ \ I ~ ~ -24 : , l I , !
o
-32
.8 .4
-.8 1.6 -.4 3.6 2.8 2.4 2.0 3.2 -1.6 -/.2 -2.0 ~ Q) t:: u
r;S' 8
.~ ~ <;.: ~
-+:1.2 ;t '-:::, ;g
o
.!
-.4 -./ -.3 -.5 "to
l-.2 ~ () o (,) ~ ~
I . ~
'I-.~ ~ ~ 'I-.
> "'! > ..... !:tj "'! o ..... t"' '=' ~ >
CfJ ~ ~ ~ !:tj ~ o ..... ::0 ::0
I
z » 0 » en U'I (,.) 0 - Q)
~ /.6 LZ
I, /1 !7 I-
/
rr "J p
II I)J i.>'-' .8
p- h'
c, QZ8 .034 .QJZ .J.
b' , 1.0· ,4 y/c.
J..1 -0.045 '-.~34
position ±
l~
-- --
coefficient, ;rIc a.c.
.8 0.262 j2~Z: lift 1-
r--- i--
1\1,-
-
i-- I~ .6 -- -.40 rw _- R Section ..
6.0 0.0xI0· ~ 1- ~o.lll ol'/c A
o o '?
.4 -.8 M chord.·
"
1\ "\ "\ \ 10 24-inch ['\t .2 -/'2 "1--1"-'
-- --
~ - ?
1k-:" 1 } 4
o
-1.6 airfoil section, .2 -. -.5 -.2 -.4 -.2 ti ~.; ~ .024 .016 .0/2 .004
~o " §
..... .~ .\J ~ ~ ~
..;:: ~
i:J ~ () & \J
~..,.020 ~ (lJ 8 *'
-j:: Jl.008 ..... :t: 65,3-018 NACA \
- -
r- l- I- I i - +- the -l .j.
-1 -f-- of , I- b-- E- r"'-Q.
iL ~J
-, .J
r\ 1::'\' l'>.
I ~ : j /6 ~ T';Q; -{, ,J I ..
deq cc" .~ If Aerodynamic cWilcwri.tics .iJ.
I- A l' attach, Ii IL of i II angle II -8 .
.d: :Jl] Section \ -16 1..10.
1'-1 \
em--: lllt
;: " -24 I I
T
r r
r r r, I, I r r , r
IJ I
i
, , 1
4 9 o J .4 -32.
.8 1.6 -.4 -.8 J.6 3.2 2.8 Z.4 2.0 -I.e -1.6 ~1.2 -2 ~ \J ~ \J c:: ..... .~ . ..;:: ~ ~
~ .~ ~
o
./ -.2 -.3 -.5 -.1 -.4 ~ ~ <.,)E ~ 8 f::
..... .~ .0 ..;:: .... .g
\
I
> -a 0 Z > t" > '=' <! ...... rJ] ::c ~: 0" 0 ~ ~ ...... >-3 >-3 l'J l'J IT! 0 !:U > l'J !:U 0 Z > d ,.., ...... 0 rJ]
tv g
,~ II I ~ ~ 0 0)
Z > 0 > en 0'1 W ... ,0)
1.6 I~ I I l2 II ~ I ~ 1/·11 ",.~J'1 h .JI .8 A
"1
.OJ2 .036 .028 W .4
-
lr:; ylc -.098 -.086
I-
iJ...
coefficIent, .8 0 -L i~.069 position .'
;;;0.
r--
I'--.. r:..:. L.-, lift :rIc a.c.
-..... .L6B 1.266, O.~69 -<: "( I -.4 I'-.
Seciion J\.. I'n ~I' I x/D' :rIc \ F<> R 3.1 6.01 10 h ::0 .4 o 09.0 -.8 ,0 - .2 -1.2
- -
airfoil section, 24-inch chord.
;
./
, '~ 0 ~ '" ) I 4 I}
o
-/.6 -.2 -.3 -;4 -.2 -.5 a=O.8 ti ;:>l -R ~ .024 .0/6 .012 .004 ~o I!J o <::
~ \) ~ o
..... :Q ..... :::>: ' :t ".020 ... ~ c:
8 &'
..... :u :t: is ~.008
'~ 65,3-418, N ACA the /6 characteristics of deq tXQ, aerodynamic attack, of' onqle -8 Secfion -16 -24
o
-32 .8 .4 -.8 1.6 -.4 3.6 3.2 2.8 2.4 2.0 -l6 ~2 -2.0 Il> \J is 8 I::
.~ .ll .... ~ ~
l:.~1.C .;::: <t- ~ :g
/ ~4 -;5 -.1 -,2 :t ~v
<:J o II E
~
~ ~ ~
~~3 1
rp rj ~ ~ ~ >-1 o I-.:j > .... ~ I-.:j o .... t"' t:1 ~ >
"" o
-
I
z » 0 » 0) CJ1 CAl 0) - CD
1.6 , I l'[ II I I I.Z I ....z.;r IJ-III ( I.
t'l IV IP ~ I{/ I II ~ .8 1.
Ir V I l ~ c .oJZ .036 .oZ8 ,/ ~I-- 1.0 .4 rouqhness r-,.
y/C ~~ -.063 -.093
r--. r' :-...,'\
coeffiCietlt, .8 -+--, £.020 position ~ ~ t- lift
r-- ,.....- i'.. ./ :rIc
a.C.
Standard Or-.. .276 .273 I", I'Ll Ivr- 0.279 I-- .0 -.4 0- Section
"'-
x/D' zjc Il-. r--r R 3./ 6.2 62 .4 o o 090 I).
-.8 chord.
.2 -1.2 V
V
, 0 r
'~ I 1 I , ) ~
~ :;
o
' -£6 .2 -.
-.e -.2
u eO
.oe4 .016 .00e .008 .004
~o (,) ~ ~ ~ ~
~ :\:l :::: .... ~'
,.:'3.020 ~ I:> 8- § \J
...: .~ .13 :t {j :,::: ~
65,3-618 airfoil section, 24-inch NACA ct"rr the Ih.
_\ \V ,:)- ....
~ f'\ ~p \
ru r"-
r>...
r'-- ....
:Y' J- /6 b-- I".- f?-
l/i :r iJ
deq -/.i A V.....- ~ DC., j !"" ,Acrodynamic characteristics of If/
I//:{ u
attock, 1/. J/ of III. !II N IJ.
'(// A 11/ ~ ~ i ol7<Jle II; ~ I -8 iNC / '\. ,/ Sr:r;tipn D- ~ -/6 't ( , -:?4 __ -..
~ 1 t 7 - , 1 7 1 1 ) ) 7
.8 -32 /.6 -.
3.2 3.6 28 2.4 20 -I.
-I.
-2
CJ Q) 8 S ~
-+-.~1.2 .~ <t-.. ....
.g ..... ..::::: .;:: ~
.I -.I -.3 -.4 -.5 It.
~ QJ 8 ~ Ii:
l-.2
..... .~ ~ "-- ..... ~
I
l:!l t"J '"d 0 l:!l >'3 2! ~ 00 ~ "'" z > >-3 ..... 0 Z > t"' > t::I ~ ..... <F1 0 l:!l >-1 c 0 ~ ~ ..... >'3 >'3 t"J tz:l ""J t"J >'3
~ ~ 0 ~ > ~ 0 Z > c::l .... C <F1
I
z » 0 » en U1 .... CO ~. r+ :::T ::!: S»
en
W 'C
I I
I
·1 1 1
~ ~ "I
0.0'
t 1'1
I T I I I
II I II II I I
II I
I III 11 J
I I I I I I I I
I I I I
111111111 II I
Til I I
1111 I I
I I I I I I I I I
I
II
I I I
II
1.6
I I I I I I I I I I I I I I I I
z £1 1 [~ lill I TTl
ltlllllllllll '1
~ I
036 ~=v
',1 m ~
1.· IT I
.1.
.
/.e
rr
U m
1:I.t.'.1 J
T mJhl 71-
lIlt
I I I y."p II
,/.0 'L 7
IT I coefficient,
I A
I ~ I [
II
.8 I lift
I I I I I I I I I I I I I I I I I I I [
II I j
~~
I
~
.8
R=6XlO', I¥ ~~ I p."
I I .4 Section I ' .
~-i-0321
,~
I I... I
I I I I I I I I
sealed plain flap.
I I I I I I I I I I I I I I I I I I I I I
~
.A 'V
II
II
~
O.20r ~
I I I I I I I I I
",¢, I I
I
-;4 ., ~ f.:El
II I
-
'\. ~~
~""
1~1't> r;::;- [
I I I I I
I I -;8 0. ",l\
1.1 I airfoil section with
I I I I
11111_11111
m 4 °111_111111111111 2
.21
-.2~
,81
-~Ol .008 65,3-618 .02. .004 .02
~
i
J
i1 i ~.OI6 ~ u B'.Ole ~
. .)01
;...- :t: ~ t Jl
NACA the of I ~ - - it t- C- t- t-
I> :{~
t7 f'" TTl
'"" ~~
rr!t ~ ~
1/
~ ~
H'r- deq /6 r;<. ~~
~
+=cttii
[?' :J'I<' -f Q"o, j Aerodynamic characteristics ~ fV'~ ~~ .
~ 0-
H
~~
~ II
[7::F 8_ TH-t-+-I----L..LJ
5-- I{ ~ ~I! 1/
8: ..,,,(7 r'U'"1 It
rL . attacK, I II/ r' ,/. (I L(I/ I"I~ r" II I nf I ~ of {JI/ LA' IF -J-J-J I
l/. !I 1/_/ ~I 1,..1 V
.I( ~ 0 .If
2"1
I I -1 :of JI,....,!
I angle I
PI~
71ti I !
~
!/1I1'1111/1(" ~ ~I~WIP III
1"," 1""< (f 1/ :VI II I / -A ~8
=r=Ujjrt-H+'H-+U=ttti
A _
~U/
'IlV) Section
c r
)
0 St:!tt±±l+t-t+P=t=tlIH 0
Ill!
~rl=t!jtti-t1~~~~t1=t±i
IO~l=ttj-rt~t+~~J:t!jl 15 50 I
cO JO:U=d±H--t-1H- 40t:tl=tti+-H4-+=l=+uttrt 60
~- -cO -15 'I III '11/1/7 I - (deq)
It '~I!Jf1
~ V f7 A 'V
o 19 1£> I¢..- [ -16 JJl
~
......:-1 L...
J
L... L... \] L. L... t-+-I-...j......JI- t-+-I-..L.-l ~~_IO ~~ I---4---I-L .......... ............. ~~---.I<I ~~---.I'Q t-+-I-..L.-lq .-+--'--'- Cl=t!j~i-t+~~/~t4~I/~~~~ rr-r-r-t--H+H/.Jf-jMAi~ ,r;
, 'L... , , . t
A o -24
.,.
.8 .
-.4 -.8 2.4 20 1.8
16 3.2 -1.2 -1.2 -/.Ii
Q, a -20,
C .~ ~
.... .t .t ~ <t: "" ", .«.:: ~ ~ t-< t-< t"' ~ > ..., > ~ > ~ 0 "'l > ~ 0 Ul 0 ~ I 0 (J1 0l=Io U1 ~ en U1 J\) - - Z » 0 » - $ 1.6 I I I 1/ 1.2 Ilf!
J ,.., r 9/V II .8 '>- ; l c .OJ2 .028 .OJ8 .4 1.0 " ~l\.
r--..
y/c -.020 -.031 -Q023 coefficient, t--.h 10. o r-- .1 _ position -- .8 f::= - .. lift J:jc t"--, a.c.
r-- .267 .266 o.26~ ~ "c p:: -.4 .6 r-..
-===-- Section ~~ "- ;qc ~ R 6.0 3.lxIO· b o o 08.9 -.8 .4 I~t--..
""h ~ -I.e .2 ....-- ~ , > 0 -1.6 ,V o a=O.5 airfuil section, 24·inch chord.
-:5 -.4 -J -.3 -.2 .2 ti -.2 i ~ o Ole <U o !.J ~ ~ 020 (,) ~ ~ .008 .004 ~ "- .....
....
.024 ·415, ~o - <.J c: !.J c: \l Cl &,.
.:1. i· ~ ~ .'Ji ~.016 ......
I , I 65(216) i NACA Z4 the \ -\.
,'\(;1 ~.
" \':~ r\.
~ ~'r"\ ...-1 deg- :t r - :bJ «0' LJ lP "--- It' 'd Aerodynamic characteristics of IA Ijl attack, j }I - of IJ01 ".
- A.
1(1 - .~ '/I angle '--- J// JJ -8 f/ ~ VI ~ L- Section I I I , I : -.YU )0 ___ -16 1O I -24 -32 .8 ~.8 -.4 1.6 -/.6 -1.2 -2.0 2.0 2.4 ~1.2 3.6 3.2 2.8 \) ~ (> t:: ~ <S' II) .~ "-.: ~ <::: ~.4 .... ~ .1 -.5 -.3 -.4 -./ :t ~ o \J ~ c: ~ ~ l·.2 .~ ~ .....
.... 't
I
t;:J ~ 0 l'j >-3 ? 00 1>:1 II'> z >- >-3 .... 0 Z >- t' >- t:I .... U2 0 l'j ~ ~ ~ .... r-J r-J
~ l'j Z < (j 0 M I"l d !::j > M ~ ~ >- 0 ::! ("') '-"
0 .~ I
Z » 0 » 0) 01 0 0 0)
I 1 l 1.6 I .
1.2 I I 60° ..
I I I .8 .
I - I IP p I I c, 1}J IJt ~ .CBI5 .OJZ / .028
!~ 10' ~iN'
deflected II I I 1.0 .4 I it'!' P: 8'" I I
r- '-> roughness
I roughness L flap '" y/e ~ ~ I \ -.014 -.033 \ 0 _I ~~r s&'"' I coefficient, split 0 position .8 ~ I
r
liff 1!1'.1 I ~/c ac.
Sfandard Standard .256
"'" ~
I_I ,\ ~ 1:250 I \ \ .6 \ -.4 simulated I 1 I L "\:::i Section I
tc
J ,,~.258 UJ/c "\,..,,- ) R 3.0xI0· 6.0 9.0 6.0 o.20c 60 6.0 L..L .4 17
o o o A '\l
-.8 _ chord.
"--'-- 24,·mch .2 -/.2 section, L- 0 , 7 t 1 ~ ,
o
'--1.6 .2 -./ -5 -.2 -.4 -.2 .
<i .<1 .004 ~O .024 .016 .012 .008 (.) ~ () 0'-.3
I
"i-.~ ~ .... ~
~ ~ \,)
(,J'!1.020 c: &' o
.'1:! .~ {; ~ ~
.... :t.
65·(J(J6l1irCoil 32 CA, I ....
; ~ :>cb chllracteristics of the N ~ :13 )X deg '-1 ''1 I?l: ~~ 0i.- J1 «., ,1: ~7<;l lEI T ~ .-\.erodynllmic ~ Id ~ ~ .
1 ~ J!l ',\.
Lg attack} JE i~ I !If or
II ~ ~ g ...,..""
I P , J vi"1 i!!
I t] angle ~ II r{ f.d -8
I'\J IJ1 [\:' ru
v.
I I!.
ret ~ I' I' III :x:t S~ction -15' I I : , -24 .8 .4 ':J?
1.6 -.8 -.4 2.0 3.6 2.8 2.4 -1.2 -1.6 -3.2 -2.0
~ II! 8 t:: 0
"i-.~1.2 :~ <t:: :.::: ;g ~
.~ :t
.I 0 • -./ -3 -.4 -.5 ~
J-.2
\,! III 8 t
.~
"i-.~ . ~ t: ~
OJ) ('"j ~ ~ ;:.- ~ ;j ;:.- ~ o 1-< t"" ~
I\:) ~
I
Z l> 0 l> en 01 0 0 (0
/.6 /.2 I , / '- l- I- J 60° Ie: 4 I .8 -l- I/V IT ~ I / c II I .OJ!!
.036 .028 iSV
rv deflected
I 1 10 I .4
Ia In- rv
I I
tv Cf rouqhness
lb rouqhness Tlop J
, y/c
c1ro'
I \ -.018 -.034 , -0.004 p-f-' codficient, .8 split 0 position LI ~ lift XI}!: I o.C.
Standard Standard 1'- 0.262 1.26 1.264 1"-
1-1
.6 >t:::, } -.4 simulated I \ Section 1\ I"" !'-.. I I \ :rIc I II i"- I\:i I R 3.0xIO' 9.0 D.20e 6.0 I", I .4
1716.~
o 06.0 <> 66.0 v -.8 1'-' I~ \ chord.
l'i ,'\ 24-inch .2 -1.2 I
-
r--:::: I . 0 ! ,
o
.2 -1.6 -./ -.2 -.3 -.5 airfoil section, -.2 -.4 ci ti ~O .016 .024 .012 .008 .1J04 Ii ~ \) ~ t' :i:J 'i:: '1... ..... ~ <:
I;,TJ.020 ~ <.; &' c: '0
"-
]' .\J 'i:: i5 .Cl ..:::: .3i
Q!HJOII A I 1 I
- l l
"1 -1 l l l "1 ~ 1
A.C ~
;~ - t-
I
11 N
tM of I ! 1 I I I ~ .
deg :J.,) r\ 1,1, ::1'-l:( ~ ; I ~ «0,
I 1'-' Ih
I I Itl II ~ !.:.. Aero<iyn!\mic char&cteristics 7V 1'F('i;i II ~ 1.1" Ii:!
ottac;:k, II 1;1 of
1# IA B'
II lY-' II II ~ 0l79le III IF j -8 IA VI tYr'i ~ 'v}."
VI 1 I:J b I" 5ec;:ti()n -16 ' ..
I i ! .. I -24 I I I I~ .
.,
4 0 0 o
.8 .4 .4 -32 1.6 -.4 -.8 3.6 3.2 28 2.0 2.4 -12 -1.6 -2.
c; c: ill 8
'-.,12 ~ .g> ~ ...... ."- -::: .~ .;::: .'0 ''1...
.1 -./ -.3 -.4 -.5 't ' \)
J-.2 \J ill <:) c: ~
...... .~
~ .... ~
I
.... "'l
t;j >'3 Z ~ 00 ~ 0I>- z > >'3 S Z > t" > ~ .... '(JJ g '"'1 0 0 ~ ~ >'3 >'3 t;j t;j 0 ~ > t;j ~ 0 Z > ~ >-l .... 0 1/1
~ ;g ~
~ 0 ~ I
z » 0 en en N 0 en
»
1.6 , .'
.
1.2 -~ ;:V I .8 I L~ '-4 .J I / ~ ::6 l- 'y" c, 1111/ .
.OJ6 .OJZ .Qza V P'r" de/,~c~ed 1.0 .4
/ fir
.
lSI' -'- .A~ rouqnaess ~ rournessll f/~
'/
flIJ' 1I1e po-
, C -.0/9
'r~ "-.045 , -0.0/3 coefficient, .8 position '. ~ ~ 3:>!hl.
c.
2"- lift I" ;:t; <clc o.
standard standard \ .257 .257 ). r-....
I I \ I .6 <:: ~ J: -.4 I ).
Section 'lie R 3.~IT::J0.254 6.0 90.1 6.0 o.2~cJsi7'jlarefJpl{t 6.0 6.0' 11111/111111 .4 o o <> 6. "V -.8 .2 -1.2
, , 0 I , ! )
o .2 -1.6 -.I ai!'roil section,24,inch chord. , -.3 -.4 -.5 -.2 -.2' .; ~o .ti .024 .016 .012 .004
" ~ ~ \J ~
"-~ ~ "- ..... ~
"f::
Q".o20 II> v 8' \J
'Cl ....:- .~ .\J .;:: "- i5 .S .;::: <3l.008 65-206 NACA the of I I I ! I I ! , , ...1 .l ~ , w.
""- :J:'::l~ 1'5 ~~ M tH~ :II:!
de9
h
~~
{',tIC!
"0> ~ !Z!!
Aerodynamic chuniCteristics
"'"
...& ~ t'\.
.J ~ I"~ 1\ '\ attack, J .
Itf' '; g ~ IE ~rn. of li Ir't I 1,1 IV!:':': / .~ angle ~ ~ 't: I -8 4 ...r1
u:""
- Uj.
Section -16 - -24 ,.
L.._ 0 .6 -32 .8 • .
1.6 -4 -.8 3.2 2.8 2.4 2.0 ·.3.6 -1.2 -/ -2.01 Q ~
III a c: 0
...:-1.2 .... ..... ~ :2.4 ~
;g
.I .
-./ -.5 -.2 l ~
<! c: o \J
III ~~.4
..... ·91 .!,! :::-3 ..... ,~
,~
U2 Cl ~ ~ ~ o "'l ~ l:C >:j o ..... t" o > .., >
I).:l o ""l I
z » o » CJ) (II N o CD
1.6 il
Ig If! ry
/ 1.2
.L
'f
IL V 1/ ¥ ,- V 80· :f .....: I 1-.1 J;, I /i '.8 / .
./ l
c
2v
IP
.032 .028 .036 -.C F-f deflected I 1.0 I .4 P I ~ I v ...... ?
flap roughness roughness \
h...c:k ""' IIle
OI '\ - , -.010 -.004 I}I; JIJJL1_ coefficient, split
J. t· 0
.8 position f-. R- ~ I"" z/c Sfando;'d 0.0.
.259 .259 1\ ~b..
i.Standard I I ,5
-.+
':::::r .
I simulated I t Section/iff ~ I ;rIc Y-;v 1\ R l. c
6.0 gO D.20e 6.0 6f
J.DXR=~261 6.0
J .4- I7
o o o e. 'V
-.8 inch choru.
'?t- -1.2 :.2 Stll'tiofl, v-: ?
, I , foil
o -/.6 .2 -.I -.5, -.2 -.3 -.4 ait -.2 o ti .004 .012 ~ ....
-10 .024 ~ Qj o () t g
~ 1--- ~ <0 .....
,J.020 2 ~ &
~
.v ::;:.016 is t ~.008
;-,.- oa- A AC N , I .
- _
. the
or 24· L-L-.
_ L- be )D : /5 1: 1:::l \~ ~ ~L~ .'( I~ de(j f\-J ~l\ .l-L-~
\ "
¢--. i1'" ~ ~ N ~ Aerodynamic characteristics Ii< II J ~ 1.-1-1- oitack,«o'
II IP
d ~ of IL II ;:I
,J "'" 1-1-
!!. 1/1 ",'! angle J I; II Il\ J 1 -8 I ,
, J'V~
~II III\, IC> I I 1-1- Section ~ -16 -24 , -32 .8 1.6 -.4 -.8 3.6 3.2 2.8 24 -/.2 -1.0 -2.0
Q, 8 !::: u
t;-
.(.) ~ <:::: :.::: ;g.4
-..::1.2 .~ <"';" ~ • .I -./ -.3 -.5 ~
-
&-.2 l:: l1! C> () ~ 1:-;4 <:) •
.... .g 'Q; "I- ~
I
.., t'j '1i 0 ~ Z 0 00 ~ w:- z > j 0 Z > t"' > t:I -<: .... Ul 0 !:::! ~ 0 0 ~ ~ .... .., t'j t'j ""J 0 ~ > t'j !:::! 0 Z > c:1 .., ... 0 Ul !:::!
~ 0 00 I N , I I 1.6 I ;> I It 1,2 /, (') rt /~ / ~Oo ::1 1/ ';f/ ,; . .8
II lv ~
I I I
v" ......
I c
V- v. P
I .0.36 .032 .(J28 - I;' deflected I I 1.0 ~rY .4 :...i; Jl!:.'-', I I ~ rouqhness rou9hness (laf I
'/
life i I I ! ! i I ~ r'5' I
+ +-
-.038 -.020 , \ -0.044
r--
I split coefficient,
t o
.8 position i I I I t-., I h lift a:/c I a.G.
Standard Standard .262 .262 "-:::;'::('1 i'- ~ I \
r--- t-- _f.262
I \ .6 -.4 1 ).., ...... sImulated I ~ SectIOn I z/c L" I R c 3.~rr- 6.0 90 0.20c 6.0 6.0 Ie I .4 V [7 o o () 66.0 . -.8 .2 -/.2 - f0- ~ (J , , ,
,---
o
.2 -1.6 -3 airfoil section, 24·inch chord.
-.2 -,2 -.4 -;5 ti ti
-
~O .024 .016 .004
J ~ ~ \) r- ill
- .... :~ 'i:: '1... ...... ~
~'lj.020 ~ g v ()-. \J S
{:j.012
.... :~ ::::: ~.008
65-210 N ACA
--
r-!-- I the , i ' I ! ! : , I ...
I i I I i i I J
--t-
~ ! , I --- I -- ~ 'Y /', ;;, ~ I~ I deg A ----- ~ !:A i IV' '1 ~ r\Z Id 1:\:0, I '\ j }\ --1- V- I Aerodynamic characteristics of A I(j ',1 :;>" /, ottaclr:) i -f- VI ,ff - 0 ~ "I of ~ ~ --
VI
/l c- Itf I £ ~ II angle j-J I --- _J; II -8 p\:<.
Ih t Ir;! 11 ~~ I
•
L
rT
~ Section I I i ' i \ I , I I I i i ; I : I ! 1 I - I I I i I I I I ! I ! I -16
!
T I ' ! I I -24 .8 -32 /.6 1.2 -.8 3.0 2.8 2.0 -.4 3.2 24 -/.0 -/.2 -2.0 "'~ ~ \) c:: U
1:"
.<Ii .~ 'i:: "- ~ :j?.4 Ji
.1 -.1 -.3 -.4 -,5 -.2 ~
-
,," E:
.... 8 &
.£ ,g 't ..... ~
Ul q ~ ~ :.- ~ o :.- .... l:O "'1 o ..... I:" t:! ~ :.-
l:O t-.:) ~
->:oj I
» en ~ o
z » o U1 ....
I I , J 1.6 I b. LI- [I 1.
Q , J 1/\2. V ./ /.2 Ie< V 'I 1/ V / i I II P g~ 6?O V I-' ~t" i 1 , ,8 V >f' ' ~ j V "I .tJ36 .IJ32 .IJ28
m'O
J...L! 1 r' , deflected /.0 -'- .4 I I , ~ ~~ !::::. ! .
joy~h~e~s I roughness flap " y/c 1 I I \ ..1 \ ,- -.035 J--
b- 1't::1 ~B!
I -0.025 coefficient, spill .8 0 ~-.024 position ~ I
-
!:l:;:"", i\ 10. I '..1 I lift
t-r
\. :rlc.J S{T~ar,d O.C. Standard .259 .262
1\ Ll ["-. I"
0.258 I I ~ ~ ! I .
-\ -.4 I I"l: I I 1 I i I . I' 1 simulated Section ~ ! 1 ,- ; I I- I I ~ ;rIc ~
, I i i. . I
R
SF
6.0 9.0. 6.0 o.20c 6.0 I ! I I ! I ' I I .4 ~ I I I , ! i I , ! i I i I I ,O.']'OxIO· [0 10 -.8 _I.e. chord.
I , i I i ! ! I : ! ' ! 1 ! ! ! ! i I. I I ,
I I 1 I 1 1 l I I i j ! I i ; I I I i I' I I i
1 1 1 - i _ I ! I i ! I I i i ; I· I .2 -/.e I 1 I ! I I i Ii; I 1 1 I I 1 I f.-- I iii 1 1 1 ,! ; ! ! I I I I L .
i . ; 1 I ! I ! ! I k-::
t , , I
o
.21 -/.6 -.3 -.4 -.5 -.2 -.2 - .; ci ~O .008 .004 .024 .0/2 ~ ~ Q, () \.;) ~ § '1-..- :~ ::::: :;:: - .,...
c:: 8 ~
Cj'.020 III l::: u (lJ .... :~ ~.OI6 {j .S> ..... v, 65-410 airfoil section, 24-inch , 1 I NACA I the I i :> I
"
characteristics of I:": ::>M: It I -i;.
Fi;il:{ ,.., c;Q~ 1
deq ~, ..,.
I i ! I I o• ,j A;...
~ T \
;r '
a Aerodynamic ~ I~ W .,u I i i i 1 1 ).: ll', I~ '(' , I : .4 d Ij ~ I ; i - j, ottacA, df: (j ?f ~ J.
til of' Ijf I /, VI Il':l i ! I i /, ""K;L IJJ II ~
In (I
~ , , II Y ;,-< 'v.. -8 I , 1 I I I I. , , I 1/'
¥~ I-K
1: \ b secticm engle I ' -/6 I ! I -24 , ,
7 7 I I , ?
'> 2 q 4 -32 .4 1,6 -.
3.6 3.2 2.0 -.
2.8 2.4 -I; -I.
-2.
~~ III Cl <oJ ,...
.~ .\.;)
"[1.2 «::: '" <!:: ::::: ~ ~
o
./ -.i -.3 -.4 -.S ;t
-
l-.2 ~ 8 ~
~
...... ~ 'q; ...... ~
I
.-:l t:; ~ ~ .... >"l >"l t'l t'l ~ >- t'l ::>:l 0 Z >- c ~, ....
t'l "d ::>:l >-3 ? "0 "" z >- >-3 .... 0 Z >- t" >- < .... Ul 0 ::>:l ~ 0 0 ~ 0 Ul
~ 0 t-:l I--' 0 ·Z I
Z > 0 .... N
> 0 en UI
J 1.6 I '~ I~ I /.2 [...{' I.P..
r rJ: V / V . ~ I )1 / T I)- / .8 / y II V 1 1 I Til f:, c, V ;...-- .032 .036 .028 V InV "I.
defieCted·60· /.0 .4 g I I l l ~ ..t; IW ,~uf1e~s roughness flop , , , yle ..., , .002 I, , -.OIc
- r 0.018
coefficient,· splif ~ .8 position 'V' Q..I/\
-
~ lift ;c/c a.c. Sta~~a~d '" c:: Standard \\'~ I 1.26/ 0.255 l258 ~ I Tl .6 2: -.4 'y ~ simulofed I I Section
1\ "- I II I
~/c \ ~~ ,~ \ \ 6.0 .J.OxIO· 9.0 6.0T QcOc
"
.4 £; "76.° 17,6.,0
"'0- o o o il
~ -.8 0tQ ~l\.
If"' , 1,- b I'I~ .2 -1.2 In
-
, ~ ¢ 7
0--- ? 4 o
.2 -/.6 airfoil section, 24-inch chorq.
-,/ -.3 -.5 -.4
-.c -.2
d.
-
~O .0/2 .008 .004
.024 l
t Qj () \J ~ %
:i;? ~ - .... ~ .,..
~"'.020 ~ 8 g- {J
.~ ~
.~ ::::.016 {j .!;! .... ~
....
651-Q12 NACA the
~ r- r- r- r r
.= t-
+-
.
c:j -t-
H r-H H- H- J r-- r-r
rT H-- l:i
+-r
+-
--+-:::
r-+-t- t-+-r -;--t-
r-+-t- t- t-HH H---
t- ._
11_
Ii+- I:c'. Ili
r-
J -I----h- "Jt-H
.
'" ~~
I rttt-t- "' 11 ta-r--
-I-h--r--
deP
---+--r-t- -+-+-l
wr>~ =tttt H-t
~H
"
t-H-;--r-t-- «o,
'" t-
111'm ~ rl±t
.
-t- ~ .
t-+-"+-r-~ , I
t- t--- II
Aerodynamic characteristics of
r- ~
+ ~
.
+ c:j I' I- r-
1 rEt r- t-
ill
m +-+
ottocA,
r t- r
::: r-r-
U t-t-rr
", '!-H-rr
t-t-R= of
I 1
~
+-t-~ J +-
II d
iT f.j --r- _ '"j::j
H-~~ t-
#
H h~~r H--t-rr~
t-r-t-H
TiT J't- . angle 1 I t- ~.
1:
Jl
-8
t-+- H-rr I
I, .
t-H --t-
t- t- I- r-iB
~
~1:±±t-t-W ~
. ~ -r
r, t-t--t-t-I r- I+'r-
LJJ: . Section
~ r-
r
t-t-rftt-
t--r
~
H-t-~ I I -/6
±±:: . ttt-r-rr ,I .
;::;::;::;::;::H- >--r-T-T-1It-H-
t-t-+-+-t- <--t--y-r
-t.
r
;:::; ttt± W'
-.l oj I - -.
. +-
1·
+J-+-t-t-' -24
H-r-t-H- I 1 '1
1 ,.--rl
t- J::+-+-l±L-"'1lt-ffir
r- ilia I 1
r r t-. r-
t±±j 'II
~r
-r rt-~H-t- t- - r t-~ :- r- -f:ttt-
OTI~..1 c- c- c- I t- c- c-t-t-~;::;::tt-t-ttt~rtttt- t- t-t-H r t-t- r H- H-±:±:±±f:l:.K- t-I- I :='t:J±r-t- ~FFliet-tt t- t- t- 1tt+H r- CW-+-+
e
'r-
.F~rrD '1
'E
.8 -32
:4
-.8 -.4 3.6 3.2 2.4- 2.0 .1.6 .2.8 -1.2 -/.6 -2..0
iJ' ~ Q) a \::. 0
:~ J/
,+-~/.2 ~ t :g
';/.
-.3 -.4 -.5 .-./ ~ 8 !::
.J-.2 ~
~ .t: .S! ~ ~
'qi '"
"'1 o t" t:!
w ~ is:: is:: > ~ o "'1 > ..... l:C ..... > >"3 > t\J ...... ......
I
z ~ ~ en N ..... N
o en -
..
1.6
:r 1..1
J.
I .
{J --
1.2.
I~ IT lI..P / 1.1':
c
It' /'rf 1 1 {]O°l II ~ .8 IJl 1 : - J ~p I ,./ c I;: .036 .032 :028 I~ '7 deflected 1.L 1.0 .4 .
I I I I ~- roughness roughhels
rlqo
,
1lfc
, [VI'-" -.032 -.02b 0.037
:---
1', spiV coefficient, .8 position ~ lift
l"- LLJ~L
o.c.
Standard Standard .259 .261
"
- ~ I
_1.~ v I
.6 \
--4 i'-. C--""f I 1 simulated : Section Klt-- I ......
a:jc ,:--..'- R 9.0
3.0~~j:258 6.0 6.0,! 0.2Oc 6{
,~ '0.
.4
o <> '<76.0 (7
"Ill. o 4 -.8 1-1 '"\ '-<...IQ
t1
J 1\ .2 -1.2 I J-- V r--.
o
.2 -/.6 -.I -.3 -.5 -.2 -.4 -.2 o
£"
~o .024 .016 _004 ~ o ()
~ ~ c:- Q) 8
;g <;.: .... ~
" -..:- 0"4;020 c: ~ u (), \) c: u .'!! ~.OI2
.... .<J :£ .0 .;:: ~.008
.
N ACA 65t-212 airfoil section, 24-inch chord.
the
~
-+ --t
,
H
-i
~-I -1-1 -+ -t
-1-1 +~
=:§ _
~
r r- r >-+-I I '-+-1
~-d 11
III
-r-
r-
"i-l, ...1-1-1-
.1-1- ...1.1
, _
-I
'
!l'
4 1 -1' -+ 1 /6
...l . , '''--J-,
~
I I
-4~rr j+rJ ..- "~~
...1 ..1 ..1 ...1 ..1 f .-+ -l-:- 11\
-l-w++-
=a.
'r-t '...1 -H-
+±' ,] 'li-h
4t
ttl +-
'-.l' oco,dec; t- t-...1 lR' Aerodynamic characteristics of
~
r r I I I
jl j .1
1 ± 4-1-
-- ~ r-y
r, attach) _
1 r
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1 I
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1 +-t-- 4
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r- I no 1
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r 1
.
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t+
~
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±
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.J. .J. ,
'I
.1 Section
~ .1
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rr ~
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r, t-r -/(j
I I
r'
HY ..1...1' -Ll-ri ++--1-
.
r -I-
1 1 ~
rr .1
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1=
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i i
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J
I ±- ~
t
.
reT
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'1 1/ --~ ;:: -, or r r rrM r-rr rrr r-r--r- r-r [-...1..1 r-u ~nD
1 ct+-t-+--.lH-n=t.J.u+ 1-+-+--1;'("+1 ~~..1RIBi~r~+R±nrTU~~r I-+4-TI 1--4-..-t--1:-,. 1-+-;--"1 ~'" r [ --"" W-I-H-t-I1-rll W--I-t::t++-t-tUllt:+ >-r-,--1, °U ' 1 t-n 1-1--1-1-;---;-1
r
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II> 8 c: I)
~ ~
;g ...: <::: :g.4 ~
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l-.2 ~ \l) o ~ ~ ~
~
..... ~ .... ~
I
z :> '"" .... '=' <: .... 0 ):l:l "'1 0 ~ ~ .... >-i >-i t"J t"J "'1 t"J ):l:l 0 Z :> c:: >-i ...... 0 U2
):l:l t"J "d 0 ):l:l >-3 Z ? 00 t-:) "'" 0 Z :> t'" :> U2 0 0 ):l:l :>
Nl .....
tv I
U'I - N ~ ;:; ::T ::!:l m
Z » 0 » 0) - N
'tJ 3.2 I I I
l l 1 l
-j , t-1 r- Il
J
2.4 ~ 2.0 ..
'J' c{ rz: 1.6' 'il- ~~
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40 70 50 60 Of I I coefficient, 1.2 V- (deg) I
0 0 " A \l T
JJJ
?,:; lift (1 R=ftXl0'.
L- .8 I~ \,n Section p.:~ .J.J split Oap, O,20c .4 with E!ib.
o
~ecnoh r:;i:
rv
.J h.?-- airfoil <~ -:4
R 1'....."
r,5,-212 I .
: i I r
. 1
o 7
.5 .3 .2 .4 '-;8 -.2 -.3 -.4
-.1 -.5 -. -9
-.6 -.8 the N ACA ~ ~o " ~ Cl CJ ~ E: ~ o[
;..:- :iJ 'Q; ..... ~
::>.
~ moment charactCl'istics >-~ and I 16
~;J_ JJ
I.iIl ~ 2..
deg y •
"
o !li\ J er ''''f~ ~ l/ 1 ) 'iff' ..1, I~ ,>/, IV ~.
attac/{) If r<J.1 10 ,c: :/11 of
'Ii2
IrP I J /I.
I 'I I/, I~ J ( I
III '(f
"~I I I -8 II( I ·UI
11ft VI 1t5
( \ /7(1 I y<;., (fI1 : 1\ Section anqle II -16 I • ~ 1
• r r f 9
o
-24 .8 1.0 -.4 3.6 3.2 2.8 2.4 2.0 -I.~ -1.6' -2.
"'~ t:: Q) 8 !:: ~
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rtl c:l ::>:! 0 ~ t:::! ~ >-
t-:)
- ~
• II
Z l> ~ C» (JI - N .... N Q P C»
w' 1.6 -.rI ~
I!l
II l( .
L 1.2 II f-' ~ r:fV >7 P IL Ff 60· .
h' Iii V I .8 if' lL j l(;ir": .,... I, I .
f!
til t;-. I .032 .036 .028 ~ .d '6r1 ~:J deflected I 1.0 .4 Ok>.
I
, I , •
• roughness
flap roughnessll11 I
\, #
, -.019 -.019 , -0.066 ~ I-- I split coefficient, position 0 .8 I3;i IOl ",',' I
*t I lift
t-- a.c_
Standard Standard '. .269 :< _.
~.-.- "-
J... ~p. . ' I.
0.265 1.
-
);:- ...,!
I I l \ .6 ....... ....... simulated -.4
1\ r' 6 I 1
Secfion I"'-- I II J:/c I" R 3.0x10 6.0 o.20e 6.0 6.0 )- 6.0 '"' 10['1: ~_L
.4 o o 09.01 L;. <7
r-r
-.8 b IV", .2 -1.2 I-- airfoil section, 24-inch chord.
, 0 )~ t I 1 I 4-
o 3
-1.6 .2 .
-.
-.4 -.5 -.2 <i ~ ~ a=O.6 ~O .024 .012 .004
~ ~-2 ~ c u_ ~ ~
..... .~ .g '<-.: ...... ~
,].020 c: Q) B g-
o
...... :Q '<:::..016 il ~ ~.008
65,-212, NACA the ---j --j -
1" "1 :H h.- r
t-I t-I I t-I
·Al
"I-j"t I 16 ft II!"
~'} -0-
['D Ii ~fT ~~ deg
" I
• , b I: o I I • ~ ~ a ~ I~ ~ I}' • Ifj .J.
II Ir"r, l!f, IV' AerodynaIllic characteristics of .£ ~ I I II' 1 attock, -~ ~ :f i\Etj \.J..
of ll.iI I
J
~ III b ~ II" Ii angle ) ;'1{u -8 ~ 10 1'-' .
.
• I • • I ~ ll:~ li, Section -16 -24
l- - '- i-' I-
'- I- I- l- , !'- , .4 -.32 .8 1.6 -.8 .3.6 2.4 2.0 -.4 3.2 ;::'8 ~1.2 -1.2 -1.6 ~ 0 -20 c.i 0 c: .
.g <0.: <::: ~
..... .§ "" :g
o
.I '-.I -.4 -.5 ~ ~ c
.1-.2 S ~ o I:;: ~
.... ~ ~
'--.3 .....
I
>-3 00 ~ "'" z > >-3 .... 0 Z >. t" > t:1 < .... Ul 0 l:O >1 C":l 0 ~ ~ .... >-3 >-3 ~ t"l "1 0 l:O > ~ l:O 0 ~ > c::1 '"'l .... C":l (J) l:O t"l "d 0 l:O Z ?
t-:) ...... J+:o.
I
Z » 0 » 0) (11 ~ .- N
1.6 "~ I ) I I I '? 1 I f/ I 6.
I~ r7 r/
I /.2 I y f/ I II Vri ~ 1;1 lP laV , 60· lP .8 Jl Cf ~
Itt
I{ l c .0.16 .032 .028 deflected 1.0 .4 v
, , , I
- l ~;)r
roughness ~~U~h+~S
T/ap , ,
, I.L_L
, -.044 -.038 ,
r-- "
tl-I& I coefficient,
o
split .8 position ~o.047 r-....
I lift
- ~
1'1\ " I
;e/c.:..J- y/c a;c:
Standard )ta~Ja~d
\ .265 r-.: I,
0.263 I
-
1.264 \ c~ fion 16. ~ I- .6 -.4 simulated
I II
Sec "-
I
.;rjc \ ........ I R
3.0X(OL- 6.0 9.0 6.01 o.20c ~g
IQ Iq I"- I
.4
o o <> ~
-.8 inch chord.
..
.2 -1.2
-
~ectbll, r--
IV- . ~ I 1 t 1 ~
o
.2 -1.6 ~.1 a!rfdJ -;5 -.2 -.4 ti ~.; ~ .Ou8 .004 ~o .024 ~
~-.2 IV o U-.3 l' ~ o
..... :~ ..... ~ 412 ~ ~ 8 t>- t:: r,J'.020 c: 'll o () .~ .~ ~:OI2 ~ ~ ..... 'if:..OI6 6."
A At' the N of I I
1 -t 1
/
~
~
ro
I I I I I -p, I : I 0 I I ~
1?f6
.16 ~ lAo )d I I I deq I ~i I \,I \I ,~'I'-' ~ I ...
~ «0' IP :~ Aerodynamic charncterist.ics '/J P ,~ ' '\l '.
,if
I'
attack, ,~ '<;1 p I ~ of /I III 1--.
/J angle 1// J.
-8 (f "- C I I 'il~ /I ~'ff I
c
~ -/ ():! Ii f"'.'
< I ! Section 'C , I , / -16 -24
-
.8 -32 1.6 -.4 -.8 3.6 2.8 2.4 2.0 3.2 ~/.2 -/.2 -1.6
-e.o
~ t:: (I) ~ t:: u
..... .~ ~ <t:: :gA ~
""
.1 -,/ -.3 -.4 -;5 ~ ~
l-;2 c: 'll 'll o U c: ~
..... :~ ~ ..... ~
w. q ~ ~ ~ ~ o "l ;.. I-< pj "l o I-< t"' t;;j ~ ;..
tv I--'-
""
I
z » ~ m tn I\) o tn
....
1.6 t> ;:> ~ 1 H 10 -<' b ITII I 1.2 ~
II vv
>- )-
I
II V I I
c '<"
Ir I I
I .8 V- I "'-
IT I-f "
c, 1./ !no [() .036 .032 .028 ~ II!:r: deflected !.O .4 W I I I I 'Il ~ ')(~ 'of rouqhness :-o~qh~ei5·1 , .1 flap \ y/c .0/4 " , 0 t-- ~I r-> -0.02'4 eoefficienf,
°
.8 spIff
=t
position lift
r-- l--
.rIc.4 a.c.
Standard .257 ~~ , 1 15~o~Ja).d -1.259
- 1""
~ It. 1)11; t-0.263 .6 -.4 I
I): ts jB
simulated Section
,-- I
p; .rIc 1'-; po. Po \ R o.20c 6.0 61'7 I'. 1 1"'- .4 17 060 'V 03.0xft- 09.1 6601 -.8 l'i 1\ h ,'\ ~ 1\ .2 -1.2 '\1\ ,~ \K I-- lQ I- t-- t- r ['- l- ']
,V ) f
, '1
::Ie- airfoil section, 24-illch chOl·d.
° '
.2 -1.6 -./
-.2 c.s -.5
-.2 -.4 .; ,;
"
~ ~o .024 .016 .(}/2
.004 ~ ~. a
(.) .~ \..I, ':;; .... ~ "- .....
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8 ~ c: ",,,.020 {j .Zi ::: ~ ~.008
.... 65,.-(J15
"~ NACA the ----r ---1 - f- - - l- - ~ '-- t-r- I I J , 1:
~
~ characteristics of b :b ~ ~ I , ~ lb deq AI.
IL I'l ~ 112 ~ "0, JJ L Aerodynamic t, I ~ ~ attacir, '11): J, ~ r- J.
rt:. of ~ w..
I'Z.
[::Q I Vi anqle J -8 ~ ~ 1.
I~ j .Ji III Section
"
l.3 It lb tl J \ ..J -16 I~ JJ -24 .
- -
']J r " ,-
'r- .8 -32 1.6 -.4 -.8 3.6 3.2 2.8 2.4 2.0 ~!.2 -1.2 -/.6 ~20 <S c: ~ v c: ~ .Il!
...... ~ .... "- :.:: 2.4 Vi
.I -03 -.5 -./ -.2 -.4 't n
J c: ~ (j c: ~
.Il! .1.) .;:: '>- ~
..... .....
I
~ ..., Z 00 to:> "" ~ H Z :> t"' ~ H ~ ~ ~ ,... t-:3 t-:3 l"j t'.1 "'l ::0 :> t'.1 Z :> ..., ..... 1fl
t-' 0;, 1;0 0 1;0 ? z 0 < 1fl 0 1;0 0 0 0 ::0 0 0 0
tV I
Z en N N U1
» 0 » U1 ....
1.6 I i ~ A If 'r
'<
/1/
T
/ /.2 -JU fI I j 1/ II I If 60·
tJ v
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N c!
/ F I .032 .028 I deFlected 1.0 .4 l,..,( I 1-1--+--+-+-11---4-+4-+--+-+--1 +-1-+-+-+-+-I-I-+-+-+-1-J ~-+~+-H-+~++---"H I ~
r. ~ I
roughness roughness rlap , vIc I ,
-.043
-.033 ,I -a061 coefficient, splif 0 position .8 _'"'" l>:!.
lift .,rIc a.c. IT'!
Standard Sfandard
.269 .269
!
--'--
~ ~ RO,.268 .6 -.4 I simulated SectioR --f- ~ ~t-+l:::;: ~~H_ I .,r/c 1\ l'-I'i't--. I I R 6.0 o.20c 6.0
3.0XF'0t=0 9.0 6.0 6.0
J:; ~ !.036 .4
o <> v
~~ -.8 1\ chonl.
~ ~
5 \'
" \ \' 1\ ~ ~ IG~ 24-incA ( C ~ .2 -1.2 ~ S\lCtiOfi, --~ ,,--~ ",-- , 1-1'-+-++-+-1-+-1 1--+-+--1--1-1--4-10 1-+-1--1-+-1'-+-+6 1---4-+-1-+---1'-+-117 1-1,-+-++-+-1-+-++-+-1---4-1 , • o t , !ttjj=tttjj=tttjjj~~~~~~~'~1jj=ttl1j=w ,~~~~~~~~+-~~~+-~-+~+-~-+~+-~-+~~ ,~~~~+-~~~+-~~~+-~-+~+-~-+~+-~-+~~ o 3 .2 '-1.6 -. -5 -.4 -.2 ti ..
.004 ~O .024 .012 ~ ~-.2 Q) o 0_. ~ t::
'R' :Q ~
~ ....
-
'Il
Q".020 c: 8 ()-. o u
.~ ,~ {, ~ ~.008
.... ::::.016
A 65j-215 airfoil A,C N QUAC I I I I I
-I -t -l -t
...j 1
,~J
-t
~ I- 1O~ J: 1.
~ ~ ~ \ ", '\ .e~ D,
rv
I I
e
~ deg If'1"\ ''VoW N-.
h VI Of.)
i\ lfL Aerodynamic characteristics ..t V '.i!...
Rif' Il.
attack, It Ie of fj(
""
I{ if L / angle 'L r"- -8 ~ , I I':t~;} '(j Jr-r.
Section I '!
/Q 1'1 ..i -16 It I -24 .
r - :-
r r r r r r '--
°
-,32 .8 .4 1.6 -.8 -.4 3.6 3.2 2.8 2.4 2.0 -1.6 ~1.2 -1.2 -2.0 lIS' ~ ~ 8 t:: ~ ~
.... ~ <;..: ~ ~
'
°
.I -.I -.5.
-.3 -.4 It
l-.2 ~ o u ~
~
i-.:' ~ 'ti 1-.
Ul d ~ ~ :> ~ o ":j :> J-J r:g o J-J t"' t::J
:> 8 :>
t>.:) ...- 'I
I
z > 0 > 0) 01 J\) .p.. - 01
J .
1.6
II I I
.., J.
1 1 I I 1 I lOTl TT1
.... .1 IT
TI ~
7l1T :IV
I II
fry IYl l';r; I 11111
TIIII
JL1
1.2
~
1 1 1 II
-L J . i
1 1 I I IT II I I I I I I I I I 1111111
wr j
...,.--.---.- p
I
60·
~;g
1 I .
1 I I
.B J ;:yj
I III III
I
t9.l
J .........
1 I c,
lrJdr.'r
" 1 1 1 I I .036'-1 • J v deflected
1.1 I I I I I I I
1.0 .4
" 1 I lin
roughness
I ;:ouqhness I
flap
'IJ
lIlc
I I I I I I I I I I I I A I I I " I I
-.062 -.065 ~ -0.063
1 B I
coefficIent, split 0 -:\ .8 position
~,- I II
~/ ~ lift
I
J)/c J
O.c.
~ta.n'{ar,d .266 .268 l""'+oal I I 0.266
~
'UO" j- PStandard
\ .6 -.4
\ 'b. simulated
IIIIII~IIIII I
I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I A Section
I ~
I I FFR
'* ~~
II I 1 1 I I
R 6.0 6',0 3.OX~/06 8.9 6.0 o.20c 6.0 ~~.
1 I I
1111#1.""11111111111
T ~
.4 (7
0 () ~ 'V
-.8
II
..
I
J -1.2 .2
II
I I I I I I I I I I I I I I I I I I I I I
I I I I I I I I I I I
o
-1.6 .
.2.
-3 -:51 airfoil section, 24-inch chord.
-.2 -.4 -.2' .-.1 '1 E~ ~ ~o~ .0241 .008 .004111111111 " ~ u E:
~
.~ .u :t .... ~
....
'tJ.020 c::
': :g u ~
.S! ~.016 {j.OI2 ~ t ~
't
65t-415 NACA the J AIo '\ ~+- >.
td 1\\ \ ,~ ~~ ~ ~ tj po; \ 16 I' .,..~ I'" ('i~ <--' It ~ IN deg ~ ~ :I'tZ A;;;: -rl {7 ~ :ltri aD)
"
J, Y LlJ 17 Aerodynamic characteristics of I!i\ I~ '// ~ 14 J attack, IE Id I( of' If} ~ ~ 'I j1 f) 'V- "I IP ~ ?-1loo /I v 1(/ ~ fJ if anqle ~
IlL
I -8 III ~ fl 1// iI",", I ,
/1
I ~I 1:1 Section \ lib <;j -16 -24 , ~ 1 , .8 -32 1.6
-.8
2.0 -.
~.O 2.8 2.4 3,2 -1.2 -1.6 -2.
C 'l.> 8 c:: o
yl.2 .S! ~ 't.. <:::: ;g.4 ~
""
o
.!
-.I -.3 -.4 -.5 tl- ..
~
J-.2· ~ o \.J ~ Eo
.... :t> <i:: 'Q; ~
....
I
::tI t".J '"C 0 ::tI 0-3 Z ? 00 ~ ... z :> ...., ...... 0 Z :> t"' :> ~ ...... U1 0 !:1i ~ ~ 0 ~ ~ ,.... ...., ...., t".J t".J :6 !:1i :> t".J !:1i 0 Z :> d 0-3 ...... ~ U1
t-:) ...- 00
.
I II II)
Z » (1 » en 01 ~ .... 01 ;:a 0 01
1.6 , ~ T I.e 71 7 T- II I 17.
jll ~ ~ rlJ / .8 ~ .LJ c, .OJP .0311 .01'8 ...-' 1.0 .4 ~ lIlc -.032 ".032
--- '"9-S1
coerficient, -8 -1 ~~.018 0 .
position
r--n
--- ~ b... lift
.z:ic o.c.
...... , 1.264 1.284 0.284 "'- r- .:......
.6 -.4 ~ ~ I I Section I I ;c/c 'r-.:: f\.
R 6.0 8.9 3.0xIO·
"
"- ,'y
o <>
o -.8
,2 -1.2
--
--~!-+- airfoil section, 24·inch chord.
/""
0 , '~ , I I
o
-1.6 .21 -.3 -./ -.2 -.5 -.2 -.4 Q
~O l
a=O.5 .024 .016 .012 .004 !w ~ ~ u
i
...: :0 :::: ~
.....
.
t: CO 8 g- (J
..,"'.020 §
.9! .<J t
~- $. .:;:; ~.008 65.-415,
-j .....j
_ =1 _ 3 -l _
_ 1"""1 _ ,. ~ r r r- e- r- r-
~
~ r-[- Cc r-r- rr r~ rr
t- rr rr rr
~
NACA
-+--E
-rr r-r-r- r-Ir- Irr ~~~
""ti, r:±:q
thc \'-I-t-r,
mr--1 l 1
__ 1 ~ J.
l rK IJ I 1.1
,--J....>.,t- r
~ tE t1- .1.1 J.J. 1 :t
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±I
, r;::r-1J 't""I.
"""
f
-L ~
I ;>=1""'"'
Iii
~ -l- +++, /6
r, , , r-t-t-tl!
'TT ,,~ ~
~j -+ 1 1
++- tI
w;;r.
FH r-:=
r-r- J.
deg .IJ:u:rr..
W .
r -l.
1 1 ~ ~
-r-
_ - (;(01
-"
11.1 IJ
'
r", 8 1 J.J. J.
. 1 1 I IJ.
r Ll 1 _L.1.
rt 'It 10 r
J ~J
Aerodynamic characteristics of Ai
~
II t-r
.1 aHaclr,
-+
--r-
r-r-
ff H
~ l--.....L...JI~D 0
or
rlt I
rr
J. J. J. ~ + j
E
.J L
-
t± 1
angle
r~.;:: r~ IJ. 1.1
J. 11
H
-8
--r-
~.J
~;::r:::
-l- + ++- tt 1
..1
h-I ~
+-n
J ] cl
~
.....-,---,--,-,TTlrIIl+H·.,...-;-j Section
j j j --l ~
JEI~+=~-+=~jtl=~-=r~~t=t1=El=~
L\
.1 11 Ii. 1'\ ±-I ;1
..l
r- r- -/6
.L\.
--
i
I I I I'}
~ ± j
1 J. J. J 1..1
1 r- I
;::r 1.1
1-- J. J.-.l 1
±
,- --r-:t
II IT
-l--I-t----rT+-'<a-r-
r r-, r ~ r r r r r -24
I L r r 1 1
~TT
-+-j----J-'I r
I
]
1 ] 1
-+---t-o,
t- r~ ~ ~ Err
[ '
· I' · •
II~-l---t--n .r
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'F ' ,U o It
.8 -j2 .4 /.6 -.8 3.6 -.4 3.2 2.8 2.4 2.0 -1.2 -/.6 -2.0
rj 3 .J::: \)
.(J .CO
-..::1.2 .~ ~. <0..: G::: '" .~ ~
.....
o
,2 -./ .. ....
-.3 -.4 :t ' ( ... ~ ~ o II ~ .~. . ~
..... :t ~ ~
U2 t:j a:: a:: >- !:t! 0-1 o >- ..... "1 o ..... t"' t:I ~ >-
!:t! ~ ...... ~
."1 I
z » 0 » en en w 0 -0. (X)
1.6 I 'Ie 1.2 I I~ fV 1-+- IT'V / r'i
" ;..J
/ 60° :/ ir{ r...-.
' .8 J II I I rt In f'o If'! b. I J c, k9 1 1 gee .036 .032 I r l.d;;f ;;u.. l- f, It' deflected 1.0· .4 I V 1/ 1 f I I I ,
N I
roughness roughness Flop , y/c " -.013 -.012 t::::::: ~ Iff 1 coefficient,
o
.8 -I- split :f.007 position f I lift
I- I-- I
.:rIc ac.
Standard .268 .267 In<" 0.270 1-'. lStandard l"- I--
"
.6 -.4 I\. ~ 1\ 1<'" r"" 1 simulated
6 J
Section
,"'- 1 1 1
~
_I 1
xlc 1\ ill to: po.
R \
~ 3.0x10 6.0 0.200 6.f
~ 1 ~ .4 \ ~ V' t o 06.0 09.0 <C. '76.0 -.8 1\ I'n \ t.
\ ::> ;\ 1M I,-{ I"'
J
t-.
.2 u -1.2 1':1 ~ t-- V ['--..
o
-1.6 .2 airfoil section, 24-inch chord.
-./ -.2 -.3 -.4 -.5 ti .024 ~O .016 .004 \J
l ~ ~ ~
~
~~-.2 ;g 'i- 1-
~
rJ· Q) 8' \J
~ 8 {j.012 §
..... :G :::.: :;:: ~.008
653-018 I I NACA I the I I
t
~I-l I
~R 22 I
i\
fh I I ~ I~ I I de:;
LortJ I
~...£A\.
I~ "0, I ~ 8 I~ ~ I Aerodynamic characteristics of I f l~ ~ 'if I attocfr,
lL [g I I
of
III I
!f ~
iJJ I{f
w I
A
Vi I~ I
angle II! '11 I I
Ii
h -8 I
Vi mlllllllllllill I
):l
Iill l\d I
O-l:tll
I
GT~ I Section vtllll'( I
Jh
-16 I I!!
14?t, I ·'->li
b. 111111 I 1
-qlllll~lll~ I I I. I -24 I
IH
I I .8 .4 -32
1.6 -.8U
-.4 3.6 3.2 2.8 2.4 20 ~1.2 -1.61 -20,
,.; & \J c: ~
.... .~ V)
~ "-- -<- "-- :.:::: :g
.I 0 -./ -.5 -.4 ~
J-.2 () \J
1:
~~ .~ .\J -;:: 'Qj-.3 t ~
I
~ ~ o l:C l"j '"C 0 l:C 1-:3 Z 0 00 ~ "" z ~ H 0 Z >- t"' H
>- I::) < [Jl 0 l:tI ><l C":l 0 ~ ~ ~ 1-:3 l"j l"j I:;j 0 l:C >- l"j l:tI 0 Z >- c:j 1-:3 H C":l [Jl I
z » 0 » 0) U1 W .... ..... (X) ~ :T ~ D)
::.
't:S
"
J
.~-~
""'=fioe>, path FMp /---FlOD
(\
) I to
T\ ~
~
45° plvof deflection Flap Trom 65° Trom
.01
deflection pivof
-~~-
45· I~ FI~
~t
.212 O·fo '---"= flap.
.864;>j slotted donhle 0.309c with (a) Configuration.
airfoil section 65· fi53-118 N5 .806 N AC.'\.
deflected retracted FlOp Flap
.rol-------------------------------------------------------
(a) k k , >-3
Ui q ~ ~ :> ~ >-4 0 "':j :.- ...... ~ "'1 0 ...... t"' t1 :> :>
t-J t-J .....
I
en ~. r+ :T ~ I»
Z » 0 » en 01 w ... ...
"t:I ri' \ r>; k.f Ie:...
1"'71 ..... ...{i-- p:: fro
-
I"
-
"- 2.8 /" r-' 1/ ......
~ I II 1/ r,- 1-_ /
..- Ii'>
V ~ V 1/ V f c J
r
~ I /- ICi / -1- 1.6 / ./ ..
~
Ii r:r
IY I ~ /,'V II coeffiCient,
l.-L rr
II r-r-.
I I
I ..... 1.2
lift
/
..,..-Q "1>_ In I-- , ~ --:r- .....
.8 t) Secfion h-' h r-.., kz.
~ <l 6• ....... --I-- --r- ---p.....
.4 ~ "< R=6X10 0- double slotted,lIap.
-- --
-- ""
o
I\) f".- r--... O.30ge
r--
-o~ with K I\.
0.. -.4 "\[".
T-- airfoil section
o
-.8 -.8 -2 -.4 -.6 -/.0 .008 .032 .028 .020 .004 61i3-118 .036 .024 ~ (bl AerodYnamic characteristics.
~
G ~ <J § ~
t' t
(J ~.0/2 \) .Q) <.:: '<-: c:: Q) 8 c:
"'- t .0 15 ~
-~.016 ~ NACA I "\ R ~ ~ ~' 1\ \ iy-Q' deq
" \
f7 \ ;>..
I'i 'f ao, / I_~ J
v
I (>.....p ir~ II ;:I If / ~ II ottocx, I ~ I-,-:l-- V V II II lei. of 1/ II II I / I rl / 17 /
V /
0-8 I
c
/ II II
V
j / onq/e 17 16 II I~ I .-< ~ 7 II II I / I I Ii I I I I I -8 V II II Section I I I I I 0 I II II 3S 45 6S ( 10 55- eft I IT IT 1,( II 17 (deq) v c> <I I I 0 o 020 ~ I I j I 16 II 1/
I I I
I~ II I I .- /1 II
rz;-I I-
1M
'3 ,
Q " 4
0 o -24 ·8 .4 -.8 24 20 1.6 -.4 J.6 28 3.2 J.2 -1.2 -/.6 \)
C c:: 8 >:: -2.0
"" ~
.... :~ <;:::: 'Q; :g
<t:
I
l'1 "Ii >'3 ~ >- >'3 ~ 0 ~ Z ~ 00 ... z I-< 0 Z >- t"' >- t:; <'l I-< U1 0 ~ >-<i 0 0 ~ ~ I-< >'3 >'3 l'1 l'1 "'J 0 >- l'1 ~ 0 Z >- c:: >'3 I-< ~ ~ t~ 0 U1 .~ I
Z » 0 » en (II (..) N .... en
1.6 I In.
In II /,2
P / ~
I I / II IL r
I I
10 lI- t / I 60.o V , I .8 /,
6.. ~V I II
t. 1
rr.( .
I Cl ~ 1& l~ 1..(1 I;l; I .OJ~ .032 ./ I deflecfed 1.0.Q28 .4 I I I
. I
'I ""'1"-' I roughness :'-o'uc)h';eis flap , y/c I , -.03/ -.027
'"
..- I-C I
r--I- -0.040 coefficient, split 0
t
-I.
position .8 I lift ~ I
--
r-- .rIc
a.c.
Sta~dard
Standard .264 ~ ~ 1.263 I, I I ). I I
- I----
~~ .6 -.4 ~ r.J..,. I simulated I Section \ iti I I I J <tic ~ "\ I R
\ J.~rr-_~:264 6.0 9.0 6.0 a20e ~g
\ ~ .~ i
.4 o ~
o o ~
-.8 R I'.
\ rho ~ ( \ ,\ ,~ Iq -/.2 .2 1\ 1<:;1\
f..-- ~
'-
I~ . 0 , r , ,
o
-1.6
.;}, -. -.3 -.5
-.2 -.4 -.2 airfoil section, 24-inch chord.
ti
l
.0/6 .0/2 .008 .004 ~O
.024 (,/ ~ \I) Cl v g
~ -+-~ :~ ~ ~ ~ -+-
tJ'.020 \:; III 8 ~ § ~
..... .~ ~ {; :;:: C/) 651-218 ACA N the ! I , It ~
QJ
( J, b_ L::::,
~a::I 11- ~~
I J, .~ /6 lClEt Lt "!:b::: ~ I:L lJ:l l.& , deg ill [}J~ 12 11 I~~ :11 a• ~ tiC 1\ II( lA'~ ~ ~~ ~ Aerodynamic characteristics of 14"' lr!. ~ diE!!
-.L Ii i' I ~ ~ \ attack, j!
\
Vl , 1\ ~ ("I
I I 0 II til ;,t IV of /I /, 1 '5/ 'f' jVh:.!.
ang/e ./i VI IA H -8
~V
7ft.
U III: ::rL: ~ .1 lrlt ~ Section ~ -, -/6
'm
It.
\ -24 4 -32 .8 .
-.8 1.6 -.4 3.6 3.2 24 20 -1.2 -/.6 -2.0_
~~ iji 'b 3 Q III
....
-1..:"1.2 :~ <i::: "-.: ~ .~ "'" Ir)
./ -./ -.J -.4 -.5 ~ l-.2 ~ Cl I.l ~ E::
~
...... ..... ~ 'q; .....
~ ~ ;,. ~ o >xj ;,. .... l:d >-rj o I-< t:-' t? ~ ;,. t>:) t>:) ~
en c: I
» » 0') (.) ~ ~ Q)
z 0 (II
1.6 '- 1<;1 II h- I 1.2 'III> II r II f/ I i' I I
T I
(/ If I 60° II PI II-' .8
I
II It In-~ 1 I I c!
Ir .036 .032 .028 I I I deflefted .4 ·1.0 I I I I roughness roughness flop
\~
!lIe "- -.06u
r--_
-0.090 coefficient,
o
split .8 --I-:: position +059 ~ IO~ lift
- -, 1'-'. Iv, Fd
.rIc a.c.
c-tandard Standard I'....: 1.263 0.265 I l-265
'--- f--- f-
It.. K Q.-
.6 -.4 ,\ ~ ~ r-...
simulated T Secti0f7
-
'''i R- T ~ .rIc "- F:;: R 3'~¥1- 6.0 D.20c
-, 6.0 6.0 6.0
.4 <., 17 o o 09.0 C;. 'V -.8 chord.
.2 -1.2 ~ f-..
/' '-
o
-/.6 .2 -./ -.5 -.3 -.4 airfoil section, 24-inch -.2 ti .; £ .016 .012 ~o .024 .004 ~-.2 QJ Cl \j ~ Cl
" 12 ~
"'-" ;g "- .....
-
\.>
~.020 c: Q; 8 g'
.5!! .\.> :i: {) ! ~.008
.....
653-418 A AC N the I C>, ! I ~ :) P...
r\. P- f:l...
~ , ~ ~ characteristics of
P . :1l
I ~v- -(' 'Y lIZ r-'. I~ deg \7~ ;F IZ t> \ de V <>C" lI.f I)'
.tl
lf17
.Aerodynamic if 11l ~ J Id~ ~ { attock, I.
J
f/ IJ>
o
/, P, of w V-b.
J: j1 N VI ~ 'V'..
1// f) an91e ~ f \1:7 VI -8
// II?
/J .}
>tl "'h
IlL
II VI I'h I 1"'- 1/ Section l VI!
'tJ, I fd t t:.
-16 l t [ -24 -32 .8 -.8 1.6 -.4 3.6 3.2 28 2.4 20 -1.6 -1.2 -1.2 -20 c: ~ \.> \.> Q) <S
..... .~ "-- "- "- :...--:: ~.4 Ii]
~ .I -./ -.3 -.5 l!:.
~ ~ \) ~ E -.4 l-.2
"'-" :Q :;:: .... ~
I
"C o !::::I >-l Z o to:> >-l '=' >-l
~ ~ H::o- !::::I M CO .. z > ..... o Z > t"' > ;S U1 o !::::I >-<l (") o ~ ~ H >-l M M "J o !::::I > M ~ Z > C1 >-l H (") U1 I 1\ ~
z (,) ~
» o » en (J1 .... (X) o (J1
1 I I I I I
J
1.6 I I I I I I
~ ~ T1
I I I I I I I I i·
I I I I I I I I I
1.2 I I I I I I ll::
III L.c: I I I I I I I
I I I wb I I
fiT ..lIr I 1 I j
. I I
JJ If?
I I' '" I I I
__
.8
I I I III u 1/ v I I
A 1 I I~ IT1 IITIIII I I , , Cz I I I I .0321 .
1..0301 I I I I
I
tPP
1.0 .4 I I Iii 11 I I
1"
I I I roughness I I
i
ylc I I I I I T I I I -0.012 I I coefficient, +.
.8 position ,-.047" I I I I "fi:{!
lift I I I .rIc T ac.
1';: 267 Standard
~
I I I 1 1·266T-.052 1. I $:<!65 ~-I-+lll~J
I I __ I I I I I I I I I I I I I I i
It I
NIT _, .6 -:4 1\ I I 1 I I I I
'1
')-.....'b.I"M ~ Section II I r I 1 1 1 1 I
IT 1 I
TT~
JJ4SPF
;r/..c
~ I I I
I I 1 R 2.9xI0· 5.9 8.9
Q ~
I I I I
I
J
.4 ( 0 0 <> Ll5:9'1 I~'S -.8 I I I
I I I I I
+-+-+-+-1 .
I r I
I I
I I I I' I
.2 -/.2
I H--WLL I I I
I I I I I I I I KI I I I I I I I I I I
I I I I
I I I I I
airfoil section, 24-inch chord.
lrTl-r1
II
oml
-1.8 .2, -.3
-.41
-.2 ci ci
-
.012 .008 .00Jlllllllllllltl~jllllll ~ o
~ .0241 .016 S~21 ~ u ~ l:: a=O.5
'0 ~ ..... ~
.....
' ,Jl. c: ~ u ~ c: .~ {; ~ V)
.... .\1 :::: .0
I I I 65.-418, J ...
ACA N
> ~
the ., \ 24 \'
"
"1~
"'
\ r"-1 b ~ /6 >V y. v deg lh o:;!.
y:Y' «0.
- ..
1.!Ol .,. LP
- 'L' f9 Aerodynamic characteristics of IU Y /J attack, If .J of D"
"t
t / l angle I -8 lrJj I _- H ..
, "'"-1: Section ~ 1\,,1 't -16 CN" -24 i~
, I
o
.8 .4 -32 1.8 .3.6 3.2 2.8 -.8 2.4 -.4 "20 ~1.2 -/.2 -1.6 -20
~ S q, g c: u
..... :Q ::: <:::: '.:::: :2 ~
o
.I • -./ -3 -.4 -.5 ::t
-
c: Q) o u l-'2 ~ & li:
..... .. .\2 :::: ... ~
~ ~ ;..- "'l ;..- H o ....
rn o el o ~ t" t:1 ~ >-
~ ~ 01 I
z » o » 0) UI Co) 0) ... Q)
1.6
b. .1
V
..1 i
IL
lL
1.2 !j ~1=6 J
i
Lr.
60' p lonlb' .8 ~ h.
...... c{ .036 .032 .028 .
IA deflected 1.0 c .4
r-. ~;u~ne~s
Ie roughness flop \ vic ~ , , -.035
, -.022
r- ......
~ coefficient, split
-+
.8
"-
A.. ,l.~ lift
r--- r-
1\ a.c.position "'Ie
Sf~nda).d
Standard I J.'-,. r-..
\.278 1.274 ~t::::
- _.-
I': 1\ .6 -.4 simulated I
I't: 1'0 " I
" Section
I ~/c ......
I R , J'~rr-t-o.2751O.044 6.0 9.0 6.0! D.20c B.O 6.0 b.
"" r'::'.
.4 ( o '\? 17 o ~ 6
"" -.8
.
[\.
bK.
.2 -/.2 I--
r-
IV ) , I
1 I J .J 4
o
-/.6 .2 airfoil section, 24-incb cbord.
-.
-.2 -.
<$
~~ -
~O .012 .004 ... () 0_
.024 ~ II! ~
.... :\l <i::: ..... j-
-
......
r..".020 ~ Q) g- t:: u
.~ a {)
..... ::::.016 ;g i3L008
653-618 NACA tbe J; J. ) ~ I~ 'y) i J..
~
""X;::s Iv-
M;g:D deg
"
~ w,=t:'r"-'v/'- X'lli~ .i ~t\ I~;n IXo, ~O:!
~ Aerodynamic cbaracteristics of
A
Ie@' ~ti ~ I~I"" 19, IB attocK, ,..j _'"
:~ rr lj,l
II ~ of , h IQ tJ- d
r h
anqle JL 'fl II -8 /1 IA rt: ....
J!:J 1// I" Section ~ ,.
I I I I J',;~ -16
r-tV l'1lnl,l!Sl
, N -24
D--I-W--l--l---t++-++-HlIVH-H---r-H-:llm~~~
L
, ~ I ~ ~ , ~~++~~~~+r~TlIHIITI I'Cl~-LLL~~~~-L~~~~~~~~
?~~++++++~HH~~~IIIIID
-32 .8 1.6 -.
3.6 3.2 2.8 2.0 2.4 -I. -I., -2.
...
c" <1> Q) 8 <J Q)
• .G .;:; ..... ~ .1:: 1l/. it: ~.4 er,
o
.I -.3 -./ -.4 -.5 !t
-
l-.2 t:: ~ Q) Q o ~
1: ~
. .~ ::::: .....
I
t'"J >C >-3 ~ 00 t-:> z > >-l H 0 Z > t< t:I ~ H U2 0 !:O ~ 0 I<' .... ~ H >-l >-l to:: to:: '"'1 0 !:O > t':I ~ Z > C1 >-l H (1 U2
~ 0 !:O Z "'" > (1
~ ~ ~ I ~ II
Z 0 c» Co) c» (X) 0
,. ,. UI .... UI
I 1.6 1.2 I'" .8 I
i t'r; rv
r<>' /' c, 1/ .(J3Z .o,a .036 1/ /.0 .4 I..-l n I--. rou~hness b,\ y/c f.... I-- I-- ...
-.026
--
o coefficient, positi
.B -+-. tD.059
"" h.- ~
.......
r-- '\ lift
.rIc a.Co V'3ndard .264 .265 r-.,.. fV;:t--- 1\ 0.257 1\· L .6
t -.4
Ir)f'.
In Section /0' l"- zlc ~ Ie 1'-< R 3.0 6.0 9.0 .4 o
o <> A,6P -.8
-
.2 -/.2 airfoil section. 24-inch chord.
/'" ......
) 1 ·0 I ~ t 1 .
.
o
-1.6
-. -:2 ... -.5
Ii d 0/6 ;:,) e a=O.5 ~ .00, .004 .024 I.i ~
~ \J ~ §
...:- :Q :t: ..... ~-.4
rv~·020 III H' c:: \J
..... .£ ;g ..... {5.0/2 ~
:g
653-618, N ACA the J: 24 ~. ~~ .
"'- ..... :-"1''''9 rN cd. )-< ~ 'Y 'fA deq
~::v .r
•• ~ IX lcil l~
•
.9 Aerodynamic characteristics of lIP- )/ attaclr, I~ V d of ;r I } , If l angle I ~ -8 I I }-[4 l.): Section 1'1: .....
t '-!
\ -/6 It ["'c -24 0 -32 .8 1.6 -.4 -.8 .16 3.oE' 2.8 2.4 2.0
- -1.2 -1.2 -/.6
-20
... ~ \) S \)
..... .~ ~ .....
.... ~ ;::.4 Jl
.1 -.1 -.3 - 4 -.5 11.
-
l-.z
() \) t ~
.... ~ ~
.~ q, ...
UJ c:j a:: ~ :> ::l ~ :> H !:O "'1 o H t" I:;; :> >-l :>
I:.,;) I:.,;) ......
I
z UI
» 0 » en ~ 0 I\) .-
1.6 1.2
r
"'V r- 1.
1 L
60· I I j, 'i" I L 1 1 I I 1M rJ IT 1 !
/ Cz
iLll 1
.036 .03Z .028 Cf deflected 1 'I 1.0 .48 :1=;{( 1 1 I I I I ~
i roughness rough~ess I
lK ..... flap 1\1 .I , y/c I 1 -.025 " \ I r-- r-- -0.017 coefficient, splil 0 position .8 I c lift
-- J:/c.....l.-, I
t-- o.
SIandard Standard .267 ~ ....
.......
0.271 1.z70~.025 1 I, I) l\ ~ I
- I--
.6 -.4 simulated r-...
1 I
-
\ Section
~ '\-, ,lI ~~
1 1 1 \ iit:/c \ I~ ~ I R <: 3.0xIO· 6.01 9.0 6.0! D.20c 6.0 .4. Ii'
o o o t;,. "l6.°
h" -.8 \
\
\ \U .2 - -1.2 ~ t--.
~ "--
') , I
o
-1.6 -j .2 -.2 -.3 -:4 -.5 ti -.2
i'
.016 .008 .004 .024 .012 C.> Q) o <J
S ~ ~
~O ...... :Q ~ ...... ~ :<j.020 ~ .. t:: , t:: III C) \) ~ \) .~ i; <%
..... .'l1 :::: :g
65.-021 airfoil section, 24-inch chord_ NACA I I I , J the of 0 ;>,. Q:.rc ~ , [J 24 '-\: L'l h rq \ , - -I..; ~~ b ::,::, I"" P= .(: ~c .p r-r bC W> ,c I ~ I ~ L6 deg lftil K.. ~:t ~c Y 9l ~ ---- - «0' ~ 'f ~ ~ Aerodynamic characteristics I~ ~ lr\! { /J f !f A a/lock, '1/ ~ "'~ .L
or
Iff r; 'tr'
'l
..
III (f Itt A) IJ 1I':t;o.
U
.... ~
A! angle jj 't/ t.P ~ t; -8 IlL ~ I'2sl IL Ii \( lA di J ~ Section ..d 'n lJ -16 r-t 9::tQ \
ro i\
-24 .8 -32 /.6 -.8 3.6 3.2 2.8 2.4 2.0 -.4 -/.2 -1.6 -2.0 ~ <J t:: \) <S
-+-.-1.2 .§ ~ "- <:::: '" ~.4 ~
j -j -.3 -.2 -.4 -.5 ~
J o <J ~ 'E
'()
r <;:: 'Qi ...... ~
t".l '1:j 00 ~ II>- ~ H 0 Z > t" > t:::1 <1 H rJ). 0 ~ Ll 0 ~ ~ H >-3 >-3 t".l t".l ..", 0 ::c > t".l ::c 0 Z > cj >-3 H Ll rJ).
::c g >-3 Z ? ~
~ ~ 00 I .... ....
Z » g en UI I\) I\)
1.6 1.2 - ITI I? , -'I .8 1: ~, R 'I +-t-f-+-I I r'" 1-t-t---+-c1t--1 1/ , /..
/ 5;.~H-'+-H-~ c[ I .O~~-r~~ .032 .028H--f-+f-t-1 / I deflected /.0 .4 ...... H , +-i-t-+-t--+--+-+-I +--1--1-+-+-+-1-+--1 0 I I " I rouqhness rouqhness flap , v~ yje 043 052 050 I ~ ',: [--t- ....- -0. I coefficienT, split .8 ..l. 1-: 1-.
posifion+-i-t-+-t--+--+-+-I I ............ -- I lift "'Ie STandord ~ ,.273 ,Standard J274 I - 0.274 't~ ~ , ~lr ~ .6 \., ~ -4 simulated , 6 I Section 1\ K 10 I I ~, J:C "L,), Ox O+--t-H-+-++--t-H-++-+-+-~-+-1 I R 6.0
9.0 6.0..!, o.20c 6. 6.0
lb."
I -+-3.
.4
'\.It- 0 o
~ -.8 ~ K .2 -1.2 --I'-- '-- I./V' V 1--1--+-+--+--+-+-,1-+-:1::--+--+-+-+ a.c. 1--1--+-+--+--+-+-11- H-+++-H-t I--I--+-+--+--+-+-t..c. 1--1--+-+--+--1-+-1."7 i--H-+++-+-+fl 6 'r--.---r--'--.---r--~-'---r--~-'
o o
) !~++~-rr+~-rr+~-r.r+~-rrT~rH '~~~r+~~~4-~~4-~~+-~~+-~~-t-1 ! '~~-+~+-~~-+~+-+-H-+~~+-~-+-r~+-~~ o -I.
.2 airfoil scction. 24-inch chord.
-,2 -.2 -.3 -.4 -.5 ti
~O .012 .004 ." -
.024
" ~ Q) Cl U
i
...... :Q ~ .... ~
.::.020 t: <ll 8 8' t: u
..... .2> .>! ~.016 {; :g ~.008
651-221 NACA the ""<.)
1\.0-
"\ 5.J fI'(l~ ~ 4:.~~ IX ,...
~ ,b'dOl, :p h.<;; ~
Vv
:z l~ ':f" L,..;; II deq rJ' In ~ R ~.r lXo,
l..l / ~
Aerodynamic characteristics of
i
fN' If .y' r<
Iff ~ rr ~
attocA, IlL Ir' N;l.
II III 'V'-h., of 'I 1/) '£S F ,fl II 'I \1- IP ;p angle . ~I 1i 1.6 rv.
~ Ii A'i 'J -8 II :R A 'v.;, '-1
II
!J .rf 1'\7 Section I II In' ~~ -16 Id Lin, Ll,( t Jq IP -24 : I ,
i , I ) I 1 ~ ,. ?
.
4 0 -32 .2 .8 /.6 1 -.
2.0 3.6 3.2 2.8 2.4 -I.
-I. -2.
§
cJ' 8
·91 .U <0::. ..... '-i::.. ~
1:- ;:::: :.:::
o
./ • ~5 -.I -3 -.4 ~ ~ QJ \) ~ ~ J-.2 o
:Q ~
..... :::: .....
~ ~ ~ >-<l o ~ >- 1-1 ~ o 1-1 t" t:l >- "" >-
U1 q !:l::I I:>:) I:>:) ~ I z » » 0') CJ1 .... ~ N ....
/.6 1.2 ) I-+- Jj 1l:J IV IU~ I ~ .8 'I-~ I I I" c, I~ ).,., L_L .028 .036 .032 I~ I deflected 1.0 .4 V I I I I I ~
r--.-
roughness roughness flop ~I " y/c
, I
" -."076
I-- .--
I -0.065 coefficient, split J. 0 -+-.046 .8 t-.... ~ I
r-- -- I~ I lift
:cjc a.c. pOSition standard '\;lIS" .272 .272 l);; 0.268 l fA Ih Iri 10.
~ .0 -;4 simulated "1 IIIStandard (" Secfion I
I
IIClc '-11:4.
R 6.0 JOxIO' 9:0 6.01 0.20c 6.0 Ig ~ LI .4 V 6.0
o o o b. "V
-.B .2 -1.2 _L-
,..... -
.............
V
o
.2 -1.6 airfoil section, 24-inch chord.
-./ • -.J -.4 -.5 -.2 -.2 ti
i -
.024 .0/6 .0/2 .008 .004 ~o \,)
~ III C> \) &; ~
:~ ~ ...... ::::: ......
\,)",.020 III
8 ~ ~ o
~~ :C :;:: ~
:t is
65,-421 .32 the N ACA of
R
i '"'\~ I , I ).; :r>-, .......
::r-
/6 <'"fJ C"" -/''f- :r f7 } , \ deq 1\ l.,? V ~ .lJ tXo, l' L,d=Y Y'/ Aerodynamic characteristics k ~ ~ I"J d ]'~ ~ ,d~ 1.(1 aftack, ~ 'di7 1f:J, If ~ r< I of 6' 'I il. ~I I}' ~p...
'/ 'I n.
~ ~ angle 1/, '7 IL If Ii -8 Iii 1/ IY A JI 'I /J II / Section I / lv< -/6 M
l-
1 rrtY' \ .,.., -24 -32 .8 1.6 -.4 -.8 3.6 3.2 2.8 24 2.0 -/'2 -1.6 -2.0
\,)~ c: III 8 0
~ ~
",1.2 .1Il .\3 't... <::: "" ~.4
.!
-./ -.3 -.4 -.5 '-t "' /-;2 s::: \) Q) () \.) ~ ~
~
..... .'l,> . <i:: <0..: .....
, ::c M '"d 0 ~ >-:l Z ~ 00 I):> "'" Z :>- >-:l .... 0 Z :>- t"' :>- t:I <: .... rJl 0 ~ ~ ~ ~ .... >-:l >-l M M "i 0 ~ :>- t"1 q >-l rJl I ~ /I
Z > 0 > 0) (JI -Ii> ~ N .... 0 (JI
/;6
1.2
I _I 1,,-(
"""
~ t:C .8 ..
.r: t c .03~ .O.JZ .oc8 k::{ 1.0 .4
"
I~ roughness yfc.
~ -.011 -.004, -Q084 ~
--
coefficiet7t, \ position .8 r-.... ~ "0 lift I'" is..
r-- I:--- *1
a.c.
Standard .271 .272 "lo 0.266
l'h h ~
I.- .6 -.4 I\, I"~
-
\
- Section
\ ~ OE/C
1\
R 3.0xIO~- 6.0 89 6.0 "-l', IQ <~ .4
o o <> 6
-.8 ............
-1.2 .2 section, 24-inch chord.
-
l"-
V 'J [
I t ) ;
o
-1.6 .2 ~3 ~ -.2 ~4 -.2 6 020 I!.'; ~ .024 .0/6 .008 .012 .004 o
~O (.j S <u o ~ o
..... :.Q :::: .... ~
.... ~
c: III o <.> 8' c:- o
<,J
.... .~ .0 ::: {; .0 ..:: ~
, I
65.-421,~a=O.5:airfolJ I
I
~ACA I , , , I the
" r"-I \'"- I
It 14
IIIII
(>., ':fb-6. I ~ I
w , , , , , , , I
deq I
IIIII
"., I ~ ~ ..,..,. I I
::rv ~
Aerodynamic characteristics of tv' I~ltf I attack, J I
JIt
filii
of IJ ~ I
11I 1I I I I I I I I I I I I I I I
H
~ -.J.
I ~ angle I -8 !
1}If 1111111111, I
km
I Section I
III I -/6
I
t4
'---OTT'): I
I
I -24 "
n
I
, 1111111111
0, .8 .4 -32
1.6
3.6 3.2 -.41 -.8H
2.8 2.4 20 ~1.2
-1.21111111mf -1.6'""""""""""""""""
-2.0, <S ~ () ()
..... .~ ~ C!:: :.::: .~ ..... ~
.I
-./ ~3 -.4 -.5
~~2 o ()
~ E:
-o.!' ~ ..... ~
~ ><j ~ t" t::l
[f1 c:: ~ ~ o "i ~ "i o I-! > ,.., :>
~ ~
I
» (') en U1 N - (II .... .... 0l=Io
z »
,.... '-' 1.6
•
fO II IV '/ I 1.2 I ) I.R ;>-?
I
f
Lrl/" Id p::; I 60° ./ ;'! ......
J
il Ir-l" f\{ V 10' V I I t c V D' ~ I I .036 .032 ..
';:f I 1 deflected 1.0.
.4 ~ I ·lcB I
1::::= L1 ~
rouqhness flop ;"o~qh~~ J \1 y/c I
\
-.0.27 , t::::: I'- 1 1 1 I -0,020 coefficient, .1 1..111.111 0.
split .8 -l- posdiOn ~ lift j----- ~ I 1 l- :ric 1.
a.C.
Standard Sfo'ndard .265 jQ-b-\ ~'i. 1.264+,0.27 !
- 0.261
~ Ih Io.lh i"'" 1-' .6 -.4 ,_ t-- simulated S Section \ ~.+- b I Q...
_I .rIc
~"
L"tJ.
I R
.c
-lo.xlO 6.0.
D.2Oc I" ( 1 .4 f76.o.
o 06.0. <>9.0. ,6 "76.0.
-.8 1\ I~~ \ ~ ~ 1"-1"\, \ ~ 1\ l\ \.
\:{ .2 -1.2 I-- I---
lc::::: r-- )
0.
-1.6 .2 -.I ~3 -.4 -;5 -.2 -.2 ti
EO -
~o .016 .012 .008 .004 .0.24 (,j -114 airfoil section, 24·inch chord.
S Q) o l.J g
.... :Q .... .... t ~
".020 -
\) s::: (lJ o \J 8' s::: .... ,~ ;g 't... {) .Q ....
.... .5{ (213)
I 1
32 C A 1 1 I 1 I 'I
!!
I I N A I
I I I
the
I I I I I I of
i I 'I
I
I
I
f\. 1 1 1 1 1 I I
I I I I I!
/6 '<J characteri~tics
I I I I I I I I I I I I
I I ~ deg I ,
h-, I I I I W
JP
"0' 1711 ~ I !
1>51 IIW~I IJ/ ':[ I-I I , Aerodynamic I~ ~17fR I IQi It' ,F{ I
'fTI1'111
ttt-+-r'I--I~+-t-+
.TIK
~ w tl I
I attock,
TI
di
I l?: I I
I I I I I I 11
0.
.jfICLIl.li~L1Jf of I I
II I
flll·11
~ 11/1 ~I I
[If! 19 ITI I
---~-r-~ angle
VJ 1 1 I !
I/! 1
-8
w I I I I
~II~ III.IIIIIIIII!I'
111 1 I I
++-/._\-+-+
iii I
Section 1 1 L\' l:tj/\ I -16 J
1 1-+--4-1·1 tm
I ,-+~T I
I I I 1 I I I I -24 1 I I 1 1 1 1 1 1 I I I 1 I 1 I I I 1 1 1 1 1 1 -32
-.81111111-1-~11imttti=+=H~tll
-.4 .3.6 -/.21 -1.61 -2.0.
c: ~
;::: '" .Q "f-. II)
0 5 .I • ..
-,{ -3 -.4 - :!!.
~
l-.z ~ Q) () \J ~. ~
.... . .\,1 ...
:::: ~
-i
I
l:tI ~ ~ ~ I-f >-l >-l t'1 t'1 "'i l:tI M ~ :>- >-l H 7Jl l:tI M '"d 0 ~ Z s:> 00 "" "" z :>- >-3 I-f 0 Z :>- t" > I::i <: I-f 7Jl 0 0 0 0 :>- l:tI 0 q 0 ~ O:l Hl- I ~ :::r' :!! I»
» 0) 0) .... I\) .... 1\). ;+
Z g 'a
J
3.Z 2.8 2.4 .
.., ~
2.0 ~ l c /.P-' k'f' ~ 1.6
"r:;:.. 1'If
. LAsij ..-I'=f- 40 50 70 ;:'.. 60 Ln Of coeffiCIent, 1.2 (degJ 0 0 o /!.
lift
r--
R=6XIO".
t..-- .8 ): Section split flap.
O.20c .4 t:'"~
o
-f- akfoil section with F5 I~ -.4 66,1-212 / I~ N ACA • , , , I ~ 5 -.8
o VI .I 4- .7
.4 .2 -.3 -.7 .5 -,2 -.4 ... -.6 -.8 -.9 thO 't ~
l-. ~ 8 t: ~
....
.~ ~ "qj '\- ~
-
- ~ I- f- 0- I--
l- I-- -r- and moment characteristics of I 16
-'-
IT ,-'-
J,irt l(- / \ , II '«0,009 it:
£t-
L ~ fJ ~r{J' ){ I ":I.
IN V./I attock) l' ~ liN :tJU C5 of·
I ¥/x-i
liN I f(,11 1'fI II I ~ YJ/I?
I1(j angle I '(j<J( I ,.I 1;:'1 -8 V/, ~ fIJI . I I IK I 'IL:.L [ I~ l' I Section I l+-.
-16 - r ~- r-
r-+, r--t-1 r- r- r- r- r- r- r- r- r- r- r r- r-
r-r
)~ ) 7 o 8 .6 2 .4 8 .4r- .8 .4 -24 I.
-.4 -.8 ..'i 3.
2.8 2. 2.
-1.2 -1.8 :..2.0
...r <II 8
-r1. .1l1 ~ ..... :::: ~
.§ '\-
.;t
U2 c:::l ~ ~ :> !;d ~ o '='j I-< !;d ~ 8 t< ~ >'l
:> :> I\:) ~ .<:.11 I
z » 0 » 0') 0') I\) (J1 0 - 0')
- .-
-
/.6 .n I ""-l 1.2 V <; I 7 .1/ II 101
1"<
), 1/ t.
.rl II c .8 ;/ / /CI: c .032 .036 .028 j//:.
/.0 .4 ./~ /l ~~ VI 1J y/c 17 033 " -.022 --- -- :;s: coefficient, -0.045
J. i-·
.8 position I q.-
r-- I----" 1 lift
;rIc a.c. 261 "1 - 1.
0.263 1~ ,\) .6 "1~
'\ '"
-.4 I Section
h, "- I
x/O' 11
a:jc i'.- R 6.0
3.1 9.p
"'
.4- o o <>.
'\. -.8 1\ .2 -/.2
~ -
V I-J
airfoil section, 24·inehchord.
o
-1.6 .2 -.I -.3 -.4 -.5 -.2 -.2 ti ~o .0/2 .004
.024 Jci \J o
~ t:: -J..:' ~
;g ~
ij'.020 ~ III 8 t:l-. tJ c: I.J
-1-- :(j ~.016 iJ :g ~.008
66(215)-016 l- NACA :J 1'1. ~ ~:J \ R'- rp., \
"'"
I I I I U:I' ,1'\ I ~ - /6 ~~
rr
/' ~ deq lA l/ltJ ¥ «0' Iff, ;;r 1/ ;or. Aerodynamic characteristics of the f:l aftacK, I - of rJCI
Id ..,-
j angle J
r
-8 i7 Section rt-r l\-.
"' -/6
t \ \ L~ \=-£ ~ r~ -24 r ) ! I ) II 1]. 0 .8 -32 /.0 -.8 3.6 3.2 -.4 2.8 2.4 2.0 -1.2 -1.6 -2.0
" <U 8 c: o
-..:1.2 .~ .\J <i::: "- "
:<-. -::::: :E..4 ~
o
.I -./ -.3 -.4 -:5 ~
-
l-.2 ~ 2 ~ 1::
.... ;g ~ '- ~
!:C i:'1 "d. 0 !:C >,3 0 00 .", "'" ~ > >,3 .... 0 Z > t" > t:; <! .... rI2 0 ~ (") 0 ~ ~ ~ i:'1 i:'1 I%j 0 !:C > i:'1 ~ Z > cj 8 .... (") U1 ~ Z C/o:) 0:.
I
» » en -- 01 .... en
Z 0 en N N
-
-
1.6 I i , I IT' rY I I I I III I
P
I ..f 1.2 T I J b I .1 II II I ~
II II
r
V .8 II V V / V If c, .036 .OJ2 .OZ8 II'b: 1.0 :>- t-- y/c I.m -.096 ....
:-<,....
-0.123 ~ r-- .
coef''hcienf, position .8 I\" Ci.
I ti!'J..
lift .r/~
r-- f.--
G.C .263 \ '-:P 1.26 ,
1'-."1: tf.26
-.4 ~ It;J.. I'" 1_ Section i'.
~/c
:\ n
R 6.0.
\ 'It> 2.9~fT= \ o o 0.8;9i -.8 1\ rt. po J.
, 10 1\
'i\ J'"1i-- 1\ 1'-': \ h 1\ -1.2
--
-
r---
V
? airfoil section, 24-inch chord.
.2 -1.6 :5_ - 1 -.3 -.2 -.2 -.4 ·0 u
~ti -
~o .024 .016 .012 .004
"
~ ~ (J % §
.... :t; ..::: .... .... ~
-
~'tJ.020 t; &' t \J
B -G .g
..... .~ .;::: 't... ~.008 66(215)-216 ACA N the ) \P )..~ r\. 3- tl \
\
~~ ~'3. :\ ~ ~ I<Y 1-.0, '1' ~ l/ '.r j deq 'J ~/ "0, lfi;5 J-l/ IJ Aerodynamic characteristics of
--
Id 9' attack, ~ '( '-- j if of, f.
- 1, 1/ ~ angle VI -8 -A~ (I '--- A fl ~ ~ Section I I !
I .''1:,----, -16 Ib II' ~ ,- l[\.
L- -24 L- L.
.8 -32 /.6 1.2 -.8 3.6 2.8 -.4' 3.2 2.4 2.0 -1.2 -1.6 -2.0
~N ~ 8 8 \.J
...... .;::.4 Ji
.g ..... ..... 't... '- --
.f -.4 -.5 -./ -.2 -:3 ~ It ~
J- e ~ ~
..... ~
.~ ~ 'Q; .....
~ ,~ .~ t"' t::) > >-'3 > > ~ ~ 0 "'1 > H l:d "'1 0 H Ul d ~ ~ I ::e .... ::r = Q) 01 N - en » en en N - 'C Z » 0 - - , ~fJ'>..
:.::.
~ 1;)-:-- P" - ..... c, ~ ~ .032 .036 /.6 '1"- ,.-:' /.2.
I- ~ - !lltI ;r~ '0, 1.2 19 1/ ~ 1/ ~' coefficient; I'll III yo :!-£~:';Z ~x lift R.- k I rt;; '\t\ R=6XIO'.
{ ~ -< :;.; ,I },. Ii' ';> ',0 .8 ,~\ '.g','/'F VI ~ .8 "- .h I1c / flap.
.k?< 'V'--<:7 :r!
I /L f!f-'?
Section "- ~ - .-1- ~T I IA.
~ plain 'I::L II .4 - ~ .6 :z/c l~ (iT - I f' ;(-' sealed 7' ""!ti O.20e L o _ .4 y: I'1T t'-- with _ '\ :-c L.
~s--- \ ~~ - -.4 ~t'- .2 In: Iu f\...
'- f-.- I---.
~r- v -'<- ~ -- ~ -.8 I , I'-- 0 o ,le- -.4 -.6 -.8 -.2 -1.0 .2 ~2 ~ 66(215)-216 airfoil scction () .004 QJ () \.J ~ E: .016 ~ ~ J ->:: ..... ....
.024 .020 ...:' ·.iJ ~o t::: (,.OOB 8 8'.012 <J 'q; ·2 S5 f '.~ .;:: is NACA 32 the - -j ;r '=1 ,\1 ~ 0 f&.. B:, ~ p; fi - T characteristics of ~ hliI.
---I ~ ~~ ll..:!;\ ~.
'-' deq /6 i=C' ~ I-,.l.
r fl' J'!t ~~ '"' p..;; J - - cr., II Aerodynamic LI II ~~ ~ I>: \) p: hv~ ~/ I{ /fk' -_ ,. I I 8 .j II ~ ~ :r/r 1<1 ,/ attack, If.
II "I 1I11 ;:\:/- 'j Ilk' ~I 1/ "/ ~ ---1- I 1I1I I '" .
:>1 1.( 1I1I 1/ of rlf/ 1)1'« l? I~ V (I LYl( I VI II k' V III ~ -; II J .i I>M IIV IJ-'Irl 1;< angle 1/, ~1rT IV Id IJf1 ~ l)-' 1<H10 ~I Xc ftl~ IAI) i:Z[ I~ f{f7 I /, -I II ~ !> iI IlrZI/ II II [ -8 I 'j Section .7 'I :J~ f7. w-l..J..b~ 1/ Inl rz O· 40 50 -5 10 15 20 25 30 60 65 II r7 II 0, I j, A Q. II II (degJ <l V' '\l Do <)..-10 6...;-15 0 <> "" \1 I> ,( -/6 , IJi 1-1 - -24 - I- l- o -.4 -.8 -/.6 -/.2 -2.0 1.6 20 \.J 2.4 ,..
2.8 8,8 " ~.4 J!
3.6 3,2 C ~ \ .... ~ ..:1.2 .~
I
M 00 l>:) "'" z ::>- .., .... 0 Z ::>- t' ::>- Sl .... U1 0 ~ 0 0 a;: ...... .., .., M M "'l 0 ~ ::>- M !:tl 0 Z ::>- L! j 0 U1
~ >-t:J 0 ~ .., Z ?
t..:l CI:i 00 -a;: I ::::T Q) Z l> ~ en en N - UI N - en !1. ...... :::.
"C
-
-
2.8 i I r I 1 I , ~ .
'IL ~ 2.1.
,q '10 2.0 ., c[ 'V , f 1.6 I
L I
40 60 50 70 I lri.-l. I I 0, coefficient, /.2
h.. a- (deq) I
I I
o o Do 'V
() le lift R=6Xt06.
l~f-, If''I N'i ,~ .8 f\. I- Section DJ, LL..-.l- split flap.
I" O.20c L.J .4 w;:J-.
with R C-.J t--...
.
I."., r-
.
:-
o
N ~.
airfoil section ~
'"
..,4 r--...
66(215)-216 [V' NACA -;8 o
.5 .4 .3 .2 .I •
-.3 -.4 -;9 -.5 -.7 -.8 -.6 :t.
~ the
l-.1 OJ o () ~
~
.§ .~ .... ~
..... ::::-2
, 1 , ,J characteristics of h R f4.' 1\ .l~ \\ ~~.
11: moment [\ -H It) and L( .~ / ~ ~ Lift deg .L b§ ;j-'ri:l~ :1 ~~ I lt~, f£ V ~ 1/) II P!I ~I.;> II V L~lJl rl Vi . ).l ottoch) W I '/iSI I/il. of' il.
V II IWI Iq ')/.
~R I'll -~ 'I. II ~'lJfI I (!J(l rJl L ( -8 'I- I I ~f U<.rf jl'j ') '/ ~l II ~ I If') II, ~U, cV Section on9/e . fL.:!'1 I '( I I -/6 ): II ~
A o
-24 .8 .~ I.ti -.8 3.6 -.4 2.4 2.0 2.8 -1.2 -1.6 -2.0_
·3.2 10: a
.~ .Il ~
1€1.2 \ .t: ::::: .S ii ~
l-v ~ <:0 t"' ~ :> ~ ~ :> ~ to1 0 "'J :> ...... ~ "'J 0 ...... l:j m q II I l:I 0 . ell ell N ... U1 N ... ~ Z » 0 » ell - - /.6 I 17 ,/ 1.2 I II II II 1£;1 <;: I I II /../ .~ 1/ .8 I "" ../ i f-' r VP z C p IJ .032 .028 .036 .4 1.0 ..n~ i-ooV t-- y/c -.061 -.04J q:>o- coefficient, -\. _-0.090 _ t-- !--- position ,1\ .8 9:- lift 1::.\ N ;xlc a.c.
r-- I-- .257 .258 "'" ,u., 0.256 - \'., -.4 .6 h f"'. '" Section '" ;elc R 3.0xI0· 60 gO L o o o -.8 .4 [ -c,.....
i\ \ airfoil section, 24-inch chord.
\ 1\ -/.2 C; .2 a=O.6 I-- I ;- I ?
I .1 t 7 V r- "1 -1.0 o , '0 1 7 _ -.2 ~ ~ • ~ .. -.2 66(215)-216, t Q) o (J ~ .00< .004 " ..... ~- .0/2 .... :~ :::: .024 ~O ~ I:::: I..)
,.:".020 ~ Q) () (., g- :g (J!, :G ~.016 il ....
NACA the I , I - -I ~ 1 -+ --; I J......L .-, , ,I I " ~~ [\ ['; ?
,'rh BY p, $Ii deq 1{ :v V) 0(0' Aerodynamic characteristiCS of fp 8 J V J., 1<1 attacA, - of f- .
f 1-1- angle -8 f / Section J , , , -16 -24 - ? 5 01 4 - g 7 - -32 :1 ¢ I-.
I I-. . • I /I "'- - -/ -I. -2.
I.
2.0 3.2 2.8 2.4 Q) 3.6 o <J s:: (J '"' Q) .~ ..... V) "E!.2. .~ .1..) ..:::: .... i!:: "'" IJ • ./ -.4 -.5 -3 -./ "" ~ ~Q ~ '" -.2 Q) o \; 1': .... .... ~ .~ .~ :::: l\:) H:>- o M I ~ '"d o !:O I--'l ~ ? 00 "" .... z > I--'l ...... C Z > t"' I:j H > -< m o !:O ....; () o ~ ~ ...... I--'l ..., M M "'! §S > M !:O o Z > d ~ () m I II z » o » en en .... !:l 01 N .... en N o en ~. ..... ::::r D) - ::!: "tJ I -- .lD .17 .14 .004 .
0 0 -.83 -.92 -.79 -.62 -.46 -.33 -.21 -.08 -0.58 -1.00 -1.00 Ordinate seal Lower snrface 0.5 2.08 1.04 4.16 6.25 8.33 lD,41 12.50 14.58 16.66 18.75 20.29 22.91 25.00 27. 29.16 30.00 Abscissa ---- .---·Rubber - coordinates Flap .62 .
1.54 2.04 2.75 3.75 4.28 4.53 4.58 4.33 .015-.020 4.50 4.06 3.54 3.04 2.06 1.27 rJap a Ordinate "Slotted Upper surface 0.5 1.04 2.08 4.16 6.25 8.33 lD.41 12.50 14.58 16.66 18.75 20.29 22.91 25.00 27.08 29.16 30.00 Abscissa ~ >i plain flap, O.10c .10 and ~oo------------------------------------------------------~ slotted ,~// retracted deflected /_---.025R extended c90------------------------------------------~ ''-;---:/?~;;' 0.30c deTlected k flap flap flap flap configuration, Plain Slotted (a) Airfoil-flap configuration.
Plain Slotted Flap _ airfoil section with (b) a=0.6 .30 .~5 lIne ____ --- I/ne ce 66(215)-216, ---- ~ chord NACA referen ------------------------- __ ---- I ----:....:...
Flop -4/rf'Oil ------------ I /Pivat ---.:....: --,,---- -"L- .
, ' (b) + I ~I ~.
(a.)
~ ~ I ( . I< I [go l ~ >- ~ >-<j "'I >- ...... ~ 0 ...... t-< t::I >- t-3 Ul c:j ~ 0 "'I >- t>:> t+:- .......
I II ~
..... 0 en =E ;:; = :::!! S»
2 » 0 » en en N ..... til N en
"0
- -
2.4 I&; ~ C?
1.6 0'[ 'fr A II >J""t IJV f-< [Jt;.
.028 I :036 .OJ<' C; 7V IYl ~ c, ;;'\ I I 1.2 /,2 /.
V b.; .C rll~ II \-
Ct- 11' l/
h II kit,.
I ...............
IJ ~
1....- I
f ~ ...
1.0 I 1/1 coefftcient, -,
--
....~ "-
J
t /' YI I"L f..8 I ""L Ii l1F1- [I [..(1 't21:b L- t::::::.. lit ~ fJ I--'" } .8 .4 ["" 1(') II r..
.......
r--- ~ t- It<-II
SecfiOl7 nap.
~?'
...... fl,( ~ ~ rq Irll I"\: o .8 ~ plain ",/e !'is ~ ~ lA
tx 18
O.lOe E'€ p.[ \ < and '" ...
r-. 1<
.4- -.4 ~ j'{..: 1'-..1 n- IS": slotted !'no. l" O.3OC t---. r\. i'n .2 -.8 Id ~ Slotted lIap retracted; R=6XIO'.
r..- ~
V
~
,1....- , , , I
o
o! -1.2 .2 -.2 -.4 -.6 - -.2 -1.0 :t .0/6 .008 .004 .020 .024
l .c:: ~ Q) ~ ~ §
~ - .... ~ airfoil section with
~ ~
<J !:: 8 ~.0/2 .c:: II
is ~
...... '.~ ;.::: "Q5 :2
a=O.6 Aerodynamic characteristics.
66(215H~16.
(c) >-,
ro: ~
\ '\<ty NACA
r'\ rH\
\.l..
.
,\
~:;>... '""
R..
[ 'iI~ ~~ I}ty r\ deq 1.
;:(1./. 1/ C( {(o, ~ l.( ~:? (- I ../ 11-, 1/ j if, L( J [7 II 1/,110 ;:> ;t( arrach, ~~ V
VV V 'UJ
)
:1 of
>C>, ,,.< [2.111 I( Id I - ~ 1!j2J ~~ ~
o
J II
il 19 II Ir ~
angle j II I/j ~ ItJ II l I ,J, III
VI II II
j ) IJ If,
"
I '(f II -
( [
I A -8 II r7 Section [ (
V
30 ) ~ -10 -eo (deq)
0 o o I:>. q (). A-3D
;q
\ \ , -/6 'n (e)
I , I
i 3 ¢ -24 4- :> .8 -.4 -.
2.4 2.0 1.6 28 -I. -I.
3.2 -1.2 -2.
3.6 Q) !::
~ c: 8 .~ /}j
.~ .g ~ "to
..... "'- <!::: \':"
•
I
~ ~ I-< 1-3 1-3 t'J t'J "'l 0 pj :> t'J ~ 0 Z :> c:I 1-3 I-< a Ul
II>- Z :> .., 0 Z :> t"' :> t:l < I-< Ul 0 ~ ><j a 0
~ t'J '"d 0 ~ .., Z P 00 ~ I-<
b:) ~ b:) I II Q)
~ :§. r+ :r ~
CJ) N .... tTl N .... CJ) 0 CJ) "C
z » 0 » CJ)
- -
3.2 / 2.8
-
::r
P= I( I-
!=!: 0' If ..,- §!' -- 2.4 .4 .3 :b .5 1>- '- tr Iq wI-
2~0
r ~ y.. --u
1.4 ~ L,'. 20 t- f1" ~ retracted
-
r- 1'-'
"-
-.-
l', flop ~ --0 111111
-
I
I~~ ~ I-c r- r- 10-
1.2 1.6 I-..\. r<:: r- It:
"
-?
.A
~ I-- r-
Sioffed p;- A.
-
o
10 II ~ -C !:':- 20 30 Of -10 -30 coefficient, -20 1.0 1.2
r- re r- t- (degJ
I-- :>-.
o () fJ. D- <>- 6.
\l lift
........., -z:'! -v l- D t-
h
-- t--
~ -
P-
.8 .8 .rIc I-- I"'- h: I'-.. ~ ....-:::
'-<r- Secfion
:--:-.. b-- *- ;?'
....... ....... Jo..... plain flap R=6XIO«.
v.
b= t-- r- rn
.0 .4 O.lOe ""-to- 220;
r--
r- r-..... I'-..
and
0- ""'
r-:-I--
N: ~ I)... b. -'--
'"
o
slotted - -(: (-.
[ r v 0.30c Vt.
rc: L_
Slotted flap deflected .2 -.4 , - -- )-.
r--
I'--- ?-.
.V 0
, , ) , r ; ) I
o -.8
.z
.2 -.
-.2 -.3 -.4 -.5 -.0 -.7 -.8 ~9 -.2 l,t
~o -
J ~ \)
Q ~. E: ...... .~ ~ 'Q; ......
~
a=0.6 airfoil section with moment characteristics.
and 66(215)-216,
_ -
r- r r-
r- ~~ (d) I.lft ...J N ACA '" Ii!, I II::: J \ /6 fu
LQa L
i§ lI1 [2. I£:L
'{.. ~
JL't<- ~C: ttl- ..L
("~ tXo,deg
lL II >+-11 lL lL ~ ~ J 'I,
lI! IL
IL .C~
[LIL L L
(£:t: lL attock}
LI
('IS1 ~
lLLJ
1illLL
of Ii .lUI [L
1. tllL L L
IL JL.
• f£.
If IL 1.1-.
~IL angle , .l.'d.
V, .J:I..
JJ ..,l~ -8 LLiIJ. ,.:
~ Jill L
'.1JU 1.1 ....s It'd w; l,{ J/...,t Section /& L jJjJ L -/6 ~ J V I
-I - ~ ,- ,- ,-- • ,-
,- '1 I r- .
.~
.tt~±ttth1=+=+fttft1itf±t$$$~~
S
4 o 6 ::> 9 4 o
d ."7 -24 I.
-.8 3. :i Z.8 2. 2.
-L -1.2 ~1.6 -Z.O
Q) 8
r.: .... .~ ~ ~
.... .... ~ ~
t..:l l+>- e:> ~ I-%j 0 .... t-< t:j ;.- >-3 ;.- ~ ~ ;.- ~ ><l 0 I-%j ;.- ....
(/) r.
II D) I ~, ~ :r :::!!
~ 0 en '0 N - en .
en en N - (J1 Z » 0 » - - 3.2 I v 2.8 v.'
.J: - I" ~ 2.1 I 17r' .3 .5 .4 I tel: I V' I --r. !t'll- z'r r' T t 2.0 refracfed I 1.4 / m I- I ~ .,6" I- - c[ flap I ;t- )--.
'" ~ I ~;- P'i-- 1.6 ~ I---- I n 1.2 "b~o,":.:::.
I I - --0_ I~ .., '-:.....
~, Slotfed I io ~ i-- I--<: a o /0 30 I I- r--: l- r'" Of coefficient, '- -/0 -20 1.2 La.. f'=S4II -'~r- I -- - (deq) /.0 ~ ~ o 0 o t:> a. 0. A-3D lift iLl ~ ..£..
.....
>0- ..,...
:;;...
r- r-- -~ h flap, .8 I====- .8 -- Section ~-- ...... --.r plain ...-- '+i-- r--- R=6XIO".
z/C r--.
r-- r->- ; I- -=~ , -- ...... O.lOc »-.. 27 .4 - s--- .6 r-- I-i and i'- 1'-.1 r-- 'IV IJ ..... , ...... ,~ slotted "'- o -'>--.
.4 In.
O.30c £ ......
r-- h Slotted flap deflected with 'U.
-A .2 I- f-- IQ ~ 1 airfoil section I( -:8 ---- V , a o -.9 characteristics.
-.7 -.8 .2 . I -.3 -.4 -.5 -.6 -.1 -.2 /r a=O.6 8 t § ... lb' ~ ;:,0 .~ .\J 'I-.
-1-- ~-2 ~ moment and 66(215)-216, "-1 "-1 Lift :., (e) NACA '\ ~ l.,.-' .
~ /6 ~t:j\r-+-+-+-I- J " ~ T I deg A 7 l'i :n' J •.
I 'f.
lI:1.~ l/ ~ « I V r' . li"l" I / ']A II . / 1/ l{ V 7Jfl ~ / rl h'l ottacl(, 'j. I r" I JlIIJI I I I I 0 TIT \I I IN! "q of I J '/1{ '" if II rn !Ill U I I lho ~ 'j.
/ Ifllll / P I -l- !hi fill II on9ie I.
S / II J I (f) IIJ -8 U I J.
/ IIA './/7 'i !), J,~ ~ lIl/ WI r v Section l ~~ [ cJ.?
-/6 ]'1/ I I C I I I (e) cv-o i ,pjjW:ti+tt+H+tmmttt ,~tttDDi++~HR~+ttw±t , _ ~ ril1-t-+--++ 1I1IT+++++44~~U~~ -24 tt~It~~~~~~~[lI1-T~~~+4-L , o .8 .4 -.8 -,4 -/.2 -/.6 1.6 -2.0, 2.0 2.4 3.6 3.2 2.8 r:: a ~ 1.0 .\:; ..... ~ :g .~ s::: 'q; ""{I.e
I
:>- t>j :>- q ~ ...... (") I>:l :>- ~ ...... 0 Z :>- t"' :>- ~ ...... U1 0 ~ (") 0 ~ ~ ...... ~ t>j t>j "'l 0 ~ l:O 0 Z U1
~ t>j '"d 0 ~ ~ Z 0 00 >I>- z
t-:) ~ ~ II I
$:I ~. ..... :::r C)
Z l> l> en en N .... (J1 N .... en 0 en =
0 'C
- -
1 I 1 I I h.
--
I 19 [,{ , po< t-t-,..
'k I I I I 1 I I I I I I I ! ! ; : 1 I ~ C I'-!"\.- L:.
2.4 [V 1
.5 I h'-
:r>
--
I 1 1 I P- t- <;< t'.
J i I I IJ I Iv 1 I I I P- , 1---1- li:\ t- In
32"_·4 )
~
I I I k k">. t- k retracted 11
- I'" \: I I I 1 I I +-
I :t- +- s=-
c, -J, I 1
I ' I I >- F to flop'
it'> ~r, I I I i I
;'~6,~·3
!.6 ; I t- I
'\
_._ -I--- I I~ ~, I .v., ()-- Sioffed i I I : l In t-- "h- +-
/0 o
I I ! i I , t t- ro t- ~ 20 30
~ ?- coefficient, Or -10 -20 -30
1.0 ---l-l- 1.2
, l=>- t- r-- (deq) 1 1 1 1 111-1
0 0 o 6. Cl. \l 0- A lift
~~ I ~ , i ~t-
I- ! . i I i I i l- b I I i i
I:::::... I I to
.8 .8 Lt-- .ric
I I ; I ! -to \'1=t- i'-t-
I I I I I ! 1\-
--
Section I I I I I I I I
r-- h
plain flap.
-=F" I R=6XIO'.
1- 1"- h- l;-- 1""1
-
I I ~ 1- O.lOc .4 .6 r--k i"- tA- and f"., ' t- ~ 'f"'y -A jQ t"- I'- I I 10' :j-..:. slotted
o
.4 ~t'L\ [U [D I" ~ O.30c i'C f'..- t--.. ~ "j-...
1.6 I' Slotted flap deflected 32°; I- I :> ;:.... Ir [U I( .2 -.4 1\ \ [)J ~ b' , 1\
V I'--
'} airfoil section with .I 0 -.8 .2 .2.
-./ -.3 -.7
-.2 -.4 -.5 -.6 -.8 -.9 -.2 ~ ~O
J
£ C) \J (:: ~
i-..' Q) a=O.6
~ ~. ~
moment charactcristics.
and 66(215)-216, I I I , ,
~
1 -t
Lift
_ -
11 -l r- r-
I r
-~ r-r- (f)
I NACA I f',)--t I
~(
I :i' I \ '.
l..l
;\~ -tV
!i I /6 J ~ S<::
I k"~
-L I r'" deq 'j.: .L I
-L
"0,
t 11 IL
'1f.j}{ I
I JjjjJI ('1' ~
I I li:!j1 ;t Iv attock, JJf.j).(,
ImA
';I c.
J of
1~ II
V I/" &'iJ J .1J ltd 1 L C
m
.
'S;
M
j II lULl!
'i'IJlL anqle '(fl'1 'I,L I ~ -8
JJ I
II c:.
U, it/,'l/ J II I IL / 'II -.S ',L I ~ Section III JJ '1> I 'J.
-/6 I \ r1' ~
]
I] DJ~+-~+I~IW~~~ 1- I r r- Uj~I/++tp~I±~~~+EBD~ W ~~=H=ttff±ft±±tt=R±ftt r-±±±+tHffiE±J±t+++t=Effiw ,(f)
,tn----l[tj~~+1=t!jJ±t=ij;MBB± ' 7 ; " / 7 1
4 o
-2.4 .8 1.6 -.8 36 3.2 2.8 2,4 2.0 -1.2 -/'2 -1.6 -2 ~ \J ~~ ..... .~ .\; .;: "- .E ~.
A '"l H ~ "'j 0 H t-< ~ ~
m q ~ !7 ~ 0 > 0-3 > J+:o. c:Jl
.
I II ~ en N - 0 en ::IE _ .... :T .... . Q;
Z » 0 » en N - t.n ~ . 'C
- -
3.2
-
-
~
-
2.8 ~ -, I~ .-b- 'Yp
"t: )- ::>
- -
....
loA \.I. 'U I ()-<
--
,
.3 -- 2.4
.5 .4 I .
b.
- ~t7-7t 1& 16.
I r-r-.
:- I
:17"- 71
L: "'"
..... 2.0
n-
l
'"
- fT-, flap'refroctfto c
I -6/ '+6,
--.,.- - -
~ I 1.6 ~ r'- -'l'l,.
(':- ~/.4 t<;:~ ~, h )....
L Sloffed r- t- N I ':
o
F
t'A J.. 10
['v' r-h 20 30
-10 -30 coefficient, /.0 1,2
- (dG'qJ
o 0 0.
o A 0-20 t::.. \l
- """--t- -~
~ lifT
...... --n.
r
:::::::... ~ c:...'
.8 .8 flap.
A_
r-.. r-
Section r-- ~ -I'. )0-. C>....
zjc plain ~ ~ ~
-
.= fb. "'t: ........
0.10c .6 .4 'i.; r-.. r-..
"- ~ and
~ r-.. t'Y
I" ~ I"-h ~
o slotted
.4 t.
c\... ~ K
<r-
0.30c '\.
P Slotted flap deflected 37°; R=6XIO'. with ~I'., I I -:4 It;: 'rl.
t1 \ I-- c-
"
~
airfoil section
vI'-- 20 C / 0 / c J 4 5' 7 '3 <j -:8
.
-.3 -. .., -: -. characteristics.
::t ~o ;:;; ~
~ ,,- ~
k .~ .i3 ~ \-: ~
-t-- .~ a=0.6 moment I and 66(215)-216, Lift p., (g) NACA h / b ~ ~ "K "
. ill
,.
~ ~.
if 1/ decj .
,r
h t-'- jW If r II
a o, ~-t I{ V
O?/u II
l7Cf f7 rt~ II V
C1~ I 1# II 1/ A } Ilk, (jr;J IIII II Id attack, / I ) I/JI I} II )..( 0 I of VIH ' I
~ll II r- 7 IF
I )
IZ ,!/,r! II
) I I -< fill 1/17 ~nl V II I' II < anfjle I I} '!&'I. 7' I I -8
1/1/'1 III} fI
If} ¥liT/ ' 'j , IK/C VI!. II
m
III '[ ,~
m WI I/, ' II
Sect/on
t <
I
,i
Ie.
II m V
I -16 'Ir T}
U7 AI
1,\ cf' {~l I
' o
-24 .8 .4 1.6 -.8 -.4 .3.6 3.2 2.8 2.0 -1.2 -1.6' -12 -20 ~~ s:: Q, ~ !:: ~ "l ,~ ,~ ~ :.::: ~ ..... ;:::
I
I':f ~ 0 l:tJ ...., ? 00 I>:> 2: :> ...., >-< 0 2: :> t"' :> t; >-< 0 l:tJ ~ 0 ~ ~ >-< ...., ...., I':f 0 l:tJ :> I':f Z· :> ...., >-< l:tJ 2: tl>- <: U1 ("") I':f "1 l:tJ 0 q ("") U1 ~ ~ 'FO I
Z ):11 ~ 0) 0) I\) .... U1 ~ .... 0)
- -
1.6
-f
I
io i .I
b J
I I
~ II I I 1.2 b II J I }; IP II S I J 'Til rr: .8 )J I:J
If [,V
l c .036 .032 .028 1.0 .4 -,,>- I-- y/c -.123 -./05 R
t-- r-
coefficient,
+- 0
f~/37 .8 position ~ c.
lift
t--- ~ ~ 1'0
zlc o.
f' !'- 1.266 1.265 0.266
-
t::t p: In .6 -.4 ~ I t-- Section I,.. I Q.. l"' xlO ~/c R 6.0 8.9 "\ ......
.4 03.1 o () -.8 r-< \ ::r- ; ,
r-
.2 -1.2 -l-~ ~
V r--
- __ 'J. airfoil section, 24-inch chord.
o
-1.6 .2 -.I -.5 -.2 -.2 -.3 -.4 ~ ~o .024 .012 .004 (1ti
~ ~ \.) ~
~
~ ;...~ :Q ;::: ~
;...
~".020 ~ QJ o l> \.)
~ §
.... :u ~.016 ~ :;::: ~.008
66(215)-416 N ACA I I the '1 ~ 1-1:> ).
'\
\. ~
r
\, ~ i!h"8 ~J' 7, deq .
'/J1
Jr
~ (xo, ~ !J .
.p Aerodynamic characteristics of B attock, /I ~ /I 1/ ~ of !J 6J A ltf>1 anqle
W
/ -8 /I Section 1/ I I J6 -24 .8 -32 J.5 3.6 3.2 -.4 -.8 2.8 2.4 2.0 -1.2 -1.6 -2.0 ~ QJ ~. \.) <lJ
.~ .!:! ~ ;::: :..::: V)
-+<1.2 :;:::.4 .I -./ -.2 -.3 -.4 -.5 It ~v 1:1 c:
.Il! 8 ~
...:- ;g "Q; .... ~
w C1 ~ ~ ;.- ~ o "'l ;.- ..... "i o ..... t" t:j ;.-
::0 :;:
~ '"'-l
•
z » ~ 0) 0) o o 0)
1.6 1.2 .
60~ .8 z J C L,.: .036 .032 .028 I Ale' deflected I J.O I .4 I .l- ""~ roughness roughness flop \f;~ ri.-!
ylc ~ \ .0(J8 , '-llf-Ll ldI coefficient, spJ;f
.8 -+-
position "" f-,c, ffi .~ c.
'IIi: lift .rIc Standard Standard a .258 .252 IJ- 1)- 0.2551"027 Z1 !'-I'r .(J II \ -.4 b I simulated Section rnl 1'Xth.
_I .rIc
III f'S
R 6.0 0.200 6.0 .4 f76.0 03.0></0 06.0 09.0 V' Ll.
-.8 - - .2 -/.2 ? , , : , 4 2 4 S .I
o -/.6
.2 airfoil section, 24·inch chord.
-.2 _ _. -.4
-,2 u ~ .024 .016 .012 .004
~o tJi ~ §
~ 8 ~
'<3 ~ ....
~ ....
"",.020 ~ III 8 8' c: u
~ ~.008
.... :g :::: :g
66-006 NACA the of IU !H; p>f'!
deq Il 1"E'l. oe., v "v j!.
Aerodynamic characteristics ~ l"J \.
, ~ IA IF 1,\ attock, , If' llli'
ijll 'Yli of
~.
I 9 I I onqle 18- -8 ~ 'Ll c~ l.
u'1 '"CQ M Section u -16 -24
! , r I
,
o
.4 .8 -32 1.6 -.4 -.8 3.2 .3.6 2.8 2.4 2.0 -/.2 -I.(J ~1.2 -2
\)~ c: Q) 8 t
.... ·91 .\.) .;;: .... .<J ..::: ~
~ ..::::: -.1 -.1 -.3 -.4 -.5 It.
l-.2 ~ <J \j ~ 12
-I..~ ~ .... ~
'Qj
I -3
~ ~ ~ ,." trl trl "'J 0 :> trl 0 Z :> c:1 ,." I-< (")
"d ,." Z ? 00 t.:> """ z ). ,...., 0 Z :> t" :> I:j ~ I-< W 0 pj (") 0 I-< ,." pj pj W
pj trl 0 pj tI:) ~ 00 I
z » (") » c» c» 0 0 (0
1.6 1.2 11" 1/ I 60°1- 1/ I .8 1/ I J I
rr I
I ( 1 C II V k;f 1 .032 .028 .036 lJ IR: def"lecfed I 1Y , I 1.0 .4 I _I I " ' "y V flop roughness roughness _I
\ J.yP""
Y/C' I , .002
, NT
-.025 ~ I ~ coeffiCient, split -l- 0 +0.063 .8 position
&x I
c.
i--- ~l I liff xlc Standard Standard a .258 .259 f', ~ . I I
I" I'" I
.
.6 I'c !'@ simulated -.4 1\ I I I I \ Section I I I I L ~/c ~ L
P
R 3.~Xf1-_~.255 6.0 9.0 6.0 D.20c 6.0 6 .\ .4 o <:> tI \l o ::L 1(71 In 11 ICl -.8 1- L- .2 - -1.2 I-
-
-- L.-- -l- 'J
o
.2 -1.6 -.I -.3 -.4 -.5 airfoil section, 24-inch chord.
-.2 -.2 .; .;
-
•
~o .024 .004
Q (J q, () " ~
~ ~ ..... _~ ~ ~ 1].020
{J 'Il 8 Q
.~ g' 4-..~ . ~.016 i·OJ2 .~ .;:.: q1.00B 66-009 ACA N the : I /6 ..--,
p- r
~ I~ ~
Iii
deq 'i ~ IJsI.
~\\. '", lXo, v A ~1vs:i)7 ~ ~ Aerodynamic characteristics of ~ Id 1,\ rf
a
f' attock, ~ }' 9"
"
II It" of" II ~ j
/
1// I I anr/e Id h{{ I '.
-8
rv.. ~
\ A II ex l Section -16 -24 .8 -32 1.0 20 -.8 3.6 3.2 28 2.4 -.4 -1.2 -1.6 -20
~ c: ill 8 t::: Q
.~
.~ :::: ~ :.:::::. as
-....-1.2 :g.4 .I • -./ -3 -.4 -.5 It.
/-.2 ~ ill 8 ~
~ ~~ :~
:f "i- ~
UJ c:j ~ ~ g; l-<j o "1 > >-< l:O "1 o >-< t"' t:::) ~
;,.
t-:l CD .t+:- I
z » (') » en en I\) 0 en
1.6 /,2 / 60· .8 )J-J.< c 1/ l/ .036 .032 .OE?8 Vr:~p f7 I.
deflected I /.0 I .4 ~ ~p 5 I 1 I I r roughness roughness , flap \ ylc t-c. '(ff \ r \ , -.048 -.017 !
coefficient,
o
.8 t024 sflif
+
position f5:.5"l:
""
\ c.
1\ il"t-[); lift oXic Sfandard a.
.255 .257 l:Jtandard 0.253 ~ .6 -.4 I I simulated Section 1< ,1 <~ oXic R 3.0x/O" 0.20c 6.0 6.0 8 .0 .4 V o 06.0 'V 09.0 6 -.8 .2 -/.2 0 ! , , , : !
, ,
°
.2 • -1.6 -: -.2
u •
airfoil section, 24·inch chord.
.024 .0/6 .012
~o (.f .... Q) <:; ~ \.J ~
i..
-1-" .;::: ...... ~
".020 u
... III S ()-. Cl. ~
~ .;::
-I-~ .\.) ::: il ~
6C,..206 N ACA the of --'--- lil:,,- ~ ~h.
deg '-~ .~
g
1£1..,.
cx., iJi r.t ~ ~ Aerodynamic characteristics I\{ Id IF P:!: '\ i'r"li: \\ . ~ 1,-4 ~ attocfr, I.f I~ IJ! It' 1.11 liS 1.-.. of I II
7 rv
I.Q anqle \I ~.
I~ -8 1< 1,# Iv\..-, I I ' 1 \ " ~ " Ii) r !'ii'1 I'{..W::l I" Ii I f'-JR.
t'- Section -/6 -24 I I , ,
" 1
4 Ii 4 0 2 -32 1.6 3.6 3.2 2.8 2.4 20 ·1.2 -/.
-/.
-2, \,)~ c: ~ v
.... .9; .... ~
...:::: :.:: Jl
./ • -.1 ~3 - 4 -.5 =t.
~
.l-.2 s § ~ EO
.... .Q :::: .... ~
I
~ trJ '"d ~ >-3 Z 00 t>:> "'" >-3 t::1 I-< ~ ~ ~
C)l t'>o? 0 ? z > I-< 0 Z > t" > ~ U2 0 0 0 ~ I-< >-3 >-3 trJ trJ "':I 0 ~ > trJ ~ 0 Z > q I-l H 0 U2 . t>:) I
Z » 0 » en en ... N
1.6 !
J1
.Ll 1.2
, ~
IQ. ld.V ~ -r--
I-': J.
60~ e. f';I.
1 11
.8 J 1 L
J. 1
,/ ~ lIT I'-" , 0, / r-::: .OJ6 .o3Z .DZB _I
~ r
V , deflecfed _I, /.0 .4
~
V , I I 1 I ~ 1 , roughness roughness flap , .J~ \&;l ylc -.004 " t--- r- , coefficient,
,I o
.8 split
+
position --., - , c.
lift !-- , c~ .rIc
" Standard
a .259 .258 0259-f008 15tandard
- .K
"
.6 -.4 ~ t-o I!;), l": simulated
1 1 1
Section ~ oXic 4 10, ~ I'" I '\ R 3.0x106 o.20c 6.0 60 Ll ~ I .4 o 06.0 09.0 'V V' 66.0"" ~ 8.. )l -.8 chord.
1\ ~ \.
~ 1\ -
...s ~
[ .2 -1.2 c- V- I"- 0 ! I . , ,
o
.2 -/.6 -.3 -.5 Rirfoil section, 24·ineh -.2 -.2 -.4 .; Ed ~o .024 .0/6 .012 .004 (,)
(lJ C) l! ~
:~ .... ~
..... .~ ::::
~",.020 ~ 8 g> t:: \J
iJ ....
..... :<:; ..::: 'I- .Q. Jl.008
6111-012 ACA N the of ...J I\> /6 characteristics ~ I~ IQ I ~r-...: ..."
H. .
\ deq l.t "I "" IV<; II lXo, (~ 8 11 il / Aerodynamic I /'.
Li "l (f attack,
.n
lil VI A VI ~ k of ~ ~ }5 '"""'"\ VI
u I
angle 1# fr l/!
d', -8 .1
I{ 1/1
i k "\ I Y
r
0.
10' Section b.
I -16 -24
-
! , ! r > I
. r ! I , o .4 .8 -32 /.6 -.
-.4 2.0 3.6 3.2 28 2.4 -/. -/.6 -2
~~ 8
S ~
~./.I: .~ .0 -.::: 't <t: :.::: :j:: .1 -.1 ~3 -.5 -.2 -.4 ~ ~ ,J t:: ~. I.J t:: ~
.... .~ .~, :::: ..... ~
t"
1fJ c:j ~ ~ > ~ o ":j > ..... ~ o ...... t:! > io'l > t-.:l ~
I
z » » en en - I\) .... I\)
/.6 I !Tl ~ J / 11 V< '/
Ir 1/
I 1.2 '1=1 /~ II J IM~ / 60"-t---f--+-+-+-+ V I/~ .8
./
~ JJ;!V
V V
I c, ~ [,.( tiTr.....- I 0321-+-l--+-+---Jf--+--+-+-
.O~ro-r'-"rT'-'- . .0281-+-l--+--t---f-+--+-+-
f2t:0" r deflected .
/.0 BY .4 IC/ I +--Ir-+-+--l--t-+--+--I-+---l-l--I +-If-+---+--+----I-t--+--+-t-t---t--t +-If-+--+--l--I-t--+--I-+---l-+--+ I , ~ fL roughness I f'lap rouqhness-+l-lt-tl-+-+-+-t-+-t y/c I " cOl6 eOl5 tJ I coefficient, split 0 .8
position, +
~ ) I c.
r--t:-:-- ~~ lift
~/c I _ 0.
.259
',standard 1.257 (?--, 15tandard I) I"l:~ I .6 -.4 f't
l\ \ I simulated 11
_ Section ~ ~~
I_I \ '.1
J +-H-+-+-+-H-+-+-+-H-+++-H-+-++-+
.rIc -L ~, I':::: 20e R 3.c;~I-O.255+-0037 9.0 6.0 6. 6:0 0.
~ .4 o 06.0 -.8
1"-
chord.
~~ I~.~ 1-L1-J
\
2Hnch .2 -1.2
r--
r--r--.-~--.---.-~--.--.--.--. I~ r--- H-+-+-+--t-+-+-+-+-t-+-+-+ I-+--+--+--+-+--+-+b. I-+--+--+--+-+--+-IV'
IJ H--+-+-+-t-+-l-....j H-+-+--H-+-+O 1-+--+--+--+-+--+-1'7 L..I...-L-L-..l-l-
.2 o ,;irfoil section, -.1 -1.6 -.2 cZ -.3 -.4 c5
ti •
016 004 ' ~o .024 . .012 .
1.1£ ~
8 ~ §
..... ~ ~
;..
<.>",>.020 8 ~ (J
;..~ .ID .\J .;:: .....
{j .~ ..;:: Jl.OOB 661-212 32 ACA the N of ~, f't) ~ 1...iR de9
I Iml~ I I)..q I I Id'
(f"11I'1'\ ~ 0(0' IT ~ I [7'/\""-,
r
'N
It' I I 1fT
Aerodynamic characteristics g
II
I ITI_~ 111
~ \a'
VII I r
attac/r,
All ~
~' LU~IJ~ 10 f'V\
&' of I~
WI I I I I I I
ret
Iff I I I I I
liLI IJ If' r-y..., anqle
Iff! I If
-8 '~
YI I IN
1-.f JJ.
VL
,~ ">: II
r~ li
Ui
q\;iY Section -q
rb&d
-16 EL3i~Y -24 .8 -32 1.6 -.4 -.8 3.6 3.2 28 2.4 2.0 -1.2 -1.6 -2.0 t.i \.) \.) QJ £: ~ ;..~1.2 ·21 :~ ~ ..... ~ It)
].4
.I 0 -.I -.2 -.3 -.4 -.5 :t .
rJ s::: Cl \.) ~ ~
........ .....
.\Q .~ Qj ..... ~
I I
>-3 ..... ~ ~ .... >-3 >-3 t>:I t>:I 0 ~ ;..- t>:I ~ 0 Z ;..- d >-3 ..... (") [fJ
'" t>:I "C 0 '" >-3 Z S' co ~ ... Z ;..- ..... 0 Z ;..- t"' ;..- t:l < [fJ 0 ~ ~ (") 0 I:rj t...:l C1 0") I 1.6 i I ~ II 1 I .1 ;> 1.2
I J
rr' II II ? ~ if I I I 60° j ~/ r1 /' I I .8 10 ~ I V '- If 1,1 I c, /' I .036 .O.l .028
"
deflected ~.
1.0 .4 I I I I I ~ .....", ~ L<v roughness· flap . I , ..
I , 0.106 -.088 -.073 , J:l v
r- <-
coeffiCient, J
sf!it o
position .8
i'- I'-, t-- I
c: ....., 6. lift
-- - r. k52
*~
,.!
2 ~t~n~0'J!l:~Ughness n.
\ standard 1. I
1"- r- I
-1.259
14 rr ~
.6 -.4 simulated ~ f' I I I
II
Section '\. ;l...,I'-. I I I
I
;rIc \ R
6.0 o.20c 6t
\ 3.0XIOj_,-0.258 9.0 6.01 6.
\ .4 o o " vI
<> 6
...
K -.8 t-\. 2"" .2 -1.2 f.--
V r---
'}
o
.2 -1.6 -./ -.3 -.5 -.2 -2 -.4 q ~ 10 .024 .012 .004
l III () u c: ~
.~ I\)
.... .~ ~ .... ~
~ c:: () U ~ ~.020 III () c:: u .il? .!::' {j ~.008 .... :;::.016 .0 1:: 66.-415 airfoil section. 24-inch chord.
NACA the
r-
S\ \ ~- f-' ~ ~ n~ ;'l"l K:J ~ characteristics of d ~ J:J.
/,;1"' ,j ~ I\, O('.J l< deg
1\ ~ .s:t"
...:"'t;> \ L.~ ~. ~ «0, A IF Aerodynamic 'V"V
IJ' I~ ~ ,-
~ I :y attack,
IY .I ~
of II II '--- j I ""': :s- If J '1 I angle 1/ Id' i'-;
-
1 k
I -8 !1 I J.
II ~I 'I Section q,..
'IN lIt
[ t \ 1\ -16 ~ -24 .8 .4 -32 1.6 3.6 2.0 -.4 -.8 3.2 28 24 -1.2 -1.6 -20
~ III 8 c: u
...:-1.2 .~ .!:! ~ ;:::; :.g c35 ~ ./ -./ -.3 -.4 -.5 "t ~
c:: ~ o \.) ~
l-.2
..... . i;:: ~ ..... ~
\
\
U2 q ~ ~ > ~ o I:rj > ...... ::tI I:rj o ...... t"' lj
~
~ Cll ""-l I
z » 0 en en CrJ 0 ... Q)
»
1.6 , ~ I 1.2 II
:>
~ it II I I I P II V.?
~ I / I!- / I >: .8 l¢>- Id 1 1 1 1 1 I /J II f- 't c,
/6
.036 .032 .028 61)- tJ.1 I «ei'/eftfd(6f 1.0 r:~ Ji. .4 tV 1 I 1 (:
I'-~ v ~
;-o,uq~~pss rouqhness flap , y/c. 112 K
.'
-.027
I-- Ie-- 1-0
-0.
coefficient,
.8 o
.-<... +
+.058 Sflit position ~ c.
r-- V tv liff
!f
.x/c Sra;o;ar,d Standard a .264 r.::: I
-
-
~ 1;;"'" H:266 1 1 liT .6 "';\ -.4 Simulated, 't- Section ""\ 1\ Ar\\' 1 1 1 \ \ _, ;rIc
"
1\ r\ h
f
\
it
K I\, 3.(j;,ior·~r-o.268 5.9~f1 6.0 D.20c 6.0 16.
.4 0. 'V (; o 09.0 C!. 'V 1\ -,8 t>., \ chord.
II 1\ ,~ r
"
( )..
24-,inch .2 1\ -1.2 < I--- 1---
l"-
,V > 0 z ) , !
/
o 5
.2 -1.6 airfoil section, -.4 ....
-.2 -.3 -,2 ti £0 ,024 ,0/6 .0/2 ,008 ~ ~o .004
" ~ \) ~
o
~ ...... .~ :g :::: 'i- ~
,,'0,020 ~ Qj \) \J (), \) \.)
~
...... ~ 'i-- ~ ~
:g
~18 AcA N the of C4 d<Jg «0' Aerodynamic characteristics attack, of angle -8 Section -/6 -24
o
.8 -.32 1.6 -.
3.;:: 3.6 28 2.4 2.0 -/. -I.
~1.2 -2.
~ \b 8 ~
.~ .~ '<..
... .;:: <::: :.::: :;::.4 ~
o 2 4
.I -./ .... .... -.
-.5
J
~ ti \j III U o !::: ~ ~ ..... .~
~ .... ~
I
>"l ....
t>:> 01> z ~ .... 0 Z :> t"' :> tj -< .... Ul 0 ~ >1 l.l 0 ~ ~ ~ >-3 l':I l':I '"l 0 ~ :> l':I ~ 0 Z :> q l.l Ul
~ '"d 0 ~ Z ~ (¥:J
~ 00 1'-:) I
z 0 » en en w N ... (X)
»
1.6 I I I : I I I I
J J J J J .1
1.2
1 L
"t
tJ
/ ,/ ''r'j f- I I~ 11 1/ II P ~
, I
60' IL
II i
A .8 1/ 'I c5 I I I f!,.
r!/~ it. ~]
j 'r c
.036 .032 .oze deflected /.0 .4
: I
\! roughness r-" (~u~hlns flop y/c " K --.085 -.084
~ -0.120
---I- 'P
splif coefficient,
o
poSition $
.8 c.
liff M::i r- f-- r--..
x/c Sfancbrd ~ta,n~alt a .260
~ ro 'Q
:v "1.261
-
\.
in 1-:0.264 .6
r u -.4
simulated ,\.. \->.,~ I I Section 1\ ~ r--..
I I 't i.,' z/c 1\ ~ R
3.ljfZr 6.0 6.0 020c' 6.0 6',0
\.
,,-\ .4 }- 17, \ o o 6. 'V 09.0 -.8
C\. c..cr--~
\ - r\ \ \ '\.
~ .2 -1.2 '- ~ --..
r '-
o
.2 -:1.6 airfoil section, 21-inch chord.
-.I -3 -.2 -.4 -:5 -.2 <i 016 £0 ~ ~o .004 .02.4 (,j Q)
~ e ~ ~ Cl
~ ~
.... .....
\J
~.020 (::' 8 g' c: Cl
Jli .\3 'i:: ...... {s.OI2 . ~ ~.008 ....:- 663-218 NACA the ~ L ~ en:> "'::i\
~f '\~
~ ~ ~ ::0 I\.
/6 \ ~ f7 I:.lI 'FP I\. ~ rY.: dec; '\
L&rr
~ 7- 1\ La h / ('iT' ''0' :fj; J.W
II rr rt
Aerodynamic characteristics of '// ~ A ~'1J II IJ.? M'P athck, ~ 111 ~ f"', III of ,~ /I 'I ~ !J
""
fj I~ Ii
1 angle
IlL , "-
~ -8 1.
III rl W
~ II V' i I I I I (f I~\a
'I! 10
Section ,.k 'h'ii -16 ( '-'tDr '.
-24 !
-32 .8 1.6 -.4 -.8 3.6 3.Z 2.8 2.0 2.4 ~1.2 -/.2. -1.6 -2.0
1,,- S ~ \J c: \J (IJ
..... ~ ~.4 V)
:12 .... "- '- -
.f -.1 -.3 -.4 -.5 It.
-
./-.2 S CJ \) ~ E
.....
:\3 .;:: Qj .... ~
[f1 c::: ~ ~ ::>- ~ o ":J ::>- .... l:d ":J o .... t" ~
t\j Cl \0 I
z »> 0 l=- 0) 0) ~ .... (X)
(..)
- 1.6 I I '- <"> 1.2 - J II II f- I J 'r' .
I
I 60' I ' I '~ :8 k{ '~~ I / 'nN ~ z J C .(),36 .032 ,028 deflected 1.0 .4 I I roughness
r- nap i~u~h'7efST
\~ lI/e ~ 0/34 -.096 -.090 , '\"
r-- ...-
split coeff'icient,
.8 o
position $
C c.
V'['(j lift
~ f'n RJt-t: :.ct
1]111111J1J~ a Standard .264 ~tor~a~d .26Z
-r--, ~ 0266
I\?- ~~ .6 \ -.4 simulo(e1 ['-,,1'-" f', I . I C 5ection 1\ Ib In \ \ J:lc 2~ 1\ 1\ )-
F
II _.
3.0(!-1- 6 6.0 9.0 6.0 0.
r>-..f\ Lr 1 Iq
.4 \ 11
o 'V6.(} 17 o <> t1 -.i!
1\ 1\ p .2 -1.2
-
~
f-..::: 1
V 0 t
, () .2 -1.6 airfoil section; 24-inch chord.
-./ -:3 -.2 -:2 -.4 -:5 .; ,; ' .024 .0/2 .008 .004
~o t!
~ ~ ~
U § ~
' ..... ~ ..... "'-
QJ B B' § u
<.),".020 .... .!,i :f.0/6 is
.~ :;:: Ji
663-418 A C N A the ;.> p ~~ ~ \ \ ~ [, ~Q ~
IA ~ ,.,.... rv
/6 Characteristics of &.
.'"
:;t $.fW" ~~ ;fV., ;yo deg ,
r'<;> i\ ~ r V
!(' ~ ~ ~ ff" «0; y ~ Aerodynamic
1/, f( rl :;e
~
'II I~ ;r"I .-
/, attoclr,
1.4 Wit n
H of, IlL (f j f1 .......
III 'fj J IJ '(j Ii A ongle 1/ II r-..
j ~8
g
;z If) '(f A fL
IJJ
~ h VI V Section f'Ji !/ i<L
t
'\ 'rV. "'"-P; -16 I'r 1\ -24 .8 '-32 1.6 -.4 -.8 3.6 .12 28 2.4 2.0 -1.2 -1.6 -20 IS ~ C:.
\J \J
-l--.,1.2 .~ ;l:: "-- ..... ~ ~.4 ~
./ -.4 -.5 -./ -.3 :t r.l-.2 c: o \J ~ ~ "-"'
.Il> .C .;::: 'Qj "i- ~
I
~ 0:. 0 ::0 t'l "d 0 ::0 ..., Z ? 00 I>:> "" z > ..., ..... 0 Z > t" > ~ ..... 1]2 0 ~ ~ ~ ..... t'l t'l
Ll 0 ..., ..., t'l ";I 0 ::0 > ::0 0 Z > d ..., ..... Ll 1]2
I
Z ""
» 0 » en en 0 N -0.
1.6 1.2 I I I I I I I .
.8 ~ I I I ~I I
11=±±J
z I I ~ C
.03~ .032 J
Ir deflected .
J, 1.0' I .4 l.J.
1:...Li
: I
"
"T , I 1/ I
roughness L} roughness
flap' , t/ .L -.235 J.
-.024 ,
r-- V ! I
coefficient, split .8 )- +360 I I c.t;ifion \
r--. I--- lift
.r/cy/c J.
Standard \ ~ Stondard a.
1.275 1.257 l
t.
-
--- hi\"
1\ I 'r-o.280 .6 ~ :J.,V It I simvlafed -.4
r 1
Section [;( 1 1 -' ~lll \ ;rjc IU
r\
1-' [ R I e.0 o.2Or:: 3.0xI06_ 9.0 6.0 6.0
K fC 1
.4
o o <> L'> "76.0
-.8 chord.
24~inch '1.2 .2 ~- I---
l"-
,V
) 0 I ) t 1 , ~
o
-1.6 .2
-. -..3
-,2 -.5 airfoil section, -.2 o .008 .004
~o .024 .016 .012 l (J
~ q, o S ~
:Q "'" ::::: .... ~-.
~
Q"d.020 t:: ~ 8 g- t:: (J
·91 .~ {,
.... ::::: :g ~
OO,-{)21 N ACA I the ~ E4 h \ \ \
¥~ ,., ~
') ~ ;c.
i,Q 1'« [r , (u Ig w deg \ A'A (), I'.
~ 1>1 I.n «., ~rd Ih~ f~ Aerodynamic characteristics of F Iff ~
III IW 197
ottocx, 11 ...tf I:§J r-...
~/ of II. IE!
Ii /,l VI fB ~ ill fl !Q /; '\ angle II }l i"" I/~ III -8 [1'\ '), 1" Section r -16 \ ; -.Ii,..( [ -24 t 1 I ?
9 4
" 0 0 8
.8 -32 1.6 3.6 3.2 2.4 2.0 -.4 2.8 -.8 ~1.2 -1.2 -:1.6
8 t:: -2-
IS a1 III (J ~
..... :u .;:: .... ;::: :.::: :2.4
o
.I
a-
~
J-:E q, o \) ~ ~
.~ .~ ~
..... ~ .....
(J2 c::1 a:: a:: >- ::ll >-< o "'1 >- I-t ::ll "'1 o I-t t"' t::) >- ~ >-
t~ 0;, I-' I z » en en .jIo ...
0 » N N
1.6 I.e i I J 60°
J
" .8
LQ 11 1 I ~ c,
If V
I I .032 .0315 .028
II 1 I
deflecfed I 1.0 .4 II If, I I I I I I' I \j I r-- I I flop roughness roughness "- ylc V " I .0/8 .088 ,
:----........ k----
I I split coeffiCient,
-+ to.OOO
.8 position ~ 1 I c.
r--- I---- ~ I lift
:rIc 2 Standard a Standard ~ ,~ ».1- 1.257 , 1.260 0.266
---- \
i"--
III
.0 , -.4 simulated \\ 1\\ "\ I I
I
, Secflon
\1\
I" 1'-1 1\ I
IV J I
~/c -" p: J 3.0)(/0' 6.0 9.0 6.0 Q20c 6.0 .4 o o <> t;, "
(,of -.8
.2 -1.2 ~ t---
l"-
,V 0 ,
o
-1.6 .2 airfoil section, 24-inch chord.
-.1 -.3 -.4 -,5 -2 -.2 ~ ~O .016 .Olc .001 .024 ~ti ~ 'IJ () (J
~ g
-;.." :Q ~ ~ ....
",..,.020 c: <!J 8 B' !;:: u
-a
"'!-.:' :~ :t :g ~.008
66.-221 NACA , , I o{the .>-,.~ ~H ..-'~ \ ~~ b- ~~ V' \ '"'"t:'t: ~ :\ ~ tjJ ~ ~ I I I
v
\
liJ :r'
deq '< I /,U , I\. -r:; Ii o l\ » ~
r IX
~\ ~ Aerodrnamic characteristics A IV' !i!H.
I!
~ W 11'V attach, ~< 1// 51 d of (7 III (f f) ~I I~ 5f r"--: II IJ angle A III !f '(f A
'"
-8
VI V, If
II , '(j ,e '\/P"' F\ II Section ~ ,'( I eif-' r't -16 ID 0... IDr i -24 -Jc .8 /.0 -.8 3.6 3.2 -.4 28 c.4
e.o
-I.e -/.o -20
I.I~ !;:: co 8 (J
§
"h.~/.2 .~ .... '" ~
.g ~ :t:;.4
./ • -.I -3 -.4 -.5 't 0 ~-.2 c: <!J () \) c: ~
~
-..- .'!! ." ~ ....
I
i:e I:'J '"C ,." ~ oc t-:> z >- ,." .... >- t" >- t:; ><: .... ,." ,." ,."
~ 0:> ~ 0 i:e Z "'" 0 Z -< .... w. 0 i:e 0 0 ~ ~ I:'J I:'J I-:<j 0 i:e >- I:'J ::0 C Z >- d ..... 0 rF- I
Z » 0 » 0') ...... .... N .... UI
1.6 i )
r
If 1.2 d I / II r.
1/ V ) 't P I I .8 j
II
z IV C 1/ tp' .032 .028 .036 I T 1/
I
1.0 / .I .4 ~ ---< t-- I--Il" roughness ~ y/c ['-l: -.0.65 t-- L..-- ....., coefficient,
.8 -t iD.o.46
u~ position to..
"'"
c.
--- _\ lift
~ :ric ( ...
a .248 .....
, i U?37 .
~a'7ci.ard .6 -.4 Section ;c/c ..
R
6.0 6.~11
2.9xIO' .
.4 o o 6 -.8 -, .2 -1.2
~ -
V
r,,"----- . 0 ! , , , I
o airfoil section, 24-inch chord
.21 -1.6 -.3 -:-.4 -;5 -.2 -.2 ti EO .004 ~o. .024 .0/2
CJ ~ ~ v ~ ~
:Q ::::: .... ~
- -.-=
".020
.. t: ~ 8 l5' ~
~.008
.... .~ .~ ::::.016 ~ ~
67,1-215 NACA I'
j
--l --l --l -t
-I -I
::l
r--1 H H H r--1 :::) --t the
I I H P I I
+-'-1 T-1
H
-+ ~.-,
+-I 1
.,...,., _ .......,.,
t---H r- c-t-i
I f ,..
1 1 .L
...J.++, j
--t-1 +-+-, -r-r --t-
r---r---;---1 It- t-t-,.....-r-t .L.L.L
~ rll "+1
r +--+-r-r .1 :..t J
'" I
r--T'""'-+ H .I--.l+-t-.---ru
-I
...l
~
I I ,-l---f--r-T 1.
f
I
rr-±-.l
...1J l deq
+ 1
-I--+-t-,.,- 112.
I
1 ...1 -+to j .1
.
aD, r I r I 0rr--t-1 -t +1 8 ;!
-
f1-
-l -l J-l I II: Aerodvnamic characteristics of
r-
j -I-t-r---'...1nFf +--+-+-, .L.L
~ -.J. -+
.1 .
I r"" I I
+--1 -.l attock, A r 1Il 4-1-1 ,-
" rr Tl 0
...1.1 -I J , -: of
t~~~tr1r=t±=tt~fi~t~
r-rl '~ t--t-r
.L
J J -+--r-l...1
-i ..1
IJ
J J +
-l.
r
rr rr+--+-r-T ;-..!
", rr
.\ + .L
angle
~ H-r
1---r-1 I tTl r
1.
I -8 -
,- r r l I I
r
-.l ~
J J:
~
I
r-
~
j-
~
I
~~~~~:ti=E±=t~~~§5;t±=
+-
J=t' Section
J i
I-I-i U-
llll
.LJ..J
ii
-+-+--1-, -16
rGI3Ef{=~f-~r-+±jeI~++ti±Ej
+-+-+-TI I D ~
L
~
r- r-
, rr rr rr ' I
-!.t ...I- J....j.. ...l.
, r I r I r
~ j ± -.l
J. ...1
t+
j .1 -24
t
...1 1J-I
I ~ r---t- II-- ,1
r-
n
J.l.l-l.-+.l ~
-+-t-t-r-,-
--I-+-iro
r ,r-
.. ;::r
-l-t-T,r --+-'-'11
- ~, r 2...~-,-:- r r;- r
R ~ I II r. I, r r r Ettitt-~Bl~"t·,t++~R=Er=DM '-+-+-lT r, rr rr r rr1
, '1':-,.
, 'rr Or 4 .~:::~ 5 ,
?r, 4 -32
-.
1.0 2.0 -I.
3.6 :12 2.8 2.4 -I.
-2
<I> <:) ~ §
~~
.~ ~ .... ....; ~ t ~
"f1.2
o
.
./ -.3 - -;5 -.2 .-./ ...
....
-
\J ,J 'Il Q t: q, E:
..... ~
.... .~ :::: >
U2 q ~ ~ o ":j >-< ll:i ~ >-< t' tl ~
t-:l ~ O!I
z > 0 > .... ~ .... > (,.) -0. UI
1.6 :
l'
V I
IL
1.2 II I I
'L Y
I 'r II 1/ V
I r
j; 60° I V .8 f
r.>
>1 J YO.
l
""' ~ ~ c
I I .036 .032 .028 ...n deflected 1.0 n..r .4 I I I '\I v- roughness roughness flap , :<:i.{"'::C vic .OJ4 .012 0'066 " ~c.
- '--T '.
coerticient, sflil o ~ positIOn .8
r-. tl. hl
c.
~ lift r--....
<r'C Standard Standard a .262 )..
I'-'
- 1\ t-, I
.6 --.4 simulated ,\ .....;:
-
SectIon ~ mi261
"
;rIc R .J.OxIO' 6.0 9.0 6.0 020c 6.0. 60 c "'" r-.- 'V [7 \ o O o L; (' [\. -.8 >, 1\ 1<;- ,
t:i r\
.2 -/.2 ~ .
t-
r--
0 ,t-
IV t I I 1 / ?-- " 5
o
-/.6 airfoil section, 24-inch chord.
.2 - _ -.
-.2 e
,"
.004 .0/2
~o .024 .0/6 " £: q, () \J ~ S
'd ~
-...:- ~ -...
A315 \J () \J <0'0.020 ~ u (iJ () ()-. § .;:: ~.008
-.-:- :::: iJ
NAOA I - -
l- I- ~ f-, I f- f-- the
l--
+-
f-t- 'J -l- L...J t- '--') K ' T.\- -<.; ~ "",po; r\ /6 -'H
J....11---t-~-Pcl~~ '
\ \ ~~~p:~h..
I~ ~ 1'7
deq
~9':l
tA f7 I/l ~ ~
~
I;:; r1 I VI
«0, },'1\ IH <";\!--,<
t+-H-:t+!1jji
AerodYnamic characteristics of
~
II 11 i i: I -.),¥, 5'
.• rv---;-v
attack, :;.
'17 - Itl
J'
I! ~~ of II j
r7 '8! rv
angle If
:~ -8
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III
~
M A
/1 I f I I
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Section
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...,-1.2 ~D <0::: "-- ;t:: " ~.4 U)
./ .
-./ -.2 -.4 -.5 -3 :t
,,"
IJ () ~ III a ~ 1:::
~ :Q ~
::::: ......
"
t;..J "d l:ll >-3 (Xl oj:>. > 8 1-1 0 Z > t"' > ~ 'fJl ~ ~ ~ ..... >-3 >-3 ~ t;..J "'.I l:ll Z > q >-3 1-1
l:ll 0 Z 9 to:) ~ ~. C 0 0 1; ~ C 'fJl
~ 0') Mlo-
Z » g .... .p.. .... ~ .... en
I I 1.8 I i/o 1_
T C
1/ II : -v ~ 1.-< I / I.C I ~ (:
D. I
I -
fri~ I l
" I
I 60~ ~ .....
( / r'ff .8 I I riUr' 'Ap;:r ,......
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I .036 .(J32 .028 rv deflected 1.0 .4 1 I I I
" ~
I roughness i...LL l- flap roughness
", H
yje I"-<
, J
, -.00/ -.001
t-- :"\ N
c(>efficient, sf'it 0
-+
:rO/8 .8 position
- l--.-L
c.
liff
t-- 1
xjc
D, '"'r--. V-.....j
Sfandard a .260 .260 "0 1\
1\ i'- ~
t--- 8."4 Li.staroard
.6 -.Lf simu(ated
I.r. ra ~ ,-
-
Section h: ,"-
EE
z/c .......
b..
~ I ) 6.0 9.0 G.20c 6.0 6:CjI 30,/0- 6:';
" ,,~ r
J .4- 17 o <> 'V
o L>
r\. K. b.. I"R -.8 chord.
,_ - jM-inch .2 -1.2 ~ I IL- i"'--
V 0 'f
r I I o airfoil section, -1.6 -J .2 -.2 -.3 -.4 -.5 -.2 <l
.(}I2 i
.016 . .OOB .004 .Ot?4 "->
io ~ III o o ~ ~
~ .... ~ "- ....
~
~ ~ 8 '" tl c:: tJ
'W..,.020
.... :(J :t: ~ :g Jl
747A415 , I I I NACA the t-' f:l. t> '~i.~;h~;.~,
-'
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fl- r-:~
~a I:c /6 ~ :t" P<-f"' ~ I ~ ...... k-I
1J. p-
deq )II' In ~f-I' o, t( r7 1\ oc If-' i"
~ ,v'
![ , Aerodynamic characteristics of lr4 II ~ it ~ '£ IQ !f attach, I~ ~ kr I'<J If ~ of tL '.L rv-~ If L ~ ,J J 13 kz.
~ I -8 I ~ A Jfz ~ '/ 'I f', I)" J-L Secfion angle I I l-q f\ -/6 f\ 1\ [ , [ -24 .8 -:32 -.8 1.6 3.6 3.2 2.8 2.4 2.0 -.4 -1.2 -1.6 -2.0
fS ~ ~ 3 ~ 0
-t-~I.t? :c:; .;::: c,..; ..... ~ .::::: ~.4 ~ .I -./ -.2 -.3 -;5 -.4 ~ ~
J 8 ~ ~
..... ~ ,~ ~
<;:: 't ......
" s ?' o ~ '" ~ .: ~ >-l "' iOl z ~ o o ", :'l () '" m m w m ~ w ~ .. o
,0
\
"-
"-
.....
......
......
z
Positive directions of axes and angles (forces and moments) are shown by arrows Axis Moment about axis Angle Velocities I Force (parallel Linear to axis) Sym- Sym- Positive Designa- Sym- (compo- Designation symbol Designation Angular direction nent along bol bol tion bol axis) Rolling _______ LongitudinaL ______ RoIL _______ X X L Y----+Z u p LateraL _____________ PitclL ___ . __ I) Pitching. __ . __ Y Y M Z----+X v q
'"
N ormaL _____________ Yaw ________ Yawing. __ . ___ Z Z N X---+Y w r
'"
Absolute coefficients of moment Angle of set of control surface (relative to neutral
L M N position), o. (Indicate surface by proper subscript.)
0,= qbS Om= qcS
O"=qbS
(rolling) (pitching) (yawing) 4. PROPELLER SYMBOLS D Diameter
p
Power, absolute coefficient Op= fD6
Geometric pitch pn P Pitch ratio
p/D 6/ V
0,
Speed-power coefficiellt= -V ~nj
Inflow velocity
V'
Slipstream velocity V.
Efficiency 1/ Revolutions per second, rps n
T Thrust, absolute coefficient OT= ;D4
pn V I Effective helix angle=tan- (2 ) 7rrn Q Torque, absolute coefficient Oa= ~nr.
pn LF 5. NUMERICAL RELATIONS 1 hp=76.04 kg-m/s=550 ft-lb/sec 1 Ib=0.4536 leg 1 metric horsepower=O.9863 hp 1 kg=2.2046 Ib 1 mph=0.4470 mps 1 mi= 1,609.35 m=5,280 ft 1 mps=2.2369 mph 1 m=3.2808 ft