The Development of Cambered Airfoil Sections Having Favorable Lift Characteristics at Supercritical Mach Numbers

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NATIONAL ADVISORY COMMITFEE

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FOR AERONAUTICS

L) z TECHNICAL NOTE No. 1771 THE DEVELOPMENT OF CAMBERED AIRFOIL SECTIONS HAVING FAVORABLE LIFT CHARACTERISTICS AT SUPERCRITICAL MACH NUMBERS ByDonaldJ. Graham Ames Aeronautical Laboratory Moffett Field, Calif.

Washington December 1948 CON1. irATE L1T BUSINESS, SCILNCE DEP'T.

TICHNOLOGY & DEC 13 1948

NATIONAL ADVISORY COl tLim FOR AERONAUTICS TECHNICAL NOTE NO.

177].

THE DEVKLOPMENT OF CAMBERED AIRYOTh SECTIONS KAVING FAVORABLE LIFT CHARACTERISTICS AT SUPERSONIC MACH NUMBERS By Donald J. Grnhmn STJMMARY Several groups of new airfoil sections, designated as the NACA 8-series, are derived analytically to have lift character- istics at supercritical Mach numbers which are favorable in the sense that the abrupt loss of lift, characteristic of the usual airfoil section at Mach numbers above the critical, is avoided.

Aerodynamic characteristics determined, from two-dimensional wind- tunnel tests at Mach numbers up to approximately 0.9 are presented for each of the derived airfoils. Comparisons are made between the characteristics of these airfoils and the corresponding character- istics of representative NPiCA 6-series airfoils.

The experimental results confirm the design expectations in demonstrating for the NACA S-series airfoils either no variation, or an Increase from the low-speed design value, In the lift coeffi- cient at a constant angle of attack with increasing Mach number above the critical. It was not found possible to improve the variation with Mach number of the slope of the lift curve for these airfoils above that for the NACA 6-series airfoils. The drag characteristics

of the new airfoils are sonlevhat inferior to those of the NACA 6—

series with respect to divergence with Mach number, but the pitching- moment characteristics are more favorable for the thinner new sections In demonstrating somewhat smaller variations of moment coefficient with both angle of attack and Mach number.

The effect on the aero&ynamic characteristics at high Mach numbers of removing the cusp from the trailing-edge regions of two 10-percent-chord-thick NACA 6-series airfoils is determined to be negligible.

NACA TN No.

The use of a negatively deflected plain flap at supercritical Mach numbers on an NACA 6—series airfoil is indicated to be a feasible and promising means for obtaining on demand the favorable variation with Mach number of the lift coefficient at a given angle of attack, characteristic of the NACA 8—series airfoils, while retaining at all other times the superior drag characteristics of the NACA 6—series type of airfoil.

INTRODUCTION The usual positively cambered airfoil sections exhibit two particularly undesirabl characteristics at supercritical Mach numbers. The angle of attack corresponding to the design lift coefficient increases rapidly with increasing Mach number above that for lift divergence, and the lift—curve slope decreases sharply at these Mach numbers. The effects of these character- istics on airplanes employing such wing sections is to alter, respectively, the longitudinal trim and the longitudinal stability and controllability in such a manner as to promote serious airplane.

diving attitudes, recovery from which may be extremely difficult with normal controls. (See reference 1.) On light highly maneu- verable aircraft, these characteristics can be avoided or satisfacto- rily coped with by the use of symmetrtcal airfoil sections and special controls. Neither of these means is advisable for large heavily loaded aircraft, however; the first, because in this case the air- foil must of necessity carry some design lift, and the second, because, as is stated in reference 1, the.trim changes occur so abruptly that the aircraft would be subjected to dangerously high accelerations before the controls could be reset. The logical means for avoiding the trim and stability changes on large airplanes is the employment of airfoil sections having no adverse changes with Mach number of the angle of attack for the design lift coefficient and of the slope Of the lift curve. The development of airfoils having such character- istics at supercritical Mach numbers has accordingly been made the subject of an intensive search.

Although it has not yet been found possible to control the variation with Mach number of the lift—curve slope, a means for achieving a favorable variation with Mach number of the lift of a positively cambered airfoil at the design attitude has been conceived by H. Julian Allen of the Ames Aeronautical Laboratory.

This principle has been employed to derive analytically a new group of airfoil sections, designated the NPLCA 8—series. The aerodynamic characteristics of these airfoils have been determined experimentally

NCA N No. 1771 in the Ames 1— by 3-1/2--root high—speed wind tunnel, and the results have, in most cases, confirmed the design expectations. An account of the airfoil development, analy-tical and experimental, is the subject of the present report,.

It was observed early in the course of investigations of compressibility effects on airfoil characteristics that the initial loss in lift (sometimes termed the "shock stall") experienced at supercrltical 'Mach numbers was associated with the formation of a compression shock wave on the upper surface of an airfoil before the critical Mach number of the lower surface had been exceeded. It has been concluded that the loss in lift results from an effective change In the airfoil camber occasioned by a suddenly thickened boundary layer behind the shock wave on the upper surface while the boundary layer on the lower surface remains sensibly unchanged.

Previous research has been aimed at continuously increasing the Mach number of occurrence of the compression shock so as to delay the shock stall. In the present development the upper—surface shock wave is accepted, but the associated loss in airfoil lift is forestalled by inducing a corresponding shock, with accompanying boundary—layer growth, to occur on the lower surface.

It was reasoned that If the flow over both surfaces could be- kept similar at supercritical Mach numbers the net lift of an air - foil could. be maintained at an approximately constant design value.

To effect this result the respective minimum pressures on the upper and lower surfaces would have to be equal. Because the drag charac- teristics at supercritical Mach numbers would be adversely affected by simultaneous occurrence of compression shocks on the respective surfaces, It would be desirable to obtain the highest possible air- foil critical Mach number. It was further realized that, to produce a positive lift force on the airfoil at supercrltical Mach numbers under this condition, the position of minimum pressure would have to be located further aft on the lower surface than on the upper surface.

The respective upper— and lover—surface minimum pressures being equal, a more severe adverse pressure gradient would thus be imposed aft of the minimum pressure position on the lower surface, forcing a greater thickening of the boundary layer on this surface at Mach numbers above the critical. The effect of the thickened lower—surface boundary layer should compensate, to a degree depending upon the respective upper— and. lower—surface velocity distributions, for the upper—surface boundary—layer growth and result in either no change or an effectively positive change in the airfoil camber at Mach

14.

NACA iN No. 1771 therefore concluded that, by numbers above the critical. It was suitably choosing the velocity distributions over the upper and have, at a lower surfaces, airfoil sections could be designed to given angle of attack, an approximately constant or an increasing lift at supercritical Mach numbers.

Following this line of reasoning, an initial group of three NACA 8—series airfoil sections, 16—percent--chord-thick, having different respective positions of minimum pressure (or maximum local velocity) on the upper and lover surfaces was designed and tested. The sections were derived in essentially the same n1mer as were the later families of NACA airfoils by combining mean camber lines with basic thickness forms to produce a desired veloc- ity distribution. Velocity distributions were selected to provide the desired aerodynamic characteristics at supercritical Mach numbers, and the airfoil shapes corresponding to these distributions were determined by the method of reference 2.

The first airfoil was proportioned to have equal upper— and lower—surface minimum pressures occurring at 30 and 50 percent of the chord, respectively. The base profile was obtained by combining proper fractions of the thickness forms of the NACA 63— and. 65—series airfoils and the "double—roof" profiles of reference 2. A mean camber line satisfying the condition of equal minimum pressures for the upper and lower surfaces was determined by combining suitable proportions of the NACA a=0.3, 0.5, and 1.0 mean lines. (See reference The ordinates of the mean camber line were adjusted to produce 3,) the desired design lift coefficient. The actual airfoil shape was then obtained by combining the base profile with the mean camber line, using the methods of references 2 and. 3.

In the described manner three airfoil sections were derived with different respective upper— and lower—surface minimum—pressure positions so located as to permit the effects of a variation of the severity of the lower—surface pressure recovery to be observed. The airfoils were designated as follows: NACA 835A2l6 NACA 836A216 NACA 811.7A216 The shapes and velocity distributiori.s for these airfoils are illus- - trated in figure 1.

The numbering system for these airfoila is identical with that for the NACA 7—series airfoils and is suimnarized given in reference

NACA TN No. 1771

as follows: 1st digit - Airfoil series number 2nd digit - Position of minimum pressure on upper surface in tenths of chord from leading edge 3rd digit - Position of minimum pressure on lower surface in tenths of chord from leading edge Letter - Serial letter distinguishing airfoils having the same thicknes design lift coefficient and minimum pressure positions but different camber or thickness distributions Itth digit - Design lift coefficient in tenths 5th and 6th digits - Thickness—chord ratio in hundredths Tests of the Initial three airfoils revealed variations in lift coefficient with Mach number in the vicinity of the design lift coefficients which at supercritical Mach numbers differed in impor- tant aspects from the type of variation normally observed for airfoils.

The lift coefficient at a constant angle of attack increased markedly with Increasing Mach number above that for normal lift divergence as contrasted with the usually noted opposite variation. Instead of decreasing with increasing Mach number above that for normal lift divergence, the lift coefficient at a constant angle of attack increased, markedly with Mach number. This result confirmed the design expectations to a greater degree than anticipated, and indicated that great difficulty would be experienced in trimming an airplane using such wing sections at any but positive lift coefficients at.

supercrItical Mach numbers, an important safety feature for large heavily loaded aircraft. This characteristic was unfortunately accompanied by erratic variations with Mach number of the slopes of the lift curves, which are very undesirable fron the standpoint of airplane controllability.

The desired type of supercritIcal speed lift characteristic having been realized, efforts were directed toward the derivation of thinner sections with modified camber so as to produce less powerful lift changes at supercritical Mach numbers. A group of 10—percent--chord—thick profiles was accordingly derived from the NACA 836A2l6 airfoil, this section among those tested having the most favorable characteristics at low and moderate lift coefficients.

NkCA Til No. 1771

This second group of airfoil sections was composed of the following: NACA 836Allo NACA 836Bllo NACA 836db NACA 836Dllo The NACA 836A110 airfoil was scaled down in thickness from the NACA 836A216 airfoil and the camber line ordinates were adjusted to give a design lift coefficient of 0.1. Tests of this airfoil disclosed the need for modification of both the thickness and the camber distri- bution, the gain in lift at supercrltical Mach numbers still being greater than desirable.

It wasreasoned that, by decreasing the negative lift carried over the rear portion of the airfoil at subcritical Mach numbers, the change in the total lift of the airfoil at supercritical Mach numbers would be reduced. The NACA 836B1l0 airfoil was designed to effect this result by modifying both the mean camber line and the thickness distribution of the NACA 836All0 airfoil. The mean camber line for the former was obtained as the sum of equal propor- tions of the ordinates and slopes of the mean line for the latter aIrfoil, and of a uniform load (a=1.0) mean line, The upper— and lower—surface minimum pressures were maintained approximately equal by adding to one—half of the base profile ordinates of the NACA 836A1l0 airfoil, one—half of those for the NACA 66-oio airfoil.

The resulting changes in profile and velocity distribution may be noted from an examination of parts (d) and (e) of figure 1.

The NACA 836db airfoil was designed to investigate the effect on the supercritical speed aerodynamic characteristics of the NACA 836B110 airfoil of removing the cusp from the rear portion of the profile. The former differs from the NACA 836B1lo airfoil only in that the profile is linear over approximately the last two—tenths of the chord.

Tests of the NACA 836Bll0 airfoil indicated that the profile modification from the NACA 836All0 airfoil was effective in reducing the magnitude of the lift—coefficient increase at supercritical Mach numbers in the viôinity of the design lift coefficient.

Further improvement was still felt to be desirable, however, particu- larly in the slope of the lift curve at lift coefficients greater than the design value. A decrease in the severity of the pressure recovery over the lower surface (by decreasing the negative pressure peak) was indicated as a possible corrective measure. To test this

NACA TN No.. 1771 . 7 hypothesis, the NkCA 836D1lo airfoil was derived by combining the thickness form obtained as the sum of equal proportions of the NACA 836AOlo and. 63-olo profiles with the mean camber line of the NACA836BuO airfoil. The difference between the NACA 836BllO and 836Dno airfoils may be seen from figure 1 to be principally in the magnitude of the lower-surface minimum pressure.

To investigate the possibility of realizing Improved character- istics from more rearward minimum-pressure positions on both surfaces, three additional 10-percent-chord-thick sections were derived from.

the NACA 814.7A216 airfoil and were designated as follows: NACA 81i.7Allo NACA 814.7Bfl0 NACA 81i.7cllo The NACA 8'VTAllO airfoil was derived from the NACA 811-7A2l6 airfoil by reducing the base-profile ordinates o? the latter in the ratio of flfteen-slxteenths times the quotient resulting from the division of the ordinates of the NACA 66-oio airfoil by the ordinates for the NACA 662-015 airfoil, and. by reducing the camber line ordinates and slopes in the ratio 10:16. The NACA 81i7B1l0 airfoil was obtained by combining the sum of one-half of the base- profile ordinates of the NACA 8#7A110 airfoil and one-half of those for the NACA 6 14oio airfoil with the mean camber line consist- - ing of equal proportions of the slopes and ordinates of the mean line for the NACA 8 1 7Allo airfoil and of the uniform load mean line.

The NACA 8i.7C11O airfoil consists of the NACA 8Wuo airfoil with the cusp removed from the trailing-edge region of the latter by substituting straight lines for the portion of the profile from approximately the 80-percent-chord position to the trailing edge.

The ordinates of all of the airfoils investigated are given in tables I to X. The shapes and theoretical velocity d.istribu+,ions for all but the NACA 836db and. 8 1 7Cll0 airfoils (which differ but slightly from the NACA 836B110 and. 8147B1l0 airfoils, respectively) are Illustrated In figure 1.

It Is to be noted that negative d.eflections of a plain trailing- edge flap on an ordinary airfoil section would produce lower-surface velocity distributions approaching in character the distribution previously described for the new type of airfoil section with the reflexed mean camber line. The results of an investigation (also conducted in the Ames 1- by 3-1/2-foot high-speed wind tunnel) of an NACA 65-210 airfoil with a 20-percent--chord negatively deflected flap accordingly are presented and compared in the present report with those for the NACA B-series profiles.

NACA TN No. 1771 -- SYMBOLS a mean—line designation, fraction of chord from leading edge over which design load is uniform a0 airfoil section lift—curve slope, per degree c chord, feet section drag coefficient Cd c1 section lift coefficient c 21 design section lift coefficient section moment coefficient about quarter—chord point

I

N Mach number V free—stream velocity, feet per second v local velocity, feet per second x distance along chord, feet y distance perpendicular to chord, feet a0 section angle of attack, degrees aj section angle of attack corresponding to design lift coefficient, degrees flap deflection, degrees § APPARATUS AD TESTS The tests were made in the Ames 1— by 3-1/2—foot high—speed wind tunnel, a low turbulence, two- .ditnensional—flow wind, tunnel.

The airfoil models were accurately constructed of duralumin and were of 6—inch chord and 12—inch span. The models completely spanned the narrow dimension of the tunnel test section. Two— dimensional flow was assured through the use of sponge—rubber /

NACA T? No. 1771 gaskets (to prevent end. leakage) compressed between the model ends and. the tunnel walls.

Measurements of lift, drag, and. quarter-chord pitching moment were made as nearly simultaneous as possible at Mach numbers ranging from 0.3 to as high as 0.9 for each of the 0airfoils at angles of attack increasing by 20 increments from —6 to a maxm'mi of 120.

The Reynolds number variation with Mach number for the tests is expressed graphically in fIgure 2.

Lift and. pitching moments- were evaluated by a method similar to that described in reference 3 from integrations of the pressure reactions on the floor and. ceiling of the tunnel of_the forces on the airfoils. Drag values were determined from. wake-survey measure- - ments made with a rake of total-head tubes.

RESULTS AND DISCUSS ION Section aerodynRm-ic characteristics in coefficient form are presented as functions of Mach number in figures 3 to 47 for the NACA B-series airfoils, two representative NACA 6-series airfoils, and. the NACA 65-210 airfoil with a 20-percent-chord plain trailing- edge flap neutral and negatively deflected through 6°. All of the characteristics are shown corrected for tunnel-wall Interference by the methods of reference 4. The-dashed portions of the airfoil characteristics curves serve to indicate the extent of possibly unreliable data obtained in the close vicinity of Mach numbers for which the flow In the tunnel test section was choked, that is, for which the Mach number of unity was attained locally across the test section.

Characteristics of Initial. Three Airfoils It is seen from fIgures 3, 4, and.

5 that the respective variations

with Mach number of the lift coefficient at constant angles of attack for the NACA 835A216, 836A2l6, and.

81 4 . 7A216 airfoils differ markedly from the vacriatións generally observed for ordinary airfoil sections.

An abrupt increase of large magnitude occurs in the lift coefficient at angles of attack within the normally useful range at Mach numbers above those for lift divergence in place of the customary decrease in lift coefficient. The difference in characteristics is emphasized in figure 6 which illustrates the variation with Mach number of the angle of attack required to maintain the design lift coefficient of 0.2 for each of these NACA 8-series airfoils, and for the NACA 652-217,

10 NAA TN No. 177].

a=0.5, airfoil (reference 5), a representative NACA 6—series airfoil.

The explanation for the radical lift characteristics of the new airfoils is to be found in an examination of the theoretical low—speed velocity distributions of figure 1. At Mach numbers above the critical the strong adverse pressure gradient aft of the minimum—pressure position on the lower surface promotes a rapid thickening and separation of the boundary layer from this surface, resulting in the loss of an extensive portion of the negative lift carried over that'part of the airfoil.inimed.iately aft of the lower— surface minimum—pressure position. The Increasing extent of the separation on the- lower surface with increasing Mach number produces the increasingly positive variation of lift coefficient at constant angles of attack observed in figures 3, ii. , and 5.

The variation with Mach number of the lift' coefficient at low positive and. negative angles of attack, although favorable in the sense that the lift coefficient increases with Mach number rather than decreases, is. so violent for these three airfoils as to cause very erratic and. undesirable variations in the slope of the lift curve at the higher Mach numbers. (See fig. The variation of

7.)

the angle of attack necessary to maintain the desii lift coefficient of 0.2, although in the direction to promote safety at high Mach numbers for an airplane employing such airfoils as wing sections, has already been observed in figure 6 to be undesirably large. For these reasons it was concluded that the first airfoils were cambered too severely and that a modified amount of camber as well as a change In the distribution would produce less drastic changes in the lift coefficient with increasing Mach number.

The drag characteristics of the three airfoils (figs. 8, 9, and 10), as was expected, are much inferior to those of the NACA 6—series airfoils, as represented in reference 5 by the NACA 652-215, a=0.5 and NACA 66,2-215, a=0.6 airfoils, with respect to divergence with Increasing Mach number at low and moderate angles of attack despite allowance for the small difference in thicimess of the airfoils.

The variation of pitching-noment coefficient with Mach number for the NACA 835A216, 836A2l6, and 8 Ii.7A216 airfoils, shown in figures 11, 12, and 13, respectively, is consistent with the varia- tion of lift coefficient. The moment coefficients vary from positive values at low Mach numbers where negative lift is carried over the rearward portion of the airfoil to negative values at high Mach numbers where this negative lift is lost.

NACA Til Eo.

Characteristics of the NACA 836-110 Airfoils NACA 836A110.- Because current design trends indicate thinner wing sections for high Mach number applications, it was considered desirable to further the investigation on airfoil sections of 10- percent-chord maximum thickness. The effect of halving the amount of camber and decreasing the profile thickness of the NACA 836A216 airfoil may be seen from an examination of the characteristics of the NACA 836A110 airfoil.

The variation of lift coefficient with Mach number for this

airfoil (fig. 114 . )

is much less drastic at supercritical Mach numbers than that noted in figure 3 for the NACA 836A216 airfoil. The lift- curve_.slope variation with Mach number Is considerably improved for

the NACA 836A110 airfoil (cf. figs. 7

and 19), and the angle of attack required to maintain the lift coefficient at the design value (cf. figs. 6 and 20) is correspondingly reduced for the thin- ner lower-cambered profile. The latter variation is still uxidesir- ably large, however.

The differences in the lift characteristics of the NACA 836A110 airfoil and the NACA 614ll0 airfoil, as representative of the best NACA 6-series sections for high Mach number applications, may be seen from a comparison of figures 114. and. 18 to lie in the variations of lift coefficient with Mach number at small positive and negative angles of attack. The departure of the characteristics of the former airfoil from those usually observed for airfoil sections at supercritical Mach numbers is more strikingly illustrated in figure 20, depicting the variation with Mach number of the angle of attack required to maintain the design lift coefficient of 0.1 for the NACA 836-110 and 611110 airfoil sections.

The drag and pitching-moment charncteristics of the NACA 6i-ll0 airfoil section at high Mach numbers being unavailable at the present writing, these characteristics for the NACA 836A110 airfoil must be compared with thcse for the NACA 65-210 airfoil as the next most representative profile of the NACA 6-series airfoils available. The drag characteristics of the NACk 836A110 airfoil (fig. 21) compare unfavorably with those of the NACA 65-210 airfoil (fig.

25), parti- cularly at the angles of attack corresponding to the lover lift coefficients. Divergence not only occurs earlier for the former, but the drag coefficients at a given lift coefficient are higher.

From figures 26 and. 30, respectively, the pitching-moment coefficients for the NACA 836A1l0 airfoil, in additIon to being

NACA TN Ro. 1771 more positive, exhibit a generally smaller variation with Mach number than do those fOr the NACA 65-210 aIrfoil.

NACA 836B110.- The NACA 836Bll0 airfoil was derived from the a mrer as to reduce the negative lift NACA 836AllO airfoil in. such on the afterportion of the airfoil at subcritical Mach numbers at the design lift attitude and to retain the approximately equal critical Mach numbers of the upper and. lower surfaces. The effect of these profile modifications on the lift-coefficient variation with Mach ]A arid. The variation number is shown by a comparison of figures 15.

in the vicinity of the desigLi lift coefficient is seen to be small for the NA.CA 836BU0 airfoil as compared with that for the NACA 836A110 airfoil. At angles of attack appreciably above and. below the ideal angle, the lift-coefficient variation resembles that observed for the NACA 6-series type of section. (See fIgs. 18 and 14.2 for the NACA 6l_li0 and. 65-210 aIrfoils.)

the slope of the lift curve at the A considerable increase in design lift coefficient Is observed in figure 19 for the NACA 836B110 airfoil over that of the NACA 836A110 airfoil for Mach numbers between 0.75 and. 0.85. The variation with Mach number of this parameter for the former airfoil is closely comparable to that for the NACA 614-110 airfoil. From figure 20, it can be seen that the variation with Mach number of the angle. of attack necessary to maintain the design lift coefficient for the NACA 836B110 airfoil is greatly reduced from that observed for the NACA 836All0 airfoil.

The drag characteristics of the NACA 836B110 airfoil (fig. 22), although considerably Improved, over those of the NACA 836All0 section, are still Inferior with respect to divergence with Mach the vicinity of the design lift coefficient to those of number in the NACA 65-210 airfoil when compared on the basis of equal lift coefficients for the two airfoils.

The variation In pitching-moment coefficient with Mach number (fig. 27) for the NACA 836Bll0 airfoil closely approaches that for result of the camber modification the NACA 65-210 airfoil,, as a frdm that of the NACA 836All0 airfoil.

NACA 836db.- This airfoil section was tested to determine the effect on the aerodynamic characteristics of the NACA 836Bll0 airfoil of removing the cusp from the trailing-edge region of the airfoil. Comparison of the respective variations with Mach number of lift, drag, and. pitching-woment coefficients (figs. 16, 23, arid.

28, respectively) for the NACA 836db airfoil with the corresporid.- lug variations for the NACA 836Bll0 airfoil reveals no significant

NACA TN No. 1771

differences in the characteristics of the two sections.

NACA 836D110.- The NACA 836Dll0 airfoil was designed to investi- gate the effect of both decreasing the amount of negative design lift from that of the NACA 836A110 airfoil and raising the lower- surface critical Mach number above that of the upper surface (note theoretical velocity distribution, fig. 1) in an attempt to obtain a more favorable variation with Mach number of the slope of the lift curve. Figure 17 indicates the NACA 836D1lO airfoil to be the most promising of those yet discussed, virtually no variation with Mach number being manifest in the lift coefficient near the design value. This characteristic is reflected in the small variation with Mach number in the angle of attack required to maintain the design lift coefficient of 0.1. (See fig. 20.) With reference again to figure 17, the variation with Mach number of lift coefficient at constant angles of attack other than that. corresponding to the design lift coefficient indicated lift-curve slopes closely resembling those of the NACA 6-series airfoils as represented in fIgure 18.

Except at the higher lift coefficients, no inprovement in drag characteristics from those of the NACA 836B110 airfoil resulted from this profile modification.

The variation with Mach number in the pitchIng-nioment coeff I- óients of the NACA 836Dll0 airfoil (fig. 29) does not differ note- worthily from those for the other airfoils of the series.

811.7-110 Airfoils Characteristics of the NACA NACA 811.7A110.- The NACA 814.7AllO airfoil was derived from the NACA 8 l 4.7'A2l6 section by decreasing the thicimess and the camber-line ordinates in the same rner as was done in the case of the NACA 836Al10 airfoil. The lift-coefficient variation with Mach number (fig. 31) closely resembles that for the latter airfoil. The varia- tion with Mach number of the angle of attack required to maintain the design lift coefficient (fig.

is similar to that observed 35) for the NACA 836A1l0 airfoil (fig. 20).

The variation in drag coefficient. with Mach number (fig. 36) is more favorable for the NACA 811.7Afl0 airfoil than for the NACA 836A110 airfoil from the standpoint of divergence with Mach number at angles of attack in the vicinity of the ideal angle.

C) The pitching-1oment-coefficient variation with Mach number for the NACA 811.7AllO airfoil (fig. 39) is similar to that for the NACA 836A110 airfoil, but the moment coefficients are of smaller magnitude.

lii.

NACA TN No. ma NACA 814.7B110.- The lift characteristics of this airfoil, developed from the NACA 8 14.7A110 airfoil by decreasing the negative contribution to the design lift distribution and by decreasing the lower—surface pressure peak below that of the upper surface, are seen from figure 32 to be considerably improved over those of the latter airfoil. As in the case of the NACA 836D110 section, the variation with Mach number of the lift coefficient in the vicinity of the design value is indicated to be very small, and yet a reasonably satisfactory lift—curve slope is retained. (See fig. 314..) The variation with Mach number of the angle of attack for maintenance of the design lift coefficient (fig,, 35) is as favorable as that observed for the NACA 836D110 airfoil. A marked improvement In the variation of drag coefficient with Mach number at zero lift is noted from a comparison of figures 36 and 37 for the NACA 811-7A110 and 81VBli0 airfoils, respectively. For this condition the drag—coefficient variation of the latter airfoil is superior to that of the NACA 836D110 airfoil.

The superiority Is considerably reduced at the design lift.coefficient and disappears at the higher lift coefficients.

The variation in pitching—noment coefficient with Mach number for the NACA 814-7Blio airfoil (fig. 14.0) is observed to be very small in the vicinity of the design lift coefficient and parallels the characteristics of the NACA 836Dll0 airfoil in this respect.

NACA 814.7Cllo.— This airfoil was designed to investigate further the effects on the characteristics at high subsonic Mach numbers of removing the cusp from the trailing—edge.region of an airfoiL From figures 33, 314.

35, 37, 38, 14.0, and 14.1, the characteristics of the resulting airfoil are seen to be essentially the same as those of the cusped NACA 814.7BllO profile. From this and the similar result observed in the case of the NACA 836-110 airfoils it is concluded that for 10—percent—chord—thick airfoils of this type of section the aero- dynamic characteristics are not materially affected by removal of the cusp from the afterportion of the profile.

It is to be noted that in the case of both the NACA 836-110 and the NACA 814.7-110. airfoil developments the sections having the most favorable lift characteristics are those for which the negative portion of the design lift is small and for which the minimum pres- sure is somewhat lower on the upper surface than on the lower surface of the airfoil. The latter result is in contradiction to the design assumption that the upper— and lower—surface pressure peaks should be equal. Although it was not found possible to improve the lift- curve—slope rariation with Mach number for these airfoils over that

N&CA TN No. . 15

characteristic of the NACA 6—series airfoils, the . NA2A 836D110 and 8 11.7BiiO profiles are indicated to be the. equal of the NACA 6—series type in this respect. The drag characteristics of the best NACA 8—series airfoils thus far derived are not as favorable as those of NACA 6—series airfoils in that the drag—divergence Mach numbers are lover for comparable lift coefficients in the vicinity of the design lift coefficients. The pitching-noment characteristics of the more promising airfoils of the NACA 8—series are, if anything, superior to those of the NACA 6—series in that the variations of moment coefficient with Mach number are generally smR.11er for the former.

Characteristics of an Airfoil With a Negatively Deflected Flap From an examination of the lift characteristics of an NACA 65-210 airfoil section with a 20—percent--chord plain trailing—edge flap, a marked similarity was noted between the variation with Mach number of the lift coefficient at various angles of attack for a small negative flap deflection and the characteristics previously observed for the NACA 8—eeries airfoils. It would be very desirable to be able, by negatively deflecting a plain flap, to effectively reflex the camber of a wing section on an airplane in flight from the uniform load type at subcritical Mach numbers to something approaching that of an NACA 8—series profile at supercritical Mach numbers. To permit an appraisal of the characteristics of an airfoil with a negative1- deflected flap at high Mach numbers, the aerodynamic characteristics of the NACA 65-210 airfoil section with a 20—percent—.

chord plain flap undeflected, and. negatively deflected 6°, are 14.3, 14.6, and. 14.7, presented in figures 2, 25, and 30 and in figures respectively, for comparison with those of. the NkCA 8—series airfoils investigated.

The similarity between the respective variations with Mach number in the lift coefficient at a constant angle of attack for the NACA 65-210 airfoil with the flap deflected -6° and the NACA 836D110 and 814.7BllO airfoils is readily apparent from a comparison of figures 14.3, 17, and 32. The lift characteristics of the three .1 1. and li.5 depicting the airfoils are further compared in figures 14 respective variations with Mach number in the lift—curve slope and the angle of attack required to maintain the lift coefficient of 0.1.

The similarity between the latter characteristics for the airfoil with the negatively deflected flap and the two NACA 8—series airfoils is unmistakable.

NAA TN No. 177].

The drag characteristics of the flapped airfoil (fig. 1 .

4 6) are similar to those of the NACA 836D110 and 814 . 7Biio airfoils. The pitching-moment characteristics of the three airfoils (figs. 14.7, 29, and. 40) also bear a close resemblance to one another.

The principle of reflexing the camber line by negatively deflect- ing plain trailing-edge flaps on NACA 6-series airfoils at super- critical Mach numbers to produce favorable variations in lift coeff 1- dent v1thincreasing Mach number on the strength of the results contained herein has already found important app1ication (in an expedient sense) on several high-speed airplanes and. merits further investigation.

C0NCL]DING REMARKS A new group of airfoil sections, de1gaated the NACA 8-series, has been developed having favorable'lift characteristics at super- critical Mach numbers. Through the use of negative camber over a portion of the airfoil chord it has proved possible to hold the lift coefficient of the new type of airfoil approximately constant at some design value with increasing Mach number to at least 0.9 Mach number, the limit of the present investigation. By suitably choosing the camber and thickness distributions for the airfoils, a particular variation with Mach number of the angle of attack required to maintain a given design lift coefficient can be obtained.

No means has been found for improving the lift-urve-slope character- - istics of .the NACA 8-series airfoils beyond. those of' the NACA 6-series sections. Although some control can be exercised over the drag and pitching-.moment characteristics of the former airfoil sections without adversely affecting the lift characteristics, it is generally necessary to accept drag characteristics somewhat poorer with respect to divergence with Mach number than those of the NACA 6-series air- foils presently used for high Mach number app1ication. The pitching- moment characteristics of the NACA 8-series airfoils are generally more favorable than those of the NACA 6-series airfoils in that the variations of pitching-moment coefficient with Mach number and angle of attack at supercritical Mach numbers are somewhat smaller for the former. - Flat-sided profiles may be used. in place of the cusped trailing- edge profiles on 10-percent-chord-thick NACA 8-series airfoils with- out significantly altering the aerodynamic characteristics of the airfoils at superOritical Mach numbers.

The lift characteristics of the NACA 8-series airfoils at supercritical Mach numbers can be approximated with NACA 6-series

NACA TN No. 177].

airfoils through the use of negatively deflected plain trai1ig- edge flaps. This application appears to be a promising means for obtaining on demand the favorable variation with Mach number of the lift coefficient at a given angle of attack, characteristic of the NACA 8-series airfoils, and yet retaining at all other times the superior drag characteristics of the NACA 6-aeries airfoils.

Ames Aeronautical Laboratory, National Advisory Committee for Aeronautics, Moffett Field, Calif.

1. Hood, Manley J., and. Allen, H. Julian: The Problem of Lon.gi- tudinal Stability and Control at High Speeds. NACA Rep. No.

767, 1914.3.

2. Allen, H. Julian.: General Theory of Airfoil Sections Having Arbitrary Shape of Pressure Distribution. NACA Rep. No. 833, 1914.5.

3. Abbott, Ira H., von Doenhoff, Albert E., and Stivers, Louis S., Jr.: Summary. of Airfoil Data. NACA ACR No. L5CO5, 1914.5.

14. • Allen, H. Julian, and Vincenti, Walter G.: Wall Interference in a Two-Dimensional-Flow Wind Tunnel With Consideration of the Effect of Compressibility. NACA Rep. No.

782, 191414..

5. Graham, Donald J., Nitzberg, Gerald E., and. Olson, Robert N.: A Systematic Investigation of Pressure Distributions at High Speeds Over Five Representative NACA Low-Th?ag and Conventional Airfoil Sections. NACA Rep. No. 832, 194.5.

18 NCA 'nr No. im TABLE I. - ORDINATE FOR THE NACA 835216 AIRFO]1 - [Stations and.. ordinates given in percent of airfoil chord] Upper surface Lover 8urface Station Ordinate Station Ordinate 0 0 .321 1.090 .679 -.890 1.311.2 .550 .950 -1.0514.

1.11.70 1.030 1.7311.

-1.292 2.268 2.11.74 2.732 -1.714 4.793 5.207 -2.391 3.591' 7.342 4.453 7.658 -2.979 9.901 5.162 10.099 -3.528 6.264 15.034 14.966 -4.4 20.186 7.036 19.8111.

-5.506 25 . 369 24.631 7.523 -6.405 30.589 7.7114. 29.411 -7.246 35.832 7.567 34.168 -8.023 40.881 7.1511.

39.119 -8.748 45.574 6.687 44.426 -9.203 50.234 6.189 49.66 -9.239 54.967 5.666 -8.812 55.033 59.778 5.123 60.222 -8.037 64.675 4.567 -7.027 65.325 69.6611. 4.'oio 70.336 -5.831 74.713 3.436 75.287 -4.552 2.8311.

79.793 80.207 -3.246 84.883 2.205 85.117 -1.989 89.963 i.54o 9O.O37 -.878 .821 95.009 94.991 -.072 100.000 0 100.000 0 L.E. radius: 1.121 Slope throu€h L.E.: 0.191

NACA TN No.

TABLE II.- ORDINATZ FOR THE NACA 836A216 AIRFOIL [Stations and ordinates given in percent of airfoil chord] Upper surface Lover surface Station Ordinate Station Ordinate 0 0 0 .303 1.145 .697 -.927 . 530 1.402 .970 -1.092 1.006 1.790 1.4911 -1.3111.

2.232 2.768 -1.715 2.541 11.735 3.680 5.265 -2.326 7.266 4.580 7.734 -.2.852 5.311.11.

9.816 10.184 -3.354 14.945 -4.276 6.522 15.055 20.094 19.906 7.359 -5.139 25.279 7.904 24.721 -5.960 8.134 29.467 30.533 -6.742 8.004 34.220 35 . 780 -7.484 40.900 7.620 39.100 -8.186 45 . 910 7.085 44.090 -8.803 50.822 6.453 49.178 -9.273 5.769 54.375 55.625 -9.533 60.311 5.067 59.689 -9.469 64.912 4.374 65.088 -8.9311.

69.627 3.703 -7.855 70.373 74.560 3.100 75.440 -6.450 79.605 80.395 -4.873 2.515 84. 735 --1.945 85.265 -3.239 1.3112' 89.873 90.127 -1.688 94.981 .707 -.433 95.019 100.000 0 100.000 0 L.E. radius: 1.183 Slope through L.E.: 0.208

NACA T No. 1771 TABLE III.- ORDINATFB FOR TBE NACA 8147A216 AIRFOIL (Stations and. ordinates given In precent of airfoil chord] Upper surface Lower surface Station Ordinate Station Ordinate 0 0 .280 1.169 .720 -.9149 1.1430 .501 -1.108 .999 .962 1.873 -1.359 1.538 2.176 2.677 2.8214. -1.763 14.673 3.830 5.327 -2.2914 7.192 14.738 7.808 -2.732 5.14.90 9.720 10.280 -3.118 i14.8oi 6.692 15.199 -3.816 19.905 7.606 20.095 -4.14614.

25.027 8.279 214.973 -5.079 30.170 8.7314.

29.830 -5.670 314.. 611 .8 8.968 35 . 352 -6.21414.

14.0.605 8.920 -.6.810 39.395 14.5.825 1414.175 -7.358 50.931 7.962 149.069 -7.880 7.226 55 . 909 514.091 -8.328 60.786 6.1406 59.2114.

-8.626 5.5145 614.11.39 65.561 -8.697 70.189 14.697 69.811 -.8.14.09 7)4. 761 3.873 -7.525 75.239 3.110 79.569 80.1431 -6.050 814.635 2.367 85.365 -14.261 89. 790 1.602 90.210 -2.401# 914.952 -.714.0 .820 95.0148 100.000 0 100.000 L.E. radiuB: 1e2142 Slope through 0.220

NACA TN No. 1771 TABlE IV,- ORDINATES FOR THE NACA 836Ano AIRFOIL [Stations and ordinates eiven in percent of airfoil chord] Upper surface lover surface Station Ordinate Station Ordinate 0 0 14 .26 .

.700 -.569 .575 .662 .861 .831 -.672 1.1 145 1.110 1.336 -.816 2.378 1.578 2.585 -1.078 14.881 2.262.

5.076 -1.480 7.392 2.818 7,563 -1.810 9.909 3.297 10.045 -.2.117 14.958 114.997 4.033 -2.680 20.017 4.564 19.940 -3.217 25.089 4.917 24.869 -3.732 30.189 5.072 29.772 -4.226 35.286 5.009 34.677 -4.703 40.334 4.788 39.632 -.5.162 45.342 4.475 44.627 -5.5514.

50.309 4.108 . 14.9.663 -5.870 55.255 3.699 54.721 -.6.047 60.113 3.279 59.867 -6.025 64.956 2.866 65.026 -5.707 69.843 2.14.65 70.144 -5.052 74.816 2.093 75.174 -4.177 79.835 1.722 80.157 -3.187 1.3115 84.888 . 85.107 --2.145 89.946 90.052 -1.144 . .935 914.992 .1493 -.318 95.007 100.000 0 100.000 0 L,E, radius: 0.498 Slope t1irouh L.E.: 0.119

NACA TN No. 1771 TABLE V.- ORDINATES FOR TEE NACA 836B110 AIRFOIL [Stations and ordinates given in percent of airfoil chord] • Upper surface Lower surface Station Ordinate Station Ordinate 0 0 0 0 .7314.

.371 .11.81 - .611.3 .600 .893 .7211. - .765 1.075 1.1143 1.215 - .943 2.302 1.580 -1.238 2.4511.

4.808 2.228 14.9511. -1.679 7.1463 7.321 -2.035 2.750 9.850 3.195 9.968 -2.311.7 1 14..92 3.902 -2.889 14.991 20.015 14.1428 20.019 -3.356 25.125 4.808 25.053 -3.768 30.277 5.053 30.101 -4.144 35.425 5.153 35.147 ...4.14.95 40.501 5.128 40.167 -4.779 14.5.516 5.001 14.5.168 -4993 50.469 4.794 50.149 -5.123 55.391 4..519 55.119 -5.145 60.181 4.180 60.045 -5.017 3.784 64.963 -.4.690 64.949 69.781 -4.1311.

3.326 69.905 714.892 74 .740 2.828 -3.423 79.768 2.282 -2.617 79.902 84..84 1.707 84.931 -1.771 89.929 1.110 89.963 • - .956 94.993 - .280 94.99]. .525 100.000 0 100.000 0 L.E. radius: 0.659 Slope through L.E.: 0.095

NACA 'IN No. 1771 23 TABLE VI.- OBDINATES FOR THE NACA 836cuo AIRFOIL (Stations and ordinates given in percent of airfoil chord] Upper surface Lower surface Station Ordinate Station Ordinate 0 0 0 0 .7311 .li.81 .371 -.611.3 .600 .893 .7211.

-.

1.075 1.1113 1.215 -.943 2 .302 2.11511 -1.238 1.580 4.808 2.228 4.9511 -1.679 7.321 2.750 7.463 -2.035 9.850 3.195 9.968 -2.347 111.925 -2.889 3.902 14.991 20.015 11.428 20.019 -3.356 25.125 11.808 25.053 -3.768 30.277 5.053 30.101 -4.111.11.

35.11.25 -11.1185 5.153 35.147 110,501 5.128 40.167 -4.779 115.516 5.001 11.5.168 -4.993 50. 11.69 11.7911. 50.1149 -5.123 -5.114.5 55.391 11.519 55.119 60.181 4.180 60.0115 -5.017 64.963 3.784 64.949 -4.690 69.781 3.326 -11.131l.

69.905 714.892 74.740 2.828 -3.1423 2.335 -2.670 79 . 767 79.903 811.8112 811.9311.

1.8144 -1.907 89.925 1.325 89.967 -1.171 .711.5 -.499 94.991 94.994 100.000 0 100.000 L.E. radius: 0.659 Slope through L.E.: 0.095

214.

NA(A TI là. 177].

TABLE VII.- ORDINATES FOR TEE NACA 836Duo AIEFO]1 [Station and. ordinates given in percent of airfoil chord] Upper surface Lower surface Station Ordinate Station Ordinate 0 0 .14.43 .769 -.678 .558 .681 .812 -.808 .937 1.166 1.2014. 1.315 -1.0014 2.400 1.6914. 2.5611.

-1.352 2.1106 4.898 5.070 -1.858 7.1106 2.959 7.560 -2.244 3.1122 9.913 10.0111 -2.574 14.942 4.135 15.014 -3.122 19.977 4.650 ' 19.981 -3.578 25.016 5.004 24.942 -3.964 30.071 5.204 29.891 -4.295 35.124 5.238 34.840 -4.570 40.149 5.119 39.817 -4.770 4.884 114.8i 45.155 -4.876 50.139 11.9.833 -4.887 4.160 54.862 55.114 -.4.786 60.051 3.709 -4.547 59.929 3.251 64.995 -4.157 64.997 69.941 2.793 70.047 -3.601 74 .932 2.325 -2.920 75.058 79.911. 1 1.847 80.051 -2.182 84.964 1.366 -1.430 85.032 89.985 .885 90.013 -.732 95.001 .427 94.999 -.182 100.000 0 100.000 0 L.E. radius: 0.618 Slope through L.E.: 0.098

NACA TN No. 177].

TABLE VIII.- ORDINATES FOR THE NACA 847A110 AIRFOIl [Station and. ordinates given in percent of airfoil chord.!

Upper surface Lower surface Station Ordinate Station Ordinate 0 0 0 0 . 14 .29 .738 -.628 .5'].

.669 .902 .831 -.740 1.3143 1.157 1.173 -.915 2.396 1.652 2.60)4.

-1.19)4.

4.896 2.324 5.104 -1.556 7.403 2.877 -1.853 7.597 3.298 10.088 -2.112 9.912 14.938 4.010 15.062 -2.572 19.970 4.563 20.030 -2.991 25.008 4.975 24.992 -3.375 30.053 5.267 29.947 -3.735 5.437 34.890 -4.075 35.110 40.190 5.454 39.810 -4.398 45.259 5.297 44.4i -4.697 50.293 5.001 49.707 -4.959 55.286 4.620 54.714 -5.172 60.248 4.188 59.752 -5.298 3.729 65.178 64.822 -5.305 3.246 70.060 --7.102 69.940 74 .923 2.753 75.077 -4.579 2.236 79.859 80.141 -3.706 84.880 1.697 -2.645 85.120 89.930 1.134 90.070 -1.536 94.984 95.016 .548 -.508 100.000 100.000 0 0 L.E. radius: 0.590 Slope through L.E.: 0.104

26 NA.CA TN No.

814.7B110 AIRFOIL TABLE IX.- ORDINATES FOR TEE NACA [Stations and. ordinates given in percent of airfoil chord] Upper surface Lower surface Station Ordinate Station Ordinate 0 0 O 0 .1414.5 .792 -.711 .555 • .688 .9611. .812 -.814.7 1.179 1.239 1.321 -1.055 2.11.20 1.723 2.580 _1.11.00 2.11.13 14.918 - 5.082 -1.871 7.421 2.947 -2.233 7.579 9.926 -2.537 10.0711.

3.390 -3.038 14 .9142 4.095 15.058 19.962 20.038 -3.447 4.631 214.985 25.015 -3.784 5.031 30.011 29.989 -4.057 5.309 5.11.67 -4.270 35.044 34.956 5.11.90 39.913 -4.14.26 14.0.087 145.1214. 114.876 -4.498 5.346 4.14.99 49.859 50.141 5.072 54.863 -4.433 55.137 4.705 4.273 59.881 -4.293 60.119 3.796 64.913 -4.069 65.087 69.963 -3.722 70.037 3.280 74.982 2.711.1 75.018 -3.207 2.179 80.043 -2.517 79.957 -1.714.4 84.965 1.606 85.035 90.018 89.982 1.033 -.976 .14.77 95.001 -.299 94.999 100.000 0 100.000 0 L.E. radius: 0.693 0.080 Slope through L.E.:

NA.CA T No.

TABLE X.- ORDINATES FOR TEE NACA 847db AIRFOIL [Stations arid ordinates given in percent of airfoil chordi Upper surface Lover surface Station Ordinate Station Ordinate 0 0 .1114.5 .792 .555 -.711 .688 .961i. .812 -.814.7 1 .179 1.239 -1.055 1.321 2.14.20 1.723 2.580 -1.14.00 2.14.13 4.918 5.082 -1.87].

7.421 2.947 -2.233 7.579 9.926 10.074 3.390 -2.537 114.942 4.095 15.058 -3.038 4.631 19.962 20.038 -3.447 24.985 5.031 25.015 -3.784 30.011 5.309 -4.057 29.989 35.04 1 1. 5.467 34.956 -.4.270 40.087 5.490 -4.426 39.913 45.124 5.346 44.876 -J..498 50.141 5.072 -4.499 49.859 55.137 54.863 4.705 -4.433 60.119 4.273 59.881 -.4.293 65.087 3.796 -4.069 64.913 70.037 3.280 69.963 -3.722 74 .982 2.741 75.018 -3.207 2.210 79.957 80.043 -2.514.8 84.963 1 .716 85.037 -1.853 89.979 1.218 90.021 -1.161 94 .998 .684 -.506 95.002 100.000 0 100.000 L.E. radius: 0.693 Slope through L.E.: 0.080

N&(A TN No. 177].

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11.0 N&CA TN No. 177].

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NCA TN No. 1771 - _I

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NACA TN No. 1.5 iT'll •aIBN _.-iiiiiii : UURMI1IiIIIIB_ I_II_1I1fflh1U_ IL .1R I_IIiiiøI'III .1 II III_WIFIJJI_ _•-muiiiv mu_ rum- • mma_mumm--iiitauim_ m•_ummuiiiivuim_ •mmmmmmuiiariiumm.

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NACA TN No. 1771 li6 • 4 .3 E / -4 -c .6 .7. .8 .9 ID .3 4 5

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Figure II.— The variation of section quarter-chord moment coeff/c/enP

• with Mach number at various angIe of attack for the NA CA

835A216 airfoil. • ..

NACA TN No. 1771 17 IIUUUUUUU UUUUUI IUUUUU_UUU UUUUUR IUUUUU_UUUUU_U U !UIIUUU-UUUUU_UUU I._UUU__!UUUU UUU U_ UIUIUUUUU SC,) U UIU_U UUUUM\1I1U UU_UUU_ U _ UU_U-UUU_1-.

U ':U_U ..- UU_Ill U .

U_U_U U_11 U I __ u-u_U U UI_U U UU_UU-__U_U U UUU_U-UU -UU U UUUUUUUU1N ! UUUUUUUU UU Mach number, M Figure /2.- The variation of sec//on quarter-chord moment coef-.

f/dent with Moch number at various angles of attack for the N4CA 83642/6 airfoil.

NA.CA TN No. 1771

.3

• .1

.

IC

-

-'.-

I I I I I I I I _61 9 10 5 .6 .7 .8 .3 .4 Mach number, M Figure /3.- The var/allen of section quarter-chord moment coef- ficient with Mach number at various angles of altaok for the NACA 847A2/6 airfoil.

NACA TN No. 1771 li.9

1.2 _u. a a.aaa_a_u _ a..

I-

IL

•a...._-a

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Ill U.l dll -5 lINI#lUUAU _.IPJ.I_ ll_•..U, -8 U._I...._l _ UI UUIIU-U U' -to 0 3 .4 .5 .6 .7 .8 .910 Mach number, U Figure /4.- The var/ct ion of sect/on 1/ft coefficient with Mach number at various angles of attack for the NACA 836A I/O airfoil.

50 NACA TN No. 1771 /2 'U _ 1.0 U..,.__ _ • nil__ _ .8

UUEUL1HU

.6

I....._lPU!Ui

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l•lU• _ _

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.6 .7 .8 .9 10 .3 4 .5 Mach number, M Figure /5.- The variation of section Ii!t coefficient with Mach number at various angles of attack for the NACA 8368/10 airfoil.

NACA TN No. 1771 51 1.2

I I - I

Va:

0a-6° 4°

_40

o

6° - - - - - -2° /00 N 0° ii 1 . 2°V /2° .6 'S .,..

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.5- 'S '3 CI).4 -.6 -.8 I I I I I ..,q .9 1.0 .7 .8 4 5 .6 '0 .3 Mach number, M I/f / coefficient with Mach Figure /6.- The var/at/on of section NACA 6'36C//0 number at various angles of attack for the airfoil.

NACA TN No. 1771 1.2 1.0 .8 .6 :4 C.)

-8 I I I I I I I I I I I I 3 .4 5 .6 .7 .8 .9 1.0 Mach number, U Fiqure /7. The variation of sect/on I/f I coefficient s'/th Mach number at various angles of attock for the N4C4 836D1/0 airfoil.

N&CA T& No. im /2 .4.

_••I17.

'I_ _ I'; i ...

_ B _•suuiiumu •• .6 _iuuu•ui _•ui•u••u•ua AU _..__••••_..

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-86 J 4 .5 .6 7 .9 /0 .8 Mach ,rnrnber, M Figure /8.- The variation of sect/on I/ft coefficient with Mach number of vor/ous angles of a/tack for the N4C4 64-I/O airfoil.

51.

NACA TN No. 1771 .32

_11111111

__ NACA 8364/10 - ----- NACA 836B//O - .28 NACA 836C//0 - ¼ --- NACA836D//0 NACA 64-I/O -4 -20 / -' -

---------------,c------

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-

¼

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--'L--

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-

rsi V 1.0 •4 .5 .6 .7 .8 .9 Mach number, M lift- curve Figure /9. - The var/at/on with Mach number of the section slope at the design lift coefficient for the NAC4 8364/10, 836B110, 836C1/0, .836D/l0 and the 64-I/O airfoils.

NA.CA TN No. 1771

H __

NACA 83641/0 -NAA 83681/0 4 - NACA 83601/0 - --- NACA 836Dll0 NACA 64-I/O a i,2 - - -_ ---- - __ '3 - .-.

C.)

(2 -3 .3 .4 .5 .6 .7 .8 .9 .L0 Mach number, M Figure 20.- The variation with Mach number of the sect ion angle of attack for a I/ft ccefficient of 0.! for the NAC4 83641/0, 836B/l0, 8360/10, 836D//0 and 64-I/O airfoils.

NACA TN No. 177].

.22

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20 0a-6° ----- 0 $ 4e -- —2° -- -------- ------ 0° .18 -- 6°

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8°

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.08 .06 .04 .02 .3 4 .5 . .6 .7 .8 .9 10 Mach number, M Figure 2/.— The var/cf/on of section drag coefficient with Mac/i number at various angles of attack for the NACA 836A//0 airfoil.

NACA TN No. 1771 .22 .20

'1

.18

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; •

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i••u••uawwui ••

.02 .3 .4 .5 .6 .7 .8 .9 Mach number, M Figure 22.- The var/c//on of section drag coefficient with Mach number at various angles of c/tack for the NACA 8368/10 airfoil.

58 NACA TN No. 1771 .22 IIU.IUIUUIUI.IUI II IIUIIPNIIUU : II URIIUFIIIU• /8 II I.II..II.

UI 111.1w_UI i .16 IIPIWUIII II: ll11111 II 4- ll!lPllllIJRUll ImdIIIIIUIII IIIIIIIIIIIJI!ANIU I.)

UllllUlllh1IUll llIlllllllII4UlU .08 llllUlllliM1Ul (1) •lUlUPIalmtP1•l .06 IflllUFililll .04 llllUlW!JW4WJllI .6 .7 .8 .9 10 .3 .4 .5 Mac/i number, M Figure 23.- The var/ct/on of sect/on drag coefficient with Mach number at various angles of attack for the VI4CA 836C/I0 .a/rfoI/.

NA.CA TN No. 1771 .18 uuumurnwiuuu /6 I.: •RRRRIV1....

US :UUUU5iPJUS us uuuuviuuvuu S.MU.UU.U•U.1UIU UUUSUUNIVUSU urns uuuunsiiuu uiumumuuruviuvum UUUUUU.RU,I,11UU.

•uuuuuu twruiiu_ .04 UIUBSIWALI- uu•uuurw;triuuu .02 P'! .

Mac/i number,M Figure 24.- The yap/a/ion of sect/on drag coefficient with Mach nz.imber at various angles of attack for the NI4CA 83601/0 airfoil.

60 NACA TN No. 1771 _ • -

____ ii huN

;

i liii- _Al_

—fI_ __ 1 I II RI_Ill_ -: ;; I 11-11101 -A S.._ ,I11 Ill_ _ _ II Ul_Mull U_4 _.l_I _ uuy_u_ •l uiu • u—• waiiuu_ ___i_II I11W1.

S _•ll VIffll_ _••_ii__ _ i_i_a _ I lE1tiJ1Z1l II I uuuiuirirariwii__ _ a iuri•__ Jl_I IldlWi11J'i1PiU — ___ K, .- wi 1jw .uI_,A!jlèIIl,

NACA TN No. 1771 61 .5 a asma_aaaaaaaaa 11U1_aaaauaaaaaa mama_aaaaaaaauaa mama_aaaaaaaaaa .3 __aa_am..._aa_a mama_mamma_mam_a •..a_mamma ariaa iiaa a..._aaaaa mlamai•i iiaa mmaaumiiaima • aaaamm-aamma a ':U•U•••_arncimm : I;

amaaaa_aama

-3

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4 .5 .6 .7 .8 .9 10 .3 Mach number, M Figure 26.- The variation of section quarter-chord moment coef- ficient with Mach number at various angles of attack for the NACA 836A I/O ohio/I.

NACA TN No.

177].

.5 U..._._._...._ U..._I--._U......

IUUUUU iUUUUUU IUUU_UT _UUUUUUU .....I1..

•UUU_UU UUUU_UI_Ii Will UIU•_UU__UUUFAUII UUIUI Sill lUll '3 MUllU UI --u .

•uai_.i1S.0

1-i

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.uurn _ 5 O .3 4 .5 .6 .7 .8 .9 i0 Mach number, M Figure 27- The var/of/on of section quarter-chord moment coef- f/dent with Mach number at various ongles of attack Por the NACA 83681/0 airfoil.

NACA TN No. 177].

i _a_I I II a• _I—IIIII• 11111_III...

lull_I_I_.II_I _II•II_IIC a_U_I IUI_III uriiiuu

I

I_Il_IàlI_I

I

I:II_u_u.rniuu_I l__I_IIIII_I IIIIlIIII '.3 •!IIII_III_II Il—II_III III uu—IuuuI lull 11111 II_uIu_ UI 1111111111 Mach number, M Figure 28. - The variation of section quarter-chord moment coef- ficient with Mach number at various angles of attack for the NACA 836C110 airfoil.

61 NACA T No. 177].

.4 U UUUUUUUUUUUI RU-U_U-UUUUUUUUU Ii ii__U-UUUUUUUU UUUUUUIIUU UU_U U RU_UUU_UIUUUUU .q) UUUPUUU RU_U_U .c .1 UUUUUWA1IUU _U_U U_:UUURk!?UU .

U_U-UUUPUUU I_IUTUUUUUU U__UUUUtUUU I_ UUU UUU_UUUUU RU_UUUUU UUIU_U_ lUll_U_UUU_UUUUU -.

_ UUUU-U_UUU _U U UUUUU UUUU .6 .7 .8 .9 10 .3 4 .5 Mach number, M The variation of section quarter-chord moment coef- Figure 29. — ficient with Mach number at vorious angles of attack for the NACA 8360/10 airfoil.

RWA TI No. 177].

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T: Mach number, M Figure 30.- The var/of/on of sect/on quarter-chord moment coef- ficient with Mach number at various angles of attack for the NACA 65-2/0 airfoil.

NPCA N Nci. 177].

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.4 .5 .3 3 .7.8.9 f0 Mach number, Al Figure 3/.— The variation of section lift coefficient with Mach number at various angles of attack for the NACA 8474/10 airfoil.

NACA Th' No. 1771 1.2 I. C .8 r------ - .- Uu..

•uraII1•uu _ .6 uti•u uuu•iu .4 U._UU___.U.,..

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1.00 4 5 .6 7 .8 /0 - .9 Macb number, M Figure 32.- The variation of section lift coefficient with Mach number at various angles of attack for the NACA 8478110 airfoil.

VJN

68 ICA No. im 1.2 1.0

'1jul.11

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-.6 .1803 4 5 .6 .7 8 .9 10 Mach number, M Figure 33.— The variation of section I/ft coefficient with Mach number at various angles of attôck for the NACA 847C110 airfoil.

NkCA No.

177].

.28 ¼

HH

11:111=1

NACA 847Al/0 .24 .-. NACA 8478/10 - - - - - C.)

N4CA 847C//0 NACA 64-I/O - -- I -------- - - --- - - ¼

IIi!i1IIIIII

---- .08 - - ----- - --7'- ni 2 .3 .4 .5 .6 .7 Mach number, M Figure 34.— The variation with Moch number of the section• lift- curve slope at the design lift coefficient for the NACA 84741/0, 8478/10, 847C1/0 and 64-110 airfoils.

NACA TN No. 1171

11111111

NACA 84741/0 NACA 8478/10 -----NACA 847C//0 -----NACA 64-I/O - - C.)

'I- l) ---------------- 7 3 4 .5 57 tr Macb number, M Figure 35.- The variation with Mach number of the section angIe of attack for a I/f I coefficient of 0.1 for the NACA 84741/0, 84781/0, 847C/I0 and 64-I/O airfoils.

RACA ¶1W No. 1771 __i _a_ _ a .1UU _ !IJa a- __ __law...

__ __a 1111 a..

a__•,I._a _RU • IUaa a-: - aarniva A - _a_riawi_a _ __aa a•,1.I._a • Ia_ _ ,j. U na -hi_a __•_a W R1111111 •Iji_U I _ ___

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NACA TN No. 1771 .22 iIii___1TIU_I.

u.a_ _ria-_I_

.18

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Figure 37.- The variation of section drag coefficient wi/h Mach number at various angles of attack for the NACA 8475/10, airfoil.

NACA TN o. 1771 • .22 Us.._U._•_ U _ U.s. US_ RUCU- .20 _ UIU_U!URU : • :5R FARUU .18 •_ :RI _UI •_ USUUUPIUU- .16 •_ _ •_RUWIU_II-IRRU-- U_IUAU_UUMWI__ Ui•U_U_UUUIVUU /2 ••U_..••rn.r,4s_ •_RU UUU-IWAI1U_ U_RU •UUUUTIIZIUUU_ U_RU U-UVWIUUU_ I.).

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V .3 .4 .5 .6 .7 .8 .9 1.0 Mach number, Al figure 38.- The variation of section drag coefficient with Mach number at vorious angles of attack for the N14CA 847C1/0 airfoil.

711 NACA TN No.

177].

_••........I _.

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_ _ Q) UURUUUU UU U_UUUUUUUUEIUUU _ _U _

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liiUUUUmiiuU _•_UUUUUUULUU UU_UUUUUUIUU _U 0 .,.- __U•U_UUU_•U•U _••••• •••_••••• _UUUUUUUUUUUUUU -J

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3 4 5 .7 .6 .8. .9 /0 Mach number, •M Figure 39.- The var/at/on of section quarter -chord moment cóef- ficient with Mach number at various angles of attack for the N4CI4 8474//U airfoil.

NACA TN No. 1771 _ __•• _ _ UN U- •aw.amuuuiiuu -_ r ••a••••••a•_U LU s.m.iuuvau_ - w. - - - -.. e .

tIUU_UUsiUU •••_•••••_••• U : UUUUUUUU -IN - ••UI..IUIU_UN UUUUUUUUUUUU_III UUIUUUUUUUUU_ 1, !'!UU!U!U

NACA TN No.

1TT1 •uUUlU.IU.RR INI iuum•ui•_ii_i•• uiu_••• ••_••• im•um_iu •1_••u .1.1_u •1ll•_U..

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u..i.ui•u_uuu -4 ..u-U....... _ .3 4 5 .6 .7 .8 - .9 /0 Mach number, M Figure 41.-The variation of section quarter-chord moment coef - ficient with Mach number at various angles of ctock for the NACI4 847G//0 airfoil.

NAA TN No. 1771 1.4 INN_N UNNII NUN_U NUNS- _N NNU INN U NPUiNN •NiPAUNE Ni.• Nutriøi- • NRNNUN_ __N uaU•ii- .

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N I N NNNNN_ -.2 •NUNNNNNUN_ • NN NtININ_ -.4 NNNN'!UN_ U_•_ _ -.6 N_N_N_NNNu-iu N N N_NNNUIII .80 .4 .3 .5 .6 .7 .8 .9 10 Mach number, M Figure 42.- The variation of sec/ion lift coefficient with Mach number at various angles of at tack for the NI4CA 65-2/C airfoil with a 20-percent chord plain flap. 8, 0°

NACA TN No.

•iu•uumumii•m 1.0 •uuauu#iuiuu .8 uuuuuuwivauu• .6 •uuaia•• iuu

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NACA TN No.

Li _

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84781/0 8 —

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Mach number, M Figure 44.- The var/at/on with Mach number of the sect/on lift- curve slope at a lift coefficient of 0.1 for the N4CA 65-210 (c -6), 8360110 and 8478110 air foils.

NACA TN No. 1771

I

LI LI I .L 111111

NAOA 83601/0

--

NACA 847B//0 - - - - - -

--- NACA 65-2t0 (8t-6

Li .

- - - = - - —= -

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-3 - .3 4 5 .6 .7 .8 .9 1.0 Mach number, M Figure 45- The variation with Mach number of the angle of attack for a lift coefficient of 0.1 . for the NACA 65-2/0(9jr=-6°,), and 8478/10 airfoils. ....

83601/0

NACA TN No. 177].

8i I

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Mach number, M Figure 46.- The variation of sect/on drag coefficient with Mach number at various angles of attack for the NACA 65-2/0 airfoil with a 20-percent chord plain flap. 8, -6°

NACA TN No. 1771 UU•UUIIUUUU _ I U -_I LI IUU_U_UI_---II.

BlEW U_U_au__III .- UUU_lUll II 111111_U_•U • C.)

UI_UI..

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I II_11111__II_III U___U_IU UI_III II_11111 I..IJ_U UIUIIIU III _U .9 1.0 .3 .4 .5 .6 .7 .8 Mach number.. M Figure 47- The variation of section quarter-chord moment coef- f/dent at various angles of attack for the NI4CA 65-210 airfoil with a 20-percent chord plc/n flops , -6°