airfoil designed lift of - nasa technical memorandum 81912 low-speed aerodynamic characteristics of...

NASA Technical Memorandum 81912

Low-Speed Aerodynamic Characteristics of a 14-Percent-Thick NASA Phase 2 Supercritical Airfoil Designed for a Lift Coefficient of 0.7

Charles D. Hrrris, Robert J. McGhee, and Dennis 0. Allison L,atigIey ResearA Ceriter Hamptoir, Virgt'tiia

National Aeronautics and Space Administration

Scientific and Technical Information Branch

1980

https://ntrs.nasa.gov/search.jsp?R=19830013896 2018-05-28T04:29:56+00:00Z

Airfoil characteristics varied systematically Over the range of test conditions and behaved in a conventionally accepted manner. lift coefficient was about 2.2 and occurred near M = 0.15 ber of 12.0 x lo6. Mach numbers and at all Reynolds numbers exhibited behavior generally associated with gradual trailing-edge separation in the prestall angle-of-attack range, but the stall itself was more abrupt than would be expected of classical trailing-edge stall. The stall became more gradual at the higher Mach numbers. Drag remained essentially constant over a lift range which extended frtn near zero to beyond the design lift coefficient of 0.7 for a constant Reynolds number with fixed transition.

The maxirun aeasured at a Reynolds n u t

The variation of lift with angle of attack at the lower

INTRODUCTION

Continued development of supercritical airfoil technology has resulted in the design of family-related phase 2 supercritical airfoils. 10- and 14-percent-thick, designed for a lift coefficient of 0.7 have been tested at transonic speeds in the Langley 8-Foot Transonic Pressure Tunnel, and the results were reported in references 1 and 2.

Two such airfoils,

The 14-percent-thick airfoil has also been tested at low speeds in the Langley Low-Turbulence Pressure Tunnel, and the results are presented herein. Included are the effects of varying Reynolds nunber from 2.0 x lo6 to 18.0 x lo6 at a Mach number of 0.15 and the effects of varying Mach number from 0.10 to 0.32 at a Reynolds number of 6.0 x lo6.

SYMBOLS

Values are given in the International System of Units (SI) and in U.S. Cus- tomary Units. Measurements and calculations were made in U.S. Customary Units.

cP P - P, pressure coefficient, -

&,

Cp,sor.ic pressure coefficient corresponding to local Mach number of 1.0

C chord of airfoil, cm (in.)

CC 4j cp $) section chord-force coefficient,

Cd section drag coefficient determined from wake measurements, f /h\

Cd ’ Cl

cm

Cn

h

1 5

M

P

Q

R

t

X

Y

z

=C

a

point drag coefficient (ref. 3)

section lift coefficient,

section pitching-aosent coefficient about quarter-chord point,

Cn(C0S a ) - cc(sin a)

section normal-force coefficient, -($% d 6 )

vertical distance in wake profile, a (in.)

c2 section lift-drag ratio, - Cd

free-stream Mach number

static pressure, Pa (lbf/ft2)

dynaric pressure, Pa (lbf/ft2)

Reynolds number based on free-stream conditions and airfoil chord

airfoil thickness, cm (in.)

abscissa measured along airfoil reference line f r m airfoil leading edge, an (in.)

spanwise distance fraa centerline of tunnel, Q (in.)

ordinate measured normal to airfoil reference line, a (in.)

ordinate of airfoil mean line, cm (in.)

angle of attack, deg

Subscripts:

1 lower surface

M X maximum

min minimum

U upper surface

a, f ree-stream conditions

ORIGINAL PAGE IS OF POOR QUALITY

2

Airfoil designations8

SC (1 ) -071 4 supercritical (;.hare 1) -0.7 design lift coefficient. 14-percent-thick

SC (2) -071 4 supercritical (phase 2) -0.7 design lift coefficient, 14-percent-thick

A m r t e d effort within the National Aeronautics and Qmce Achiniatra- tion ( M A ) over the past several years has been directed toward developing practical trro-dirsnaional airfoils with good transonic behavior while retain-

to as the supercritical airfoil. acceptable law-sped characteristics and has focused on a concept referred

An early phase of this effort was successful in significantly extending drag-rise characteristics beyond those of conventional airfoils. (see ref. 4.) These early supercritical airfoils (denoted by supercritical (phase 1) prefix), however , experienced a gradual increase of drag (drag creep) at nacb numbers juat preceding the final drag rise. This gradual buildup of drag was largely associated with an intermediate off-design Mcond velocity peak (an m l e r a - tion of the flar over the rear portion of the airfoil just before the final recompression at the trailing dge) and relatively weak shock waves above the upper surface at these speeds. (See, for example, ref. 5.)

Improvements to these early, phase 1 airfoil. resulted in airfoils with significantly reduced drag creep characteristics. (See, for example, refs. 6 and 7.) These early phase 1 airfoils and the improved phase 1 airfoil8 of references 6 and 7 were developed before adequate theoretical analysis codas were available and resulted frolr iterative contour rodifications during wind-tunnel exper inents. The process consisted of evaluating experimental pressure di8trf- butians at &sign and off-design amditions and physically altering the airfoil profiles to yield the best drag characteristics over a range of teat conditions.

During the experimental develomnt of these phase 1 airfoils, design cti- teria were recognized which provided guidelines for the design of supercritical airfoils and are briefly discussed in references 1 and 2. Based on these criteria, two phase 2 (denoted by supercritical (phase 2) prefix) supercritical airfoils were designed: the 10-percent-thick airfoil reported in reference 1 and the 14-percent-thick airfoil reported in reference 2. The deaign lift coefficient was 0.7 for both airfoils. An iterative design process wa8 used which consisted of altering the airfoil coordinates until the recently rkvel- aped, viscous, airfoil analysis program of reference 8 indicated that t.% design criteria had been satisfied.

3

We1 I

Geometric characteristics of t h e SC(2)-0714 model are presented i n fig- u r e 1 , and section coordina tes are presented in table I.

The model was oons t ruc ted wi th a metal core around which plastic f i l l and two t h i n layers of fiberglass were used to form the contour of the airfoil. O r i f i c e holes of 0.lO-cm (0.040-in.) diameter were d r i l l e d perpendicular to I

the model s u r f a c e i n t o embedded p r e s s u r e tubes. The model had a chord of 61 a (24 in.) and a span of 91 CB (36 in.). The airfoil s u r f a c e was sanded i n t he chordwis5 d i r e c t i o n with number 400, dry sili;.on-carbide paper to provide an aerodynamically smooth f i n i s h . kO.10 lllp (k0.004 in.) .

1

f

The rodel contour accuracy was g e n e r a l l y t ' h i n

W i n d Tunnel

The Langley Law-Turbulence Pressure Tunnel (ref. 9) is a cl red-throat, s ingle-return t u n n e l which can he operated a t s t a m a t i o n p r e s s u r e s f r a a 10.1 to 1010 kPa (0.1 to 10 atm) w i t h tunnel-empty t e s t - s e c t i o n Mach numrbers up to 0.42 and 0.22, respec t ive ly . per meter (15.0 x l o6 per f o o t ) a t a Mach number of about 0.22. test s e c t i o n is 91 cm (3 f t ) w i d e by 229 cm (7 .5 f t ) high.

The maximm u n i t R e y n o l d s number is about 49.0 x l o 6 The t u n n e l

Hydraul ica l ly ac tua ted c i r c u l a r plates provide posit' .ning and at tachment for two-dimensional models. The plates are 102 ca (40 in.) i n diameter, rotate with the airfoil, and are f l u s h with tbe tunnel w a l l . ment plates i n the c i r c u l a r plates hold the nodel i n such a way that the c e n t e r of r o t a t i o n for angle-of-attack a d j u s t n e n t w a s a t 0 . 2 5 ~ on the nodel r e f e r e n c e l i ne .

Rectangular model a t tach-

A s k e t c h showing the r e l a t i o n s h i p between the ends of t h e model, t he tun- ne l walls, the model attachment p l a t e s , and t h e rotating c i r c u l a r plates is shown i n f i g u r e 2. W i t h t h i s mounting system, the model completely spans t h e t u n n e l with each end a n f i n e d i n a recess i n t h e tunnel sidewall. A t t he side- walls, i n s i d e j o i n t sea- between mating pieces were sealed and f a i r e d smooth with model plastic and silastic rubber to minimize the e f f e c t of a i r leakage.

Measurements

Surface-pressure measurements.- S ta t ic pressures were measured on t h e s u r f a c e of the m o d e l and used to determine local surface-pressure c o e f f i c i e n t s . The surface-pressure measurements were obtained from a chordwise row of orifices located approximately 5 cm ( 2 in . ) f r a n t h e t u n n e l c e n t e r l i n e . O r i f i c e s were concentrated near t h e leading and t r a i l i n g edges of t h e a i r fo i l to d e f i n e t h e pressure g r a d i e n t s i n these regions, and a rearward-facing o r i f i c e was included i n the t r a i l i n g edge. I n addi t ion , three spanwise ~ O W R of o r i f i c e s , loca ted a t 1 - , lo- , and 75-percent chordwise s t a t i o n s , were included to e s t a b l i s h t h e two d imens iona l i ty of the f law over t h e model.

4

Wake measurement.- Drag forces acting on the airfoil, as determined by the m e n t u m deficiency within the wake, were der:ved from vertical variations of the total and static pressures aeasured across the wake with the wake survey rake shown in figure 3. positioned in the vertical centerline plane of the tunnel, one chord-length behind the trailing edge of the -1. The total-pressure tubes were 0.15 a (0.060 in.) in diameter and the static-pressure tubes were 0.32 cm (0.125 in.) in diameter. *

The fixed rake, aounted from the tunnel sidewall, was

The total-pressure tubes were flattened to 0.10 u (0.040 in.) for 0.61 Q

(0.24 in.) from the tip of the tube. flush orifices, drilled 900 apart, located eight tube diaraeters from the tip of the tube in the measurement plane of the total-pressure tubes.

The static-pressure tubes each had four

Instrumentatior

Measurements of the static pressures on the airfoil surfaces and the wake pressures were made by an autoaatic pressure-scanning system utilizing variable- capacitance-type precision transducers. with precision quartz mananeters. brated digital shaft encoder operated by a pinion gear and rack attached to the circular @el attachment plates. tion system and recorded on magnetic tape.

Basic tunnel pressures were measured Angle of attack was measured with a cali-

Data were obtained by a high-speed acquisi-

Reduction of Data and Corrections

Calculation of cc, cn, and %- Section chord-force, normal-form, and pitching-manent coefficients were obtained by numerical integration (based on the trapezoidal method) of the local surface-pressure coefficient measured at each orifice multiplied by an appropriate weighting factor (incremental area).

Calculation of ca.- To obtain section drag coefficients, point drag ooefficients were computed for each total-pressure measurement in the wake by using the procedure of reference 3. summed by numerical irtegration across the wake, again based on the trapezoidal method.

These point drag coefficients were then

Corrections for wind-tunnel-wall effects.- In the linear portion of the lift curve, corrections for lift effects and solid and wake blockage based on references 10 and 1 1 are small, on the order of 2-percent or less, and are usually neglected. As the model approaches maximum lift conditions where the lift characteristics become nonlinear and the viscous effect8 become significant, the assumptions underlying the corrections based on references 10 and 1 1 begin to break down and become inadequate. For these reasons, the data presented herein is uncorrected for tunnel-wall effects. however, to indAcate the effect corrections would have on the aerodynamic char- acteristicc if they were applied.

Figure 4 is presented,

5

?ZST CONDITIONS

Tests were conducted over a range of Reynolds numbers based on the wdel chord from about 2.0 x l o6 to 18.0 x l o 6 at a Nach number of 0.15. was also varied frola about 0.10 to 0.32 at a Reynolds number of 6 . 0 x lo6 . airfoil was tested smooth (natural transition) and 4 t h transition trim fixed along the 5-percent chordline on the upper and lomr surfaces. The traneition trips consisted of sparsely distributed 0.13-cla-wide (0.05- In.) bandp of car- borundum grains, sized according to the technique in reference 12 and attached to the surface with clear lacquer.

Mach number The

PRESENTATION OF RESULTS

The results of this investigation are arranged in the following figures:

Figure Effect of angle of attack and Reynolds number on the up;?er-surface spanwise pressure distributions. M . 0.15 . . . . . . . . . . .

Tuft photographs of the effect of angle of attack on the upper-

Effect of fixing tranuition on section characteristics. M = 0.15 Effect of fixing transition on chordwise pressure distributions.

M = O . l S . . . . . . . . . . . . . . . . . . . . . . . . . . . . Variation of pressure coefficient on upper surface at particular chordwise stations with angle of attack. M - 0.15 , model smooth

Variation of pressure coefficient on upper surface at particular chordwise stations with angle of attack. M = 0.15 , transition fixed . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

M - 0.15, transition fixed . . . . . . . . . . . . . . . . . . . . . . . .

surface flaw pattern. M . 0. 15; R . 2.0 . 106 . . . . . . . .

Effect of Re-ynolds number on section Characteristics.

Effect of angle of attack and Reynolds number on the chordwise

Effect of Reyr.olds number on base and upper-surface pressure

Variation of minimum drag coefficient with Reynolds number.

Variation of m a x i m u m lift coefficient with Reynolds number.

Effect uf Reynolds number on chordwise pressure distribution.

Variation of sectlon drag coefficient with Reynolds number.

Effact of Mach number on section Characteristics.

Variation of maximutn lift coefficient with Mach number.

Effect of Mach number on the chordwise pressure distributicc for R = 5 . 0 x l o 6 , transition

pressure distribution. M . 0 .15 , transition Fixed . . . . . . . coefficients. M . 0.15 , transition fixed . . . . . . . . . . . M t O . l 5 . . . . . . . . . . . . e o e o . . . . . . . . . . . .

M . 0.15 , transition fixed . . . . . . . . . . . . . . . . . . . M . 0.15 , a . 8O, transition fixed . . . . . . . . . . . . . . M . 0.15, cl . 0 .0 ?nd 0.7, transition fixed . . . . . . . . .

R . 6 .0 . l o 6 , transition fixed . . . . . . . . . . . . . . . . R = 6 . 0 x 196,

transition fixed . . . . . . . . . . . . . . . . . . . . . . .

angles of attack near maximum lift. fixed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . M . 0.32,

Rffef-t of angle of attack on the chordwise pressure distribution. R . 6 . 0 . l o 6 , transition fixed . . . . . . . . . . .

. . . 5

. . . 6 0 . . 7

. . . 8

. . . 9

. . . 10

. . . 1 1

. . . 12

. . . 13

. . . 14

. . . 15

. . . 16

. . . 17

. . . 18

. . . 19

. . . 20

. . . 21

6

Figure Bffect of Nach nwrb.r on the chordwire prerrur, dirtxibution.

R - 6.0 x lo6, a - 8O, tranrition fixd . . . . . . . . . . 12 Effect of Reynold8 number on c l , m X . N (I: 0.15, modal smooth . . . . 23

DISCUSSION

Two Dimensionality of Flow

The spanwire prerrure dirtributionr shown in figure 5 indicate that the flow over the model renained two dimenrional to angles of attach at which muti- mum lift coafficientr occurred. For example, the flow at R = 2,O x 106 and a = 16O figure 5(a) to be two dimnrional in character even though a considerable region of reparstion over the rear upper surface is rhown in the tuft phot+ graphs of figure 6. the other test Reynolds numbers (fig. 11) with the corre8ponding rpanwi8e prer- sure distributions in figure S indicate similar rerultr.

(shown in fig. 1 1 to be near maximum lift conditionr) is Jn'licated in

Correlation of the angles of attack for mxiraw lift for

In addition to the flow over the airfoil remaining tm, dimenrional to angler of attack beyond maximum lift, the tuft photograph8 for (fig. 6) show '%it the flow on the tumel sidewall above the rodel remains attached until an &.lgle of attack of 17O, which is alao beyond daxirum lift.

R - 2.0 x lo6

srtperimntal Resultr

In brief, airfoil characteristics varied systcrlratically over the range of test conditions and behaved in a conventionally accepted manner. ime largert measured value of ~ 1 , ~ ~ ~ was about 2.2 and occurred near .! A 0.1s at a Reynolds number of 12.0 x lo6. (See figs. IS and le.) The vrriatiot: c* lift with angle of attack, at the lower Mach numbers and all Repno1.d~ nunherr. exhibited behavior generally associated with gradual trailing-edge separation in the prestall angle-of-attack range, but the stall itrelf was =re abrupt than would be expected of classical trailing-edge stall. (See figs. 1 1 and 18.) The stall became more gradual at the higher Nach numbers. Drag r e w i n w essentially constant over a lift range whlch extended from near tcro to beyond the design lift coefficient of 0.7 (figs. 1 1 (c). 17, and 18) for a constant Reynolds ..rWer with fixed transition.

Effect6 of fixing transition (fig. 7).- Because of the vircous decaabering effect of the thicker boundaty layer, fixing transition near the leading edge generally rarulted in decreared lift, increased brag, and lers negative pitching -anta. The effects were nore pronounced at the lower lift coefficientr and Reynolds numbers where the greatert extent of laminar flow would exirt for the smooth model. Shown in figure 8(a) are the presrure dirtributioas for a = -2O and R = 2.0 x 106, where the greatest affect of fixing t;rnsition was obrerved. The low Reynolds number and the near set0 prerrure gradient8 wuuld encourage a long run of laminar flow before natural tranrition would OCCC. for the smooth mode!.

7

i

These effects generally decreased with increases in either angle of attack or Reynolds number. The reduced effects w i t h increased angle of attack are associated with the steep adverse pressure gradient of the leading-edge suction peak at high angles which causes natural transition to accur near the leading edge on the upper surface. The reduced effects with increased Reynolds number are due to the thinner boundary layers, the tendency for natural transition to occur nearer the leading edge, and the higher wind-tunnel turbulence level which would also induce earlier transition.

The small differences at the higher angles of attack may not be completely explained because of the inability to determine the relative positions of boundary-layer transition between free and fixed transition on the lower surface. This inability to determine lower-surface boundary-layer transition at high angles of attack is due to the favorable pressure gradient over a large portion of the lower surface and to the fact that the lower-surface transition strip was ahead of the stagnation point at high angles and, therefore, MY not have properly trippet the lower-surface boundary layer.

The general effect of fixing t:ansition and its trend with ihcreasing Reynolds number reverses at high angles of attack for in figdre 7(f) are increased lift, increased pitching moments, and reduced drag with fixed transition. These results are believed to be associated with the elimination, or reduction in size, of a small laminar-separation bubble on the airfoil upper surface by the introduction of roughness near the stagnation point on the lower surface. Tabulated pressure distributions (not presented) indicate that at a = l Z 0 , where the lift curves of figure 7(f) first diverge, the stagnation point has moved to a lower-surface location corresponding to the trip location (0.05~). The presence ,f a laminar-separation bubble can only be ass?tmed since there ere no discernible dismntinuities or plateaus in the pressure coefficients plotted against angle of attack for given chordwise stations shown in figures 9 and 10. Such disconcinuities are usually indioative of the presence of a bubble of separated flow.

R = 18.0 x lo6. Shown

Since natural transition usually occurs near the leading edge of airfoils in actual flight conditions because of the roughness of construction or of insect remains gathered in flight, the remainder of the data discussed in this report (except for the coaapaiison with other airfoils shown in fig. 23) will be with transition fixed at 0.05~ on both upper and lower surfaces.

Stall characteristics.- The shape of the lift curve (fig. ll(a)) in the prestall region is associated with a progressive separation or thickening of the iurbulent boundary layer in the region of the trailing edge where the lift curve slope begins to decrease at the onset of trailing-edge separation. appearance of trailing-edge separation is also manifested by a flattening of the pressure distribution at the rear of the airfoil (fig. 12) and by diver- gence of thc pressure coefficients over and near the trailing edge. fig. 13 1

The

(See

Although +:,e shape of the lift carve in the prestall region is characteristic of thick airfoils (thickcses ratios of approximately 0.12 or greater) which stall as a resuit nf trailing-edge boundary-layer separation, the actual

8

stall is shown in figure 1 1 as a sudden, sharp loss of lift usually associated with leading-edge separation on airfoils of more moderate thickness.

Generally, airfoils which stall as a result of trailing-edge separ-4 have relatively flat pressure distributions over a large region towar" &I& r: of the airfoil at maximum lift, but the peak suction pressures near A , - leadill. edge continue to increase after the stall resulting in a lift curve with rounded maximum. Airfoils which stall as a result of leading-edge stall, b v c r , S ~ Q W a continual increase of the peak pressures near the leading edge up to the angles of attack for maximum lift, followed by an abrupt collapse of the leading-edge pressures.

The pressure distributions of figure 1 2 show that, at angles of attack for maximum lift conditions, trailing-edge separatlm extends over about 20-percent of the chord. After stall, however, the upper-surface separation has spread to almost 90-percent of the chord. Except for the lowest Reynolds number of 2.0 x l o 6 (fig. 1 2 ( a ) ) , there is only a partial collapse of the leading-edge peak which indicates that stall was caused by a precipitous forward movement of the trailing-edge separation.

Effects of Reynolds number (fig. 1 1 ) .- As the Reynolds number increased from the lawest value of 2.0 x l o b to 12.0 x l o 6 at pitching moments became more negative, and drag decreased (minimmi drag SUR- marized in fig. 1 4 ) due to the thinning of the boundary layer. Because the thinner boundary layer is more resistant to separation, the maximum lift coefficient (fig. 15) increased from 1 .84 to 2.23 over this range of Reynolds numbers. As Reynolds number was further increased to 18.0 x l o 6 the maximum lift coefficient decreased to 2.15. In addition, higher values of drag and less negative pitching moments at high lift coefficients for compared with R = 12.0 x 1 O 6 the variation of the aerodynamic characteristics with Reynolds number are due to a greater amount of trailing-edge separation for R = 18.0 x l o 6 compared to R = 12.0 x 106. (See fig. 1 2 . )

M = 0.15, lift increased,

R = 18.0 x l o 6 are indicated in figure 1 1 . These reversals in the trend of

Alt:,;:-gh the experimental data is not detailed enough to explain this phenomenon, it is possible that increasing Reynolds number frm, i 2.0 x 106 to 18.0 x l o 6 moved the reattachment location of the laminar ieparation bubble which altered the initial thickness of the reattached turbulent boundary layer in such a manner as to promote increased trailing-edge separation.

Effects of Mach number (fig. 18) .- The effects of Mach number on the maximum lift coefficient at a Reynolds number of 6 .0 x l o 6 , shown in figure 18, are sumnarized in figure 19 . There is a decrease in the maximum lift coefficient and in the angle ot attack for maximum lift with increasing Mach number, which becomes appreciable above M *I 0.20.

In reference 1 3 , the decrease in maximum lift with Mach number is related to the appearance of local supersonic flaw near the leading edge. In figure 20, it is shown that supersonic flow was encounted near stall for Mach numbers of 0.28 and 0.32. Although the zone of supersonic flaw may be small, the resulting recompression of the flow thickens the boundary layer and increases the tendency for turbulent sr+ration at the rear of the airfoil.

9

The stall characteristics are much less ahrupt at the higher Mach numbers (fig. 18) due to a more gradual forward progression of the trailing-edge sapa- ration. (Compare fig. 21 with fig. 12(c).)

Comparison of Maximum Lift Characteristics With Other Airfoils

Maximum lift characteristics of the SC(2)-0714 are compared with those of the NASA low- and medium-speed airfoils (refs. 14 and 15) and NACA airfoils (ref. 16) in figure 23.

This comparison shows the maximum lift coefficients for the SC(2)-0714 airfoil to be about 0.1 greater than for the NASA law-speed airfoils, almost 0.2 greater than for '.\e NASA medium-speed airfoils, and from 0.5 to 0.6 greater than for the NACA airfoils. Because of the greater aft camber, the pitching- maneirt coefficient of the supercritical airfoil is about -0.13 at zero lift, compared to about -0.10 for the law-speed airfoils, about -0.07 for the medium- speed airfoils, and near zero for the NACA 23015.

Higher maximum lift could, of course, be achieved with conventional airfoils through increased camber. Maximum lift coefficients approaching that of the supercritical airfoil are reported for a NACA 6716 four-digit airfoil in reference 17, for example, but those higher maximum lift coefficients are accom- panied by pitching-moment coefficients of approximately -0.20 and separation over the trailing edge at law angles of attack.

CONCLUDING REMARKS

Wind-tunnel tests have been conducted to determine the low-sped two- dimensional aerodynamic characteristics of a 14-percent-thick NASA supercritical airfoil with a design lift coefficient of 0.7. This report documents the experimental results of these tests. Free-stream Mach number ranged from 0.10 to 0.32, and the chord Reynolds number varied from 2.0 x l o 6 to 18.0 x l o 6 . Analysis of the results indicate the following general conclusions:

1. Maximum lift coefficient increased with Reynolds number from 2.0 x l o 6 to 12.0 x l o 6 (M = 0.15) and attained a value of about 2.2. in maximum lift was observed with further increase in Reynolds number to 18.0 x l o6 .

A small decrease

2. Maximum lift coefficient decreased with increased Mach number (R = 6.0 x l o 6 ) , the decrease became significant above M - 0.20.

3. The application of a transition strip near the leading edge resulted in only small effects in maximum lift since natural transition occurred near the upper-surface leading edge due to the adverse pressure gradient associated with the suction peak at high angles of attack.

4. The shape of the prestall lift curve was characteristic of a gradual trailing-edge stall, but the stall itself was abrupt and attributed to a precipitous forward movement of the trailing-edge separation point.

10

5. Drag remained essentially constant over a lift range which extended fran near zero to beyond the design lift coefficient for a constant Reynolds number with transition fixed.

Langley Research Center National Aeronautics and Space Administration Hampton, VA 23665 November 26, 1980

11

1. Harris, Charles D. : Aerodynamic Characteristics of the 1 0-Percant-Thick NASA Supercritfcal Airfoil 33 Designed for a Normal-Force Coefficient of 0.7. NASA 'M X-72711, 1975.

2. Harris, Charles D.: Aerodynamic Characteristics of a 14-Percent-Thick NASA Supercritical Airfoil Designed for a Normal-Force Coefficient of 0.7. NASA !lW X-72712, 1975.

3. Baals, Donald D.; and nOUrbess, Mary J.: -erica1 Evaluation of the Wake-Survey Equations for Subsonic Flow Including the Effect of Energy Aaition. NACA WR L-5, 1945. (Formerly NACA ARR LSH27.1

4. Whitamb, Richard T.: Review of N S A Supetciitical Airfoils. ICAS Paper No. 74-10, Aug. 1974.

5. Harris, Charles D.: Wind-Tunnel Investigation of Effects of Trailing-Edge NASA BI 1-2336, 197l. GeoPetry on a NASA Supercritical Airfoil Section.

6. Harris, Charles 0.: Aerodynamic Characteristics of an Imprc red 10-Percent- Thick NASA Supercriticrl Airfoil. NASA 'Iw. X-2978, 1974.

7. Harris, Charles D.: Transonic Aerodynamic Characteristics of the 1 0-Percent-Thick NASA Supercritical Airfoil 31. NASA RI X-3203, 1975.

8. Bauer, Frances; Garabedian, Paul; Korn, David; and Jameson, Antony: Supercritical Wing Sections 11. and Mathematical Systems, Springer-Verlag, 1975.

Volume 108 of Lecture Notes in Economics

9. Von Doenhoff, Plbert E.; and Abbott, Frank T., Jr.: The Langley Two- Dimensional taw-Turbulence Pressure Tunnel. NACA TN 1283, 1947.

10. Pope, Alan: and Harper, John J.: Low-Speed Wind Tunnel Testing. John Wiley 6 Sons, Inc., 1966.

11. Pankhurst, R. C.; and Holder, D. W.: Wind-Tunnel Technique. Sir Isaac Pitlaan &I Sons, Ltd. (London), 1965.

12. Braslow, Albert L.; and Knox, Eugene C.: Simplified Method for Determina- tion of Critical Height of Distributed Roughness Particles for Boundary- Layer Transition at Mach Numbers From 0 to 5. NACA TN 4363, 1958.

13. Wootton, L. R.: The Effect of Compressibility on the Maximum Lift Coef- ficient of Aerofoils at SubsonIs Airspeeds. J. Roy. Aeronaut. Soc., v01. 71, July 1967, m. 476-486.

14. McGhee, Robert J.; and Beasley, William D.: Low-Speed Aerodynamic Charac- teristics of a 17-Percent-Thick Airfoil Section Designed for General Avi- ation Applications. NASA TN D-7428, 1973.

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I S . W h e e , Robert J.; Beasley, William D.; and Uhitcorb, Richard T.: N M A Low- and Medium-Speed Airfoil Developent. NASA Rl-78709, 1979.

16. AbbOtt, Ira H.) and Von Doenhoff, Albert E.: Theory of W i n g Sections. Dover Publ., Inc., c.1959.

17. Binghaa, Gene J.; and Noonan, Kevin W.: Low-Speed Aerodynaaic Characteris- tics of NACA 6716 and NACA 4416 Airfoils With 35-Ptrcent-Chord Single- Slotted Fla?s. NASA TM X-2623, 1974.

x/c

0.0000 .oo; .005 .010 .020 .030 .040 .OS0 .060 .070 .080 .090 . l o o .110 .120 .130 .140 .150 .160 .170 .180 .190 200

.21 0

.220 ,230 --

TABLE 1.- SECTION ODORDINATES POR SC(2)-0714

E = 61 .O cm (24 in,) ; leading-edge radius = 0.0304

( z / c ) u

0.0000 .0108 . 01 67 .0225 .0297 .0346 .0383 .0414 .0440 .0463 .0484 .OS02 .OS1 9 .OS35 .OS49 .0562 .OS74 .OS85 .OS96 .0606 .0615 .0624 .0632 .0640 .0647 .0653

0.0000 -. 01 08 -.0165 -. 0223 -. 0295 -. 0343 -. 0381 -. 041 1 -. 0438 -. O461 -. 0481

-. os1 7 -. 0533 -. 0547 -. 0561 -. 0574 -. 0585 -. 0596 -. 0606 -. 061 6 -. 0625 -. 0633 -. 0641 -. 0648 -. 0655

-. 0500

0.240 .250 .260 .270 .280 .290 .300 .310 .320 .330 .340 .350 .360 .370 .380 .390 .400 .410 .420 .430 .440 .450 .460 .470 .480 .190

0,0659 .0665 , 0670 -0675 .0679 .0683 .0686 .0689 .0692 .0694 .0696 .0698 .0693 .0700 .0700 .0700 .0700 .0699 .0698 .0697 .0696 .0694 .0692 .0689 .0686 .0683

-0.0661 -. 0667 -. 0672 - . 0677 -. 0681 -. 0685 -. 0688 -. 0691 -. 0693 -. 0695 -. 0696 -. 0697 -. 0697 -. 0697 -. 0696 -. 0695 -. 0693 -. 0691 -. 0689 -. 0686 -. 0682 -. 0678 -. 0673 -.'0667 -. 0661 -. 0654

14

ORIGINAL PAGE Is OF POOR QUALITY

~-

0.500 .510 .520 .530 .540 .550 .560 .570 .580 .590 .600 .610 .620 .630 .640 .650 .660 .670 .680 .690 .700 .710

.730

.740

.750

.720

( Z / C ) u

0.0680 . 0676 . 0672 -0668 -0663 .0658 .0652 .0646 .0640 .0634 .0627 .0620 .0613 .0605 .OS96 .OS87 .OS78 .OS68 .OS58 .OS47 .OS36 .OS24 .OS1 2 .0499 .OA86 .0472

TABLE I.- Concluded

-0.0646 -. 0637 -.0627 -.OS16 -. 0604 -. 0591 -. 0577 -. 0562 -. 0546 -.OS29 -. 051 1 -. 0493 -. 0474 -. 0454 -.0434 -. 041 3 -. 0392

-. 0349 -. 0327 -. 0305 -. 0283 -. 0261 -. 0239 -. 021 7 -. 01 95

- . o m

X / C

0.760 .770 .780 -790 .800 97 . L L d

.830

.840

.850

.860

.870 -880 .a90 .goo .910 .920 .930 .940 .950 -960 .970 .980 .990

1 .ooo

0.0457 -0442 .0426 -0409 .0392 .0374 .0356 .0337 .0317 -0297 .0276 .0255 .023’ .02 .0186 .0162 .0137 .0111 . 0084 .0057 .0029 .oooo -. 0030 -. 0062 -. 0095

-0.01 73 -.0152 -. 01 32 -.0113 -. 0095 -. 0079 -. 0064 -. 0050 -. 0038 -. 0028 . 0020 -. 001 4 -.0010 -. 0008 -. 0008 -.0011 -. 001 6 -. 0024 -. 0035 -. 0049 -. 0066 -. 0086 -. 01 09 -. 01 36 -. 01 65

i L

5

Figure 1.- Profile, thickness distribution, and camber line of 14-percent-thick supercritical airfoil (SC(2)-0714).

1 6

c A

Figure 2.- Typical airfoil mounted in wind tunnel. All dimensions are in terms of airfoil chord; c - 61 . O cm ( 2 4 . 0 i n . ) .

1 7

Stat< - pressure

3 .042c

- - . 0 4 2 c I

t-----.2k+

Figure 3.- Wake survey rake. All dimensions in terms of airfoil chord. c = 61.01 cm (24.02 in.).

1 8

ORIGINAL PAGE is OF POOR QUALIV

.04

.03

02 Cd

01

0

I6 :.

c.

I 2 '--

0 * -

I

-4 t

-0 -I 2 - 8 -.4 0 . 4 .8 I 2 I 6 2.0 2.4 2.8

CI

Figure 4.- Canparison of corrected and uncorrected aerodynamic characteristics; M - 0.15, R = 18.0 x lo6 , transition fixed.

19

- 24.

- 20.

-16.

CP -I 2.

- 0.

- 4.

0

- 4.

- 3.

-2. CP

- I .

0 - 1.6

- 1.2

- .8 CP

- .4

0 -1.00 -.80 -.60 -90 -.20 0 .20 40 .60 .80 1.00

y / c

Figure 5.- Spanwise pressure distribution on upper surface: M = 0.15, transition f ixed .

20

- 24

- 20.

- 16.

cP - 12,

- 0.

- 4.

0

- 4.

-3

- 2, cP

- I .

0 -1.6

cP -1.2

- .0

4

0 - -1.00 180 -.60 4 0 720 0 .20 90 .60 .80 1.00

y/ c

(b) R = 4.0 X l o 6 .

Figure 5.- Continued. 21

i 1 Tunnei watt I -20. i

16. 1

~

i I cP I

12. ;

8 .1 t

1

4. *

n I -4. . ' I /

t - +

! I

-3. t t

c;P - 2. 1 .i I

- 1 . I 0

- .8 CP

t

f I i

b 1

i

I

?

t i l l

I / C = 0.75 . . . 1

-I.W -.80 ,60 -90 720 0 .20 .40 .60 .80 100

Y / c

( c ) R = 6.0 x l o 6 .

Figure 5.- Continued. 22

ORlGlNAL PAGE ?S OF POOR Q U A L I n

- 24.

70.

- 16.

-12.

cP

cP

- 8.

- 4.

0 - 4.

- 3.

- 2,

- 1.

0

- 1.6

-1.2

-.8 cP

- 4

0

. - . * ,

I I I I I I

+ *

* *

. * 4 1

1 .

-1

1 I 4

-1.00 -.80 -.60 -90 7.20 0 .20 .40 .60 .8U 1.00

v ' c

(d) R = 9.0 x l o 6 .

Figure 5.- Continued 23

ORIGINAL PAGE 1s OF POOR QUALITY

- 24.

- 20.

CP - 16.

- 12.

- 0.

- 4.

0

- 4.

- 3.

cP - 2.

- 1.

0

-1.6

-1.2

cP - .8

- .4

0 -1.00 -.80 -.60 -.40 72C 0 .20 90 .60 .80 1.00

y/c

(e) R = 12.0 x 106.

Figure 5.- Continued.

24

ORIGINAL PAGE IS OF POOR QUMW

cP

- 24.

- 20.

- I 6.

- 12.

- 8

- 4.

0 - 4.

- 3.

- 2. cP

- I .

0 -1.6 1 I

! -.4 j

780 760 -40 y20 0 .20 90 .60 .80 1.00

y/

(f) R = 18.0 x lo6.

Figure 5.- Cor.cluded. 25

ORIGINAL PAGE !$ OF POOR QUALITY

. rg 0 c

- x

(Y

I

a n Q -

I I 1 - I

3"

29

ORIGINAL PAGE 1s -OF POOR QUALW

30

ORIGINAL PAtiE 'rs OF POOR QUALITY

. . . * . . , . . . . . . .

31

ORIGINAL PAGE IS OF POOR QUALIP

. . . . 1 - I I C U r u N - - - - C

1

3"

32

ORIGINAL PACE IS OF MOR QUALtTY

I I I I I I I I I I

E u

33

0°/

r l n 0

0 0 0 ~ 0 0 0

LI

( D F r a m m . . . . e . .

34

ORIGINAL PAGE IS OF POOR QUALI'IY

ORIGiNAL PAGE IS OF POOR QUALITY

cu cu

0 cu

a0 - u) - f I

cu - 0 - m

u3

f

cu

0

nJ I

f I

u3 I

(G I

0

I &

I I I I I I I I I I /

E u

36

.

! r-

37

. . . . c u ( u c u - - - - - I I I -

I

J"

38

ORIGINAL PAGE lg OF Poan QUALITY

Ai

39

ORIGINAL PAM E OF POOR QUALITY

. . . . . . . . , s a . . I . .

W 0 c

X

0

cy

I

p:

. c

CI

Q) Y

OREINAL PAGE fS OF POOR QUALITY

42

. . . . . . . . . . .

4.3

OWilNAL PAOE is OF POOR QUALm

n a

'. r-

44

ORlGlNAL PAGE Is OF POOR QUALITY

E u

45

ORIGINAL PASL k OF POOR QUALln

.

46

ORIGINAL PAGE ES OF POOR QUALITY

. R I

I

a . w 0

X

0

(Y

c

II

0:

A

cp Y

-A w win 0 - u o 0 0 ) u

.

. $ I.

(D

al

47

ORIGINAL PAGE 13 OF POOR QUALITY

48

- 13.

- 12.

- I 1.

-I 0.

- 9.

-0.

cp -7.

-6.

- 5.

-4.

-36

- 2.

- 1 .

ORIGINAL PAS',: i . i OF POOR QUALITY

0

Figure 9 . - Variaticn o f chordwise s t a t i o n s

4 8 12 16 20 24

a , deg

(a) R - 2.0 x 106.

pressure c o e f f i c i e n t on upper surface a t part icular with angle of attack; M = 0 . 1 5 . model sniooth.

49

ORIGINAL PM'dE OF POOR QUALIW

- 14.

- 13.

-12.

-I 1 .

- I 0.

- 9.

cp -0.

- 7.

- 6.

- 5.

- 4.

- 3.

- 2. 0 4 8 12 16 20 24

0 , deg (b) R = 4 . G x l o 6 .


50

ORiGiNAL PAGE OS OF POOR QUALITY

- 13.

- 12.

-I 1 .

-10.

-9.

p -8. C

- 7.

- 6.

- 5.

- 4.

-3.

- 2. 0 4 €3 12 16 20 24

a, deg ( c ) R = 6.0 x lo6.

Figcre 9.- Continued.

51

+

-5. ’

--4. ’

--3.

? -L.

0

ORIGINAL FAGE 1s OF POOR QlJn-.LlTY

- 14.

- 1:3.

-12.

- I I.

-10.

- 9.

cp -8.

-7.

-6.

4 8 12 I6 20 24

a , de9

(d) R = 9 .0 x l o 6 .

F i g u r e 9 . - Continued.

52

CRlGiNAL FAG: IS OF POOR QUALiW

- 14.

- 13.

-i2.

-I 1.

-I 0.

-9.

-8.

- 7.

-6.

-5 .

-4.

-3.

- 2. 0 4 8 12 16 310 24

Q, deg R = 12.0 x IO'.


(e)

53

-13. ;

I -12. i

/ - I 1 . I

-10. I

-9. i i

I

I

!

!

I

-0. ' j

0 4

j i 1 -

i

I

8 I

12 16 20 24

( f ) R - 18.0 y l o 6 .

Figure 9.- Concluded.

54

- 13.

-12,

- I I.

- I 0.

- 9.

cP - 8.

- 7.

-6.

- 5.

-4.

- 3.

(a) R = 2.0 x 106.

Figure 10.- Variation of pressure coefficient on upper surface at ?articular chordwise stations with angle of attack; M = 0.15, transition fixed.

55

- 13.

- 12.

-I 1.

-10.

- 9.

-8.

- 7.

-6.

- 5.

- 4.

- 3. 0 4 8 12 16 20 24

a , deg (b) R = 4.0 x l o 6 .


56

-14- ~

i 1

- 13.

-12.

-I 1 .

-I 0.

- 9.

- 8-

I - 4

,

,

-7. '

-6.

-5. .

-4. '

0

- t

0 0 0 4 h

t

t

- T

- *

t

+

24


-14.

- 13.

-12.

-I I.

-10.

- 9.

- 8.

- 7.

-6.

-5.

- 4.

a , deg (d) R = 9.0 x lo6 .


58

ORIGINAL PAGE OF POOR QUALITY

-14.

- 13.

-12.

-I 1.

-10.

-9.

-8.

-7.

- 5.

- 4. 0 4 8 12 16 20 24

0 , WI ( e ) R = 12.0 x I O 6 .

Figure 10 .- Continued. 59

- 14.

- 13.

-12.

-I 1 .

-I 0.

- 9.

- 8.

- 7.

-6.

- 5.

- 4. 0 4 8 12 16 20 24

a, deg

( f ) R = 18.0 x l o 6 .


60

a’ a a 1

2 3

N

u

a Y

.. \u w I.

z c c

1 01 ‘rl 14

61

ORIGINAL PAGE \ti OF POOR QUALln

62

ORIG:NP*i PAGE IS OF POOR QUALIW

I I I I I I I I I I

E u

a' m m 1 Lc

d

A

P,

63

E u

64

H

u

m 1 m $4

P a u

0

65

ORIGINAL F.hGE id OF POOR QUALlTV

66

ORIGINAL PAGE 13 OF POOR QUALM

w \ H

I I

.. .

67

0001 I

I I

68

mlNAL PAGE IS OF POOR QUALITY

. W 0 - X

0

cy

I

p5

n Q -

69

I

- - I I I I I I I I I I I 1 I

70

PAGE CS OF POOR QUALtlV

e W 0

2.

0

0

I a

P

. W 91 1 E 4 u

! (Y

91

c

71

ORlGINAL PAC35 iS OF POOR QUALITY

72

ORIGINAL PAGZ ;;i OF POOR QUALITY

73

-e-

I l - - - I 1 I

I I I I I

74

0"

16

t"" I " ' "" " I ' I " '

-

77

ORIGINAL PAGE 55 OF POOR QUALITY

. w 0

X

r

9 (Y c

n p5

n 0, Y

. a c 4 U

u

8

0"

79

ORIGINAL PAQE 1s OF POOR QUAL^

80

ORIGINAL PAGE hS OF POOR QUALITY

W '3 r

X

0

QD

II

p:

F

81

.

A

tt!

82

ORIGINAL PkLS !3 OF POOR QUALtTv

-. 3

-.2

-.I

0

. I

-.3

-* 2

CP -.I

0

.t

r

t P

-4 0 4 8 12 16 20

a, deg !a) Base of model.

Figure 13.- 3ase and upper-surface pressure coefficients; €4 = 0.1 5 , transition fixed ,

83

CP

-. 3

1 .- -

- . I

0

R

0 2 x IO6 0 4 O b 5 9 b 12

I I 1 I

-. 3

-. 2

-. I

0

P P

I I I A 1 I .I - 4 0 4 8 12 16 2 0

a , dec,

(b) Upper surface; x,/c = 0.99.

Figure 13. - Continued.

84

-. 3

-. 2 cP

-.I

0 2 x IO6 o a 0 6 A 9 b 12 h 18

. I I 1 I

cP

- 4 0 4 8 I2 16 20

a , deg ( c ) Upper surface; x/c = 0.95.


85

ORlGiNAL ?AGE E$ OF POOR QUALiW

L CT (s [r

0 0

P I

I I

I I

t .- E si 0

86

0 9

89

ORIGINAL PAGE IS OF POOR QUALITY

E u

'. W

al Lc 1 P 4 b4

c

91

I. m c

92

‘1. max

2.4

2,2

2.0

I .8

I .6

I .4 0 ,I *2 .3 *4

Mach number

\ I I u

Figure 19.- Variation of maximum l i f t coefficient with Mach number; R = 6.0 x l o6 , transition fixed.

94

ORIGINAL PAW I$ POOR QUALITY

95

'. 0 rv

96

*ORIGINAL PACE IS OF POW @:JALllY

m

w o O \ d A I -

d

0

97

ORIGINAL PAGE Is OF POOR QUALITY

CL 0

98

ORKiINAL PAGE ls OF POOR QUALlTV

c1, 10

I

100

airfoil designed lift of - nasa technical memorandum 81912 low-speed aerodynamic characteristics of...

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