m —— ___QJ /

umcCf

~~

NATIONALADVISORYCOMMITTEEFOR AERONAUTICS

TECHNICAL NOTE 3036

THE FLOW ABOUT A SECTION OF A FINITE-

ASPECT-RATIO NACA 0018 AJRFOIL

ON A TRANSONIC BUMP

By Jack A. Mellenthin

Ames Aeronautical LaboratoryMoffett Field, Cal.if.

Washington

October 1953

-- .< . --- - --- . . ...-. —.. —.. ..- .. .. . . .- . . ..- ------------- -. -.=... -------- . ...

[w

TECHLIBRARYKAFB,NM

.

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

TECHNICALNOTE3036

THE FLOW ABOUT A SECTION OF A FINITE-

ASPECT-RATIO NACA 0015 AIRFOIL

ON A TRANSONIC BUMP

By Jack A. MeZ1.enthin

suMMARY

Pressure distributions on a semispan rectangular wing model havingan NACA 0015 airfoil section and a moderate aspect ratio were measuredat one spanwise station in tests made on a transonic bump in the Ames16-foot high-speed wind tunnel. The free-stream Mach number range wasfrom 0.4 to 1.o6. The results showed that at a fixed angle of attack aregion developed over the airfoil wherein the Mach number at each pointremained essentially constant as the free-stream Mach number wasincreased above the critical. This region covered essentially thewhole chord of the airfoil at free-stream

INTRODUCTION

Mach numbers near unity.

An understantig of the adverse aerodynamic characteristicsexperienced by airfoils in the transonic region is dependent upon acomprehension of the flow changes which cause such characteristics.Toward this end, considerable work has been done in the study of thenature of transonic flow. Of special.interest are the investigationsreported in references 1 to 5. In these references it has been recog-nized that at transonic Mach numbers a region exists on the surface ofan airfoil wherein, at a fixed angle of attack, the local Mach numberremains essentially constant with increasing free-stream Mach number.The present investigation supplies additional information on the chwac-teristics of this cunstant local Mach number region through the rangeof transonic free-stresm Mach numbers up to 1.o6 where data have here-tofore been rather meager.

The results of the present investigation do not correspond strictlyto two-dimensional flow since a finite-span model having a moderateaspect ratio was used. Chordwise distribution of surface pressures

— -- — ——-.———.—— —— —. —— .——- -

2 NACA TN

were measured at only one station. To reduce the effect of aspecton the data, this orifice station was located far inboard from themodel tip.

a

c

cdP

cl

%

M

Ml

P

Pm

P~

P

~

R

v

‘zx

Y

a

local.speed of sound,

airfoil chord, ft

section pressure drag

NOTATION

ft/sec

coefficient, section pressure dragqc

section lift coefficient, section liftqc

section pitching-moment coefficient, about the qusrter-chord

point, section pitching moment

qcz

vaverage free-stream Mach number, _

VIlocal Mach number. —

“a

pressure

pressure

P~-Pcoefficient,

~

coefficient at which

a

local sonic velocity occurs

local static pressure, lb/sq ft

free-stream static pressure, lb/sq ft

free-stream dynamic pressure, lb/sq ft

Reynolds number

free-stream airspeed, ft/sec

local airspeed, ft/sec

chordtise distance belrhd the leading edge of the airfoil, ft

airfoil ordinate, ft

geometrical angle of attack, deg

3036

ratio

“

NAcA TN 3036 3

APPARATUS AND TESTS

‘.

A semispan wing model mounted on a transonic bump in the Ames 16-foot wind tunnel was used in the present investigation (figs. 1 and 2).This model had the NACA 0017 section, a 6-inch chord, a semispan ofapproximately 12 inches, an aspect ratio of about 4, and was the sanemdel used in the two-dimensional investigations reported in references1 and2. For the present tests, the mcdel was modified by adding a tipfairing and by putting orifices on both the upper and lower sumfaces at0.25-percent and 1.25-percent chord, so that additional data could beobtained near the leading edge. (See table 1.) To decrease the effectsof the bump boundary layer on the flow over the wing, a fence wasinstalled on the airfoil 3/4 inch from the bump surface. A woodenfairing was included between this fence and the bump surface to housethe mciielsupport bracket. The mciieland wooden fairing (figs. 1 and2) were mounted on a turntable flush with the bump contour, and theangle of attack was changed by rotating this turntable. The effect ofthe wooden fairing on the data is believed to be small.

The row of pressure orifices on the mciielwas 2 inches from thefence. Pressures were measured on the upper and lower surfaces of theairfoil using a mercury-in-glass, multiple manometer.

During this investigation, only pressures on the upper and lowersurfaces of the model were measured. The force and moment data wereobtained by mechanical integration of the pressure-distribution plots.The pressures were measured through an angle-of-attack range from -4°to +15°, and the Mach number range was from 0.4 to 1.o6, while thecorrespondingReynolds number range (fig. 3) was from appro~tely1.3 million to 2.1 million. The dynamic pressure in the test region washeld within 0.5 percent of its predetermined values, and the angle-of-attack measuring device was accurate within W.l”.

Mach number contours over the bump @ the absence of the model areshown in figure 4. The heavy dashed lines in this figure indicate themodel location. The reference free-stream Mach nwbers shown are aver-ages of the local Mach numbers over the orifice station.

RESULTS AND DISCUSSION

Pressure Measurements

Representative chordwise pressure distributions for angles of attackof 0.5°, 4°, and 80 are shown in figures 5 to 7. (The test resultsshowed that the section lift was equal to zero at an angle of attack ofabout 0.50.) Values of the critical pressure coefficient are shown in

-— — —.

4 NACA TN 3036

the figures for free-stream Mach numbers below 1.00 where mixed flowwas present.

The changes which took place in the distributions of pressure asthe free-stream Mach number increased through unity are quite apparentin these figures. At singlesof attack of 4° and 8° it is evident, asindicated by the locations of abrupt increases in the pressure coef-ficient at low supercriticalMach numbers, that the shock wave on thelower surface moved downstream rapidly as the free-stream Mach numberwas increased, while the shock wave on the upper surface moved down-stream much less rapidly. Consequently, over the rearward portion ofthe airfoil, greater negative pressures occurred on the lower surfacethan on the upper surface. This sequence of events produced a reduced,or in some cases a negative, lift on the airfoil, as will be discussedlater in this report. As the free-stream Mach number was increasedfurther, the shock wave on each surface moved downstream to the trailingedge, eliminating the region of negative lift. The accompanying recoveryof the lift will.also be discussed later.

Region of Nearly Constant Local Mach Number

Local Mach numbers over the airfoil surface for angles of attackof 0.50, 40, and 80 have been determined from the pressure coefficientsof the present investigation, using isentropic relations, and sre pre-sented at selected chordwise stations in figures 8 to 10 as a function offree-stream Mach number. As the test Mach number was increased abovethe critical, a small region developed over the forward part of the air-

.

foil at an angle of attack of 0.5° (fig. 8) wherein the Mach number ateach point remained essentially constant with increasing free-stresmMach number. For example, at 5-percent chord the local Mach numberincreased o~y about 0.02 while the free-stream Mach number was increasedabout 0.20 (from 0.85 to 1.05). At angles of attack of 4° and 8°, theregion of constant local Mach number formed first at low supercriticalMach numbers at about the ,~-percent-chordstation on the upper surface,and, as the Mach number was increased further, this region spread bothforward and rearward on the upper surface, reaching the leading edge,and becoming established on the forwaxd portion of the lower surface ata free-stream Mach number of about 0.9. At all three angles of attack(figs. 8 to 10), the region essentially covered the airfoil from the ,leading edge to the trailing edge at a free-stream Mach number of unity.

It is also observed in figures 8 to 10 that, as the free-streamMach number was increased, the local Mach number for rearward chordwisestations increased rapidly before the relatively constant region wasformed. On the upper surface, at angles of attack of 4° smd 8°, thisincrease was more gradual thau at an angle of attack of 0.5° (figs. 9(a)and 10(a)).

NACA TN 3036 5

Force and Moment Chsxacteristics

The force and moment characteristics,which were obtained bymechanical integration of the pressure-distributionplots, are presentedin figure Il. From this figure, it may be seen that the airfoil exhib-ited undesirable variations of section lift, pressure-drag, and pitching-moment coefficients at the high subsonic Mach numbers.

The adverse variations of lift and pitching moment resulted fromthe previously discussed rearrangement of pressures on the upper andlower airfoil surfaces at high subsonic Mach numbers. Most of the liftreduction took place between free-stresm Mach numbers of 0.75 and 0.85and was so severe at a Mach number of 0.88 that the section lift coef-ficient actually was negative at a positive angle of attack of 4°. Asa Mach number of unity was approached, the lift recovered somewhat, asmay be seen in figure Il. This recovery, which resulted from the pre-viously mentioned behavior of the pressures, occurred principallybetween free-stresm Mach numbers of 0.90 and 0.96. The unusualJ-ylargevariations of the pitching-moment coefficients at high subsonic Machnumbers, shown in figure lJ near O angle of attack, indicate corres-pondingly }arge variations in the chordwise center-of-pressure locationin this Mach number range. The mount of pressure drag contributed bythe parts of the airfoil forward and reward of the point of ma@mnnthickness has been determined at an angle of attack of 0.5° and is pre-sented in figure E together with the total pressure drag for comparison.The initial total pressure drag rise with increasing transonic free-stream Mach number resulted principally from decreasing pressure coef-ficients over the rear part of the airfoil. However, at Mach numbersnear 1 an increasing part of the total pressure drag was due to theincreasing pressures over the forward part of the airfoil.

CONCLUDING~

As the free-stream Mach number was increased above the critical.,aregion of nearly constant local Mach number formed nesr the leading edgeof the airfoil at zero lift. At higher angles of attack, this regionformed first at about the 5-percent-chord station on the upper surfaceat low supercriticalMach numbers. The region spread both forward andrearward as the free-stresm Mach number was increased, reaching theleading edge and becoming established on the forward portion of thelower surface at a free-stream Mach number of about 0.90. When theMach number of the free stream was increased further, the region ofnearly constant Mach number expanded re?mnard for each angle of attack

6 NACA TN 3036

until essentially the entire chord of the airfoil was included at afree-stresmMach number of unity.

Ames Aeronautical LaboratoryNational Advisory Committee for Aeronautics

Moffett Field, Calif., Aug. 27, 1953

REFERENCES

1. Graham, Donald J., Nitzberg, Gerald E., and Olson, Robert N.:A Systematic Investigation of Pressure Distributions at HighSpeeds Over Five Representative NACA Low-Drag and ConventionalAirfoil Sections. NACA Rep. 832, 1945.

2. Nitzberg, Gerald E., and CrandaJl, Stewart: A Study of Flow ChangesAssociated with Airfoil Section Drag Rise at Supercritical Speeds.NACA ~ 1813, 1949.

3. Bryson, Arthur Earl, Jr.: An Experimental Investigation of llhmnsonicFlow Past Two-DimensionalWedge and Circular-Arc Sections Usinga Mach-Zehnder Interferometer. NACATN 2560, 1951.

4. Tsien, Hsue-Shen, and Fejer, Andrej: A Method for Predicting the!Iknsonic Flow Over Airfoils and Similar Bodies from Data Obtainedat Small Mach Numbers. WIT, 1944.

59 Gul.lstrsnd,Tore R.: A Theoretical Discussion of Some Propertiesof Transonic Flow over Two-Dimensional Symmetrical Aerofoils atZero Lift with a Simple Method to Estimate the Flow Properties.KTHAero TN 25, 1952, Royal Inst. of Tech., Stockholm, Sweden.

NACA

“

.

TN3036

TABLEI.- MODEL COORDINATES AND PRESSURE-ORIFICE STATIONS

[Stations and ordinates in percent of airfoil chord]

‘ACA0015 coordinate

Station

o1.2502.5005.0007.m10.0001~.00020.00025.00030.00040.00050.00060.00070.o(x)80.00090.Ocm95.000100.000100.000

Ordinate

o2.3673.2684.4435.2505.8536.6817.1727.4277.5027.2546.6185.7044.5803.2791.8101.008(.158)o

Pressure-orifice stations

Jpper surface

o.250

1.2502.5005.00010.00015.00020.00025.00030.00035.00040.00045.000W. 00055.00060.00070.00080.00087.~0

Lower surface

o.250

1.2502.5005.000

10.00015.00020.00025.00030.00035.00040.00045.00050.00055.0006Q.00070.00080.00090.000

7

——. .—. .-—-.—-— -—_ .- ... —.——

8 N.ACATN 3036

NACA TN 3036 9

.

(Axis of

Wooden fuiring\

Orifice stution

rotution)

Fence

Hun vie w

/

Tip foiring,

33T +

(c)k\\

Front view

___ -—- ---< 1

‘---—_ ----

Nofe: All dimensionsSide view v in inches.

Figure 1.- Three-view drawing of the model on the transonic bump of theAmes 16-foot wind tunnel.

.—— ——.—.——-..——— __ ———_ .—> 1::

10 NACA TN 3036

l?iglm

(a) Three-quarter front view.

(,b)Three-quarter rear view.

2 2.- The model on the transonic bump of the Ames 16-footwind tunnel.

.

U.

.

2.0

1.9

1.8

/.7

/.6

/.5

/.4

/.3.4 .5 .6 .7

Moth number, M.8 .9 /.0

Figure 3.- Variation of Reynolds number with Mach number during the16-foot wind-tunnel test of the NACA 0017 airfoil model.

/./

.— — ———_ —— .—. _—— -. — —-——

12 NACA ~ 3036

20

/0

-o

20

/0

o“

70 80 90 /00 //0 /20Bump sfution, inches =-S=

Figure 4.- Typical Mach number contours over the Ames 16-foot tunneltransonic bump.

.

.

.

NACA TN3036 13

-L6

M

-/.4 0 0.40

-L?

-LQ

-.4 T ,..

-.2

.2

.6

12

---— - Lower surfuce1.4

0 10 20 30 40 50 60 70 80 90 /00

Percent chotd

(a) M, O.hO, 0.60, and O.70

Figure 5.- Effect of Mach number on the chordwise distribution ofpressure coefficient, a = 0.5° (approximately zero lift).

...—————-—.— - ——

NACA TN 303614-/.6

-/.4

-L2

-Lo

-.8

-.6

.4

.6

.8

Lo

L2 <

[4

M

0 0.75

A .85

--- .

=Pcr fo r M, o.75—

-yPcr for IV, O.8P=

1

Upper surface––-—– Lower surface

o 10 20 30 40 50 GO 70 BO 90 /00

Percent chord

(b) M, 0.75,0.80,0.82,and 0.85

Figure 5.-Continued.

NACA TN

-/.6

-/.4

-42

-Lo

-.8

-.6

k -.4%%c

0

.2

.4

.6

.8

/.0

L2

/.4

M

o 0.90

El .96

0 /.02

A /.06

I I I

Upper surface-– –-– Lower surface

o 10 20 30 40 50 60 70 80 90 /00Percent chord

(C) M, 0.90, 0.96,1.02,andl.06

Figure 5.-Concluded.

15

—. ———— .— —..

13616 NACA TN 30

-1.6M

o 0.40

-f.4 ‘ El .60A .70

0 .75-12 2

\

-t :) ~ .- Pcr f or M, O.70 ‘–

-.6

k&J*

>Q“- -.2<:

2

$

2 .2$

.8

12

Upper surfoce–––– – Lower surface

/.4 t

o /0 20 30 40 50 60 70 80 90 /00

Percent chord(a) M, O.@, 0.60, 0.70, and O.7~

Figure 6.-Effect of Mach number on the chordwise distribution ofpressure coefficient, a = 4.0°.

NAC!ATN

-16

-/.4

-12

-1.0

-.8

-.6

.4

.6

.8

1.0

$!2

/.4

Upper surface---- – Lower surface

I I I

44

(3 0.78

El .80

.82

$ .85

—..

17

0 /0 20 30 40 50 60 70 ~0 90 100

Percent chord

(b) M, 0.78,0.80,0,=82,mdO.85

Figure 6.-Continued.

-- __— .—. . -—— —_—. —_..

NACA TN 3036

M

o 0.88

.-.

1?

I

I

Upper surface-––—— Lower surface

I I I

o 10 20 30 40 50 60 70 80 90 /00Percent chord

(c) M, 0.88,0.90,an&O.94

Figure 6.- Continued.

NACA l!N

-/.6

-/.4

-L2

-1.0

-.8

-.6

.4

.6

.8

/.0

/.2

/.4

3036

-=M

O 0.96

El /.02

O /.06

I Lo-b‘d. . ...”

\1,,1 w 1I 1 I m I I —

1

–-–– – Lower surface

o 10 20 30 40 50 60 70 80 90 10

Percent chord

(d) M, 0.96,1.02,=dl.06.

Figure 6.-Concluded.

19

0

—- —..— .—.-—. ——. _.— ..— —— —-—

Upper surface---—– Lower surface I I

NACA

:

=/& for M, 0.65=

-H--Pcr for M, 0.75 —

w

.65

.75

o 10 20 30 40 50 60 70 80 90 /00

TN3036

Percent chord

‘(a) M, O.kO, 0.65,and 0.75

Figure 7.-Effect of Mach number on the chordwise distribution ofpressure coefficient, a = 8.00.

.

NACA TN 303621

M

0 0.80A 0.85

El 0.88

i

I \\\ i

1 I I I I I I I

— Upper surfuce Ill Ill‘— Lower surfdce

=+!$=

o 10 20 30 40 50 60 To B. go100

Percent chord

(b)M, 0.80, 0.85,and O.88

Figure 7.-Ccmtfiued.

—. . . . —_____ _____ ._._ .._. ..— —.— ._ __<

NAcA ~ 3036

M

o 0.90

~ /.06

!

1., t #

I Id

~~

JI “

II’IIa

I

Upper surface-- -– – Lower surface

1 I ! I

o 10 20 30 40 50 60 70 80 90 /00

Percent chord

(c) M, 0.90, 0.96, andl.06

Figure 7.-Concluded.

NACA TN 3036 23

.

Percentchord

o 0❑ 0.250 /.25- . .~ Z.au~ 5.00 . . .b /0.00Q20.00 I I I IQ 30.00450.00& 70,00v 87.50

.4 .5 .6 .7 .8

J40chnumbe~ M

(a) Upper surface.

Figure 8.- Variatim of local Mach number withat selected chordwise stations for an angle(approximately zero lift).

.9 Lo 1/

free-stream Mach numberof attack of 0.5°

. .—— —-—- .—_.. .—— .— -..

/J

IWC~TN3036

Percentchord

o 0E 0.250 1.25A 2.50~ 5.00Q /0.00Q 20.00Q 30.00#50.00 .Q ?0.00v 90.00 . . .

.“ ““

r

v

> () t.. . . ) e..>. “<) > ao 0’ O* ~ ““. “

“.4 .5 .6 .7 .8 .9 10 /.Ihfachnumbe~ M

(b) Lower surface.

Figure 8.-Concluded.

NACA TN 3036 25

.

.

.

.

Percentchord

o 0EI 0.250 /.25A 2.50~ 5.00k 25.00n 40.00950.00087.50 . .

)~

<).-. “., u c

L y

o

.4 .5 .6 .7 .8 .9 10 1/Mach numbe~ M

(a) Upper surface.

Figure 9.-Variation of local Mach number with free-stresm Mach numberat selected chordwise stations for an angle of attack of 4.0°.

.— -—-- ~———-—— —— ——. —.— —__—— . . . .

NACA TN 3036

Percentchord

o 0E 0.25010.00A 20.00k 40.00h 70.00Q 90.00

0

.4 .5 .6 .7 .8 .9 /.0 /./Mach numbe~ M

(b) Lower surface.

Figure 9.-Concluded.

NACA TN 3036 27

“

Percentchord

o 0❑ 0.250 2.50A 5.00L io.ooh 20.00Q 30.00060.0008Z50

/

/ A

T

/ m/

i

<~

om

y!=

.4 .5 .6 .7 .8 .9 10 IvAfoch numbe~ M

(a) upper surface.

Figure 10.- Variation of local Mach number with free-stream Mach numberat selected chordwise stations for an angle of attack of 8.00.

–.. ————..- .–——-—-—-——- ._————-——-—-

NACA ~ 3036

Percentchord

o 0❑ 0.25@ 2.50A 5.00k 20.00k 40.00Q 90.00”

),

,

/

..” “

-y=

.4 ~ .5 .6 .7 .8 .9 10 /./

Mach numbe~ M

(b) Lower surface.

Figure 10.- Concluded.

5WNACA TN 3036 29

.6

.4

.2

0

72-4 0 4 8 /2

Angle of attack, a, deg

aOfbbbbhb bbr!l ++;+

for A40f0.:0 .;0 .;0 .;5 .;0 .;5 .;0 .8!7 .:8 .90 .96 102 L06

.6$-

.4 <PQ

h / 7

.2

0 .{ ~ !i . ‘f ~ ,, (. / {i

%? IJ o -J -.2 73

Section pitching-moment coefficient, cm

Cmof+ + : + { + + } + + + + :

for Mof 0.40.50 .60 .65 .70 .75 .80 .85” .88 .90 .96 102 L06

Figure Il.- Section lift, pressure drag, and quarter-chord pitching-moment characteristics.

.—— ——..——.——.— .—— —— --- —. —-———. — — — ——.

30 NMYLTN3036

./6

./4

./2

(j? ./0

-.02

-.04

-.06

/

Rearward of maximum th~cknesI

\ ~

— — — / //

/./

/ -. —— __ __ _ .

/Forword of moximum thickness

.3 .4 .5 .6 .7 .8A40ch number, M

.9 /./

Figure 12.- The effect of Mach number on that part of the total sectionpressure drag coefficient due to the components of the airfoilforward and rearward of the point of maximum thickness.

NACA-1.aI@W-lO-~-5~- Iwo

——