NACA airfoils post stall.pdf

8/10/2019 NACA airfoils post stall.pdf

1/23

NATIONAL

ADVISORY

COMMITTEE

FORAERONAUTICS

TECHN ICAL

NOTE

3361

AERODYNAMICCHARACTERISTICS

OF

NACA0012AIRFOILSECTION

AT

ANGLESOFATTACK

FROM

0TO

180

ByChrisC.

Critzos,

Harry

H.

Heyson,

and

Robert

W.

Boswinkle,

Jr.

Langley

AeronauticalLaboratory

Langley

Field,Va.

occursa t

an

angle ofattack of

aboutlk. Asecondlift-coefficient

peak,havingavalue

of

I . 1 5 ,i sshown a t

an

angle ofattack

of

about V ? .

The

second

lift-coefficient

peak

i s

much

lessabruptthan theinitial o n e .


6/23

NACA

TN336l

As would

be

expected

for

a

symmetrical

airfoil

section,

initial

and

second

lift-coefficient

peaks,having

valuesnegativeto those

obtained

at

l>

and

V?,

ar e

obtained

at

3k6

and

315,

respectively.

orthe

sharp edge

foremost,

initial and

second lift-coefficient

peaks

having

magnitudes

of

0-77and 1.07,

respectively,

also

occur.

The

minimum

value

of

thesection

dragcoefficient,

although

not

shown clearly

in

figure1 ,was

found

to

be

about

0.007

with

. t h e

rounded

edge

foremost;

however,with thesharp edgeforemost,

a

minimum

value

of

about

0-01^wasobtained.eyondthestall,thesectiondragcoefficient

increased

with

angleofattack

untila

maximum

value

of 2.08

was

reached

atanglesof

attack

of

90

and

270.

The

section

pitching-moment

coefficient

is

shownin

figure

1

to

become

negative

after

the

stall

(a

lk)

and

to

remain

negative

to

a

=l80.

he

variation

of

pitching-moment

coefficient

with angle

of

attack isantisymmetricalaboutan angle

ofattackof l80.

Thevariationsof

theforce

and

momentcoefficients

with

angleof

^attack immediatelybeyond thestall ar eshown in

figure

1 tobe

functions

of the

direction

of

changeof

angle

of attack;

the

direction

in

which

theangleof

attackwaschanged inthisregion

is

indicated

infigure

1

by

arrows.s

the

angle

of

attack

was

increasedbeyond

the

stall,

the

lift

coefficient

is

higherandthe

dragcoefficient

is

lower

than

the

values

obtainedwiththeangleof

attackdecreasing from some

higher

angle.

Cross

plots

of

figure

1

yielded

the

drag

polar

offigure

2(a)

and

the pitching-moment

polar

of

figure

2(b) .

tmaybenoted in

figure

2(a)

that

thelift

coefficient

hasa

positivefinite value

atan

angle

of

attackof

90and anegative

finitevalueat270.he

finite

values

of

the

lift

coefficientat

these

angles

of

attack canprobablybe

attrib-

uted

to

the

fact

thatsome

lift

is being realizedovertheroundededge

of

the

airfoil.

he

finitevalue

ofthe

pitching-moment

coefficient

at

zero

angleof attack in

figure

2(b)is

probably

the result ofa slight

asymmetry in

the

model.

Effectsofapplying roughnessandofreducingtheReynoldsnumber.

Theapplication

of

roughness, at

the

leading

and

trailing edgesof

the

airfoil

at

aReynolds

number of 1. 8

x10

isshown infigure

3(a)

to

have

only

small

effects

on

the

lift

coefficients

obtained

at

angles

of

attack

from

2 5to125.owever,

at thestall

with

therounded edgeof the

airfoil

foremost,the

effect

ofroughness

was

to

reduce

themaximumlift

coefficient

from

I.33toI.07.oughness

is

also

shownto

reduce the

initial

and

second

lift-coefficient

peaks

obtained

with

the

sharp

edge

foremostand

to

reduce

slightly

the

lift-curve

slopenear=l80.

At

anglesof

attack from

0 to

165

,reducing

the

Reynoldsnumber

from

1. 8

x10

to 0. 5

x106 with

the

airfoil

surfacessmooth is

shownto


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NACA

TW

336l

haveeffects

ontheliftcoefficientsimilar

to

those

obtainedbythe

applicationof roughness.

owever,

ataReynoldsnumber of 0. 5X10,

the

initial

lift-coefficient

peak obtainedwiththesharp

edge

foremost

was

slightly

higher than

that

obtained ataReynoldsnumber

Of

1. 8

X10 .

Also,at the

lower

Reynolds

number,

thevariation

of

theliftcoefficient

with

angleof

attack

near

an

angle

of

attack of

l80is

quite

different

from

that

obtained

at

the

higher

Reynoldsnumber.etails

of

this

varia-

tion

ar e

discussed in thesubsequentsection.

The

application

of

roughnessata

Reynolds

number

of

1. 8 x10is

shown

(fig.3(a))to

reduce

the

drag

coefficient

at

=90

rom

a

valueof

2 .08

to

2.02;

reducing

the Reynolds

number

from1. 8

x

10 to

0 . 5 x10

with

theairfoil

surfaces

smoothresults

ina

reduction

of

thedrag

coefficientat=90o

a value

of

1-95hus,

the

surface

condition

and

the

Reynolds

number

ar e

shown

to

have

a

noticeable

effect

onthe

drag

at

an

angle

of

attackof

90.

t

might

be

expected

that the

lower

dragcoefficients

obtained at=

90

7

with

roughness

or

with

the

lower

Reynolds

number,mightbe

the

result

of

delayed or

incomplete

separation

at the

rounded edge.

f

this werethecase,

thelowerdrag

coefficientsshould beaccompanied byhigherliftcoefficientsthanwere

obtained

for the higher

Reynoldsnumbercondition

with

the

airfoil sur-

faces

smooth.owever,the validity

of this

expectationcannotbecor-

roborated

by

the

present results

since

the

differencesin

the

lift

coef-

ficientsfor the threetestconditionsatan

angle

of attack of 90ar e

smallandwithin the

experimental

accuracy.

Application

of

roughnessandreduction of theReynoldsnumberar e

shown

in

figure

3(a)

to

have

only

small

effects

on the

pitching-moment

coefficients.

Theeffects

(shown infig.3(a))

of

reducing the Reynoldsnumber

andof theapplication of

surface

roughnesson theforceandmomentcoef-

ficients

for

an

angle-of-attack

range

from

-2 to

32

ar e

presented

in

greater detail in

figure

3(b);these

effects

ar e

typical

ofwhathas

been

obtained

with

many

airfoil

sections

inthepastandthereforear enot

discussed

further.

Detailsofliftcurvesnear=l80. The

lift

curvesobtained

near=

l80

ith

the

model

surfacessmooth(fig.3(a))

ar e

presented

ingreater

detail

in

figure

k.

At

a

Reynolds

number

of

1. 8

x

10

(fig.

4(a)),

the

lift

coefficientisshown

to bea

continuous

function

of

angle

of

attack

through

an

angle

of attack

of

l80.

he

small

dis-

continuity

in

lift

coefficientat=182

sprobably

dueto

small

angle-of-attack

errors

in

alining

the

modelprevious

toone

orboth

of

thetestsin

which

datawere

obtained

at

this

nominal

angle

of attack.

At

aReynolds

number

of0-5x10^(fig.

4(b)),

it may

be

noted

that

not

only

doestheliftcoefficientappeartobea discontinuousfunction


8/23

NACA

TN

336l

of

angle

ofattack,

bub

also_the

dij^nj^ui;txJlccur&.-.a_^^

greaterL.t]aan.X8Q..for

the

increasing

angle

of attack

and

atan

angle

slightly

less

than

l80

for

the

decreasing

angle

of

attack.

t

may

also

he

noted

that

some

differencesar e

evident

in

the

data obtained fromtwo

testsin

which

the

angleof

attackwasincreased

from

different angles

ofattackbelow

l80toanglesofattack

beyond

l80.he

dataof

fig-

ure

4(b)

indicate that

the

hysteresis

effect

persists

all

the

wayto

the

stall.

omparison ofthe

data

of

figure4(b)

with

thoseof

figure4(a)

indicatesthat

larger

values

oflift

are

obtained

at

the

lowerReynolds

number allthe way

from

thediscontinuityto the

stall.

The

phenomena

just

described

were thought

to

have thefollowing

explanation:

ith theairfoilproducingpositivelift

near

an

angle

of

attack

of

l80,

the

flow

over

the

sharpairfoil

edge produces

a

small

separated

flow

region

ontheupper

surface,after

which

the

flow reattaches

to

thesurface;

the

boundary

layer

is

turbulent

from

thepoint of reattach-

ment

to the

downstream

separationpoint.

n

the

lower

surfacethefavor-

able

pressure

gradient,

which

exists

on the

surface

for

a

great

distance

from

the

upstreamedge

of the

airfoil,

is

conducive to

a

laminar

boundary

layer

from

theupstreamedgeoftheairfoilto theseparation

point.

n

the

region ofthe

high

adversepressure

gradient

at

the

rounded,

downstream

edgeof

the

airfoil,

the

laminar

boundary

layeron

the

lower

surface

would

beexpected to

separate

ata

more

upstream

location

than

theturbulent

boundary

layer on

the

upper

surface.nder

such

circumstances,

the

flow

inthe

vicinity

of

the

downstream

edgeof

the

airfoil

would

be

somewhat

similar to that overanairfoilhaving

a smallpositive

flap deflection.

The

sudden

change

(fig.

4(b))

froman

effective positiveflap

deflection

to

an

effective

negative

flap

deflection

would

occur,

of

course,

when

theboundary

layer

on

the

lower

surface

becameturbulent

and

the

boundary

layer

ontheuppersurface became

laminar.

he

hysteresisshown

infig-

ur e

4(b)

probably

results

from

the rathercomplicatedrelationship

between

the

pressure

field

aroundthe

airfoil

and

the

region

ofseparated

flow.

Separation

points

ar edependent uponthe

distributionof

surface

pres-

sure;

however,

the

pressure

distribution in turn depends

not

only

upon

the

boundary

shape

of theairfoilbutalso uponthe extentand location

of

theregions

ofseparation.

On

the basis

of

the

preceding

explanation,

the

presence

of

turbulent

boundary

layers

on

both

theupperand

lowersurfaces

ofthe

airfoil

should

eliminate

thediscontinuity in

the

lift

curve.

he

absence

of

the

dis-

continuity

at

a

Reynolds

number

of

1. 8

x

10

(fig.

4(a))

suggests

that

transitionhasoccurredonthe

surface

upstream

of

the

point

at

which

laminar

separation

tookplace

at

the

lower

Reynoldsnumber(fig.4(b)).

In

an

effort toobtain

some

additionalinformation

onthese phenomena,

lift

data

near

=

l80

ereobtained

at

aReynoldsnumber

of0. 5XI06

with

roughness

on

the

leadingand

trailing

edgesofthe

airfoil.he

purpose

ofapplying

the roughness

wastoestablishturbulent

boundary

layersonbothairfoil

surfaces

which,

again,

would

be

expected

to


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NACA

TN 356l

eliminate thediscontinuity intheliftcurve.heresultsobtained

were

rather

inconclusive,

however,

in

that

some

of the

effective-flap

effects

were

still

evident

in the

data.

he

roughness

in

this

experimentwas

the

same

as

that

used

inthepreviouslydiscussed

tests

atthe

higher

Reynoldsnumber,

andthe

possibility

existsthat thissizeof

roughness

at

the

lower

Reynoldsnumber

was

insufficient

to

cause

complete

transi-

tionto

a

turbulent

boundary

layer

on

both

surfaces.

Comparison ofpresent results

with

thoseobtainedinother facili-

ties

.

The

present resultsobtained

with

theNACA 0012

airfoil section

in theLangley

low-turbulencepressure

tunnel

(Langley

LTPT)

ar ecompared

in

figure

5with hithertounpublished data obtainedwith the

NACA0012

airfoil

section in

the

Langley 300

MFH

7-

by

10-foot

tunnel

(Langley

7

x1 0 )

and

with data from

reference

k obtained

with

the

NACA

0015

airfoilsection.

The

data

of

the present

investigation,

shownin figure

5 >were

obtained

at

aReynoldsnumber ofjL*8.X

1 0 . 6

withtheairfoil

surfaces

smooth.

urves

ar e

shown

for

the

data asobtained

(uncorrected for

tunnel-

wall

effects),

corrected

for

tunnel-wall

effects

bythe

method

of

refer-

ence

8

(samedata asinfig.1),andcorrected

for

tunnel-wall

effects

by

the

equations

of

reference

9-

pplication

of

the

tunnel-wall

correc-

tionsof either

reference8 or 9

to

the

presentliftanddragdatais

shown in

figure

5

t o

yield essentially

the

same

result

even

for

the

drag

coefficientatan

angle

of

attack

of

90

The

investigation

in

the

Langley300

MFH 7-

by

10-foot

tunnel

was

madeataReynoldsnumberof

I.56

x10and a Machnumber of about 0.20.

The

1-foot-chord

model

used

in

these

tests

completely

spanned

the

7-fot

dimension

ofthe tunnel so that theratio of

airfoil

chordtotunnel

height

was

1.5

timesthe

ratio

for

the

present

investigation.

he

Langley300

MPH

7- by10-foot

tunnel datawere

correctedbythe equations

of

reference

9

a n d -

the

corrected

data,

as

shown

in

figure

5 >

SJce

i

n

good

agreement

with

the

correcteddata

of

the

present

investigation.

The testsof

reference

k

ere made

at

a

Reynolds

number of_123.X.

10

with

a1.5-foot-chordNACA0015

airfoil section

spanning

the

shorter dimen-

sion

of

a

2.5-

by

9-foot

tunnel.

heindicatedairspeed

of

the

tests

is

stated in referencek to have

been 80

milesper

hour;thelift

anddrag

characteristics

were

determined

from

both

force

andpressure

measurements.

Onlythe

results

of

the

forcemeasurementare

presented

in

figure5

conclusion,

based

on

some

experiments

and assumptions,

was

reached

in

reference

4that tunnel-wall

correctionsto

the

datapresentedtherein

were

unnecessary.

The

lift

and

drag

coefficients

for

the

NACA 0015

airfoil

section

asobtained

inreference

k

ar eshown

in

figures

5(a)and

5(b)

to

be

much

less

thanthosefortheNACA 0012 airfoil discussedpreviously.he

differences

in

the

data obtainedwith

the

two

sections

appear greater

than

could beattributedtoa change

in

thicknessratio

from

12

to15 per-

cent.

seof the

pressure measurementsfromreferencek

for

thecomparisons


10/23

NACA

TN

336l

would

yield nobetter

overallagreementinthelift

variationsandwould

yield

pooreragreement

in

the

drag

variationssince

thefriction

drag

would

not

be

included.

pplication oftunnel-wall

corrections

(for

example,

those of

ref.

9 )

wouldresult

in

even

greater

disparityin the

data

obtained

with

thetwosections.

Thedragcoefficientfor

a

flat

plateof

infinite

aspect

ratio

inclinednormal

to the

flow

is

found

in

the

literature

(for

example,

refs.

10

to1 3 )tobe

very

nearly

2.0.hisvaluecomparesfavorably

with thedrag

coefficients

obtained in

the

presentinvestigation

with

the

airfoil

atan

angle

ofattack

of

90.

The

data ofreference10showamarked

effect

of

aspect

ratio

on

the

drag

of

a

flat

plate

at

=90.or example,

the drag

coefficient

of

a

flatplate

having

an

aspect

ratio

of

20

is

shown

to

be

about

1.48

in

comparison

withthe

two-dimensional

value

of

2.0.

s

pointedout

in

reference

6 ,this result emphasizes

a

basic

question,notye t resolved,

as

to

how two-dimensional data

shouldbeappliedtoa rotatingwing

for

those

cases

in

which

the

flow

over one

surfaceis

characterized

by

exten-

sive regions

of

separation.

CONCLUSIONS

Thefollowing

conclusions

maybe

made

regarding

the

resultsof

an

investigation

of

theaerodynamic

characteristics

of

theNACA 0012

air-

foil

section

at

angles

of

attack from

0

to

l80:

1.After thestall

with

the

roundededgeoftheairfoilforemost,

a

second lift-coefficientpeakwasobtained

atan

angle

of

attack

of

about

V?.

nitial

and

second

lift-coefficient

peaks were

also obtained

with the

sharp

edge

of

the

airfoil

foremost.

he

values

of the

lift

coefficientattheinitialand second

peaks

with

the rounded edgeof

the

airfoil

foremost

and at

theinitial

and

second

peaks

with the

sharp

edge

foremostwere1.33,

1.15,O.77,

and

1.07,respectively,

ataReynolds

number

of1. 8

x10 with the

airfoil

surfaces

smooth.

2.

A

small

finite

value

of theliftcoefficient

obtained

at

anangle

of attackof

90

was

probably

theresult of realizing

somelift over

the

rounded

edge

of

the

airfoil.

3.Application of

surface

roughnessat

the

leading

andtrailing

edgesand

reductionof

the Reynolds

number had only

small

effects

on

the

lift

coefficients

obtained at

angles

of

attackbetween25and

125.

k

.

At

a

Reynolds

numberof

0. 5X

10 with

the

airfoil

surfaces

smooth,

a

discontinuous

variation

of

lift

coefficient

with

angle

of

attackwasobtainednearanangle of attackofl80;thisresultis


11/23

10 ACA

TN

536l

believed to

have

been caused

by

adifference

in

the

chordwise

locations

of the

separation

points

on the

upper

and lower

surfaces.

. 5 .AtaReynoldsnumber

of1. 8 x10 with

theairfoil

surfaces

smooth,the

section

drag coefficientatan

angle

of

attack ofl80

was

about

twice

that

atan

angle

of

attack of0.

6.

The

drag

coefficients

obtained

atan

angle

of

attack

of

90

and

a

Reynoldsnumber of1. 8

x

10

were

2.08

and

2.02

with

theairfoil sur-

facessmooth andrough,respectively;thedragcoefficient

obtained

at

an angleofattack of90ataReynoldsnumber of0. 5x1 0

with

theair-

foil

surfaces

smoothwas1.95-hese

values

compare

favorably

with

the

dragcoefficient

of

about2.0obtained fromtheliteraturefora flat

plate

of

infinite

aspect

ratioinclinednormal to

the

flow.

7-

The

quarter-chordpitching-moment

coefficient

became

negative

after

thestallandremainednegativeuntilan

angle

of attack ofl80

was

reached.

8.Thedata of the presentinvestigationwerefoundto bein good

agreement

with

results

obtained with

a differentmodel of

the

same

air-

foilsection in

anotherfacilitywherethe

ratio

ofairfoil chordto

tunnel height

was

1. 5

times

larger

than

thatfor

the

present

investigation.

Langley

Aeronautical

Laboratory,

NationalAdvisory

Committee

for

Aeronautics,

Langley

Field,

Va.,

October

11,

1954.


12/23

NACA

TN

336l 1

REFERENCES

1 .Knight,Montgomery,

andWenzinger,

Carl

J.:

ind

Tunnel

Tests

on

a

Series

of

Wing

ModelsThroughaLarge Angleof

Attack

Range.

PartI- Force

Tests.

ACA

Rep.317,

1929.

2.Lock,C.N.

H. ,

and

Townend,

H.

C .

H. :ift

andDrag

of

TwoAero-

foilsMeasured

Over360

Range

of

Incidence.

.& . M.No.

958,

British A.R.C.,1925.

3 .

Shirmanow,P .

M.:

.-

Yawing

Moment

ofan

IsolatedAir-Foil.

II.- Test

ofan Air-Foil at

Angles

of

Incidence

From0

to

360

.

Rep.No.

271,

Trans.

CART

No .

36

(Moscow),

1928.

k.Pope,Alan:

he ForcesandPressures Over anNACA0015 AirfoilThrough

l80

DegreesAngleof

Attack.

eorgia

Tech.

Rep.No.E-102,Daniel

Guggenheim

School

of

Aeronautics,Feh.19Vf.Summary

of

paperalso

availablein

Aero

Digest,

vol.

58 ,

no.

i t Apr.

1949,

PP- >

78,

and

100.)

5.

Loft

in,Laurence

K. ,

Jr.,

andSmith,HamiltonA. :

erodynamic

Char-

acteristics

of15NACA

Airfoil

Sections

at

SevenReynoldsNumbers

From

0. 7

X

106to9.0

XK)6.

ACA

TN

19^5,

19^9

6.Loftin,

Laurence

K. ,Jr.:irfoil Section

Characteristics

at High

Anglesof

Attack.

ACA

TN

324l,

195^-

7.Loftin,LaurenceK. ,Jr.,andVon

Doenhoff,

Albert E. :

xploratory

Investigation

at

High

and

Low

SubsonicMachNumbers

ofTwo Experi-

mental 6-Percent-ThickAirfoilSections

Designed

To

HaveHigh

Maximum

Lift

Coefficients.

ACA

RM

L5IFO6,

1951.

8 .Abbott,Ira

H.,

Von

Doenhoff,Albert

E.,and

Stivers,

Louis

S. ,

Jr.:

Summary

of

Airfoil

Data.ACA Rep.

824,

19V?-Supersedes NACA

WRL-560.)

9.Allen,H.Julian,and

Vincenti,

Walter G.:

all Interference

in

a

Two-Dimensional-Flow

Wind

Tunnel,

With

Consideration

of

the

Effect

ofCompressibility.

ACARep.782,

I9V4.

Supersedes

NACAWRA-63.)

10.Wieselsberger,C:irplane

Body

(NonLifting System)

Drag and

Influence

on

LiftingSystem.ol.

TV

of

Aerodynamic

Theory,div.

K,

ch .II,

sec.1 ,

W.

F.

Durand,ed.,

Julius

Springer(Berlin),

1935*

pp.lVl-lVo.

11.

Wieselsberger,

C:

ersuche ber

denLuftwiderstand

^gerundeter

und

kantiger

Krper.

rgebn.

Aerodyn.Versuchsanst.

Gttingen,

Lfg.II,

1923,PP -33-3^-


13/23

12 ACA

TN

336l

12.

Prandtl,

Ludwig:ssentials

of

FluidDynamics.afnerPub.Co.

(New

York),

1952,

p.

182.

13.

Fluid MotionPanel of the

AeronauticalResearch

Committee

andOthers:

ModernDevelopmentsin

Fluid

Dynamics.ol.I ,

c h .

1 sec.

10,

S.

Goldstein,

ed.,TheClarendon

Press

(Oxford),

1938,p.37-


14/23

NACATN

36l

13

0

O

to

O

to

to

0

0

to

0

t>-

a )

1

C O

f

/

^

ra

1

0 -p

H

N

\

i

l

e

c

t

a

t

a

p

o

i

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)

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/

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H

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15/23

Ik

NACA

TN

35^1

0

0

O

0

0

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+3 +3

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f\

t-

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--0

^

in

C \J

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s

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in

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c

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16/23

NACA

ra 361

15

.n

J

s

h

225

5

r

h

f

315

0

/f

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\

30

0

.?

\k5

-,li

9

3

i**

r

1.2

-. 8 -

u r

Section

liftcoefficient,

( b )

Pitching

moment.

Figure

2.-

Concluded.


17/23

16

NACA

TN

561

sN

0

C O

m


18/23

NACA

TN

36l

17

1

to

oa

p

a

H

H

fto

R Airfoilurfacecondition

0

1.8

0M

0.5

x

10J

O

1.8

0

Leading

nd

railing

dge

ough

Flagged

ymbols

denote

decreasing

ngle of ttack

r

)

J

l-

Vjs

\T ~\/

\

\\

V

/^ SSr

1.2

1.0

p

p

a

o

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o

p

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12

6

0

Angle

of

ttack,

a,

deg

24

28-

3 2

( b )

Angles

ofattack from -2

to

32.

Figure

3.-Concluded.


19/23

18

NACA

TN 36l

1.2

- h

-1.2,

164

68

72

76

80

84

88

Angleof ttack, a, deg

0

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182

3

)

31

js

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y

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192 196

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176

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Angle

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lfil

,

deg

188

19 2

196

(b) R .5 06.

Figure

k.-

ariation

of

section

lift

coefficientwith

angle

of

attack

for

NACA

0012

airfoil

section

in

smoothcondition

near

an

angleof

attack of

l80.


20/23

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Documents

NACA airfoils post stall.pdf