Investigations on reductions of friction on wings, in particular by means of boundary layer suction


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Investigations on reductions of friction on wings, in particular by means of boundary layer suction
Series Title:
Physical Description:
iii, 92, 58 p. : ill ; 27 cm.
Pfenninger, Werner
United States -- National Advisory Committee for Aeronautics
Place of Publication:
Washington, D.C
Publication Date:


Subjects / Keywords:
Drag (Aerodynamics)   ( lcsh )
Boundary layer -- Research   ( lcsh )
Airplanes -- Wings -- Testing   ( lcsh )
federal government publication   ( marcgt )
bibliography   ( marcgt )
technical report   ( marcgt )
non-fiction   ( marcgt )


The drag of an airplane consists of the induced drag, the frictional and form drag of wing, fuselage, tail unit, and, occasionally, radiator drag. Investigations have shown the frictional drag to be the main portion of the drag. Thus the reduction of surface friction has gained considerable importance during the last years. Since the laminar friction is, in general, considerably lower that the turbulent friction, the frictional drag could be reduced by a laminar boundary layer as long as possible. The aim of the tests described here was to keep the boundary layer completely laminar up to the trailing edge of the wing.
Includes bibliographic references (p. 77-83).
Sponsored by National Advisory Committee for Aeronautics
Statement of Responsibility:
by Werner Pfenninger.
General Note:
"Report date August 1947."
General Note:
"Translation of "Untersuchungen über reibungsverminderungen an tragflügeln, insbesondere mit hilfe von grenzschichtabsaugung" From Mitteilungen aus dem Institut für Aerodynamik an der Eidgenössischen Technischen Hochschule in Zurich. Herausgegeben von Prof. Dr. J. Ackeret, Nr. 13."

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A( A SH\R1

NA2CA TM Nro. 1181




Chapter 1: Introduction, Abstract . 1
1. General rlemali~rks . . 1
2. Earlier reports puzbl~shed on the reduction of frictionlal
and profile-drage and on related fields .. ... 1
3. Influence of the trru-anstion-polut position on the
profle drag for largerl Reynold~s numbers;
statement of the purpose .. .. .. .. ... 4

Chanter 2: Causes of Transition .. .. .... .. 6
1. Influence of t~he exctelnal pressure g-radcient on the
-tmanrit/ on .. ... .. .. .. .. 6
2. Influence of the external turbulence on the
transition; tu~rbu~lence of theJ atmosphere .. .. 10

Chapter 3: L~amilnar Profiles with the Trasr~ition Ta~kinzg Place
Fa~r Tourarl thle Rea;- (with-out Bound~L~-ary-Le r Suction) .. 16
1. Gener-al considerations ... .. .. ... ... 16
2. Prelimninary~ tested on leminar profiles for-
Bemanda -e.ntrance ... ... .. ... .. 16
3'. Laml~nse profiles for propellers .. .. .. .. .. 18
Ib. Laminar profiles for winds .. .. .. .. .. 21

Chapterl It: Laminar Bounrdary-Layer Sulction, Gen~eral Rema~Rks 26
1..Aime for further levelopmenat ... .. .. .. .. 26
2. Effect of eactionZ of lamilnar boundary layer on the
flowr characteristics .. .. .. ... .. .. 26
3. Statement of the problem 27
4. HisftoYry f' deveilopmeint of~r the IslainarL bo)undEary-layet~r
suic t iorn .. .. . . 28
Chapter 5: Investigastion of the L~aminar Pressure Increase
with Boundary-Lasyer iuc-tion for Smnall~er anr? Medium
Reynolds~ Nu~mbere .. .. .. .. .. .- .. .. 36
1. Laminar suction tests with three suction slots
arranged one after the other ... .. .. 36'
2. Tests with~ lamirnar bound~ary-lalyer suction writh a
single suction slot .. .. .. .. ... .. .. 38

NACA T1M No. 1181


Chapter 6: Investigation of the Slot Flow for Laminar
Boundary-Layer Suotion with Single Slots .. .. 45
1. Laminar suction tests with straight suction slot .. 45
2. Investigation of the slot flow for lamrinar boundary-
layer suction with suction slot (i) curved forward
(definitions, see beginning of chapter 5) ... 47
3. Investigation of the elot flow for laminasr boundary-
layer suction with the rearward curved suction.
slot (h) .... ... .. .. .. 48

Chapter 7: Tests about KeCeping a Boundary Layer for Bigh Reynolde
NWumbers La-m~inar with the Aid of Boundary-Layer Suction .. 50
1. Purpose of the tests .. .. .. 50
2. Test apparatus ... .. .. .. .. .. .. 50
3. Measurements .t .. .. ... .. .. 51"
4. Symbols and evaluation of the section tests ... 52
5. Test results .. .. ., ... .~ .. 59
6. Extension of Schlichting's theory on the laminar
boundary.-layer development with area suctioh in
thae case of the acceleration of the sucked air to
the undisturbed free-stream velocity Uo *. ** ** 62

Chapter 8: Investigation of a Slightly Cambered Laminar
Suction profile of 10.5-Percent Thickness8 . 65
1. ~Purpose of the investigation ,. .. 65
2. Profile, test arrangement .. . .' 65
3. Measurementis with laminar boundary-layer suction .. 66fj
4. Test results .. .. .. .. 6
5, Conclusions from the tests of chapters 7 and 8 for the
desig)n.of larminar suction profiles with the lowest
posasble.drag for high Reynolds numbers .. .; ." V
6. Prospects for application of lampinar boundary-layer
suction in flight for high Reynolds numbers .. 73

Appendix *. *. .. ********* 5

References ....... ... .. r... .,... ,. .. .. 77


NACA TM No. 1181


The present retort deals with the reduction of frictional trag
by maintaining a more extended laminar bound.aryr layer, particularly
with the aid of bounda~r-1y-ly suction. The first chapters treat
publications in this field, the causes of the boundasry-layer transi-
tion and. a few laminar profiles without bouindasry-layer suction.
Next, tests with laminiar suction profiles are ?es-cribed. The
behavior of the suction slots for laminar bounda~ry-layer suction
was separately examined..

The present report was beguwn in 1.940 and financially supported
by the Committee for Study of A~viation. I feel obliged and am
glad to express here my sincere grati-tude to this committee and
particularly to its president, Prof. W. J. Ackeret for energetic
support of my work.

Digitized by the Internet Archive
in 2011 with funding from
University of Florida, George A. Smatherys Libraries with support from LYRASIS and the Sloan Foundation unit

2. Earl-ier Pleport~s Published on the RedFuction of

Fric tional and Pr-ofile Drass and on PEelated Fields

(a) The anssihili~ty of redjucin:- friction by maintnininG~ a
laminar boundary layJer for a longer time has been mentioned by
B. M. Jones (reference 1). Bi M. Jones prIoved later (reference 2)
that on finished wina Drofiles in fligh~lt there m-ight appear
laminar boundary layers of much greater extent znd with the
"Uhtersuchunge~n liber Peibungsvermindlerung7en an Tragfligeln,
insbesond3ere nu t il~e von Grent schichtabsaugung.z~" Mittel luneen aus
dem Institut fiir Aerodyn~amik an der Eidg~enissischen Tec~hnischen
Hochschule in Ziir~ich Herauag~eqeben von Prof. Dr. J. Ackreret Nr. 13.

NACA TM~ No. 1181





By Wderner Pfenninger



1. G-eneral Remarks

Thn diRag of an airplane consists of thle inducedt drarq, the
frictional and fo!m dra; of wing, fu.selaige, tail unit, a~nd,
occasionally, r~adilator drag. Investications haeve shown the
frictional drag to be the main portion of the drsl. Thus1 t~he
red~ction of surface friction hasz ained conrsiderable importance
during the last years.

the lemin~ar frictionr is, in IGeneral, conaSderabtl~y
the turbuleint friction, the fri~ctional d~aq could be
a laminar bDoundary layer as l~onig as possible. The
tests described here was to keep the boundaryr layer
l~ainar up to the trail~inf: edg3e o~f the w-ing.

lover than
reduced. by
aim of the

transition point lying farther to the rear than was expected so
far (compare Serby, Morgan, and Cooper (reference 3)). B. M. Jones
(reference 2), Squire-Young (reference 4), Pretsch (reference 5),
and Serby, Morgedi and Coopjer (reference 3) -investigated to what
degree the transition-point position affects the frictional and
profile drag.

According to these investigations a farther rearward position
of the transition point should make low profile drag possible
even for thicker profiles at; higher Reynolds numbers, Re. In
fact, tests on "laminar profiles" of this type resulted in con-
siderably smaller profile draqe, particul~ailyfor weak external
turbulence and larger Re referencess 6, 7, and 8). During the
second world war these laminar profiles were thoroughly investigated
in various countries.

(b) The position of the transition point depends mainly on
the exter-nal pressure distribution, the external turbulence, and
the nature and curvature of the surface. The influence of the
external pressure distribution an the trahisition-point position
was investigated by B. M. Jones (reference 2), Serby, Morgan,
and Cooper (reference 3), Eall and Hislop (reference 10), G.i I. Taylor
(reference 11), Faye and Preston (reference 12), Schubatuey
(reference.13), Faye (reference 14), etc. Flight tests by jones
(reference 2), Serby, Morgan, and Cooper (reference 3), and the
NACA (references 9 and 15) showed that for "clean" su-rfaces treasi-
tion in fligiiht generally takes place in the region of the point
of separation, even for higher Re~ynolds .numbers. Wind-tunnel tests
at moderate Re showed that the transition after a. slight pressure
increase, in general, takes place shortly after the separation
point of the laminar layer as long as the external turbulence does
not affect the transition (for instance, Hall and Hielop (refer-
ence 10)). A similar behavior in transition was found o~n bodies
of revolution (reference 16).

(c) Uhder the influence of an external turbulence the transition.
for higher -:Re occurs sanetimes at a considerable distance before
the point of laminar separation. The dependence of thie transition
on an external turbulence wa~s studied by G. I. Taylor (references 11,
17, and. 18), von Ka'rm n referencess 1-9 and 20), D~ryden (references 21
to 24), L'. Prandt1 (for instance (reference 25)), Schlichting
(references 26 and 27), Tollkien (references 28 and 29), Schubauer
(reference 13), Faye and Preston (reference 12), Faye (reference 30),
etc. For weaker external turbulence, in general, higher critical
Reynolds numbers are obtained at the point of transitioni compare,
for instance, B. M. Jones (reference 2), Hall and Hislop referencee 10),
Tani (reference 7), ~Lewils (reference 6), and also (reference 15).

N~ACA TM No. 118'1

NACA TMr No. 11815

For accelerated flow, too, there ressul~t higher critical Re,
compare Dr~yren (reference 21) and Peters (reference 31).

(d) For application of lamninar profile in flight for higher
Re, knowle9ee of atmospheric turbulence and its influence on the
transition in7 important. The atmospheric turbulence was investirgated
in flight wilth hot vires, among others by Stephens and Hall (refer-
enc~e 32). These two au!thors rePached the conclusion that in their
tests the Influence of the atmospheric turble~lnce on the transition
was nerpligilble.

American flight tests on a laminar profile of 15.9-percent
thickn~ess at Re = 17 x 104 confirmled thlis result (reference 15).
The influence of the vrall curvat~ure on the tranrsition was investigated
exPerilmenltall~y, for instance, by M. ardc F~. Clause; (reference 33)
and theoretically by HT. O~irt~ler (refer~ence 34). The measurements by
Clauser showed at the transition for convex (or concave) wall
curvature hieberl (or lower) Re -than fr a straight wvall comparee
also L.. Prandt1 (referecnce 3';) and RaylelCh (repferencre 86)).

(e) The influence of surface d.ilsturbances on the trasnsitfon
(roughn~esres, turbulence vTires, trip vJi~res, rivets, ~unevenesese in
the sheet-metlal skin, surface leaksweset) sinsige,
for nstnce byYou (reference 387), andl Faye and Preston referencese 12). Tests3 of
the author PhovJed that 'lafminar n~rofiles at hiubehr Re ase sensitive
to vavziness of the surface sadr that sulrfice leaks must be avoided
in order to nrev~ent thle air f'rom bein; sucked fomm the wing
interior and the bounrlary layer thus :;rom becoming turbulent.

(f) The pressure distribution of suitable profile shapes6
may be calcurlated with the aid of conformal mapping;, for inst~ace,
according to Theolorsenr preferencee 3c)) (rompare also references 40,
41, and. i62), The s-ingula~rity methods which replace the profile
by vortices, o1u~rces and sinks, takte.lesse time f'or the detailed
numerical calculation, bu.t are less accurate. From th~e pressure
distribution one may7 calculate the development of the boundary
layer in the laminar and turbulent part of the profile.

(g) The laminar boundary-layer development may be determined,
for instance, according to Pohlhausen (reference 43), Falkn~er and
SkaLn (reference )ch), Falknuer (references 45 and 46), and Hovertrh
(reference 48). For moderately accelerated and slightly retarded
flow the approximation method of Pohlhausen is well applicable as
demonstrated by a c~omlparison with Falkner's method (referencesI 45
and 46) for an external velocity distribution U = kxBm. According
to Pohlhausen, the laminar-plate friction (pressure-gradient zero)

NACA rm No. 1181

is overestimated by 3.3 percent. .(Comparieta with Blasius (refer-
ence 47).) According to Bowarth (reference 48), the laminar
separation 'starts in the case: of the external velocity distribution
U = UO bx at X = 62 u' -7.5j instead of X = -12 according to
Pohlhausen;' For the case U = kx-0.0904 the method of aalkoner
(references 4T eWL 46) results 06i boundary-layrer profiles with
vertical, tangenti already at X1 -5,

Howarth (reference 5l) gave a compilation of various methods
for calculation of the' laminar boundary-layer development.
L. Prandt1 (reference 49) and H. Gijrtler (reference 50) investigated
the laminar boundary-layer development by -exact calculation and
compared the different.tsown-solut;ions with each other.

Tomotika (reference 52).d8evelop~ed a method corresponding to
that of P~ohlhau~sen for the three.-dimensional case.

(h) The turbulent bound~ary-layrer .development may be calculated
according tto A. Buri (reference 53), Gruschwitz (reference 54),
Keh1 (reference 55), Squire-Young referencee 4), and Young (.refer-
ence 56). In most cases, the turbulent shear stress at the wall
is the same as that for the flat plate with turbulent flow without
pressure gradient for equal Rep. The ratio -" = was frE~quently
assumed. constant (H = 1.4 for not too large pressure8 increase).
Otherwi se, Hcould be determined. from turbulent boundary-layer
measurements as a function of the pressure increase (compare, for
instance, Gruschwitz (reference 54)). The turbulent boundary-
layer development along the fl/zt plate without pressure gradient
was investigated. by Th. von Karman (references 57 and 58) and
L. Prandt1 (reference- 59) .

The boundary-layer development of the wake may be calculated,
for instance, according to Squire-Young (reference 4).

3. Influence of the Transition-Point Position on the

Profile Drag; for Larger Reynolds Numibersj

Statement of the Purpose

For slightly cam ered profiles of'various- thickness ratio
d/t at Re = 15 x 10D the profile drag was calculated according to
Squire-Young (referedee 4) for various popiftions of the transition
point x/t (compare fig. i)- 'Fur~thermo~re,, the profile drag was

NACA TM No. 1181

calculated for a profile of 16-Dercent thickness for various P~e
and~ different positions of the transition point. (See fig. 2.)

For a rearw~ard movement of the transition point ait larger
Re, CgO vould drop to low values. In thle ideal case (boundary
layer kept, completely laminarl up to the trailing ; rcde), even
thicker prcf-iles would gi~ve very low dragq for larger De's.
Thicker p~of~iles at-e structurally more favrorable and permit larger
spans, thereby reducing the induced draq which wNouLd.~.reS;?in
significance due to'thle decrease of the frilction-l drag Moreover,
thicker profiles permit, a me.-e favorable installation in the win::
of fuel tanks, power plantsE, other loads, and finallys, suction
ducts. Furrthermore, thicker profiles possess hiihezr raximuma lift
with suitable high-lilft. devices. For faster airplanes the maximum
profile thi4ckness~ is probntlp dependent on the stipulat~ion of
sulfficientl~y small superst~remn vrelocit-ies in order to reach high
Mach numbers without compression shoclrs.

Ac3cordinr: to these considerations, th7e followsInP: aim was set:
Development of thicker profiles rwith snal~l superstream velocities
where the boundary layer remains., in flight, at, high P~eynol3s
Numbers, Imminar as far toward the rear as possible, if possible
as far as the trailing~ edte. Thre anaxJmurn Itft for take-off and
landiin.g is to be as hig,:h as possible.

The p~ossibililties of kieepine thle boundarqy layer laminar
for a long time are, for instance:

(a) Use of profile forms where, by specal design1 of the
contour, the transition is shifted rearward (f~lst pressure
distribution with small superstr~eam velocities ann delayed pressure

(b) Preventing the boundaary layer from becominT tulrbulent
by means of boundary-layper suction, possibly in combination with~
profile forms with pressurre distribution.

In order to study the methods for keeping the boundary
layer laminar for a longer time, one must know the pr~esumable
causes of transition.

NACA TM-No., 1181



:11. Influence of the lbdmerand Pressure GrLadierit

on the Transition

Transition tests showed that, in .general, the transltion-
takes -place after a slight pressure increase in the neighborhood
of the point of laminar separation, as long a~s .an external ture-
bulenco does not affect the transition (reference, -2, 3, 1j, 12,
13, 15, 16, and 60O). These observations were confirmed by tests
of the- author, for instance, on at NI~ACA 0010 profile (fi~g. 3), on
lamninar profoilees of 10lperPcent and 14-perceht thickness (figs.,le
and 5), on a body of revolution (fig. 6) and, later, by 3aminar.~
suction tests. With increasing Re the transition point moves
forward more slowly for laminar profiles than for so far conven-
tional ones.

These tests, as well as those of chapter 3 and the later
suction tests of chapter 4, 4C, -chapter 6, 3, chapters 7 and 8,'
w~ere preformed in the large wind tunnel of the Institute described
in reference 61a. The wind-tunnel turbulence was

u1 ~ iPT = 0.0040 to 0.0045
Uo Uo

The pressure gradient in the adjustable test section is very small.

In. figure 3 the prnessure distribution 1- is plotted. on an
NACA-0010 profile (.so far conventional profile fond) for ca = 0
andvarousRe =- along the chord (Uo free-etream velocity,
t = chord = 0.60 m)* The static pressure p was measured through
0.(5 millimeter q, bore holes compared to the s~ta~tic pressure po
in the test ~section without wing. The freers~tream stagna~tioti

NACA TM Ho. 1181

pressure was go = So Po (eo = Lundisturbedl total prepsurel).
The positions of the transition point (slrrows) were ascertained
from the break In the orerssure-distributiono curves at th~e transi-
tion (compare transition tests on NACA 0010-p;rofie-with soot.
coatina and stethoscope (reference 6510)), For larger Pe the
determination of the position of the transition point, f ~om the
break in the pressure distrbbutionse became u~nreliable. With
increasing Pe the transition point shifts rapidly forward.

`In figc~uree 6 and Cj the pressure r3istribut-ion p/q, of two
lam-inar profiles (fi'. 12~) ics -lo-ttfed versus the chorda with
6/t = 10 percent, t = 0.76r m;l/t = ill Der~cent, t = 0.70 m.
The static pressure p wass meaeured wlith 9 ste~tic-oessure tube
of dl = 2.0-millimeter' 11xmeter, wh~ich was 'Jut tangentiilly to
the test point in the flavc~ -irretion versus po(static pressure
in the test section for "tunnel-emot;." condi tion) The fouxr
0.4 millimeter r9 connecting static-Pressure holes of the static-
nressure tu~be were located 10 mil'limeter-s behind the semicircullar
head and 1.00 millimeters ahead of the stinrg of 3.0 ad.llimeter~s rp.
A. comPerlron of pressur7e-dli t~rsbut~ion measuremnents~ on a laminar
profile of 16.7-per~cent t~hickn~ess a~nd3 2.6-percent, camb~es of the
profile mean line showed~ t~hitr p/o on the uopFe.-. and~ lower side,
respect~ively~, was measuredj with the stastdc-pressure tube on the
average by 0.007 and. 0.006, resfpectively, to~o h~iph, as compared
with the measulrement by means of connectin-i asttic-?iressure holes.

The transition (arrow) was determined by means of the
stethoilcope (reference 61c) and fr~om the break in the pressure-
distribution curves preferencee 16).

IEy means of anl ann~ular equalizer ooenin: at the end. of the
closed test section one succeeds in establishinr there atmoslJheFric:
pressure (f or model present andi tunnel emptyv condition). The
undisturbed static psre~ssre for modell present" very far in front
of the model is a~ppr~oximately equal. to the static pressure po in
the test section at the location of the mod3el if the tunnel is
empty neglectingg of the vake behind. model ani suspensionn.

It is self-evident that. t~hreb~y the profile properties are
,riven in such a manner as if the profile werce workin: in a closed
tunnel, not in the unlimited air stream. Since the models investi-
gated here were small relative to the tunnel cross section (2.1 x 3m,
octagronal), the respective jet correction was omitted (about a
corresponding correction of go, Uo> po for larger model dimensions
for two-dimensional flow. (See chapter 8.),

NACA TM No. 1181

The pressure distribution -? is plotted on figure 6 for a
:bodyT of revolution of D=0.212 oat 0.47 xL& from the front
for asymmetrical flow at various ReL q -fi~-, (.D = maxcimum diameter
of the body of revolution, L = length of the body of revolution
= '0.85 m).

The body of revolution was held from behind by a cylindrical
sting, suspension vire-s on t~he~body were avoided. p was measured
versus ~po (static pressure in the empty tunnel) by mesan of a.
2.0-millimeter static-press~ure tube.

The position of the transitionl point (arrow) was ascertained
with the stetho scope and from the break in the p~res sure -di stribution
curve. The sliight superstrearn velocities result in a pressure
increase and transition far -to the r~ea-r.

Drag~ measurements: At the end of the body of revolution at
the juncture to the sting the Boundary-layrer profiles were measured
for symmetrical flow at various ReL. The total pressure ,e in
the boundary layer was determined by means of a, flat total, head
tuibe of 0.2-millimeter x 2.5-millimeter inner crose section and
0.5-miilimater external height. The static pressure in the boundary
layer was measured by means of a 2.0-mi~llimeter static pressure.
tube. From the boundary-layer measurement at the end of the
body the momentum-loss area Xmfar to the rear was calculated
for the undisturbed static pressure poaccording to Young
(reference 56).

X, = 2xL wak 1 ) dr far toardn rear

The drag coefficient c 0 and cw, eferred tothe body
surface area 0) = 0.369m21 and the maximum cross-section area

-x D2 = 0.0294 m2, reepictively, then becomes:

2X, 2Xo 2'Xoo 2/3 = 0.0570 mn2
c,,~ = -O~-, .cwH --2 eq/3 = v2/73' (V = Volume)

CW0 (ReL) is plotted on figure 6. cw0 decreases with ReLi at
first similarly to the friction of the laminar flat plate

NACAP TM HTo. 1181

c-1.328 see Blasiuls (reference 47) and incresees- again for
higher ReL, due to more pronounced shifting forward of the
transit-ion point (observntions by stethoscope), caused by the
tunnel turbulence. The minimum drag coefficient referred to the
maximum c-ose section resulted as cH = 0.01r06. For sJmaller
Re, stethoscope observations showed that thecl boundary layer at
the end of the boiy undlergoe~s lasminar separatson andl does not
readhere with the properties of a turbulent boundaryr layer, therby
greatly increasing pressure and; total dlraeG.

The position of the transition point can be determined in
different ways:

(a) From hot-wire observavtionsp (comparsle, for instance,
Dryden referencese 21))

(b) By measurement of the total surface pressure along the
chord ri~th a fine total head tub~e (refere-nces 2 and 10)

(c) By toonrda.,-rlyleer mreasurement~s (compare, for instance,
(referencenrE 16))

(d) Fromr thF t2e brea in the pressure-distributionn curve at
the tranrsition, causedl by ~the su~dcen decreased; of the
dli sTl ac emen t thi cknesE 5'' at the transition ( cormpare
A., Fag.e ( :e ference 14) and. t~he pres sure -distribut~ion
curves of figures 3, 4, 5, /I, 39, h0, 77-80, etc.,
compare also calculations by A. Betz (reference 77)
for discontinuous change of thle cuirvature)

(e) Acourtically by stathoscope observations (Irefe:ence 61c)

(f) By soot coatinrr (reference S10)

(5) By measurement of the total head of the boundary lawyer
along the chord at a greater distance from the wall
(reference 2)

As9 long as the transition is caused by an external pressure
increase and not by an ex~terncll turbulence, it takes place in a
narrow comparatively well-defined zone. The methods indicated
then yield, In general, the transition-point position reliably.

Presumable cause for transition: According to Ralyleigh
(reference 62), Tietjens (reference 63b), and Tollkien preferencee 29)

NADA TM No. 1181

lamninar boundary-layer profiles with inflection point, as they
originate with rising pressure, are unstable. Transition tests
showed that the tranaltion occurs wrhen laminar boundary-layer
profiles having a sligh~t-reaarvad flow-ib~the neighborhpod of
the vall exist comparee, for instance (references 10, ~13, and 78),
this was confirmed by transition observations of the author with
soot coating~ for med~ian Re. The originating of sufficiently
strong vortices in the immediate neighborhood of the wall (as they
form, for instance, for larger lamninar pressure increase, for
discontinuities in the surface, by an1 external turbulence, or for
tube flove with sharp-edge inlet) seems to be required for the
transition (compare L. Prand~t1 (reference 63a), Jr. Pat~ry (refer-
ence 64), and L. Schiller (reference 79)).

The Reynolds number Rel x referred to the distance 1
between the laminar separation point and the start of transition
resulted as Rel = 40,000 to 70,000 as long as the external
turbulence does not affect the transition comparee references 9
and 10, confirmation by measurements of the author). U = mean
velocity at the edge of the boundary layer between laminar
separation point and start of transition.

For very weak external turbulence and larger Re the tran-
sition for laminar profiles is probably to be expected in the
region of the laminar separation point.

2. Influence of an External Turbulence on the

Transition; Turbulence of the Atmosphere

(a) Various drag measurements and transition measurements on
flat plates, profiles, and bodies of revolution had shown that for
larger Rieynolds numbers the transition, under the influence of the
external turbulence, takes place considerably farther in front and
that the drag increases again. The same observation was made in
tests of the author on laminar profiles without boundary-layer suction
(see, for instance, Cc4m (Re) of a profile of 3.35-percent
thickness (chapter 7* fig. 88, curve a, and corresponding position
of the~transition point, fig. 89, also figs. 7, 8, and 10) atnd on
a, body of revolution with flat pressure distribution and transition
lying far back (fig.' 6, chapter 2, 1). For larger Reynolds numbers,
individual turbulent bursts wrere determined in the bou~nd~ary layer
with a stethoscope which occur more and more frequently downstream
with increasing boundary-layer thickness until the boundary lawyer
becomes fully turbulent. The transition takes place in a Idore or

NyACA TM No. 1181

less wide transition regain; its positio3nie less re--dily ascertained
according to the mrethods indica~ted. in ch'pter 2, 1 than in the case
of transition due to external pressure incrlesee. A renewed increase
of the profile drag due to the external turtulen~e cre alsoi obseuvtrve
for laminar profiles with boundrlsLy-layer suction, foir lar::er

Re = U (fi3s. 18, 88, 9'i or chapter Ir, b4c, chapters 7 and 8).

(b) Causes of transition for tra~nsfuncn ?iue to an external
turbulence: By; the external tulrbolent velociti y fluctuations
considelrable veloc~ity fluctUa~tion: n fl~~ow lizec~tion origiinate In
a boundary layer (see Tollkien (referecnce 28b), G. J. Taylor (refer-
ences 11 and 18) and Drydren (refernce 21.)), th~us causing temporarily
unstable profiles with inflection point: and finally
reverse flowl in the neighborhood of the vall. The transition then
occurs, as in the case of larger laminar ,rereure increase, due to
the formation of vortices at t~he surface. There o~rioinate isolated
turbulent discontinuities which become orruse anl mor~ie frequent
downstream with increasing : bou~ndar~y-layer thilcknees until the
boundary layer is fully turbulent.

G. I. Taylor (reference 18), Hall and-. Hislop referencee 10),
Fage and Preston (reference 12), and Farce (r~efernce '0) attempted
to find a criterionz for the transition du~e to an evte;'na3 turbulence.
They assumed that the transition starts when there orlyinate, under
the influence of the external tubrblent pressure fluctuations,o
momenta-rilg boundary-layer profiles wirth vertical tanP;ent, that is,
X Pohlhausen = -12.. or, respectively, its mean sqaere value

6' =1~ =Acitical

According to G. I. Taylor referencess 11 and 18g) turbulent
pressure fluxctuastions originate in the case of' Ln external tur-
bulence, their mean square value is, for isotrope turbulence,


NACA TM No. 1181

Sis a parameter connected. with the diffusion in a turbulent
fow (1q = length of the diffusion).

(X? and A (mean magZnitude of the srmallest vortices
significant for the dissipation) were, for instance, experilmentally
determined. by Hall and H~islop (reference 10) from temperature
distribution measurements behind a heated wire and from measlurementsq
of correlation behind grids.)
Taylor deduces by insertion of --1for isotrope turbu-
lence behind a grid (mash vidth M) that the critical :Re~yolds

number Re critical i critical is a function of

0 .2 i s: fu 'cr ti 2 c r ti c n 2

wherein A = constant,

according to Taylor, This relation was experimentally tested byv
Hall and Hislop (reference 10). According t ity, Xe
I1increases wi th~I dIiminihingI~l~ vaLtlue \g) -6 somewhat more slowlyjcil a
than according to the theory, that is., X decreases with
increasing Rexcritical. This^circumstance may perhaps be explained
from the fact that a sufficiently strong reverse flow at the w:!ll
(X4-12) and not only X = -12 I(vertical tangent to- the velocity
profile) is decisive for .the transition. For small Reft larger
negative h-values ar~e necessary for the transition if one assumes
the reverse flow required for transition to be of equal strength
(with equal circulation).

Turbulence crriterion of Fage (reference'30): thder the
assumption that the transition starts when momentar-ily a critical
negative h-value Is reached due to the t~urbul!ent-pressu-re
fluctuations, Fage finds that the value

a ~ ~ ~ ~ ~ / 3.5R riia00.2

NrACA 'TMt No. 1181

ought to be constant, with

L turbulence scale B d

Ry correlation for velocity fluctuations in 2 points at dizt~anc~e
y perpenficular to flow

Pep~llp. = -- cti at the transition

Fage finds that EL decreases with increasing Regoritical

For accelerated or for retarded flow the transition occurs,
due to external turbulence, at hi!-her or -3+ consid;elrably lower
Recrtia-values, respct~ively, camnpare, for instance, figures 3
and b. Influence of the atmospheric tu:bulelnce o~n th2e transition:

The turbuxlence measurements in flight, bi Stephens clnr Hall
(reference 132) had aiven the reFsult that, in their tests, the
influence of thle stmospTheric tulrbulcnce on th~e transition w:as
negligible. In most cases, the tulrbu!elencyie-eee of the Lltmosphere
was very Lsmall: "- n .0003I 'forl U L 5 m/sec in calm air,
U = velocity ou.tsides the boundary~ lOve;. Considlerabl~y larger
turbulent fluctuations were foundm for unstablp stratification of
the air (u'/t 0 .003i for UI = 66 mI/Eec), howevr;V~ Idue to the
turbullence scale which wacs Eo much lar.ev than in thie windl tunnel,
their effect on the trnzs-iti~on ase, accord~in-~ to Sfetephns and Hal~l,
of secondaryr importaince. (Magini~t.ud of th16e smllect vorLices n.3 m).
Stephens and Hall fou~ndl the fluctualtions in t~he bounrla;ry player to
be considersbl;- sma!.le-r in flight thanl in wrin J-tlunnel teat.s (refer-
ence 21). The same cheervation was made in tests of lamilnsr bounlndiar
layers i~n tunnels of low turbulence: (irefernc~e C;) and in tests of
the author with a tube flowJ at low turbualence. American measurements
in fli10ht on a laminar profile of 15.o-per~cent thickne~s seem to
confirm that the atmospherice turbulence a~ffectsl the transition only~

sliEghtly, even for larger Re (reference 15): For Re = U = 1 .
and a Mach number M = 0.52 the measured value vae c0~ ,= 0.00~3n,
which from the viewpoint of calculation would correspond to a mean
position of the transition point (mean of upper and lovwer side) of
about 0.68t from the front in the rerdon of the point of lami~nar
separation. From boundary-layer calculations, shortly before the
transition give a Reynolds number Reg = = 2600?, referred to the

momentum loss thickness 8 and the Pelocityr U at the edg~e of
the boundaryr layer, that is, essentially more than was observed
in conventional wind tunnels,

6 = momentum lose thickness = 1 '- ; dy

s" = displacement; thickness = 1 u g


8 total bound~ary-layer thickneess

u velocity in the borundtary laveri at; the distance from the vagli y

Uo flight oelocit~y
Tests of the author with a laminar-tube flow in the starting
region shotted, ~thkt considerably higher Re can be reached for a
freei stress free of turbulence.

Tests with a laminarr-tube flow in the startingr resgion.- The
purpose of these tests was to ob-tai~n hi-h laminar Reynoldsr numbers
'by means of a latminar-tute flow in the startling re;gion.

_Test arangeent. A conic~al inlet, furnnl of 0.9 meter length
maximm a diamneter at the btntrance 0.18 m) was fixed to one end of
a cylindrical anticorrosion tube of 6j meter lenoth and 0.025-meter
inner diameter. The transition to the tube was smooth. At the
entrance of the inlet funnel was a rectifier consistin. of circular
tuibules of 3 miflli~meters 93, about 0.1-millimeter wall1 thickness
and 0".2 meter l~en ;th in stacks. A Laval nozzlih was attache-1 to
the other end of th~e tube which was connected with the evacuated
su;per'sonic tunnel of the Institute (reference 61a). The air was
'sucked from the spaCe. The air motion at the entrance of the inlet
funnel was kept as anall as possible.

Measuremnents.- The' static pressure along the tqbe was mesasred
with 0.8 millimeter q, connecting static-pressure holes wiith the
atmosphere as reference level, Furt~her,~ the state of the boundary
layer a~long the tube was tested wjith the stethoscope (refernce 61c)
which was attached to the s~tatic-pressur~e holes. Constan~t air
vel'ci~ty was obta'nedi -througESh thei Lcval7 n.!zzle a~t the rearu End of

NACA .TM NJo. 1181

NACA TM No. 1181 15j

the tube~ and dlisturbances from upatrFma ef~fectin" th~e b~oundlary
player of the tu~be wer~e avoided. WTith this telt: arrsngement the
tube flow could be topt laminar up to considerable Reynolds3F
nurmbers, it is true, arccelerated. flor extisted.. The3 madmur
stagnlation pressure at the end of the tube, in its center, with
laminsar boundar~y layer up to +he end of the tube, amnountei! to
180kl~Filoorms per meter", corresponding to~ a velcityg of j6.5 meter-s
Per second in the center of the tube. The numerical evasluntlon,
according to L. Schiller (reference 83), resuilte2 in a total
bowunary-layer thickness 8 = 6.2lillimeter;s at the tube end nd
thle following Reynolrds numbers:

Repl = 2600 Regl = 6"50


Tl~ = Pei 2 13.9 x 103


U velocity at the edle of the bouindary 13 or

Rei 1 ~ 1 4B
u vrelocityl in bounrdsar layer
r variable radius

R tube radius (0.0125 m)

So far, sphere or hot-vire measurements we-re performedl
frequently for determination of the turbulence. The eph)ere test,
in particular, becomes unreliable w~hen the exter~nzl tur;bullence
is veryr wea~k. For ma~ny purposaes, for instanrcp, thle application
of wind-turnnel tests on laminur profiles to fli~ht conditions,
the Ree-values, wJhich can be: obtained] with a la~minar boundary
layer and a flat pressure distribution, are the main object of
interest. The manner in which these values are reached is often
of little importance.




1. General Considerations

Since the transition occurs mostly after a slight pressure
increase, probably those profile forms will be favorable for
keeping the boundary-layer laminar for a longer time in which the
pressure distribution is flat and the pressure increase lies far
toward the rear. The flat pressure distribution results in smaller
superstressm velocities anid thus higher Mach numbers without
compression shocks) of course, separations in the r'egion of the
pressure rise must be avoided.

Profiles with such pressure d~istri`butions have the maximum
.thickness at (0.4 to o*5)t from the front.

2. Preliminary Tests on Liaminar P~rofiles

for Smooth Entrance2

A few lightly cambered laminar profiles of various thicknesses
were designed according to these considerations (fig. 9).

The profile drag was determined for smooth entrance for various
Re = ---by memas of the momentum method. (See, for instance,
reference 61c) (fig. 10).)

The laminar profiles investigated here are, for larger Re,
superior with respect to drag to profiles of equal thickness used
so far. (Compare with the NACA profile 23012.) Vith increasing
Re, cwm decreases more rapidlyr than for conventional profiles.
The drag incr~ease-for larger Re (caused by the faster forward
travel of~the transition point due to the external -tu~rbulence)
ata~rts~at higher Reynolds numbers than it, did for profile used
so far,

2Smooth entrance: No flowr around t~he profile mean line at
the nose of the wing.

NACA TM No. 1181

NACA TM No. 1181

For smaller Pe the laminar p-rolf:es become more unfavorable
with respect to dlrag since the bourndery layer remains lami~nar for
too long, therefo-re undergoes laminar- separ~ation in the -re-r part
of the profile and does not readhcre- ?alan in a purely turbulent
manner. (Observations by stet~ho~scope.) The pressure dracq is
thereby et~ro~ngly increased. By artificially cre--ting a turbulent
boundary layer In the region of the point of laminstna secpanat'on
it, is possible, in many cases, to prevent a aere extensive lamina7r
separation and to obtain a turtulent readhering of the btoundaryp
layer connected with a correspondiingz drag decr~ease. Thr!E, the
drag olf the thin ]amlnar p~rofile number 7 for emaller Fep could
be essentizllyr reduced by bl.owing-off of air fro7m fine blowing
holes which rendered the bollndary layer in the region of~thie point
of lamilnar secara~tion tu~rbu-lent.

The blowing holes of 0.8-millimeter 9 and ?-millimeter
length vero placed vertically to the ving surface, th=y were an
the upper side 133 millimeters, on the lower side ?.1 millimeters
ahead of the trailing edge. The ving chord- wasP t = 0.60 meter.
The spacing of the holes was 26 millimeters. The total energy~ of
the air at the entrance of the blowing~ hol4es :;s practical~ly equally
to the undisturbed total head in the turnnel, Turibulen~t vedges
originated behind the blowing holes (obse;rys.tions by tthbocope and
with soot coating) which rapidly fused, thus c~ausinr the boundalry
layer to become turbulent over the entire span.

The mean profile drag c, over- the series of staed holes
was meas~uredl by t~he mlomenturm method Clt vrari-one Fe The
follov~ing drag resurlteil:

Re = 1.h65 x 106 ,1,13 x 10 ;,0.75 x 10h 0.52P x 106

c, = 0.003,3,0.00.00 .0T, 0 2)~,0. 0053

eve was reduced mainly in the region Po = (0.h to 1.7) x 106

A furrt~her slight drag3 reductions fTor .snaller P~e (0.3 x 10~
to 0.6 x 106) vlae obtained by placing the 0. millimeter
blowin3 holes farther to the front of the inr~ (on t~he ulpper
surface 155 millimeters, on the lower surface 110 millimeters ahead
of the trailingg edge) for equal specinp of the holes. Test results:

NACA TN No. 1181Z

Re = 0.33 x 106, 0.T0 x 10q 0.68 x 10q 1.10 x 106

cs = .0.0063j, .0051, ,.005, 0.00395i

For larger, Re, cm is larger than for the case of the blowing
holes lying farther toward the rear.

The dashed c m-c~urve of figure 10 of profile number 7 Rives
the optimum c m-values for the~ most f~avorable position of the
blowing holes in the direction of the chord.

Further tests showed that the spacing of the holes, the hole-
diameter, and the total head at the entrance of the blowing holes
may be widely varied in order to make the ~boundary layer for
smaller Re. artificially tur~bulent, For smaller Re, bounmdary-
layer measurements at the trailing edge of the wi-ng, witn and
without blowing~ holes, .resulted in considerably thinner and fuller
boundit~y layers largerr 8 l9 values) when air was blown into
the boundary layer.

Similar tests showed that a laminar boundary layer can be made
turbulent in a desired place by other measures, too (stepE in the
surface, considerable rougshnesse's, etc.), compare, for instance,
the tests with the profile number 32 of 6-percent thickn~ese with
di sturbanc e s. (See the appendi3E.) An increasee of the anglea of
attack beyond the angle of smooth inflow has the same effect In
obtaining for smaller Re a turbulent readhering on. the upper
side and, hence, a. smalle-r profile drag. (See profile-d~rag polars
of the propellor profile number 11. for smaller Re(ig11)

3. Laminar Profiles for Propellers

On the basis of the preliminary tests described above, propeller
profiles with highest possible lift-dra~g ratios were developed at
moderate ca and with small superstream velocities.' A few test
results on a propeller profile of 9-percen~t thickness (number 11)
are shown asE an example. (See fig. 11.)

Thbe corresponding profile coordinates may be seen from the
table of coordinates.

NAC.~, TM~ Ho. 1.~11

The investigatedi wrng was a rectanul-ar wing of b = 1.50! meters
span, t = 0.2'52 meter chiorrd auld F = 0.376 met er2 area of p~ro~jec~tion
(= area of ref'erence). The wingt ende, seen fro2m thle f-c0at, wefre
rounded semicircurlarly. (See, for instance, r~eferenlce 1El br 99

Moasu~rements.- (a) Defermination of:

A W = ~tjk
CR = -~-, CW = 0' m/ = TFt

for various ang~les of Itta1ck by meanlP of threep-co~mponent, measurements

(b) Momentum-measuirements ( tort al !1 PrEsse and.; ItatiS prE-ssur)
0.146 meter behind the vring in a wGinsj OScfion 0.22 mleter latterally
from the wi-ng plans of s~ymmetryI outside of the sulspensi-n fittings
for~ various Fle =--V7 ad c,

(c) Transition measurements with the stPthoscope 0.22 moeter
latera~lly from thle w-;ng plane of symnmetrjy for vzrlione T-e ulnd cat
(start of transition sad beginnings of the. fuDlly develo~ped tu~rb~l~ent
boundary loyer)

The chord of the ce~mber line was chosen ss line of refeirenc~e
for a. The noint of reference for cmt/ \lmkc, /, 0,a-incr'eas!ng
lies at a distance of t/k from~r the front. on` this Iline.

The jet corrections For the downwa~sh and the induxced drag for
the closed tunnel vere calculated according- to de H~aller- (reference 86~).

The momentsn measuremente were evaluatedl in the custolmiryl
manner. The local lift coefficient c, at the momentum measurirnl
station was put equa~l to 1.10ca*

A few test reen~lts can be seen f om figure 1.1: variation of
c, wi th cyat various E momentumm measurements), snd. the
beginning of the developed turbulent boundary layer on upper and
surfaces for vrarious ca aLnd Pe,

C"'/~(ii for Re = 0.76 x 100

NA4CA TM ~o 1181l

In an optimum eg-range, decreasing with increasing Re,
the transition on both wing surfaces occurs far toward the rear
which results in low profile drags. Wihen this range is exceeded
the drag increases and the transition polet on upper and, lower
sides travels rapidly forward, the reason is the appearance of a
suction peak at the wing nose due~ to the angle of attack. For
larger ca .a slight local turbulent separation occurs on the
upper side (tufts and stethoscope observations). cw decreases
with increasing Re somewhat more slowly than the laminar friction
of the flat plate. For larger Hte 'at modaera~te ca-values, more
favorable profile drag-lift ratios result. For smaller Re and.
smooth inflow ove deteriorates since the boundary layer of the
upper wt~ing surface undergoees laminar separation and does not
readhere with the properties of a turbulent boundary layer
(observation by stethoscopee. Only for larger c, the lamninar
boundary layer is disturbed so strongly by the incipient suction
peak at the wing nose that the transition occur to time .to bring:
about a turbulent readhering of the boundary layer for smaller
Re (obser-vation by stethoscope). cym then decreases with
increasing ca, 'and Cw~p lies at considerably larger ca than
would correspond to the smooth inflow. A simlar reduction of oz
vith increasing ca had resulted also in earlier measurementst on
ordinary profiles for smaller Re (Gijttinger Lieferungen I to IV
(reference 80), F. Schmitz (reference 831), NACA measurements
~(reference '82) etc.)

The lift and pitching moment distribution Cmtl shOW standard
behavior in the optimun c,-region. For larger angles of attack,
discontinuities appear in ca(&) the variations of c, with a
and %af (a.), due to the change of the effective profile camber
by the thickening of the boundary layer on the upper wing surface
which is caused by the forward shiiftcing of the transition point
after exceeding the optimum ca-region.

The use of laminar profiles for propellers reduces their
friction losses. Due to the eneller super~stream velocities, com-
pression shocks start at higher Mach numbers than for conventional
profiles. The use of laminar profiles for propellers will probably
rather lead to wider blades of smaller thickness ratio with
relatively low cacvalues under standard flight conditions. Bence,
there results again higher admissible Mach numbers and a larger
starting thrust.

NADA UK No. 1131

Ic. Lamninar Profiles for Wings

A basic requirement for ,pood wineT profiles is aL larcel ran.:P
,that is, low dral: under stWndard fli,'bt co0nditio~na and
highest poselble maximum lift with extended landing ailS. F1.rt~her
desired characteristics are small suTretream vel~ocit.ies !.n border
to attain high Mac~h numbers without compressio~n -ishcks7, lowr-divin-
moments8 Cmo, steady7 pitchinrj-m~oment beha~vior, i'avo~rable estair:
conditions for bend-inn und torsion of the vinr-, andi favorabDle
conditions of installation for landlingr flaps. Thepse la~tel Pre to
improve not, onily the landin; bu~t also the take-off considerably sn3
should, if possible, exte;nd, over the entire span.

Essed on these stipulations, a few laminar proFiles were
developed which7 are probably' appropriate or wings. The profile
shapes of Ireminar p;-ofiles of 10-Fsrcenl: anj. of h-p~ercent,
thickness can be seen From ftil~ ar l? n' the of coo~rdlinntes.

Profile data:

Thickness ratio d/t = 0.10 in 0.1;9 t f'rom theF fr.ont culrvatur1~e -ati-o
.of the mean line of the pr~ofile,
f/t = 0.006 in 0.5 t from thle front-
PFot = 0.009 (EO, = nose enlrva.tu~re ---dilus)
Chord t = 0.'71e meFter
d/t = 0.15 In 0.11h t fi-ro thle Ciroit.
f/t = 0.0265 in 0.!1 tI from the frocnt
Eo:t. = 0.01'7
t = 0.70 meter

The reear p-art o~f the4 profile iPrsainally converg~es into a
pointed trailin, edge:. Helncc, the pit~chinr-momr-nt behavior~ vas
improved as conpar~ed -\it~h lamlns:~ nrof~ilrjf wlth~ blunlt
edge, as tests of the --uthor on Isminor profbiles
trailln[;-elge angles had demonstratl-~'ed. The airofile o~f 1hr-per~lcent
thickness shove a~n S-shlape eomber line far tnward the rear whrich
made it possible to keep emo low;, without considerstle fo-rwiard
shifting of the transition point on the u!pper~ enrface.

%emi was determined for Variion1 Re b.-- mapns of" thle
momentum method. (See fiig. 7.) The creasuzre ;istributinns
p/qlo along the chord and the position of the t!-ansition poirlt can
be seen from figure9 L mad 5.

NACA TM N~o. 1181.

The pressure distributions of both profiles are flat.
Pressure increase aznd transition occur far toward the rear, thus
causing low oz For l.arger Re (>2 x 10 ) the drag. is increased
by the influence of the tunnel turbulence which is caused by the
more rapid forward moving of the tran~sition point due to an external
turbulence Ifig. 8). Verifying calculations of the profile drag for
;Re = 2 x 106 on laminar profiles of various thicknesses for the
positions of the transition point, which had been experimentalsly
determined, gave the following result The drag increase with the
profile thickness is primarily due to earlier transition for a
larger profile thickness, only seconda~trilyt~o higher form drag and
increased skin friction because of the higher superstream velocities
for thicker profiles.

In order to test suitable landing aids and ailerons for
Iamminar profiles, three-component, measuzrement-is were performed for
a laminar profile of 14 -percent thickness (fiP. 12) on a r~ectanglu~l~ar
wing of b = 1.50 meters span and t = 0.250 meter chord. (The
ving ends seen from the front were rounded off semicircularly.)
Figure 13 shows the landing aids investigated in retracted condition.
A Fowler flap C of a chord of 0.348 t which extended over the
entire span was used as landing aid. It extoras somewhat beyond
the main ving toward the rear~ and hence makes possible the attach-
ment of a trailing-edge aileron D of 0.125 t chord which also
extends over the entire span. For standard flight conditions,
this aileron serves as twist with resulti~n- mall flap control
moments whereas for low-speed-flight~t the extended Fowler? flaps
could assume the lateral control, perhzaps iti combination with the
trailing-edge aileronr. In this manner the Fowler flap may be
constructed Tx) as to extend over the entire span.

A frurther camax-increase for extended Fowler flaps coulld be
obtained by additional use of the trailin-edge aileron D), by
extending of slate B and. A at the Fowlelr fla.p C and at the
main profile. The retracted front eleat A would cause, under
standard flight conditions, an early turbulence of the boundiaryr
layer on the upper wing surface, in order to avoid this, the ala~t A
was built partly into the upper surface of the main wing and thus
a smooth surface obtained.

Momentum measurements showed that the profile Itree was not
measurably increased by the installation of landing aids at
Re = 1.07f x 106 when the slot between Fowler flap and trailing-edge
aileron was sealed. In the fucll-.acale model, probably the grooves
between main ving and the two slate also ought to be sealed.

The receding corner on the ubper wing surfacn between main
vinE and t~railing-edge a!.1eron also does no~t measurably increase
the drag~ in hig~h-spee? fli~ght (acPording to momentum measurements).


(Be = (0.70 to 0.75) x 106):



State c a,,

Fowler flap C extended, both slats A and E r~etrac ted (1)

BF = 520, D$ = 00, trailing-edge siler~on slot open 2.82
PF = 520, F, = 00, trailing-edge a!.leron slot closed 2.79
By = 4-FoI Cg = 310, trail~ing-edy~e alleranr slot open 3.12
By = 1;go, Pp = 300, trailing-edge alleto1n slot closely 2.885

Fowler flap C and Fowl!Er flap slat B extended, slat At at the
main v~ingr ret~rae cte

BF = 690 C = 00, 150, trailing-edge aileron slot. closed 3.h?
BF = (90?or, = 180, trailing.-edge alleron slot open 3.605
Fowler flap C and bothl elats A znd B extend ed

By = 90, Bg =00, 0, trailingi-edge aileron slot closed 3.93
Gy = 690, aq = 140, trailing-edge aileron slot open '6.05
Fowler flap C and frontal slat A extended, slab B retracted

By = 47", Fp = 00,s trailing-edge alleron slot closed 3.40
By 47, B = 0,trailing-edge aileron elot open 36
BF = 44", Dq = 180, trailing-edge aller~on slot closed 3.b9
By = 44o, Bg = 270, trairlin3-edile alleron alot open [=.683

NACAi TIM N\o. 1181

Ic m x

is referred to the wing chord For retracted laznding flalps.

The effect of the trailing-edge alleran of. 12.5-percent
t chord wras investigatted separately for~ retracted landing aida
(three-component and momentuml measurements for vari-ous control
surface Reflections P&).

ca and cmt/ (>0a-increasing) vere measured for various
a by means of three-component measurements, cw, by momentum
measurements 0.15 meter behind the wing at 0.11 meter lateral
distance from the wing plane of symmetry. The carnber line was
chosen as line of reference for a, the point of reference for
cm~ lies in t/4 from the front on this line. The local. lift
coefficient at the momentum test point was put equal to 1.10ca*
The Reynolds number in the three-component and momentum measure-

ments was .Pe = 0 .80 x 10 and 1.07 X 10 ,~ respectively.
The test, results can be seen fron figures 14 to 1.6.

Owing to deflection of the trailing-edge aileron extending
over the entire apan, a favorable envelope polar with. low profile
drag results for standard flight conditions in a considerable
ca-rang~~...~~ e (See fig. 14.)

The transition-point position in the optimum ca-rang~e is
only sightly shifted. forward by moderate deflections of t~his
nlasrrow control surf ace (observiations by stethoscope). At the
recedingc corner on the upper wing surface between usin ving and
control surface, a local separation on the upper control ourfa~ce
is avoided in the optimum ca-range up to pQ = 200 according to
observations by stethoscope and tuf~ts, hence, a low profile drag
for larger positive control surface deflections is attained.

For Bg = -5o and .-100 the boundary layer at tche nose of
the control surf ace on the lower wing surface was made artificially
tuzrbu-lent by providing a receding step (fig. 13) in order to avoid
a more extensive laminar separation on the bottom of the control
surface and to obtain a turbulent readhering.

Observations by stethoscope showed that the boiundary layer
of the lower wing surface for Pg 5 -50 underwent laminar
separation unless the bourndary layer was artificially. controlled,
and would not readhere completely turbulently. Corre sponding~ly,
there resulted (according to moments .measurements) relatively
large profile drag. By preov~iding the receding step F on the
lover surface 23 millitmeters ahead of the flap nose, t~he boundary

NACA rm No. 118i

NACA TM No. 1181

layer became turbulent at Re = 1.07 x 10 0 to 5 millimeters
behind the latter accordingq to stethossope observations) and
the profile dratg decreased momentumm meaeur~ements fig. 1L).

Thel variation of c, and c9/ versus a for various
control- surface deflections Oq can be seen From figures 15

Bdy deflection of the trailinq-eige aileron ca undergoes
a relatively considerable change which will probably cause
rolling moments sufficient for normal1-flight conditions. For
pQ = 00 the maximum lift without landing aids is cnmax = 1.06.
The anlee of attack~ at th-e stall is rather largor than for con-
ventional profiles of equal thickness. The lift increase is
normasl for the optimum ca-rainge and decreases sharply for larger
hence, nne can attain rather emll11er; posi-tive gust loads than for
earlier conventi-one.1 profiles of equal thick~ness and camber. The
zero moment for pg = 00 is emo! = -0.027 anld may be kept
arbitrarily small by small negative control surface deflections.
The pitching-moment distribution for moerats.le 09 probably will
be sufficiently uniform.

NADA TM No. 1181

0114PTERl 4


1. Aims for Further D~evelopment

The tests of chapter 3 showed that the profile drag oan be
conlsidersbly lowered by a suitable profile shape. With increasing
profile thiickness the drag will increase relatively strongly, the
main reason being the earlier transition for greater profile
thickness. Considerably lower dirsgs would be possible by main-
taining: the boundar~-y layer up to the trailing edge completely
Laminar. For larger Reyvnolds numbers very low drag would~ then
result even for thicker profiles. (See figs. 1 and 2.)

Thus the following tests were undertaken which aimed at the
development of thicker profiles where, for large Reynolds numbers
and in flight, the boun~day lawyer remains laminar up to the
trailing edge for a sufficient ca-range. In order to obtain
high Mach numbers without compression shocks, slight maxifman*
superstream velocities are desirable, that is, the pressure dis-
tribution is supposed to be uniform far toward the rear.

Bo~uncary-layer suction made it possible to maintain in the
present tests the boundary layer completely laminar up to the
trailing edge.

2. Effect of Suction of Laminar Bounuda~ry Layer

on the Flow Characteristics

The suction of a laminar boundary- layer has various effects:

(a) Aiugmnenation of the laminar pressure increase: According
to chapter 2, 1, the transition occur for a weak laminar pressure
increase caused by laminar separa-tijon of the boundary layer. By
'elimination of the greatly retari~e(" portion of the boundlaryr layer
in the neighborhood of the wall. with rising pressure, by means of
suction (through separate slots or by area suction) the laminar
separation and hence the transition could be avoided even, for
considerable pressure increases. A favorable eff~e is created
in many cases by the so-called sink effect: The, location of
suction acting as sink: produces in its neighborhoods an additional
pressure field with accelerated flow. Its superposition on the

fNCA TN N8o. 1181

external flow nt thre location of s~ct~ion results in a steeper
pressure increase fr~om th~e Pflay conditions immediately ahead of
the suction point to the stanPntion point T7ehind the suction slot.
Hence, the Icremainin pressur-e !ncrea~s along: the vall is
correepo3ndin:;l.y reduced.

(b) Ptta4inment of' higher Reyn~clde numbers Re Uot wi th
luminar~ bounda y layeFre: According to chapter 2, 2 influencee of
an external turbulence on the trans~ition) one obtains for a g~5iven
ext~ernal tu~crbulence withl a laminard bounda~ry layer, a mriaxium
critical ReSnolds numiber Re -~3 (U = velocityr
at the edge of the bou~ndiary layer.) Sinlce the increasing of 9
and T~ep, respectively, along t~he chord can be reduresd by to ndtary-
layer suction, it shold~l be 3oiesible to o~btEin higerPe-
for eunrl simissibl~e max~imum Heyncl~i: numbers Roy~critical

(c) Py toundal.: r~r r-3aler suctiojn and~ reaccelerationm of the suction
air to free-stirem vrelocity, a par't of the r~elatively large !;inetic
wake energy of ther 1mIT~nar tonn~ldary7 layer can be recoveredl anrd the
powr~'L realuir)edj for peopu~.leon can thus bre redu~ced:. Professor Ack~eret
(re~fFerenCe 65) was thle first~ to poinlt out the possibility of
redLucing the powder Yrequired for propujlsion of' airplan~es byr ut~iliza-
tion of the kinetic vsklc energy.

(8l) Withi the -aid of bountairy-lay-er srlcticn the boundary layer
could be mlaintrme~d completel:' l-amirlar for a larpier rangeP of angqle
of attack in spit~e of thes ooccurrinE greaste pressure increases
which wroulr! ext'end thle ca-r'ane- withl low drag.

?. Statement of the Problem

AC boundary 17.yer on thiicker profiles for larger Reyjnolkl
nurmber kept crmp!.s.tely lamainar with the sid! of bounitary-lsyer
snction is equlivslent-l- to th~e maintenance of laminar bolundar-y layer
forhig ew-ithn sim~t~tneous ct~ang pressure increase. In
o!rler' to obtain lowest. noscsibl.e drass +he losses in the suction-
slote lulst be redulce~. The boun".ai-ry lyer continuing behind the
su~ctionn slots is not to be addiitionally difsturbed by the suction

Besiidee, the customa:7r stipullations for vings must be observed.
(See cha~pter 3 and 6.)

NAGA TM No. 1181

A wing with laminar boundary-layer suction of equal strength
and rigidity should not became much heavier that customary vings
without suction.

In carrying out the tests the problems connected with laminar
boundary-layer suction were at first investigated separately,
step by step, and then gradually combined. The test sequence was
as follows:

1. Test; with laminar bounda~ry-layer auction on a lightly
cambered profile of 6.75-percent thickness with a stagle suction
slot in 77 percent t from the leading edge on the upper surface to
investigate the basic aptitude of boun~dary-layer suction for
maintaining a boundaury-layer lamninar (chapter 4, 4).

2. Study of the laminar pressure increase with boundary-
layer suction in separate slots for smaller to medium Reynolde
numiber~s Re (chapter 5).

3. More detailed investigation of the flow in the suction
slots for laminarr boundary-layer suction (alot losses, sink
effect, slot flow, etc.) (See' chapters 5 and 6.)

4. Test for obtaining higher Reynolds neaterps with laminar
boundary layers with the id. of bound~ary-layer suction for
intentionally small external pressure increase and normal vind-
tunnel turbulence (chapter 7, Tests on a T~hin Symmastrical Profile
with Suction).

5. Investigations of the lamninar boundar~y-layer suction
for larger pressure increase and higher Reynolds numbers on a
slightly cambered profile of 10*5-percent thickness with con-
ventional thickness distribution (without extended flat pressure
distribution ahead of the pressure rise). (See chapter 8.)

4. History of Development of the Laminar

Boundagr-Layer Suction

(a) In 1928 the assumption was expressed for the first time,
by B. M. Jones, that a boundary layer Edgh~t perhaps be maintained
laminar for a layger time with the aid of boundary-layer suction
which would reduce the frictional drag.

(b) L. Premit1 calculated the laminar boundary-layer
development with suction for pressure increase, under the
supposition APohlhause, = -12 (reference 72, pp. 117 and 118).

N10CA TM No. 1181

(c) laminar anrction tests on a alightly: cam'bered profile of
6.75-percen~t thickntress. The first, test, mrade for the purpose of
orientation, on a El11 :htly cambered profile of" 6.75-percent. thick-
ness with a single suction slot in 77 percent t. from the leading
ed3e on the upper su-rface (autumn 1960 anl winter 1900-b1) showed
that it 1 i)spssible, by boundary-layer Puct~ion, to mainta-ln thie
boundrliy Isfer comaplet~ely- leminar ani! to obta!.n low profile drags.

The investigated profile (2) with the suction slot can be
seen from figure 17. The wing was nounlted between end dislks in
the vind tunnel of the Inst~itote.

Ch~or1 t = 0.451 meter

Profl~e thickn~ess Z/t = 0.0675 in 0.306 t from
the leading edge

Nose cu~rvature radius Roft = 0.0035

Slot in 0.77 t f'rom th~e l.eadinc; edgie on the upper surface,
mirtumnr slot wiltlh s = 1.3 m-illimeters and a = 0.3 madllimeiter,
slot directed tor-rard the ~ear at arl angle of L53 to the surface.

Thn suction slot wass developed as a diffuser with an eo opening
angle in order to conv vt part of the kinetic energy~ of the
sucked air into pr~essurei hnd bence obtain smaller negative pressures
in the suction chamber and! smaller pressures and power recuire-
ments ior~ thle evation blo-:er. Thle development of the suction! 910%
as a dli."user f'o. thle purposes of a Dal-tial conversion of the
veloc-itl- enery of the suck~ed air into pre~ssure was for the first
time ,cucesefu~llyr applied. in tur~bulent suction tescs byv A. GerbDer
(reference 69?) on a suggcestioni by Professor Ackeret.

The sl.ot inlet was desilned on the basis of a thects by
H. Bleuler, inlveutiigat~i a~ JTy Poe3sso Acker.:t. The: old-f ilow was
treated by H. Blenller a,?s a ree j~et with constant pressu:e naOlong
the jiet edge, uner the assumption of frict-ionless flow.


DraiT measulrements:

t wing chord (0.lrj1 m)

NACA TM No. 1181

3 slot length (0.324 m)

F area of reference (t3;) (0.166 m2)

a minimum alot width

go undisturbed total. pressure outside of boundary layer or
wake, measured. with atmosphere as reference level

po static pressure in center of side wall of tjind tunnel at
begkibnitng of' test section, measured with atmosphere as
reference level

go free-streaml dynamic pressure, kilograme per meter2 (go Po)

Go free-streamn velocity meters per second g

p, static pressure in center of suction chamber, measured
compared with 'po, kilograms per meters!

D;PE suction blover pressure, kilogr~ans per meter2

UbL mean velocity of sucked air perpendicular to wilng in suction
chamber at location of static pressure measurement
= ~Suction quantity a h 81-
=~ ~ ~ o atths16ato
Cross section of the suction chamber

uA 'exit velocity of sucked air rearward in free-st~ream direction,
meters per second

Qa suction-quantity (m3/a) measured by calibrated ourturi nozzle
of 17.4millim~eters at narrowest crose section and
28.0 millimeters P ahead of nozzle'

6" displacement thickness ahead of- sulction slot

QgiY US*3 w~here U = velocity at edge of boundary layer ahead
of slot

Re Reynolds nurmber -


total profile dlrag decisive for propulsion

dlrag contribution of wake




NACA TM No. 1181 ?1

Vg drag cont;rution of sulction blower
Dimensionless coefficients:

cQ suction-quan~tity coefficient

ePG suction-blower pressure coefficient (;9

CP = --

e~ coeffid~ent of total profile drsg --

c coefficient of *1r-acr contributions of waske (det~ermined4 by

c, coeff'icient of drag conrtributi;onl of suction blower -F

Me asursments : Thle asttic pressure p, was measured, for
varliousB Xe = a nd suction quantities ain th~e center of
the anlcticn chlamber withini 0.S mill.?meter pressures holes).
Momentum measurements were performed 0.173 meter behind the wing.
The slolt vii~th a wje 1.3 millimeters and~ 0.1? millimeter. The
state of the bou~ndairry laer on thle upper enr~fsce behind the slot
was verifiedl with a hlot vilre (references 21 and 61b), by soot
coatiny and acousticallyv by stethoscope (reference 61c).

Dra~g evallatiojn for" Eondasrry-12er suction (compare also
0. Schrenk (refETPHOO 75)).

Let the sulcked air b~e Recelerated, rearward in flight, direction
to the vselocity uA The motive power L, of thle suction blower
(efficiency q ) then is

L,"A = au2Pa uL2j


A 2 Pa L~ = Suction-blover pressure

NACA TM No. 1181

The kinetic energy BuL2 of the sucked air in the suction chamber
may still be utilized. Pressure losses in the suction ducts were

When uA Uo> the sucked air fatr behind the wing has 'another
momentum thanr far toward the front; the change of momentum is

~J = pBa(Uo uA_

(for uA < Do, llJ signifies drag, for uA> o thrust.

The propulsive power requirement for overcoming the drag
contribution of the vake V,' and P~a(Uo uA), will be, if a
propeller in flight or installed in a tunnel with the efficiency
rip is provided:

/rW I 09a(",Uo A o

The entire power requirements; for overcoming of the profile
drag becomes:

ae u2 Pl "L2) '+DaU AU
L = Lg + L~P = -43~ LIU--
rg 3p

Depending an the exit velocity uA of the sucked air, L is
distributed unequally between Lg and Lp. The minimum total
power requirement min Results for =U O, qg and ri may,
in the general case, also be functions of uA2. It is now assumed
that ?g and rip are independent of uA. Then there becomes for

aL up o:
=U 0:- =-
EE~ 'o n

NACIA TM No. 1181

~e-t fulrthermnore an equal effi;cienrcy of propeller end sulctiocn blower
be rssumed: 'I = 'Ip = ?.

For Lmin thin uA = Uo and

UfT 2 p -~ 'DL o Uo

that is,

Zo aTV pL o x e

where aP

by = 2 _pa 9uL= Sction-blo~wer pressure for acceleration
a L of the sucked air to Uo

Under the 9ecumptions given above (equal efficiency of suction
bloveVi and mropell-er, a~ccelerstion of the suckiedl air to Uo toward
the rear) one may calculate for the d-aq evaluation precisely as
If the suction blower hail a 100-percent effici~ency~. The total
d~rag rccefficienzt decisive for the propulsion becomes:

cy -p + cw = eqcp + cyo = cW + C
O: q,F UoaoF g

The IraFg mensurements witih bounlrdaly-layer suction were eval~u~iate
accordilng to the alCvIE formu3as rUnfer the alsslumption that propeller
and suction blower havre eqadl effPiciencyr and the sucked air is
accelerated rearward in fli~rht direction to iio. For. comnarison,
the minimum profile d~rag Withzolt s1ction (suct~ion slot closed) of
the profile (2), d/t = 0.067 and the profile (1), d/t 0.09 was
measured by means of the momen~turllmethod.

NACA TM No. 1181

The measurement results can be seen from figure 18 to 29.

c'"pt(Re) vith a = 049 net Fig. 18
c, (Re) without suction

owmi c ) as 1.3 mm FigT. 19

c, (cq Fig. 20

CoptCp(Re)(e a = 0~ ms-I.9 at and~ Fig. 21

2 hot-wire photographs with and without suction (f-igs. 22
and 23).

2 photographs with soot coating (1aminar and turbulent)
(figs. 24 and 25).

~By boundary-layer` suction it was possible to maintain, for

Reynolds numbers up to Re = = 800,000 on upper and lower
surface a completely laminar boundary.1a~yer (hot-vire photographs,
fig~s. 22~ and 23, anrd photographs, figs. 24 and? 25 with soot
coating, confirmation by stethoscope observations). Generally,
only weak and slow laminar velocity fluctuations were ascertainable
behind the suction alot.

The laminar pressure increase on the upper surface from the
pressure minimum to the trai~ling edge~ amounted to 35 percent of
the pressure difference between stagnation point and~pressure

The lowest total drag was, with a = 0.9-millimeter slot
width for .Re = 0,9 x 10 : c, = 0,0035 (power reanuired for
suction included). For smaller R~e, cz increased somewhat
more slowly than the lamilnar friction of the flat plate. For
larger Re, c, Increases a~a~in since, owing to the influence of
the tunnel turbulence, the treasitions on the upper side occur
already ahead of the suction slot (as was shown by stethoscope
and soot-coaiting observations). Hence, the boundary layer remain

NACA. TM~ No. 1181

turbullent behind the clot in spite of increasedf suction. The dlral7
is definitely larger ri~thout. suction thanl with it.

The action quantities cg required. for keeping the
boundr~cs~y player laminar wrere veryr unall, particularly for a elot
width s =0.9 millineteri in general only ;I fraction of the
displscemnent th~ickness 5 hadr to be sulcked off the laminar
boundary layer ahead of +.he slot (15 percent to 50 percent for
Re = 0.2 x 10~ to 0.9 x 106). (See fis. 21 Coopt(Re) for
a = 0.a nmm and a = 1.3 mrm.) The required minimm'm suction quantity
for a suction slot of the width a = 1.3 mllimaters was larger
than the one for a = 0.9 millimeter (fif. 21). If the suction is
too weak, a l~cal leminar separlation occurs at the suction point
and the: boundary layer behind the elot blecomess rapidly turbulent
(alcco-cdingF to observations by stethoscope, soo~t coating, srnd hot
vire). Corr~esonlingly, for very welk suction Q increases
again (fig.S 19). For an optimum3 suction qua~nti~tyr eq results
the lovent. drse c"-ln ,cob ~ increases wlith increasing Reynolds
nulmber (fig. 21). For eg > "crnt the sk-in friction increases
behind the slot thinnerr b7unilary layer), iience, c,, inc reases
with eq.

In 17encral, thekr corresponding negativer suction pressures
c~ap in the suction chamber were small (fli 21) although~, for
laminar nvotion, the conversion of the velocity energy of the
sulcke! ir into pressure in the slot, dliffuser vae not particularly
good (ccmpare,1a~ter tests).

(d) In connection w~ith these lamninar suction tests with a.
single sl.ot, H'. Ras perforzed!, on the suggast~ion of Professor Ackeret,
in the Institute for Aerodynamics, Zurich, tests with laminar
bounrlary suxction bly means of a sort of .nrea suction consisting of
35 nParrow slots arrangfed one after the other. Wlith this area
suction one obt~ained. latinarP nrecsure increases of about 5= percent
to 5'i percent of +he pressurep difference between sta nation point
anl p-esE~ure minimum for smaller Rcynolds numbers preferencess 66,
67, and 683).

(e) H. Schslichting calculated thse lamin-r boundary-layer
development. on, a flat, plate wi~th area Puction for constant suction
intensity.: (reference 70). Furt~herlmore, Schllichting calculated
recenntly the leominar touxndary-layer development wsith area suction
on a Joukowsky profile (referencei 71).

NACA TM No. 1181





1. Laminar Suction Tests with Three Suction Slots

Arranged One After the Other

Tlhe suction tests with the 1aminar profile of 6.75-percenlt
thickness fo~r Re = 0.8 x 106, showed a lam~inar pressure increase
of 35 percent with a single suction slot. With area suction a
laminar pressure increase of 53 percent to 55 percent WasE attained.
More laminar suction tests were performed, with three suction slot
lying one after the other in the suction tunnel used by M. Bas
with the aim to obtain larger laeniar pressure increases with
relatively few lots.

Definitions for the suction tests with single slot in the .
suction tunnel (chapter 5, 1, chapter 5, 2b, 2c, and 2d,
chapter 6, 1 and 6, 2).

q dynamic pressure at edge of boundary layer

qm maximum dynamic pressure at edget of boundary layer at narrovrent
section of tunnel

Ap static pirefssure` at test plate or at1 walls of slot diffuser,
meat;scu red th 0.5 :nilliverter bore holes with static
r"reasure at test plate as- narr'owest place of tu~nnl as
refernce level

PA stat~iSc pressure in suction chamber, measured with 0.5 millimeter
bore~r holes with static pressure at test plate at
narrowsetfQ plaice of tunnel as reference level

t slot length (0.40 m)

s minimum slot width

8~ displacement thick~ness ahead of slot (8* vas caleviasted cvery
time from measurred pressure distribution for mean suction
quantity according to Poblhausen (reference Ic?)).

QgF = UB 3~, with U = velocity at edge of toundar-y layer ahead of
suction slot

h displacement of slot trai'ling edge with respect to slot inlet,
h > 0 for inward displacement

Qa suetion quant~ity measured with calibrated measurincgc nozzles

V, velocity of sucked air at anc' of slot diffuser, determined
from difference of total gat alot exit measuredd with
1.0 mm total) hlead tube) andi static pressure pA In
suction chamber in neighboarhoodc of test, point for velocity
distri~buction. (The test ar.angement canl be seen from

The flet test plate was provided with three suction slots
the shape of which can? be seen from figure 27.

The slots were perpendicular~ to thle surface and. were developed
as diffusers wIct~h emai1 opening angle (alot shape a) Ey adljust-
ment of the opposite wall NoI. V the i.nterna ~lvidth of the tunnlel
and the pressurre distribution at the test plate could be changed.
The boundary player of the opposite wall also wasB kept. laminar by
suction. The vriath of the test plate and the slot length were
400 millimrE~ters.

The pressure distribu~tion along the test, plate was determined
for various suction ausntities n_ ani tunnel w~ilths. Th~e state
of the bo~undary layer behind the slot wais ver~ified b:' hot vire and
stethoscope. The test results can be seen from figures 28asnd 29.

A9t the arrow behind. the third slot the boun".ar3 layer was
eti11 laminart. Consil.era-blyr larf~er lamninar pressure increases
resulted with boundary-le.yer suction than wlithout it (mnaxinma
63 percent withI 40 mm tun~nel. width).

The sink effect makes an essentia3. contribution to the
lacminar ~r~essure increase, particularly for stronger externs1
pressure increase. The signlificanlce of the sink effect for
turbulent boundaryr-layer suction was pointed out repeatedly
(Ackeret, Frandtl, 0. Schrenk, Geber (references 69, 72, 73,
74, and 75)).

NACA TM No. 1181l

NACA TM No. 1181

2. Tests w~ith Lamninar Bound~ary-Layer Suction

with a Single Suction Slot

The purpose of these testss was a more detailed study of the
flow phenomena in the neighborhood of a suction elot for laminar
boundary-layer' suction (sink effect, laminar pressure increase
down to the transition; flow in the suction slot and pressure
losses in the slot).

(a) Testsi with lamainar botundary-layer suction in the water
tank.- (a) Tests on a single slot: A few lamninar suction tests
for the purpose of orientation were performed with various alot
shapes. The suction slot was placed on a syrmmetrica~l profile of
LO-percent thickness (at 0.43 t from the' leading edge) and t = 1.21 m
chord at a distance of 0.71 m from the leading edge. The slot
wats d~eveloped as diffuser with small opening anglZe.

SThe model was toured through the water with a velocity of
0.1 to 0.2 mater per second. The flow in the region of the suction
slot vae made visible by sprinkled-on aluminum powder and
pho~tographed. (Compatre the flow pictures figs. 30 to 34.)

In general, the larminar boundary-layer suction operated
faultlessly even for very different slot shapes and suction of
various strength. The laminar bouniaryr layer, continuing behind
the alot to which no suction was' applied, was mostly undisturbed
by the suction. The flow in the suction alot separated on one
side of the slot.

With increasing suction quantity the sjlot must be made wider
in order to avoid high velocities at the slot exit and corre-
apondingly large slot losses.

For analler suction quantities the slot width must be
reduced since otherwise a local laminar separation occur~sat the
slot inlet (fig. 31(g)). The tests with the suction wing of
6j.75-percen~t thickness (chapter 4, 4) and later laminar suction
tests showed that in such cases the boundary layer behind the
slot became~ rapidly turbulent (obse~rvation by stethoscope).~ Due
to outward .curvature of the profile surface ahead of the suction
slot (fig. 31) and a stronger local pressure increase ahead of the
slot, a laminar separation may occur for weak suction; thereby
the boundary- layer behind thie alot also becomes rapidly turbulent.
Forwaridccurved suction lots (figs; 30)(e), 30(f), 31(a), 31(b),
and 31(c.)) are especially ~sensitive in this respect.

NACA TM No. 1181

In a few cases the slot trailing edge was shifted various
dlegrees towa 4 the outside and the inside, respectively. (See
fjes. 30(c), 31(d), and J1(e).)

A too great inward shifting may eventually also lead to a
local l-aminar separation at the slot inlet (fig. 38) and thus
render the Ilaminar suction ineffective. (See later testsl chapter 5, 2
test F vith slot (b) vith h = 1.3 rmm inward

((3) A test with several narrow suction slots placed one
after another (a sort of area suction) shows how the lsaminar
boundary layer oozes into the winfG interior (ficg. $)

(r) On as thin symmetrical profile, elc~tioan was applie4 to
the laminar bound~ry l7ayer at the traziling edge of the wlng:.
The first suction tests wP1re perfolLirme without a partitiojn vaall
in the suction slot and (-ave a neC;ative result: for star~tinn
conditions a stanat~ion point orig~inated in the free stream behind
the suction point which ~traveled~ at the slightest d.1sturbance
toward one or the other wing slde. (See ficg. 373(b).)

By means of a naartition wsll in the clot, the rear stagnation
point was fixed on that Wll.1 (See figs. 33(a) anda 33(c).)

The boundary la.:er behind the slot, was mostly very thin.

By rotation of the part~itioning metal sheet and by suction
of different magn~itude on upper and lower surface one may shift
the rear stagnation point and thus (as remarked by Professor Ackeret)
change the circulation around the wing.

The laminar suction of the tralling edgoe is very sensitive
to the shanP of the win surface shortly- thead of thre slot., For
a slot inlet whlich is too around the boundalry layer shecrdr of the
slot undergoes a laminar separation, if thre inlet is too pointed:
there result lasrge negat~ive pressures at, thle slot inlet, high
velocities in the suction slot, and large slot losses.

(8) Another test with laminar boundary-leiyer solution wa~s
performed in the water tank with a slotted flap wing of chordl
t = 0.61 m (dlt, = 0.15:7). The boundary layer could be maintained
completely lamirnar by means of suction on main wing and flap for
various flap deflections (fig. 3k). The Reynoldls number was
Ret 2 100,000 to 120,000.

NA0A TM NTo. 1181

(b) Laminar suction tests with the slot (a)l see flig. 27
and for definitions see beginning3 of chapter 5).- Vith the test
apparatus used for the laminar suction tests with three slots
(chapter 5, 1), laminar su~ctiodi tests with only the suction
slot farthest to the front were performed (alot shape(a)). The
test plate was plane. Suction was applied over the width of the
tunnel of 400 millimeters.

The pressure distribution on the test plate was measured,
with the minimum alot width s, suction quantity Qa, tunnel
vidth and free-stream velocity being varied. The trailing edge
of the slot could be adjusted at different heights with respect to
the slot inlet(displacements h & 0 ,for' inward shifted trailing
edge of the slot).

The influence of the suction quantities Qg and QaS"
respectively, (8n = displacement thickness ahead of the alot)
of the slot vidth and of the shifting of the slot trailing edge
on the sink effect were investigated. 3Figures 35 and 36 show
a few p-essure distributions on the test plate in the region of
the suction, elot. In the investigated cases, the boundary layer
remained laminar -behind the slot accordingg to stethoscope

The laminar pressure increase due _to sink effect increase
with growing suction quantity Qa and increasing inward shifting
of the slot trailing edge and vice versa, whereas a change in the
slot width produces practically no effect. With increasing elot
vidth, stronger suction must be applieda in order to avoid a
local lam~inar separation at the suction point and a rapidly
becoming turbulence of the boundary layer behind the alot accordingg
to stethoscope observations).

The measured laminar pressure increase due tx> ainkr effect is
considerably larger than it would be under the assumption of
frictionless flow by the boutndary-layer suction the displace-
ment thickness 83 becomes suddenly smaller behind the, auction
slot, and, correspordndingy, the effective surface (at the distance 8"
fran the wall) is s~hilfted nearer toward the wall. With an
effective surface of such vavy development a stronger local
pressure increase would result as shown in respective calculations
by A. Betz (reference 77) and pressure-di stribution measurements
at transition due to external pressure increase, where the
displacement thickn~ess also decreases suddenly. (See figs. 3, 4,
5, etc'.) By superposition of a stak at the suction point, there
originates a considerably larger pressure Increase due to sink
effect than for frictionless flow. By inward or outward shifting

NACA TM No. 1181

of the trailing edge of the slot this effect may be still increased
or ?ecreased.

In further measurements (c) the attempt was: made to increase
the laminar pressure increase, due to sink effect, attl1 further by
wavy develounment of the wall in the neighborhood of thep slucion
point, particularly for larger suction quantittee 0a/93** It, is
true that the required minimum suction quantities, far which the
boundary layer behind the suction slot still remains completelyi
1Evninarr, wl~l probablyr increase: the Plow~est, baundaery layer pzrts
to which no su~ctijon wae applied can be ret~ardedl at thea meet to the
velocityr zero at the stag~nat4on POinTt beinrl. the slot. o lre
external pr1esgsure increase (_Por instance, due to estrolnerS sink
effect because of~ waviness of the su~rfie.ce ait the suction slot) and
for smaller suction quanTtitieA sometimes a strongc local nressure
increase and a rapid transition of thle bointcry player ocenars
behind +.he slot. (See later mleasurements: fic. 39, test 8A, fig. Ir0,
test 1009

Ic) Investgaio of the suction elot (b~) wifth the test -Mate
(b) for lamninar boundryss-lsyer suzctfon.- The form of the test
plate (b andi the suction slot (b)~ c(;n be sean from~ figurer 37
and 38. (See betginnines of chapter T for deafinitions.)

The slot (b) was perpendicular to the surface Ilnd was also
design~ed as diffuser wit~h mall opening angle. T~e eIxt~ernal
presecroc Sistribu~tion could be varied by sdju.s'tment of the opposite
wall V, the bou~ndary lur-er olf which was msintainerl laminar by
suction. To the test plate, suction was applied over the tunnel
width? of 400 millimeters.

The pressure dliotribution at. the test, plate and thle transition-
point position were determined for various suction quan~tites Q
and a. S, repectivelyr, fo~r various. free-str~eam velocities and
tunnel vridths.

The trans~ition-~o~int position evidenced by the break in the
pressure-difstribution curve which~ in turn was cus~ied by th~e sudden
reduction in displacement thick~ness due to the t-ransition, was
also determined from observations by ste~thoscope.

At the transition, considerable fluctu.ations9 in static pressure
could be ascertained byv means of a sensitive mancometer.

The results of the pressure distribution and transition maeasure-
ments can be seen from the figures 39 t-o rc4..

NACA rm No. 1181.

The experiences described- in~the last section with regard to
sink effect were confirmed:' increase of the laminar pressure
increase, due to sink effect, with growing su~ction quantity Q,
and Qg/Qae, respectively, and with increasing inward shifting
of the trailing edge of the slot, and vice versa. An excessive
inward sh~ifti~ng (h = 1.3 mm) did not augmen~t the stik effect any

Higher laminar pressure increases due to sinkr effect result
for larger external pressure increase (40-m tunnel width) than
for 80-millimeter tunnel. wi~dthj however, the minimon suction
quantities Q&/Q 6" were increased.

With increastag Reynolds -number, the pressure increase due
to sink effect decreases allthtly aind vice versa, for equal
suction quantity Qg/QsBM (compare fig~s. 40 and 4l).

9, =16.3 ke/m2 ~8.15 kg/'m2
m" = 32.6 kg/m2! Fi8* 39, 4-m =' 6l~.28kgm2
The sink effect is increased by the wvariness of the surface in
the region of the suction slot. Of course, as was to be expected,
larger mdnimu suction quantities Q./Qgw are required to maintain
a laminar boundary layer behind the slot.

The total lamina~r pressure increase up to thze tranzsition
point is, generally, considerably larger than without anotion.
It increases with growing suction quantity as well as with rising
external pressure increase (comparison of the tests with 40-mmn
and 80-mm tunnel vidth).

For equal Qa 8i, the transition occurs earlier with growing
Reynolds number, similar to the case without boundary-la~yer
auction comparee fig. 40 and 4l)

m = 16.3 k9/ n and 32.6 kg/m2
also figure 39

16. kej

NACA TM THo. c1181

By inwJard shifting of the trailing edge of the slot (stronger
sink effect), the laminar pressure increase before the transition
is slightly suce~nted and vice versa, under the assumption of
eaurzl suction pover.

h =-0.3 see .fi-. )r2
h = -0.3 see fie 40, 43
h = -0.1 see figs. L0, L?

Tests lb' with h = 0.9 to
Tests 17 vith h = 0.9 and
Tests 183 with h = 0.9 and

An excessive inward. shift~ng of the t~railinge edge of the slot
(h = 1.3) deteriorat~es the larminar pressure increase again
(figs. L2 and 4k).

Incrleasing of th~e sink: effect -nd the lamin~ar pregssure increase
by a slight~ly wavy formation of the surface in the region of thle
suction slot, similar to the form of the test plate (b), .As n~eeful
probably only for larger suction quanti~tie.s Qa Qge; it ie lose so
for smaller ones.

(d) Lamninar suction tests with slot (b) and~ test plate(d)
(See fii 37'.).- By Sncreasing the vaviness in the regionn of the
suctcion slot (b) (Fsee form olf the test plate (d)) thie Einkr effect
and the lamuinar pressure increase before the transition was7~
augmner.ed still further for stronger slct~ion; see the pr98ssuIre
distributiion at the test plate (fig. 4T-).

Of cou-rse, still 'larg-er minimum sulction quantities than for
slot (t) wit~h t~est plate (b) are required in order to kreep thle
boundary layer behind the slot completely laminar.

Further laminar suction -Fests! with thle single slot (g)
(fig. 46) gave similar results (fig. L7).

The suction tests wilth the plate (I') showed that a boundaryT
layer for a flow around a s~lightl.,i protrudling: corner may be maintained
laminar by means of' the boundary-layer suction, if a suction slot
is placed in the corner.

44 NACA UK~ No. 1181

.Sammary Regarding Laminarr Pressure .Increasle with Boundary

Layer Suotion for Small up to Medium Reynolde Numbers

For small up to medium Re (N = 1/3 R a2 ahead of the slot
varied between 0.4 and 0.8 x 106) leninar boundary-layer suction
makes high3 laminar pressure increases with relatively few suction
slots possible; generally only a fraction of the respective
displacement thickness 8* ahead of the .slots must be sucked off.

In most cases the sink effect makes a considerable contribution
toward the laminar pressure increase.. For larger suction quantities
Qg/Q6+, mne augmenting .of the pressure increase due to sink effect
by suitable shaping of the surface is useful; it is less so for
small ones.

NACA TM No. 1181




1. Laminar Suction Tests with Straig~ht Suction Slot

The test apparatus for the tests of chapter 6 was the same
as for the tested of chapter 5.

(a) Tests with slot (a) (see fly. 27 1Jnd for definitions see
beginning of chantter 5).- Thle test plate was plane. Suction was
aPplied over the tunnel width of h00 millimeters.

Measurement+s: The static pre~ssure in the sulction tank was
measured for various suction quantities Qa and Qg/Qge, rsetie
(6" = displacement thickn~ess ahead of the slot), slot widths
and s~agna~tion pressures or Peynold~s numbers, respectively. Further-
more, the trailing edps of the slot was made to sh~ift to various
extents ou~twardr or invard~ with respect, to thle slo't inlet. The
distribution of the st~at?.c pressure in the suction chamber with
the varying suction qluantity can be seen'from filgulres lr3 to 53.

WJith growing suction quantity, the negative pressure in the
suction tank incr1esae. For larger Re, the suction qulantity Qg/QaY
being equal, they are smaller, the same holds true for vider suction

If the trailingp edge of the slot is shifted inward, larger
negative pressures result ait the slot intet and In the suction
tank du~e to the stronper sink effect. Invard or out'. of the trailing ed-;e of thle slot causes, conditions otherwise
being equal, only unessential changes in the pressure difference
between suction chamber and the place directly ahead of thle suction

In order to obtain a uniform sulction along the span, the slot .
width and the' extent of the inward or ou~tward shift of the trailing
edge of the slot along the span must, as far as possible, remain
constant. If the trailing edcre of the slot is somewhere shifted
further inward, less air is sucked there', sometimes this fact may
cause a local laminar separation shortly ahead of the slot; also,
the boundary layer behind the slot may become rapidly turbulent
at that location, as vae shown in observations by stethoscope

NACA TM No. 1181

(confirmed in the lasilnar suction tests of chapters 7 and 8). From
the point of transition a turbulent wedge spreads toward the rear
in the usual. manner. If the slot is lightly narrower in some place,
less bouxndary-lay~er air is sucked there. The laminar boundary layer
behind the elot then thickens; sometimes an earlier turbulence" may
occur, unless the preceding or the following 'alot is widened

(b) Tests with slots (b), test plate (b) (fins. 37 and 38).-
The static pressures in the suction tank and along the slot di~ffuser
were measured, together with the velocity3 distribution at the elot,
exit for various suction quantities, slot widths, and stagnation

Other conditions being equal, practically thle same pressure
differences between suction chamber and the place directly ahead of
the slot resulted as for the slot (a).

By lenTt~hening the alot diffuser from 16j millimeters to 24 milli-
meters, the negative pressure in the suction -tank was slight reduced
and the pressure increase in the slot diffuser slightly augmented
(fig. 544 see also ;later tests wi~th suction slot curve rearward,
chapter b, 3).

For a clot wilth of a = 0.8 mm ant qm = 1B.'J kgp/m2, a wJeak
uniform pressure increase occurs in the slot diffuser which
increases with growing suction quantity (fic. 54).

The velocity distributions at t~he slot exit for slot width
a = 0.83 mm, qm = 8.15 kg/:m2, 16.3 kgf/m 32.6 kg/m2, and various
suction quantities can be seen from figures 55 to 5,7.

For wesk suction the velocity distribution at the slot exit
is laminar wi~th a weak reverse flow on the front side of the slot.
For stronger suction the slot flow separates on the rear side of
the slot and the suction air flows, at a relatively, hith velocity,
Into the suction tank near the front side of the slot. Accordingly,
the conversion of the kinetic energy of the suction air into pressure
Is not particularly favorable in the slot diffuser (see figure:
Distribution of the static pressure in the slot diffuser). The
minimum slot losses result for small suction quantities Qga/Qg <,
suction is applied only to the innermost lowest parts of the
boundary layer.

The conversion of the velocityr enercy of the suction air into
pressure in the slot diffuser would improve if the elot flow would
e~his fact vaEa, in cases of inaccuratelr alijustedi suction slots,
established by observation by stethoscope repea~tedlyg.

NACA TM1 Ho. 11:71

readhere turbullentlyr before the end of the slot and the velocity
di~stribut-ionx at the slot ozit v~.:ali teac~me uifkt'..rm oce aia
suction tests vilth the slot (b) .Rhowred that this aim wasf a~ttainedt
byr incr~eset. slot~ wi:th, larger auction quahntities, and sag.~nation
pressures, t~h7t. is, fo?' larger Reyvnolds numbers, referred. to the
slot flow (lieste K, L, a = 1.3-ann slot vrid+h, q, = 32.5 k,8/m)

The velocilty distribution at the slot exit '.:as for s = 1.1 rmm
andi qm = !2.5 k/m? relat~ivela unitform. (See flr;. 5".)

The transition-point position behind the slot shows that the
boundlar; layeuahead o" nd behind the slot w~as lan~ina~r.

The pressure distribution alon~ the clo~ t 11ffuer~ (fic~-. 5o)
sho~w et: the presuxmable tleransitlon point of the slot flJOv a7 rsnidl
pressure increas-e an' st~zron fluctcuatilen in the static pressure,
similairly to: the conditions fou~nd at the treasition of a laminar'
boundary layer for increrrsing pressure.

For 3 = 1.3-nalII slot wid~thl and m km the slot
losses and ne68tive oressures in the Fuction tank are all~ht.
(See fiS. 60.)

Slummr~y Pe -ardintg the L~osses in the Straiht .Syueion Slnt-

for Lamlnar Bo3undasry Lg)yer Suction

Desirn of the suzction elat~e as diffusers will make it possible
to convert part of the kinetiib energy of the auction air inl them
into pressure. For weak suction or emall. Rieyno~L- d num~ber~s
the E101 f~low Sefa:'Cats bnli thie {.Oe 428i ir.CleBBEi 18 the 3135r
diffu'ser is carrersp. .ndine.1~c.v~ :--sl. Frr st~-r ger ;-u~cti n, vlider'
slotes and~ Metrl: 2 otognott ia proileu~lo tJhat is, for? larger

rcdheres-c, the!reby cr~ausin a co:nsiderablel Ipresr-i in~clresco ini the
clOt dif~ifuser.' TIhe slo.t 1-.:E.'L. thcanZj.:~ te re lw.

2. ICnvestigation olf the SClot Flow for Laminar Boundzrry-

Layer Suotion with the Su~ction .Slat: (i) Clurve'!

Forwalrd ~efinit~ionslSee Ber~nninS of .Chaptsr 5i)

Vith the suction slot (i) curved forward (see -fig. 38), laminar
suction tests were performed in the same manner as with elot. (a)

NACA TM No. 1181


and (b). The static pressures in the 'suitionj tank and the velocity
distribution at the alot exit were determiined.

The velocity distributions at th~e slot exit for =1.5m
alot viath and Ra 32.7 kg/m~ 'andt the corresponding static
pressures in the suction tank for various suction quantities can
be seen from the figures 62 and 6j3.

For weak suction, that is, exaller Re of the elot flow, the
slot (1) behaved like the straight alot (a) or (b)..

For stronger suction, that is, larger Re of the slot flow,
the flow continued in the investigated cases on the front side of
the' slot flow under laminar sepaseation; accordingly, the resulting
elb1t losses and negative pressures in' the suction tank were larger
than ~for .the .straight slot (a) or (b).

The transition of the alot flow is probably retarded by the
stabilizing effect of the outward curved front side of the slot.

3. Investigation of the Slot Flow for Laminnar

Boundary-Layer Suction with the Rearward.

Curved Suction Slot (h)

On the basis of the larminar auction tests with straight and
Forward curved suction slots, one may assume that the slot flow
for: a rearward curved suction slot (h) (fig~. 64) 'becomes sooner.
turbulent than for the sltTaigh~t alot, due to the concave curvature
of the front side of the slot. A laminar boundary layer for concave
or conviex curvature becomes turbulent sooner or la'cer, respectively,
than on a plane surface; measurements by F. and M. Clauser (refer-
ence 33) and M. Fazuconnet at the Institute for Aerodynamics,
E. T. H. Zurich (not yet published) demonstrated this fact.

In order to examine the assurmption above, laminar suction
tests were performed with a single rearward curved suction alot (h)
(~fig. 64) whichm:as located on the bottom of a slightly cocmbered
profile of 10.5-percent thickness (chapter 8). The suction slot
was placed at a distance of 0.73 meter from the wing nose. Provi sions
were made for auxiliary auctions on both sides of the test auction,
which had a length of 0.18 me~ter.

M~easuremeints: The static pressurese p, 'in the suction tank
and .p on the" front aide of' the' suction alot were measured for

NACA TM No. 1191

various stagnation p essures and suction quantities for a = .0-rmm
and. s = 1.2~5-mm elot width, with the ving engle of attack remai~nina
constant. pa and p vere ascertained by means of 0.5-maillimeter
p~preassure holes with the static pressure p as reference level
(see chapter 8, determination of U)

The distribution of the static pressure ahead of the slot and
in the slot diffuser for various slot widths, stalnetion pressures,
and suction quantliteP 09/08* can be seen fr~om figuree 66 to 72.
8* = boundary lawyer displacement, thicknless ahead of thle suction slot
calcul.ated? accordlin; to Po~llharusen fromn thle measured' pressure
distribution (fiP. 65).

The static pressure increased sharplyI at thle -presuma'ole
transitioon ninlt in the stri-nht rear part of the slot d.iffusers
adljoininl,, a further weak pressure in~creese; took place. Lengthening
of the slut diff'user provedl ,enerally favorable.

The pressure increase in the slot diff'ruser 7nd hence the con-
versio~n of the kinetic ZneTry of t~he auct~io!n i-r intor pressure are
for the slot (h) superior to those for the strailcht slot (a) or (b),
particularly for smal3er Re of the slot .flow. Co!ospondin~, lyr
the resulting negative preseures in the suction tanrk are smaller.
(See fiPS. 73 and Th.)

NA4CA TM No. 118;1




1. Purpose of the Tests

The purpose of the tests is study of the laminar boundarsy-
layer development with suction for larger .Reynolds numbers, normal
wind-tunnel turbulence and at first alight external pressure
increase on a symmetrical profile of 3.35-percent thickness for
zero angle of attack (profile shape see fig. 75S(a)). The laminar
boundary layer was sucked off through eight consecutive slot.

Furthermore, the problem had to be investigated, whether one
could attain, for similar external pressure distribution, equal
maximum Reynolds numbers for laminar flow Ree or Reg*C referred
to the momentum loss or displacement thickness with or without
boundary-layear suction.

2. Test Apparatue

The wing constructed of w~ood of) t = 2.032 m chnord, was
erected vertically in the closed wind-tunnel test section.Th
span was equal to the height of the test section (2.12 m). For
reasons of measuring technique, suction was applied to one wing
surface only the comparison for conditions without suction was
gained fran the meazsurements on the opposite sjide without suction.

The suction elota I: VIII may be seen from the slot drawing 76.
They are straight and have a rearward inclination of 60o,

As before, they were designed as diffusers with small opening
angle. In order to intensify the stak effect, the surface vae
made slightly wavy to the slot region, according to the tests of
chapter 5 (fig. 76). In the first tests the va~viness was exaggerated,
consequently, the boundary layer remained lamilnar only with vcry
strong suction. Slot width and relative position of slot inlet
and slot trailing edge were asgusted along the span as constant as

The laminar bounrdary-layer development with suction vae
investigated at the wing center, the location of the auction slot

NACA TM1 No. 11811

with the pertinent suction chambers (fig. 7'S(b)). Auxiliary sulctions
laterally to the test station mai'e St. posilbleF to keen the boundary
layer in the entire test section laminor.

Suction w~as applied nega.racely through each~ suction slot. The1
appertaining auziliary suctions were adjust.ed, in the best: po~ssble
agreement writh thle test slot. The length1 of the test su~ction ele~ts
vae for the slots I VI 0.62;0 met~er, for thle RactP VI, V~III
0.208 meter. All suction chslmbers are coniacl Lnd at, an sn-sle to
the wing snd have an amply dimensioned! cross ~sctiin Hence the
pressu-re losses in the sulctio~n chambers are sli:ht and the suction
along the span uniform. The suctilon qualcnititY ;or each test suction
elot was, determined by calibr~ated mneasurtng nozzlex which wlere
attached to the lower end of the respective su~ction ch~amber. The
measuring nozzles used~ may be seen from figur;le 75(c). Th~ey were
calibrated by measurement, of the vnlocityi dis.tributio~n at~ thle
nozzle end by, means of a fl?,t tota~.l-eadl tube of n.11-ai~ll-:mete~r
by ?-milli-meter Inner opening? and 0.2-m.limeterr ou.ter Ileight anA
of a 1.0-millimeter static pressure turbe; the t~wo latee~l
0.4 -mill 1ime ter 11 test holes of: the la~ttr ::ere ot a dista-n~ce
of 9 millimetero fsrom the semic-r~cular hea and of 90 mi-llimeter:
from the sting (2 13m 1). Co;7ate lij.- etefr in th~e c-yljn:r~iral pert:
for the slots I VIT (4) = 17 Inm, for VII, VTIII d = 13 ne. The
static pressure in the suction chambrers me7 me-sur-ed t their
lower end at sufficient distEnce from- the meas~uring~ noteles~. The
suction air was Fguided, thlrough ducto (':;ce) from the meae1Ewing
nozzles; to the Lsuc~tion venti.lator. Thle suetion uan-t~ityv of the
slot vas adjusted~ by throttllng f the~se hons- 'jln to~ varying the
rpm of the vlent-ilsaor.

3. Meaaurementse

I. Suct'ionr side:

1. Pressure tistribultion alo~ng the vIin(: chonrd, measure? byr
static pressure~ holes of 0.5 millilaet~er up to the slot VIII,
from there to thle t~railinEr edge wit~h 1 .0-!:dllinmetjer st attic-
pressure Utube

2. Suction qulantityr of the eirht test eiuct~inn slots, mesasuredl
with 17-millimeter a nozzles for the slots I tr VI end with
13-mill.imeter r1 nozzles for the slot VTIT: and VIIII
4The static pressure tube of 1.0 millimneter used. here was
built like the one used for the calibrations of the' measu-ing
nozzles. Conltroil measurements with connectingr pressure holes
showed that the static pressure p/4o had. been measured within
an accuracy of 10.005.

3. Static pressure iin the eight t~et suction chambers, measured
with 0.5rmillimeter pressure orifices
4. Boundary-layer profile at the point G 9 millimeterss behind
the alot VIII: in 1.790 meters distance from the wing nose
5. Transition-point position by stethoscope and from the
break in the pressure-distribution' curve at transition
The suction was regulated in such a manner that the boundary
layer remained laminar as longT as possible with a minimum of
total suction power, The mneasurements were performed for vari~ousi
Reynolds numbers and, accordingly, different suction quantities
Some measurements were repeated with wider suction lots.

Slot Tests 8 to 37, 19" Tests _18 to 71, 14",. 62"
I s = 0.5 to 0.55 mm s = 0,65 mm
II s = 0. 30 mm s = 0. 35; hmm
III s = 0.36j mm s = 0.40 to 0.4-5 mm
IV El = 0.45 to 0.~48 mm a = 0.45 to 0.50 mm
V s = 0.35 to 0.4 mm a = 0.4 mm
VI a = 0.4 to 0.45 mm a = 0.4 to 0.115 smi
VII s = 0.38 to 0.4 mm s = 0.4 mi
VIII s = 0.3 mm a = 0.35 um~

II. Opposite wall without suction
1. Boundary-layer plrof~ile at the tirailing edge for various Re
2. Transition measurement by stethoscope for various Re
The boundary-layer profile at the point G on the suction side
was measured by means of a flat total-head tube of 0.12-millimeter
by 2-millimeter inner opening and:0.2-millimeter outer height from
the side where no suction was applied. By a micromet~er the wall
distance of the static tube could be adjusted within an accuracy
of 0.01 millimeter. The static ~pressure at the point G vas
measured by a 0.5-mill~imeter pressurre hole which was located
next to the total-head tube.

The boundary-layer profile at the trailing edge'of the oppr~osite
wall, to which no suction had been applied, was determined by
measurement of the tco~tal head. and the static -pr~essure by means of
a total-head tube and a static-pressure tube' the profile dr~ag of
the 0990isite side without suction calculated for both wing surfaces,
results therefrpom according to known~ methods. (Squire-Young,
RM 1838 (reference Ic)).

4. Symbols and Evaluation of the Suction Tests
t wing chord (2.032 m)
b span of test suction, for slot X to VI: b = 0.6520 m, for
VII, VIII: b = 0.208 m,. area of reference

NACA! TM ;No. 1181.

NAIPCAl TM No. 1181

.F projected area of vinlg (bt)

go undist~urbied total head ouvta~rle of bCkinder;~ '.vayer or Pets~e,
masurel writhi respect to stl@m.ile::Pha k;! T per: meter2

14, static pressure in center of' one of sile walls at, be~rrilnnin,:f
closed test section. 0.90 mneter shead of wJing nose,
kilogr;ams per mreter;'

go free-stream d~ynanic pressure, k-ilograme per metclr2 (p.

Uo free-tes velocity to)

p static pressure ai- slurfuee of pro-file, measured id~rth r~espect
to po, kilogramrs per. meter?~

Pa static pressure in onrction cha~mbers, measurerrd wijth re-pect;
to por Frilo3rams per mefter2'

Ap4 sulctioln-btlnwer p eossur,fnr acceleration of suction si:'
to U~o, kil.oira-ms per meters

UL mean air velocity in suction chamber at: nC~le t" w~n(: at
point of: static-pre 8EurT mO89UTnnent, IretSTOJ p87 2007*3l

v velocity in bounla~ry ltyer o'r va3ke. meters per srchJnA

TJ velocity at 8000~ of' bota~ltar_ 1 :::r1 or m~e, in r
ner' sercond

Displacement thlicknlessr 8 = i .1-u 17, wlth1 y = unll} dista~ce.

8 total boundaryi-laye th~ickne:~
Momentum-lose thickcnese 9 = 1 -I -u

H -

8, momentum loss thickness of one of win e idzes far toward'
rear at static pressure po
Re = --

Reg = -

NACA 131 No. 1181

Q&i suction quantity of different test suction slots (1) = I to VIII
for one wing side

We total profile dlrag significant for propulsio~n

V / drag contribution of wakre

Wg drag contribution of suction blowers
Dimensionless coefficients:

Suction-quantityr coefficient of the slot

icqi = 2---

calculated for both wing surrfaces.

Total suction-quantityr coefficient for both wing cur~faces;

Cet = eqiCo,

Suction blower-pressure coefficient,

c = c e- c
Ws 40F' "m qoF k8 90

The drag and coefficients, respectivelyr, wtere calculated for both
wing surfaces.

NACA TM No. 1161~

Evaluatiorn of the Suction Tests

For evasluazting the drag, one securmes t~hac the ruct~ion blowers
efficiencyr q ) accelerate the suction air to thlc free-stream
velocity Uo in flight dilrectiqn toward the r~e'r. A propeller
(in flight or intstlled in the tunnel) (efficiency ?p) Is pre-
supposedl for overcoming the dres~ contribut~ion of thle ;roke. The
efficiencies of the suction blowers andl of1 thie propeller are to
be equal: qg = rip = q.

Then the total profiila drag sirgnificant fo~r the proTpulsionr
becomes, accord~ing to chapter 6, 4(c)

SPa r "



;/ i~


6p = rT 'C02 P D L2

The factor 2 in the first term st-ems from thle fa~ct thlrat
measu~red for one w-ing eider only~.



i=1 F




29 ,
cw = = draCS contribution of t~he v7ake
x t
(9,for wr~ing sidO)

"~ L~

w,=2 VITT LI ,(T-*
Woo = 2 Uo

NACA? TM No. 1181i

Vith the suction-quanti~ty coefficient

2Qq Uo"

and the suction blower-pressur-e coefficient

1 oD

there becomes:

C~,' oil ceic 81

+ CWm = eq~ + cyl


cy= ,T~ cficp = drag contribution of the suction blowers
B 1=1

In order to determine ago thie drarg contributions

c, of

the suction blowers and cw 'of the wake were ascertained.,

c, :: For the various tests, the va~lues Cgi and CPPi of
.the slots I to VIII were dfetermined according to (6)j, (7), (3) and
hence c, according to (9). of2uL2 was in most cases neglibibly
small compSared to 0/2Uo2 ~ Pa*
In order to determine o~'m = --]? ec was ascertained as follows:
The boundary-la~yer profile vasaemasured. at station G9 millimeters
behind the slot VIII at a distance of 1.790 meters from the front.

N\ACA TM No~. 1161

From G on, the lamninar boundary-~layer developmen-t (mementum-loss
thickn~ess 8 and Re, UG as calcu~trlate with the _.eeasu~red

pre~ssure dist~ribut-ion up to t:_e expe> ?rlentallys deterimined1 tran:1!tion
point according to Failkner and. Howarth (refereFnces 4, 4i anr3 kPf
At the transit~ion p-oint 9 was assumeI1d constant. .Frorm the transition
to the t~ral'ing~ e-dge 6 was determninedr byr means o~f a methodl of
differences accordinrl to S~u-ire-Youlngi referencese 4); well shearinrl
stress To equal .to the one of tIhe turbulent flat plate without
pressure gradient for eaual Rep =U (U = velocity zt the bound~ry-

-s I
as p-u2

.U- 3(H + 2)

(boundaryv-layrer momentum eq~uation)



H + 2 = 3.&; U'

s along sur~face.

The wall shearins stress is

pUJ2 2`~

where f = 2.557 x In (h.0175Beq)
(reference 4).

according to squire-YounR .

Reg400 600 800 1.000 31500 2060
R~ I18..91 19.94 20 .68 21cl..26 22.29 23.02

Ree 3000 '1 6000 To000 ..7000 .1 10000 j 15000
S26.06 24.80 25.36 26.23 27.14 28.18

NACA TM No. 1181l



2 x lo







5 x 105

The variation of 9
toward the rear was
(reference 4), with

in the wake from the trailing edge to very far
determined also according to SquireYoungT
the assumption that in the wake


(see also reference 610).

The momentum loss thick~ness 9, far toward the rear becomes:


with the index h
Go3 becomes

referriag to the trailing edge. With

8,FB( 3.2

hence follow a '

and a .

/v \
In t O I

NACA TM' No. 1181 59

5. Teet ResulltR

The test results can be seen from the figures 77 to 90 and
the test tsbles whl'ich contain the 3.nts (.?el at thle end
of the reportt.


ilSI To, t, Re = --, ct; cq

and ep for' thie eirht enction slots, the mome-ntum-loss
thickness 9 and the static p~ressurec at the bouni~rdary;-lsyel' test
point C.

The pressure distributions salnf, th~e cho~rd, t~hs static pr-esoures
in the suction chambers and the tra6nsition-point positions (a~rroucs)
for various Re and suction qluantities are plotted .In fiigurev 77
to 80.

The pressure distri~but~one are! flat ,and shnow a Eli. ht prTe.Sllre
increase only in the r~ear par~t olf the proafile;. The lani~n~ar pr-ensure
increase before the t~ranlsition amounts to 17 to 21..5 percent of the
press~r~e difference between stanar~tion! Dint and pressurre minimum.l
At the sulctionr points alPPeaS" the t.7:pjSic. 1re;7sure inc~rease due to
sink effect which inc-reases with :r'.1~'ing~ suctio~-n ?uantity. Generally,v
the boundaryi layer i s accelerated between the slocts. Onl! behind
slot VIII occurs a str-on!:er lamjnar pre~ssureo increase which finally
leads to the traznsi~tionT inI fienerol. cha~t~l, a)ead o the traslling
edg~e. In~ most casess the PiPresure diftibu~itio~n rlurves show the
typical treak during trEansition.

The negative pressures in the euctionm chaimbere increamz with
grrowing suction qusnf-ity-. TheyJ are smaller f'or lar :er Re, also
for vider suction slots.

In the cases of larger1 Re or of emaller suction nuant~ti0s the
transition point travels f~orward. Thle sutione quantities eqt
required to keep the bound4try layer lumrilna up to a place shortly
ahead of the trailing edge are generanllyr smnll. Witl incr~easing
suction qu-?nti-ty more .sucti-on has to be applied, see optimum cir~ve
cqtopt (Re) for the minimum total dra;. (Se~e fig. 81.)

For suction quantities Esmller than cQtopt one119 ma acertain
by stethoscope, isolated turbulent bursts in the r~ear part of the

NAlCA TM No. 1181

profile, which, with decreasing suction quantity rapidly become
more frequent and start ~a~rther toward the front. The transition
starts earlier and takes place in a, more or less wide tremsition
region which is no longer sharply d~efined1. In these cases thze
boundary lawyer becomes turbule8n~t by the effect of the wind-trunnel
The critics .Re~ynolds numnbers R7eC3 = q during transition,
where the boundary layer under the effect of the external turbulence
barely remained larminar, were for

Re = 2 to 4 x 10 Regkr = 93 to 880

with the narrower suction lots as well as with the wider ones.
For larger Re resulted soilewha~t lower Reper.

The critical Ree-values are slightly Iarger on the sueti.on
side than on the opposite wall to which no suction had ~been applied
'and are practically of the same magnitudae as for t-he lamnarln profile
lk-percent thickness in figure 12.

It may be concluded that a laminar boundary layer with suct~con
reaches the transition point, due to an external turbulence, for
equal Ree-values, as without suction for identical flat extelrnal
pressure distribution if the slots are correctly adjusated.

Smaller Repor resulted only for veryg weak suction. The
suction slot probably- were too wide for~ very small suction quantities
(as was shown by a veri~fyrinsj calculation of the lwzlinar boundrly-
layer development with suction), hence, a local laminar: separation
occurs at the alot inlet thus causing the boundary lawyer continuing
behind the slo-ts to be disturbed observationn by s~tethoacope).

The bouxnda~ry-layer profiles at the station Gr 9 millimeters
behind the slot VIII for.various, auction quantities and dynmamic
pressures can be seen frosl figures 82 to 87.

For weak auction, ~the boundary-layer profiles: are similar as
on a flat Plate without pressure .grkdient- (B1lustus (~reference 42))
and become f~uller with inbrbasing suction quantity. (See also
measurements by M. Bas (references 66 to 683) with laminar area
suction and calculations by 8. Schlicbhting7 (reference 70) about
laminar area, suction on a flaZt plate.)

NACA. TMI NoI. 1181~

In figure 88 the optbumm total dr'agl cw with suction (cuv.;e b)
is plotted. versus Pe = -- nd compared '-.ith thle opporsite side
(curve a) to whiich no? suction had~ been Ipp!-ied.

FrT larger Re the dr--, is considerable reduced b;T the~
laminar bo~undary~:-l.ayer suction. The increase o~f the dro6 rlue tor
tunnel tulrbulenc~e starts with stcct~ion for consl21rab'l .7 lar-iei 9re.
Without suction cy decreases witl Ee .sim'1*r~lyr tio thie 11s~inrl
plate friction to Re =- 105 aind increases again for Ilrrer Ee.
owing to the forward travel of the transition dus to the wi~nd-
tunnel turbulence (flia. 89)-

Up to R~e = 4 x 10(; the disrag wth boundari73y-layer r sCfjion was
only lightly larger thun the 1.eldna! plate fric1~tion; the P~eynlolds
numbers e were relativelr3y low. For~ Rr- = L x 10" resul~ted
cys = 0.00167. For larger Re (.:-. x 100) cW inclreasedl at-:ainI 'TUe
to the eiffect of the tunnel turbu?lence: The'"nuct-ion! in thie su~ction
slots must be mc.e! str~onger andr str.Cnger in ;Irder to avoid -
at larger Fe tu~rblulent disco~tlnui~ties which illncrease the-sk-in
friction a-ndi star-t at E~egcl. Henrce, lay;'er slot, losses and -very
thin laminar boundary layrsr rEsu~lt, d"irectl; behind the suc i:on
slots, causing an increase of the ur~fT.ce rri~ction. Thlereby the
dr~aq c, for larger Re, --in Incre--ses, al-thougv.h it was by
means ore nuct~ion possible toe maiints..o th~e boundlar;- la-er laminar- up
to Re = 5.k X 10'-'. For Te the; lorWest iF~rags resulted
precisely !rith the +strt, of a few iso~lte turilblle~nt bIurst~ which
did not ye~t cause a lasrge increase in elskin frictioun.

The plot of the va:i-tion ofr c,~ andor. versu-Je the suction!
quantity c, for Res 3.0 z: 10" can be seen~r from -figure 0.

cwi. varies w-ith cyt sim~ilarlly as for thle first lamina~r
suction ving of 6.75-percent, thicknes9 (c~hapter h, L):

c,; is anallest for the optirmam su~tion quan1tity COptot For
smaller eqi:-values turbulent d'isconiuti ies!_le appear in the rear
part of the profile, increasingj c ~. For1. lager opt the IsminarI
boundatry layer becomes thinner (see bosundairy-layer measurements
figs. 82 to 87) causing: an incr,esse in skin. "lic'-ionl and1 CW

Widening of the slots resuxlted in sunller negative pressures
in the suction taik anld thus in scrmewhat lower cw and c
(test 38, 14*, 5b).

NACA TM No. 1181

The drag could be decreased still farther by placing more
suction slots behind slot VIII, thus

1. Maintaining thle booundarc:y lawyer la~minar up to the

2. Recovering a considerable part of the kinetic-wake energy~
of the boundary Ilayer at~l the trailin:.:edge (pr?ns~lppo~sitioni
acceleration of the suction a~jir to Ug)

The das=eT cluve c (fig. 88) showe the "Ir:R which mayr be
attained by placing two more slots near the trailing edge. Thle
curve c wras calculated from the test values with the aid of a
theory of the laminar boundar!3y-1E.laye development with suction.

The fact that the dr-ag siignificanllt for the propulsion can1 be
lower thatl the laminar plate friction can be explained by .the
partial recovery of relatively large krinetic wake energy7 of the
laminar boundary layer. By wake~ utilization the .Tr? of thle
laminar flat plate could be daecreased, on prinpcipl~e, by 21.3~ percent
(fig. 88, curve d).

6. Extension of Schlichting's Theory on the Laminarn

Boundary-Layer Development with Alrea Suction

in the Case of Acceleration of the Su~cked Air'

to the Uhdisturbed Free-Stream Velocity Uo~

B. Schlich-ting cazlcula~ted the lamninar: boundary-layer development
on a flat plate -wi~th area suction at constanlt suction, velocityl vo
perpendicularly to the plate (reference 70). Schlichting und3e the
following formulation for the velocity dl.s~t~ribu~tion in the boundary

= e +Kq", r

thu,"he obta-ined the distribution of' the skin friction .c0 shown
in. figure 91 for various Re, and suctioni intehaitie& --oV (Uo =
uandistur~bed free-strea~n ve;locityg). To,

N/-CA TLINo. 1181

The condition on the plate infinitely far toward the reasr is
given with strict accuracy whereea for smaller Re, eqi~ is
overestimated. W!ithout suction the result. is, accrrdi~nlr to

c, 16 instead of ."-3 5 according to aRlsius
R / Pe /5

If the suction a-ir is accelerated- to Up, witl~houJt less, the
drag c'2 whlich is significaFnt for' the propulsion becomes con-
siderably lowrJ~ than the skin tfriction c the reason Is the
recovery of the kinetic vake energy, of the suc!ed air which
otherwise moves forward with the 01sts?, (See; fig. 37.) Fra

inf-initely long plate c2bec ome 2 c' 0.

The drag eq for acc73leration of the suction to U,
decreases with P1e considerably mere str~onely t~han the surface
friction c uner thle assumption of Iunifoirm s.Icti~n intenait,- -

Hence, low drags3 c'2 become possible at l.arger Re for r~el:t~ively
thin larminlar bouindarl layer whlich are, wi~th respect to stabil.ity,r
better cnntrolla~ble.

Conver~sely, it: In probal\yl~ posble to o~btai-n, for h1Zh~er
admissible laminar Pe-va~l.veT (with ryeak ex:erna~l turbulence) cith
the aid of the lamninir boul~ndsr ~-l'ayer suction very large P'e
under lam~inar conditions, th~e dri cu would become only sl!ghtlyp
larger than the 12~minar-plate .fr'ictiron.

A further draG reductions wou?~lld reBul't i one would, moreoenr-cr
accelerate the boundary layer a, thle endl f theF plate with-out loss
to Uo, thus possibly recovering its wake enerisy. (See fig. 93.)

The utilization of the wake r-nergy of the boundaryi layer on
a laminar~ suction profile could be attained with relatively small
losses by gradual suction of thle boundryn lawyer in the realr par~t
of the profile ~t~hroah .seve-ranl uctio~n s!ots plascedl one bohin?. the
other for a stattic. pressure increasing :-;var'1 the rear, and byv
reaccelera~tion of the suck~edl ir to 80. I each slr-t, suction
need be applied only to a fraction of !i~e respective boundary layer,
so that mall slot losses reSul~t. In this manner a thin lamninar
boundary layer would. result at the trail~ing edge of the wing ite

64 NACA TM N\o. 1181

wake energyi would now be only slight; moreover, it la parrtiallyr
recovered during the following accelerat-ion in the wa~ke. An
arrangement of pressure propellers also would make it possible
to recover a emall1 part of the bounrdary-layer watke energy (as in
the similarL case of vake propellers of shipe) .

N.^CI.TM No. 1181




1. Purpose of the In~vestigation

The laminar pressure increase with boundary~-layer svctlion was
studied for higher Ro, at normal wind-tunnel turbulence and for'
various ca, on a slightly cambe-red. profile of 10.5-percent
thickness and conventional thickness dietribultion. Frhroe
the drag reduction by the laminar boundary-layer suction was

2. Profile, Test Arran~gement

The investigated profile with the suction slot can be seen
from figure 94.

Profile thickness d/t = 0.105

Curva~ture of the mean line f/t = 0.019

N~ose curva-ture radius Rof/t = 0.00197

With the selected thickness distribution, the pressure increase
on the upper surface starts relatively far to the front.

The w~ing whlichl was made of vood was erected ver-ticall: between
floor and ceiling of the closed~ wrind-tl.tmanel test section. In the
central wing~ section wrere the test suactio~n slots of 0.18 rmeter
length with auxiliary suction slots on bothl sides. Suction was
applied to the boundary layer on the upper sujrfsce t~hr'oughlb e4lots
and on the lower surface through 10 slots. For most tests the
foremost lots of the upper surface wJere sealed wilth pul*tty and no
suction was applied to them (see test tables). A4 separate auction
chamber, in which the static pressure was measured, was connected
to each suction slot. The suction quantit,; of' eachz test suction
slot was determined by calibrated measurin, nceeles a~ttachedl to the
lower end of the suction chamber. The sect-ion chambers were conically
developed in suction direction Ind had a cross section writh ample
dimensions. The anx~iliary suctions were adjustedl as similarly as
possible to the test suction.

NACA TM No. 1101

Form and position of the suction alots can be seen from
figure 94. The narrow lots which are developerl as diffusers have
a~rearvard inclination of 600. The forward or rearward curvature
of the slots served only for deflection of- the suction air, not
for an artificial. production of turbulence of the slot flow.

In order to intensify the lamninar pressure increase by sink
effect, the surface in the slot region was made slightly wavy,
onl the basis of earlier tests, for test 21 considerably less
than1 for the test 27 to 55 (see profile sketch).

The suction slots along the span were as far as possible
identically adjusted (constant; slot wid~th and shifting of the
trailing edge of the slot relative to the slot inlet along the

3. ]Measurement with Laminar Boundary-Laeqr Suction

(a) Pressure distribution along the chord, measured with
0.5-mllimneter Q pressure holes and with 1.0 mjd111meter
static-pressure tube; (b;), (c) suction quantity and static pressure
in the different test-suction chambers (vi-th callibretedl measuringq
nozzles and with 0.5 amu bore holes), (d) nomantum measurements
in the wake.

The boundary-layer condition -in their suction .ei-~on was
verified byr stethoscope. 'The boundiary-\~lay'er` SUCtin Was recgulat~ed
to the loveat possible total l-1 D.

The tests were preformend for various c, (fromn pressure
distribution) and Rie ,in a few cases the suction quantity
was varied. For the tests 27 to 55 the el.ot width were enlarged
as compared to test 21.

For comparison, GW without suction in the part of the
wings to which no suction had been applied was dletermined fo7r
various Res byl mears' of the namentum method. (See fig. ?5.)

NA;CA TM No. 1181

Symbols and Evaluation of the Suction Tests

Wing chord t = 1.035 m

Span of the test suction b = 0.18 m

Reference area F = bt = 0.186 m2

The investigated wing was relatively large, compared to the
cross-sectional area of the test section (3 x 2.12 m, octalgonal~).
The profile then operates between side walls or in a wing cascade
(mirroredl wings). In order to obtain from the test values the
properties of the investigated wing as individual profile in the
unlimited air stream, the undisturbed free-stream dynamic pressure go
and the free-stream velocity Uo, respectively, and the static
pressures at the profile evaluated as follows: The tunnel side
walls and the wings mirrored on them, respectively, cause the test
wing in the center of the tunnel to be subjected to a stream of air
with a velocity which is by AUlargSer than the free-stream
velocity U'c far toward the front: Uo = U00 + DU. (The index G
refers to the cascade.) OU is the incremental velocity due to the
mirrored wings at the location of the test wing. The undisturbed
static pressure po at the location of the test wing is by
0/2 (002 Uo 2) lower than the static pressure pop far ahead
of the wing.

The incremental velocity CAU at the location of the test wing
under the influence of the mirrored wings was calculated by replacing
each of them by a source and s-ink at 0.8-meter distance in free-
stream direction and a vortex. The strength was chosen so that the
maximum thickn~ess of the mirrored wings equalled the one of the
test wing. The mirrored vortices do not cause an incremental
velocity in free-stream direction in the tunnel center at the test
wing location, on the other hand, w~ith growing ca the effective
profile camber is increased by 6~f (see P~randtl-Betz (reference 85):

Af/t = 0.0023ca. Therb resulted AU-0.006, thus Uo = 1 .006Uo,
The following quantities were measU id: the Gstaic pressures p .
and pur on both tunnel side walls at the location of the neximum
profile thickness for installed test wing, furthermore, the total
head go and the static pressure pog' at the location of the
test wing (for "emrptiy-tunnel" condition) (py'1 2' and go were
were measured with the atmosphere as reference level). In general,
py. and p,2 are different due to the circulation around the wing.
Because of the displacement effect of the wake behind the ving po
is slightly different from poC

68 NACA 'lM No. 1181

The static pressures p1 at the profile surface and p in the
suction chambers were determined wi-th po ae~ reference leval..

Evaluation of po> go, Uo from Puts, Pw2.- The incremental
velocity aU, at the tunnel side walls at the location of the
maximum profile thickness under the infIluence of the test wing
and the mirrored vings together was calculated, all wings were
replaced by ,a source and a Sjink at 0.8 meter distance each. From
nv,, au, and p, = p 2- 5 folows the undisturbed static
pressure po at the location of the test; wing squaredd terms of
bU and nt, neglected):

Po pV 2(.dity aU)



90 = 80- Po

The further syvmbol~s and the drag evaluation are the same as
in chapter 4, 4. Acceleration of the sucked air- to Uo and
equal efficiency of propeller and suction blower were presupposed.
The kinetic energy urL2 of the sucked air in the suction chramber
was included in the blower pressure by .

Qai suction quantity of slot (i) measured with calibrated .
measuring nozzles of same shape as in tests of chapter 7

Ppi suction blower pressure for slot (5) 10 pai uL~Z

pal static pressure in the auction chamber (i), measured with
O .5-millimater 6 holes

NACA dN No. 1181

p static pressure at the surface
Re v

SuctiLon-quantity coefficient cqi = -- of the slot, (i). Total
suction-auantity coefficient o-

09 = 04

for uplIer side,

cq, = /' e

for the lower side,

Drag contribution of the suction blower:

Dlrap, contribution of the wake c0 = --, determined by momentum
measurements. Total profile drag significant for the lift

Cwo = cW,' + CW,

eg was evaluated from the prenssre-distribution measurements.

The test results can be seen from the test tables and the
figures 95 to 103.

b. Test Results

The minimum drank with laminar boundary-layer suction (pover
required for suction included) is plotted versus Re for various ca
in figure 95 and is compared w~ith the measurement without suction.

NACA TM No. 1181

In spite of the thickness distribution which is not particularly
favcorable with respect to drag, the resulting total drag for
Re = 2.2 x 106 is cw, = 0.0023 with suction, comlpared to
opt Wt uto
crot=0.00535 without suction for Re = 1.4 x 10 .Wt uto
eq decreases with Re up to 2 x 100 sim~ilarly to the laminar
friction of the flat plate and is only slightly larger than the latter.

For Re> 2.2 x 10 cm4 increases again due to the wind-
tunnel turbulence (starting of isolated turbulent burate similarly
as in the sucti.on tests of chapter 7).

By widening of the sutction slots (tests 27 to 55) the slot
losses and cwu were reduced (compar~ison of thle tests 27 and! 21).

Figure 96 shows the optiman profile drag polars with laminar
boundla~ry-layer suction for various Re.

Since the boundary layer was krept~ completely laninar up to
the trailing edge on both w~ing surfaces, c, remained low in a
considerable ca-range, thus causing favora lee profile-drag lift
ratios. In order to maintain, for larger ca> the boundary layer
still laminar up to the tr'ailing edge in spite of the growing
pressure increase, stronger suction must be applied on the upper
side, whereas the suction of the lower side may be correspondingly
reduced, and vice versa.

For still larger (or smaller, respectively) ca one wing
side finally becomes turbulent, in spite of stronger suction, due
to the incipient sunctilon peak at the wing nose, hence, the drag
increases accordingly.

The influence of the suction quantity eg on cown, cym *
and Cy can be seen from figure 97 (Re = 2.2 X 10 ).
~for cwop the boundary layer remains laminar down to~ th-e

t~railingi edge with small suction quantities Ceg ~ on bo~thl wing sur-
faces. For larger suction quantities eg> c optj thielaminar
skin friction and the total drag cWC. increase. For weaker suction

cq<' cqapt isolated turbulent bursts start in the boundary layer,
in the region of the pressure increase, which become raPidly mor~e
frequently with decreasinr suction quantity and increase the skin
friction and c,

The percenlt-draP contribution of the suction blowerp to the
total drac is considerable, particularly for larger Re.

Fig~ure 98 shows a comparison of the wakes with and without
suction. The drag contribution of the wake with suction is very
much smaller than the drag without suction.

The pressure distr~ibutiojns along the chord can be seen from
figures 99 to 103. The test; points obtained, byv pressure holes and
the static pressures in the suction chambers are plott~edl. The
pressurre-li stribution curves were supplement-ed by measurements
with a 1.0 statsic-pressurye tube; the corresponding test points
do not appear in the olot.

For larger Re considerable lamninar pressure increases were
obtained with the ai'l of boundary-layer suction, for instance.. in

No. ca Re Laminar pressure increase
( pe;ecent)

27.1 0.232 2.16 x 106 4?

55.3 .587 1.5 3.3

The sink effect makes an essential. contr~ibu~tion to the laminar
pressure increase.

For a few cases the 3amnan~r bounilsry-layer development was
determinc-d alone the chordj down to the trailin:5 edge. The boundar~y-
lay~er development alonc, the wall between the slots wIAs calculaitedl
according to Prohlhausen preferencee 43). The boundary-layer mo~mentum-
loss thickness directly behind the locations of suction was
determined accordingn toe Bernloulli from the momentan-lose thickness
of the boundaryr-layer part ahead of the suction point to which no
suction had been applied (milxinG within the boundary lawyer with
pressure increase as a result of sink effect neglerect). There-
from resulted. 9 along the chor~d a~nd Reg = .2, respectivelyI
(U = vel~cit~y at the edge of the boundaryr layer at the particular
station). Those critical Reynollda numbers Rpe were designated
by Reeqp where the bouxndasry layer just remained laminar on both
surfaces down to the trailing edge.

Influence of the external pressure distribution on Reper
For not too small ca-values (ca 0 .27) similar Reper resulted

JHACA TP-I ~o. 118:1

NACA TM1 No. 1181

on the lower wilng surface (flat pressure distribution) as for the
symmetrical suction profile of 3.3,5-percent thickness of chapter 7
(Regc, = 800 to 850). On the other hand Raer at the upper surface
in the pressure increase region was essentially lower, particularly
for larger ca (Rego = 550 (at start of pressure increase) to
400 to 450 at te end. of the pressure increase in the neighborhood

of the trailing: edge). Thus, similarlyl to the case without suction,
higher (or essentially lover) Reper than for pressure gradient
zero result for accelerated (or retarded) external flow for; laminar
boundary-layer suction.

Influence of the sin~k effect on Reeor.- The tests 21 and 27
with sink effect of different strength resulted in practically equal
Reg-values, with cOopt remaining thle same. Thus it seemsg that
primarily the total external pressure increase is decisive for the
stability of a lamilnar boundary layer with suction with increasing
pressure. Whether this total external ~pressure increase is to a
larger or smaller extent created. by sink effect or by the flow
along the wall seems, within certain Itmite, of lesser importance.

If one calculates from the velocity gradi~ent u' of the
external pressure distribution as it would result; without sink
effect a quantity X = 62!Vur (which corresponds to the Pohlhausen
method), the following values result in the region of increasing
pressure on upper and lower side: X = -2 (start of the pressure
increase) to -8 (at the end of the pressure increase). Therefore,
similar negative X-valures are obtained for laminarp pressure increase
with boundary-layer suction as without suction, under the assurmp-
trion of equal external turbulence and equal Rep.

5. Conclusions fran the Tests of Chapters 7 and 8 for the

Design of Laminar Suction Profiles with the Low~est

Possible Drag for High Reynolds Naumbers

In order to obtain for laninar suction profiles highest possible
Ree over a, large range of thie ving chord, and ~thus a low surface
friction and small dra:= for larger Re, one should use profile shapes
with uniform pressure distribution and a pressure i~ncrease occurring
far toward the rear, as they were developed for laminar profiles
without suction (chapter 3). This corresponis to the' combination of
the test of chapters 7 and 8.

NACA TM~ Ilo. 1151

Weak suction only wnuld have to be applied in the region of
flat pressure distribution. Boundary-layer thicknless and Ree
would bsvre to be reduced byl suction! ahead of" the pressure increase.
So much suction would have to be applied in the region of pressure
increase that, for the pr~eent turbulence = 0.004, Feg
remais Re ( 50 an 1 = u' dose not become excessively

negative (X 2 -6) (u' = velocity gradient of the external pressure
distribut~lion without sink effect).

6, Prospects for Application of Lamnar .Bound~ar-Layer

Suction in Flight for fI!Fh Peyvnolds Nurmbers

The caclculation for' fl!sht measurements on a laminst profile
of 15.9-percent thickness (15) resulted, 9lhead of t~he transition,
in a Reynolids number Ree UO = 2600 in the region of thle point of
laminar separastion chapterr 2, ~2), that is, about three times more
Than for the pres~ent. wind-tunel tests. Hence, pre~sumably, for
I lminar suction! profiles in f~light, with a boundary~ layer kept
completely lumin-, about three times smalller' drass for nine times

higher Re o are possible than had been measured in the wind

For higher flight velocities the percent atmospherice
turbulence u-decreases; thus one could then expect higher
laminar flow Re and lower drage, rt least as lon6 as no com-
pressibility disturbances appear.

The drag of fuselage and tail unit also could be considerably
reduced by maintaining the boundary layer laminar with the aid of
suction. The fsirings from vaing to fuselage, etc., also could
in principle, be kept laminar by bolundar~y-layer suction.

Thre induced drag whl-ich now gains renewed importance may be
reduced by enlarging of the span and increasing of the flight
velocity, possibly by sta~ggred flight arrangement (as used by
migratory birds). The optimumv drag/lift ratio wcoul~d result for
mall ca, that is, for high flight speed at not extreme altitudes.
Large pans require wings with sufficiently thick profiles and

NACA TM No. 1181

adroit design of the wing structure. In order to obtain high Mach
.number~s without compression shocks, the superstream velocities ought
to remain as samll as possible.

For laminar suctio~n profiles in flight exists the possibility
to adapt the boundary-layer thick-ness aind Ree, respectively, by
means of boundary-lay-er suction to ~the respective state of turbulence
of the atmosphere. F'or higher atmospheric turbulence stronger suction
would have to be appied far-thero toward the ~front in or~der~ to keep
Ree suf~ficienltly Low and to obtain a lower sensitivity with respect
to variations in angle of attack due to gu~st8, etc. Conversely,
weaker suction could be. applied in case of the air being~ very calm,
resulting in larger Ree and correspondingly lowerl drag.

Translated by Mary L. Mabler
National. Advisory Committee
for Aeronautice






It had been shown on the propeller profile number 11 of
9-percent thickness (fig. 11, chapter 3, 3) that the boundary
layer of the upper wing sturface undergoes, for small Re and.
smooth inflow, lsminar separation and does not turbulently readhere
(stethoacope); hence, the profile drag is sometimes considerably
increased. Only for larger on the boundary layer of the upper
wing surface is disturbed so much that the transit-ion occurs in
time to obtain here a turbulent readher~ing of th1e boundary layer
for smaller Re.

A report on tests on a medium-camberedl profile of $-perc~ent
thickness (fig. 10k) follows. The pr~ofile-drag polars on this
profile were, for smaller Re, imp~rovedrl by artificially produced
turbulence of the boundary layer on the upper w~ing surface in the
region of the point of laminar separation with the aid of surface
disturbances (steps in the surface, see fic. 106: disturbances 1
and 2),

For various Re and c, momentum measurements were performed
on the smooth wilng and with the disturbances 1 and 2. (See profile-
drag polar figs. 10k to 107,) For low R~e the profil.e-dlrag
polars are definitely improved by use of the disturbances, the
boundary layer after the disturbance generally turbulently readhering
and undlergoing laminsr separation only for very small Re
(observation by stethoscopee. The disturbance then would have to
extend farther to the front.

The weakcer or stronger 11sturbance 1 or 2, respectively,
improves the drag polar mainly in the Re-region of 250,000 to 300,000
(or 200,000, respectivelyy. For larger Re the boundary layer
becomes unnecessarily early turbulent due to the disturbance
(observations by stethnoscope), thereby corresspondinglly increasing
the profile drag. The Reynolds number Re2 =- referred to the
distance I from thle start of the disturbance to the start of
transition resulted for the present case as Reg = 48,000

76 NJACA TM No. 1181

(3 = distance from start of disturbance to start of transition;
U =- mean velocity at the boundary-layer edge between start of
disturbance and start of transition) .

The method of producing an artificial turbulen~ce of a laminari
boundatry layer in the region of the point of laminar separation
is, in general, successfully applicable if otherwise a stronger
lamina-r separation occur and the boundary lawyer does no~t turbulently

NA~CA TM no. 118;1


1. Jonesa, B. M.: Sk~in Friction and the Drag~ of Str1eamlline Bodies.
R.M. 1199, 1928, 1929.

3. Jones, B3. N: FliGht Experimuents on the Boucnday Layer.
J. .'t. Sciences 1938, Jan., v;ol. 5, no. 3, p. 81L.

3. Serby, J., Morgan, M;., tind Cooper, E.: Flight Tecsts on the
Profile Drag of 14 Percent andh 25 Percent Thrick Wing~s.
R.M. 1F26, 1937

4. Squire, H. B., and You~ng, A. D.: The Calculation of' the Profile
Drag of Aerofoils. R.Mu. 1E38, 1938.

5. Pretsch, J.: Z~ur theoretischlen Ber~ech~nung des Profilviderstandes.
Jhrbruch 1938 der Dut~schen Luftifablrtf forschung, PP. 1-60.

6. LeWis, G. Ir.: Eiome Moder-n Methods o" PeSE-~rch in the Problemns of
Flight, Low Tvr-oulence W~lind Tuninel. The Joulrnal of the
Royarl A~eronautical Society 1939, p. 779.

7. Tani:, 1., and Natu-isi, 5.: Contrib~ut-ions to the Design of
A~erofoils Suiltale for H~igh Spe~ede Rep. of the Aeronautical
ReEserch InstYitu.te Tol:io, NoT\. 198, Sepit. 19/40.

8. Pfenninger, r.: Vbe-r die serodlynamische Durchbildun2 von
Flii'3l;'t;eblenanklischUssn. Flu23,*ebt Lund-Techinik, Sept. 194~2.

9. Jacobs, F'est,!crun and ;-on Doenhoff, A. E.: T'ransition as it
Ocetre ~societed- wit~h anld Following Lamnninr etparastion.
;th Inter~n. C'ongr. fo~r appl. Mech., CEombr., Man.Z 1938,
p. 311.

10. Hall, A., and Hisclop, G.: Ex~per~iments on thle Transition of' the
Laminar Bound.ary- LOayer on a Fla:t Plate. HR.M. 183,~ 1938.

11. Taylor, G. I.: Somne R~ecenrt Devel3Fpmlents in the Study of Turbulence.
P-rolc. /5th Inte~rn. CinEgr. for aipFl. Nech., Camnbr., Masc. 1938.

12. Fag~e, A., lan Pr.eston, E.: Exp~erimeInts8 on Transition from Laminar
to Turbulelnt Flojw in the Boundary Lay:er. Proc. Roy. Soc. A,

13. Schubauer, r.: Air-flow In the Boundar~y Layer of an Elliptic
Cylindler. IJACA R~ep. TTo, 652, 1939

78 NACA TM No. 1181

14. Fage, A.: The Airflow Around a Circular Cylinder in the Region
where the Boundary tLayer Separates from the Surface. R.MI. 1179,

Fage, A.: On Reynolds Numnbers of Transition. R.M. 1765, 1937.

Page, A.: Experiments ch1 a Sphere at Critical :Reynolds
Numbers. R.M. 1766, 1937.

15 Amerikrenische Profflwriderstandame asungen im Fluge an ei~nem
15*9 percent dicken Laminarprofil mit einer "KingT Cobra"
(nalch Reis~eberichten).

16. Lyon, R. M.: Flow in the Bound~ary L~ayer of Streamline Bodies.
R.M. 1622, 1934.

17, 18. Taylor, G. I.: Statistical Theory of Turbulence, Parts I IV,
Proc. Roy, Soc. A, vol. 151, no. 873, Sept. 1935, Part V,
Proc. Roy. Soc. A, vol. 156, no, 8088 ,Aug. 1936.

19. Von Kearman, Th.: Turbullence and Skin Friction. Journal of the
Aeron. Sciences Jan. 1934.

20. Von Karman, Th.: Turbulence. Journal~ of the Royal Aeron. Soc.,
vol. 41, no. 324, p. 1109, Dec. 1937.

Von Kalrman, Th.: Some remarks on the Statistical Theory of"
Turbulence. Proc.-of the 5th. Int. Contr for Appl. Mech,
Ca~mbr., Mass. 1938.

21. ]Dryden, H. L;.: Airflow in, the Boundary L~ayer neazr a Plate.
N.A.C.A. Rep. Lj62, 1936.

22. Drydren, E. L.: Turbulence and the Boundary Layrer. Journal of
the Aeron. Sciences Jan. 1939.

23. Drlyden, R. L;.: Turbulence, Companion of Reynolds number.
Journal of the Aeron. Sciences, April 1934.

24. Dryden, K. L.: Turbulence Iinvestigations at thneNational
Bureau of Standards. Proc. of the 5th Intern. Coner~. for
Appl. Mech. Cambr./Mass. 1938.

25. Pranatl, L.: Beitr ~ge sumn Turbulenzsymp~losion. Proc. of the
5th Intern. Congr. for Aprpl. Mech. Cambr.,MNass. 1938.

~Prandtl, L.: itber daie.Entatehung der Turbulent. ZAMM 1931,
p. 407.

NAsCA TM NT-. 1161 79

26. Schlichting,H.: Ber-echnung der A8nfachuvng kleiner Stiorunnsen
bei dier P)lasttenstri~mu~n: Z.A..M.]R 1?33, Ed. 13, Heft ?.

27. Schlicht~ing,H.: Zulr Entatehnlng der Tu~rlenz bel der
Plattcnstr~i mungng. 1S--chr. Ocs Wiss. Gottingen, Math .-Phys.
Klansse 193=, p. 181..

28~a. Tollmi~en, !.: it7er ~ii Entste~hung ier Tu1~rbulenz Es~chr. Ces. Wise.
Gi~tt;in,:en, Ma~th,-Phys. Klasse: 102)~, p. 21..

28b. Tollmien, W.: Obrp die; Korrrelation der Ge~schwindlikeits-
kcnponenten in nor.:C;~ch sChZwankendenI Vilrbe'lverteilung en .
Z.A.MI.M. 1935, n. ?r6.

29. Tollmien, W.: Ein a'l~PlueineS K'rite;'rium dler In~stabilit'at
lamiCnR1'E: r GS ah\inli ght~i .70trtel lan en. Eiarhr. Ges. MIrSS.
Gitting~en? Malth.-Phys. Klasse 1935, Fach.Trup'pe 1, p. 79.

30. Fage, A4.: Traneition in the Boundary Lsyer Caused by Turbulence.
F(.M. 1896i, 19612.

31. Peters, H.: A Study in Bounda~ryl Leyers. Sth Intern. Congr.
for Appl. Mech. Cambr.,M~ass, 19J38, p. 393.

32. Stephens, A. V. an' Kall, A.: Hot.-WJ'r~es In Flight. Proc.
Sth Inltern. Congr. for Apl. Mec~h. Quan~lm,Mass. 1033, pq.336,

33. Clauser, M. and~ F.: Thze Effect of C~urvaturx E On the Transition
from L~aminar to Turbulent~ .Bonda~rp Lleve. N.A.C.A.
T.N. 613, 1937.

34. C~jrtler, H.: Instcbil~it~t lamroinar~er Grenzac~hichten an konkahven
Wiinden gegenuber -twissen 3-dimensionelen Stirungen.
Z.A9.M.]M. 19!L1, p. 250. Machr. Ges. Wics. bt~t-ingen, Math.-
Phy-s. Klasse 1960O, p. 1.

35. Prandtl, L.: Einflues stabilisierender Krif'ts auf die Turburlenz.
Vor~tr~ge us dem Gebiet de~r Aerody~namaik !und vercrandter
Gebiete. Aachen 1929, p. 1.

36. Young, 4. D.: S~urfacr; Pinish and Perfor~mcnce Aircraft
Eng. Sept. 1939.

37. Hoodl, M. J.: Surface R~oughness and 1JJ~ng Drag. Ni.A.C.A.
T .N. 695 Aircraf-t; ng. Sept. 1939.

NACA TM No. 1181

38. TadF, I., Hamna, R, and Miituisi, S.: On the Permissible Roughness
in the Laminar Boundary Lay-er. Rep. of the Aeron. Research
Institute Tokio, No. 199, Oct.~ 19340.

39. Theodorsen, T. and Gatrrick, I.: General Potential Theory of
Arbitrary Wing Sections. N.A.C.A. Rev.452, 1.933.

40. Kochanowsky, V.: Zur Berechnunr3 der Druckvertellung "uber dan
Umfang beliebig geforat~er Fliigelchnmit~te. Jahrb. 1937
der Deut~schen Luft~fahrtforechung, pp. 1-58. D.YV.L.-
Jabrbuch 1-937, P. 139.

41. Kocha~nowslry, W.: Veitere Ergebnisee von Druckverteilunga-
rechaung~en an beliebigen Fliigelechnitten. Jahrbuch 1938
der Deutschen Luftfahrt~forachung~, pp. 1- 82.

42. Pinkerton, R. M.: Calculatted and ]Measured Pressure Distribution
over the MidapEan Section of N.A.C.A. 4412 Airfoil.
N.A.C.A. Rep. 563, 1936.

43. Pohlhausen, K..: Zur ndh53erungsweisen Inteo~ration der
Differentialgeichung der laminarenz Grenzechicht. E.A.M.M.
1921, P. 252.

44. Falkner, V. Mc. and Skan, S. W.: Some~ Approximate Solutions of
the Boundairy Layrer Equaltions. R.M. 1314, 1930.

45. Falkner, V. M.: A Further Investigation of Solutions of the
Boundasry Layer Equations. R.M. 1884, 1939.

466 Falkner, Y. M.: Simplified Calcultation of the Lamninar Boundary
Layer. R.M. 1895, 1941.

47. Blasius, E.: Gr~enzechichten in F'ltsaiglkeiten mit kleiner
Reibung. Zeitschr. f. Math. u. Phys. Ed* 56, p. 1, 1908.
Dies. G~ttingen 1907.

48. Howarth, L.: On the Solution of the Laminari Boundary Layer
Eq~uations. Proc. Roy. Soc. A, N'o. 919, vol. 164, 1938.

'49. Pran~dt1., L.: Zur 13erechanwg der Grenzechichten. Z.A.M an.
1938, p. 17.

50. Go~rtler, H.: Weiterentvilckl~ung eines Grenzechichtprofils bei
geg~ebenem Druckverlauf. Z.A.M.M. 1939, P. 129.

NACA TM No. 11I1

51. How~arth, L.: SteadY FlowJ in the Boun~aryr Laye- near the Eurface
of a Cylinrder in. a Stream. R1.M. 1L;32, July~ 1P34.

52. Tomotika, S.: The Lamninar Bunvndaryu Layer on thre Surfacre of a
Sphere in a Uniform Stre~m. Ri.M. 1;7.3, 19357.

53. Buri, A.: Eine Berechn~un-scrundlage fiir die turbulente
Grenzschicht bei beschleunilter and verzdferter Grund-
strdjmung. Dise. Z'djric~h 1931.

54. G-ruschwitz, E.: Die~turbrulente Eeibunrsechicht in ebener--
Strimunr bei Druckatfall uxnd Druckunsties. In.-. Archiv,
1931, p. 321.

55. Kehl, A4.: Uhtersuchungen iiber konverer.ente und ~divergente
turbulente ReibunEsach~icht~en. Inc. Archiv, 1943, p. 293.

56. Young, A. D,: Ther Catlculation of' thlelToth1 Zrnd Skin Friction
Drage of Aodies of Pevolu-tion at Go' Incidence. R.M. 1874,

57. Von Ka'rmn, Th.: iibeiur lminare und turbu~lente R~eibung.
Zlectchr. f. enger. Math. u. Mec~h. 1, 19?31, p. 233.

58. Gidttin~er N~achrichten! 1930, p. 58S, Yortray allf dem ',. Int.
Kongr. f. techn. Mech. in Stockholmn 1930 (Velrhan41uvngen
dieses K~ongr. R? 1, p. CS).

Von Karmanj, Th.: Theorie des Rieibu~n~seidp~~ners ane. In
Kydrolcmech. Probleme dles SchiffslantrieFe 193'2.

59. Prandtl, L.: Ergebnisse ler~ A.V.A. Odattingen No. 3 (1927), p. 1.

60. Von Deeinho~ff, A.: A Preliminary InvJesl~sigt ion of Boundary
Layer Transition Along a Flat plate with Aryverse P-ressur-e
Gradient. NI.A.C.A. T.N;. 639, 1938.

61'a. Ackereet, J.: Das Institut, fiir Aerodlynamik an d~er Eldg.
Techn. Hochschule.

61b. Datwyler, G,: Eine Apparatur tur Messan; turbuilenter
Schwasnkvngen in Strl~mung~en.

61c. Pfemainger, tV: Verfleich der Impualsmethorl.e mit der TVafgun
bei Profilwi derstandamea sun~gen Mi-ttellung No. 8 des
Inst. f. Aerodlyn. E.T.H. Zurich, 19643.

NACA TM No. 1181

62. Lord Rayleigh: On the Inqstability of Certain Fluid Mortions.
Proc London Math. Soc. 11, p. 57, 1.880 und 19, p. 6I7, 1887.
(Scientific Papyers vo;l. I, 1. 4711. und- aol. III, p. 17*)

63a. Prandtl, L.: Bemarkung~en Eb'er die En-ta.tehung. der Trurbuilenz.
Z.A.~M.A. 1921, p. 431.

6:3b. Tie~tjens, 0.: Beit~rtge :jber die Entatehlung der T~r~bulent. Diss.
G~tting~en, 1922. Z.A.M.M. 192), p. 200.

64. Patry-, J.: Ins~tabili~tel.d~'une range'e de ~touribillons d~e long d'unet
paroi. Helvetica Phnysica Acta 1943, ~p. 88.

65. Ackeret, J.:_ Probleme ,des Flugzreugantriebs in Gegenwart und
.: Zukulflt Schweiz. Baureitung~, BA. 112, No, 1, 1938.

66. Ackleret,. J, Ras, M., and Pfenrninger, Wl.: Yerhinderung des
: .Tarubu;Lentwerden s einer Grenzaichicht durllch Abeaugung.eZ DPie
Naturvliesenschaften 19111, 29. Jabrg, H~eft Ic1, p. 622.

67. Rsea,;-M,.,and Ackeret, J.: jober V~erhin~der~ung der Grentachicht-
.turbulenz durch Absaugung. Helvletica Physica Acta 19'41.":

68. Ras, M.:. D Iiss.Paris:. Contributions al'k4tut~e de la i..ouch~e
limite aspijriee 194f5

69. Gerber, A.: Un~tet;~rsuchu~ngen 'iber; GrCenzchiichtababsa,~t ugung. Mittleilung
.'Not 6 dea-Instl. fir Aer~od~ynamikr E. T. H. Ziirich, 1938*

'70. Schlichting, H.: Die G~reneschicht an der ebenlen Platte mait
S.A :.bsayguag. undl Auablasen, Luftfahrrtforachung 1942 p. 179,
p. 29)3.

71. Sch3ichting, E.:. Berechnung der laiminaren Grenzschsich~tentvicklun
mit Abeaugung: an eineta Jouzkowr~sk~yprofil. Cahiers d'Ae'orodnamique,
No. 3, Oct;.-Nov. 19415.

72. Prandtl, L.: The Mechanics of Vciisc~ous Flu~ids. Aero~dynanict Theory,
Durand, vol. III, Div. G.

73. Prantl1, L.: Stfrijmun~gelehre, 1942. :

7Lz. Sc~renk, O.:-...Grenzeschichtasbeaugung iAnd Senken?.eirkunG*,~ Z.A.M.M.,I
B d., 13, 1933, IP. 1830.

75. Schzrenk, 0.: Yjersuchle mit Abea~suglerl~igeln.` ZI.' M. I 31, Heft 9.
Luftfatfs~i lorachung 1935, P. 10.

NACA TM H~o. 1181

76. Ackeret, J.: Grentschichtabsaugu~ng. Zeitschrift des V.D.I.,
Ni. 35, p 1153, Sept. 1926.

77. Betz, A.: Verlaulf der S'rimoa~ngsgenchwinigke-it in der ~N he einer
He~nd bei anrsteti~ger ~Anderu~ng der- Klnrammun Luftf abr tfor~echung
1942, Liefg. by p. 129.

78. JTordEan P.: Auftri~ebs'6 TteChrecng rand Et~rj~~~lr~imngswreg beim
tfbe schr~eitenl des Max-rmalauftr~ieb;. Dilss. G~ttinoen 1939
(Lu~tf abrtf olschlung, Ed. 1.6, 1939, p 1 8e) .

79. Schiller, L.: V/erh. des 3. intern, Kong~. i'ir techn. Mechanik I,
p. 226, 1931. Z.A.M.M. 14, p. 36, 19314 Froc. of the $th
Intern. Co7ng:-- for Applied MeI~chanics, p. 315, 1938.

80. Ergebnisse der Aer;odyn. Versucheane~talt zu C-'i~ttingen,
Lieferunge~n L, iii, IV.

81. Schlmitz, E'. W.: Aerod~ynaikil des Phymuodells (Tr'jagfluigel-
messvungen I).

82. Jacobs, E. 1'., and ,Cheraer, A.: Ajrf'oil Eection Chaact~eristics
as Af'fect~ed bjy variations of the Reynolde Number. NACA
Rep. No. ;?6, 1437.

93. Schiller, L.: UsntersuchungryIn 'dbe;- lamilnare und tlurbulente
Stirrdmeng, Fo,sc~hunsarbeiten auf dem~ Gebiet des Ing.-Wesens
No. 268, 1922.

eB4. Hal~le--, ,. de: L'inrflue3nce des limited de la vei~ne fluide sur
les caracterijstiques Rerndynamniques d'une surface porthnte
Mittellungi des Inst. f. Perodynamik E-T.H ZiE ich, No. 3,

<8(5. Prandt1 7ndl Betz: tr A'bh~anilungen zur iydrodynamtk und Aerodynamik,

86. Rayleigh: pr'oc. Roy. Soc. A, 1916, p. 148.


Laumnar proftle Laminar profile'
Propeller profile No. 11 d/r = 0.10 In 0.40 1 f~rom the front dit = 0.140 in 0.44 t From one frons

Ro/t =0.008 f/t = 0.00525 In 0.50 t From the front f.'t = 0.0245 In 0.41 t from tne front














O 0.0030
.025 .0324
.05 .443
.1 .0600
.2 .0788
.3 .)892
.4 .0940
.5 .0936
.6 .0868
.7 .0736
.8 .0492
.9 .0)140
.95 .0040
1 0

0 0

1 0

R, 't = 0.009

Ro/L = 0.019

~To the laminar profile d..t 0-14 (fig. 121. According to measurements of F. Feldmami in one nigh-speed
trmwel of Ue hIstitute (description,61r. e.compression shocks for this profile start, For shokldess enonume, at
a Yach number M 0.71. For M 0.76 0.77 and shockless entrance the lift decreases and disturbances
in longitudinal srability appear. (Re = 570,000)

FUPLCA 'T'PI I'~o. 1 181

Slot Cp/2o/0 oox Op0/2o/oo OPg Cpx10-32 Op% Cq/Soloo Cpg Cp/2oloo Cpg

1 0.0752 1.102 0.0643 1.098 0.0552 1.106 0.0466 1.092 0.0637 1.111
2 .0376 1.143 .0318 1.127 .0267 1.142 .0233 1.110 .0308 1.153
3 .0342 1.105 .0293 1.096 .0243 1.117 .0216 1.083 .0286 1.125
4 .0342 1.078 .0294 1.074 .0254 1.074 .0220 1.063 .0292 1.079
5 .0352 1.165 .0303 1.149 .0258 1.150 .0220 1.119 .0298 1.166
6 .0453 1.118 .0395 1.107 .0336 1.102 .0290 1.089 .0389 1.115
7 .0344 1.120 .0283 1.105 .0237 1.097 .0207 1.085 .0288 1.109
8 .0288 1.123 .0242 1.106 .0198 1.092 .0173 1.059 .0226 1.117

Cqto/oo 0.018 0.556 0.4690 0.401 0.514

C~o/O .725 .613 .523 .441 .612

No. 14 16 17 18 15

qo 32.2 32.0 32.0 32.0 32.2
Uo 23 .73 23.63 23.66 23.63 88 .73
1/16.20 x 10-6 16.30X 10-6 16.15 X10-6 16.10 X10-6
Re 2.982 x106 2.960 x106 2.942X 106 2.964)( 106 2.908% 100
B9mm 0.494 0.416 0.372 0 .308 0 .456
for p/qo .036 .049 050 .056 .043

Slot C/2ofoo Cp CQ/2o oo, Pg Cq/2o o Cp C/2ofoo Cp cq/2o/oo CpB

1 0.0750 1.118 0 .1084 1.145 0.1282 1.165 0 .17 03 1.225 0.0912 1.128
2 .0351 1.182 .0534 1.243 .0551 1.276 .0843 1.386 .0449 1.207
3 .0335 1.146 .0491 1.190 .0591 1.207 .0773 1.276 .0412 1.167
4 .0338 1.083 .0498 1.094 .0598 1.101 .0770 1.116 .0419 1.088
5 .0353 1.191 .0521 1.259 .0628 1.299 .0817 1.390 .0429 1.222
6 .0462 1.125 .0672 1.160 .0809 1.194 .1042 1.263 .0560 1.140
7 .0341 1.125 .0481 1.175 .0581 1.211 .0740 1.369 .0399 1.149
8 .0286 1.140 .0393 1.229 .0497 1.297 .0602 1.444 .0352 1.179

Cq ofoo 0.644 0.936 1.130 1.458 0.786
C, o/oo .732 1.104 1.370 1.885 .908

Translator's note: A value of T2.2o o or Cq 10- 2 of 0.0752 denotes aleo

Tg of 0.0000752.

RACAL Th14 No. 1181

15.85 x10-6
3.015 x 106

16.10 x 10-6
2.998 x 106

16.25X 10-6
2 .981X 106


2.982 K 106


for p/qo

16.00<~ 10-6
3.OO4 x 106

qo 45.07 45.6 45. 0 45.1 45.3
U, 28.22 28.32 28.19 28.22 28.17
1/o 16.43 4 10-6 16.30 10-6 16.41 < 10-6 16.41 x 10-6 16.01< 10-6
Re 3.488 < 106 3.530/ 106 3.49 4 106 3.496 r 106 3.575 x 106
B 0.422 0.472 0.456 0.376 0.276
for p~m .035 .028 .032 .037 .049

Slot Cp/2o oo Cp C9/2ooo Op C9/2ofoo Cp C9./2o oo, Cp Cq/2o oo Cp

1 0.0807 1.107 0.0645 1.094 0.0713 1.097 0.1032 1.113 0.1554 1.152
2 .0402 1.155 .0317 1.127 .0358 1.134 .0506 1.174 .0764 1.232
3 .0866 1.122 .0292 1.098 .0327 1.101 .0470 1.121 .0696 1.146
4 .0369 107 .0292 1.070 .0330 103 .0467 1.080 .0691 1.089
5 .0388 1. 174 .0003 1.144 .0344 1.152 .0495 1.195 .0736 1.258
6 .0195 1.122 .0393 1.100 .0434 1.109 .0633 1.141 .0942 1.208
7 .0060 1.119 .0286 1.097 .0316 1.104 .0470 1.139 .0691 1.212
8 .0290 1.139 .0238 1.094 .0260 1.115 .0394 1.188 .0575 1.320

Cpto/oo 0.696 0.552 0.616 0.894 1.328
C o/ .785 .611 .641.018 1.585
Wg oo

No. 37 39 41 43 40

Qo 45.4 11.15 10.96 11.29 11.40
Do 28.20 13.93 13.81 14.01 14.12
V 16.01 x 10-6 16.00 x 10-6 16.00x 10-6 16.00 1 106 16.15 x 10-6
Re 3.580 r 106 1.768 v 106 1.753 106 1.779 r(106 1.775 r 106
emm 0.230 0.870 0.830 0.760 0.842
for p/q, .051 .001 .009 .019 .004

Slots Cp/2o oo Cp. Cgh/20 Cpb Cq//2o oop Cgq/2o, on p cq/2o. ooC

1 0.1958 1.202 0.0268 1.094 0.0428 1.111 0.0623 1.126 0.0314 1.099
2 .0963 1.297 .0144 1.115 .0213 1.140 .0006 1. 177 .0164 1.121
3 .0876 1.176 .0137 1.089 .0200 1.104 .0286 1.129 .0160 1.091
4 .0867 1.100 .0137 1.073 .0200 1.082 .0286 1.093 .0151 1.0715
5.0929 1.321 .0137 1.116 .0200 1.146 .0293 1.190 .0151 1.125
6 .1172 1.283 .0175 1.090 .0261 1.118 .0385 1.135 .0203 1.105
7 .0862 1.289 .0138 1.083 .0200 1.105 .0274 1.136 .0163 1.080
8 .0752 1.463 .0113 1.071 .0157 1.099 .0226 1.136 .0124 1.081

C4 ofoo 1.676 0.250 0.372 0.536 0.286
C o/oo 2.104 .273 .414 .610 .315

N\ACA. TT I 11. 1181

No. 44 46 45 42 32
go 11.31 11.49 11.30 11~.37 45.3
U, 14.03 14.13 14.03 14.04 28.22
u' 16.00 x10-6 16.00 x 10-6160 16.0016.00x106 16.30 x*10-6
Re 1.780 x106 1.793 w 106 1.780 x106 1.783 x 106 3.515 x 106
8mm 0.700 0.598 0.668 0.780 0.486
for p/qo .027 .048 .035 .015 .024

Slot Cd2/2ooo Op, CP/aooo OPg C/2o oo Cpg Cq/2o/oo Opg G4/2o/oo OPg

1 0. 0812 1.143 0.1258 1.184 0.1010 1.162 0.0514 1.117 0.0571 1.094
2 .0399 1.217 .0618 1.316 .0500 1.260 .0257 1.160 .0282 1.122
3 .0374 1.153 .0573 1.209 .0464 1.177 .0245 1.117 .0258 1.097
4 .0377 1.104 .0579 1.129 .0461 1.115 .0238 1.086 .0263 1.066
5 .0093 1.239 .0601 1.347 .0484 1.285 .0245 1.164 .0271 1.133
6 .0508 1.157 .0772 1.216 .0629 1.181 .0316 1.123 .0352 1.088
7 .0363 1.174 .0584 1.265 .0457 1.212 .0232 1.119 .0256 1.089
8 .0292 1.186 .0466 1.326 .0373 1.248 .0183 1.112 .0220 1.070

Cgto/oo 0.702 1.088 0.872 0 .444 0.494
C, o/oo .822 1.352 1.048 .501 .543

No. 55 54 35 -56 57

Qo 45.4 32.5 45.5 45.1 57.8
U 27.97 23.64 28.23 27.87 31.73
Vo 15.85 X 10-6 15.8 x 10-6 16.01 x 10-6 15.85 #10-6 16.08KX10-.6
Re 3.583 x106 3.030 x106 380 06 3 .570 106 4.216
inn 0.420 0 .440 0.316 0.328 0.442
for p/qo .034 .038 .046 .045 .081

Slot C/2o o C~p C/2ooo Cp~ Cd2/2ooo CPg CZ/2oQo OPg Cdo oo CPg
1 0.0792 1.095 0.0931 1.110 0.1284 1.126 0.1228 1.118 0.0678 1.087
2 .0400 1.133 .0463 1.166 .0638 1.193 .0610 1.187 .0386 1.112
3 .9036 1.097 .0426 1.117 .0587 1.128 .0559 1.128 .0308 1.084
4 .0366 1.073 .0480 1.081 .0581 1.084 .0559 1.083 .0310 1.064
5 .0384 1.157 .0447 1.193 .0618 1.219 .0590 1.215 .0326 1.128
6 .0487 1.115 .0573 1.133 .0775 1.169 .0748 1.154 .0423 1.090
7 .0344 1.112 .0407 1.141 .0673 1.172 .0536 1.164 *0805 1.086
8 .0298 1.123 .0353 1.166 .0469 1.238 .0456 1.234 .0268 1.072

NACA TM No. 1181

CQ o/oo
C, o/oo






.916 1.284 1.217


No. 58 60 61 5952
4, 57.9 ,58.3 58.1 57.9 78.9
U, 31.75 31.71 31.68 31.76 37.02
V16.03 r 10-6 15.90 x 10-6 15.90 x 10-6 16.03 x 10-6 16.06 x 10-6
Re 4.02 '106 4.05 x 106 4.045 x 106 4.02 x 106 4.68 r 106
8,, 0.385 0.276 0.218 0.327 0.219
for p/qo 0. 036 .047 .053 .043 .047

Slot Cd2oloo Cp Cqh/2olo Cp6 Cp/2oloo Cp Cq/2o/oo Cp Cd/2ofoo Cp

1 0.0786 1.090 0. 1338 1.L127 0. 1708 1.163 0.1025 1.100 0. 1452 1.132
2 .0090 1.122 .0662 1.207 .0818 1.258 .0515 1.147 .0649 1.186
3 .0360 1.091 .0606 1.129 .0760 1.157 .0468 1.100 .0645 1.121
4 .08163 1.069 .0604 1.079 .0753 1.087 .0464 1.072 .0637 1.076
5 .0373 1.144 .0634 1.232 .0863 1.308 .0490 1.171 .0728 1.207
6 .0483 1.105 .0822 1.152 .1091 1.204 .0628 1.127 .0922 1.166
7 .0048 1.098 .0601 1.161 .0775 1.225 .0437 1.126 .0653 1.165
8 .0005 1.099 .0511 1.224 .0647 1.336 .0385 1.163 .0542 1.226

C to/,, 0.682 1.154 1.480 0.882 1.246
C o(/oo *749 1.338 1.791 .989 1.443

No. 64 70' 65" 661 71'

qo 79.3 109.1 79.5 79.6 109.9
U, 37.27 44.07 37.32 37.34 43.91
vo 16.32 x 10-6 16.80 x 10-6 16.32 r 10-6 16.32 w 10-6 16.38 x 10-6
Re 4.63 r 106 5.33 x106 4.64 *106 4.645 *106 5.45 r106
B, 0.250 0.231 0.294 0.328 0. 186
for p/qo .044 .038 .041 .036 .039

Slot Cph/2ooo Csg Cda/2oo Cpp Cq/2ofoo Cpg Cq/2o/oo Cp cd2/ofo Cp

1 0.1216 1.111 0.1356 1.114 0.1017 1.099 0.0895 1.093 0.1903 1.151
2 .0544 1.252 .0611 1.126 .0464 1.132 .0411 1.120 .0833 1.166
3 .0544 1.106 .0626 1.090 .0462 1.095 .0406 1.088 .0864 1.103
4 .0537 1.074 .0632 1.069 .0460 1.071 .0400 1.068 .0868 1.079
5 .0610 1.199 .0678 1.161 .0512 1.175 .0458 1.157 .0938 1.211
6 .0772 1.143 .0857 1. 126 .0651 1.123 .0575 1.110 .1204 1.177
7 .0542 1.133 .0600 1.110 .0653 1.109 .0403 1.102 .0860 1.167
8 .0444 1.173 .0496 1.166 .0383 1.130 .0334 1.101 .0690 1.262

N~Polit ICA TJo. 1181

C ofoo











Transition 10 milllmeters ahead of trailing edge, Individual turbulent
2Indlividual turbulent bursts further Forward.
3Frequentt tuirbuilent bursts Further forward.

Lbursts earlier.

4Transition 5 mrillimeters ahead of trailing edge, individual turbuilent bursts earlier.

0.610 5** 84.2
1.600 1.6809
1.936 9.085
0.500 0.500

2*526 9.595



Nlo. 30.2


No. 32.8
0.90 0


On both sides lamina~r up to the traillag edge

NACA TM No. 1181

qo 69.8 69.5 85.0 5 4.8 103.8
ye 35.15 35.3 39.0 31.2 42.8
v 16.60 10-6 16.88 X 10-6 16.98 x 10-6 16.55 x 10-6 16.45 M 10-6
Re 2.188 x 106 2.160 x 106 2.382 x 106 1.946 X 106 2.690 x 106
C, 0.232 0.232 0.232 0.232 0.232

Slot Qg 410-3 Cgp Qg x 10-3 Opg Qg X 109 OPg 4, x lo** Cp qax 10"














0. 548







C oi/oo

C1 n

0.784 No.
0.506 21.1 21.2 21.3
1.290 1.340 1.470 1.572
1.557 1.622 1.790 1.928
0.838 0.755 0.678 0.657

2.395 2.377 2.468 2.585





I 1 I I

Stat Qg a10-3 Cp 9ga0~ xp IO 0ra 9 43 c pa x 10-3 Cp Qa g 10" Cp

10 0 0 0 0
2 0 10 0 0 0
3 0. 192 1.467 ~. 168 1.480 0.395 1.624 0.488 1.600 0.302 1.615
4 0. 1471 1.527 0.131 1.562 0.2718 1.635 0.040 1.654 0-216 1.635
5 ~. 193 1.501 0.170 1.534 0.350 1.7 .440 1. 580 0.285 1.574
6 0.335 1.448 0.296 1.467 0.566 I.511 0.694 1.511 0 .449 1.505
7 0.398 .5 0.356 1.480 0).691 1.47 .640 1.490 0.542 1.487
8 0.297 1.360 0.260 1.378 0.472 1.385 0.574 1.085 0.374 1.085
9 0.428 1.2901 0.376 1.304 0.6BS 1.311 0.822 1.318 0.543 1.307
10 0.254 1.229 0.218 1.246 0.393 1.236 .421.235 0.323 1.236
II 0.403 1.145 0.323 1.16 0.616 1.154 0.754 I.155 0.496 1.155
12 0.253 1.126 0.210 1.146 0.288 1.115 01.350 1.118 0.267 1.124
13 0.300 1.036 0).2701 1.053 0.40j3 1.019 0.484 1.027 0.361 1.027
14 0.400 0.981 0.431 1.019 0.335 0.939 0.40)0 0.9411 0.386 0.965
I' 0.048 1.189 0.038 1.212 0.025 1.025 0.032 1.028 0.022 1.023
2' 0.146 1. 172 0.124 1.203 0.121 1.049 0.292 1.048 0.095 1.048
3' 0.104 1.136 0.090 1.156 0.064 1.05rl 0.076 1.051 0.055 1.053
4' 0.093 1.138 0.081 1.162 0.6*1.052 0.072 1.053 0 .053 1.053
5" 0.170 1.113 0.148 1.135 0.099 1.0)58 .6 1.049 0.080 1.056
6' 0.228 L.119 0.196 I.143 0.149 1.043 0.180) 1.043 0.119 I.01;2
7' 0.281 1.078 0.242 1.095 0.217 1.028 01.262 1.028 0.176 I.027
8' 0.366 1.018 0.316 1.0)35 0.355 0.981 0.410 0.979 0.310 0.982
9' 0.256 1.019 0.220 1.043 0.262 0.979 0.014 0.978 0.239 0.989
10' 0.726 1.006 0.642 1.0301 0.;56 0.980 0.858 0.954 0.708 0.975



16.45 < 10-6
1.965 r 106

10.52 L 10-6
2.206 s 106

16.60 10-6
1.639 1 106

C coloo 0.733 0.816
Cp 0.492 No. 35.2 0.534
C 1.225 1.300 1.050
C 1.454 1.555 1.632

I 1.020 0.910 1.110
C 2.472 2.465 2.742

On both sides larmnar up to the trailing edge

0.940 1.062 1.000
0.362 0.306 0.331
1.302 1.468 1.380 1.310
1.618 1.828 1.7371 1.681

0.748 0.668 0.693 0.707
2.066 2.496 2.430 2.368

No. 38.2 38.3 38.4

0.398 No. 30.2 39.3
1.412 1.375 1.486
1.761 1.710 1.857

0.657 0.692 0.8635
2.418 2.402 2.492



NACA TM No. 1181

16J.62 I 10-6
1.3009 r 100

16.55 10-6
1.655 r 106

Qo 1 25.5 55.8 71.80 39.9 25.8
Vo 21.2 31.4 35.70 26.46 21.24
S16.36 x10-6 16.47 x 10-6 16.60 x 10-6 16.25 Xt10-6 16.25x 10-6
Re 1.337 x106 1.968 w 106 2.220 x106 1.682 X 106 1.352 X106
C, 0.372 0.480 0.480 0.480 0 .480

Slot Q x 10-3 Cpg Qgx10-3 Cpg Qax10-3 Epb garO~ x 0 pg ga108 Cpg

1 0 0 0 0 0
2 0 0 10 0 0
3 0.218 1.617 0 .461 1.736 0.530 1.718 0.354 1.786 0.257 1.726
4 .157 1.657 .324 1.769 372 1.786 .251 1.749 .184 1*745
5 .208 1.593 .439 1.668 .500 1.645 .845 1.665 .258 1.666
6 .338 1.510 .662 1.577 .754 1.563 .518 1.568 .892 1.564
7 .406 1.505 .805 1.558 .9)20 1.536 .610 1.549 .471 1.552
8 .288 1.397 .548 1.431 *628 1.418 .430 1.429 .885 1.427
9 .411 1.015 .786 1.353 *894 1.842 .622 1.348 .470 1.847
LO .245 1.246 .465 1.368 .526 1.255 .367 1.266 .281 1.268
11 .374 1.163 .630 1.166 .718 1.159 .494 1.165 .870 1.171
12 .215 1.140 .310 1.184 .345 1.122 .263 1.136 .220 1.166
18 .280 1.047 .426 1.045 .454 1.0833 .370 1.046 .300 1.058
14 .392 1.008 .822 0.940 .330 0.934 .404 0.964 .400 1.004
1' .015 1.036 .017 .958 .020 .940 .DU@ *945 .000 0.950
2' .069 1.058 .082 .993 .092 .988 .068 .990 .048 .994
3' .020 1.061 .050 .933 .057 .987 .036 ,996 .027 ,990
4' .018 1.062 .031 1.011 .036 1.005 .017 1.010 .011 1.010
5' .060 1. 056 .063 1.007 .074 1.001 .053 1.013 .043 1.018
6' .090 1.047 .100 1.015 .115 1. 010 .078 1.013 .061 1.013
7' .138 1.033 .125 1.007 .141 1.004 .120 1.005 .008 1.000
8' .256 0.991 .227 0.967 .257 0.963 .280 0.967 .249 0.976
9' .195 1.005 .219 *&75 .248 .9&D .232 *978 .193 *988
10' .580 0.988 .794 .948 .887 .943 .660 .953 .585 *904




C4, "

C, "
C, "

.364 No. 41.2
1.260 1.315
1.551 1.624
1.13 1.10
2.681 2.724

0.293 No. 43.2 43.3
1.350 1.390 1 .4300
1.770 1.828 1.882
0.735 0.695 0.660
2.505 2.528 2.542


1.020 No. 45



0.680 .

On both sides laminar up to the trailing edge

NACA TM' No. 1181

NACA TM No. 1181

No. 54 55.3 56.1 57.3

No. 55.2 55.4 0.264 No. e
1.430 1.405 1.480 1.550 1.C
1.985 1.936 2.045 2.154 2.5
0.95 0.97 0.92 0.730 0 .7
2.935 2.906 2.965 2.884 2.5

;h sides laminar up to the trailing edge


4s 02 ga Go Q cs co e t

Figs. 1,2

NACA TML No. 1181

Figure 1.- Influence of the transition-point position on the profile drag

for various profile thicknesses; Re = 15 x 106.

Position of
~t~s~ Itransition point

Laninar flat plate 5 eg


240' sto I a

Figure 2.- Influence of the transition-point position on the profile drag
for various Re; d/t= 0.16.

NA.CA TIM No. 1181 Figs. 3,4



the chord on the profile
(The pressure distributions

0 01 0,2 03
NAlCA 0010 urb

Figure 3.- Pressure distributions along
NA~CA? 0010 for various Re; ca = 0.

for the diifferent
by Ap/g = 0.1.)
t = 0.60 m.

Re are every time shifted vertically downward
The start of transition is denoted by arrows,

Figure 4.- Laminar profile d/t = 0.140 (fig. 12), t = 0.70 m. Pres-
sure distributions along the chord and transition start (vertical
arrows) for smooth inflow. The lower figure shows the pressure
distributions in the rear part of the profile for various Re on
an enlarged scale.

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