Effects of wing and nacelle modifications on drag and wake characteristics of a bomber-type airplane model

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Material Information

Title:
Effects of wing and nacelle modifications on drag and wake characteristics of a bomber-type airplane model
Alternate Title:
NACA wartime reports
Physical Description:
27, 72 p. : ill. ; 28 cm.
Language:
English
Creator:
Neely, Robert H
Fairbanks, Richard W
Conner, D. William
Langley Aeronautical Laboratory
United States -- National Advisory Committee for Aeronautics
Publisher:
Langley Memorial Aeronautical Laboratory
Place of Publication:
Langley Field, VA
Publication Date:

Subjects

Subjects / Keywords:
Airplanes -- Nacelles   ( lcsh )
Bombers   ( lcsh )
Aerodynamics -- Research   ( lcsh )
Genre:
federal government publication   ( marcgt )
bibliography   ( marcgt )
technical report   ( marcgt )
non-fiction   ( marcgt )

Notes

Summary:
Summary: An investigation of a model of a large four-engine bomber was conducted in the Langley 19-foot pressure tunnel to determine the effects of several wing and nacelle modification on drag characteristics and air-flow characteristics at the tail. Leading-edge gloves, trailing-edge extensions, and modified nacelle afterbodies were tested individually and in combination. The effects of the various modifications were determined by force tests, tuft observations, and turbulence surveys in the region of the tail. Tests were made with fixed and natural transition on the wing and with propellers operating and propellers off. Most of the tests were conducted at a Reynolds number of approximately 2.6 x 10⁶. The results indicated that application of certain of the modifications provided worth-while improvements in the characteristics of the model. The flow over the wing and flaps was improved, the drag was reduced, and the turbulence in the region of the tail was reduced.
Bibliography:
Includes bibliographic references (p. 22).
Statement of Responsibility:
by Robert H. Neely, Richard W. Fairbanks, and D. William Conner.
General Note:
"Report no. L-114."
General Note:
"Originally issued October 1944 as Advance Restricted Report L4G15."
General Note:
"Report date October 1944."
General Note:
"NACA WARTIME REPORTS are reprints of papers originally issued to provide rapid distribution of advance research results to an authorized group requiring them for the war effort. They were previously held under a security status but are now unclassified. Some of these reports were not technically edited. All have been reproduced without change in order to expedite general distribution."

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University of Florida
Rights Management:
All applicable rights reserved by the source institution and holding location.
Resource Identifier:
aleph - 003633521
oclc - 71753149
sobekcm - AA00006268_00001
System ID:
AA00006268:00001

Full Text

Nfi-,


-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS





WART' IME REPORT
ORIGINALLY ISSUED
December 1945 as
Advance Restricted. Report L5J05

EFFECTS OF WING AND NACELLE MODIFICATIONS ON
IRAG AND WAKE CHARACTERISTICS OF A
BOMBER-TYPE AIRPLANE MOIEL
By Robert H. Neely, Richard. W. Fairbanks,
and D. William Conner

Langley Memorial Aeronautical Laboratory
Langley Field., Va.








A .. '


WASHINGTON
NACA WARTIME REPORTS are reprints of papers originally issued to provide rapid distribution of
advance research results to an authorized group requiring them for the war effort. They were pre-
viously held under a security status but are now unclassified. Some of these reports were not tech-
nically edited. All have been reproduced without change in order to expedite general distribution.


L 114


DOCUMENTS DEPARTMENT


ARR No. L5J05





































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


http://www.archive.org/details/effectsofwingnac001ang








7TACA ARR No. L5J05 RESTRICTED

"ATTOV'AL ADVISORY.: COMMITTEE FC?3 LEROL-JTICS


ADVANCE RESTRICTED REPORT


EFFECTS OF '.ING AMD IJACELLE ?ODIFI ATIOiT3 01

DRAG, ATrD '.-J"E CHARACTER IST ICT OF A

MEByER-T YPE J.IRPLtJE MODEL

Ly RoObrt H. Ieely', Richard ;,. Fairbanl:s
and D. ;illiam Conner


S U i, M A R Y


An investigation of a model of a large four-engine
bomber was conducted in the Langley 1o-fooc pressure
tunnel to determined the effects of several wing and
nacelle modifi'atio-ns on drag characteristics and air-
floa characce:riL tcs at the tail. Leadin.-edgo gloves,
trailing.-ed.,e *:.xt'nsicns, and modified nacelle afterbodies
were tested. ir.:iivdually and in combination. The effects
of the varic... r-olifications wver6 deccrmined by Izorce
tests, tuft o, servations, and turbulence surveys in the
region of the -ail. Tests were made with fixed and
natural transition on the w'ing and with propellers oper-
ating and proplirs off. M.'ost of the tests were con-
ducted at a Re-nolds number of a proximate ly 2. x lO'.

Th3 results indicated that application of certain
of the modifications provided worth-w-hile improveme-ints
in the c'-iara _,2ristics of ti-. r:iodel. The flow ov.r the
wing and flaps was irmprovcd., the d.ra was reduced, and
the turbu.len.nc in cnc r,_ion of the tail was reduced.

Trailing-edge extensions were the most effective
individual rodifi cazion In ir--provin- the flo over the
wing wit-h viAr. flaps ne'u.tral, cc,,l and intercooler flaps
closed. [1'cdified nacellc afterboJ.i-s weree th: mIr.os
eff-ctiv'-j individual modification in reducing drag with
either fixed or r-:ati.ural transition on the wing; how 'ever,
trailin--e-d- extinsion.s were slic htly more effective
with f:-ied tr aisicion. Combinations of either leading-
or traiL ng-e g3e e:,-.cens ons and i:.odified afterbodies were
more effecti-: thai either rmcdifli.ation alone. With cowl
and irter-cooler flaps open, urailin:-edge extensions with
modified afterbodies provided substantial improvement in


RE STRICTED










flow and drag characteristics. With wing flaps deflected,
enclosing the flap behind the inboard nacelle within an
extended afterbody or cutting the flaps at the nacelle
appeared to be the most promising methods of improving
the flow over the flaps and the tail. Although the
results of hot-wire-anemometer surveys were not conclusive
in regard to buffeting characteristics, the modifications
did reducce the turbulence at the tail with wing flaps
both net-ral and deflected. The modifications, as a rule,
were favorable to maximum lift. Appreciable reductions
in longitudinal stability of the model were caused by
addition of leading-edge gloves and trailing-edge
extensions.


I N T R ODUCT I ON


Separation of flow over a wing increases the drag
and has, in a number of instances, caused tail buffeting
because of the irregular nature of the flow at the tail.
Several wing and nacelle modifications, designed with a
view to improving the flow over the wing, were tested on a
model of a large four-engine bomber to detr-in-.e the
effects on drag characteristics and air-flow character-
istics at the tail. Leading-edge gloves, wing trailing-
edge extensions, and modified nacelle afterbodies were
tested. The characteristics of the basic and modified
model were determined by tuft observations, force meas-
urements, and measurements of turbulence and dynamic
pressure in the vicinity of the tail. Turbulence was
measured by means of a hot-wire anemometer. The hot-wire-
anemometer equipment was furnished by the California
Institute of Technology and was operated under the direc-
tion of Dr. Hans W. Liepmann of its staff. The investiga-
tion was conducted in the Langley 19-foot pressure tunnel.


SY B OLS


The coefficients and symbols used herein are defined
as follows:

CL gross lift coefficient (L/qS)


increment of maximum lift coefficient


ACLmax









ITACA ARR No. L5J05


CD drag coefficient corrected for jet-boundary
interference (D/qS)

C0 gross pitching-momert coefficient (M./qS')

CD induced-drag coefficient (C- A/A)

CDp parasite-drag coefficient (CD CDi)

ACDp increment of parasite-drag coefficient

Te thrust disk-loading coefficient for one
propeller (Te/pV2D2)


rv root-mean-square value of the deviations, per-
pendicular to wind ax:is, of instantaneous
local velocity from its mean value

U nean value of instantaneous local velocity
along wt-nd axis

Rt local dmnaiic rresslure at tail

q free-stream dynamic pressure ( -V.V2

a angle of attack of wing corrected for jet-
boundary interference

6f wing-flap deflection measured from neutral flap
position

R Reynolds number (p',./i)

where

L lift

D drag; propeller diameter

M pitching moment about center of gravity

Te effective thrust of one propeller

S wing area (22.219 sq ft)








NACA ARR No. L5J05


c mean aerodynamic chord (1.l4O ft)


A geometric aspect ratio = 12.8)

b wing span (16.875 ft)

V velocity in free stream

p mass density of air
Coefficient of viscosity


MODEL


The general arrangement of the model is shown in
figures 1 and 2. The model was of wood and metal con-
struction and was finished with lacquer. Pertinent
dimensions are given in table I.

The wing had Davis airfoil sections, 22.9 percent
thick at the root and 9.5 percent thick at the tip. The
maximum camber was approximately 5.4 percent of the chord
at the root section and 1.4 percent at the tip. The
location of the maximum camber was constant across the
span at 31.5 percent of the local chord. The geometric
aspect ratio of the wing was 12.8 and the taper ratio
was 3.077:1.

The horizontal tail had no movable elevator. The
vertical tail was off during all the tests.

The installation of the cowl flaps and intercooler
exit flaps is shown in figure 3 for their deflected posi-
tions. The cowl flaps extended from the top of the
cowling to a point slightly below the nacelle center line.
The intercooler cooling-air exits were located on the
upper surface of the wing.


STANDARD CONFIGURATION

The standard nacelle afterbodies (afterbody 1) are
shown in figures 3 and I with wing flaps neutral. No
part of the nacelle afterbody is on the upper surface of








NACA ARR No. L5JO5


the ".'ing near the trailing edge. The extreme and of the
standarJ. afteroodies ,was attached to the lower surface
of the wing flans and deflected with the flaps, as shown
in figure 5. The wing flaps were of the Fowler type and
consisted of inboard and outboard sections. Ordinates
for the inboard section are given in table II. At the
inboard nacelle, the nose of the flap was 2.4 percent
wing chord ahead of and 2.9 percent wing chord below the
trailing edge of the wing.


rODIFICATIO;IS


In an attempt to delay separation of air flow on
the wing in the cruise condition, the original wing chord
was extended in order to reduce the peak pressures and
adverse pressure gradients. Leading-edge gloves (figs. 6
and 7) were built to rACA OL-series ordinates modified to
fair into the original Davis airfoil section. The gloves
extended the wing chord 10 percent between the fuselage
and the inboard nacelle and tapered to the original
leading edge at the outboard nacelle. Because these
gloves were added to the original wing, a perfect contour
could not be formed where the gloves faired into the wing.

The trailing-edge extension was a thin metal strip
attached to the flans and deflected down 110 from the
lower surface of the flaps. The extensions were deflected
down 10 to 20 from the wing chord lines. The installation
of the extensions is shown in figure 7. Extensions of
two different spans and three different chords were tested.
Extensions attached only to the inboard flaps are desig-
nated 0.3' span, and extensions attached to both inboard
and outboard flaps are designated 0.6 span. The chords
were 1-, 2 and 5 inches. Unless otherwise specified,
2' 1
the term "trailing-,;dge extensions" designates the 1r-inch-
chord extensions attached to both the inboard and outboard
flaps.

The installation of the various modified nacelle
afterbodies with wing flaps neutral is illustrated in
figures S to 10. Drawings of the modified nacelle after-
bodies are given in figures 11 to 13. Afterbodies 2
and 5, shown in figure 8, were attached to the inboard
nacelles only and differed mainly from the standard
afterbody in that they had a fairing on the upper surface








NACA ARR No. L5JO5


of the wing. Afterbody 4 (figs. 9 and 12) was a beaver-
tail afterbody with a small fairing on the upper surface
of the wing. Afterbody 5, shown in figures 10 and 15,
was faired on the upper surface-of the wing forward to
the intercooler air exit. The -lower part was extended
in order to obtain a better afterbody shape.

The installation of the deflected flaps with inboard
nacelle afterbodies L and 5 is shown in figures 14
and 15, respectively. With afterbodies 1, tests were
made with the flaps cut out below the afterbody as shown
in figure 14 and also with the flaps not cut and extending
below the afterbody. Inboard afterbodies 5 enveloped the
center part of the deflected flaps, and the flap-nacelle
juncture was faired with plasticine as shown in figure 15.
The tips of the modified afterbodies did not deflect with
the flaps.

A double slotted, or vaned, flap (fig. 16) was
tested in an attempt to improve the air flow over the
flap. The outside contour and the installation of the
flap and vane combination were the same as for the
original Fowler flap. The ordinates for the double
slotted flap are given in table II.


APPARATUS AND TESTS


Tests of the model were made for two basic conditions:

(1) Cruise condition wing flaps neutral,
cowl and intercooler flaps closed

(2) Landing condition wing flaps deflected 4o0,
landing gear down, cowl and intercooler
flaps closed

For the cruise condition, a few tests were also made to
determine characteristics with cowl and intercooler flaps
open.

For the cruise condition, tests were made with the
model in the standard configuration and with leading-edge
gloves, trailing-edge extensions, and modified nacelle
afterbodies. For the landing condition, tests were made
with the model in the standard configuration and with
modified inboard afterbodies and modified inboard flaps.








?TACA ARR No. L5J05


The cc;::l flaps in the ope.n position were deflected 10.
.'hien the intercooler flaps were opened, the exit gap was
increased 7/16 inch, which corresponds to the maximum
deflection of the intercooler flaps. The air flow through
the nacelle was adjusted, with the model at an angle of
attack of 50, to provide a pressure drop of approxi-
mately 0.75q throu,:- the cowling with cowl flaps open
and a pressure drop of approximately 0.-'q through the
intercooler ducts with exit flaps closed.

For a flap deflection of 400, the main landing gear
was down; however, no provision was r:ade for simulating
open landing-gear doors. For' flap deflections of 0
and 100, the landing gear w.as removed. rost of the tests
were made w-ith the horizontal tail off. The vertical
tail was off for all tests.

Power conditions were simulated by matching the
thrust coefficients of the model and airplane at each
lift coefficient. The variation of thrust coefficient
with lift coefficient for sea-level power conditions is
presented in figure 17. The thrust coefficients for 0.1
normal rated power at sea level correspond closely to
those for cruising power (C.6 normal rated power) at
25,000 feet. The )ropeller blade angle at 0.75 radius
was 50.

It was believed that transition from laminar to tur-
bulent flow on the airplane wing /would occur at approxi-
mately the location of the front spar. In an attempt to
make the results of the model tests more representative of
flight conditions, most of the tests were made with the
transition fixed at a chordwise station corresponding to
the spar location. The transition was fixed by placing
a strip of 60-grain carborundum on the upper and lower
surfaces at the 10-percent-chord station of the original
wing section. The width of the strip between the fuse-
lage and the outboard nacelles was approximately 5/8 inch
and tapered to approximately 1/4 inch at the tips.

The character of the flow over the wing and the
nacelle afterbodies was determined by observing the
behavior of tufts, which were attached to both the upper
and lower surfaces of the wing and nacelle afterbodies.
No tufts were placed ahead of the 20-percent-chord sta-
tion of the wing. These tests were generally made with
fixed transition and vilth propellers operating at
0.4 normal rated power; however, a few tests were made
with natural transition and with propellers off.








NACA ARR No. L5J05


Force and moment characteristics were measured by a
six-component automatically recording balance system.
Lift and drag were measured for all configurations. The
effects of leading-edge gloves and trailing-edge exten-
sions on the longitudinal stability characteristics were
determined with the horizontal tail on (elevator neutral)
for several power conditions.

Measurements of the air-stream turbulence were made
at several spanwise stations along the elevator hinge
line by means of a hot-wire anemometer. The basic prin-
ciples of operation of this instrument are described in
reference 1. The turbulence measurements were supple-
mented by measurements of the local dynamic pressures at
the tail obtained from surveys with a rake of six pitot-
static tubes. All surveys were made with fixed transition
and with the propellers operating at 0.4 normal rated
power.

All tests were made with the air in the tunnel com-
pressed to an absolute pressure of approximately 35 pounds
per square inch (p z 0.00558 slug/cu ft). Most of the
tests were made at a Reynolds number of approxi-
mately 2.6 x 10 and a Mach number of 0.12; however, a
few tests were made at a Reynolds number of 5.9 x 10
and a Mach number of 0.18.


RESULTS AND DIS CUS S I ON


The results of the investigation are discussed from
the standpoint of (1) flow over the wing and flap,
(2) flow at the tail, (3) drag and lift, and (4) longi-
tudinal stability. The characteristics of the standard
model and the effect of the various modifications are
shown for the cruise and landing conditions.

Jet-boundary corrections have been added to the
angle of attack, drag coefficient, and the vertical posi-
tion of the survey points with respect to the elevator
hinge line as follows:

Aa = 0.654CL

ACD = 0.0106L 2


Az = -0.153CL









NACA ARR No. L5J05


where aa is in degrees, Lz is in inches, and z is
the vertical position of the survey points with respect
to the elevator hinge line. Hio corrections have been
applied to the data for the effects of model-support tare
and interference or for air-stream misalinement.


FLO V C-HARACTERISTICS AT THE WI.TrG


The results of tuft observations that show the char-
acter of the flow' over the wing, flaps, and nacelles are
given.


Cruise Condition

"he stall progressions of the model in the original
and modified configurations with wing flaps neutral are
given in fiEures It to 22 for covwl and intercooler flaps
closed. The values of lift coefficient at which separa-
tion first occurred on the wing are riven in table Ill.
Diagrams showing the flow over the wing at a lift coeffi-
cient of approximately 0.8 are presented in figure 25.
Stall progressions for the model with corl and intercooler
flaps open are given in figure dh. The propellers were
operating at 0.L normal rated power and the transition
was fixed at 10 percent wing chord except for two te-ts
of the ssa-cdar-i icc.- i Diagrams are presented to show
the effect cf over :nd the effect of removing che transi-
tion strip cn the wing.

St.;ndard c-,onfi,'-ur:t:1on.- .iith co'.l and intercooler
flaps clo e', npra ..1- operating at C..L n-crmrl rated
power, ani tranE.siti :t :cd, thCe initial sull on the
wing occi.urred at a c-: coffic ent of abor. 0,0> with
the rmIo6. in th- st:::.-d or i rginal c n.fi- u'ation
(fi:, 1 :'a a) The i.i.ial ser -_.cion occur' : L J ,n t!1e
rear part of the ,ing io the left of each ::- 1 ... The
flowI over the -jinr directly behind each nacelle was rough
but not separated for most angles of attack. .'ith
natural transition (fig. l6(b)) the stall patterns were
about the same as with fixed transition but the initial
separation occurred at a lift coefficient of about 0.91.
With propellers off and transition fixed, the initial
separation occurred at about the same lift coefficient
as with the propellers operating, but at higher lift
coefficients the area directly behind the inboard nacelles
was stalled (fig. 187c)).








NACA ARR No.. L5J05


Separated flow like that indicated in figure 18
produces an increase in drag and could cause buffeting.
The purpose of the modifications was to delay this sepa-
ration and to cause a general improvement in flow through
the cruising range (CL = 0.6 to 0.9). A substantial
improvement in flow over the wing was obtained, as shown
in figure 25(a), by deflecting the wing flaps 100. The
drag, however, was increased.

Chord extensions.- Chord extensions delayed the
initial separation and improved the flow over the wing
at higher lift coefficients.

Leading-edge gloves (fig. 19(a)) delayed the initial
separation to a lift coefficient 0.12 higher than for the
standard model. About the same improvement was realized
with the 11-inch-chord 0.3-span trailing-edge extensions.
2
The 0.6-span trailing-edge extensions were the most
effective individual modification in improving the flow
over the wing. The Il-inch-chord 0.6-span trailing-edge
2
extensions (fig. 19(c)) delayed the initial separation
to a lift coefficient 0.21 higher than for the standard
model. More improvement resulted from the 2 -inch-chord
0.6-span extensions. Subsequent tests, however, were
made with the l1-inch-chord 0.6-span trailing-edge exten-
sions because the greater improvement in flow with the
21-inch extensions did not seem sufficient to warrant
the additional structural changes necessary to the air-
plane. In evaluating the improvement due to either
leading-edge or trailing-edge extensions in terms of the
increase in lift coefficient at which separation first
occurred on the wing, it should be noted that the gain
in lift coefficient was partly due to added wing area.

Modified nacelle afterbodies.- Modified nacelle
afterbodies caused only a slight delay in the initial
separation but improved the flow over the wing. This
improvement is shown for afterbodies 2 and 3 by comparing
figure 20 with figure 18(a) and for afterbodies 4 and 5
by comparing figures 21(a), 21(b), 22(a), and 22(b) with
figures 19(a) and 19(c). Afterbodies 4 and 5 appear to
be most effective. No flow separation occurred on the
lower surface of either the standard or. modified nacelles.









NACA ARR No. L5J05


Comb nations of chord extensions and rricdified nacelle
afterbodies.- Chord extensions, either leading: edge or
trailing edge, in combination ;ith :odified nacelle
afterbodies were more effective in delaying. separation
and in improving the flo'.- over the winog than were chord
extensions or afterbodies alone, The cojbinations of
leading-edge gloves with afterbodies 2 and leading-edge
gloves with afterbodies 5 (figs. 21'a) and 22(a') delayed
the initial separation to a lift coefficient a.proxl-
mately 0.16 higher than for rlhe standard model. 4 ith
trailinz-edi-e extensions in combination with either
afterboiies 4 cr c5 .'- 1-s. 21(b) an 22(b)), the initial
separation occurred at a lift coefficient of about C.-l1,
which is approximately 0.2, higher than for che standard
model. As shown in fi.sur-e 21(c), a greater improvement
in flow was obtained with a combina.Lion of leading-edge
gloves, trailino-edge extensions, and afterbodies 4.
Separation on the inner 'ing sections was delayed to a
lift coefficient of over 1.0. A similar com-bination with
afterbodies : was only slightly more effective: than the
combination of trailing-ed'e extensions '.ith afterbodies 5

Cowl and intercooler flaps open.- '.'Lith the model in
the standard configuration, opening the cowl and inter-
cooleri flaps caused separation of the flow over the wing
directly behind then at all lift coefficients (fig. 21 (a)).
The addition of trailing-edge extensions reduced the
extent of the stalled area, as shown in figure 24(b).
WVith afterbodies 4 and trailing-edge extensions on the
model 'fig. 2!:(c)), the initial separation behind the open
intercooler flans "'as delayed until a lift cotficient of
about 0.'-.5 had been reached. 'j.ith inboard afterbodies 5
and trailing-edge extensions cn the model, no separation
occurred behind the open intercooler flaps of the inboard
nacelles (fi.'.. 214(d)).


Landing Condition

Stall characteristics of the standard and modified
model with wine: flaps deflected are shown in figures 25
to 27. For these configurations the propellers were
operating at C.14 normal rated power, the transition was
fixed, and the cowl and intercooler flaps wi;ere closed.,

Standard, config.ur action. W'itft the standard configu-
ration and flaps deflected 400 (fig. 25'b)), the initial
stall on the wing occurred ahead of th- ailerons and was









NACA ARR No. L5J05


followed by separation between the nacelles. For all
angles of attack and all flap deflections, separation
occurred on the part of the flaps blanketed by the
nacelles. The flow over the lower surface of the nacelle
and the part of the afterbody that deflected with the
flap was not separated. (See fig. 26(a).) Removing the
afterbcdy tips from the standard model did not improve
the flow over the flaps. It was thought that separated
flow over the inboard flaps combined with irregular flow
created by the afterbody tips would probably contribute
most to any tail buffeting. Modifications for the landing
condition were therefore directed toward improving the
air flow at the inboard nacelles.

Modifications.- With .double slotted inboard flaps
deflected 40 (fig. 26(b)), the flow over the right flap
was not separated but the left flap was stalled, as was
the standard flap. From tests made with the afterbody
tips removed and with power off, the separation over the
double slotted inboard flaps at 0.4 normal rated power
appeared to be caused by the afterbody tip,.which
deflected with the flap, and the dissymmetry appeared to
be associated with the rotation of the slipstream. When
the standard flaps were continuous and were deflected
through afterbodies 4 (fig. 26(c)), the lower part of
the afterbody and the surface of the flap below the
afterbody were stalled at all angles of attack. With
the flaps cut out at inboard afterbodies 4, as shown in
figure 14, no stall occurred on the flap or nacelle
(fig. 26(d)). The same flow existed with trailing-edge
extensions on the flaps. With the standard flaps
deflected within afterbodies 5 (fig. 26(e)), the flow
over the flaps and afterbodies was not separated at any
angle of attack. Adding trailing-edge extensions had
little effect on the flow if the extensions were cut out
below the nacelles (fig. 26(f)).

The most promising methods of those investigated
for improving the flow over the flap were enclosing the
flap rear of the inboard nacelle within afterbody 5 or
cutting the flap at the nacelle. Trailing-edge exten-
sions and modified afterbodies, to a lesser extent,
delayed separation on the inner wing panels and thus
aggravated the tendency toward early tip stalling indi-
cated by stall studies of the standard model. This
effect could be minimized by reducing the wing-flap
deflection.









NACA ARR No. L5J05


FLO.l CHARACTERISTICS AT THE TAIL


The results of turbulence surveys at the tail are
presented in figures 28 to $1 for the standard model and
for several modifications. Diagra-is showing the flow
characteristics over the wing for conditions at which
surveys were made ar-e given in figure 52. The turbulence
data are presented as the variation of root-mean-square
value of the vertical velocity deviations with vertical
distance for several spanwise stations. Axial velocity
deviations were of the same order of magnitude as
vertical velocity deviations. I.aximnum values of the
velocity deviation are several times the root-mean-square
values. The vertical velocity deviations may be inter-
preted as angle-of-attack changes; for example, a value

of / of 0.04 is equivalent to a root-mean-square
angle deviation of slightly over 20.

Buffeting tendencies are difficult to evaluate
quantitatively because the root-mean-square deviation
indicates neither the large fluctuations that may occur
nor the frequency. Both of these factors play an impor-
tant role in determining buffeting characteristics. The
main value of the data presented is the indication of
the effects of the modifications on the turbulent wake.
The curves indicate the normal wake of the wing and
nacelles by an increase of turbulence. Beyond this main
turbulent wai-:e, there are small peaks that define the
edge of the slipstream.

The variations of local dynamic pressure with ver-
tical distance for the standard model are indicated in
figures 33 to 35. The curves show increases in dynamic
pressure due to the slipstream and depressions due to
wing or nacelle wake. The point at which the maximum
depression occurs has been assumed to be the center of
the wake. The variations of wake-center position with
lift coefficient are given in figures 36 and ,57 for the
standard and modified models. Except for displacement,
the modifications changed the profile of the dynamic-
pressure wake very little.

The vertical position of the peak values of turbu-
lence agree closely with the position of the dynamic-
pressure wake centers. The vertical extent of the main
turbulent wake is roughly the sane as that for the








NACA ARR No. L5J05


dynamic-pressure wake. A comparison of the turbulence
and dynamic-pressure-survey data indicates that only a
slight amount of turbulence is produced by the slipstream;
the greater part of the turbulence is produced by the
wake of the wing and nacelles.


Cruise Condition

Standard configuration.- For the standard model with
cowl and intercooler flaps closed (fig. 28), the largest
turbulent wake and the maximum turbulence occurred at
stations 13 inches right and 27 inches left of the fuse-
lage center line, which are behind stalled parts of the

wing (fig. 32). The maximum value of \/v2/U obtained
with cowl and intercooler flaps closed was 0.04. As shown
in figures 28 and 33, the variations of turbulence and
dynamic pressure are different on the left and right
sides of the fuselage center line. The growth of turbu-
lence on the left side from the high-speed condition
through the cruise condition is illustrated in figure 29.
At stations to the right and left of the inboard nacelle
(15 in. and 27 in. from fuselage center line) the turbu-
lence increased with increasing angle of attack. Directly
behind the nacelle, the turbulence decreased slightly as
the angle of attack increased from 4.4 to 7.60. With
increasing angle of attack, the wake centers moved up in
relation to the elevator hinge line (figs. 29 and 36).

Combinations of chord extensions and modified nacelle
afterbodies.- The effects of modifications, in general,
were to reduce the extent and magnitude of the turbulence
and to displace the whole turbulent wake downward. As
shown in figure 28, the greatest reduction in turbulence
was obtained with afterbodies 4 in combination with
trailing-edge extensions. The combination of after-
bodies 5, trailing-edge extensions, and leading-edge
gloves was somewhat less effective; and the combination
of afterbodies 5 and trailing-edge extensions was the
least effective in reducing turbulence. The greatest
reductions in turbulence were obtained at stations
15 inches -right and 27 inches left and right of the fuse-
lage center line. These reductions are apparently due
to modifications delaying separation on the wing at these
stations. At the station directly behind the nacelles
(20 in.), afterbodies 4 caused a definite reduction in








:.ACA AR. No. L5JO5


turbulence whereas afterbodies 5 caused little change.
The downward displacement of the wake due to modifica-
tions wcald be --reater than shown in figure 23 if the
data were co!:..ared at the samne lift coefficient. (See
fi2. 36.)

Cowl and intercooler flaps open.- Co.'iparison of
figures 2- and 50 shc-.os that opening the cow.l and inter-
cooler flaps caused large incr-sses in the turbulence at
the tail both wv.ith th: standard model and 'ith after-
bodies 5 A-nd trailin2-ed-.,e extensions. The mag-nitude
and extent of the velocity fluctuations 'ver, however,
much lower for the model with aft-rbodies 5 and trailing-
edge extensions.


Landinn Condition

The results of surveys for the landing condition are
given in figure 31 for a sranwise station behind the
inboard nacelle. ::ith both the standard and n-iodified
I -
models, the maximiun value of &/v2/U occurred belo;v the
elevator hinge line.

With the flaps deflected within afterbodies 5, the
turbulence was less than for the standard model. With
flaps continuous and deflected through afteroodies 4,
slightly greater turbulence was obtained than for the
standard model. The increased turbulence was evidently
due to the stall that occurred on the flap and lower
part of the afterbody (fig. 2oc)). 4ith the flaps cut
out below the nacelle as shown in figure 14-, the turbu-
lence would probably be less than for the standard model.


DRAG AIT! LIFT CHARACTERISTICS


The variations of the parasite-drag coefficient with
lift coefficient for the standard and modified model are
shown'. in figures 38 to 45. The lift, drag, and pitching-
moment characteristics for the cruise and landing condi-
tions are given in figures 46 to 52. Table III giv,-s the
numerical values of the drag changes due to modifications
at several lift coefficients and the increments of
maxim-um. lift coefficients due to the modifications







NACA ARR No. L5J05


obtained in the cruising condition. The data have not
been corrected for support tares or air-stream misaline-
ment and therefore should not be considered -as absolute
values nor should the shape of the curves be considered
correct. It is believed, however, that the changes in
drag and lift due to modifications would not be materially
affected by the application of such corrections.


Cruise Condition

Inasmuch as fli ht Reynolds numbers are much greater
than that at which the tests were conducted, the inter-
pretation of the drag reductions due to the modifications
is difficult, particularly in the range where separation
occurs. The value of drag reductions obtained at a given
lift coefficient of the model may not be in agreement
with reductions that would be obtained from tests of the
full-scale airplane. The results of the model tests,
however, are believed to be indicative of the results
that would be obtained from installation of the modifi-
cations on the airplane.

Figure 38 presents the data obtained from runs made
near the beginning, middle, and end of the investigation
with fixed and natural transition. The displacement of
the test points gives an indication of how closely test
conditions (primarily model surface condition) could be
duplicated. Fixing the transition at 10 percent wing
chord increased the drag coefficient at all values of
lift coefficient and also decreased the lift coefficient
at which the rapid increase in drag occurs.

The effect of Reynolds number on drag characteris-
tics with transition fixed is shown in figure 59. The
drag for each model configuration was lower at all lifts

at a Reynolds number of 3.9 x 10 than at a Reynolds

number of 2.6 x 10 and the knee in the drag curves
occurred at higher lifts with the higher Reynolds number.

Chord extensions.- The effect of span and chord of
the trailing-edge extensions in reducing drag is shown









ITACA ARR No. L5J05


in figure L1O. At a lift coefficient of 0.7, reductions
with (.:.-span and 0.6-span extensions were approximately
I
in the ratio of 5:4. The 2Z-inch-chord extensions
4 1
reduced the drag scmewhat more than the 1--inch cxten--
sions. No further reduction was realized by increasing
the chord to 5 inches. Subsequent tests were made with
the 0.6-sparn 1-inch-chord trailing-edge extensions.
2

Changes in drag caused by trailing-edge extensions
and leading-edge gloves were dependent upon the type of
transition. The 1--inch-chord 0.6-snar trailing-edge
extensions reduced the drag coefficient of the m..r.del
by 0.005 :ijth t -rsrsition fi;-ed at a lift ce -ff.i ient
of 0.7 (.fi .. rIi(al .ith nacural tras.sit:.. ( -,-. 41(b)),
the trailing-.dJe, extensions Licre .sid the `S 0.0005 at
a lift coeffi.-ent of 0.7 oit at lift c(effLcrents
above 0.8 appreciable reduce i. ns v.,ere obl; tr1nd. Leading-
edge -gloves r2.u':.ed the j r- i. ce. fic lernt bc i .,02.S at a
lift coeffici-ni of 0.7 .Lith J.xe- tra0. ition and
by 0.0012 wiLh natural transition (fliz. .,1 Because of
the imperfect cc^.K cur fe med where the- ].o- s aired into
the wing, the rc- si1.. is ob gained w: rh rh; ',,i :mI''i fied in
this -manner are pr.c.=bI7 not so ;o:'. s .''i-d bo obtained
if the wing v.ere built to the revised dimensions.

!odified nacelle -ftetod.ics.- All modified inboard
nacelle after".oil.. r ..n v :!, e C-.-a g ci the model at
cruising lifts (.is. 1_2) buT had little effect at low
lifts. The order of inrcreaaing effectiveness was afuer-
bodies 2, 5, 4, and 5. Modified .fterbodies on all four
nacelles reduced the drag at all lift coefficients. Four
afterbodies 4 reduced the drag coefficient of the model
by 0.00h. with fi;ed transition and 0.0017 with natural
transition at a lift coefficient of 0.7 (fig. 43). Four
after'odies 5 (fig. U4) were so:.ewhat less effective in
reducing the drag than four afterbodies 4. Modified
nacelle afterbodies were the most effective individual
modifications in reducing drag when considering both
fixed and natural transition on the wing; however,
traillng-edge extensions were slightly more effective
with fixed transition.

Propellers-off stall studies of the standard model
(fig. ]S(c)) show that at a lift coefficient of about 0.7








ITACA ARR No. L5J05


the flow behind the inboard nacelles was separated but
the flow behind the outboard nacelles was smooth. These
studies and a consideration of the shape of the drag
curves indicate that the modified inboard afterbodies
reduce the drag by delaying separation on the wing,
whereas the modified outboard afterbodies reduce the drag
by i.:iproving the afterbody form. This explanation
probably accounts for the differences between the effec-
tiveness of the inboard and outboard afterbody
modifications.

Co1,binations of chord extensions and modifiedd nacelle
afterbodies.- The combination of chord extensions and
modified nacelle afterbodies generally was more effective
in reducing drag than either modification alone. 'Some of
the more effective combinations with the drag-coefficient
reductions at C,- = 0.7 are as follows:
L,


"7odTifLation F ixed Natural
transition transition

Four afterbodies b. and trailing-
ed-... ec-xtensions 0.00,5 0.0015

Four afterbodies 5 and trailing-
ed-e extensions 0.0053 0.0005

Four afterbodies 5 and leading-
edge gloves o0.OoL6 0.0015

Four afterbodies 5, leading-edge
gloves, and trailing-edge
extensions 0.0048 0.0017



There appears to be no advantage in combining leading-
edge gloves with afterbodies 4 for reducing drag.

All modifications increased the maximum lift coeffi-
cient of the model with either fixed or natural transi-
tion, and most of the modifications increased the slope
of the lift curve.

Cowl and intercooler flaps open.- With the model in
the standard configuration, opening the cow'.l and









:-.sCA AER No. L5J05


intercooler flaps caused a lar-e increase in drag. (Corn-
pars figs. Lh5 and 5. Th, resulting. high drag coeffi-
cient 'was decreased )0.007'2 at a lift coefficient of 0.7
by the addition of intoard afterbodi-s 5 and trailing-
edrre extensions or four after-bodics i. and t.railinr-edge
extensions (fig. P15). It should be noted that these
rodificaticns reduced the drag coefficient approxi-
mrttsl- 0.0010 with cowl and inrtercooler flaps closed.


Latiding Condition

The lift characteristics for the standard and modi-
fied models with wing flaps deflected 1;.0 are presented
in figure 52. \With transition fixed and propellers off,
the same maxJ.imum lift coefficient was obtained with
double slotted inboard fla)s or with standard flaps
deflected within aftertodies 5 as was obtained with the
model in the standard configuration.

In oraer to eliminate flowv s-eparation that occurred
on the flaps and the rear part of afterbodics iL, part of
the inboard flaps were cut away. This change r- sulted
in a reduction of about 0.1 in the :m:imur:i lift coeffi-
cient. An addition-al reduction would result if the out-
board flaps were cut. With flaps deflected through
either afterbodies 4 or 5, the addition of trailing-edge
extensions increased t-=e ..i..xiniLui lift coefficient by
about 0.2: the result'ir_ lift coefficient was greater in
both cases than for the strn-dard model.


LONfT- T UDINAL STABILITY -C ARACTERISTICS


Ditchling-:nomrent curves for the. standard and modified
model with propellers removed and horizontal tail off are
presented in figures 1h6 to 11. Ponjer-on pitchin.g-moment
curves with horizontal tail on and off are presented in
figures 9 to 5;. i1o corrections have been applied to
the ,;tchi, moments but th. results presented indicate
the eOffect of the various modifications. W-ith tail on,
the large differences in trim and tne fact that large
parts of the curves are considerably out of trim make an
accur-ate evaluation of stability changes difficult.
;.]oman.t c-.ves for several elevator or stabilizer settings
would Le required.








NACA ARR No. L5J05


As indicated by changes in the slopes of the pitching-
moment curves, leading-edge gloves and trailing-edge
extensions caused appreciable reductions in longitudinal
stability.


Cruise Condition

With wing flaps neutral and the horizontal tail on
(fig. 55), the pitching-moment curves show that the model
with chord extensions and modified afteroodies was less
stable than the standard model for all power conditions.
Up to a lift coefficient of 0.6, the combination of
trailing-edge extensions and inboard afterbodies 5 changed
the moment-curve slope dCm/dCL approximately 0.O4
to 0.06 from the standard configuration. The combination
of leading-edge gloves, trailing-edre extensions, and
afterbodies 5 changed the slope approximately 0.08 to 0.10
from the standard configuration. Above a lift coefficient
of 0.6, the modifications were more destabilizing,
probably because the delayed separation on the inner wing
panel results in increased downwash. A satisfactory
compromise between the adverse stability changes and flow7
and drag improvements due to trailing-edge extensions
could probably be obtained with an extension having a
smaller chord than those tested.


Landing Condition

With wing flaps deflected, horizontal tail on, and
propellers operating at 0.4 normal rated power (fig. 54(b)),
the slope of the pitching-moment curve was approxi-
mately 0.04 less negative (model less stable) with the
combination of trailing-edge extensions and afterbodies 5
on the model. With leading-edge -loves, trailing-edge
extensions, and afterbodies 5 on the model, the slope of
the moment curve was 0.07 lzss negative than with the
model i:. the standard configuration. With propellers
operating at zero thrust (fig, 54(a)), only a slight
change in the slope of the moment curve was caused by
either modification. The adverse effects on stability
could be minimized by reducing the flap deflection.









NACA ARR No. L5J05


C ON C L US ION S


From an investigation of a model of a four-engine
bomber-type airplane that was made to determine effects
of wing and nacelle modifications on drag and air flow at
the tail, the following results were shown-

1. Worth-while improvements in the characteristics
of the model were obtained with certain modifications,
The improvements were indicated or the basis of improved
flow over the ':Jing and deflected flaps, reduced turbulence
in the region of the tail, and reduced drag.

2. Trailing-edge extensions were the most effective
individual modification in improving the flow over the
wing with wing flaps neutral, cowl and intercooler flaps
closed. Modified na.el.l afterbodies were the most effec-
tive individual modification in reducing drag with either
fixed or natural transition on the W'ing; however, trailing-
edge extensions '.vere slightly more effective with fixed
transition. Four afterbodies 4 (a beaver-tail type) alone
were superior to four afterbcdies 5 (an extended conven-
tional afterbody) alone in reducing drag. Combinations
of either leading- or trailing-edge extensions and modi-
fied afterbodies were more effective in delaying separa-
tion and reducing the drag than either modification alone.

5. iVith the model in the standard configuration,
opening the cowl and intercooler exit flaps caused sepa-
ration on the wing behind the intercooler air exit,
increased the drag considerably, and increased the tur-
bulence at the tail. These conditions were greatly
improved by addiyn3 modified nacelle afterbodies and
trailing-edge extensions.

41. W'.ith wing flaps deflected, enclosing the flap
behind the inboard nacelle v.ithin nacelle afterbody 5 or
cutting the flaps at the nacelle appear to be the most
promising methods of improving the flow over the flaps
and reducing the turbulence at the tail.

5. Although the results of turbulence surveys .made
with a hot-wire anemometer do not indicate definitely
that buffeting would occur with the standard model or
that the modifications would eliminate buffeting, the
modifications did reduce the turbulence at the tail with
wing flaps either neutral or deflected.








NACA AFR No. L5J05


6. Appreciable reductions in the longitudinal sta-
bility of the model were caused by leading-edge gloves
and trailing-edge extensions. In the landing condition,
chord extensions also aggravated the tendency toward
early tip stalling obtained with the standard model.

7. All modifications increased the maximum lift with
wing flaps neutral and gave a maximum lift equal to or
greater than that for the standard model with wing flaps
deflected except when the inboard flaps were cut out
below afterbodies 4.


Langley Memorial Aeronautical Laboratory
National Advisory Committee for Aeronautics
Langley Field, Va.



REFERENCE


1. Dryden, H. L., and Kuethe, A. Mi.: The Measurement
of Fluctuations of Air Speed by the Hot-Wire
Anemometer. ITACA Rep. 3o. 320, 1929.






NACA ARR No. L5J05

TABLE I


GENERAL SPECIFICATIO'S OF NODEL


t .o
. .
* .
. .
* .
* .
* .




t .


Wing.
Airfoil section . .
Root-section thickness, percent .
Chord, ft . .
Tip-section thickness, percent .
Chord, ft . .
Taoper ratio . .
Span, ft . .
Area, sq ft . .
Aspect ratio . .
Mean aerodynamic chord, ft .
Centcr-of-gravity location, percent ?..A
Above root chord, ft .
behind leading edge of root chord, f
Incidence (with respect to fuselage
center line), deg .. .
(-eometric twist, deg . .

Fuselage:
Over-all length, ft . .
maximumm diameter, ft .
'Maxitium frontal area, sq ft .

Nacelles:
Trontal area (each), sq ft .
Incidence (with respect to wing chord),

Horizontal tail-
Area, sq ft . .
Span, ft . .
Incidence (with respect to fuselage
center line), deg . .
Elevator hinge-line location (fuselage
horizontal)


Horizontal distance rear of leadin- edge
root chord, ft . .
Vertical distance below leading edge of
root chord, ft . .


. Davis
. 22.9
. .2.0
*.I ?.5
. b.650
3.077:1
16.`75
22.219
12.8
1.
21.6
0.185
0 o.525


. 5.0


10.386
. 1.187
. 1.107


. 0.1450
deg -5.0


. 5.201
. 5.575


of
* .


-1.0



5.866


. 0.051


Propc Ilers

Diameter, ft . .. 2.082
Blades .. .. . 14.
Blade design . Curtiss Wright 10l6-16


NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS


* .







NACA ARR No. L5J05


TABLE II

ORDT:ATES FOR F~.7LER A'ND DOi.VL SLOTTED FLAPS

stations and ordinates given in inches and are measured
from leading edge and reference line of Fowler flap]


6.25"


Fowler flap

tto +- Fpner I Lower
station surface surface I


.078
.156
.?12
.469
.625
.937
1.250
1.875
2.500
5.125
3.750
4.575
5.000
5.625
6.25


0.196
.5364
.442

.632
.688
.769
.819

.821
.760
.522
.5366
.220


0.196
078
062

.oW6

.026
.010
.006


L.E. radius: 0.125


7 AT IO,'AL .ADVISORY
COaMMITTEE FOR AEROGTAUTICS


Double slotted flap
Vane

Station Lower
surface
0.1i6 0.06
.28 .12
.L- .5o

.78 .64
1.05 .75
1.26 .o2

Nose of flap
Station Upper Lower
surface surface
1.05 0.20 0.20
S1.11 .8 .05
1.1 .5 .05
1.50 6 -------
1.66 .70
1.97 .78 -------
2.28 .82 -------
2.50 .82 ------















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NACA ARR No. L5J05


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NACA ARR No. L5J05


INDEX TO FIGURES
a. Illustrative Figures


Figure Material presented
1 Three-view drawing of model
2 Model mounted in test section
3 Photograph, standard nacelle afterbodles, Of = 0
4 Drawing, standard nacelle afterbodies
5 Photograph, standard nacelle afterbodies, Of = 1400
6 Drawing, L.E. gloves
SPhotograph, L.E. gloves and T.E. extensions
Photograph, nacelle afterbodies 2 and 3
9 Photograph, nacelle afterbodies 4, Of = 00
10 Photograph, nacelle afterbodies 5, Of = 00
11 Drawing, nacelle afterbodies
12 Drawing, nacelle afterbodies 4
1Drawing, nacelle afterbodies 5
Photograph, nacelle afterbodies 4, Of = 400
15 Photograph, nacelle afterbodies 5, Of = 400
16 Photograph, double slotted flap
17 Calculated thrust coefficients for B-32 airplane


b. Stalling Characteristics; 0.4 Normal Rated Powerl Transition Fixed

8Figure Configuration Of Cowl and inter-
Figure TConfi.gurati_ (deg) cooler flaps
18(a) Standard model 0 Closed
al8(b) ------do --------------------------------- 0 Do.
bl8(o) ------do--------------------------------- 0 Do.
19 Chord extensions 0 Do.
20 Modified inboard afterbodies 0 Do.
21 Inboard afterbodies 4 and chord
extensions 0 Do.
22 Inboard afterbodies 5 and chord
extensions 0 Do.
23 Comparison of stall patterns at CL 0.8 0 Do.
24 Standard and modified model 0 Open
25(a) Standard model 10 Closed
25(b) ------do--------------------------------- 40 Do.
26 Standard and modified model (flow over
Inboard afterbodies and flaps) 40 Do.
27 Modified model (flow over wing) 40 Do.
aNatural transition.
bPropellers off.


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS







NACA ARR No. L5J05


TINE TO FIGURES Concluded
c. Tall-Survey Data; 0.1 Normal Rated Power; Transition Fixed

Figure Material presented Form of data (d Cowl and Inter-

28 Standard and modified model, a = 7.60
and 8.7o 2 0 Closed
29 Effect of angle of attack against vertical 0 Do.
30 Standard and modified model, a = 7.50 dista wc 0 Open
31 Standard and modified model, a = 12.7 d0tane Cloed
32 Stall patterns for survey conditions ----------- ---- ----- ...---
33 Standard model q 0 Closed
S *----do----------------------- against vertical 0 Open
35 -...---do----------------------- distance 40 Closed
36 Posltlon of wake center I Vertical dIstanoe Do.
37 ---------do------------------------ against CL 0 Open



d. Aerodynamic Characteristlos with Propellera Off

Figure Material presented PFor of data or n o finter- Transition

38 Several rune with standard model C against CL 0 Cloned Plxed and
S9 Effect of Reynolds number --- -do ------- 0 ------ do------ Fixed
0 T.E. extensions ------do------- 0 ------do------ Do.
41 Standard afterbodies with chord Fixed and
extensions -----do------- 0 ------do------ natural
12 Effect of various nacelle
afterbodle ------do------- 0 ------do------ Fixed
45 Aftirbodles 4 with chord Fixed and
extensions ------do------- 0 ------do------ natural
14i Afterbodles 5 with chord
extensions -------do------- 0 ------do----- Do.
45 Effect of modifications with
cowl and intercooler flaps
open ------do------- 0 Open PFixed
46 Effeet of transition CD, Can 0 Closed Fixed and
against CL natural
l7 T.E. extensions ----do------- 0 ------do---- Fixed
8 Standard aterbodles with chord Fixed and
extensions ------do------- 0 ------do----- natural
49 Afterbodles 4 with chord
extensions -------do------- 0 ------do------ Do.
50 Afterbodles 5 with chord
extensions ------do------- 0 ------do------ Do.
51 Effect of modifications with
cowl and Intercooler flaps
open ------do------- 0 Open Fixed
52 Effect of modifications on lift CL against a 40 Closed Do.


e. Longitudinal Stability Characteristics; Transition Fixed
uf Cowl and inter- Tall power
Figure Configuration (dog) cooler lap- Tal Power

553(a) Standard and modified model 0 Closed on Te = 0
53(b) ------------do------------- 0 ------do------- -do- 0.4 normal rated power
53(c) ------------do------------- 0 ------do------- -do- Normal rated power
54(a) Standard and modified model 40 ------ do ------- On To 0
54(b) -----------o-- ----- 40 -----do------ -do- 0.4 normal rated power
55(a) Standard and modified model 0 ------ do---- off 0.4 normal rated power
55(b) ------------do------------- 40 --- do--------do- 0.4 normal rated power


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS








NACA ARR No. L5J05


2


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS


Figure I. Three -vew draw/ng of model.


Fig. 1









NACA ARR No. L5J05 Fig. 2



























bo
V.





0



c



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NACA ARR No. L5J05 Fig. 3


Figure 3.- Standard nacelles. bf = 00.







NACA ARR No. L5J05


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Fig. 4a











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NACA ARR No. L5J05


Figure 5.- Standard nacelles.


Fig. 5


5sf = 40.


I 1 i







NACA ARR No. L5J05 Fig. 6










2 / *
















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NACA ARR No. L5J05


"'

LMA 4 2


np-"


Figure 7.- Leading-edge gloves and LLaiiing-
edge extensions.


Fig. 7


/






NACA ARR No. L5J05


(a) Afterbody 2.
w 'it


(b) Afterbody 3.

Figure 8.- Inboard nacelle afterbodies 2 and 3.


Fig. 8ab







NACA ARR No. L5J05


Figure 9.- Nacelle afterbody 4. 5f = 00.


Fig. 9






NACA ARR No. L5J05


Figure 10.- Nacelle afterbody 5. Sf 00.


Pig. 10





NACA ARR No. L5J05 Fig. 11




ici- / /K^ '' q :i
^~~~~ o-^-? ^-
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NACA ARR No. L5JO5













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NACA ARR No. L5J05


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Fig. 13a












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1',
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Fig. 13b






.-

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NACA ARR No. L5J05


Figure 14.- Inboard nacelle afLerbody 4. Sr = 400.


Fig. 14







NACA ARR No. L5J05


Figure 15.- Inboard nacelle afterbody 5. 6f = 400.


Fig. 15







NACA ARR No. L5J05


--nw
r (, "
^~~~ 1 ^^,


Figure 16.- Double slotted inboard flap.


Fig. 16







NACA ARR No. L5J05 Fig. 17



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NACA ARR No. L5J05


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Fig. 22a-c


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9 ST LLED UN5TALLE D


ITEPMIT- [) L ECTiON
TENT 5T;-LL O) FLcWJ


3o ny f I I L r nw'f g rc R org I I .f o uf n
A ,-erbodotes 4; 7. .rercoaes s-;
a = z75 ileeac,.;-ig eab e.- = L 5 aea,.aeglors
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ar = 6 5" o g ed-e s t /on r"- ..J... ~e 4 /a1J
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65 &l^ e qb^sr'1'qt 0 75 5 J ea g ok i//ng-
5.0.80C: 4' eerensons *c -o7Q '~edye 7iens.'oots
S= 8 1 NATIONAL ADVISORY
COMMITTEE FOP AERONAUTICS



Fig.,_re 23. -Co .cor- so c oo.e 5 nW f /-o or D-0 .,rco. fe'
...oe o at '.F ::oe-' ier.. c:.pr.~.C:.e.r / .8. -=04; cc 0o-s
'-Zercor ,O00.o cc.sso; rcrms r ."; r.ans Z.c 6'00e0d
P %,600,000.


Fig. 23








NACA ARR No. L5J05

















S i

































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NACA ARR No. L5J05


Fig. 2?a-d


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NACA ARR No. L5J05


Standard configuration; a = 7.50; CL = 0.73
Inboard afterbodies 4 and trailing-edge extensions; a=7.6o;CL=0.91
Inboard aftercooles 5 and trailing-edge extensions; a=7.60;CL=0.91
Leading-edge gloves, inboard afterbodies 5, and tralllng-edge
extensions; a = 7.60; CL = 0.90.
13 in. 20 in. 27 in.


20


Zzz
0



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0








QQJ





-o
4-


Left of fuselage center line








,0_P -





NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS


.02


0 .,
}/1/u


.04-


Right of fuselage center line
(a) a = 7.6.
Figure 28.- Vertical velocity deviations in region of horizontal
tall for standard and modified models. 6f = 0; cowl and
Intercooler flaps closed; transition fixed; 0.4 normal
rated power; R = 2,600,000.


/0



0



-/0


Fig. 28a


^


.0a







NACA ARR No. L5J05 Fig. 28b
















0 Standard configuration; a = 8.5; CL = 0.82
o Inboard afterbodies 5 and trailing-edge extensions; a=8.f; CL=1.01
7 Leading-edge gloves, Inboard afte6bodies 5, and trailing-edge
extensions; a = 8.7; CL = 1.01


U

15 In. right of
fuselage center line


.02


27 in. left of
fuselage center line


0 O .04-
NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS


.04


(b) a z 8.70
Figure 28.- Concluded.


/0
5*%C







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v7/u






NACA ARR No. L5J05


13 in. left of
fuselage center line
20





/0


20 in. left of
fuselage center line


27 In. left of
fuselage center line


fa) Standard configuration, a C
(deg) L
& h.5 0.45
4 .-- o.43


10 f --.-.,-- .-- -
/0






-/0 ., I

0 C
(deg) L
(b) Inboard afterbodies .and h,~4 0.5b
trailing-edge extensions. 0 7.t .91
9 t.7 1.01
0 .02 0 .02 0 .OZ .04
V71/U NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS
Figure 29.- Effect of angle of attack on vertical velocity
deviations in region of horizontal tall for standard and
modified models, 6f = 0; cowl and Intercooler flaps
closed; transition fixed; 0.4 normal rated power;
R z 2,600,000.


Fig. 29a,b





NACA ARR No. L5J05


20 in. left of fuselage center line
0 Standard configuration; a = 7.40; CL = 0.66
F Inboard afterbodies 5 and tralling-edge ex-
tensions; a = 7.60; CL = 0.88


Vt

(-
cA
C--4-


Figure 50.- 'vertical velocity deviations in region of horizontal
tail for standard and modified models. 6f = 00; cowl and
Intercooler flaps open; transition fixed; O.4 normal rated
power; R: s 2,600,000.


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS
cit .06 .6
7r__


Fig. 30






NACA ARR No. L5J05


(a) Standard configuration.
a = 12.70; CL = 2.52.


Q)



(-)
-I--

C,

0s


(b) Inboard afterbodies 4,
flap continuous.
a = 12.70; CL = 2.51.


Distance left of fuselage center
(In.)
SE 0 17
S0 20 --
o 23

W ,!

"l^


(d) Leading-edge gloves,
Inboard afterbodies 5,
and trailing-edge
extensions. a = 12.9 ;
CL = 2.95.

O .02 .04
NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS


Figure 31.- Vertical velocity deviations in region of horizontal
tall for standard and modified models. 6f = 400; cowl and
intercooler flaps closed; transition fixed; 0.4 normal rated
power; Rz2 2,600,000.


line


(c) Inboard afterbodies 5.
a = 12.70:' C = 2.55.

0 .02 .0 4

^//7-/u


Fig. 31a-d








NACA ARR No. L5J05


Fig. 32a,b


STALLED INSTALLED

SITERMIT- DIRECTION
TENT 5TALL OF FLOW


Bough -- [ *
= 400 ?20 .u-
C /0 t \ 7 i_ 1 no--'es
'Thz C '.1I -


(b) 4fYerboc' ", -
NATIONAL ADVISORY
COMMITTEE FOR AEIONAUTICS


Figure 32.- Stall patterns for conditions of
turbulence surveys. Transition fixed;
0.4 normal rated power; R ,--- 2,600,000.







NACA ARR No. L5JO5


Fig. 32c,d


L i CALLEDD UN5STALLE D


L, IrTEPMI T- DLECT-iON
TENT 5TLL OF FLOW


V* I
71,
-r


(c) Af+erc'c'les 5. .,'.,-
goge &Csr'iS,0C


n ---, r-- NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS


Figure 32 ." Concluded.







NACA ARR No. L5J05


c =4.3




/0-


0

-I0





20

/0






t -/


I -
20
/0



( Z -
0 ^


.8 /1.0 1.2


= 7So er= 70











13 Ln. from fu7e4&ae center Ulne
13 Ln, from fume &ge center ].ina


1:=2.70

F-./


20 in. from fuselage center line


-bI







27 in. from fuselage center Line
.8 1.0 12 /4
. /0 .2 14 .8 /.0 12 1.4

/q t/ NATIONAL ADVISORY
C/ Iq COMMITTEE FON AEIONAUTICS


Figure .3.- Dynamic-pressure varigtion In the region of the horizontal tall.
standard configuration; 6b = 00; cowl and Intercooler flaps closed; transition
fixed; 0.6 normal rated power: R ". 2,600,000.


Fig. 33






NACA ARR No. L5J05


.6 .8 10 12


.6 .8 1.0 /Z /.4
b .8 10 /I /.4 .6 .8 /.0 12 ,4

I,


Fi.-ure L.- Lynarlc-c.re-asur.a vr ilertl In in to region of tr.e r.orlzontal tall,
a20 Inches from fuselage center line. ?tsr.dard -nrflgurat Icr; = O'-
coal and Intarcooler flasp open: trar .!r_.I n Ii-ed,; O., normal rat6d power:
Fi h 2,600,000.


Fig. 34






NACA ARR No. L5J05


-/
,13 in. from fuselage center line

-20 -- -- t
Le t,













20
So -Tr .. -"











20 In. from fuselage center line



... -0 12 14 8 /.0 2 14'
k ." A -_ ." f 5. .- _






/.0 12 /4 &. /.0 12 /4

,'_ l/ NATIONAL ADVISORY
t1 COMMITTEE FOR AERONAUTICS


Pe-ure 35.- Dyrenaic-ore .r- .arlEaclr, Irn rne regionr, of tro riorl r :ntal tail.
.iEr, dard confl'.iuratlor,; or = 400; cal and Irntercocler flaps closed; transition
fixed; O.4 normal rste d po-er; R = 2,600,000.


Fig. 35







NACA ARR No. L5JO5 Fig. 36






---Standard configuration
---Inboard afterbodies 5 and trailing-edge extensions
/0 ----Leading-edge gloves, inboard afterbodies 5, and trailing-
I_ ~edge extensions I I I I / I I I


15 in. from fuselage center line


Right


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS


/ I /

o. .-- ------

7-5 __ l

27 in. from fuselage center line


Figure 36.- Variation of location of dynamic-pressure wake
center with lift coefficient. 5f = 00; cowl and intercooler
flaps closed; transition fixed; 0.4 normal rated power;
Rz: 2,600,000.


u QI


Left






NACA ARR No. L5J05


--- Standard configuration
--- ----- Trailing- edge extensions
---Inboard afterbodies 5 and
trailing-edge extensions


20 in. left of 20 In. right of
fuselage center line fuselage center line

11111_I I I I _


.8


Figure 37.- Variation of location of dynamic-pressure wake center
with lift coefficient. 6B = 0 ; cowl and intercooler flaps
open; transition fixed; 0.4 normal rated power; R 2,600,000.


Q)
b-






0


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0


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NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS


I


Fig. 37







NACA ARR-No. L5J05 Fig. 38














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Fig. 39a


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Fig. 39b


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Fig. 40b


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Fig. 41a


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Fig. 41b








NACA ARR No. L5J05


Fig. 42


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Fig. 43a


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NACA ARR No. L5J05


Fig. 43b


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NACA ARR No. L5J05


Fig. 44a


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NACA ARR No. L5J05


Fig. 44b


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NACA ARR No. L5J05


0 'c












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Fig. 45







NACA ARR No. L5J05 Fig. 46








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NACA ARR No. L5JO5


___ ___ -& __ __ 0 ^ ^ ._ ,.^^' .^1* ",V .... __ _




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Fig. 48a


















II
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NACA ARR No. L5J05


Fig. 48b


a
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NACA ARR No. L5J05


Fig. 49a










NACA ARR No. L5J05


Fig. 49b


o *~
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-4
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Figure 52.- Lift characters tion with wing flape deflected for standard and modified
models. 6f = 400; cowl and intercooler flaps closed; fixed transition;
R= 2,600,000; propellers off.


Standard configuration
Double slotted Inboard flaps
Inboard afterbooles 4, flaps cut out at
Inboard nacelles
Inboard afterbodiles 4 and tralling-edge
extensions; flaps and tralling-edge
extensions cut out at Inboara nacelles
Inboard afterboales 5
Inboaro afterroales 5 and tralling-edge
extensions
NATIONAL ADVISORY
COMMITTEE FOR AElONAUTICS


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