Determination of the stability and control characteristics of a straight-wing, tailless fighter-airplane model in the La...

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Title:
Determination of the stability and control characteristics of a straight-wing, tailless fighter-airplane model in the Langley free-flight tunnel
Series Title:
NACA WR
Alternate Title:
NACA wartime reports
Physical Description:
16, 16 p. : ill. ; 28 cm.
Language:
English
Creator:
Seacord, Charles L
Ankenbruck, Herman O
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, Tailless   ( lcsh )
Aerodynamics -- Research   ( lcsh )
Genre:
federal government publication   ( marcgt )
bibliography   ( marcgt )
technical report   ( marcgt )
non-fiction   ( marcgt )

Notes

Summary:
Summary: An investigation to determine the stability and control characteristics of a straight-wing, tailless fighter model with a pusher propeller designed by the NACA has been made in the Langley free-flight tunnel. The investigation consisted principally of force and flight tests of a powered dynamic model. The effects of tail configuration, center-of-gravity location, and power on the stability and control characteristics of the model were determined. Tests were also made in the Langley 15-foot free-spinning tunnel to determine whether the model would trim at very high angles of attack.
Bibliography:
Includes bibliographic references (p. 14-15).
Statement of Responsibility:
by Charles L. Seacord, Jr. and Herman O. Ankenbruck.
General Note:
"Originally issued February 1946 as Advance Confidential Report L5K05."
General Note:
"Report date February 1946."
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."

Record Information

Source Institution:
University of Florida
Rights Management:
All applicable rights reserved by the source institution and holding location.
Resource Identifier:
aleph - 003804667
oclc - 123894119
System ID:
AA00009655:00001


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N~'~ & -IQ~i


X


ACR No. L5K05


/


NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS





WArll TIME REPORT
ORIGINALLY ISSUED
February 1946 as
Advance Confidential Report L5K05

DETERMINATION OF THE STABILITY AND CONTROL CHARACTERISTICS
OF A STRAIGHT-WING, TAILLESS FIGHTER-AIRPLANE MODEL IN
THE LANGLEY FREE-FLIGHT TUNNEL
By Charles L. Seacord, Jr. and Herman 0. Ankenbruck

Langley Memorial Aeronautical Laboratory
Langley Field, Va.








NACA


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 199



































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/determinationofeng









NACA ACR `?o. L5K05


nTATIOIAL ADISORY O. TTlE p: 1. AERONAUTICS


ADVANCE CONF7IDLETIAL REPORT



DETERUTTTATIOrT OF THE STABILITY ATD CONTROL CHARACTERISTICS

OF A STRAITGHT-W~II-, TAILLESS F' ': i : -AIRPLA P I:ODEL I!:

THE LANIGLEY FREE-FLIGHT I'T1;EL

By Charlis L. Seacord, Jr. and Herman 0. Ankenbruck


SUT:- UARY


An investigation to determine the stability and
control characteristics of a straight-wing, tailless
fighter model with a pusher propeller designed by the
NTACA has been made in the Langley free-fliht tunnel.
The investigation consisted principally of force and
flight tests of a powered dynamic model. The effects
of tail configuration, center-of-gravity location, and
power on the stability and control characteristics of the
model were determined. Tests were also made in the
Langley 15-foot free-spinning tunnel to determine whether
the model would trim at very high angles of attack.

The results of the investigation may be summarized
as follows: The general flight characteristics of the
model were good and compared favorably with the flight
characteristics of good conventional airplane models
previously tested in the Langley free-flight tunnel. As
the angle of attack was increased, the longitudinal
stability of the model increased instead of decreasing as
that of tailless airplanes with swept-back wings usually
does. Power caused a slight reduction in the longi-
tudinal stability measured at constant power. This
reduction in stability, however, did not affect the longi-
tudinal steadiness of the model in flight tests. The
model did not show the tendency to trim at very high
angles of attack (above the stall) that has been a char-
acteristic of some swept-back tailless airplanes. 'The
lateral flight characteristics of the model with both
vertical tails installed were good. The directional
stability of the model was satisfactory and was improved
by thu application of power. The effective dihedral was
desirably small and was not appreciably affected by power.









NACA ACR No. L,5 K05


The control surfaces of the model provided adequate longi-
tudinal and lateral control.


INTRODIJCT ION


Previous investi- tions of the stability and control
of tailless airplanes with swee':,sck (references 1 to 4)
have indicated that the sweepback is the cause of the pcor
longitudinal stability and the loss of control near th6
stall which are often characteristic of such airplanes.
In order to determine the effects on stability of elim-
inating the sweer .' ack, a straight-wing, tailless fighter
airplane has been designed by the :UACA and a model of the
design has been tested in the Langley free-flight tunnel.

The present investigation is one phase of the tailless-
airplane research T.ro-ram being carried out in the Langley
free-flight tunnel to determine the relative merits of the
various types of tailless aircraft and includes results of
both force .'d flight tests of a dynamic powered model with
a pusher propeller. because some tendency has been noted
for tailless airpl-..s to trim at very high angles of
attack, 900, brief tests wore also made in the Lanr.ley
15-foot free-splnnrinr tunnel to investigate the trim
characteristics of tho model at large :.niles of attack.
The force tests were made with the :odel equippId with
two different sizes of vertical tail surface, with pro-
p-llurs off and with propellers on, and with power oijusted
to simulate that typical of modern fightter airplanes. The
model was flown with two different sizes of vertical tail,
w.ith various center-of-Cravity- locations, r.ad :1th various
amounts of power.


SYMBOLS A2"D :-CATI "1S


CL lift ccefficient ( 1.

CD dra coefficient Drpr

C pitchin-morent coefficient about normal
n*g center-of- -.avit- lcatlcn ['/q73i

Cy lateral-force coefficient (LateMal re)
CC .'l .!'IAL


CO -IDE7'T'IAL









NACA ACR 'o. L515:;:i


CL rolling-noment coefficient (L/qbS)

C, yav.ing-moment coefficient (N/qbS)

L rolling moment, foot-pounds

M pitching moment, foot-pounds

N yawing moment, foot-pounds

q dynamic pressure, pounds per square foot ( pV2)

p mass density of air, slugs per cubic foot

V airspeed, feet per second

W weight of airplane, pounds

S wing area, square feet

b wing span, feet

c wing chord, inches

5 mean aerodynamic chord (M.A.C.), feet

a angle of attack of fuselage reference line, degrees

angle of yaw, degrees = -)

p angle of sideslip, degrees

I angle of roll, degrees

CT rate of change of rolling-moment coefficient with
P angle of sideslip, per degree (dC7/dp)

C rate of change of yav:ing-moment coefficient with
angle of sideslip, per degree (dCn/dp)

-T thrust disk-loading coefficient (T/pV2D)

T thrust, pounds

D propeller diameter, feet

5aR right-aileron deflection, degrees
C TIDEIT IAL


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IACA ACR 1Jo. L5K05


6e elevator deflection, degrees

kX radius of gyration about X-axis

ky radius of gTration about Y-axis

kZ radius of gyration about Z-axis

The parts of the model are designated as follows:

, wing

F fuseleve, incluiinr pilot's enclosure and wire
landing gear

P propeller

V1 lower vertical tail

V2 upper vertical tail


APPARMaTUS

Wind Tunnels


The investigation was carried out in the Langley
ifr -flight tunnel, which is equipped for testing free-
flying dynamic models. A complete description of the
tunnel and its operation is given in reference 5. Force
measurements were made on the Langley free-flight-tunnel
six-component balance described in reference 6. The
forces and moments are measured on this balance with
respect to stability axes. T.'.e stability axes of an air-
plane are defined as an orthogonal system of axes inter-
sectin, at the center of gravity in which the Z-axis is in
the plane of symmetry and perpendicular to the relative
wind, the X-axis is in the'plane of symmetry and perpendi-
cular to the Z-axis, and the Y-axis is perpendicular to the
plane of j:Tr.Lrtry. A sketch showing the stability axes of
an airplane is -.resented as figure 1. A photograph of the
test section of the tunnel showing the model rbeling tested
in fl-igh is presented as figure 2. The tests to deter-
r:-ine the trim characteristics of the ic.del at high -ngles
of attack were made in the La..lv-'y 15-foot freu-spinning
tunnel, a description of which is given in reference 7,


C O'!F' IDE-T,' IAL


Sr 7r-- T'," rp, T AT









NACA ACR No. L5K05


MTodel


The test- model was' designed and constructed by the
NACA and corresponds to a- 1 -scale model of a hypo-
thetical tailless airplane with a 4O-foot span. It is
a high-wing design with the 50-percent-chord line straight
and has a small fuselage,.a pusher propeller,.and conven-
tional vertical tail surfaces. A drawing and photographs
of the model are givenn as figures 3 to 6.

Longitudinal control for the model was provided by
elevators that extended over the inboard portion of the
wing, and lateral control was provided by conventional
aileron and rudder surfaces. For power-off windmillingg)
tests, the model was fitted with a four-blade propeller
that was allowed to windmill freely. For power-on tests,
the model was equipped v:itfh a l-horSepower electric motor
'driving an 11-inch-diameter three-blade propeller. The
three-blade propeller was installed in place of the four-
blade -propeller because the characteristics of the motor
made. it possible to obtain higher thrusts with this
arrangement.

The- model wing had a Rhode St. Genese 55 airfoil
section reflexedd) because this section has a high maximum
lift coefficient at the low Reynolds numbers at which the
tests were run.

The physical characterististics of a full-scale air-
plane based on scaled-up values (10:1) of the dimensions
of the model are:


Weight, pounds .. .
VWing
Area, square feet .
Span, feet . .
Aspect ratio . .. .
Sweepback of 50-percent-chord line, degrees.
Sweepback of 25-percent-chord line, degrees.
Incidence, degrees .
Dihedral angle of midthickness line, degrees
Taper ratio . ..
M.A.C., inches .
Location of M.IA.C. behind L.E. of
root chord, inches. . .
Root chord, inches . .
Tip chord, inches. .. .
CONFIDENT TI AL


. 8050

. 266.67
. 40.0
. 6.0
. "0
S. .2
. 0
* 0
. 2:1
. 85.9

. 12.0
. 107.8
. 55.9


CO11 DENT'IAL










NACA ACR :o. L5K05


Wing loading, W/S, pounds per square foot .
Aileron
Type. . . .
Area, percent wing area .
Span, percent wing span .
Chord, percent wing chor .
Elevator
Type. . . .
Area, percent area .
Span, percent wing span .
Chord, percent wing chord .
.Iormal c.g. location
Behind L.E. of root chord, inches .
Pi-hind L.E. of root chord, percent T.I.A.C.
Above thrust line, inches .. .
Above thrust line, percent M.A.C. ..
.Ratios of radii of gyration to wing scan
kx/b . . .
k . . .
kz/b . .


Vertical tails
Total area of each, percent wing area .
Rudder area, percent total vertical-tail area
Aspect ratio (each tail ) .
Distance from c.g. to rudder hinge line,
percent wing span. .


* *


. .Plain
* S 7
. 15
.. 20

. Plain flap
. 7.7
. 45
S. 15

. 28.8
. 20.0
. .0


. 0.158
. 0 .133
. 0.175


. 5.0
23
. 241.85

2 ,2


In -' i r n T
rrJ. ^J3


Force Tests


Most of the force tests were made at a dynamic "rr-s-
suri of 4.09 pounds per square foot, which corres bonds
to a test Reynolds number of approximately 240,000 based
on the mean aerodynamic chord of 0.699 foot. The force
tests consisted of angle-of-attack runs made to determine
the effects of power and various modifications to the
model on longitudinal stability and control a.d yLa runs
made to determine the lateral stability and control char-
acteristics of the ':odkl in all conditions. A suizary
of the force-test conditions is given in table I. As
shown in table I, power-on force tests were made to idtermine
the static longitudinal stability of the .ndcl operating
with power simulati:;-i zero thrust and 1200 brake horseIpower
for the -v 'othetical full-scale airplan-. In the late ral-
stability tests the model operated with power simulating


CO f DENT IAL


CO'Tr ID iT IAL









IA.CA ACR'"To. L5K-05


COF IDE n' IAL


zero thrust, 1200 brake horsepower, and 2000 brake horse-
power for the fall-scale airplane,

Values of thrust coefficient required to simulate
1200 and 2000 brake horsepower over the lift range of the
model tests are shown in figure 7. These data are based
on an assumed propeller efficiency for the full-scale
airplane of 75 percent and a wing loading of 30 pounds
per square foot.


Flight Tests

Model flight tests were made to determine effects
of lift coefficient, center-of-gravity location, vertical-
tail area, and power on the stability and control char-
acteristics of thLe model. In the power-off condition,
flights were made over a range of lift coefficients from
0.52 to 0.95 for center-of-gravit locations ranging
between 15 and 25 percent I..A.C. Most flight tests were
made with the model equipped with both upper and lower
vertical tails, but a few tests were made with the model
equipped with the upper tail only. Power-on flight tests
were made for a lift-coefficient range with the normal
center-of-gravity location (20 perce-it M.A.C.). The highest
power simulated in flight was 1200 brake horsepower.


Free-to-Trim Tests

In the free-to-trim tests the model was supported
in t'he air stream of the Langley 15-foot free-spinning
tunnel on the stand shown in figure 8. The model was free
to rotate in citch about its center of gravity and lad
a possible travel of about 2000. The model was restrained
until the airspeed had been adjusted and was then released
to trim. The model "was released at angles of attack from
00 to + 900 with the elevators set to trim the model at an
angle of attack of 8.


RESTCLTS AND DISCUSSION

Longitudinal Stability

Power-off force tests.- The results of force tests
made to determine the longitudinal stability characteristics


TCI 'IDE2TIAL









:TACA ACR No. L5KO5


of the model with power' off are presented in figure 9,
The effects of the variohis cor'pon-nt parts of. the :rmcdel
on the stability are also shown in this figure.

The data of figure 9 show that the model was longi-
tudinally stable up to the stall. This characteristic
is desirable and is not usually possessed by tailless
designs incorporating sweepback. The stability of the
model increased with increasing lift coefficient; the
static margin, as indicated by the value of -dCm/dCL,
varied from about 0.02 at low lift coefficients to about
0.07 at high lift coefficients with the.normal center-of-
gravity location (20 percent M.A.C.). Reference 8 shows
that this change in the slope of the pitching-moment
curve is characteristic of a high-wing arrangement on a
round fuselage and indicates that the pitching-moment
curve could probably be straightened by lowering the wing
to a high mi:- ing position.

The data of figure 9 show that adding the fuselage
to the wing caused a reduction in static margin (-dCm/dCL
of about 0.01. The data also show, however, that" the
stabilizing effect of the windmilling pusher propeller
counteracted the destabilizing fuselage effect, so that
the stability of the complete model was similar to that
of the wing alone.

Power-off flight tests.- In power-off flight tests
with tle center of gr'vity at 20 percent M.A.C., the
longitudinal stability of the model was satisfactory at
lift coefficients from 0.50 to the stall, at which the
static margin was .0.05. At .lift coefficients less than
0.50, however, the longitudinal motion of the model was
unsteady a-id frequent elevator control was required to
keep the model flying. This unsteadiness was attributed
to the small static margin (about 0.02 or 0.03) at low
lift coefficients previously indicated by the force tests.

"e-'en the static ,.rg;in was increased by 0.02 by moving
the center of gravity ahead to 18 percent MI.A.C., steady
flights were obtained over the entire lift range from a
lift coefficient of 0.32 to the stall (flights could not
be made at lift coefficients lower than 0.32 because of
tunnel airsFpci limitations).

Decreasing the static margin by shifting the center
of gravity to 22 percent ::.A.C. caused the longitudinal


COC:;'DE; IAL


CO'FIDE-'- IAL









TACA ACR TNo. L5K05 CCO IDETIAL


flight behavior of the model to become completely unsatis-
factory at lift coefficients less than 0.50 and only fair-
ly satisfactory at lift coefficients greater than 0.50.
In some flights iade at lift coufficients above 0.50 with
a static margin of about 0.02 or 0.03, the model was very
unsteady and difficult to control and the longitudinal
characteristics were very similar to those obtained in the
flight tests made at lift coefficients below 0.50 and with
the center of gravity at 20 percent M.A.C.

Previous tests in the Langley free-flight tunnel
(reference 9) have shown that conventional models had
longitudinal steadiness characteristics which were
essentially the same as .tlose of tha straight-wing,
tailless model with corresponding values of static margin.
In this respect the results of the present investigation
are in agreement with the results ,of reference 9, which
showed that variation of da-:iping in pitch has little effect
on longitudinal steadiness as long as the static margin is
satisfactory.

Povwer-on force tests.- For purposes of discussion,
static margin has been assumed equal to -dCm/dCL. This
assaLmption should be nearly true in the case of the model
tested because the model has no horizontal tail and
because the wing is not in the slipstream. The force-test
data of fi-gure 10 show that power- caused a reduction in
static margin which, though appreciable (;0.'0 or 0.01),
was not so .great as the reduction often caused by power
on conventional single-engine airplanes with tractor
propellers (reference 10). At 1200 brake horse power with
the center of gravity at 20 percent M.A.C. the model had
a static margin of only about 0.01 over most of the lift
range.

The results of calculations made to determine the
cause of the decrease in stabiliPty with application of
power are presented in figure 11 in the form of incre-
mental pitching moments provided by the propeller normal
force and propeller thrust (figs. 11(a) and (b)). The
combined calculated effects of propeller forces are also
compared (fig. 11(c)) with the measured power effects
taken from the data of figure 10, The calculations show
that, although the normal force of the pusher propeller
provided a slight stabilizing effect, the propeller thrust
provided a much greater destabilizing effect. Figure 11
shows that the measured destabilizing effect of power was
about twice the calculated effect of direct propeller


CC TI DL -IAL








,TACA ACR No. L5K05


forces. The additional unstable moments may have been
reduced by the inflow effects over the wing and the
rear portion of the fuselage. The data of figure 12
show that if the center of gravity of the .nodel were
shifted vertically downward from 0.08 048'.A.C. above the
thrust line to 0.011 ;'.A.C. below the thrust line, power
would not affect the static longitudinal stability of
the model.

Power-on flight tests.- Although the force-test
results indir,-ted a decrease 'in longitudinal stability to
a static r.ar.gin of 0.01 as the porer was increased from
zero thrust to 1200 brake horsepower, the longitudinal
steadiness in flight tests of the model was not appreciably
changed by power application. Flights made with power
simulating 1200 brake horsepower were as steady as flights
made with zero thrust. These results thus appear to
disagree with the results of the power-off flight tests,
in which a reduction of the '.ov;-r-off static margin from
0.05 to 0.02 caused the longitudinal steadiness of the
model to become definitely worse.

An explanation of this apparent discrepancy is that
in the nower-on force tests the thrust was varied with
angle of attack to represent constant-power flig'.t at
different airspeeds that is, Tc was varied with CL
and thus with airspeed, as shown in figure 7 'whereas
in the power-on'flight tests the airspeed did not vary
immediately with angle-of-attack changes that is, Tc
and airspeed remained constant when the model pitch-ed up
or down. If the thrust coefficient Tc instead of the
power had been kept constant in the force tests, there
would likely have been little or no change in stability
from the zero-thrust to the power-on conditions. The
assumption is here made that curves of pitching-moment
coefficient against lift coefficient at constant thrust
coefficient would have remained parallel for any value
of thrust coefficient; that is,
dC /dC,

L TO 0- dL/ Tc Any value
Since longitudinal st-adi.ness is largely dependent on the
rapid pitching motions or short-period oscillations that
cause no appreciable change in airspeed, the steadiness
i pears to be. affected principally by stability changes
that occur at conditions of constant thrust coefficient
and constant airspeed and very little by changes that occur
at conditions of constant powur and varying airspeed.
COI1FIDEITTIAL


COCI DE LIAL









XACA ACR ,T o. L5K05


Longitudinal Control

1 n.he loniitudira.l-control data obtained in the force
tests are shC'.n in fi.c' 1 0 Tese dana indicate that
with the nor-,al center f-r;. vity location the model could
be trei:.~d crc- se.o lift coefficient to ,axir..um lift
coefficient ,-ith a total elevator :over:ent of about 200
The elevator cffectivweness did not change noticeably when
pov..er -'as applied, which indicated that there was little
effect of inauced flo- over the elevators. Flight tests
show-.ed that the elevator was powerful enough to trim the
model over the entire flight range with the center of
gravity at l8 percent MoA.C.

Trim at High Angles of Attack

In the free-to-trim tests in the Langley 15-foot free-
spinning tunnel, the model -upon being released in the up-
right or Znverted position (at angles of attack of 900
or -90,) assu2lr-.d iedite.ly the angle of attack for which
the elevators had been set, T-Undr no conditions did
the model slov the tendency to trim at high angles of
attack that has been exhibited by so-me swept-back tailless
designs.


Lateral Stability

Force tests.- The results of tests made to determine
the lateral stability characteristics of the model are
presented in figures 1 and .1,. .These results are
su)ammnaized in figure 15 in the fort.: of a stability chart
that is a plot of the direct ional-stabilityv parameter Cn
against the effective-dihedral arar-eter C7
P

The data of figures I3 and 15 show the effect of the
various com onent parts of t he mo el on lateral stability.
The wing-fuselage combinat ion had slight directional
instability but *aes made slightly stable oy the addition
of the pusher propeller. A-dition of the vertical tails
increased the directional stability rith pororeller
v:'ind:iilling to a value of Cn of about 0.0007. The
effective di edral was small for all.conditions, about 20
for the wing-fuselage combination and about 10 for the
complete model.


CCI?':T. TIAL


CO, _T IDE 7IIAL









12 COI IDJ CIAL INACA ACR To. L5K05


The force-test data of figures 14 and 15 show a
noticeable increase in directional stability with appli-
cation of power. Increasing the power from idling to
2000 brake horsepower increased the directional stability
by approximately 65 percent. The data of figure 14 also
show that applying power with the vertical tails off did
not increase the directional stability appreciably. The
increase in directional stability at high power with tails
on a pears to be caused primarily by the inflow effects
upon the vertical tails rather than from action of the
direct propeller forces. The effective dihedral was appar-
ently not affected by an increase in power. (See figs. 14
and 15.)

Flight tests,- The lateral flight characteristics of
the model with both vertical tails installed were good for
powers ranging from zero thrust to 1200 brake horsepower,
The directional stability appeared to be satisfactory in
flights at zero thrust and improved with the application
of power. The small effective dihedral shown by the force
tests was noted in the flights by the absence of any appre-
ciable rolling motions when the model was disturbed in yaw
and by the negligible effects on aileron control of the
adverse yawing produced in rolling maneuvers. This small
effective dihedral was considered a desirable characteristic
for a tailless design because of the relatively lo. direc-
tional stability of this type of airplane.

In flights with the lower tail removed, the lateral
flight characteristics were not so good as those with both
tails installed. The adverse yawing due to aileron control
was greater and the yawing motions of the model damped out
more slowly after disturbances in yaw. The lateral flight
characteristics were considered not quite satisfactory with
this tail configuration.


Lateral Control

The force-test data showing the aileron effectiveness
are presented in figure 16. These data show that the
ailerons were effective at all angles of attack up to the
stall (aE,' :.::. 120).

In the flight tests adequate lateral control was
obtained by using abrupt aileron deflections of 150
Rudder deflections of & 120 -s-d in conjunction with the


C rlT DLE.'AL









V.'.,A ACR No. L5KO5 CO'TIDr'IAL


aileron control were usually sufficient to balance out the
adverse yawing moments caused by aileron deflection and
rolling velocity.


C OUICLUS IONS


The results of tests in the Langley free-flight
tunnel qf a straight-wing, tailless fighter model with
a 'pusher propeller may be sumLarized as follows:

S1. The general flight character-stics of the model
were good and compared favorably with the flight character-
istics of good conventional airplane models previously
tested in the Langley free-flight tunnel.

2. As the a ngle- of attack was increased, the longi-
tudinal stability of the model increased instead of
decreasing as that of tailless airplanes with swept-back
wings usually does.

3. Power caused a slight reduction in the longitudinal
stability measured at constant power. This reduction in
stability, however, did not affect the longitudinal stead-
inessof the model in flight tests.

4. The model did not show the tendency-to trim At very
high angles of attack (above the stall) that has been a
characte-ristic of some swept-back tailless airplanes.

5. The lateral flight characteristics of the model
with both vertical tails installed were good. The direc-
tional stability of the model was satisfactory and wds
improved by the application of power. The effective
dihedral was desirably small and was not appreciably
affected by power,

6. The control surfaces of the model provided ade-
quate longitudinal and lateral control,


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


CON'TD I`TTIAL








:TACA ACR No, L5K05


1. Stability Research Division: An Interim Report on
thre Stability and Control of Tailless Airplanes.
NACA ACR No. L o19, -194A.

2. Jones, Robert T.: '"otes on the Stability and Control
of Tailless Airplanes. NACA TN To. 837, 1941.

3. C.s:'bell, John P., and Seacord, Charles L., Jr.:
Determination of the Stability and Control Character-
istics of a Tailless All-Wing Airollane P'odel with
Sweepback in the Langley ?ree-Flight Tunnel. NACA
ACE No. L5A13, 195.

4. Trouncer, J., and Wright, D. F.: 'Iind Tunnel Tests
on the Effect of Variable Incidence Tips and Tip
Slats on Tailless Gliders. TN No. Aero 1496, British
R.A.E., Aug. 194i.

5. Shortal, Joseph A., and Osterhout, Clayton J.:
Preliminary Stability and Control Tests in the NACA
Free-FliLht Wind Tunnel and Correlation with Full-
Scale Flight Tests. NACA TN No. 810, 1941.

6. Shortal, Joseph A., and Draper, John W.: Free-Flight-
Tunnel Investigation of the Effect of the Fuselage
Length and the Aspect Ratio and Size of the Vertical
:1i on Lateral Stability and Control. rACA ARR
"To. D17, 19!,.5.

7. Zii-.r:?n, C. f.: Preli.lin ry Tests in the :TACA
Free-Spinnln.- Wind Tu.irl. :;ACA Rep. eo. 557, 1936.

8. House, R.f;s 0., and Wallace, Arthur R.: %Wind-Tunnel
Investigation of Effect of Interference on Lateral-
::.bility Characteristics of Four NACA 23012 ;iings,
an Elliptical and a Circular Fuselage, and Vertical
Fins.. ",..CA Rep. No. 705, 1941.

9. Campbell, John p., and Paulson, John W.: The Effects
of Static 7:ar-in and Rotational Dar..pin in Pitch on
the Lon-ritudinal Stability Characteristics of an
Airplane as Determined by-Tests of a modell in the
NACA Free-Flight Tunnel. :ACA ARR ;:.j. L4F02, 1944.


CC'TI C-7'IAL


C-OF"ID:E FIIAL









:".CA ACR I.o. L5K05 CO'!IDa:rfIAL 15


10. White, :'-.r.rice D.: Effect of Power on the Stick-
Fixed ;'Teutral Points of Sevcral Single-Engine
rMonoplanes as Determined in Flight. JTAC CB
Ho. Li~O1, 19 4.


C ON IDEIT IAL










NACA ACR No. L5K05 16




CONFIDENTIAL

TABLE I

FORCE TESTS OF STRAIGHT-WING, TAILLESS MODEL
IN THE LANGLEY FREE-FLIGHT TUNNEL

a e R
Type of data Configuration Power (deg) (deg) (deg) (deg) Figure

CL, C, Cm W Off Range 0 0 0 9
against a


Do-------- F ---dodo--- ---do--- 0 ------- ------- 9

Do.--.---- WP ---do--- ---do--- 0 0 0 9

Do-------- WFPVV2 T = 0 ---do--- 0 0 0 9

Do-------- WFPV1V2 T = 0 ---do--- 0 0, -15 0 10, 12

Do-------- WFPV1V2 1200 bhp ---do--- 0 0, -15 0 10, 12

Cn, Cl, Cy F Off 8 Range 0 0 13
against *

Do-------- WF ---do--- 5 ---do--- 0 0 13, 15

Do-------- WFP Wind- 5 ---do--- 0 0 13, 15
milling
Do-------- WFPVI ---do--- 5 ---do--- 0 0 13, 15

Do-------- WFPV1V2 ---do--- 5 ---do--- 0 0 13, 15

Do-------- WFPV1V2 ---do--- 12 ---do--- 0 0 13

Do-------- WFP ---do--- 5 ---do--- 0 0 14

Do-------- WFP 1200 bhp 5 ---do--- 0 0 14

Do-------- WFPV12 T = 0 5 ---do--- 0 0 14, 15

Do-------- WFPV1V2 1200 bhp 5 ---do--- 0 0 14, 15

Do-------- WFPV1V2 2000 bhp 5 ---do--- 0 0 l4, 15

C,' Cn WPPV1V2 Wind- 4, 8, 12 0 0 Range 16
against aR milling

NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS
CONFIDENTIAL











Fig. I NACA ACR No. L5K05











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CONFIDENTIAL


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NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS







0 S /O

c/c. in.


CONFIDENTIAL


Fgare3. Drcaigy oJScc/e sfcact-i-win/, ia///ess fyarer mode/
free- fight / urne/.


/es/eui'/n he Larg/ey,


NACA ACR No. L5K05


A,/eron -

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CONFIDENTIAL


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CONFIDENTIAL
0 4 a /2
Ang/e of a /ac/ cX deg


0 -
C/O fcn -m/, ent
coef9/c/ent, c,


Figure 9. Aerodyniamnc characdrnss/cs of s/ra/ghtf-wlMg a///ess ~ /er
mode/ and i/s components parf/3 esked /n longley free-f/gh// /,nne/.
Center-of-gra^1/fy /ocaf/on, 0.20A1.A.. ; q, 4.09 pours per
5qu/Lae )bo. NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS


^ K3





.0
0 \3-
o


I-


-c~-c~-9


-4


Fig. 9








NACA ACR No. L5K05


R --- 1 5- 0 /o0
S--- -- --A--- -/-S 120l







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S CONFIDENTIAL
:3
-I 0 4 8 /2 6 0 -
Angl/ of attacA,, or, atd9 P/tch~i-/omnen-t
ccffic-cenI, C/r
Fguqq /lO. Effecl of power on /ong//uadonal charalerls//cs ot
rg-wing fai//es5s fgher f rxe/ /eskd /n LiA'y /r ee-h -Y g/
-unnel. Cenr-of-qrvIty location 0oo046 alove /ru5t
I/ne and a/ 0,209 .A,.;q 4.09 pound- per ire
foot NATIONAL ADVISORY
COMMITTEE FOR AEROINAIIC


Fig. 10







NACA ACR No. L5K05


CONFIDENTIAL



7=0-o


SOO ___ /x bhp


(a) A, caused by prope//er norra/ ar-e.





/200 hhp

(b) AC, caUed by prope//er /hrus/.





Calculated
/easurea (data from fig. /0) colcI

o .2 .4 .6 .8 /0O
Liff coefFic/eni, CL
(c) Tofta ACm cauCed by opero fng prope//r ;
/200 broke horsepower. NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS
Fi/ure //. EF/-ec/ of prope//er /brces on //) /orfg -
UCd/ina/ s5/ab/iy of //e /rvg/9ht- wing, /a///ess
fighMer model tested //n ihe zang/ey ree -f//gh
tunnel. Cenier-of-gravriy /ocallon 0.20/A). .C.
and 0. 048 above /hrusl //ne .


CONFIDENTIAL


Fig. lla-c







NACA ACR No. L5K05


CONFIDENTIAL


(a) Cen-er-o/-gralty /ocaQOn, 0.048 C
above /hrusf //ne.


O .Z .4 .6 .8 /0
L/ff coeff/c/enf C_
(b) Cer <7 of-grayvfy /ocaf/on, 0.0//c
be/oPy /hrusf // 7le. NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS
r/gure /2.- Comparson of /he effects of power
for /wo center-of- grwivy /ocaifons. Da. fromn
/ests of s'ra/gh -w/ng, /a~//ess r~gh/er model/ /
//e Lanrg/ey free -f/ghl // Aunr/. Cener-of-grovIly
/ocaoCN 0.20 /A.. FL.
CONFIDENTIAL


Fig. 12a,b








NACA ACR No. L5K05


(_ (I-4C fL/ ///l/
El- S WY-
0---- F,

0---- WFP
72 1 A----5 t/FPV
V- -- 2 WFPV,











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17 A? E2 __

















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1~ -------------






2 _____ __-___ ___



.0/





S- NATIONAL ADVISORY
0 COMMITTEE FOR AERONAUTICS
-4 O 4 /Z /6 2 O
> .0 / -- -- -1t




Ang/e of yaw, deg

Fgure /3.- Leral / slabv//y char2czcier/s//C of ara/oh/- /ng
Ia//ess f/gh/er moode/ /eed /n LZoQ/ey 'ree-f//gi/ unnel
q, 4.09 pounds per qualre foo/ ; four-blade ,rope//er.
CONFIDENTIAL


Fig. 13








Fig. 14 NACA ACR No. L5K05

CONFIDENTIAL
.2








1 __/,a// Rower
>o0 ----TF'^--- --


So---O-ff Tc-O 0 4
S-- Off /120ZhO 4093
o-- O ---0n i=0 409Q
A---O- /20oMp 4093
2 V_ ---.On 29M/o 1902

02












.02
S.0/ .- ___











-0/
-2O -O1 0 /0 20 30
A 'e of yaw, degq
NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS
Figure 14. Effect of power on laema/l jal/i/ly ,-fc2raCder/s,'cs
of sthiwght- wing, ull/ess fighter n 4e/ fested in /Fe
Longley free- ,.i/ht lannel. oC. 5 ; C, 0.5.
CONFIDENTIAL







NACA ACR No. L5K05


.00/2


.0006


C



.0002


.A902



0



-.0002


CONFIDENTIAL


CWF PV, V 2000 o bhp


8)WFPV V, /200 bh




j- IFPV V ; prope//er ilm///ng-
Tor 7 =0



*14'VFPJ porpe//er wfdrm///ng







WFP prope//Ar wVIdnd/n/9

NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS
.WF, power of
___T~I I I _


,roz


- C02
- C2.


.0006 -.0O8 .00/0


Fi9ure /I -_ Va/es of /a/Xera/- s/ab/hly parare/ers Cn,
and C fo'r various conf/,ura/1on5 of IheC ,/r6gh/ -
wing Ar///es5s fgh/er mode/ 9es/ed in //e Lang/ey
free-f/7/gh/ unne/. CONFIDENTIAL


Fig. 15








NACA ACR No. L5K05


CONFIDENTIAL


-/0
RIght al/eron


I NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS
I I I
0 /0 20
clef/echon, 5a R c


gure. /6. A//elron effect/veess of sf/19ht- v/ng,
/tj//ess5 fighter mtxode/ tested / Lang/ey free -
f/ight /unne.l
CONFIDENTIAL


0




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.02



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Fig. 16


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UNIVERSITY OF FLORIDA

3 1262 08106 454 4

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CD:.;AEi ITS DEP.-\FTMEINT
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