Determination of the stability and control characteristics of a tailless all-wing airplane model with sweepback in the L...

MISSING IMAGE

Material Information

Title:
Determination of the stability and control characteristics of a tailless all-wing airplane model with sweepback in the Langley free-flight tunnel
Alternate Title:
NACA wartime reports
Physical Description:
16, 16 p. : ill. ; 28 cm.
Language:
English
Creator:
Campbell, John P
Seacord, Charles L
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 power-off stability and control characteristics of a tailless all-wing airplane model with sweepback has been made in the Langley free-flight tunnel. The results of the free-flight tunnel tests were correlated with results from force tests made at high Reynolds numbers in order to estimate the flying characteristics of the full-scale airplane. The investigation consisted of force and flight tests of a 4.3-foot-span dynamic model. The effects of flap deflection, center-of-gravity location, and addition of vertical-tail area were determined. The following conclusions were drawn from the results of the investigation: The full-sale airplane will undergo a serious reduction in stick-fixed longitudinal stability at high life coefficients unless early wing-tip stalling is eliminated. The directional stability of an all-wing airplane without vertical tail surfaces will be undesirably low. The effective dihedral of an airplane of this type should be kept low. An elevon and rudder control system similar to that used on this design should provide sufficient control.
Bibliography:
Includes bibliographic references (p. 14).
Statement of Responsibility:
by John P. Campbell and Chalres L. Seacord, Jr.
General Note:
"Report no. L-50."
General Note:
"Originally issued February 1945 as Advance Confidential Report L5A13."
General Note:
"Report date February 1945."
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 - 003616137
oclc - 71293720
sobekcm - AA00006255_00001
System ID:
AA00006255:00001

Full Text

pofr-
ruI
r


NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS





WARTIME REPORT
ORIGINALLY ISSUED
February 1945 as
Advance Confidential Report I5A13

DETERMINATION OF WTE STABILITY AND CONTROL CARACTIRISTICS
OF A TAILLESS ALL-WING AIRPLANE MODEL WI= SWEEEBACK
IN TEE LAGLE! FREE-FICGHT TU'H L
By John P. Campbell and Charles L. Seacord, Jr.

Langley Memorial Aeronautical Laboratory
Laenley Field, Va.











WASHINGTON
I.
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-
iically edited. All have been reproduced without change in order to expedite general distribution.


'OCUMEINT, DL.PARItM W j


7


L .SO



































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/determinationofs001ang








TACA ACR No. L5A13


NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS


ADVANCE CCuTIrTF'TTIAL PORTT

DETERMINATION OF THE STABILITY A.TD CONTROL CHARACTERISTICS

OF A TAILLESS ALL-WING' AIRPLAfE MODEL WITH SWEEPBACK

IIT THE LANGLEY FREE-FLIGHT TUNrIL

By John P. Campbell and Charles L. Seacord, Jr.


SUMMARY


An investigation to determine the power-off stability
and control characteristics of a tailless all-wing air-
plane model with sweepbac': has been made in the Langley
free-flight tunnel. The results of the free-flight-
tunnel tests were correlated with results from force tests
made at high Reynolds numbers in order to estimate the
flying characteristics of the full-scale airplane.

The investigation consisted of force and flight tests
of a 1.5-foot-span dynamic model. The effects of flap
deflection, center-of-gravity location, and addition of
vertical-tail area were determined.

The following conclusions were drawn from the results
of the investigation: The full-scale airplane will
undergo a serious reduction in stick-fixed longitudinal
stability at high lift coefficients unless early wing-tip
stalling is eliminated. The directional stability of an
all-wing airplane without vertical tail surfaces will be
undesirably low. The effective dihedral of an airplane
of this type should be kept low. An elevon and rudder
control system similar to that used on this design should
provide sufficient control.


INTRODUCTION


The desire to obtain improved performance for mill-
tary airplanes has recently increased the interest- in
tailless-airplane designs. One of the most promising
tailless designs, from the considerations of performance,


C O NF IDE NT I AL


COUFIDE!TIAL








NACA ACR No. L5A13


is the large all-wing airplane or "flying wing." Inherent
in the all-wing airplane, however, are certain undesirable
stability and control characteristics that must be elimi-
nated before this design can be considered satisfactory.
In or'ter to study these stability and control character-
istics and to find means of improving them, an investi-
gation is being conducted in the Langley free-flight
tunnel (designated FFT) of a free-flying dynamic model of
a tailless all-wing airplane with sweepback.

The present report gives the results of force and
flight tests of the model with windmilling propellers.
Tests were made with the lift flaps retracted and
deflected. For some tests, auxiliary vertical tail sur-
faces were installed on the model. The effects of changes
in the center-of-gravity location and trim lift coeffi-
cient on the flight characteristics of the model were
determined.

In order to estimate the flying characteristics of
the full-scale airplane, the test results were correlated
with results of force tests of a similar design run at
high Reynolds numbers in the Langley 19-foot pressure
tunnel (designated 19-ft PT).


SYMBOLS


The following symbols are used herein:

CL lift coefficient (Lift/qS)

CD drag coefficient (Drag/qS)

C, pitching-moment coefficient (Pitching moment/qc-S)

C( rolling-moment coefficient (Rolling moment/qbS)

On yawing-moment coefficient (Yawing moment/qbS)
Cy lateral-force coefficient (Lateral force/qS)

c chord, feet

t mean aerodynamic chord, feet


CO NFIDE NTAL


CnN IDENTICAL









NACA ACR I'o. L5A13


S wing area, square feet

b wing span, feet
q dynamic pressure, pounds per square foot (pV
V airspeed, feet per second

p masa density of air, slugs per cubic foot
p. angle of sideslip, degrees

,r angle of raw, degrees (for force-test data,
= -0)

a angle of attack, degrees

h static margin, distance i.n chords from center of
gravity to neutral -o'nt
pb
2- hel.x angle generated by wing tip in roll, radians

p rolling angular velocity, radiars ner second

C7 rate of cl ange of rolling-moment coefficient with
helix angle


Cn rate of change of vvawing-:noniant coefficient with
V angle cf 3ideslip, npr degree 6C,/6

C1 rate of change of rollinf-moniant coeffic'iont with
angle of sideslio, crr degree (aCL/6)

6f flap deflection, degrees

6e elevon deflection, positive down, degrees (with
subscripts r and L to indicate right and
left elevon, respectively)

6r rudder deflection, positive down, degrees (with
subscripts r and I to .-ndicate right and
left rudder, respectively; if both right and
left top rudder surfaces are ,Jlflected
simultaneously as longitudinal trim flaps, no
subscript is used)


CON 7TDENTIAL


CONFTDEDITIAL










PAACA ACR No. L5A13


Reynolds numnleer


APPARATUS


The investigation was made in the Langley free-
flight tunnel, which is described in reference 1. A
photograph of the test section of the tunnel showing
the model in flight is presented in figure 1. Force
tests to determine the static stability characteristics
were made in the Langley free-flight tunnel with the
model mounted on the six-component balance, which is
described in reference 2.

The mass and dimensional characteristics of the
model are as follows:


Weight, pounds .. .
Wing area, square feet . .
Scan, feet . .
Aspect ratio .. . .
Wing loading, pounds ner square foot .
Radius of gyration in roll,- kX, foot .
Radius of gyration in nitch, ky, foot .
Radius of -'rration in y.'w, kz, foot .
Mean aerodynamic chord, foot . .
Sweepback of 0.25-chord line, degrees .
Dihedral, degrees . .
Taper ratio (ratio of tip chord to root chord)


Root chord, foot .
Tip chord, foot .
Elevon:
Type . .
Area, percent wing area .
Span, percent wing span .
Rudder:
Type .. .
Area, percent wing area .
Span, percent wing span .
Vertical tails:
Type . .
Area, percent wing area .
Asnect ratio .


. .


.
" .
. .
. .


. 2.55
. 2.51
. .5
. 7.56
. 1.02
. 0.78
. 0.55
. 0.82
. 0.655
. 22.00
. 0
. 0.25
. 0.937
S. 0.234


Plain

5. 00
50.co


. Slit, drag
. 2.86
S. .. .. 20.00

.* Twin center fins
S. .. 4.00
S. 2.00


COFMTD71UTTAL


CON r IDEN TIAL








NACA ACrl 'o. L5A15 CONFIDENTIAL 5


Airfoil section . .. .Modified NACA 105
Root, percent thickness .. .. 21
Tip, percent thickness . 15
Geometric twist, degrees . 6
Aerodynarmic twist, degrees .. (approx.) t

The component parts of the model are identified in
the tables and figures as follows:

Wing . . W
Propeller shaft housings. ... .. H
Propellers . . P
Vertical tails; two tails mounted on nacelles,
each tail having 2 percent of wing area V
Split flap (center-section lift flap, 6f = 60) F

Combinations of these letters represent the combination
used in the tests. The standard configuration is desig-
nated WHP. A three-view drawing of the model is
presented in figure 2. Photographs are given in figures 3
and L. In plan form the wing has both sweepback and taper
and has a split flap that extends from the center line of
the airplane to the inboard ends of the elevons. For all
flap-down tests, the flaps were deflected 600,

The control surfaces consist of elevons that extend
b b
from 0.55- to 0.71- and split rudders (fig. 5) that
2 2
b b
extend from 0.712 to 0.91-. The split rudder is
so linked with the eleven that in flight tests the lower
surface of the rudder moves down with the downgoing elevon
and the upper surface moves up with the upgoing elevon.
This linkage arrangement provides additional effective
aileron- and elevator-control-surface area as shown in
figure 6.

The upper surfaces of the split rudders can be
deflected upward simultaneously to serve as trim flaps
to provide pitching moment for longitudinal trim when
the lift flap is deflected. The lower surfaces of the
split rudders remain at zero when the top surfaces are
deflected as trim flaps.

The controls of the model were operated in flight
by electromagnets in the same manner as described in
reference 1.


CONFIDENTIAL








IIACA ACR No. LSA13


For some tests vertical tail surfaces having a
combined area of h percent of the wing area were mounted
on the propeller-shaft housings to provide additional
directional stability. (See firs. 2 and 4.)

For propeller-on tests the model was equipped with
two freely windmilling two-blade pusher propellers.

A modified NACA 103 airfoil with a thickness of
21 percent at the root and 15 percent at the tip was used
on the model. The trailing edge was reflexed enough to
give a slightly positive pitching moment at zero lift.
This airfoil was used to obtain a maximum lift coefficient
in the free-flight (low Reynolds number) tests more nearly
equal to that of a full-scale airplane than is possible to
obtain with other airfoils (especially low-drag airfoils)
at low Reynolds numbers.

The free-flight-tunnel model was almost identical
in plan form to the model used in the tests at higher
Reynolds numbers in the Langley 19-foot pressure tunnel.
The models differed in airfoil section, geometric dihedral,
and geometric twist. The airfoil sections of the model
tested in the Langley 19-foot Dressure tunnel were
NACA 65(318)-019 at the root and 65(518)-015 at the tip;
the geometricc dihedral of this model was 20 compared
with 0 for the free-flight-tunnel model. The model used
in the langley 19-foot pressure tunnel had 40 geometric
twist, whereas the free-flight-tunnel model had a geometric
twist of 6. The aerodynamic twist for both models,
however, was approximately 40.


TESTS


Force tests were made to determine the stability and
control characteristics of the model with flaps retracted
and deflected. The moments were comrputed with the center
of gravity at 0.25 mean aerodynamic chord and are referred
to the stability axes. The stability axes are defined as
an orthogonal system of axes 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 perpen-
dicular to the Z-axis, and the Y-axis is perpendicular to
the plane of symmetry. The conditions in which force
tests were made are given in table I.


COIFIDE JTIAL


CCON IDEN IAL









NACA ACR No. L5A13


Flight tests were made at lift coefficients varying
from 0.3 to 0.8 with flaps retracted and from 0.6 to 1.1
with flaps deflected. The center-of-gravity position was
varied from 20 to 25 percent of the mean aerodynamic chord
for flight tests in both the flap-retracted and flap-
deflected condition. Table II gives the conditions for
which flight tests were made.


RESULTS AND DISCUSSION


In interpreting the results of the free-flight-tunnel
tests of the tailless all-wing model the following points
were considered:

(1) The tests were made at very low Reynolds numbers
(150,000 to 550,000); hence, the results of the tests of
a similar design made at high Reynolds numbers
(about 6,600,000) were used in estimating the flight
characteristics of the full-scale airplane from the free-
flight-tunnel test results.

(2) The controls of the model were fixed except
during control applications; hence, no indications of the
control-free stability of the design were obtained.

(3) In determining the control effectiveness of the
design, no consideration has been given to control forces.

(4) No power was applied to the propellers during
the tests. The results, therefore, cannot be used to
predict power-on stability.


Longitudinal Stability

Force tests.- The results of force tests made to
determine the longitudinal stability and control charac-
teristics of the model are shown in figures 7 and 8. On
these figures, data from tests of the model of similar
plan form tested at high Reynolds numbers are also plotted.

The slope of the pitching-moment curve for the flap-
retracted condition of the free-flight-tunnel model
changes from negative to positive with increasing lift
coefficient. This change in slope indicates a change to


CONFIDENTIAL


CONFIDENTIAL









ITCA ACR No. L5A15


static longitudinal instability at high angles of attack.
This change in stability is characteristic of swept-back
wings because of the tendency of the wing tips to stall
first. The instability appears to be much greater for
the free-flilht-tunnel model than for the similar model
tested at high Reynolds numbers. This difference is
probably explained by the fact that the difference in the
Reynolds numbers at the root and tip sections of this
design causes a much greater difference in stalling char-
acteristics on the small-scale model than on the model
tested at high Reynolds numbers.

For the flap-deflected condition (fig. 8), the
pitching-moment curves for the free-flight-tunnel model
were very similar in shape to those obtained with flaps
up but did not turn up at high lift coefficients as much
as the curves for the flap-retracted condition. The data
of figure 8 indicate that most of the change in shape of
the pitching-moment curve from flap up to flap down was
caused by the upward deflection of the trim flap. The
flap-deflected pitching-moment curve from high-scale tests
(fig. 8) indicates practically no change in longitudinal
stability with increasing angle of attack.

The difference in the angles of zero lift indicated
in figures 7 and 8 for the two models is probably caused
by the difference in the location of the chord line from
which the angle of attack is measured. The difference in
the slopes of the lift curve is probably a result of the
difference in the Reynolds numbers of the tests. It is
unlikely that these differences in lift characteristics
would cause appreciable differences in longitudinal flight
characteristics.

Flight tests.- The longitudinal stability as noted
in the free-flight-tunnel tests was satisfactory up to a
lift coefficient of 0.7 with flaps retracted and 1.1 with
flaps deflected with the normal center-of-gravity location
(25 percent M.A.C.). Above these values of lift coeffi-
cient, however, difficulty was experienced in flying the
model because of a tendency to nose up and stall after
disturbances in pitch. This behavior was believed to be
a direct result of the change in longitudinal stability
at high angles of attack, which was indicated in the
force-test results by the change in slope of the pitching-
moment curve. Although at times the pilot could prevent
the nosing-up motion by applying down-elevator control,


C ONFIDENTIAL


CONF IDENTICAL








NACA ACR ro. L5A15


the nosing-up tendency was considered a very objectionable
characteristic that would probably prove dangerous for a
full-scale airplane. This nosin -up tendency should be
expected on any airplane having pitching-monent character-
istics similar to those of the model. (See fig. 7.)

The longitudinal stability of the free-flight-tunnel
model was satisfactory at those lift coefficients at which
the static margin h was 0.04 or greater (C = 0.7,
flaps retracted; CL = 1.1, flaps deflected) and flights
were possible at conditions at which the static margin
was as low as 0.02. On the basis of the force-test
results it appears that the static longitudinal stability
of the corresponding airplane at hi-h angles of attack
would be greeter than that of the free-flight-tunnel model.
The data of figures 7 and 8 indicate that the airplane
with the normal center-of-gravity location would have a
static margin of 0.04 up to a lift coefficient of 1.0 with
flaps retracted and up to the stall with flaps deflected.
The stick-fixed longitudinal stability of this particular
airplane design, therefore, would probably be satisfactory
for all power-off conditions except at high lift coeffi-
cients with flaps retracted.


Longitudinal Control

The force-test results presented in figures 7 and 8
indicate that the longitudinal control provided by the
elevons was sufficient to trim the rodel over the flight
range for flap-retracted or-deflected condition with a
total elevon deflection of about 200. Inasmuch as the
force-test results of the model tested at high Reynolds
numbers indicate much more powerful eleven control than
was obtained with the model at low reynolds numbers, it
is probable that the elevator control for the full-scale
airplane will be satisfactory in flight.

In the flight tests, the model could be trimmed over
the speed range with a total eleven deflection of about 20q
For the flap-retracted condition, the upper surfaces of
the split rudders were deflected with the elevons for
longitudinal trim. Abrupt eleven deflections of *50 from
the trim setting provided adequate longitudinal control
for keeping the model flying for all stable conditions.


CONFIDENTIAL


CON FIDENTIAL








NACA ACR No. L5A13


On this design it is possible that the most critical
condition for elevator control will be at take-off.
Unless careful attention is given to the location of the
landing gear, the elevens alone may not be powerful enough
to meet the Army requirements for getting the nose wheel
off the ground at 80 percent of take-off speed. Use of
the trim flaps in conjunction with the elevons will help
provide enough longitudinal control to meet this
requirement.


Lateral Stability

Force tests.- The lateral stability characteristics
of the 'lodel as determined by force tests are shown in
figures 9 to 11. The values of the effective-dihedral
parameter Cp and the directional-stability parameter Ch
obtained for the different test conditions from these
figures are plotted in figure 12 in the form of a stability
diagram. The values of Cn a and C1p for corresponding
conditions for the model tested at high Reynolds numbers
are also presented in figure 12.

The values of C for the flap-retracted condition
at angles of attack of 0 and 60 are relatively low
(about 0.00030). Increasing the angle of attack to 120
with flaps retracted caused an increase in Cn to 0.00055.
np
This increase in Cnp with increase in lift coefficient is
characteristic of a swept-back wing.

The lower values of Cn shown in figure 12 for the
model tested at high Reynolds numbers are attributed to the
lower drag of this model. For an all-wing tailless design
with low dihedral, the drag of the wing contributes a
major part of the static directional stability.

The values of C~ shown for the free-flight model
in figure 12 correspond to an effective dihedral angle
between o2 and 40. The value of Ct increased with
increasing lift coefficient as expected for the swept-
back wing. The higher values of CI for the model
tested at large Reynolds numbers is caused by the fact


COIFIDEITIAL


C ON F IDEN TRIAL









I'ACA x CA i0. L5A13 CONMID.?.AL 1i

that this model had 20 geometric dihedral whereas the
free-flight-tunnel model had 0o geometric dihedral.

Flight tests.- The lateral stability characteristics
of the -.-odel noted in flight 'were fairly satisfactory
except for low directional stability in the flap-retracted
condition. This low directional stability 'as shown
principally by slow lightly damped yav:ing oscillations
that '..ere started by gust or control disturbances. The
directional stability was not dangerously low, however,
inasn:uch as neither divergences nor unstable oscillations
were noted. The adverse ya' in- noted in flights in which
aileron control alone was used was quite snall because
the elevons were deflected upward to-ether for longitu-
dinal trim and therefore operated as "trimmed-up" ailerons,
which usually produce only small yawing moments.

Deflection of the flans or addition of the vertical
tails caused noticeable ir.lrover.ment in the oa:nring of the
yawing motion of the model, and t'-e lateral stability
characteristics at these conditions -:ere considered
generally satisfactory.

The effective dihedral of the 'hodel appeared to be
satisfactory, inasmuch as no excessive rolling during
sideslip v.as noted and the li-htly damned ya.vinr oscil-
lations were accompanied by very little rolling. Previous
free-flillgt-tunnel investigations have shown that, for an
airplane with low directional stability, low effective
dihedral is necessary to avoid a :,oorly damped rolling
(Dutch roll) oscillation.

It is probable that the lateral stability character-
istics of a full-scale airplane of the design tested
would not be so good as those of the free-flight model
because the values of Cnp of a full-scale airplane will
probably be lower than those for the free-flight model.
At the higher lift coefficients, which could not be reached
in the free-flight-tunnel tests because of longitudinal
instability, the requirements of the airplane would be more
severe for directional stability and the airplane would
probably be considered unsatisfactory in this respect. In
order to secure satisfactory flying characteristics with a
tailless all-wing airplane of this type, it appears
desirable to maintain a low value of effective dihedral and
to supplement the directional stability of the wing by
means of vertical tails or an automatic stabilizing device.


CON FTDE r T AL








lACA ACR No. L5A15


Lateral Control

Aileron control.- The aileron control provided by
the elevens appeared to be weak in the flight tests.
Abrupt eleven deflections of *150 did not provide satis-
factory aileron control in flight. Previous free-flight-
tunnel tests have shown that, if aileron deflections
greater than *150 are required for satisfactory control
on a model, the ailerons on the corresponding airplane
are likely to be weak.

A better quantitative indication of the weakness of
the aileron control was obtained in the force tests, the
results of which are presented in figures 13 and l and
which are summarized and compared in figure 15 with
results of tests at high Reynolds numbers. Computed
values of the helix angle pb/2V produced at different
lift coefficients by various eleven deflections are shown
in figure 15. The values of pb/2V were obtained by
multiplying the force-test values of rolling-moment coef-
ficient by 0.8/C, (See reference 3.) The high
Reynolds number data of figure 15 indicate that the
flying-qualities requirement for a :;Aininum value of 0.07
for pb/2V is not met by this design at lift coefficients
above about 0.l with *150 elevon deflection. The free-
flight-tunnel force tests indicate even weaker aileron
control but this result is partly attributed to the low
Reynolds number of the tests, to the wing section used,
and to the initial reflex of the trailing edge of the
win7. The free-flight-tunnel test results do indicate,
however, that linking the rudder surfaces to move as
ailerons with the elevens provides a substantial improve-
ment in aileron control.

In order to obtain satisfactory aileron control with
eleven surfaces located well inboard of the tip as on this
design, larter-chord surfaces than those on the free-
flight-tunnel model should be used or the rudder surfaces
should be linked with the elevens in order to provide
greater effective eleven area.

Rudder control.- The split rudders on the model
provided sufficient yawing moments to balance out the
adverse yawing moments encountered in the flight tests
during aileron rolls. Inasmuch as the yawing moments
caused by aileron deflection were small (fig. 14) because
of the initial upward deflection of the elevens for


C O' .TTLE~' TAL


COT.' r FDT!NTT IAL









tlAC.: .-',C No. L5A15


longitudinal trim, the rudder yawinr moments only had to
oponse the adverse yawing moments caused by rolling. The
adverse yawing moments caused by rolling were apparently
small for the model, as indicated by the small amount of
adverse yawing in flights with rudders fixed and elevons
alone used for control. These results indicate that the
rudder control of this all-wing airplane should be adequate
during normal flight.

Usually the most severe requirement for rudder con-
trol of nmrltiengine airplanes is that the rudder control
balance the asymmetric yawing moments introduced by the
failure of one engine during a full-power climb. Calcu-
lations based on the force-test data presented in figure 16
indicate that, with rudders of the size and type used on
this design, an airplane of this type having a 150-foot
span and two 5000-horsenower engines would meet the Army
requirements for maintaining steady flight with 100 or less
sideslip at 120 percent of the stalling speed with one
engine inoperative and the other engine operating at full
power.


CO TCLUSIONS


The following conclusions concerning the power-off
stability and control characteristics of large all-wing
tailless airplanes with sweepback were drawn from the
Langley free-flight-tunnel test results and from a corre-
lation of these results with results obtained from force
tests made at high Reynolds numbers:

1. Stick-fixed longitudinal instability at high lift
coefficients,or at least a serious reduction in longitu-
dinal stability,should be expected for airplanes of this
type unless the premature stalling of the wing tips is
eliminated. The upward deflection of a trim flap at the
wing tip will reduce the tendency of the tips to stall
first and will thereby improve the longitudinal stability
at high lift coefficients.

2. The directional stability of this type of airplane
without vertical tail surfaces will be extremely low.
Although the airplane will be flyable, it will probably
not be considered entirely satisfactory because of the
tendency to sideslip to large angles following slight gust
or control disturbances.


CONFIDEi7TIAL


C ON F ITE i'T!AL








":PCA ACR No. L5A13


5. The effective dih-dral of an airplane of this
type should be kept low in order to minimize the amount
of rolling accompanying the lightly damped '"awing oscil-
lations that are likely to be encountered.

4. An elevon and rudder control system similar to
that used on the model in these tests should provide
sufficient longitudinal and lateral control for an
airplane of this type.


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


RE 2"3 i":S'


1. Shortal, Josenh A., and Osterhout, Clayton J.:
Preliminary Stability and Control Tests in the
NACA Free-Flight Wind XTnnel and Correlation with
Full-Scale Flight tests. NACA TN No. 810, 1941.

2. 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
Tail on Lateral Stability and Control. NACA ARR
iio. 3D17, 1943.

3. Kayten, Gerald G.: Analysis of Wind-Tunnel Stability
and Control Tests in Terms of Flying Qualities of
Full-Scale Airplanes. NACA A-rR :To. 3J22, 1915.


CON;PIDT1 TIAL


CONPIDNT) IAL








NACA ACR No. L5A13 15


CONFIDENTIAL

TABLE I
FORCE-TEST CONDITIONS FOR TAILLESS ALL-WING AIRPLANE MODEL
IN THE LANGLEY FRLE-FLIGHT TUNNEL


a Configu- 6, 6, 6r
Test ration e 5 r Figure
(deg) (deg) a (deg) deg) (deg)


1 -4 to 20 0 WHP 0 0 0 7
2 -4 to 20 0 WHPF 0 0 0 8
3 -4 to 16 0 WHP -10 -10 -10 7
4 -4 to 16 0 WH PF 0 -40 -40 8
5 -4 to 16 0 WHPF -10 -4o -410 8
6 0 -30 to 30 WHP 0 0 0 9
7 6 -50 to 50 WHP -10 0 0 9
8 12 -50 to 30 WHP -20 0 0 9
9 8 -50 to 50 WH.PF -10 -40 -40 10
10 6 -50 to 30 WHPV -10 0 0 11
11 6 -30 to 30 W -10 0 0 11
12 0 t12 to2 0 IEP -10 0 0 13
(Right only
15 0 to 12 0 WHP (R o0 0 13
(Right only
11^ 0 to 12 0 WHP (Righ20only3 0 0 13
-20
15 0 to 12 0 WBP (Right only) 0 0 13
15 0 to 12 0 WHP 0 0 10 13

18 0 to 12 0 WHP 0 0 -20 14
18 0 to 12 0 WHP 0 0 -20 114.
19 0 to 12 0 WHP 0 0 20 14
20 0 to 12 0 WHP 0 0 20 16
21 0 to 12 0 WHP 0 0 o40 16
22 0 to 12 0 WHP 0 0o 60 16

aExplanation of configurations is given in section on
"Apparatus."
CONFI DENTI AL
NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS








'TACA AC? !o. L5A15


TABLE II

PLI(rHT-TEST CO(WDITTOYNS OF TAILTLSS ALL-'H71TG AIRPLANE

TIODEL IN LAITGLEY FR E-PLIGHT TUNNEL


Lift coefficient Configuration Center-of-gravity location
(a) (percent M.A.C.)


0.5 to 0.8 WHP 0.25

.6 WHPV .25

.6 WHPV .22

.5 'NHPV .20

.6 to 1.1 3VHPVF .25

.6 to 1.1 W-PF .25

.7 WH PVF .22

.7 'T PVF .20


explanation of configurations
"Apparatus."


given in section on


NATIONAL ADVISORY
COI!ITTEE FOR AERONAUTICS


CONFIDENTIAL


CO[NFID' .TIAT.











NACA ACR No. L5A13 Fig. 1







-,



,-4

t o



i;;r%--n~- 1--:*1* ----I


-4




bo
.rq
be





























bo
-9
rz.
!i
.



-c~!








NACA ACR No. L5A13 ig





\ *'
^ ___L-i ^
J51~ig /J ^

L U- i.







to
^I -- aUi1









"Nby
IN
b- \ |



'I'n ?1 a
',~


W''3
N)

td 1 \ \\ \ \ EbI e
I \ I 0 i






'ii \
I ~ N s





-M1 I i 1


I I I \
bl k \






j \ LiV



-.\ \ -


a fCs


Z











NACA ACR No. L5A13 Fig. 3















O

*MC
E-4


CC






'-b
z














00












o
C
E












,4
'.
0


C)(
.1-










NACA ACR No. L5A13 Fig. 4








,-I









!-4 0)
o
*H 0




'"C






E-4



0'






S. .,-4

.4
La E
















S,-14







ou a) l







r .
cO
0 )h
0)0)



O |cl .4 )



















r ..
ta
*4b











NACA ACR No. L5A13 Fig. 5










',-
ba











C)
tz
(U


0


E

m













to



be
Cd
I
,-.
































b









-4.
-l





-I
rl





















CONFIDENTIAL


-7


Rign all
left lo-er rudder surface flongrtudina
adds to eievon area


eron control
L trin flarE 001


PLgr, ulTer rudder surface
adai [E: e alvon 5rea


Right aileron and ruaaer control
I lng tudinal trim flaps O' I


If rudd3er deflection is greater
than elr von *efl-ction, deflected
rudder Eurface &ar- not moved by
elA i; ns


Fgcnt allern control
lloir, itudinal trim riapE -41 1 I'Per rudder surfaces deflect
itr. elevone only for elevon
deflectir.ns greater than -40


Right aileron and rudder ccirrol For rulder de'flectlon, upper
Ilongi tu1. na tria flapE -40" rudder Surfac- deflects from
-4C' tr, flait position


When elevens are used as -levators,
tie upoer rudder ourfaceF reflect
Wrih the eiaonr, NATIONAL ADVISORY
CUMMIIILE FOR AERONAUIICS




Figure 6.- Flevor and ruader arrangement uzed to obtain ddaittonal ailerron -ffect. eness.


CONFIDENTIAL


NACA ACR No. L5A13


Fig. 6






Fig. 7 NACA ACR No. L5A13







-e Source CONFIDENTIAL
/14 0 () of "aL
0 0-FIT
-- 0 /.9-f PT
__ n _/p 4/ RT r


/0
10 4 A I I I" rr






0 O --------j
s ^ -- -^ ------- a-

0 0/ ^&1


-4 4 8 / /6 20 0 -/
Ang/e of c/iac/r, o(C <.,.eff/ce.'~;'/i C-n
figure 7.- Zf/, drag, oaXc py/c//no-r men c/ x.acter-
/s/cs of n7ode/s of /~a//ess c//- nwin9 a/rp/a e W/
sweeplctur. F/Aaps rerac-it (&VHP).






NACA ACR No. LEA13 Fig. 8


(d ef r CONFIDENTIAL'







0 -40 -fT C_





/0/
-- -/O -40 F7
16 ---A O -jo 9- PT <
-0 -9 -7. -5 f P7- <




/2

// /


.4_.



1' --- -^ -'-" --- -,t-
'I6 2 b' w : V^ Z


Figure l///, drag, and p//c/n/ng-momrent character-
isT/cs mo ces ofm /at//ess oal- w/in orpo/ne
With sweepb~ack. F/aps deftec ted c(WHPF).







NACA ACR No. LEA13


Z .02
- 02
O


SC -.2







-.02




.04


u *102






a -.02


-04-_


- --- -c/-20--
CONFIDENTIAL






-0 / -0






------











---- -




CONFIDENTIAL
I I I


-20 -/O 0 /0 20 J0
Ang/e of yaw, ~u, ds
NATIONAL ADVISORY
COMMITTEE FOR AERONAUnCS


Figure 9. -Effect- of aro/e of atftck on the blaeral
stability chQracter/sf/cf of /e lA e.ng/ey free-f//hb/-A/n/
mode/ of 0 he fat//ess a/7-nq a/rp/ane w/th sweep-
tack Fl/ps re f raced (WHP O.=


Fig. 9


30








NACA ACR No. L5Al3


S- CONFIDENTIAL -




-3 ______OC e ,
__ C- -/0 0 0

S02 8__ [- ---08-/0O -40 60



^-S .0_ -- -

C-02



.04 --




' .0o2 -- --
0 ,




.0- CONFIDENTIAL

-.04 1 I I I I
JO -20 -/0 0 /O 20 O3
Angle of yaw, P dg
NATIONAL AWSOY
CODMITTM FOR AERONAUTICS
Figure /0- Effect of f/qs on l/oerat / s/abiify c/trac-
ter/sf cs of Lon'?y free -flh/' -unrn/ mode/ of
ta//ess5 a//-w/nq airplane w/lh sweepbacA.


Fig. 10







NACA ACR No. L5A13


CONFIDENTIAL



S- --- Nacelles of
--O Nacelles on
0 7_-,jl nn



-. 0 2










0 0-
P~oz ----- ---[










-30 -20 -/0 0 /0 80 30
,ng/e of yaw ) deg
CONFIDENTIAL

Figure //.- Effect of prope//er /us/ngs arncid verticol
ftols on /oteral sa2b/lt/y c/taracter/,stcs of langley /ee-
f/ght'- funnel/ model o/f t/lless o//-wing oarpolne
with sweepbach. OC = 6 CL 0.70.


Fig. 11






NACA ACR No. L5A13


li









-I Z





I ;*j
_ > '


'9LI


Fig.- 12







NACA ACR No. L5A13


. CONFIDENTIAL


lZ



I Q

I,,I


.0/



0



-.0/


-.02


U~t


-.0/


0 4 8
Ang/e av/AoocA, oc a69


CONFIDENTIAL


/2

NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS


figure /3. Varoion of elevon effect/veness with
ang/e of afiac/ for Zoarn y free -F//hft tunne/
mode/ of /lao//ess oil -wng aorpl/neQ with sweeptback.


Fig. 13








NACA ACR No. L5A13


CONFIDENTIAL


cZ























CO
u













ct


-O/ 1 1 1_
0 4
Ange cf aptck
NFIDENTIAL


.GS 0^9


/8

NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS


F)ure /4.- io//f/ng ord yaw/ng moments prodaceac
by Ydef/echng as e/evons fte p/aj/ rudders on
lhe l angy free-f7ght-uanne/ m cel fct /aJ///ea a//-
wing curpkone with sweepbtcck.


5rr
o---- -/0
-----/ZO
o -- 10
A-- 0


Fig. 14






NACA ACR No. LSA13


-ztL'^'it -

ptyw '4 /y/qd %/6uo xq%/


1a
to




ot






0 1 so
-i, >U


zc,


Fig. 15






NACA ACR No. L5A13


CONFIDENTIAL





C
.0/ 0.30







O

0 -2D 40 4. to
Rukder def/ecof n dr. deg

NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS
Figure /6.- >' wi/rg Poarnfets produced
by /ght sp/draofddar def/ec*oam on
the Langley ofrew -fh ghf-tunne/ mwode/
of w f~//ess a/l-wlng o/ro/efpI wi/h
sweepback. CONFIDENTIAL


Fig. 16





















































j




F'




UNIVERSITY OF FLORIDA
I III I Ill I 11111 111111
3 1262 08104 988 3




UNIVERSITY OF FLORIDA
DOCUMENTS DEPARTMENT
120 MPRSTON SCIENCE UBRARY
P.O. BOX 117011
GAINESVILLE, FL 32611-7011 USA