Effect of wing modifications on the longitudinal stability of a tailless all-wing airplane model

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Material Information

Title:
Effect of wing modifications on the longitudinal stability of a tailless all-wing airplane model
Alternate Title:
NACA wartime reports
Physical Description:
8, 14 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 -- Wings -- Testing   ( lcsh )
Aerodynamics -- Research   ( lcsh )
Genre:
federal government publication   ( marcgt )
bibliography   ( marcgt )
technical report   ( marcgt )
non-fiction   ( marcgt )

Notes

Summary:
Summary: An investigation of the power-off longitudinal stability characteristics of a tailless all-wing airplane model with various wing modifications has been made in the Langley free-flight tunnel. Force and tuft tests were made on the model in the original condition, with the wing tips rotated for washout, with rectangular and swept-forward tips, and with various slat arrangements. Flight tests were made with the original wing and with the original wing equipped with the most promising modifications. The results indicated that changes in tip plan form or rotation of the wing tips did not appreciably reduce the instability at high lift coefficients. Addition of wing slats, however, improved the longitudinal stability at the stall when the slat extended far enough inboard to cover the area that tended to stall first.
Bibliography:
Includes bibliographic references (p. 8).
Statement of Responsibility:
by Charles L. Seacord, Jr., and Herman O. Ankenbruck.
General Note:
"Report no. L-42."
General Note:
"Originally issued September 1945 as Advance Confidential Report L5G23."
General Note:
"Report date September 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 - 003616165
oclc - 71300599
sobekcm - AA00006275_00001
System ID:
AA00006275:00001

Full Text
riCA L-q2


NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS





WA TIME RE PORT
ORIGINALLY ISSUED
September 1945 as
Advance Confidential Report L5G23

EFFECT OF WING MODIFICATIONS ON THE LONGITUDINAL STABILITY
OF A TAILLESS ALL-WING AIRPLANE MODEL
By Charles L. Seacord., Jr. and Herman 0. Ankenbruck

Langley Memorial Aeronautical Laboratory
Langley Field, Va.


;WASHINGTON
N: ';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-
Sviously 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.

', Lr4. DOCUMENTS DEPARTMc .-


V





































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NACA ACR No. L5G23 CONFIDENTIAL

NATIONAL ADVISORY COMM ITTEE FOR AEROiAUTIC3


ADVANCE CONFIDENTIAL REPORT

EFFECT OF WVTIIG MODIFICATIONS OH THE LOHGITUDIITAL STABILITY

OF A TAILLESS ALL-WING AIRPLANE MODEL

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





An investigation of the power-off longitudinal sta-
bility characteristics of a tailless all-wing airplane
model with various wing modifications has been made in
the Langley free-flight tunnel. Force and tuft tests
were made on the model in the original condition, with
the wing tips rotated for washout, with rectangular and
swept-forward tips, and with various slat arrangements.
Flight tests were made with the original wing and with
the original wing equipped with the most prom-ising modi-
fications.

The results indicated that changes in tip plan form
or rotation of the wing tips did not appreciably reduce
the instability at high lift coefficients. Addition of
wing slats, however, improved the longitudinal stability
at the stall when the slat extended far enough inboard to
cover the area that tended to stall first.


INTRODUCTION


Sweepback is often incorporated in the design of
tailless airplanes in order that high-lift flaps may be
used on the center sections of the wing to increase the
over-all maximum trim lift coefficient of the airplane.
(See reference 1.) Quite often, however, the sweepback
defeats its own purpose by causing premature tip stalling
and longitudinal instability at high angles of attack and
thus making it impossible for the airplane to attain its
maximum lift coefficient in flight. A model of a tail-
less all-wing airplane with sweepback and taper recently
tested in the Langley free-flight tunnel (reference 2)
showed this tendency. The maximum trim lift coefficient


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NACA ACR No. L5G25


of this model with flaps retracted, as measured by force
tests, was about 1.2; but because of the poor stability
and control near the stall the highest lift coefficient
at which it could be flown was 0.7.

In an attempt to improve the longitudinal stability
characteristics of swept-back all-wing tailless airplanes,
an investigation of various means of preventing tip stall
has been made in the Langley free-flight tunnel. The
model used in the tests of reference 2 was also used for
the present investigation. The test program included
force and tuft tests, power off, of the original wing,
of the original wing with the wing tips rotated for wash-
out, of the wing with modified rectangular and swept-
forward wing tips, and of the original wing with four
slat arrangements. Flight tests were made with the origi-
nal wing and with the original wing equipped with the
most promising modifications.

SYMBOLS


L lift, pounds

M pitching moment, foot-pounds

N yawing moment, pounds

CL lift coefficient Lift
\IqS

Cm pitching-moment coefficient about 0.20c"
Pitching moment.
qES

CD drag coefficient (2r)

S wing area, square feet

R Reynolds number

c wing chord, feet

c mean aerodynamic chord, feet


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b wing semisspan, feet

V airspeed, feet per second

a angle of attack degrees

angle of yaw, degrees

p angle of sideAlio, degrees

$ angle of bank, degrees

6 rotati on of wing tip, degrees

6e ele:on deflection, degrees

6r rudiitzr def 1ectinn, degrees

q dyn:r,.ic pressure, pounds per square foot (ipv2)


P mass density of air, slugs per cubic foot


APPARATUS


The tests were conducted in the Langley free-flight
tunnel, '.].2.-'I1 is ( escrtibed 1.n reference 5. A photograph
of the to.-: steinn of the tunnel showing the model in
flight is pt sented as figure 1.

Force 'tests made to determine the static stability
characteri ?ci cs of the model were made on the Langley free-
flight tunnel six-component balance. (For a description
of the b....'ce see refcrencc 4.) All forces and moments
meaZsure'-.: on this balance are talken with respect to the
stability ayzs, which are shown in figure 2.

The mod?l is the one that was used in the tests
reported inl reference 2. The model is of a tailless all-
wing airplane having an aspect ratio of 7.'6, a taper
ratio (r:t:io of tip chord to root chord) of 0.25, and
swepbl:.ac cf the quarter-chord line of 220. A three-view
drawing %.f t;he moJel is presented as figure 5; and plan-
view and three-quarter front-view photographs are pre-
sented. as figures 4 and 5, respectively. For the present


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NACA ACR No. L5G25


tests, the wing tips were cut at the outboard end of the
elevons and altered so that the angle of incidence of the
tips could be changed or the tips removed entirely. (See
fig. 6.) Two additional sets of tips extending 28 percent
of the wing semispan one rectangular and one with 50
sweepforward of the quarter-chord line were built to
fit the wing where the original tips were cut. Leading-
edge slats for the outer part of the span were constructed
in three sections, any of which could be attached to the
wing separately.

The model tested in the Langley free-flight tunnel
(designated FFT) at low Reynolds numbers was, in its
original condition, identical in plan form to a model
that was tested in the Langley 19-foot pressure tunnel
(designated 19-ft PT) at high Reynolds numbers; data for
the tests in the Langley 19-foot pressure tunnel are
given in the present paper for comparison with the result
of tests in the Langley free-flight tunnel. The two
models, however, differed in airfoil section and number
of propeller-shaft housings. The model tested in the
Langley 19-foot pressure tunnel had an NACA 65(518)-019
airfoil section at the root and NACA 65(318)-015 section
at the tip; and the model tested in the Langley free-
flight tunnel had a modified !ACA 103 airfoil section
with a thickness of 21 percent chord at the root and
15 percent chord at the tip. The aerodynamic washout for
both models was approximately 4.


TESTS


Force tests were made to determine th- static sta-
bility characteristics of the model in each of the test
conditions. The force-test data for each arrangement
were based on the area and the mean aerodynamic chord of
the particular wing plan form tested.

Tuft studies were made of each model configuration
to determine the stalling characteristics of the wing.
For these tests., the model was mounted on the balance
strut.

Force and tuft tests were run at a dynamic pressure
of 4.09 pounds per square foot, which corresponds to a
test Reynolds number of about 240,000 based on a mean


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MACA ACR HoI. L5'-,25 C


aerodynamic chord of 0.655 foot. All force and tuft
tests were made with flaps retracted, vertical fins off,
and the elevon and rudder control surfaces set at 0.

Flight tests were made with combinations of slats 1
and 2 and 1, 2, and 5. (See fig. 6.) These tests were
made with the center of gravity at 20 percent of the mean
aerodynamic chord and over a range of lift coefficients
from 0.5 to 1.0. All flight tests were made with flaps
retracted and with vertical fins installed. These fins
were added to improve the directional stability; previous
tests have indicated that they had no effect on longitu-
dinal stability.

All tests were made with power off and propellers
removed.


RESULTS AND DISCUSSION


In interpreting the results of the tests made in
the Langley free-flight tunnel, the following points
should be considered:

(1) The tests were made at very low Reynolds numbers
(150,000 to 350,000).

(2) The controls of the model during the flight
tests were fixed except during control applications;
hence, no indication of the effect of the modifications
on the control-free stability of the design was obtained.

Results of the force tests are shown in figure 7.
In figure 8, the curves of pitching-moment coefficient
against lift coefficient are replotted to compare the
stability characteristics for the various wing modifica-
tions. Data for the model of similar plan form tested
at high Reyrnolds numbers in the Langley 19-foot pressure
tunnel are also shown in figures 7 and 8. Results of
tuft surveys in the Langley free-flight and 19-foot pres-
sure tunnels are presented in figure 9.


Original Wing

The force-test data of figure 3(a) and the tuft-test
data of figure q(a) illustrate the usual effect of


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NACA ACR No. L5G23


sweepback and taper upon the static longitudinal stability
and the stalling characteristics of a wing. These data
show that the premature stalling over the elevons near
the wing tip caused the original wing to become neutrally
stable at a lift coefficient of about 0.90.

The data of figures 7 and 8(h) also show that at
high lift coefficients the longitudinal instability of
the free-flight-tunnel model, tested at low Reynolds num-
bers, was greater than that of the pressure-tunnel model,
tested at high Reynolds numbers. The results of the
force tests made in the Langley free-flight tunnel are
believed to be conservative in that the necessary improve-
ment in longitudinal stability at high Reynolds numbers
is less than the improvement indicated by the tests at
low Reynolds numbers. It is interesting to note, however,
that in contrast to the dissimilarity of the pitching-
moment curves for the free-flight-tunnel model and the
pressure-tunnel model (fig. 8(h)), the stalling charac-
teristics as indicated by tuft surveys are quite similar
for the two models (figs. 9(a) and 9(1)). When flown,
the model showed a tendency to nose-up and stall after
disturbances in pitch at a lift coefficient of about 0.65,
and it was not possible to fly the model at lift coeffi-
cients above 0.7. (See reference 2.)


Effect of Wing-Tip Modifications

A comparison of the curves in figure 8(a) shows that
rotating the wing tips -100 had little effect on the
longitudinal stability and did not prevent instability
at the stall. The tuft-survey results in figures 9(a)
and 9(b) show that, although the stalling of the tip was
improved slightly by deflecting the tip, the stall inboard
of the tip was relatively unaffected. Correlation of
these results with force-test results indicates that an
improvement of the stall over the elevons as well as over
the tips is necessary to eliminate the longitudinal insta-
bility at high angles of attack.

Force tests of the swept-forward and rectangular
tips (figs. 8(b) and 8(c)) showed no improvement in the
stalling characteristics.


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NACA ACR fo. L5G23


Effect of Slats

Addition of various slat arrangements caused definite
improvements in the longitudinal stability characteristics
at the higher lift coefficients. This effect is shown in
figures H'(d) to 8(g). The tuft tests showed that at high
angles of attack the slat arrangements cleared the stalling
on the tip in approximately direct proportion to the span
of the slats, and the premature stalling over the elevons
was improved only by the slat arrangements that extended
in front of th.i elevons. (Sec fig,. 9(e) to 9(h).) The
slight roughness and stalling within the span of combina-
tions of slats 1 and 2 and 1, 2, and 3 is attributed to
slat supports, which are located between the individual
slats.

With the 50.5-percent-se.nispan slat and the 70.5-
percent-semisoan slat installed, the model could be flown
to a maximum lift coefficient of 1.0 an increase of 0.5
over the maximum lift coefficient with the original wing -
and did not show the nosing-up tendency noted in flight
tests of the original wing.


COIHC JUSJIONS


The following conclusions were drawn from tests of
a tailless all-wing airplane riodel with various wing
modifications in the Langley free-flight tunnel:

1. Changes in wing-tip plan form over the outer
23 percent of the wing semispan caused no appreciable
improvement in longitudinal stability at the stall.

2. Decreasing the angle of incidence of the wing tip
(28 percent of the wing semiispan) by 100 had little effect
on the longitudinal stability and did not prevent longi-
tudinal instability at the stall.

5. The use of partial-span wing slats eliminated the
longitudinal inst.--bility at the stall when the slat span


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NACA ACR No. L5G23


was great enough to extend inboard in front of the part
of the wing that tended to stall first.


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




iPEFEREIrCES

1. Pitkin, Marvin, and Maggin, Bernard: Analysis of Fac-
tors Affecting Net Lift Increment Attainable with
Trailing-Edge Split Flaps on Tailless Airplanes.
NACA ARR No. L41l8, 1941.

2. Campbell, John P., and Seacord, Charles L., Jr.:
Determination of the Stability and Control Charac-
teristics of a Tailless All-Wing Airplane Model
with Sweepback in the Langley Free-Flight Tunnel.
NACA ACR No. L5A13, 19L5.

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

4.. 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
No. 35D17, 1943.


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NACA ACR No. L5G23





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NACA ACR No. L5G23




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NACA ACR No. L5G23


Figure 9.- Results of tuft surveys of tailless airplane model In
Langley free-flight tunnel. 8. = 8r = P = 00; q = 4.09 pounds
per square foot. NATIONAL ADVISORY


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Fig. 9a,b


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NACA ACR No. L5G23


Figure 9 .- Continued.
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NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS


Fig. 9c, d







Fig. 9e, f


NACA ACR No. L5G23


Figure 9.- Continued.
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NACA ACR No. L5G23


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Figure 9 .- Continued.

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NATIONAL ADVISORY
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Fig. 9g,h






NACA ACR No. L5G23


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(a) Original wing, FFT.


Figure 9 .- Concluded.
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(1) Original wing, 19-ft PT.
R 6.6 x 106.


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS


Fig. 9a,i











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