7~ MR No. L5F13a

/,IJ;' .~ \ .... ('.: . '. } 1'1;-3

NATIONAL ADVISORY COMMITTEE FOR 'u" "'J 1(" AERONAUTICS vi; /.. (~ ,'~l:;~j

C •. /

UNCLASSIFIED

GeI-QPIPlBPJ1XEA l ~ lfBR/{RY COpy

MEMORANDUM REPORT

for the

Bureau of Aeronautics, Nayy Department

IliVESTIGATION OF STABILITY AND CONTROL CHARACTERISTICS

OF MODELS OF THE CONSOLIDATED VULTEETAILLESS AIRPLANE

IN THE LANGLEY FREE-FLIGHT TUNNEL

III - DAMPING IN YAW AND PITCH

TED NO. NACA 2324

By Hubert M. Drake

Langley Memorial Aeronautical Labopatory Langley Field, Va.

;;,),.:ofg~:~\t§'P:H:c-,,'" , r -1 •• ;.:_~~ ,_, __ .. ,,,- ,,'

JUN ~3 1945 :; '. - .-- " c'_· .'....J .'

",- "\ :~;':~-"'- •. -"(~

G@}QF;L~' I ltd!"

Od /13J7:Jb UNCLASSIFJEO

-------- -------------------------------------------------------------------

-,.

MR No. L5F13a CONFIDE:NTIAL

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

"",

MEMORANDUM 'REPORT

for the

Bureau of Aeronau~ics, Navy Department

INVESTIGATION OF STABILITY' AND CONTROL CHARACTERISTICS

OF MODELS OF THE CONSOLIDATED'VULTEE TAILLESS AIRPLANE

'" . "

. . . IN THE LANGLEY FREE-FLIGHT TUNNEL

III. - DAMPING IN YAW AND PITCH

TED NO. NACA 2324

By Hubert M. Drake

SUMMARY " ....

At the request of the Bureau of Aeronautics, Navy Department, an investigation t6 d~termine the stability and control characteristi.cs .of, ·models of the original and revised dSs~gns ,of ,the Con~olidatSd Vu1tee tailless a'irp1ane has been conducted in, the Langley free-flight tunnel. The present report gives'the results of free­oscillation tests and calculations made to det~rmine the rotary damping in yaw and pitch for models of the original and revised designs for both the flaps-retracted and flaps-deflected configurations.

The results show that the values of the damping in pitch, Cmq' ranging from -1 to -3.75, and the damping in yaw, Cnr , ranging from -0.01 to -0.03, for the f1aps-retrac teo. conf.igura tion were abou.t one-third to one-tenth the 'values normally measured for conyentiona1 de signs. '. Def'lecting the f'laps 400 and extending the hori~ont~l surfa~es (flaps-deflected configu~ation) in mo'st c~ses almos,t doubled the damping in pitch and yaw of both designs~ The addition of the vertical tails caused an increase in Cnr of from about -0.01 to -0.015. The data indicated that the revised model in general had slightly greater damping in pitch and slightly smaller damping in yaw than the original design in both the flaps­retracted and flaps-deflected configurations.

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2 CONFIDENTIAL MR No. L5F13a

INTRODUCTION

An investigation to determine the stabllity and control characteristics of the Consolidated Vultee tailless airplane has been conducted in the Langley free-flight tunnel at the request of the Bureau of Aeronautics, Navy Department. The results of the flight and force tests of the original and revised models have been reported in references 1 and 2. The present report gives the results of free-oscillation tests made to determine values of the damping derivative Cnr , the rate of change of yawing-moment coefficient with yawing angular velocity, and Cmq' the rate of change of pitching-moment coefficient with pitching angular velocity. In addition, calculated values of Cnr and Cmq for the free-flight tunnel models and a hiGher scale model are presented.

Free-oscillation tests of the original and revised models were made for both the flaps-retracted and the flap-deflected configurations. The original design was tested with two vertical tail designs and one flap setting. The revised design was tested with one vertical tail and three flap settings. All tests were made with propellers removed. The extensions added to increase the control effectiveness for the flight tests were on for the oscil­lation tests.

COEFFICIENTS AND SYMBOLS

All coefficients, forces, and moments treated in this text are referred to the stability axes which is a system of axes with its origin at the center of gravity and in which the Z axi sis in the plane of syrmnetry and perpendicular to the relative wind, the X axis is in the plane of symmetry and perpendicular to the Z axis and the Y axis is perpendicular to the plane of symmetry ••. Positive directions of these axes are= X axis forward, Y axis toward the right, and Z axis downward. A sketch showing the forces and moments of the axes is c;iven as figure 1.

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MR No. L5F13a CONFIDENTIAL 3

Cn

h

m

M

L

The symbols used in the report are:

rate of change of

yawing angular

yaw~n~-mom~,ent o~oe~ffici ent

veloclty In o (rb 2V)

with

rate of change of pitching-moment coefficient

wi th pi tching angular velocity f- oCm \ . ~(qc/2V»)

rate of change of yawing-moment coefficient with

angle of sideslip (coC:). . yawing-moment coefficient ~ N ~

2:pv2Sb . 2

pi tching-moment coefficient f: M ~ 'lfpv2s")

lift coefficient (, L2 \ \:!pV S)

prof'ile-drag coef'f'lcient S D02 ~ \~PV S;

static margin, rate of change of pitching­moment coefficient with lift coefficient

(at trim) O~i)

slope of lift curve (d?i\ per radian au; yawing moment, foot-pounds

pitching moment, foot-pounds

lift, pounds

profile drag, pounds

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4

q

r

v

b

s

-C

y

w g

CONFIDENTIAL MR No. L5F13a

pitching angular velobity, radians per second

yawing angular velocity, radians per second

velocity 6f relative wind, feet per second

angle of sideslip, radians

.angLe·, of sweepback of one-quarter chord, degrees

aspect ratio (b2/S)

taper ratio (TiP chord) ,Root chord

~

wing span, feet

wing area, square feet

horizontal-tail area, square feet

vertical-tail area, square feet

mean aerodynamic chord, feet

tail length, distance from center of gravity to 0.25 chord of tail, feet

distance of aerodynamic center of wing behind aerodynamic center of root chord, percent of mean aerodynamic chord

constants obtained from l~eference 3

density of air, slugs per cubic foot

( vV ) relative density factor qpSb

weight of airplane, pounds

acceleration caused by gravity 32.2 feet per second2

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MR No. L5F13a C ONF I DENT IAL 5

APPARATUS AND MODELS

The investigation was conducted in the Langley free-flight tunnel on a strut which permitted only one degree of freedom (freedom in ya~or freedom in pitch). A description of the test apparatus is given in refer­ence 4. A photograph of the flaps-down configuration of the revised model mounted on the yaw-free stand is presented as figure 2.

The models were constructed by the Consolidated Vultee Corporation and were prepared for testing by the NACA. Four models were used in the tests: One each of the original design, flaps up and flaps down, and of the revised design, flaps up and flaps down. The models of the original design were constructed to 1/30 scale while those of the revised design were 1/35 scale. Complete descriptions of the models are given in references 1 and 2 and a table of the dimensional and mass characteristics of the airplane and of the models scaled up to full scale is presented as table I. Three-view sketches of the four models are shovm as figures 3 through 6.

TESTS'

The stand tests were made at a velocity of 37.5 feet per second for the original design and a velocity of 46 feet per second for the revised design corresponding to test Reynolds numbers of 109,000 and 103,000, respec­tively, based on the average chords. The effective Reynolds number was 174,500 for the original design and 165,000 for the revised design based on the tunnel turbu­lence factor of 1.6. The test procedure followed to determine Cnr was identical with that described in refer-ence 4. A similar procedure was used to obtain Cmq •

For all tests the axes of rotation passed through the design center-of-gravity location (2l percent of the mean aerodynamic chord for the original design and 19 . percent of the mean aerodynamic chord for the revised design). The controls were set neutral for these tests. The rates of oscillation employed in the tests were approximately 120 oscillations per minute for the Cnr tests and 300 oscillations per minute for the Cmq tests.

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6 , C01:rFIDENTIAT .... ·· MR No. L5F13a

A surirrnary of the conditions tested is given as table II.

CALCULATIONS

Damping in Pitch

The calculations of Crnq were based on an expres-sion derived in reference 3 whic.h defines Cmq as

( ..._/ ~ / v~ ~ C},\ -Cmq GI~d-- Z \ h + !!! h2

1 (1) o 2'

where

{l + (1 _. ,,-) 21, A2

,-'

A-~ .l mo tan A j 1 - A- lA tan qo -'3~ 2j D {, 1 + A. 2 \.~. , (1 + A.) ~ I I '- ..,;

and rrA Bl - (1 - A.). 1 J

Zo = ""2 1 + A. C 1 - Y Bl + 2A C 1 tan A ,

The additional damping caused by the extended hori­zontal tail was calculated by the formula

12 SH tlCmq = 2 11t c2 S CLa:t

T.he value of 'llt was assumed to be 0.80 for both designs based on the data of references 1 and 2.

Calculations were 'n-lade to determine the damping in pitch of the flaps up, 'horizontal tail extended and retracted configurations of both the original and revised designs. The calculations were made using a value of 5.16 for the section lift curve slope, mn" This value was obtained from the higher Re'ynolds number data of references 5 and 6. Inasmuch as the--value of mo, measured in the low-scale tests (references 1 and 2) , did not diff,er appreciably from the data at higher Heynolds numbers in references 5 and 6 the calculated values of Cmq would apply to either the free-flight tunnel models or full-scale configurations. ,"

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MR No. L5F13a " CONFIDENTIAL 7

sian

Damping in Yaw

The calculations of Cnr'were based on the expres­presented in reference 4:

Cnr ;"(:2 ~ IICnptan) - ~.33 (~ : ~R) CDo . ,

+ 0.020 0 ~ (\; 6)_ (12:t»CL~ (2)

Equation (2) shows that the calculated. value of. Cnr is dependent upon the value of CDo' the profile-drag coefficient. Inasmuch as this parameter is considerably affected by scale effect, ,separate calculations of Cnr were made for the free-flight tunnel model and the higher scale models of references 5 and 6. The values of CDo for the models were estimdted from references 1 and 2 whereas the CDo values of the higher scale models were,' obtained from the higher scale data of references 5 and'.6. Calculations of Cnr for the model s were made for the.' ... flaps-up configurations of the original and revised des'igns with tip tails off and on. :' .

RESULTS AND DISOUSSION

Rotational.Dampingin Pitch

, 'The results of the Cmq tests and calculations are . presented in figure 7 for the original model and in '. . figure 8 for the revised model. The results are. plot.ted against -Cmq because a negative value of Cmq means positive damping. The results show that the damping in pitch for the flaps-deflected models ,was greater then for the flaps-ret~acted models except in bhe case of the original model at lift coefficients·below O~7. (See fig. 7.) In figure 8 the eff~cts of flap deflection on C~ for the rev~sed model are shown to be slight while

-: ~

the extension of the horizontal tail contributed a rela­ti vely large· increase in damping •. The reduction in Cmq ' a t low lift coefficients caused by extending the flaps and,.:.·

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8 CONFIDENTIAL MR No. L5F13a

auxiliary surfaces of the original model might, there­fore, be attributed to interference effects of the forward auxiliary surfaces incorporated in the original design.

The revised design had greater damping than the original design in all configurations. This would be expected because changing the taper ratio from 6:1 to 4:1 and increasing the aspect ratio from 10 to 12 (for the same sweepback angle)' increases the wing-tip area. An increase' in the area of, the, wing tips .increase s the damping because of the relatively 'longer moment arm of the tip sections. The change in the center-of-gravity position between the two models also tended to increase the value of Cmq.

In general, the Cmq data did not show any consistent variation with lift coefficient although the early stalling (see tuft tests of reference 2) caused a reduction in Cmq for the revised model at high lift coefficients. The

calculated VruU6S of Cmq for flaps-retracted configu­rations are in fair agreement with the experimental data but the calculations for the horizontal tail extended configurations indicate values considerably greater than those measured.

The values of Cmq for the Consolidated Vultee tailless models were considerably smaller than those measured for conventional models. The values of Cm for the conventional model of reference 7 ranged fro~ -10 to -14 for normal tail sizes. The values of Cmq measured for the Consolidated Vultee designs ranged from about -1 to -3.75 or only about one-third to one-tenth the damping of this .conventional design.

Rotational Damping in yaw

The results of the tests and calculations made to determine values of Cnr for the Consolidated Vultee~, models are presented in figures 9 and 10 for the original model and figures 11 and 12 for the revised model. Beqause a negative value of Cnr means positi1~e damping the data are plotted against -Cnr •

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MR No. L5F13a CONFIDENTIAL

The data show that -the' values of -Cnr for the flaps-retracted model configuration with ,the vertical

9

tail off were approximately tripled by_putting the flaps down 400 and adding vertical-tails on the 'fuselage or on the wing tips. The data also show that the fuselage fin was less effective, flaps up, and more effective, flaps down, than the wing-tip fins in increasing the damping in yaw. The values of Cnr measured for the original model were, in general, slightly greater than those measured for the revised design.

The values of, Cnr calculated for the models are in fair agreement. wi'th the experimental data. The calcu­lated valu~s of. _,' Cnr for the higher scale models in the flaps-up configuration (figs. 9 and 11) are considerably smaller than the calculated free-flight tunnel model values because of the smaller profile drag of the higher scale models.

A comparison of the values of - Cnr measured for the Consolidated Vultee tailless models with the values of Cnr measured for the conventional model of reference 4 indicate that the damping in yaw for the Consolidated Vultee model is only about one-third to one-tenth as great as for conventional designs. The value of Cnr for a conventional model may vary from about -0.09 in the cruising condition to a value of -0.14 in the landing condition. -

Summary of Results

From the results of free-oscillation tests and calculations conducted on the models of the original and revised designs of the Consolidated Vultee tailless air­plane the following conclusions may be drawn:

1. The values o~ Cmq ranging from -1 to -3.75, and of Cnr' ranging from -0.01 to -0.03, for the flaps­retracted configurations were about one-third to one-tenth the values normally measured for conventional designs.

2. Deflecting the flaps and extending the horizontal surfaces approximately doubled the damping in pitch and yaw of both designs.

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10 CONFIDENTIAL MR No. L5F13a ........ J';

3. The iAdition of vertical tails caused an increase in Cnr of from -0.01 to -0.015.

4. The revised model had slightly greater damping in pitch and slightly less damping in yaw than the original design.

Langley Memorial Aeronautical Laboratory National Advisory Committee for Aeronautics

Langley Field, Va.

J~~vf ?1!. £J/~~~ Hubert M. Drake

Aeronautical Engineer

Approved: . .ftiAY7i'~ r:. /.f~.\../~ ~n- Hartley A. Soulel

Chief of ?tability Research Division

dgb

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MR No. L5F13a CONFIDENTIAL 11

REFERENCES

1. Maggin, Bernard: Investigation of Stability and Control Characteristics of Models of the Consoli­dated Vultee Tailless Airplane in the Langley Free­Flight Tunnel. I - Original Design - TED No. NACA 2324. NACA MR No. L5B03, Bur. Aero., 1945.

2. Paulson, John W. and Maggin, Bernard: Investigation of Stability and Control Che.racteristics of'Models of the Consolidated Vultee Tailless Airplane in the Langley Free-Flight Tunnel. II - Revised Design -TED No. NACA 2324. NACA rIm No. L5E07 a, Bur. Aero., 1945.

3. Glauert, H., and Gates, S. B.: The Characteristics of a Tapered and Twisted Wing 'with Sweep-Back. R.&M. No. 1226, British A.R.C., 1929.

4. Campbell, John P., and Mathews, Ward 0.: Determination of the Yawing Moment Due Contributed by the Wing, Fuselage, and of a Midwing Airplane Model. NACA ARR

Experimental to Yawing vertical Tail No. 3F28, 1943.

5. Garbell, M. A.: Report on Wind Tunnel Tests of 0.0639 Scale Model Two-Engine Tailless Design. Rep. No. ZT-021, Consolidated Vultee Aircraft Corp., Dec. 3, 1943. (Ref. GALCIT Test No. 425.)

6. Rogers, 1 ... 1.: Wind Tunnel Tests of a 0.058 Scale Model Two-Engine Tailless Design. Rep. No. ZT-029, Con­solidated Vultee Aircraft Corp., May 2, 1944.

7. Campbell, John P., and Paulson, John VI.: The Effects of Static Margin and Rotational Damping in Pitch on the Longitudinal Stability Characteristics of an Airplane as Determined by Tests of a Model in the NACA Free-Flight Tunnel.' lJACA ARR No. L4F02, 1944.

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CONFIDENTIAL MR No. L5F13a

TABLE I

DIMENSIONAL AND MASS CHARACTERISTICS OF CONSOLIDATED

VULTEE TAILLESS AIRPLANE AND SCALED-UP

CHARACTERISTICS OF MODELS T3STED IN

THE FREE-FLIGHT TUNNEL

Propeller:

Original design

Diameter, ft • • • . • . • .. .• 14.8 Number of blades:

Airplane - dual rotation . • • •. 6 Model - single rotation • . • . 4

Weight: , Airplane, Ib • • • • •• •••• 90,000 Simulated by model, lb ••..... 60,600

Wing (airplane and model) : Area, sq ft • • • • • . • • . • • • • 1800 Span, ft • • . . • •. • " 134 Aspect ratio . ... .. . 10 Washout, deg . ... • • . 1.5 Dihedral, deg . • . . • • . • . . 4 Sweepback, L.E. of wing, deg • • • •. 15 Taper ratio •. . . . . . • • • . 0.167 Average chord, ft . . . . • . . • . . 13.6 Mean aerodynamic chord

Length, ft ••••.••• .• 15.9 Location back of L.E. of root chord,

ft . . . . ...•.. \\ ... 6.90 Root chord, ft . • .• • •••.• 23.3 Tip chord, ft .•• . • . •• 3.9

Wing sections: Model

Root • • • . . Tip . . . NACA 103 with 150 reflex of

• RSG-35 with 100 reflex of

Wing loading: Airplane, Ib/sq ft •••.• Simulated by model, Ib/sq ft

. . . . . . 50 33

Revised design

15

6 4

90,000 80,000

1800 147

12 2.7

2 14

0.25 12.24

13.74

7.3 19.67 4.83

O.ZOc o .10c.

50 44

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CONFIDENTIAL MR No'. L5F13a

TABLE I - (Continued)

Original design

Center of gravity (airplane and Percent of M.A.C.

Flaps-up configuration • • Flaps-down configuration

Aileron (airplane and model):

model) :

• • • . . . . 21 19

Type •• • • • • • • • . • • • •• Plain Area, total

Square feet • • • • • • . • •• 110.0 Percent wing area • • • • • •• 6.02

Chord Percent wing chord

Outboard .• • • • • • • • • • • 16 Inboard . • • • • • • • • . • •• 17

Span Feet (each aileron) ••. Percent wing span (total)

Elevator (airplane Area

Square feet Percent wing

Chord

and model):

. . . . area . . chord

· .

· . . .

35.4 53.3

132 7.24

Percent wing Inboard . Outboard

. . . · . . . . 14 14 . . .

Span Feet (each elevator) •.• Percent wing span (total)

. . Aft auxiliary surface (horizontal tail):

(airplane and model): Total area, sq ft • • • • • • • . • • Percent wing area . • . • . . . . • . Span, ft .. . . . . . . . . . . . .

Vertical tails (airplane and model) :

26.4 39.4

312.4 17.35 60.5

Type • . • . . . . .. Two wing-tip fins Area (total of two), sq ft .••••. 180 Percent wing area • . • . • •. 10 Aspect ratio ••• .. . •.•.• 1.5 Rudders (total of two)

Area, sq ft • . . • • • • • Percent total tail area

. . 90 50

Revised design

18 19

Plain

98 5.45

15.2 15.2

37.4 50.9

156 8.7

13.7 13.7

35.3 48.1

335 18.6 46.1

same 216

12 1.5

108 50

CONFIDENTIAL NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

CONFIDENTIAL MR No. L5F13a

TABLE I - (Concluded)

Original design

Vertical tails (airplane and model) - Continued: Fuselage tail (model only)

Area . . . ,. . . . . •. .. ... 135 Percent wing area •. 0.075 Aspect ratio • • • • . • •. • • • 3 Rudder • • • . . • • • . . • • • • none

Forward auxiliary surfaces (airplane and model) : Area (total of two), sq ft ••••.• 145.0 Percent wing area .••••.••• 8.05

Ratio of radius of gyration to wing span:

Revised design

none

none

Original design

Revised design

kx/b Flaps up • • • • Flaps down · · ·

ky/b Flaps up · • • · Flaps down · • ·

kZ/b Flaps up · · · • Flaps down · · ·

Relative density factor II · • ·

• •

· · •

· ·

•

· • •

· •

·

Airplane Model Airplane Model

0.115 0.185 0.194 0.175 0.176

0.064 .0.083 .0.056 0.072 0.085 0.058 0.091

0.145 0.203 0.197 0.161 0.196

4.89 3.28 3.95 4.43

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

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"-

•• • • •• • •• • •

•• •• • •• • •••• •• • • •• • •• • •

• •••• • •••• • •••••• • •

Nodel

Original model

Revised model

Model

Original model

Revised model

Or (deg)

00

400

00

0°

200

40°

Or (deg)

00

0°

0°

400

400

40°

00

00

0°

0°

200

200

400

40°

MR No. L5G13a

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TABLE II

MODEL CONFIGURATIONS TESTED

Horizontal 8urraces

Retracted

Extended

Retracted

Extended

----do----

----do----

Horizontal surfaces

Retracted

----do----

----do----

Extended

----do----

----do----

Retracted

----do----Extended

----do--------do----

----do--------do----

----do----

Cmq Tests

Vertical CL . ; tail

Tip 0.03, 0.5, 0.7 tfIl

---do--- 0.2, 0.5, o.a. 0.95, 1.4

Tip 0.05, 0.25, o.~,· 0.73

---do--- 0.16, 0.39, 0.59

---do--- 0.36, 0.72, 1.1

---do--- 0.55, O.BB, 1.2

Cnr Tests

Vertical tail

orr Tip tail

Fuselage tail

ort

Tip

Fuselage

Oft

Tip

Off

TIp

ort

Tip

otr

Tip

CL

0.05, 0.3, 0.52, 0.7, 0.92

0.05, 0.3, 0.52, 0.7, 0.92

0.05, 0.3, 0.52

0.5, 0.75, 0.9B, 1.23, 1.4

0.1, 0.75, 1.23

0.1. 0.75, 1.23

0.05, 0.25, 0.52, 0.71

O.OS, 0.25, 0.52, 0.71

.0.06, 0.16, 0.3B, 0.5B

"0.06, 0.16, 0.3B, 0.5B

-0.05, 0.30, 0.69, 1.03, 1.22

.0.05, 0.30, 0.68, 1.03 ). ,.

0.2, 0.53, 0.B5, 1.2, 1.29

0.2, 0.53, 0.B5, 1.2

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

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Figure

7

7

a a a B

-Figure

9 ,

9

9

10

10

10

11

12

11

12

11

12

11

12

"

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L

CONFIDENTIAL /';,;'t<re I . -6ysfem or .:Jfc:7b/i/ty c:7X'e;s. Arrows /lJdtCc:7te />os/fl

ve c/;recflon or moments Cll7d IOrces.

. NATIONAL ADY/SORY COMMITTEE fOil AERONAUTICS

• • •• • • • • • •••• • • ••• •

" •• •• • • • •

• • • ••• • • • • • • • • • •••• -, •••

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.. •• III • • •••

: II • .....

• II

• to •• .. . .

1I ••• • III •

It " • •••• • •

Figure 2.- Photograph of model of revised Consolid&ted-Vultee tailless airplane mounted on the yaw-free stand in the Langley free-flight tunnel. Flaps down (40 0 ), rear horizontal surface extended, wing tip fins and slats installed, propellers removed. CONFIDENTIAL IA1'IOliL ADVIIOILT CO»lI1'1'1I roa ARae.AutrOI

""!lout! IIlIIIOIlUL AIIIlOUUfIULLiBOUfOIl'l - LUOLU rIllllD, Vi.

•• e •

GO • •

• •••

•••• •• • • • • •

•• •• • •• • •••• •• • • •• • •• . :. .

• •••• • •••• • •••••• • •

, .

j . control surface extensIOns

Tip Sla.t J -used In free­Flight tunnel lest6.

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,2310% span

MR No. L5F13a

2690' 5pon l 17% chord -- ------ ---

J----------------5372·'----------------1

Cb:20.28N~

~ Tail 1 _,= __ _ " -+ - "--- \

- --- - - :::"';:- - NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

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Wing Area span Aspect Ratio M.AC. Root ChOrd Tip Chord Vertical-Tail Area Center of Gravity

2.00 -sq. ft. 5} 72 in.

10 6-37 In. 933 in. 1.55 in. 0./05 0.21 HAC

Figure 3.- Three- view drawing of lJo-scale model of the Con-solldated Valtee tailless airplane. Flaps-up configuration

• •••• • •••••

•• •• • •• • •••• •• • • •• • •• • •

• •••• • •••• • •••••• • •

MR No. L5F13a

A CONFIDENTIAL

. ,. I 13.75% chord ~ elevator extensionz

Aft auxiliary surFace ---- ---- -

-1- _ '5:. ~o~e~o~ e:~~/on -~

t I

~ -t

1j. f

8-B

1-0,'----------------53.720.------------------1

Dime sian

a h C

FLAP Inboorc1 Outboard

"ToC %C 70.0 80.0 25.0 20.0 3.3 4.0

Flap Detail

Full-span balanced .split Flap

~----2028.~"-----~

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NATIONAL ADVISORY COMMITT[£ rOi AEROIIAUTICS

Wing Area span

2.00 sq. fl. 53.72 in.

Aspect Ratio MAC. Root Chord Tip Cnord Vertical-Tail Area Center of Gravity

10 6.]7 in. 933 in. 1·55 in. 0.10 S 0.21 MAC

Figure 4. - Three-vi~w drawing of -130 -scale model of the Consolidated Vultee tailless airplane. Flops-down configuratiOn.

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End of Document