RESEARCH MEMORANDUM

FLIGET MEASUREMENTS OF DIRECTIONAL ST4BILITY TO A MACH

NUMBEROF 1.48 FOR AN AIRPLANE TESTED WITH THREE

DIFFERENT VERTICAL TAIL CONFIGURATIONS

By Hubert M. Drake, Thomas W. Finch, and Jsmes R. Peele

(

High-Speed Flight Stat~,on

.~~=~, .a~~~~~f~:k~~ ~,fj&

‘m _.....{$&#=&

J“: ~.:Ju“’l.NmuWIYw —“.”.-==-------@ -.. J.’

LIBRARY MWA “ Ilmls

CIASSIFISDDOCUMENT

Thfs materialcontafm informationaffectingthe NationalD9fenseof tiw UnitedStateswithin& meardngof the espionage laws, Title 18, U.S.C., WCS,’793and’794,Uietramsmlssionor revelationof wkdchfn any—r to anunautlwrbxsdperson is prohibitedby law.

NATIONAL ADVISORY COMMITTEEFOR AERONAUTICS

WASHINGTON

CON~ENTIAL

● ☛☛☛

● ● 0● * ●

“00 ●

● ● W

● ● *

●

● ☛☛☛☛☛

w ●

● ● ☛

9*O● ** ●

NACA w Ii551G26 CONFIDENTIAL

NATIONAL ADVISORY COMMITTEE

LIBRARYNACA -HSFR

FOR AERONAUTICS

RESEARCH MEMORANDUM

FLIGHT MEASUREMENTS OF DIRECTIONAL STABILITY TO A MACH

NUMBEROF 1.48 FOR AN AIRPLANE TESTED WITH THREE

DIFFERENT VERTICAL TAIL CONFIGURATION ;

By Hubert M. Drake, Thomas W. Finch, and James R. Peele

suMMARY

Flight tests have been performed to measure the directional sta-bility of a fighter-type airplane over the Mach number range from 0.72to 1.48. The tests were made at altitudes of 40,000 feet and 30,000 feetand employed three different vertical tails of varying aspect ratio orarea, or both.

These tests showed that the directional stability for all tail con-figurations increased with an increase in tail aspect ratio or area, orboth, over the entire Mach number range and decreased with increasingsupersonic Mach number above 1.15.

\

INTRODUCTION

The decrease in direction:.1 stability with increasing supersonicMach number during flight tests of research airplanes has been discussedpreviously in references 1 to 3. In general this decrease results from

the fact that with increasing supersonic Mach number the lift-curve slopeof the vertical tail decreases while the unstable directional moment ofthe fuselage remains essentially constant. If the directional stabilityof the airplane becomes sufficiently low, deterioration of the dynamic

stability can result (refs. 1 to 3) and, if the directional stability

becomes zero, an actual divergence can occ~.

With the advent of fighter airplanes capable of appreciable super-sonic Mach numbers, the problem of adequate directional stability at

supersonic speed has become of immediate importance. A fighter airplane

having a 45° swept wing and supersonic performance capabilities is beingutilized for flight research by the NACA High-Speed Flight Station atEdwards, Calif. During rate-of-roll tests this fighter airplane exhib-

ited violent cross-coupling behavior as reported in reference 4. Low

● *O.● ● 0

.* 9

b. ●

● ● *● ● o

●

● **9**● ●

● ● O● 0.● 00 ●

2 CONFIDENTIAL NACA RM H55G26

directional stability was considered to have contributed to the violenceof the behavior. An investigation was undertaken, therefore, to deter-

mine in flight the directional stability of the airplane. During the

investigation two additional vertical tail configurations with increasedarea and aspect ratio became available and were included in the tests.

This paper gives the results of the measurements of directional

stability for these three tails within the Mach number range from 0.72to about 1.48.

A

b

CNA

Cn

c%

IY

IZ

SYMBOLS

aspect ratio

wing span, ft

Rolling momentrolling-moment coefficient, qSb

airplane effective dihedral parameter,~, deg-l@

airplane normal-force coefficient, ~qs

Yawing momentyawing-moment coefficient, qSb

acn -1airplane directional stability parameter, ~’ ‘eg

chord, ft

mean aerodynamic

acceleration due

altitude, ft

chord, ft

to gravity, ft/sec2

moment of inertia about X-axis,

moment of inertia about Y-axis,

moment of inertia about Z-axis,

CONFIDENTIAL

Slug-ftz

Slug-ftz

Slug-ftz

NACA RM H55G26 CONFIDENTIAL 3

● 0 ●

● ● *

● . ●

09=● ● *.

● ● *

●

● 00000● ●

● ● 0● 00● OO ●

1=

it

M

n

P

P

~

q

r

s

T1/2

t

w

a

B

5at

br

‘c/4

A

product of inertia, slug-ft2

angle of tail incidence measured from line parallel to longi-tudinal axis of airplane, deg

Mach number

load factor, g units

period of lateral oscillation, sec

rolling angular velocity, radians/see

dynamic pressure, lb/sq ft

pitching angular velocity, radians/see

yawing angular velocity, radians/see

wing area, ft2

time to damp to half amplitude of lateral oscillation, sec

time, sec

weight, lb

indicated angle of attack, deg

indicated angle of sideslip, deg

total aileron deflection, deg

rudder deflection, deg

sweepback at 0.25 chord, deg

taper ratio

AIRPLANE AND INSTRUMENTATION

The airplane utilized in this investigation is a fighter type with

a single turbojet engine and a low swept wing and tail. A three-view

drawing of the airplane with the original vertical tail is shown in

CONFIDENTIAL

● 000● ● *

● 0 ●

O* ●

● ● 0● ● O

●

● 00000● ●

● ● 0● 00● 00 ●

4 CONFIDENTIAL

figure 1. Figure 2 presents a photograph ofand mass characteristics of the airplane are

The tests utilized three vertical tailsareas and aspect ratios as follows:

NACA RM H55G26

the airplane. The geometricgiven in table 1.

characterized by differing

I I Area, sq ft I Aspect ratio

Tail A 33.5 1.13Tail B 37.3 1.49

Tail C 42.7 1.49

Drawings of the three tails are shown in figure 3. Figure 4 pre-

sents a photograph of two airplanes showing tails A and C. The same

rudder was used with all tails.

Complete stability and control instrumentation was installed for theflight research reported in this paper. The angle of attack, angle of

sideslip, airspeed, and altitude were sensed on the nose boom. The Mach

numbers presented are based on a preliminary calibration of the airspeedinstallation and are considered accurate to ~0.02 at subsonic speeds andto _kO.01at supersonic speeds. The angle of attack and angle of sideslip

are presented as measured.

TESTS AND DATA REDUCTION

Rudder pulses were performed to determine the period and damping ofthe lateral motions. The maneuvers were performed by abruptly deflecting

the rudder pedals, returning them to neutral, then holding them fixed.The stick was held fixed throughout the maneuver. Representative maneu-vers are presented in time history form in figure 5. Maneuvers were per-

formed in level flight as shown in the following tabulation, except forthe maneuvers at Mach numbers greater than 1.35 which were performed indives at altitudes above 33,000 feet.

M Altitude, ft Vertical tail

0.72 tO 0.74 30,000 A, B, end C

0.84 to 1.34 40,(2QO A

o.78to 1.39 40,000 B

0.72 to 1.48 40,000 c

CON3?IDENTIAL

● ☛☛☛● ✘

9BOO**O

● 0

NACA RM H55G26 CONFIDENTIAL

It was found that the simplified methodof the directional stability parameter Cn

P

of determining the valuegiven in reference 5 was

inadequate for this airplane. Therefore, the following expression was

used (see appendix A for derivation):

5

It may be noted that this expression includes the single-degree-of-freedom relation of reference 5 modified to include the effects ofdamping, product of inertia, and an angle-of-attack term. The expres-

sion gives the value of Cn as measured about the body axis. TheP

product-of-inertia term is very small since the principal axis is esti-mated by the manufacturer to be inclined only about 1/2° down at thenose for an average test weight. Unpublished wind-tunnel measurementsof C2B for Mach numbers up to 1.0 were used in the equation. Since

above M = 1.0 the angles of attack were generally. less than 3° and

C2B is quite small, the angle-of-attack term is considered small enough

to-be neglected at supersonic speeds.

RESULTS AND DISCUSSION

The time histories of maneuvers at M = 0.74 and at M = 1.38 in

figure 5 were performed with tail B but are representative of all threetails. In most cases the pedals and stick were held fixed subsequent tothe pulse; however, small control surface movements did occur.

Figure 6 shows the measured period and damping for the three tailconfigurations. These data are presented for an altitude of 40,000 feetwith the exception of the points above M = 1.35 which, as mentioned

previously, were obtained in dives between 40,000 feet and 33~000 feet”For any given tail there is very little scatter in the periods measured,indicating the small control movements did not unduly influence theperiod. The measured damping, however, shows considerable scatter. Thiscondition reflects the difficulty of measuring the damping and possiblythe effect of the small control motions. For the most part, the periods

show a general decrease with an increase of Mach number at subsonicspeeds and show an almost constant value at supersonic speeds. With each

of the three tail configurations, the measured variation of damping of

the airplane indicates a region of reduced stability (increased time todamp) at transonic speeds with less time required to damp to half ampli-tude at Mach numbers above and below this region.

CONFIDENTIAL

6 CONFIDENTIAL NACA RM H5i5G26

● ☛☛✎

● ● 0● 0 ●

●

● 0,0,0● ●

● 9*● **● 00 ●

The values of the directional stability parameter Cn were com-

puted from the ‘~aluesof P and T1 */ in figure 6 P. Figure 7 presents

the Mach number variation of Cn determined for each of the tail con-P

figurations. The value of C‘B

increases with increasing tail size as

would be expected from the increase in aspect ratio or area, or both.

The decrease in C%

anticipated with increasing supersonic Mach

number (ref. 2) is quite pronounced particularly above M = 1.15. Withtail C, for example, the airplane lost about half its directional sta-bility between M = 1.15 and M = 1.48. A measure of the improvementin Cn~ (producedby increasing tail area and aspect ratio) is shown

by the fact that a value of CnB

= 0.001 was reached at M . 1.22 with .

tail A, but with tail C the value of Cn = 0.001 was not reached untilP

about M = 1.48.

An unpublished value of CnP

= 0.(Xl19measured in the Langley 4- by

4-foot supersonic pressure tunnel at M = 1.41 is shown in figure 7.When this value is corrected for aeroelastic effects estimated by themanufacturer

(Nn = -0.00052) and the effect of turning the air flow

P

(%at the intake duct & = -o.mo25), a value Of Cn

Pnearly equivalent

to that measured in flight is obtai~ed.

The required value of Cn for reasonable handling qualities is,B

of course, not indicated in these tests but must be determined by theperformance of other maneuvers such as the aileron rolls presented inreference 4. That Cn

Pmay become very small or neutral at high Mach

numbers, however, indicates that under some conditions poor dynamic sta-bility

to 3),

may be encountered,or an actual static

as for the Douglas D-558-II airplane (refs. 1directional divergence may occur.

CONCLUDING REMARKS

Flight tests of a fighter-type airplane with three different verti-cal tail config’uations over the Mach number range from 0.72 to 1.48indicate, as would be expected, that the directional stability increasedwith an increase ip tail aspect ratio or area, or both, over the entireMach number range and decreased with increasing supersonic Mach numberfor all tail configurations.

CONFIDENTIAL

● 000● ● 0

● 0 ●

● 0 ●

● ● *● ● 0

●

● 00000● ●

NACA RM H55G26 CONFIDENTIAL 7

The Mach number at which a value of the directional stability param-eter Cn had decreased to 0.001 was increased from a Mach number of 1.22

Pto a Mach number of 1.48 by increasing the tail area 27 percent andincreasing the tail aspect ratio 52 percent.

● ● ☛

● 00● *O ●

High-Speed Flight Station,National Advisory Committee for Aeronautics,

Edwards, Calif., JulY 20, 1955-

Hubert M DrakeAeronautical Research Scientist

Thomas W Finch

Aeronautical Research Scientist

James R. Peele

Aeronautical Research Scientist

Approved:

k

C.>”>--yHartley A. Soule

Res rch Airplane Projects Leader

mld

CONFIDENTIAL

8

● *O*● ● m

● 0 ●

CONFIDENTIAL NACA R14H55G26

APPENDIX A.

● m .● ● O

● ● m

●

● ☛☛☛☛☛

● ODERIVATION OF Cn EQUATION

P● 9*● **● 00 ●

The three lateral equations of m~)tl”r containing

to the derivation of Cn are writtep. ~:B

,

&

IXZ . qSb2‘b Cn$ + ~qSb2c r+~— p + 2VIZ‘=Iz ‘r Iz

terms pertinent

Cnij

Taking the time derivative of the @ equation assuming a is

Substituting and collecting terms gives

( )(qsb2cnr j + ~ %P - ~ cZ@ +.. qSb2c. +— IqSb aqSb

P- % Cyp .—2VIZ ‘~ 2VIZ

It is assumed that r . -~ and the damping factor, ~, is equal to

CONFIDENTIAL

● **O● ● *

● 0 ●

● ☛ ●

● ● 0● 9*

●

● s0000● ●

● ● 0● 00● 00 ●

NACA RM H55G26 CONFIDENTIAL 9

The natural frequency of the oscillation o is determined from theimaginary part of the roots.

/(2 4& lXZ qsb ~iL- J@J’q +——

‘P Ix P Iz Ix 2P)

2=0

Squaring and simplifying

IXZC +GU21Z Izc =— —Cx -—‘P qSb ‘aIx ~ Ix ‘~ hqSb

()2

()

2Inasmuch as up = ~ and ~2=4_ , substitution and

rearrangement of terms gives

Cn =B

where

c =nr

/dcn d Q

2V

Cn. = dcnp$B

‘Cy = Side forceqs

CYB .dC@

mass

= &p/at

= dr/dt

velocity

= dppt

—— d2p/dt2

10 CONFIDENTIAL

● ee -● ● *

● 9 ● REFERENCES● 0 ●

● ● 0● ● 0

●

✎☛✎ ☛✎☛

● ● 1. Williams, W. C., and Crossfield, A. S.:● ☛:00

● ** ●

Speed Airplanes. NACA FYIL52A08, 19>2.

NACA RM H55G26

Handling ~ualities of High-

2. Ankenbruck, Herman O., and Dahlen, Theodore E.: Some Measurements

of Flying Qualities of a Douglas D-558-II Research Airplane DuringFlights to Supersonic Speeds. NACA RM L55A06, 1955.

5. Ankenbruck, Herman O., and Wolowicz, Chester H.: ~teral MotionsEncountered ~~iththe Douglas D-558-II All-Rocket Research Airplane

During Exploratory Flights to a Mach Number of 2.0. NACA RM H>4127,

1954.

4. NACA High-Speed Flight Station: Flight Experience With Two High-Speed

Airplanes Having Violent Lateral-Longitudinal Coupling in AileronRolls . NACA RM H55A15 , 1955.

5. Bishop, Robert C., and Lomax, ~r~ard: A Simplified Method for Deter-mining From Flight Data the Rate of Change of Yawing-Moment Coeffi-cient With Sideslip. NACA TN 1076, 1946.

CONFIDENTIAL

NACA w H55G26 CO~WIDENTIAL

TABLE I

11

● ☛☛☛

● ● 0

● O ●

● ☛ ●

● ● ☛

● ● 0

●

● ✎☛✎✎✎

● ● PHYSICAL CHARACTERISTICSOF AIRTLANE● ● O

● 00

● 00 ●

. .wing:Atrfoil section . . . . . . . . .Total area (including aileron and

NACA 6hAO07. . .Sq ft

. . . .covered

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

376.0236.5811.3315.864.76o.~o3.5645000

by fuselage), sq ft . . . . .Span,ft . . . . . . . . . . .Mean aerodynamic chord, ft . .Root chord, ft . . . . . . . .Tip chord, ft . . . . . . . . .Taper ratio . . . . . . . . . .Aspect ratio . . . . . . . . .Sweep at 0.25 chord line, deg .Incidence, deg . . . . . . . .Dihedral, deg . . . . . . . . .Geometric twist, deg . . . . .Aileron:

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

19.327.8125~15

Area ~>earwardof hinge line (each), Sq ft ..

.

Span at hinge line (each), ft . . . . . .Chord rearward of hinge line, percent wing chordTravel (each), deg .-. . . ~ . . . . . . . . . .

Irreversible hydraulic boost and artificial feelAerodynamic balance . . . . . . . . . . . . . . . . None.

Stati; balance . . . . . . . . . . . . . . . . . Internal lead weightsLeading-edge slat:Span, equivalent, ft . . . . . . . . . . . . . . . . .Segments . . . . . . . . . . . . . . . . . . . . . . .Spanwise location, inboard end, percent wing SeITIiSpan .

Spanwise location, outboard end, percent wing semispanRatio of slat chord to wing chord (parallel to fuselage

. 12.71....

.

.

.

.

.

5. 24.6

94.1

. 20

. 15reference line), percent . . . . . . . . . . . . .

. . . . . . . . . . . .

.

.

.

.

.

.

.

.

.

.

.

Rotation, maximum, deg .

Horizontal tail:Airfoil section . . . . . .Total area (including 31.65

Sqft. . . . . . . . . .Span,ft . . . . . . . . .Mean aerodynamic chord, ftRoot chord, ft . . . . . .Tip chord, ft . . . . . . .Taper ratio . . . . . . . .Aspect ratio . . . . . . .

. . .

. . .Sq ft

NACA 65AO03.5. . . . . . . . . . .covered by fuselage),

. 98.86

. 18.72

. 5.83

. 8.14

. 2.46

. 0.30

. 3.54

. 45

.

. ;

. 25

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

. .

. .

. .

. .

. .

. .

. .

. .

. .

. .

. .

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

. .

. .

. .

. .

. .

. .

. .

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

Sw;ep at 0.25 chord line, deg .Dihedral, deg . . . . . . . . .Travel, leading edge up, deg .Travel, leading edge down, degIrreversible hydraulic boost and artificial

CONFIDENTIAL

● ☛ ●

● ● ☛

● ● ☛

●**9*:● ●

● ● ☛

✎ ✎ ✎

● 00 ●

12 COIJFIDENTIAL

!L”L-lL. I.- L.nc.uie,d

PRYSICAL CHARACTERISTICS OF AIRPLAIJF

NACA RM H55G26

Vertical tai~: AAirfoi: section . . . . . . . . . . IiAcA65A~5.5Area (excluding dorsal fin and area blanketedby fuseiage)jsq ft... . . . . . . . . . . . . . 33.5

Area blanketed by fbselage (area between fuseiagecontour line and line paral:el to fuselagereference line through intersections of leadingedge of vertica. tai: and fuseiage conto,n-line) . . 2.11

Span (u.nblanketed),ft... . . . . . . . . . . . . . 6.14Mean aerodynamic chord, ft . . . . . . . . . . . . 5.83Root chord, ft....... . . . . . . . . . . . . . 7.75Tipchord, ft...... . . . . . . . . . . . . . . 3.3?Taper ratio . . . . . . . . . . . . . . . . . . . ,. o.b28Aspect ratio . . . . . . . . . . . . . . . . . . . . 1.15Sweep at 0.25 chcrd line, deg . . . . . . . . . . . 45Rudder:Area, rearwerd of hinge line, sq ft . . . . . . . LJ.3Span athingeline, ft.. . . . . . . . . . . . . . 3.35Root cilord,ft....... . . . . . . . . . . . . 2.27Tipchord, ft...... . . . . . . . . . . . . . 1.50T’ravel,deg . . . . . . . . . . . . . . . . . . . . ~~(J

Sparrwiselocation, inboard end, percentvertical tailspin.. . . . . . . . . . . . . . . 4.5

SpanWise location, outboard end, percentvertical tailspin.. . . . . . . . . . . . . . . 58.2

Chord, percent vertical tail chord . . . . . . . . . 50.0Aerodynsmricbalance.. . . . . . . . . . . . . . . 0v5rban,-ing,

,417>,.2Jc>,:

BNACA 6>ACQ3 .5

2.117.455.517.752.32

0.3011.49

45

6.53.332.271.50

3.7

cNACA 65AC05 .5

42.7

2.45

7.?35.908.282.49

0.3011.49

45

6.53.332.271.50

*2O

3.1

44.828.4

Overban:in~,u.sealed

Fuselage:Length (afterburner nozzle closed), ft . . . . . . . . . . . . . . . . . . . . . . . . . . 4>.64Maximunwidth,ft. . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.58Maximum depth over canopy, ft . . . . . . . . . . . .

. . . . . . . .

Side area (total), sq ft.. . . . . . . . . . . . ...1 l:::””::””::::: 123~n~Fineness ratio (afterburner nozzle closed) . . . . . . . . . . . . j n . . 1 1 . . . . . . 7.86

Speed brake:SUrfacearea, sq ft. . . . . . . . . . . . . . . . .Maximmm deflection) deg.. . . . . . . . . . . ...1 :::::::::””””:: ;:

14.14. . . . 50

Power plant :Turbo jet engine . . . . . . . . . . . . . . . . Omc Pratt & ‘XhitneyJ57-P7 with afterbmnerThrust (guarantee sea level), afterbwner, lb . j : . . . . . . . . . . . . . . . .Military, lb . . . . . . . . . . . . . . . .

. . 15, (XO

Norrml,l b . . . . . . . . . . . . . . . . . . . . . . “:.::””<””””””” ““9,220

. . . . . . . . . . . . 8,cx30

Airplane weight, lb:Basic (without fuel, cil, water, pilot) . . . . . . 19,662Total (full fuel, oil, water, pilot ) . . . . . . . . . :~,c::~:jlj ::::::24,~o

Center-of-gravity location, percent ?:Total weight -gear down . . . . . . . . . . . . . . . . . . . . . .Tctalweight -gear up.. . . . . . . . . . . . . . . ..,....::;:::: ;:::

jl.m31.8C

llo~?ntsof inertia (estimated total weight) :

~x~s~u+-ft2........ . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . li, lo3

Iy , Slug-fop . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . >9,246

IZ, slug-ft2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .67,2’75

IxZ, slw-ft2. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 941

Inclination of principal axis (estimated total weight) :

B’:O,.’referencea xisatnose, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.6

-m.--m T-,. -r-. T , ,

CONFIDENTIAL

●

● ☛

9*

●

.000●

● ☛●

:0 r 557

—Q----- — cL-——$— —/

. Ccd

I I--9

— 1

\ \ \

I AII

H

1.- Three-view drawing of airplane withdimensions in inches.

vertical tail A. All

COHFIDENTXAL

● 0 ● :

● ● 0

● *9*● 0

● * ●

● O *● *

● ● *

●

● ***O

● ●

● ☛

● ☛

● 9 ●

CO1?FIDEXTIAL

.%. .. mcT\

.

2’+J

%o

I

m“

CONFIDENTIAL

● 9 ● *

900● 0

● boe* ● 9

● * ●

● 0 ●b ● 0

● 9*

CO1?FIDEI?TIAL

Tail A

/

Tail C

7

/

/

t

/

zA’r~cI blanketed by fuSeh9e(tOil

c/4 for tolls A and B

c/4 for toil C

Tall At/4 J A IA Area,

1Span, Blanketed area,

cieg 1 Sq !-t ft Sq ft(1) (2)

I1

Ii 45° 1 1.15 1 0.42R I ~~.~ 6.14 2.111

B 450 ~ 1.149 i 0.501 57.5 : 7.45 2.11

c I 45° ! 1.49 ~~ ; ‘“X1

42.7 7.93 2.451I

(1) Area not blanketed by f’uselage

(2) Span not blanketed by fusel. ge

Figure 3.- Sketch of vertical tails A, B, and C.

CONFIDENTIAL

..

-..

●✛

●●

☛;0;

●●

●0

●:0

:00

so:

/

J“.

t

Figure

4.-

Photograph

L-89

376

of

two

airplanes

showing

tails

Aand

C.

/

COi?FIDEWIWIL

8009*

● e

● 0● *

9*

● ☛

*O

90

upRight

Iv4,r,radians/see

upRight

a,~,deg

.4

.<P.+

/-. ●- y~ * - —.4

\\/

.4

8

‘- ‘L”;’\‘L-tiy ‘---- ----- ‘-–- ‘--– ‘-

4/-fl

0 — \/

—

4

Ahdonenose up4

Right————. .————

o>&-r

4 !

%~at,it, dq

8 v

120 2 4 6 8 10 12 14 16

Time,t,sec

(a) M = 0.74; hp = 30,000 feet.

Figure 5.- Time history of lateral oscillation induced by a rudder pulsefor airplane with vertical tail (B).

COID?IDENTM

● ● s● 00

● 00 ●

up ‘4Right

f-rP

P, q, r, radim\sec o - Q-’ ~-~ <“/? ‘ – -z_. --- ––Lp

● 0 ●

● ● .● ● e

●

● ☛☛☛☛☛

● ●

up 4

Right ——

+

——— /-a.———

p-.

——— —

(

I

4

Airplane nose up8

Right

E

/it–

4:

—

/sa+——— _ —---—

$-,~a+,it, deg— ——— — ——. ——

F!iii!0

4.

/y ‘

80 2 8 10b

~me, t, sec

(b) M s 1.38; ~ s 37, OOOfeet.

Figure

CONFIDENTIAL

5.- Concluded.

4

p,se

c

2 0

.,(

33

0 0

Pri

med

sym

bols

‘P=

W,0

00fe

et

1:2

1.5

I

●☛

☛●

●☛

●●✚

●●

●“e

●

●*

●●

O*

●0

●●

*:0

0●

●●

*●*

●:0:e

o●0●

(a)Vertical

tail

A.

Figure

6.-

Mach

number

variation

of

period

and

time

to

damp

to

half-

amplitude

of

lateral

oscillation

for

altitudes

near

40,000

feet.

i-+

;

●☛

☛● ●

☛

● ●

●☛

●✚● ● ●

●☛

☛

●●

o ●

●

● ●*O

9*● :0

0●

*●

P,

sec

c-l

T1/

2’

6 4

~‘a

oQ

%@

~~

. .’2

Prim

edsy

mbo

ls

‘P=

30,0

00fe

eto 6- 4

-

sec 2

-

0~ ●.I

.U

.YI.U

I.1

I.2

1.3

1,4

1,5

IM

(b)

Vertical

tail

B.

Figure

6.-

Continued.

(

T1/

2,se

c

( .

+-@

-Q

--—

i..—

I

%--

(

0’

+I

&

0(

L213

—I

M

(c)

Vertical

tailC.

Fi&e

6.-

Concluded.

●*O

●

●:●°

0●

●

●*

r●

●0

●●

**●

*●

e●

●*

●0

●●

●0

●O

.:0:**

●0

●

5%--

o-

;

CONFIDENTIAL

● O**● ● *

● 0 ●

● * ●

● ● *● ● *

●

● *****● ●

● 0:00● 00 ●

Q.

k’-P

in

3

*

-ua)

—.-0

)

+

c’”’

-7.

0

R,.

!

8

0ri

(((

(

CONFIDENTIAL

NACA RM H55G26

● 000● ● *

● * ●

● ☛ ●

● ● ☛

● ● 0

CONFID~TI&,.. .> ’”?i . .

INDEX

● Subject9.*9*9

● ●

● ● 0● O* Airplanes - Specific Types●00 ●

Stability, Directional - StaticStability, Lateral and Directional - DynamicFlying Qualities

ABSTRACT

Number

1.7.1.21.8.1.1.31.8.1.2.2

1.8.5

Directional stability characteristics have been determined from the

measured period and damping of a fighter-tYPe airplane over the Machnumber range from 0.72 to 1.48 at altitudes

of 40,000 feet and 30,000 feet.

Three different vertical tails of varying aspect ratio or area, or both,were employed.

,.

CONFIDENTIAL