research memorandum - nasaresearch memorandum fliget measurementsof directionalst4bilityto a mach...
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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
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CIASSIFISDDOCUMENT
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NATIONAL ADVISORY COMMITTEEFOR AERONAUTICS
WASHINGTON
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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
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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,
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Slug-ftz
Slug-ftz
Slug-ftz
NACA RM H55G26 CONFIDENTIAL 3
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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
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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
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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.
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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.
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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.
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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
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APPENDIX A.
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● ODERIVATION OF Cn EQUATION
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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
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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
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● 9 ● REFERENCES● 0 ●
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● ● 1. Williams, W. C., and Crossfield, A. S.:● ☛:00
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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.
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NACA w H55G26 CO~WIDENTIAL
TABLE I
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. .wing:Atrfoil section . . . . . . . . .Total area (including aileron and
NACA 6hAO07. . .Sq ft
. . . .covered
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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:
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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....
.
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.
.
5. 24.6
94.1
. 20
. 15reference line), percent . . . . . . . . . . . . .
. . . . . . . . . . . .
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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
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. ;
. 25
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Sw;ep at 0.25 chord line, deg .Dihedral, deg . . . . . . . . .Travel, leading edge up, deg .Travel, leading edge down, degIrreversible hydraulic boost and artificial
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!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 , ,
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1.- Three-view drawing of airplane withdimensions in inches.
vertical tail A. All
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● 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