Flight studies of the horizontal-tail loads experienced by a modern pursuit airplane in abrupt maneuvers

1
:. I .


1k L-13


NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS





WARITTIME RElPORT
ORIGINALLY ISSUED
June 1944 as
Advance Restricted Report L4F05

FLIGHT STUDIES OF THE HORIZONTAL-TAIL LOALS
EXPERIENCED BY A MODERN PURSUIT AIRPLANE
IN ABRUPT MANEUVERS
By Flight Research Maneuvers Section

Langley Memorial Aeronautical Laboratory
Langley Field, Va.








4NACA "


WASHINGTON
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-
nically edited. All have been reproduced without change In order to expedite general distribution.


L 93


DOCUMENTS DEPARTMENT


/


ARR No. L4FO5




































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/flightstudiesofh001ang




-1 1 2 2 Z t )


:rACA -.RR No. Irk05

i:ATTONAL ADVISORY C3Ia:ITTEE FOR AEFO3..UTICS


ADVANCE rATR IC TD RTEPCRT


FLT',IT Sr"DIES OF '7-Th HO' Z'.:TKL,-TAIL LOADS

EXPEiEPLcD d." A .W:LER' FURS IT AT-.FLA2
7Y") AAT.TT P


Dy fi.-ht -,esea:.ch ;.:r uv:rs '"ti.on





rli-t.t mea93re:p .ts were mnidE 8 "-,Drn -'ursit
Pir-l ;r.e to f t r iie fc -- .-. r:.r :: s t tud- f t'
hori -rntal te' i los., ri ? : e-le; t-: I fli t. 9.i the -se
fli ght rT~a,-i '.re Trnts, rss .rcEs a ll f : Fts '.v rE u ec
us an inr -=- of t.. t I.: '" );r ati.r ticze ores-
su ires vi t'. -or ete e ?r..re -dJ tr i t io ta bt ai i -d
:r. t-e 'ACA fui-s3 l tnrel, Ir. a- t; tlon, 3 tr in ages
ar.nd *ti )n I '.-tres f t l: deleccti n r v. ? U.sePd to
exc lore t_ "I er-, nn:.t.ire .-rd order lf ta. .it .te of the
flu 't-. ati n, tail 1. s.r in :c2elersLted zta;l .

The ;-ecv.lts inl.: z'teu. ttrt, if t a- rrlane were
not t: 1i o, t ttal ur loai of 700 can 's wo d 'e
exper-i nc r on t -. n.or! zontal t-l in;! an Cr pul.l-ip and
that, with .o:er or,, I.is icad v .ld e r lis tri .ted
uns'--etricsl ll,' with about- C -.ur.J more up loa1id n
the left stastllzer than :.r thc ric-:t. hen stalling.
ocearrred there wr.s qr. ir.ti 3] arr.upit inrease in the up
tail load of the order of 1,.- ,:-:rcerit .f the 1rev;ios1 3
loae, ;i :h wva3 follow.dc b:y e:'~ tr:d clad and stress
vari ai.ons due to tail buffietin Under the ond.tions
of thil 1 ,uffeting, the o. si -ility of eFxe -"sJ iv stresses
due t'-n rs.onaino. was in._~_i ?ted.


I!TROD"CTTO


As a resCult' of nu-eroa.s t.i'l fail'jurs of -.':r.ern
hi.h-sree' -irr-lenes i fl ffl ,ht ir.vest..-ati r
was :1.1tn ertal-'ren t- rtr t. -r.r' n- th-e c:,n-" al rn.ture .of hoiri-
zontol t1il l1oas ex -l ri.-nc ii! cr ipt .t11-il-, Prane_.vers.








NACA ARR No. L4F05


Tests were made by the NACA at Tangley Field, Va. during
the spring and summer of 1911.2. The flight-test procedure
involved the use of pressure measurements made at a few
points on the horizontal tail, which were correlated with
co-olete oressure-distribution data from the NACA full-
scale tunnel to determine the approximate tail loads.
This procedure gave satisfactory results except when
applied to stalls wherein abnormally high fluctuating
pressures, corresponding to tail buffeting, were
experienced. In order to help establish the significance
of the r.le pressures recorded, a strain gage c--able of
following the load fluctuations was installed on the
stabilizer; motion-picture cameras were installed later
to record the deflection of the horizontal-tail surfaces.

The results of the tail-load measurements obtained
are discussed in two main parts. Ci-' part pertains to
the more or less steady loads experienced in maneuvers,
for bhich the determination of loads by means of the
measured pressures is fairly straightforward. The second
part deals with the fluctuating loads experienced in
stalled flight wherein the significance of the measured
pressures was difficult to establish. For this second
case, the main dependence is placed on strain measure-
ments and photo'-r.a'.rs of the tail deflections.


DESCRIPTION OF AIRPLA- A'-.'D APPARATUS


Test airplane.- The tail-load tests were made on a
modern pursuit airplane having the plan form and dimen-
sions shown in figure 1. The gross weight of the airplane
was maintained between 11,900 pounds and 12,000 rou:-u: for
the tests. The center-of-gravity position was maintained
between 29.8 percent and 30.2 percentt mean aercdrn'.n:.c
chord.

[asic flight instruments.- Airspeed, elevator angle,
stick force, and normal accelerations were recorded during
the tests by standard NACA recoifli, instruments. The
airspeed recorder was connected to an iijCA swivelln.-
static head located 1 chord length ahead of the right
wing tip and to a shielded total hecd mounted on the
airspeed boom.









Ni.,CA ARR To. IJ.;F05


Pressiure-dt.strihution _ir st lati -n.- Tchvr a:ir s of
Grifi-s. z wcre installed on t.e 1-.cri-o-Ital Ltabl lizer to
z,,easl.ue the rjresaure iift~"tce te Let .ee.n th uopr nd
lower su-,r-faces of the stabilizer. "1. soanmise and chcr.i-
wize ?locations of t'ie orifices .e-re chczer to correscocnd
v'it h articularr orifices il.-e in t.,e pr=sJ'.u're-d. str-hution
r!eastLL u'rJ3nts made in t. '.i.Ci f.ull-.scrle tunnel. S
sl:e;c, shiov'ing the location of the ;r-fices used in the
flight t=-ts i C livn, in f'i':- ~ Pi a. s-iisues '"re
recorded for the individual orifices by an IAC.I -:echanieal
r:Iano_.eter mountedd in the baEa :c comrnsrtrient of the air-
plane. The inboard orifrices ,-er-e connected to high-
frei. .,. n..:- pressure- r-ccrders to rer- it a study of the
pressL-,ue 'luctuationz at the stall.

T n"il-deflorctic.n r.'anp.ratus.- 7-The eflections of the
hor!;ic t alt tail uniJr lo'l ,ir- :r.e. ai.ured b rhotoCraphing
the tail with t'-o 16o-milli-iet--' rot ion-picture camer-:s
moun*,f-,, one on each side rf t'e fu-el.g?, in tl.e inter-
ccoler exir ducts. 'he .ca.eras ere synchronized by
ti.min: li :!-ts oncrated by a :-.aster tli-.r that also
syn lhronized all the recordc":in Irnst,'.-ents in the air-
plane. r-cr et, 'er*: ctinted' on he tiil i :ane to
identify the snanw! '-: o t it on .n t .e :-h tc~ra-,hic
records. 'he cavern nst :lltiorn ni the tar-ets on the
hori:.or.tal tail are s.l.on by ph;ctc-r- [hs in figures (a)
and ,(b), respectiv-ely.

Strain-ra.c installation.- An electrical strain Eage
,as installed on the skin above the:0 rear spar on the
ri ht horilcntal stabilizer. .!.Z hotograoh sho-.iing the
location of the strair. gaTe and the dumny gaze or the
hori-ontal tail is Eiven in figure The orifices on
the uher surface of the tail I the lead from th
crifices on the lo'.er surface :rc also sho,-r in i inure 4.

For one flight, de P'orest scr.at'h-type strain .cages
were mounted along the front soar on the upper s::in of
the left stabilizer at 64, 60 and 74.5 inches from the
stabilizer tip. The pa-es vere rounted by gluing the
gage target and scratch arm to the skin.








4 NACA ArR No. L4F05

T:SP PROCI:~T' E


The types of tests and records obtained are sumia-
rized in the following table:


Records obtained

Flighti P of Basic Pressure Strain Tail
maneuver flight distri- rage deflection
bution

llbR Abrupt pull-ups Yes Yes No .To
15` jAbrupt pull-ups Yes Yes 1o To
138 jAbrupt pull-ups! Yes Yes Y os No
193 13800 turns Yes Yes i Yes No
21) iAbrupt pull-ups '
Sand l800turn Yes Yes Yes Yes
2413 'Abrupt pull-ups Yes, Yes Yes Yes


It is apparent frcr. the table that the test program
progressed from an installation that measured only pres-
sures en the horizontal tail to one consisting' of a
combination of pressure orifices and a strain gage and,
finally, to an installation which simultaneously measured
the pressure, strain, and tail deflection. Thv'. strain
gsge was installed to facilitate an inter-.retation of
the pressure fluctuations experienced on the horizontal
tail at and beyond maximum lift of the ving in the pull-
ups. The apr:pratus for measuring tail deflection was
subsequently added in an effort to c tainn additional data
on the motion of the tail following the wing stall for
correlation with the pressure fluctuations and the strain
measure e-ents.

The abrupt pull-ups to maximum lift were made at
various speeds, from the minimuzl spf,.ed of the airplane
to an indicated airspeed of app-roz.xi tely 21L miles per
hour. The correspo'-L,!ng normal accelerations e:--perienced
rc-.-:.d from lg to 4.5g. All tests were made at an alti-
tude of approximately 6000 feet and, except for one
power-off run, with the engine operating at 2450 rpm and
27 inches of merc.r:,* manifold pressure.








rJACA ARR TTo. LLT-'05


DETERi,'lIj'.TIOUi OF TAIL LORDS


The pressure data recorded in flight wrcre converted
to tail loads from the oire sur::e--aistritution date for
t;ie tail plane obtained in them U.CA full-,ale tunnel.
Eecauce of an unsy.nmetrical flow in lhe ful -scale-tunnel
tests, the load on the tail, as indicated Ly integration
of the Tme'-isured pressu-zres, as unsyii'.trical. The
dissymmretry of Icad is shown in fi.:ure'- 5, which is a
plot of the spanwise of5srib uion of load on the hori-
zontal tail. The v r.able 1 us-eu. in lcis figure is
the product of the r-t.tion crial-for'ce coefficient cn
and the loc-l chord c.

The normal-force coefficie-ts Cyr for each half of
the tail were plotted in figure as a function of the
pressure coefficient p/t', in vhich Ap is the dif-
ference beLween the pressures on the3 luper and lower
surfaces of th: tail plane at Che twr spanwise stations
wharc orifices vrwre lccated in the Pli.ht-test installa-
tion and c is the dy;na..l.c R'Zf5u2e. The tail loads
computed 'rorr pressures meFaecurd at the individual
orifices tlerafore assume a sy.'nm trial tail locd with a
load di stributior similar to t.at ob.-inde in the full-
scale-tuiinnel tests. The normal-fcr'ee co f'ficients for
the tail are noted to be proportional to the pressure
difference across the ta il plane and are also a function
of.tlLe elevator angle 6e. T.The tunnel data for the right
inboard orifice were considered too inconsistent for use
in evaluating the tail loads (see fiE. 6) and the evalua-
tion of tail loads for the fli.l:ht tests was therefore
based on measurements at the other three stations.

Tail loads were determined froi. the tail-deflection
data by mcans of the influence line rhown in figure 7
and the spanwise load distribution of figure 5. The
influence line was obtained experimentally by applying
unit up loads at the indicated spanwise points, whereas
the spanwise load distribution was t al.en from NACA full-
scale-t'unnel data. The tall load per inch stabilizer
deflection is obtained by the summation


b/2
Z yw a,








TACOA ARR No. L2Fo5


in which w is the running load at a s,.an,.i.se point,
y is the ordinate of the influence line at the same
point, and b is the span of the horizontal tail. This
su.rm-iation chows a load of 875 pounds per inch tip deflec-
tion on the right stabilizer and 976 pounds per inch tip
deflection on the left stabilizer.

Some question may be raised as to how the spanwise
load distribution (fig. 5) should be faired across the
fuselage, but consideration of possible changes would
not materially alter the loads as measured by tip deflec-
tion.


RESULTS A~D DISCUSSION


Loads in installed flight.- The tail loads in
accelerated flight were measured in pull-ups to maximum
lift of the wing. Time histories of airspeed, normal
acceleration, elevator position, and elevator stick force
for three typical pull-ups of varying acceleration are
presented in figure 8. The present discussion is limited
to the loads attained before the wing stalled, that is,
to the portion of the maneuver prior to tail buffeting,
as is indicated by the fluctuating normal-acceleration
curve.

The pressure coefficients Ap/q for the four span-
wise points are listed in table I. The corresponding
values of normal-force coefficient CN obtained by
reference to figure 6 are also listed for the three
stations at which satisfactory calibrations were available.
Total tail loads correspon-rin-i to the normal-force coef-
ficients of table I (tail load equals 55ql;) have been
plotted in figure 9 as a function of normal acceleration.
Lxtrapclatinic these data indicates that an up load of
about 5700 pounds would be experienced at an acceleration
of 8g.

In consideration of these tail loads, a study was
made to learn the contribution to the load of each of
the following factors:

(a) Increment of tail lIc:d necessary to balance
pitching moment of winr-fuselage-
propeller combination









1:ACA APR 'Io. 1JW05 7


(b) Increment of tail load due to hor .zontal
location of center of ;.raviit7 v'ith respect
to aerodynamic. cenze-r o'f ving-fuselac-e-
propeller combination

(c) Increment of tall lo:.d due to manipulsticn cf
elevat Dr.

At the speeds investigated, the increment of tail
load due to factcr fa) (a do'n load) vas found tc .1
relatively. small, about 5.4q or 5 pou-nds at 200 miles
per hour. At diving speeds, l.o1..ever, this incoirementL is
large enough to be of primary cr,'seiderat ior.

The increrient of tail load due to factor (b) is
alia--s an up load at positive lifts v.ith the conventional
wingv a.n tail .rrangjenent: if the aeodynas.'ic center of
the vi- nrg-fuselage- -ropeller c'ombinati. n is kn-,'n,
detlerrining this increment of tail load for an.- center-
of-_ravity position, gros s w3iht, and normal accelera-
tion resolves into a s.i,.rle .io..,ent problem. The increment
of Lil load vanlries i'e-etly as the productt of the grcss
weictt and nori:al ac *- l:rat ion .nd vnri.es linearly w':ith
center-of-rravlty lot- C'on: th't is, this increment of
ta-L loa., v-ill be zero for ,veryr flight condition if the
center c,- gravity and -arodynnaiic ccnt;r are coincident
and will increase as the center of ura-vity moves rerwv;ard.

Full-Ccale-tunrel test ndicate that the aerodynamic
center of the fuselaCe-'wing-propell r combination (power
on) of the airplane tested is at. anroximartely 15 percent
of the mean q-,rodynamic chord. titii this aerodnaiilc
center, th-e in-re-!ents of tail loa-1 calculated by the
method sufggested are in subst-ntial aprcerrient v:ith tail
loads obtained front flichit-tepst s.1aL. The ta.l loads
experienced during acceleration w,'re considerably larEer
than the loads indicated by standard idr dei n rac ti e
because the propeller and fuselse caused the aerodynamic
center to move farther forv.ard than had b.--er anticipated


A discussion of the effect on the tail loads of
factor (c) (elevator manipulation) requires a knowledge
of the control movement during the maneuver. It is
apparent from figure 8 that the elevator force is relaxed
before the maximum acceleration is reached and as a
result the stick force is approximately zero at the time
of maximum acceleration. Wnen the elevator stick force









FACA ARR No. LFP05


is zero, the elevator is floating, and the tail-load
increment due to a combination of factors (b) and (c) is
equal to that obtained in a similar maneuver, elevator
fixed, with the center of gravity at the point giving
zero stick-free stability. Computed on this basis, the
up tail load due to releasing the elevator is 135 pounds
per S of normal acceleration. Extrapolation of the data
in figure 10, which is discussed subsequently, corrobo-
rates ex'!:rimentally this calculated load increment.
This load increment is indicated by the difference
between the curves shown for elevator floating and
elevator fixed as determined from installed pull-ups
and stead; turns, respectively.

Pull-ups to maximum lift .;d installed pull-ups to
the same acceleration gave dissimilar tail-loading condi-
tions. Analysis of the data indicates that the load was
unequally distributed between the right and left stabi-
lizers during installed pull-uos, as shown in figure 10.
The total tail load, however, was the same as that
obtained in pull-ups to maxim on lift. (Compare 4-.5
pull-ups in figs. 9 and 10.) A clue to the probable
cause of the asymmetric load is obtained by a study of
the time histories of figures 11 and 12. A turn ;.ith
power cn is shown in figure 11. Immediately before this
turn was entered, the load on the left stabilizer was
greater than that on the right stabilizer and remained
greater by about the same amount throughout the turn.
-h e Dressure changes that occurred durir.- the turn were
very similar on both sides of the tail and occurred
simultaneously with acceleration chan -es. For the turn
of figure 11, which was executed with power off, the
loads were nearly equal on both stabilizers, with the
pressure orifices indicating a slightly larger tail load
on the right stabilizer. ''he .-'canes in pressure during
this turn were similar to the c-:.:es that occurred in
the p)o.er-on turn. Consideration of the magnitude of
the dissymmetry in load'ing indicates that the un--:ym-nt-
rical tail loading is attributable to a slipstream twist
which increases the angle of attack on the left stabi-
lizer 20 or 30 in a positive direction and decreases the
angle of attack on the right stabilizer by an equal
amount.

It ay:.-ars from these "ata that the slipstream twist
with power on is responsible for an a3-~rteetr'c tail-load
increment except at maxin.um lift. (See fig. 9.) The
dissyret, ,w which is inder~jdent'. of s'.ced and acceleration,









IACA A.rR lro. LiL.05


results in an un load on the left stabilizer 800 pounds
greater thln tnat on the ri-nt st a! ilizer. This
unsy~acetrical loading, if -ttained in an accelerated
pull-up of &'g, would result in a tail load of 250 pounds
on the, left half of the ta'.l or in a stress due to an
equivalent uniform tail load of 6b35 pounds.

,~.'is during stalldi fliht.- In scr-upt lull,-ups to
maxir-u lift, lar5 and e:.ra t t.ai -lc--d ircre ;m nts
were ir.diiated by sharp )rfs.s'Ere raises immediately after
th- stall occurred. T-e irtLal I.: 3.- pressures were
follow, ed by fluctuating p- esurc- tthroueg'-h;ut the period
of stall C flight, r"i- i-.itories of pull-ucs to maximum
lift ('fi r 15 and 'i show the r..3t ..2r of chese pressure
rises and fluctustio. x,, to;ethcn. wit'c si,;.ultaneous
re or dIs of strain a indies Led Ly L I::- electrical strain
fgae. These. abrupt pressure rises and flu.ctua-ions are
asci e. to fluctuations irn direction of t-e air flow at
the tr.il, vlizh are due tc stailin-. of the ving.

a-s w .s previously .r;2nticned, dar:eras vere installed
to r-ccrd the ..otion ci u'i- h-rizooital tall during pull-
ups. T'.:e pccur-ac of measurenrA.t of lead!ng-ed"e
deflr-ctioriF on t e 16-.illirr.C-r fl. :: is believed to be
within-. J.COC15 inch, ,hih is *,uj.ivalent to 0.1 inch of
actual tail deflectior. .thi.;-h a c-era speed of
apprcxi.,.ately .: fraiels p ,r seccnd v- s used, che frequency
of the tail vibrations .*wis such that the3 aximurm aImpli-
tude of the motion of the tail v.as not necessarily defined.
The data were therefore plotted rfigs. 15, lo, and 17) in
the form of instantaneous bea.n-defl-ctio:i diagrams at
ti:e incr,-ments of ap :ro.-imratel:," C.017 second during the
stalled Dort of the pull-u. In these fi:-uras, if a line
faired through the snanwist poifits at "'bich deflections
were .-eacured did not pass thlrourh zero deflection at the
center line of the tail (see 2.500 seconds, fig. 15), the
bean diagram ;,as arbitrarily si-.jfted so chat the deflec-
tion st the center line ,.ar zero. The shifted beam
curves appear in the figures as dashed lines. This shift
of the '.eam. curve is considered. justifiable on the basis
that vibration in the airplane nay have caused slight
shiftin-L of the cameras or that the zero reading, fr the
particular frame may h-ave been in error; either of these
factors would have caused a uniform rhift of the beam
line. The change in tail load, v.hlch is indicated by the
deflection of each stabilizer tip is listed at the tnd of
each beari curve. In figures 16 and 17, the total load









NACA ARR Ho. I F05


change for each beam diagram is tabulated at the center
line. Deflections of the stabilizer are also plotted as
time histories, together with airs:ec .l, accelerations,
pressure, and electrical strain-gage records in
figures 18 to 20. A marked twisting action of the
fuselage may be noted during the stalled portion of the
pull-ups. 1be deflections of the right- and left-
stabilizer tips are not, therefore, a reliable indication
of the individual loads developed on the right and left
stabilizers except during the first part of the maneuvers
before the twisting of the fusel~a- was set up. The axes
for the pressure and electric strain-gage records were so
Oraw-- that the ordinates at the beginning of the run and
at the time cf maximum acceleration are proportional to
the loads computed at these points. Because both the
electric strain gage and the pressure capsule have
straight-line calibrations, succeedi'-, peaks are also
proportional to the tail load.

The three de Forest strain sages mounted on the left
stabilizer provided a measure of str--ess on the u.,per skin
of the left stab.ilicr' dur:.nr the runs of figures 16
art 17. The d3 Forert strain-..- records are shown in
figure 21 and a photo..:. cro aph of a typical record is
shownn in fi,;.e 22. .1thouh a history of the stress
encountered v;as recorded by a de Forest scratch gage, no
time record is available. 're peak stresses, therefore,
do not indicate the frequency of the applied load and
must be inter- ;-:eted in conjunction with other records.

t1h change in load from the level-flight condition
to the -oint of maximum acceleration that occurred
immediately before the stall is indicated by ALl in
figure 15 end the change in load indicated by the first
peak on the pressure or strain-gage record after the
stall occurred is indicated by AL2. The ratios of the
load iinmediately after the stall to the load before the
stall AL2/iL1 as indicated by ri-esaure-orifice ar.3
electric-strain-gage records, as well as similar ratios
determined frc.. the tip-deflection and de FIrest strain-
-"ve records, are listed in the followin- table:









ITACA A-RR No. O1405


pressure
orifice


Right
Fi ure in-
bcard



15 i 1.5
D1 1.5
1 I 1.1
19, 211 1.2
20, 21i 1. i


Left
in-
board



1.9
2.6

2.6
1.


Load ratio, AL2/A L1

Electri- Tip Ide Forest strain
cal deflection gages, from left tip
strain -----
gage I
(root of Riht Left7 n.
rilht tip tip 17 n. 0 in. 3 in.
stabili-
zer)


21. --- --- --
1.b, --__ '--- -
i.6 .8 1.0o 1.5 1. I 1.8
.5 1.5 1.l I 1. 1.5 1.5


The tabulated data 'sho'., that i. rmediately after the
stall large and abrupt increase in thc up tail load
occurred. Althou.ih changes in load indicated by each of
the records obtained are listed in this table, the indi-
cations of the presisurn crifices are discounited, not only
because of uncertainty rtgardin, th'e d-rnamic charc.cter-
i.tics of the pressure-record'linc system, but also because
of uncertainty regardin- the ap;--)lcability of point pres-
sures In relation to total loads under these circumstances.
TILe fact should also be noted that, owving to the inertia
of t-.i tail structure, mcment",ry pressure increments would
not necessarily result in cor.iparable stress increments.
The strain-rage and deflection mcasurrnments indicate that
the initial effect of the stall may result in up loads of
the order of twice those loads experienced immediately
prior to stalling.

After the initial tail-load increment occurs because
of wing stall-ng, the tail is buffeted repeatedly by the
fluctuating downwash in the turbulent wake from the
stalled v.ing. The possibility for resonance between the
turbulence frequency and certain natural frequencies of
the tail structure exists under this condition. The
frequency of the hcri-ontal tail in crin;ary bending was
17- cycles per second and "-he frequency of the complete
2
tail in torsion of the fuselage was 10 cycles per second.
From tests in the NACA full-scale tunnel, the frequency
of the turbulence fluctuations from the stalled wing was


__ ~_









12 NACA ARR iio. L:FO5


found to be 5.5 cycles per second at 65 miles per hour.
If this frequency were a linear function of true airs'-eed,
the r'a-.e would be from about 15 to 20 cycles per second
in the speed range covered by the pull-up tests and, at
some sp.-edc, would coincide with the b-nding frequency of
the tail. The turbulence frequencies, however, as shown
by the pressure records taken at the tail, were seldom
actually uniform for more than 2 or 5 cycles. Moreover,
where definite frequencies were detectable, the turbulence
frequencies aprpeared to ran-re fro..i about 10 to 35 cycles
per second and to be inde.pendJent of the speed of flight.
This lack of regularity in the turbulence pattern was not
unexpected because both the angle of attack of the wing
and the position of the tail in the wing wake were rapidly
var'in with time. In two of the pull-up maneuvers,
however, resonance with the tail structure occurred when
pressure fluctuations of a frequency close to that of the
tail were sustained for several cycles. An example of
this condition of resonance is shown by the pull-up
recorded in figure 14 where a large periodic build-up in
stress occurred as a result of a series of regular pros-
sure fluctuations. Figure 13 shows a somewhat similar
condition at a different airspeed. Both records clearly
indicate the mechanism by which excessive tail stresses
can be produced when tail buffeting occurs.


COIC JLUSIONS


The results of the present tail-load tests with a
modern pursuit airplane show the type and the general
magnitudes of loadings encounteredJ on the horizontal tail
of a heavily loaded pursuit airplane in accelerated
maneuvers. The survey of critical conditions is not
complete, however, because no tests were made in the
high-;speed and diving-speed ranr.gs. In addition, the
measurements that were obtained are less complete and
less detailed than are required to present an accurate
quantitative picture of the loads, in particular, the
loads immediately after the stall and during tail buf-
feting. The need for further investigation of these
conditions is indicated.

The conclusions to be drawn from the present tests
are summarized as follows:










T.. (4 R? 'o. IJ4r05


(1) 1r abrupt prl1]-'us, tLh critical! hc.rizcntal-tail
loals Vwere uc lcirads -:d ".._i e ,stfar ally prrrortio.al
to the a 2xiu nucr:._al czl.at ..o. or uiistailed ll-
ups, cxtripclatiocn olf t tn test relt s:ovw chat a tctal
tail clad. of "C00 o:-C .lCds wculd U, -j,-c- rx.':enced at an
ac3.el-rate!r, gof Of t Is totil tai 1 0o=-, JTou.t
1D,'' .-ou'irds voul.d b ri tIo t. 1.. 1 :. l. c' i::. tl n of the
e l';-tcr : ri n th ul 11 -u .

(2 ) 'n n .iS: 1 Ci .,'.a v' .i :. -.o ".'er ci the sr n-
v.wie ]cdinr- 0.1 t.e hc-rizro':til sil v:-s 'u-: y .;.-trica!.
Phbo.it ':30 rounds ,n'-e ": -O "o c? rrr'rie b7 'tie left
Sta'-iliz-r t..iim by cthe rL'ht .tat iliZer. -e .,nlitude
ocf 1vis :- s ... -3 try ,.; s e7sensIaV-, r- ie .'r.ient C' the
nor..,al ac e le i tion. '.t o'.e r c.' 'e : -:-. c ,-
wv.s '- .re .3 -U C .]i

7:) I- -'.ul J-u s c t s l r. J.urt I r"J. .s in
the tc.i load occ.uir.. '. :dl- ely ter t- :t l of
th:e :i c.- D.at.- io.' tf ..:-t.l &' r line t ?te
indicate that 1'C I.r;c-'.::,- Uf tt:: c r r r cof 1,':' o- recent
of lr :c I jl ut rl.or t ctal in: ;:y : ootained.

(' ) In st-. le.1. ull--vp |:a- ";'. : the ta *.''.
buffet d re. tudj.l, -r, 1.t: tur-. l nt flov- fi-O:- the stalled
winCi. 'iL.e .-o ib lity f 0ex e. .i.E street s utt to
re ora,.' in this reLnd i. i on :'. ':-.: -tca.


Lan le em'oral _I .c rona. tical Luora.tcry
'tlor.[ L Advi.sory Cr!._-.ictc for A3ronaitic
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`19
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NATIONAL, ADVIWIRY
OONM0ITFL KUB ArulnOXICS


Figure 1.- Three-view drawing of airplane.


^ ^ ------------







NACA ARR No. L4F05


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS


Figure 2.- Horizontal tail showing pressure-orifice locations.


Pig. 2







NACA ARR No. L4F05


(al Camera mounted in intercooler exit.


(b) Targets painted on left stabilizer.

Figure 3.- Installation for photographing tail deflections.


Fig. 3














NACA ARR No. L4F05 Fig. 4





ci,













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Figure 6.- Calibration of orifices from full-scale-tunnel tests.


NACA ARR No. L4F05




NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS


N"
1%




4
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NACA ARR No. L4F05


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS


8000


6000



4000


o Leff oufboard Ocifice
A Leff /inboard or/irce
o R p/h? ou/board or/'Y'ce
x A ertve /oad










J4ff' I ----


0 2 4
Ac c eler-f/'or7,


Figure 9.- Tall loads before wing stalled, computed from
pressure-orifice measurements In pull-ups to maximum lift.


Fig. 9







NACA ARR No. L4F05


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS


o LefI oc//hoard or/fice
A Left /iboArd f orfce
o RZ1h/ oa~hoard orifice
-E/er tor f/oah/i
- -/leafor f/ 6d/


4000



2000



0


Acce/er~sa/'on, g


Figure 10.- Unsymmetrical spanwise loading indicated by
pressure-orifice measurements.


B 1
.
1^


Fig. 10







NACA ARR No. L4F05


Z,
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190
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0 _ a ^ ^ _


Fig. 11


o 2 4 6 6 10
T/ne, sec


Figure 11.- Time history 6f 1800 left turn. Power on; manifold
pressure, 30 inches of mercury at 2450 rpm. Note dissymmetry
of pressures on left and right stabilizers.


NATIONAL ADVISORY
COMMITTEE FOI AERONAUTICS


12 14 16 /8






Fi


g. 12 .NACA ARR No. L4P05
NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS








Eet Ot7/f/L 4j:-11J-IIII II W
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-- 1 -.-.-.-.-.-
n^~1- --v^^
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200


/801
o


8 /0
Time, sec


Figure 12.- Time history of 1800 left turn. Power off. Note
that pressures on right stabilizer are slightly greater
than on left stabilizer.


I I I I I I i I I i I I I I I I_ I I I


t


(o-


la
U)t









NACA ARR No. L4F05


/ a 7 1f o/ zf


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Fig. 15a


Time,
Ssec.


/. 750


0



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0
1.0


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0

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0

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L4F05


0 40 6 So 40 0
Le/' Distance from ( fusekge, in.

Figure 15.- Instantaneous beam diagrams of left stabilizer
during a 4.Pg pull-up to maximum lift. Run 1 of flight 21B.


NACA ARR No.

NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS. Time,
LoF& Wy, sec.



/.0

,93
0 -l- 12 700


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1.0

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4AI








NACA ARR No. L4F05


Load-
chare,
2.0 "

/0 A168






0

1.0
440



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0

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0 --



0 76


0
0



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2.933



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3.017



3.033



30S0



S3.067
40 0


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS.
gwche on-e, Time,
S 'sec

3.083





*~~~~ 7.3 -- -- --
l 3.100

I
"_6
3.117



3.133


3.I50



3.1I7



3.183



3.200



3.217



3233



3.250



3.267



3.283


80 40 0


Distance from ? fuselage, in

Figure 15.- Concluded.


Fig. 15b








Fig. -16a


NACA ARR No. L4F05


0 -- /, 1.50


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS.
91 /9


So-- --- -- -- ----
0 TB=e"e L" P '1 'eI'-- 1.30



./2/


1.0


.I



/ 1.0 .,

.0 8 o













0 ---- -- --- -- 1.93
04 i --3 .. -.525












/.0
Slft Dstave efrom use-lage, In. Ri ght
f ,.o



/ 1.0 _





















Figure 16.- Instantaneous beam diagrams of stabilizer ob-
tained during a 2.4g pull-up to maximum lift. Pun 1 of
flight 24B.
-68 7"~4---- --- --- --- --- --- ------ --- yS-





01 ______^t: : ^-^ 0^ 1.933



o __ _.__9S- -CiI __ l SSo
1.0 -- -- -- -- -- -- 1 -- s ^ -



120 80 W0 0 0 0 /S
left Distance from h fuselage, In. Right

Figure 16.- Instantaneous beam diagrams of stabilizer ob-
tained during a 2.4g pull-ug to maximum lift. Run 1 of
flight 24B.







NACA ARR


No. L4F05 Fig. 16d


Load chan, Total load Load chanyhe,
1.0 -/I- e Time, sea

0 __- -- 2.33

1.0 lb/3

O ---- -o- 2.450
0 -G- 2.450

0 --T O_
0 -_,' 2. 467


0 _93 y- __ ____^lS 2._83


1.0

/-o
0 -- -o -Z. 1 481
I.O



0 o- --- Z. 0/7


u1.0
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1.0 "470

- 0 "-o--o-_. r-- .-= a.533



3 0 _.7. -








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1. 0

0 a 99217
/.O I 0_ __ _


0. G






I20 80 #0 0 40 80 /Eo
Lef t Disance from fuelag e, in --- ght

Figure 16.- Concluded. NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS.










ToW/ load Time
Load clhnge change, /b Load change, s*JC
lb lb





*4719


Io 0 s I I 1 8

0


./400
S, __1_







|0 .5-33
I.0











0O -- -_ 300
0O 2.5 00











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/. ro___ __ 1
a 2.613
I. O -- _._7




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/0 2.63

40
to e 2.-70
/0 e=-- 2S83


/0 r-- a 2.60
1.0 e 3 oo








0 -.-- I 2 .667






Left Distance from fuI.relage, in. R jght
Figure 17.- Instantaneous beam diagrams of stabilizer obtained
during a 4.2g pull-up to maximum lift. Run 2 of flight 24B.
NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS.
COC(MITTEE FOB AERIONAUTICS.


NACA ARR No. L4F05


Fig. 17a








NACA ARR No. L4F05



L oad ch
/.0 /_



/0

0



0

1.0





0
tIo
0.



10 o





0
4 /.0
c I,



0 -







/.0
-kl
fO ,-S^

^3 O_

j
st 1. 6


0 I
120 80 40 0 4 80 s I20
Left Distance from t fuselage i Right
NATIONAL ADVI
Figure 17.- Continued. COMMITTEE FORAERC


Fig. 17b

Tim
/b



2.900



















0.967
c~9es























3.083



3.00




SORY
ONAUTICS.








Fig. 17c


NACA ARR No. L4F05


/ gd baa
o1e Tt:a6 /
740
/.O 1 __ chwn aeb-. __ _, -- 2 83

1.0 /7 '

0 2. % Z.3

/.0 1- ,7

0 2.17 c
o -=-= ----"- -t- z. 7/







6r __ == _=4 I t 2.875?
1.0
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/40
0 2.750



=aac from-^2. W--e



0 14 4. 783




/ /.0
1,0


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1. 0







Lo
.0 =- .857

1.0 __ ; __ _^ __ __ __ __ __ __ _s_



1.0 '3

0 0 --- f^L- -- I 2.886
120 80 40 40 80 132
lef t Distance from it fuselage, in. .fight
NATIONAL ADVISORY
Figure 17.- Concluded. COMMITTEE FOR AERONAUTICS.








NACA ARR No. L4F05


I.~

KaL~Z


-e, sec
Figure 18.- Time history of 4.2g pull-up to maximum lift.
Run 1 of flight 21B.


Fig. 18






NACA ARR No. L4F05


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS.


a-4-4
o .8 12 /6 20 2.4 28 32
T77e sec
Figure 19.- Time history of a 2.4g pull-up to maximum lift.
Run 1 of flight 24B.


S __ _j I __ __


FPig. 19






NACA ARR No. L4F05


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS.


Si :
/. vv
S ~ I I 1 I I I I I I t


"t.
- TO-


IV, V-- -


210


.E HEfF3i i -2I1I i2i2\


2,1
woo-- -
^s /9--I---
d^' o


_1th13tIL I


l e, 30 3
T/me, 5ec


I iI


3S Ij


Figure 20.- Time history of a 4.2g pull-up to maximum lift.
Run 2 of flight 24B.


Fig. 20






NACA ARR No. L4F05


2.fg pull-up_ 4.Z 9


(fig. 1) (fig.20)

r^ AA/ _Ivr


Time non uniform scale


pul l- p


Front spar


74.5 in. from tip


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS.


4.4g pull-up 4.2 g pull-up
( figk/9) (fig. 20)


_WAANv


Time, nonuniform scale
Front spar 60 in. from rip


e.49 pull-up
(fig. 9)


4.2 g pull-Ap I
(fig. 20)

,--/-\ AA-


Time. nonuniform scale


Front spar


34 in. from tip


Figure 21.- Records from de Forest scratch-type strain gages
for flight 24B. (Complete data for flight 24B are presented
in figs. 19 and 20.)


" "


'ig. 21







NACA ARR No. L4F05


Figure 22.- Photomicrograph of a typical scratch-gage record.
Gage located 60 inches from tip of stabilizer. Maneuvers:
pull-up to 2.4g at 144 miles per hour and pull-ups to 4.2g
at 214 miles per hour.


Fig. 22





































































































































































































































































































_I_










i;




UNIVERSITY OF FLORIDA

3 1262 08104 978 4




UNIVERSITY OF FLORIDA
LDOC1CI IrENTS DEPARTMENT
120 MARSTON SCIENrCE LIBRARY
RO. eOX 117011
C.i',rESVILLE, FL 32611-7011 USA










I