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REPORT No. 364
THE PRESSURE DISTRIBUTION
OVER THE WINGS AND TAIL SURFACES
OF A PW- 9 PURSUIT AIRPLANE IN FLIGHT
By RICHARD
Langley Memorial
V. RHODE
Aeronautical
Laboratory
6S5
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REPORT
THE
PRESSURE
~0,
364
DISTRIBUTION
OVER THE WiNGS AND TAIL
PW-9 PURSUIT AIRPLANE IN FLIGHT
By
SURFACES
OF A
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RICEAItD V. RHODE
SUMMARY
involved in the loads that come into play in order to
produce an acceptable airplane, but must only Iinati
The investigation n?ported herein toa8 conducted at
how to apply the design rules imposed.
Langley Field, Vs., by the National kid~isory Committee
These rules (References 1, 2, and 3) have prcwen
for A6ronaufic8 at the request of the Army Air Corps fo
themselves satisfactory,
in genend, when appIied to
determine (1) the magnitude and distribution of aeroairphmes of conventional
type and purpose.
As
dynamic loads orec the wings and tail eurface8 of a
applied to new airplanes of 1sss conventional type, or
pursuit-type airphe
in the maneurer8 likely to impo8e
to new airplanes of conventiomd type but considerably
cmlical [oad8on the rariou8 8uba8sembliesof the airplane
advanced performance,
the rules are sometimes not
8tructure, (i?) to etudy the phenomenon of center of
satisfactory in all respects.
This is usually not dispre8wre morement and normal force coejlioientrariation
covered, however, untiI a structural failure occurs. In
in accelerated$ightl and (3] to meaeure the normal
many cases it is not discovered at all, fedure having
accelerations at the tinter of grarnty, wing-tip, and tail,
been avoided by a built-in stxength in escess of that
in order to determine the nuture of the inertia force8 acting
required.
m“nauftanew81ywith the criticul aerodynamic loads.
It is perhaps needks
to say that crashes resulting
The inre8tiga#ion compri8ed eimufianeow meaewefrom structural failures in the air, even though relament8 of pre8sure at 1$0 8tatiomsdistributed orer the right
tively rare, have a particuHy
bad efleot on the morrde
upper %<ng,left lower wing, right horizontal iail 8urfac48,
of flying personnel (with some notable exceptions) and
and complete wrtical eu.rfacea in one installation and
on the attitude of the public toward aviation, and must
the came number of poini8 di8iributed owr tho8eportiane
be eventually eliminated if cordidence in the sirplane is
of the u~ng8 in the 81ip8treamand the left horizontal tail
to became deep-rooted.
It is manifest, therefore, that
&aces in anothm installation, during a ~erie8of lerel
design of airplanes must be put on an
i the structund
j?ight nm8, pull-ups, roU8, 8pin8, dire%, and inrerted
indisputably
sound basis. This means that design
jlight maneurer8. Measured also were the acceleratiana I
rules
must
be
based more on known phenomena,
mentioned abore,amgularcehxMe8, air speed, and control I
whether discovered analytically or experimentally, and
positions &ultanew81y with the pre8wre~.
less on conjecture.
The results obtained throm light on a number of imporWhile a large number of papers have been published,
tant queetwne involving 8tructural design. Some of the
both mathematical and experimental, dealing with the
more interedirg rewlt8 hare beendiwuased in some detail,
dernal
loads on airplane structures, these hare not
but in general the report iefor the purpo8e of making this
been correlated to the point where a clesr picture of
collection of airplane-toad data obtained in jlight arailphenomena ooourring in the difkent
conditions of
able to tho8e interested in airpkme 8Wcture8.
flight can be obtained, if, indeed, it is possible to do so.
The most extensive single experimental investigation
INTRODUCTION
that has been made is probably the pressure distributestson the MB-3 airphme in 1923 (Reference 4).
tion
Granting that a major factor contributing
to any
These have been criticized on the grounds that the
increase in airphme performance is a deorease in weight,
airplane vras of a ~ery speoial type, and had individit is cIear that since the structural weight of an airplane
ualities of such nature that the results were not applicaconstitutes 20 per cent or more of the total, any saving
ble to the general prcbkc..
While some of this
that can be effected in the structural parts is worth
criticism
is
well
founded,
it
is
useks
to expect, except
while. But to design a structure light yet safe, the
to
a
limited
degree,
that
complete
pressure
distribuengineer must have a through and accurate knowledge
tion
inves&~atione
on
any
airplane
-will
furnish
data
of the oharacter of the loads that his structure must
suitable
for
the
solution
of
any
particular
prcblem.
withstand.
Actually, of course, the designer need not
This is true because any airplane is neoessariIy individbe thoroughly conversant himself with all of the facto=
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688
REPORT NATIONALADVISORl COtiMTT?iT)IU
FOR AERONAUTICS
ual (unless duplicated to a very fine point of perfection) and also because, which is probably
more
important, the labor involved in these investigations
is so great that it is impossible to treat Rriy one phase
of the Ioad problem adequately if a fairly complete
picture of the whole is to be obtained.
,
distribution.
The results are of immediate interes~
to those agenck
responsible for structural
design
-,
~d even though not analyzed to any great
extent shouId also be of value and interest to airplane
designem.
To expedite its presentation
there has
been no attempt to analyze completely any one phmo
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FIGUMI.-Front viewofITf% alrplana
Thus, the present report attempta to portray the
phenomena occurrigg on a p.umuit-type airplane in
the maneuvers that it is called upon to perfow,
or
what amounts to the same thing, in special test maneuvers outlined to impose the same conditions of load
that recur at the critical times in the more familiar
of the structural phenomena that am brought to light.
Instead, the present report presents the data as obtained, worked up to the stage where they can reridily
be used in studies of design methods, for the consideration of those concerned, and has in a few instances
called to attention the more obvious points in which
FIGURE Z—TbIe8quart.ar front vfew of PW4Y airpkne
maneuvem.
To this end, pressure measurements
were made on the right upper wing axtended to include
portions affected by slip stream, fuselage, and windshield, the Ieft lower wing, and the tail surfaces of a
with amelerBoeing PW-9 airphne, eimultaneoudy
ometer readings at the center of gravity, wing tip and
tail in the maneuwm above mentioned.
The data obtained represent a ve~ extensive collection of information on structural loads and their
structural
design methods now in use me open to
criticism in light of these redts.
The flight tests were made at- the LangIey Memorial
Aeronautical
Laboratory
in 1927 and 1928, at the
request of the Army Air Corps...
APPARATUS
The airplane,-The
a slightly moditied
airplane used in these tests was
Boeing P177+ pursuit airplane.
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PRESSURE
DISTRIBUTION
OF A PW-9
PURSUIT
AIRPLANE
LY FLIGHT
689
(I!lga. 1,2, and 3, Table I.) The miIitary load, irdudI and also noticeable @ Figure 10. The effect of this
ing the main tank, -ivas removed and the top coding
gap will be mentioned in the %ussion
of results.
Pressure
orfllces and tubing.-InstaUation
photoforward of the cockpit raised sIightly w that the test
graphs are gimn in F~ures 7 to 11, inchsive.
The
instruments
and apparatus could be accan.rnodated.
The pressure tubes leading from the upper vzi~u to
ori6ce and tubing installation is essentially the same
the fusekige formed faIse struts which increased the
as those used on previous tests, with aluminum tubes
used
throughout., except for short and easiIy replaced
drag and lowered the high speed about 6 m. p. h. On
lengths of rubber tube at the manometer connections
the other hand, the weight was reduced 50 pounds
and between the fixed and movable surfaces.
A
and the stalling speed lowered about 1 m. p. h. The
diagram of the orifice locations is given in F~ge
12,
pilot reported poor directional control, but the longiand the type of ofice used is ilI@rated in Figure 13.
tudinal control and aileron action vm.re e..cellent.
On
Manometers.-The
oritlces were connected to two
the whole, therefore, the performance and maneuverN. A. C. A. type 60 recording muMple manometers
ability were not reduced suficientIy
to affect the
which were located just above the center of gratit y
significance of the rew.dts.
in the space formerly occupied by the main gasoline
The wings of the F’W-9 employ the GWingen 436
tank.
These manometers are the same in principle
airfoil section throughout the span.
They are, howas those used in the MB-3 and ~~–7 tests (References
ever, tapered in plan form, and the upper and lower
wings dfier in plan form from each other (fig. 4); I 4 and 5), differing mainly from them in that they
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FIGURE3.—Three-qmrter rear view of PIT-4 sfrpkm
in addition, the upper W@ is washed in at the center
Table II gives the actual ordinates of aU
section.
sections at which pressures were measured.
During the preliminary tests the airplane broke a
wheel while taxying in after landing, pitched over on
its back and damaged the central portion of the leading
edge of the upper m@, which remained ahghtiy deformed subsequent to repair.
OutIines of some of the
deformed ribs are shown in Figure 5, compared with
the true sections.
A structural feature of this airphme which has a
bearing on the results obtained in the tests is the lack
of flying wirw in the rear truss, redting
in a structure 1sss r@d in torsion than the normal single-bay
biphne structure.
This dews the incidence of the
ceLhde to vary with changimg load, pmticdarly
in low
angle of attack conditions of flight with the center of
pressure well back.
Another characteristic
of the sirplane that shouId
be mentioned is the slot-shaped gap between the wing
and aileron, illustrated diagrammatically
in Figure 6
accommodate 60 pressure units each instead of only 30.
A diagrammatic sketch of the attachment of a-pair
of orifices to a pressure unit or ‘tcapsule” is given in
F~e
13.
Other instruments.-In
addition to the manometers,
the fobxing
instrnmen ts were used: N. A. C.?.A.
recording air-speed meter ~eference
6); N. A. C. A.
recording turnmeter
(Reference 7); NT.A. C. A. control position recorder (Reference 8); thee single component accelerometers (Reference 9), located as shown
in Figure 14; and a timer.
A movingpicture
camera was also used to measure
an@e of attack as will be exphiined later.
31ETHOD
The method used in thwe tests does not ditler in
any essentird feature from the methods employed in
previous pr=sure-distribution
tests. As it was desired
to obtain results for all of the ying and tad surfaces
simuhaneoudy,
if possible,’ the orifices were disposed
to cover the right upper wing, left Iower wing, vertical
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690
REPORT NATIONALADVISORYCOMMITTEEFOR AERONAUTICS
tail surfaces, and right horizontal tail surfaces. Limitations of capaoity of the apparatus that could be
carried aboard the airplane prevented the investiga—.
right upper and lower or left upper and lower) was
imposed by the impracticability
of running all of the
pressure tubes into the fuselage from one side, and also
i6 ‘O”
[ —4z”—
.-.
150”
+
1
..O
~
4
.1
‘o
Q
I
1
aivodl -
. —-
1
~ titer
, strut
- t merits
1,
.1
I
1
I
I
8
1
i
t
1
t..
I
U-J.
a
?
9
.
‘..1
A
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,—
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Lower whg
A
Yxw
-—
v
c
SfobifLzer & de
vato~
a, Stay
wire a f kmbmefits
FIGURE 4,—PW+3 wings and tall mrfaceswkh sparlmstlom
tion of the remaining surfaces simultaneously
with
those mentioned, but an independent set of rune was
made later to investigate more thoroughly
the slip
stream section of both upper and lower wings and the
left horizontal tail surfaces. The right upper and left
lower wing arrangement (in place of the more desirable
—
‘Hinge &
Fin 42rudhr
and strut attachments
because it was not desired to unbalance the airplane
even slightly unless absolutely necessary.
The pressure measured at each point was the algebraic sum of the pressures on the upper and lower surfacas (see fig. 13), no attempt being made to measure
the pressures on these surfaces separately, except in a
PRESSURE
DISTRIBUTION
OF A PTV-!3 PURSUIT
AIRPLANE
691
IN FLIGHT
few cases after the main tests had been completed. I ground effect. It was feh that level horizontal flight
Pressure curves were mechanically itttegratail to obtain 1 could be maintained accurately enough in this way to
allow of the use of inclinometer readings directly +s
loads and centers of pressure.
Simu.ItaneousIy with
The method failed because the slight
the presmre measurements,
records were obtained of angles of attack.
variations in engine speed and wind veIoeity caused
air speed, normaI acceleration at the center of gratity,
left upper R~U tip and tail, angular velocity in pitch I the inchnometer to oscillate enough to make the readinge quite erratic. Further attempts were made, flying
or roLI, depending on the maneuver being investigated,
as before, with a moving-picture
camera mounted in
tind control position.
These were synchronized
by
Dw.
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GZHingen
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436
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FIGUEES.-Comparison of rib A, B, and C with trus OdttIogen 436Se&on
means of the timer, and all of the measurements made
were plotted together against time to furnish a history
of each maneuver.
It will be noted, in glancing at the center of pressure
data, that resultant centers of pressure are plotted in
terms of per cent of “centric chord.”
The “centric
chord” is here defined as the chord passing through
the centroid of the plan form of that portion of the
wing extending from the root to the tip, and it is used
instead of some other arbitrary datum, because the
position of the mean C. P. on the “centric chord” of
tapered wings corresponds
fairly closely with the
mean C. ~. on the constant chord of straight wings.
With respect to the maneuvers in-mst~~ated, speciaI
attention was given level flight and pull-ups.
A considerable number of leveI flight runs were made in
order to furnkh a basis for the study of the results
obtained in accekrated fl@t, and also to furnish data
for comparison with wind-turmeI results.
Attempts to measure angle of attack in leveI flight
faihxl. First attempts were made with a pendulum
inclinometer mounted in the cockpit, the pilot flying
horizontality close to the ground, but with a sufficient
aItitude
to eliminate
possibilities
of encountering
41630-31~
/
‘
,
I
the cockpit, the lens axis being normal to the X2 plane.
vertical reference lines on a row of hangam were
photographed,
and the wgles of thwe lines on the
picture with the frame ~Ue were taken as angles of
attack.
This method showed promise, but with the
hangars located on the south side of the field, as they
are, it was impossible to obtain c.lear pictures, inaarnuch as the verticaI reference Iines mentioned above
were aIways in the shade.
hgles
as obtained were
probably correct b within 1°, but this accuracy was
not sufficient for the purpose for which they were
desired, and all of the angle readings were thrown out.
Centers of pressure, therefore, were plotted against
normal force coefficient as the independent variable.
The extensive investigation of the pull-up was made
for severaI reasons.
First, it is one of the few maneuvers that can be subjected to reasonably close contrdj or repetition with accuracy.
In other words, it
is a simple maneuver requiring the piIot’s attention
on only the initial air speed and movement of one
control surface.
It is, therefore, possible to obtain a
graduated series of maneuvers to allow of the study of
accelerations, anggar velocities, etc., as afhxting the
distribution of load. Also, the pull-up shows the &-
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692
REPORT
NATIONAL
ADVISORY COM?sHTT.JIEFOR
k“
AERONAUIWS
“
I
FIQUM
7.--Side vfewofPW’4 afrpkmewith hstrnmmd panelsremowd
,
FSQURE8.-Detnil
view of ms.fninstrument instdh$ion
showfng tuk+ connectionsto manometers
.
PRESSURE
DRTFKBLTTION- OF A PW–9
PURSUIT
AIRPLA3E
IN FLIGHT
693
FIGUZS 9.—DetW vfew of acceIaometu LustnnatfoItInt8n Show@ alK)mkmti~
●
I
FIG~X
10.—Wing tip aeeekrometeirLnshIMon
694
REPORT
NATIONAL
ADVISORY
tribution of pressure through a large range of angle of
attack and furnishes direct information on the most
important loading condition, viz, high angle of attack.
Third, the unusualIy far forward position of the center
of pressure at high angles of att8ck in accelerated
flight indicated by tests on the WZ-7 and Tiil “i@lanes
(Reference 5), as well as the coincident high values of
normal force coefficient, made it desirable to study
the high angle of attack condition at some length with
the hope of discovering reIations which might account
for the phenomena noted.
Because of the importance
of accurate air-speed
measurements in ~obtaining normal force co~ieqts,
the air speeds as recorded: from a Pitot-static head
mou!nted on the front outer strut were carefully
calibrated against those obtained from timed runs
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eluded there that individual pressures arc correct to
tithin + !2 per cent, while values of load aro correct
to within +4 per cent. “Recent investigations
of
the effect .of temperature on the capsule calibrations
indicak, however, that those figures should probably
be incfeased by about 50 per cent for the greater part
of these tmta, and in some few cases should be doubled,
These errors, however, have a minor ~ect upon the
measured distribution of load, since the temperature
erroti ire, in the main, a certain percentage for a given
temperature
regardless
of pressure.
Furthermore,
most of the capsules, being identical in construction,
show comparable temperature
errors,
Air, speeds are correct to within 3 per cent at all
spee~ in level flight. They are probably correct
withig. about 4 per cent in accelerated flight, sinco it
FmvE~ 11.-Detd.l oforifk?afnstallntlonin wing
over a measured course, and also those obtained from
a suspended “bomb” air-speed he~d. (Reference 7.)
This calibration, it was found, sufficed only for level
flight runs and the initial air speeds of the maneuvers.
It was necessary, for the pull-ups, to metm-rre the air
speed by another method.
For this purpose an airspeed head was mounted on an outrigger about 5 feet
out and slightly forward of the lower W@ tip in
order to eliminate interference from the wings at all
angles” of attack encountered.
~
speeda obtained
from this head checked the~(bornb”
readings within
one-fourth of 1 per cent at all angIes in level flight.
It was thus considered that readings obtained from
this head in vertical plane maneuvers woild be satisfactory, and all of the air speeds for the abrupt,
power-on pull-ups were, therefore, corrected on the
‘basis of results obtained from the outrigger. head.
PRECISION
A discussion of the sources of error in. pressure
measurements,
using the methods applied in the
Tresent tests is given in Reference 11. It was ccm-
is believed that interference effects have been largely
eliminated,
although some uncertainty
still exists
on this point,
Individual values of rib center of pressure are correct to within about 2 per cent in the high angle of
attack conditions, except in the case of rib L on the
low~wing,
and within an increasing error as the
angle of ‘attack decreases until, for conditions approaching zero lift, they are quite erratic rmd unreliable.
For this reason, momente instead of centms of pressure are given for these last cases.
Imgituciinal
centi-m of preisure of” resultant forces are correct to
within + 1 per cent-at high angles “of attack, while
lateral centers of pressure are correct to within about
2 pm cent.
Accelerations are correct to withh + O. 2g except
where noted.
Control position angks are probably not correct
to within less than 2 or 3 degrees, since the instrument was connected to the control levers in the cock:
pit, and did not measure change of angle caused by
deflection of the control system under load.
PRESSURE
DISTRIBUTION
OF A .PW-9
PURSUIT
AIRPIu4NE
IN FLIGHT
695
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*per,
wing
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35 “
Stabilizer’&
Fin g rudder
elevaf=
FIGUEX lZ.-Lccatfon of orificeson wing and Ml surfacei
Location of siatiom
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696
REPORT
NATIONAL
ADVISORY
COMMITi’EE
FOR AERONAUTICS
FIGURE13.–Dlagmm showkig type of onflw and connectionto capsule
I
“C, C. G. occekrotnefer
W, Win
occejerome+er
T, TOI“? occeleromefer
7,
. . .
.
Pw-B
oirplone
showing
uccelerome+er
locdions
._—
t
FIGURE 14.—Thre8-vlswdmwfng of PW-9 airplane sbowlng nccclemmeterkxatfom
PRESSURE
DISTRIBU’IXOX
OF A PW–9
Time synchronization
is correct in most cases to
within about one-twentieth
of a second, although in
some rum, because of instrument
diflicuhies, the
synchronization
is rather poor. For this reason, any
calculated quantities depending on the records of two
or more different instruments maybe quite umeliable.
This is particularly true of abrupt maneuvers in which
the measured quantities vary tkwough a wide range
in less than 1 second.
PRESENT~TION AND DISCUSS1ON OF RESULTS
The results following are presented and diswussed
in -groups according to the maneuver under in~e@ation, tiz:
1. Level flight.
2. Pull-ups.
3. Rolls.
4. spins.
5. Irwerterl flight.
0. Dives.
i. Pull out of dive.
In addition, the tail load, slip stream, and fabric preesure data are summarized and presented separately.
Yi’bile alI of the data obtained in the tests are not
gken in their entirety, representative
examples of
the most important
cases are included, and sIsc,
where it was felt they viould be of interest, more
complete data are given.
Level fight .—The leve~ flight results are gken in
Table III and in Figures 15a to 20. Figures 15a, 15b,
15c, and 15d show the distribution of pressure in terms
of q for four representati~-e cases through the speed
range or usefuI range of engle of attack.
An inspection of these figures at once discloses sereral salient
points, viz: 1. At low speed, or high angle of attack,
the center of pressure locus is, for all practical purposes,
at. the same per cent of chord aIong the span untiI the
tip is approached, where it bends suddenly to the
rear. This is particuhdy
true of the upper wing.
With respect to the lower ving, this point is queationabie since pressures at only four points were measured
on rib L, and the accuracy of the pressure curve at
that station is poor. 2. As the speed increases, the
center of pressure moves back varying amounts at different stations along the span, the trend being farther
to the rear as the tip is approached untiI, at high
speed, the center of pressure at the tip is almost twice
as far back as it is on the inner portion of the wing.
3. In practically
all cases, minor peaks of pressure
occur od the aileron.
This rearward trend of the center of pressure from
the center Iine ta the tip at low angles of attack is
Figures 17a to 18e afford a
worthy of special note.
closer study of the center of pressure movement.
Figures 17a and 18a show the nriation
of resultant
center of pressure with norma~ force coefficient for upper
and lower wings, respectively.
Figures lib to 17h and
PURSUIT
AIRPLANE
IN FLIGET
697
18b to 18e give the center of pressure vs. CMfor each
individual rib. As might be expected, the points for
ribs in the slip .strema and near the w~u tips are
somewhat erratic, but a suilicient number of runs
were made so that fairly good curves can be drawn in
every case, with the exception of rib L. The rearward trend of the center of pressure toward the tip as
normal force coefficient decreases is clearly apparent
from these curves, and leads to the suspicion that a
considerable amount of tm-ist (washout) exists, particuhxrly in the upper wing. This suspicion is strengthened by subsequent data which show a rapid decrease
in load along the span toward the tip at low angles of
attack.
It is desirable, therefore, to study the eff~t
of twist on the load distribution at some lengt.h.
The subject is discussed wry weIl in Reference 12,
in which the author points out the importsmce of
alight angles of twist at low angles of attack, and calls
attention to the fact that the twist may be a budt-in
feature of the wing, geometric or aerodynamic, or lit
may be a torsional deflection under load. He also
gi-ies methods based on the vortex theory for calculating the load curve for wings having certain simple
conditions of taper and twist, comparing the results
with load curves computed on the basis of the “strip”
method, and pointing out that the laiter method is not
sufhiently
exact in some cases. The method is given
more completely in Reference 13, and is extended b
include any tapered wing with any character of twist.
For present purposes, it is sufficient to compare the
load curve obtained in fLight with load curves calculated on the basis of the strip and vortex theories.
TO do this, it is pecessary to choose a particular condition of flight and to determine what the actual twist
*SS in this condition.
The condition chosen is highspeed le~el flight, since sewd
runs are a~ailable for
& condition, and hence experimental errors can be
To deterreduced by taking an average load curre.
mine the twist of the wing in this condition is not difficult. The most reasonable way would seem to be by
actual measurement of the rigged incidence and adding to this the extra twist induced by the application
of the load existing in the particular condition beiig
studied.
Unfortunately
this method is not possible,
since there is evidence that the rigged incidence
changed during the course of the tests., and hence the
true rigged incidence for any particular run or set of
runs is not known.
It is necessmy, therefore, to resort to a di.flerent method.
Assumptions are made as
follows: 1. That the curves of C.. -is. a for individmd
rib sections mot too near the tip and outside of the propeIIer influence superpose upon one another; that is,
that for ribs B, C, and D for instance, the center of
pressure is at the same per cent of chord at the same
2. That scale effect does not inangle of attack.
fluence the center of pressure position except to a negligible extent.
With these assumptions,
probable
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FIGURE21.-Compiu’f.wn of experimental
loedcum
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PRESSURE
DISTRIBUTION
OF A PVf’-9 PURSUIT
values of center of pressure for ribs B, C, find D in
high-speed level flight are determined from Figures
17c, 17d, and 17e. These values are referred to
Fwure 19, which shows the variation of C. P. with a on
rib C of the PW–9 model, and the values of W@ of
attack obtained therefrom.
The differences in these
angles are taken as the twist of the wing between the
stations used. Referring to Figure 20, these diHerences are plotted against span, taking B as a starting
point.
The curre is then disphmed downward until
!ts inner end, extended, intersects the base line at 3?4
feet, the point at which the wing chord is a maximum
and which is subsequently
to be taken as the root
section for the comparison of load curves.
The outer
portion of the curve is extrapolated to the tip as shown.
This extrapolation is not particuhdy
hazardous, since
it is based on a curve of measured rigged twist (not
shown) which has a similar form, diftering only in
degree. Figure 20 represents then the twist of the
wing in high-speed level fright, and includes both the
~~ged twist and torsional deflection.
Using this twist
and an angle of attack for the root section based on
the a~erage rake of CN at rib B and the slope of the CN
vs. cc curve for a similarly located rib on the model,
the span load curve is calculated on the basis of the
voruxx theory with the resuh shown in Figure 21. The
agreement is quite good.
The desigg load curve shown for comparison is
based on the strip theory without Wowing for twist.
It is only fair to point out here that the rigged twist
cm this particular wing is rather unusual, and is greater
than the washout required in the design and specified
by the manufacturer
to compensate
for propeller
, torque.
The discrepancy between the design load
curve and esperirnental
curve is-l therefore, greater
than it should have been. But, further in this connection, it is neoessary to mention that whiIe the torsional
deflection under the applied load in this case ody
amounted to one-half of a degree at the tip the appIied load factor was 1, whereas the condition designed
for at low angles of attack is the early stage of the
pull out of a dive, in which the apphd
load factor
may be several times unity and the torsional deflection,
therefore, increased in proportion.
It is manifest then
that not onIy is the rigged incidence a matter of importance to ccmider in the. structural dwign, but the
torsional rigidity of the structure as weII, and this
latter quite apart from considerations
such as wing
flutter and eeconda~ stresses. This conclusion is not
new, but seems to have been overlooked h this
country.
In spite of the probability
that stress
analysts and designers wilI tie-w with distaste the
necessity of carrying out the additional calculations
required to obtain a more exact load curve, there
seems to be no simpler alternative at present.
It may
be remarked, however, that the solution of the load
curve, once the twist is known, is neither dif3icult nor
AIRPLANE
IN FLIGHT
703
particukdy
laborious.
In relatively flexible structures, however, because of the mutual @ationship of
twist and load, it becomes necessary to resort to trial
and error, always a tedious process.
The span moment curve is readily determined on
the basis of the assumption that the moment coefficient
varies Iinearly with the Iift ccefikient when the moment curve for the basic airfoiI section is knovmt.
The minor peaks of load occum%g on the aileron
are believed to be due largeIy to the efhwt of the slotshaped gap between the wing and the aileron, since
the control position record showed that the aikmn
was neutral or very near neutraI in aII runs.
Pull-ups,-Pull-ups
were made as folIows:
1. Abrupt, power on, through the speed range.
2. Graduated abruptness, power on, at 70, 100, and
130 m. p: h.
3. Repetition of some of the foregoing ~th power
off .
Figures 22a to 25f show the variation in pressure
distribution throughout four abrupt pull-ups at speeds
between 79 and 181 m. p. h., and Figures 26 to 29 the
corresponding span load curves.
In the puII-up at
79 m. p. h. the airplane was nearly stalled before the
maneuver was begun, and the pr~e
curves for O
time are, therefore, characteristic
of the high angle
of attack condition in steady flight. As the maneuver
progresses, however, the angle of attack of maximum
Iiffi is reached and the mean ceni%r of pressure (see
@. 30) on the upper wing movss forward to about 26
per cent of the centric chord. It W be noted from
the pressure plots for the medium and high-speed pullups and from the time history curves (figs, 30 to 55)
that the maximum forward position of the center
of pressure is the same within the experimental error
in every case, the average value being about 27%
per cent. (Exceptions: Rti 73 and Run 137.) This
vahe agrees with the vaIue found in wind-tunnel teds
on the PIT-9 celhde.
@.ference 14.) It does not,
however, check the value found in monoplane tests,
the ditTerence being about 3?4 per cent with the fuUsoale center of p~ure
farther forward.
This discrepancy is ascribed to “biplane effeot” or the mutual
interference of the upper and lower wings.
On the lower wing the center of pressure do- not
move so far forward, reaching an average value of
about 31 per oent. This value does not check the wind
tunnel value, but comparisons are not strictly valid
since the wind tunnel Iower wing model was extended
to the phme of symmetry, and hence has a greater
span, proportionally,
than the fuU-scaIe wing.
While the iduence of biplane effect on the center of
pressure is not an unknown phenomenon, thus far it
has not been taken into aoommt in the design rules.
The itnport.anco of this phenomenon should not be
overlooked.
While it can be seen from F~res
56 to
59 that the distribution of load aIong the span on both
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REPORT
NATIONAL
ADVISORY
FOR AERO?ikETICS
COMMITTEE
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PRESSURE
DISTRIENJTIOX
OF A PT–9
upper and lower wings at high angles of attack agrees
remarIiably well with the load cumw deri-red from the
design rules, stilI, the discreprmcy between the actuaI
center of pr=ure
and design center of pressure (3 I per
cent) results in an unsafe condition for the front lift truss.
Figure 60 shows the comparkon between the actuaI
gross spar load curves from Run 132, and the design
load curves (dead w~~ht not subtracted) for I@h angle
of attack on the basis of the same load factor or totaI
load. A smtdl difference in relative wing load esists
between flight and design rules (1.37 and 1.29, respectively, for this case) which ma=~es
the effect of
center of pr=ure
discrepancy on the upper wing to a
alight extent.
It is seen from the figure that the greater
portion of the front uppersparis overloaded, but that it is
underloaded at the tip. This latter condition is caused
by the rearward d.ksplacement of the center of pressure at
the tip sections, which is characteristic at high angles of
attack.
The effect of the actuaI load distribution on
the primary bending moments is shown in F~e
61.
These curves are gives for illustrative purposes only,
but are true comparisons on the assumption that a pin
joint exists in the spar at the cabane strut attachment
poin~ whichis not actualIy the case. Theyshow,however,
that a considerable discrepancy e.tits between the bending moments actually obtained and those assumed—
in this particular case, 36 per cent on the unsafe side.
In view of the importance of the l@h a@e of attack
condition in design and of the position of the center of
pressure in this condition, it would seem advisable to
shift the design center of pressure forward on the upper
wings of biplanea by the amount indicated for the
combination by wind-tunnd
tests o~ theoretical calculations, if such are available, or if not, to shift the upper
wing center of pressure forward arbitrarily by 3 to 5
per cent to be on the safe side.
It might also be advisable to increase the design
bending moment of the front spar in the middle of the
bay of externally braced types by an arbitrary amount
to allow for the rearward displacement of the center
of pressure at the tip, except in cases where the effect
of tip shape is definitely known.
The existing rules
apply an arbitrary
increase in bending moment of
30 per cent fmm the outer point of inflection i% the
tip, which takes care of possible increases in stress near
the strut attachment
point arising fmm excessive tip
Ioads. The tip loss assumed, which is considered to be
more than that actually encountered, is supposed to
take care of possible excesses of stress in the bay, but
is not sutlicient to provide for conditions arking from
displacements
of the center of pressure from the
assumed vahe.
The above discussion accentuates the
need for extensive research on load distribution over
wing tips, which must be done before the arbitrary
nature of such revisions as suggested can be eliminated.
PURSUIT
AIRPLAXE
IX FLIGHT
7’33
Figures 62 and 63 are presented to show the variation
of spar load distribution
throughout two pull-ups of
dit7erent character.
Both show the same generaI
resuks for similar portions of the maneuverl the outstanding points being the tip peaks on the upper rear
spar in the region of l@h angle of attack and the
si.darity
between front and rear spar load distribution on the lower wing in all conditions.
F~ges
30 to 55 represent histories of all of the
pti-ups
investigated
for which satisfactory
records
were obtained.
They show the relation existing between the loads acting at any instant during the
maneuver.
It is seen that the upper and lower wingg
reach maximum loads at the same time; also, that the
maximum taiI load is down and occurs early in themaneuver,
well before the wing loads reach their
ma.tium
values.
It wilI be noted, too, from the taiI
acceleration records that this down load is not accompanied by an appreciable weight or inertia load, since
the acceleration at the time is approximately
zero.
This suggests that a critical loading condition for the
fusdage is the pure maximum down load on the tail,
which agrees with the present design rules.
Referring to the abrupt power on pu~-ups in the
high angle of attack condition, the acceleration at the
taiI remains at a fairly constant ratio to the acceleration at the center of gravity, and is of the order of
one-half of the latter value (considering Ig the datum).
It would fo~ow from this that the present rules are
in error in assuming the l@h angle of attack condition
to be equivalent to a static condition with the basic
load multiplied by an appropriate
load factor, and
that they shouId take account of the fact that in this
condition there is, in most cases, an angular acceleration about the pitching axis which radti
in a varying
inertia Ioad along the fusekige. This inertia load, of
course, depends on the moment of inertia of the airpIane and the effectiveness of the controIs, and differs
In any case, however, the
with diBerent airplanes.
condition would be such that the inertia load forward
wouId be greater than anywhere else. This points
to the conclusion that the present rule giw.s design
loads for the engine mount and forward portion of
the fusdage that are too sM,
and hence m8y prove
unsafe.
Referring now to the histories of the power-off puUups, the above discussion does not apply.
The tail
accderations now are about the same as the center of
gratity accelerations and usually a Iittle higher.
In
short, the conditions in the power+ff pull-ups are in
fair agreement with the conditions assumed in the
design rules; the conclusion is therefore drawn that
two possible critical conditions exist for the fuselage
at high incidence, one corresponding to power on and
the other to power off.
.—
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a
’37
736
REPORT
NATIONAL
ADVISORT
A point of considerable interest is disclosed by the
results of the power on pull-ups.
It has been mentioned previously in this report that one reason for
investigating the pull-up at such length was to doterti’e the efFect of the pitching motion on the center
of pressure position in the high angle of attack condition, and also on. the m8xirmum normal force coefficient. It has been shown that the maximum forward
position of the center of pressure is the same within
the experimental error, regardhss of the character of
It can, therefore, be concluded that
the pull-up.
pitching dow not affect the C. P. in the high angIe of
attack condition, at least for the airfoil section and cell
used here. The maximum normal force coefficients
are erratic; that is, for different runs the values of
normal force coe~nt
are not the same, and seem
to bear no clear relationship with any other variable.
This is believed to be due to the lack of accurate
synchronization
b~tween the air-speed record and
pressure records. In general, however, the values are
greater than would be expected. The upper wing CN
from the model data (Reference 14), corrected for
scale effect reaches a maximum value of about 1.43.
The full-scrde results show some maximum values as
high as 1.8, with the average maximum about 1.66.
In only one case is the maximum CN less than 1.43.
This would indicata strongly that the normal force
coefficient of an airfoil with a positive pitching motion
attains a higher value than a similar airfoil under
This indication is further subst-ansteady conditions.
tiated by Figure 64. It will be noted that the accelerations in the power on pull-ups lie close to the theoretical curve,
NQw, since the theoretical curve is
based on the =sumptions
that the airplane is pulled
up with no loss in &ir speed, and that ON ~,c~~ti~for the
pitching airfoil is the same as that for steady fight at
stalling speed, in view of the fac~ that the air speed
actually does fall off an appreciable amount by the
time the peak load in a pull-up is reached, it can only,
be concluded that the close agreement between the
experimental points and the theoretical curve is due
to an unusually high ma&mm
normal force coefficient occurring in the maneuver.
This close agreement, therefore, could only be expected to occur for
certain conditions in which the decrease in air speed
just offsete the increase in normal force coefficient.
In these tests the conditions are met in abrupt poweron pull-ups started from horizontal flight.
In the
power-off pull-ups the air speed falls off more rapidly,
and as a result the measured accelerations in general
come below the theoretical curve.
It might. be expected from this that in abrupt pull-ups made from
steep dives in which the weight component still acts
forward at the peak load, the air speed would not fall
off so rapidIy, and as a result accelerations in excess of
the theoretical would be experienced.
While none of
the present data show this to be so, it has been demonstrated in maneuverability
tests on both an F5’C%
COMMITTEE
FOR AEROX4UTICS
and an F6C-4 airphme (results not yet published)
that this condition actually occurs.
These high normal force coefficients can probably
be accounted for on the basis of tho known phenomenon
of vortex or eddy formation.
An important consideration @ these formations over a wing beginning to stall
is that time is required for the back flow in the boundary layer to reach the stage where the &ret vortex can
be ftied.
In “the pitching wing, therefore, the flow
—
does not instantly break down when the steady fligh~
angle of rnmimum lift is reached, but continues in
force for a short timo whale the wing rotates beyond
this “tihgle, with the rwdt that the lift coatinucs to
build _up to an abnormal value. Whatever the cguse,
the result is interesting and important, as it can very
conceivably have some eflect on high angle of attack
load factors, and also on the relative distribution of
load between upper and lower wings of a biplane.
With respect to this latter point it would seem
reasonable to expect that the lower wing normal
force coefficient at peak tctal load in a pull-up would
be nearly equaI to ita steady flight value, since the
effec~ of the upper wing in the latter case is to delay
the ~~eakdown in flow over the loiver wing, and henco
the explanation given for the upper wing does not ““
apply, at least for the peak total load condition,
Fur@rmore,
the slope of the lower wihg CN curve is
progressively
decreasing
with increasing
anglo of
attack (Reference 14), and at the time of maximum
load on the upper wing has a fairly small value. This
is borne out by the relatively ffat lower wing load
curve in the present data.
The above asstimption is substantiated
by the facts,
the model lower wing C~ at the angle of attack of
maximum upper wing Ioad being 1.18 (corrected for
scale effcct), whereas the average of the maximum
full scale lower wing normal force coeflkients from tho
power on pulI-ups is 1.23, a good agreement,
If the
hypothesis concerning the high upper wing normal
force coe%icients is correct, the foregoing lower wing
values can be made to agree still better by taking the
modkl CN at a higher angle of attack.
T@ effect of the phenomenon discussed in the above
two paragraphs on the relative wing-load ratio would
be, then, to increase the ratio for the pitching high
angle of attack condition.
Comparison of this ratio
from the model data. and from the full scale pull-up
data shows this to be true, the former VRIUOat maximum upper wing load being 1.16, whereas @ latter
value (average of all power-on pull-ups) is 1,35, The
ratio of these two values, since the lower wing normal
force coefficients me essentially the same, is then
practically equal to the ratio of the steady flight upper
wing maximum normal force coefEcient to the pitching upper wing masimum normal force coefficient.
In all pull-ups, the acceleration at the wing tip is
of approximately the same magnitude as the acceleration” at the center of gravity, differences being a.c-
PRESSURE
DISTRIBUTION
OF A F’iV-9 PURSUIT
counted for as foIlows: in the edy
stages of the
maneuver, the flexibility and lightness of the wing
structure aIIow the wing tip to accelerate more rapidly
than the fuselage; in later stages of the maneuver,
slight ding
motion is undoubtedly the cause.
.4 phenomenon ewkk.need on the original records,
but not given in the histories, shows the esistence of
a high-frequency
vibration of large amplitude at the
wing tip occm-ring in each case at the iustant of mtimum load and continuing until the end of the record.
(F%. 65.) Whether this oscillation is caused by the
sudden breakdown of air flow on the wing tip, by the
abrupt change of inertia load, or is associated with the
more or less welI-knowm phenomenon of wing flutter,
Vibration tests conducted on
can only be surmised.
the airplane showed the natnraI frequency of the wing
structure as a whole to be: (a) From S.7 to 10 oscillations per second in bending, the differences being
caused by the addition of various weights up to 22
pounds at the accelerometer location; (b) 8.5 oscillations per second in torsion.
For tho upper tip alone,
a rigid support being provided at tie strut attachment, the frequencies were taken from 10 to 14 in
bending with weights added as before, and 12.5 m
In the bending tests, the higher frequencies
torsion.
corr=pond to the condition of no estra weight, which
is the condition occurring in flight.
It was diflicult to determine frequency of oscillation
from the wing accderometer
records because of the
superposition of several waves of dif%rent frequencies
and amplitudes, but in general the frequency of the principal oscillations was of the order of from 10 to 17.
Strangely
enough, the higher frequencies occurred
in the slower power-off maneuvers, thus precluding the
possibility that engine tibration was responsible.
RoIIs.-Figures
66a to 67i show the distribution of
Ioad throughout a right and a left barrel roll
The
initial stages of the maneuver up to the peak load
are identical with the pull-up, ~~cept that the loads
build up on the rudder in addition.
Beyond the peak
load, autorotation starts and is evidenced in the right
roI1 by the shapes of the pressure curves on the right
upper wing which are characteri..tic
at angles of
attack above the stall. (See Reference 14.] The left
lower wing in the right rolI does not stall, the condition
being similar to that at high angksnearnmximudift.
In the left roll, the right upper wing is not stalIed, but
is near or at ma.sirnum lift, while the left lower wing
shows etidence of being st.aUed very slightly.
The
horizontal tail surfaces show disseymmetqy of load.
In tie left roll, the distribution of pressure on the right
side is very simiIar throughout to the pressure distribution in a pull-up, but in the right roll the down load
on the elevator is replaced by an up Ioad in the latter
stage of the maneuver.
This dissymmetry is probably
caused by the influence of the fuselage.
.41RPL4X~
IN FLIGIIT
737
Span Ioad curves for the two roIIs are given in F@res
65 and 69, and time histories in Figures 70 and 71.
~ Figure 72 is a span load and inertia load diagram
combined from correspon~u
points in the right and
; left rolk at the condition of masimum dissymmetry
of load on the upper wing. This point occurs at. 1
66 and 67 that
{ second. It mill be noted from F~es
the two maneuvers are closely similar, viz, the control
~ action is practically
the samel and the rua.simum
~ accelerations at the C. G. are practically equal and
occur at the same time rdative to the start of the
maneuver.
Figure 72 is corrected for the difkwnce
in total load in the two maneuvers.
The outstanding
points in comection
with the
unsymmetrical
condition are: 1. The masinmm dissymmetry of load on the upper wing occurs shortly
after the masimum total load, and the totaI load in
the UUSPIUM trical condition is not much less than this
maximum.
2. The diss.pune~
on the lo~er wings is
of such a nature as to oppose the rolling moment due
to the upper wing. The first point is of particular
interest in that. it shows that the unsymmetrical
cabane load condition shonId be analyzed with an
average load factor equal to the high incidence load
factor.
This, of course, applies only if the arbitrary
nature of the &sting rules. is eliminated, and they are
reformulated
to include the effect of varying inertia
load across the span and the true effect of the lower
wing. The second point is of considerable interest
&o, altbough in the light of the delayed stall of lower
wings of a biplane indicated by model tests, the resuh
is not particularly
surprising.
(Reference 14. ) The
remdts indicate that at the point of masirnum upper
wing rolling rno~ent, the angle of attack of the strolled
side of the ce~ule is not beyond the st~~
angle of the
lower tig
on that side. This would explain the count--rolling
moment of the Iower wing.
The inertia load along the span varies as shown in
the figure. The slope of the line is calculated from the
angular acceleration of the airplane (5 radians per
second per second) as determined from the slope of
the angular velocity record for the right roll. The
wing tip acceleration for the right roll checks this
slope very cIosely, although the wing tip ~cceleration
for the left roll does not.
It maybe of interest to note that the angular acceL
eration as calcdated from the known rolling moments
and moment of inertia of the airplane checks the -due
from the anguIar velocity recorder and accekrometer
faidy closely. This can onIy be done, however, if the
total wing rolling moment, as determined from F~e
72, be assumed as that for the right roU. The reason
for this is that. the anguhw -velocities in right and left
rolls are not. the same on account of propeIIer torque
and because the load on the stalled wing in a roll is
quite sensitive to changes in rolling -relocity, whereas
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744
REPORT
NATIONAL
ADVISORY COMMITHZE
FOE AERO.SAUTICS
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FIOURE70.-Time bktorg of a @t
bmrel roll at 167miles w how.
(Run No. 222)
-
PRESSURE
DISTRIBUI’ION
OF A PTF-9 PURSUIT
AIRPLA3ii
745
LW FLIGHT
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●
746
REPORT
NATIONAL
ADVISORY
the load on the unstalled wing is not. Therefore, if the
right roll is chosm, we have the actual measured value
of the load on the stalled or right upper wing,
whereas if the left rolI were chosen we should have to
assume that the load on the right or stalled wing in
the right roll was the same as the lotid on the left wing
in a left roll, which would be a rather hazardous assumption,
Assuming, therefore, that the wing-rolling
moments given in Figure 72 represent fairly welI the
true wing-rolling mornant in a right roll, it is onIy
necessary to subtraot the opposing torque of the
propeller (in this case approximately
900 pound-feet,
based on a knowledge of the propeller and engine
characteristics
and the conditions at which they me
operating) to obtain the total moment causing angular
m~>terd
CO~ITTEE
FOR AERONAUTICS
they assume a major importance in the study of the
phenomenon of autorotation.
Figures 75a to 75q and
76a to 76p show the distribution of pressure for a right
the
and a left spin. In the right spin throughout,
left lower wing remains essentially unstalled, but at
an angle very close to that of maximum lift, whh the
right upper mung begins to show evidences of becoming
stalled at the tip at 2 seconds, imd thereafter the stall
progresses from the tip toward the center until at 2%
seconds the entire wing is stalled to the plane of symmetry.
This condition
continues,
with the load
becodg
smaller, to 4 seconds, where the loads start
to build up rapidly and the stall apperm to become
more pronounced; that is, the center of pressure locus
moves back, indicating an increase in angle of attack.
L?> 7.42 fi
i%eef afonq spon
Fi@JWE7G-Span I&IdOU?VH
fromrfghtand left rolls mmbfnd to show worstunsymmetricalcondItlon
acceleration.
pound-feet.
This
value
is: 9,600-900,
or
8,700
M
a, therefore, is —
I
or
8,700
~
= 6.13 radians per second per second,
2
Time histories of ~mht and left rolls at slower speeds
are given in Figures 73 and 74. These runs do not fit
together as well as the higher speed rolls, but it is
apparent from a study of the time historithat the
time of maximum dissymmetry occurs relatively late
in the maneuver with the total load from 80 to 90.
per cent of its maximum.
Spins,—Loads in the spins are relatively small and
uninteresting
from the structural point of view, but
The wholo motion appeam to be unsteady up until the
records were discontinued, although it is quite possible
that a steady condition would have been reached had
the spin been continued longer. Loads on the hori---zont al tail surfaces remain characteristic of high angles
of attack with the elevator well up until rtbout 5 seconds, when the load builds up in the po”sitive direction,
the pressure distribution becomes more irregular, and
the eIevator load changes from negative to positive.
Thesq changm occur ;Vithout movement of the elevator
and indicate a change in ah-flow conditions, probably
caused by the change in direction induced by the increased anguhw velocity.
It is interesting to noto from
the tie history (fig. 79) of this maneuver in conjunction with the pressure distribution
curves that the
horizontal tail moment changes early from positive to
negative and continues to build up in tho negative or
‘
PHESSURE
DISTRIBUTION
OF A PW-9
diving sense to a conaiderable value, whiIe the airplane
is still settling into a stronger spin. Even the load on
the elevator changes in direction from down to up
with the angular displacement remaining constant at
about 24° up.
In the left spin (figs. 76, 78, and 80) tho right upper
wing remains unstalled throughout, but the left lovmr
wing appears to be stalIed from the beginning of the
record and continues so until the end of the record,
the stall becoming more pronounced as the spin conThe motion, as in the right spin, is unsteady.
tinues.
As in the right spin, also, the horizontal tail load
increases upward and the trd moment continues to
build up in the diving sense, while the elevator remains
we~ up. The distribution
of pressure, however, is
more regular than in the right spin, undoubtedly
because the right- side is no-iv advancing into the wind
with the fuselage and vertical tail surfac% behind it
exerting little disturbing influence.
It wiU be noticed,
too, that the horizontal taii moment in the left spin
does not reach as high a vahe as in the right spin.
~ertical tail moments in both spins remain nefdy
constant and at about the same value in both.
It may be lemarlmd here that this particular airpIane was easily controlled in the right spin, but came
out of the left spin with considerable difficulty.
lmerted
flight.-Attempts
were made to obtain
records in the inverted flight condition without much
success, as the pilot found it difEcult to maintain the
condition.
The inverted
attitude
was roached by
means of a half loop, but could not be heId, the airplane losing altitude and having a strong tendency to
nose down and continue the loop. Figure 81 shows
the distribution of pressure in the inverted condition.
It will be noted thht the center of preEsure locus is
rather far forward, indicating that the angle of attack
has not reached the mgle of maximum negative Iift.
Span-load curves me given in Figure 82.
Dives,-A
representative
dive is ilktrated
in Figures 83. to 86. Leading-edge
pressures reach an
exceedingly high vcdue, as noted from Figure 83 and
Table III.
The span-load curvee (fig. 84) show the
tiect of the twist of the upper wing, the load inboard
being positive with the tip operating below zero lift.
A curve showing the variation of moment about the
leuding edge along the span is given in Figure 85.
The spar-load curves given in F~re
86 show the efkot
of the washed-in centraI portion of the wing very
clearly and also, to some extent, the effect of the
washout, the front spar load increasing toward the tip
with the rear spar load decre~~ing.
h’ormally in a dive, m in the caae reported here, zero
lift is not reached, although it could be attained if the
airphme were rosed over sufficiently.
This condition,
however, is an uncomfortable
one for the pilot, the
sensation being that the airplane is slightly over on
its back, which it really wotid be in a majority of
PUTMUIT &EtPLANE
IN FLIGHT
747
cases. Zero lift can be attained, however, with the
nose of the airphme only dightly beIow the horizontal
in a sudden push into a dive. Pressure distribution
for this case is given ia Figure 87; it is sirnihr to that
in the dive, but with the negative areas at the leading
edge in greater proportion.
The pressures, however,
are much 1sss because of the lower speed. Span-loading curves in F~re
88 show that a considerable portion of the upper wing is at negative Lift, while loads on
the central portion are positive, and that the lower
wing load is positive throughout the span.
Spar-load
curves me given in Figure 89. The effect of the twisted
upper wing is apparent in these curves, the down load
increass
on the front spar toward the tip and the up
load on the rear spar decreasing
The condition woukl
be much more severe in the case of a fast dive at zero
Iift., since the twisting moment wouId increase as the
square of the speed, and hence the deformation would
increase, remltiug in a greater proportion of load on
the outboard portion of the hont spar.
It is conceivable that the outboard rear spar load might reduce
to zero, the entire load being carried on the front spar,
in which case a form of partial inverted flight condition
would exist that would probably be critical for the
Ianding wirea and the leading edge of the wing.
Pull out of dive.-An
interwting,
though isolated,
case of a low incidence loading condition occurs in the
pull out of the dive, RurI 226. h connection with
this condition it should be said that the puIl out was
normally executed; that is, it was made cautiously with
due regard to the speed at which the airplane was
traveling, and that it in nowise represents a speciaI
t-t
condition.
The point chosen in working the
records W= the point of rmdrnum
acceleration (3.69),
which occurred early in t-he pull out and which was felt
would probably represent the most severe rear sparIoading condition that could be found without working
up a Iarge number of points.. Figures 90 and 91 represent the distribution of load for this case. The condition is seen to be characterized by moderat.dy high load
with the center of pressure fairly well back and span
load tapering off rapidly toward the tip. The spar
loads given in Figure 92 show the tiect
of these
characteristics.
Spar loads derived from the speciiied
Iow angle of attack loading condition are given for
comparison, although the total loads for the two cases
are not the same. The total load for the observed
results is the true total for this case corresponding to
an applied load factor of 3.6, whereas the derived
curves are based on a load factor of 325 which is the
low angle of attack load factor divided by the intended
factor of sdety of 2. It can be seen that the upper
rem spar load is welI with
the design load, and
although the front spar load is in excess of the specified load for low angle of attack, the condition is not
critical for the member.
On the lower wing, however,
both sDars are overloaded. the front one afmin being
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748
REPORT
NATIONAL
ADVKSORY COMMITTEE
FOR AERON.4UTICS
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PRESSURE
DISTRIBUTION
OF
A PW-9
PURSUIT
AIRPLANE
749
IS FLLGHT
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.
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12
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16
PRESSURE
DLS~UTION
OF A PW-fJ
PURSUIT
AD3PLAWE
759
IN FLIGHT
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760
REPORT
NATIONAL
ADVISORY
COMWTI’EE
FOR
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AERONAUTICS
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BIOUBE
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REPORT
NATIONAL
ADVISORY
taken care of in the high angle of attack condition.
On both win~ the center of pressure is somewhat
farther forward than is assumed in the design, the
total load is higher, and the reIative efficiency of the
uppm wing is only 1.02 as against the design twsumption of 1.29. Of these factors, the first tends to decrease the severity of the condition as compared to the
spcicifications, while the second and third tend to
increase it. It is impossible to say from one case, of
course, what center of pressure position is most Iikely
to be encountered in the critical condition for the rem
spars, since the load factor changes with the pressuro
distribution,
Defiite conclusions on this point, therefore, are not warranted.
The comparison is little
changed if both total loads are assumed to be the same,
and the effect of the discrepancy betwem the reIative
upper wing efficiencies remains as the most significant
factor.
The case, however, emphasizes the importance of extended research into the mutuaI interference
of biplane wings, and also points out that the supposed
factor of safety of 2 is not unlildy to be overworked,
even under conditions which are not considered in any
respect abnormaI.
From the standpoint
of the
operating personnel, it wouId seem that the airplane
should be handled very gently in pull-outs from fast
dives untd the low angle of attack condition is more
thoroughly known and provided for.
Tail loads.—The important tail loade are summarized
in Table IV. For each run ‘listed, several conditions
are given in chronological order to cover the following:
1. Maximum horizontal surface down load; 2. Maximum horizontal surface up load; 3. Maximum stabilizer pressure; 4. Maiimum ele-rater pressure; 5. Maxifin pressure;
mum vertical surface load; 6. Masimum
7. Wmimum rudder pressure.
Vertical surface pressuras were not worked up for the pull-ups since they are
relatively low and of little interest,
Tail surface specifications for pursuit type airplanes
impose average loads of 45. and 40 pounds per square
foot on the horimntal
and vertical surfaces, respectively.
It is interesting,
as a primary comparison
between obmrved and specified loads, to note the values
given in the columns of average loads in Table V.
It must be borne in mind while doing this that the
specifications are supposed to incorporate a factor of
safety of 2; thus, any value given in the table exceeding
one-half the specified value is to be considered. an overload and vice versa. It will be noted that the horizontal tail surface loads in the pull-ups are generally
lower than one-haIf the design load, exceeding this
value in Runs 134 and 137, and cIosely approaching it in
Run 133. It is worth noting that the applied lorid
factor (C. G. acceleration) in Run 133 is 6, making the
safety factor for both wings and tail surfaces approximately 2 (on the basis of loads only) in the same
maneuver,
In the two dives listed, however, the tafl
loading is excessive, being 26.4--and 30,7 pounds per
square foot for Runs 213 and 226 respectively,
CO.WITT’EE
FOR
AXRONATJTICS
indicating that the spetications
should be revised
upward.
Other high loadings on the horizontal
surfaces occur in the high-speed barrel rolls, Runs 222
and 225, the values exceeding one-half the design loading, but remaining less than the dive loadings.
It is
interesting to note that in the rolls the max+mum up
10ad is of the same order of magnitude as the down .._
loads, whereas in all other maneuvers the down loads
only are severe.
On the vertical’ surfaces only two loads listed excccd
the safe value, in the right barrel roll, Run 222, and in
the pull out of a dive, Run 226. Since the barrel roll
under discussion was quite severe and may be considered an unusual or test case, it would not be reasonable
to e.spect that under normal conditions vertical tail
surface loads would be so high. The load in the pull
out must be co~sidered a fair one, however, and the
possibility of the vertical surfaces receiving high loads
under normal conditions can not be denied.
In this
caso the maximum load occurs simultaneously with the
low incidence condition discussed previously in the
report.
Besides the average loading likely to be encountered
on td surfaces, the specifications must anticipate the
distribution of load with particular respect to the high
concentrations that may occur in some places, usually
the leading edges of stabilizer and fin. The present
specifications dispose the loa,d uniformly o~’er the fixed
surfaces whence itiecreases
untiI, at the trailing edge
of the ‘movable element, its value is one-third the wdue
over the fixed surface.
In addition, special leadingedge loads are applied extending from the leading edge
back one-fifth the chord of the tied surface, and having
a uniform vahle equaI ta three time9 the specified
Thus, the leading-edge load for a
average loading.
pursuit airplane stabilizer would be 135 pounds per
square foot and for the fin 120 pounds per square foot.
On the horizontal surfaces the maximum pressures
on the leading edge occur in the severe pall-ups and
exceed the specified leading edge value by a very
appreciable marginl although fortunately
they are
usually quite local in character and do not extend over
the specified area. The fact that they not only exceed
half the specified loading but actually exceed the whole
of it is well worth noting.
In all of these cases (pullups), however, the accderations
at the C. G. wero
In
the
less
severe
power+n puILups,
greater than 6.
the masimum prewres recorded are under] 35 pounds
per square foot, but in some cases considerably greater
than half that value, indicating that the specifications
should be revised upward.
In one dive, RurI 220,
the pressurewas equal to. that specified within the experimental error, md it would seem from this that to
double the present leading edge specifications would
give a much more reasonable value.
In no case does the pressure on the leading edge of
the fin exceed the specified value, although in the PUU __
out it reaches a value of 90 pounds per square foot,
.
-.
._
—
....
~..
._
,,...
PRESSURE
DISTRIBUTION
OF A PTV+
A few comparisons of some of the worst rib loads
with the specifications are gi~en in Fi=~es 93 and 94.
They are self-explanatory
and need no further comment.
Slip stream investigation.-A
number of level flight
runs and pull-ups, both power on and power off, were
made with oriflw located on sis ribs in the centd
portion of the upper wing, and one rib Dear the root
of each lower wing to determine the effect of the slip
stream on the pressure distribution.
Referring to the leveI fIight pressure plots (figs. 95a
to 95c), no pronounced ditlermce between right and
left sides can be observed in the slow-peed condition
on the upper wing, although an appreciable difference
on the Iower wing is apparent, the effect of the rotation
of the slip stream being, as wouId be expected, to reduce
the eflective angle of attack on the right side. This
effect on the lower wing occurs thoughout
the speed
range.
On the upper wing at the higher speeds, a
similar effect is noticeable, aIthough the differences are
not greater than might be expected es a result of the
deformed leading edge. The pressure plots for the
peak loads in pull-ups show no appreciable differences
in the character of the pressure distribution, with the
~~ception that the leading-edge pressure on the right
lower wing is lower than that of the left lower wing in
the power-on pulhps, whereas botk pressures are about
equal in the power-off pull-ups.
The span load curves show more clearly what differences exist. In Figures 99 and 100 it is seen that the
slip stream increases the load on the central portion of
the upper wing, while the shape of the load curve
remains much the same. The effect of the rotation of
the slip stream on the lower wing, however, is prcmmnced as the figures show.
Figures 101 and 102 show that no appreciable dissymmetry of load exists on either the upper or lower
wings at the peak loade of the puU-ups as a result of the
slip stream rotation, aIthough in the power-on pull-ups
the total load in the region fiected is greater than the
load in the power-off pulhups for the same initial air
speed. This increment of load is due to an increased
air speed, which is composed of both the sIip stream
increment and the natural increase obtaining as a
result of the power-on conditions.
From Figure 103, which represents the condition in
a dive at high speed, it is apparent from the lower wing
rib loads that the rotation of the slip stream is still
efkctive.
The low point at rib A’, which is also
apparent in the other span load curv=, is probably not
due to slip stream tiect nor to any interference from
the fuselage.
The pressure records for this rib show
violent fluctuations of pressure for the points aft of
about the quarter chord point.
This fact, in combina-
PURSUIT
MRPUXE
IW FLIGHT
765
tion ~th the knowledge that a rather abrupt discontinuity existed in the upper surface wing curve near the
leading edge of fib A’j would lead one to believe that the
streamline flow over this rib was disrupted, and hence
caused a marked decrease in the Mt. This is further
apparent from the pressure plots for the dive.
@ii.
98.) The conditicm may thus be considered as purely
locaI and having no connection with the slip stream.
Fabric pressures.-C!oincident
with the slip stream
measurements, records were also obtained of the fabric
pressurw on rib C and several other poin@. For this
purpose each pressure capsule was connected to one of
the flush ofices in the wing surface and to an open
tube terminating inside the wing near the flush oriilce.
Pressures as obtained me tabulated m Table V and
need no speciaI comment other than that minus signs
indicate downward acting pmsuree
and vice versa,
regardkss of whether the ofice is on the upper or
lower surface of the wing.
-=
.—
.-——
.--.—
..—
——
CONCLUSIONS
It is concluded from these tests that:
1. In any condition of flight characterized
by low
angles of attack with thg center of pressure well back
of the elastic ask, the wing, unless quite rigid, may
deflect torsionally to such an extent as to greatly aher
the load distribution.
This means that the load distribution is a function of the torsional rigidity of the
wing structure
and this fact should be taken into
account in the design rules.
2. Regardless of the cause, the effect of wing twist
on the load distribution
is the same and this effect
can be satisfactorily calculated.
3. The maximum forward position of the center of
pressure on upper and Iower wings is unatkcted
by
the factors involved in accelerated ~Uht..
4. The masimum forward position of the center of
pressure on the upper wing of the fulkxde
airplane is
the same as that for the model wing. This point h
questionable
with respect to the lower wing, since
strictIy comparable data are not available.
5. The mfium
normal force coefficient of the
upper wing of a biphme reaches an appreciably higher
value during a maneuver involving positive angular
velocity in pitch than it does in steady flight, while
lower wing normaI force coei3icient9 corresponding
to the mgle of attack at which the upper wing masimums occur are essentially the same for both pitching
and steady flight. This means that the relative wingIoading ratio for the true high angle of attack condition
(-which always involves a positive angular velocity in
pitch) is greater than would be expected from steady@ht (wind tunnel) data.
On the PIT”, this increase
is about 16.3 per cent of the steady-flight value.
..—
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766
REPORT
NATIONAL
ADVLSORY
COMMITTEE
60
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AERONAUTICS
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E
okmg
4
.
PRESSURE
DISTRIBUTION
OF A PW-9
6. In power-on maneuvers involving high angle of
attack, a varying inertia load exists aIong the fuselege
which is critical for the engine support.
This inertia
load should be considered in the fuselage analysis conditions.
7. The strip method of computing span-load curves
for the high angie of attack condition gives results which
check the measured load distribution very satisfactory
for wings tapering similarly to those of the PI?’-9.
S. The position of the center of pressure at sectiorw
near the tip is an important factor requiring further
study.
In the present case on the upper wing the
center of pressure locus bends to the rear at the tip,
thus increasing the front spar bending moment in the
bay for the high m@e of attack condition by an
amount which is not provided for by the present rules
concerning tip 10ss.
9. Design rules should include proper determination
of the center of pressure position on biplanes.
10. In barrel rolls the maximum dkymrnetry
of
load occurs shortly after the peak load, and the total
load when this maximum value occurs is not appreciably leas than the peak.
11. l’i’km the maximum dissymmet~
of load occurs
in a rolI, the lower wing is not shdled and the asymxnetricaI load@ on this wing is of such a nature as to
oppose the rolling of the airphuie.
12. In a spin the imide wings are stalled approximately to the plane of symmetry, while the outside
wings remain unstalled, aIthough at a high angle of
attack.
13. Leading-edge pressures on the wings reach values
of the order of 450 pounds per square foot, both in
dives and in pulI-ups.
14. Ccmditions governing the critical rear spar loads
shouId be studied at greater length.
The present
report shows that in normally executed pull-outs from
fast dives, rear spar loads may be greater than have
heretofore been considered.
15. Tad-1oad specifications should be revised upward.
This is particularly @e of leading-edge loads,
which should be at least doubled.
16. The effect of the dip stream is to cause dissymmetry of load on the two root portions of the lomr
wing, and to increase the ~oad on the central portion of
the upper wing without giwing rise to any dissymmetry
thereon.
17. Abrupt changes iD contour of the wing curvature
near the leading edge may seriody
affect the lift.. .
PURSUIT
AMPLANB
IX
FLIGHT
BIBLIOGRAPHY
Reference 1: Handbcmk of Instructions
for Airplane De&gners.
Engineering Division, Amny Air Service (1925).
Reference 2: GeneraI Spaaiticatione for the Design of Airplan&
for the United States Navy-SIM+B.
Bureau of Aerunautios, Navy Department.
Reference 3: Airworthiness
R8quiremente
of Air Commerce
Aeronautics
Btietin
No. 7–A. Aeronaut ica
Regulations.
Branch, Department
of Commerce.
Reference 4: Norton, F. H.: Pressure Distribution Over the
Wings ofan ME-3 Airplane in FIight. N. A. C. A. Teefsnical Repart No. 193 (1924).
Referenoe 5: CkowIey, jr., J. W.: Pressure D~tribution Over
in Flight.
a Whg and Tail Rib of a VE-7 and of a 7?S Airplane
N. A. C. A. Teshnical Report No. 257 (1927).
Reference 6: Norton, F. H.: N. A. C. A. Recording Air-Speed
Meter.
N. A. C. A. Teo@ical Note No. 64 (1921).
Reference 7: Reid, H. J. E.: A Study of AirpIane Maneuvers
with Speekd Reference to Ang&m Velocities.
N. A. C. A.
I&ok&al
Report No. 155 (1922).
Reference 8: Norton, F. H.: N. A. C. A. Control Pc&ion
Reeorder.
N. A. C. A. Technical Note No. 97 (1922].
Reference 9: Norton, F. H., and ‘iYarner, E. P.: AcceIerameter
Design.
N. A. C. A. TechnicaI Report No. 100 (1921).
Reference 10: Coleman, D. G.: N. A. C. A. Flight-path-angle
and Air-Speed Recorder.
N. A. C. A. Teohnical Note >-o.
233 (192tl) .
Reference 11: Rhode, R. V.: The Premwe Distribution
O\-er
the Horizontal and Vertic~ Tail Surfaces of the FtiC-4 Pursuit Airplane in Vrolent >laneu~ers.
N. A. C. A. Technical
Repart No. 307 (1!329).
Referenoe 12: Howard, H. B.: Some ProbIems in Aeroplane
%ructural
Design.
The Jaurnal of the Rayal Aeronautical
%ciety, ApriI, 1926.
of Strength Calculations, Air 31inReferenoe 13: Handbook
iatry. Air Publication 970 (His Majesty’s Stationery Office),
192&
Reference 14: Laeser, jr., O. E.: Pressure IMributian
Tests
on PW-9 Wing 3fadels from —lSO thraugh 90° Angle of
Attack. N. A. C. A. Technical Report No. 296 (1929).
TABLE
L-CHARACTERISTICS
OF PW-9
‘~pm:;$~m::&&;~T&zti-H_--:
Cantrfc Chm’d of lower W&-—-–
DfskimM llmn cedar Ik to eenfrfc_&xd---—.—
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Dihedral UQW W@ lower surface) --=
~ih&mf ~n=WhE
Iowermrface).———————
- d’%l&
adgeofmot sectfo&loWerWfnG
Abre fowar surface*
wtfoq Iowerwfng.
Distance from C. G. to centerMe of eyevntm
--------—Dfstanca horn C. Ct. to cmti Hoe of rudder
——
------------–—
&ea of ripperwfnL-__.-–—.———_—
&aaof Iow&wfna-..
Totaf wfu area .---_AreaofhdrontnI Mmrfars_ ———----—
k
of WMC91Ml SnllkxS—–.–––.
Wfght of akpkmednrfngte&._.._._R8ted hmapmrer at Znxl r. p. m-.--–––
hlEMORIW
NATIONAL
AERONAUTICAL
ADVISORY
COMMITTEE
LABORATORY,
FOR
AERO-
NAUTICS,
LANGLEY
4163c-31~
FIELD,
VA.,
Februa~ 3, 1950.
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OF
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Le’ial flight at 74 m. p. h. q~14 ~unds
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PRESSURE
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REPORT
NATIONAL
TABLE
ADVISORY
COMMITTEE
111.-RECORDED
FOR AERONAUTICS
.-.
PRESSURW&Continucd
-
No. 85. The:
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111.—RECORDED
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Rnn No. EM. Time: 1.73ermnde
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Rnn No. 130. TM
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REPORT
NATIONAL
TABLE
ADVISORY
FOR AERONAUTICS
COl@l!Fl?EEl
111.-RECORDED
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Rim No. 187, Tfme: 0.76second
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