A simple method for estimating terminal velocity including effect of compressibility on drag

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Title:
A simple method for estimating terminal velocity including effect of compressibility on drag
Alternate Title:
NACA wartime reports
Physical Description:
12, 14 p. : ill. ; 28 cm.
Language:
English
Creator:
Bielat, Ralph P
Langley Aeronautical Laboratory
United States -- National Advisory Committee for Aeronautics
Publisher:
Langley Memorial Aeronautical Laboratory
Place of Publication:
Langley Field, VA
Publication Date:

Subjects

Subjects / Keywords:
Compressibility   ( lcsh )
Aerodynamics -- Research   ( lcsh )
Genre:
federal government publication   ( marcgt )
bibliography   ( marcgt )
technical report   ( marcgt )
non-fiction   ( marcgt )

Notes

Summary:
Summary: A generalized drag curve that provides an estimate for the drag rise due to compressibility has been obtained for an analysis of wind-tunnel data of several airfoils, fuselages, nacelles, and windshields at speeds up to and above the wing critical speed. The airfoils analyzed had little or no sweepback and effective aspect ratios above 6.5. A chart based on the generalized drag curve is presented from which the terminal velocity of a conventional airplane that employs a wing of moderate aspect ratio and very little sweepback in a vertical dive may be rapidly estimated. In order to use the chart, the only data that need be known about the airplane are a low-speed drag coefficient, the wing critical speed, and the wing loading. The terminal velocities for three airplanes were computed in order to illustrate the use of the method and chart. Good agreement between the estimated terminal velocity and the measured flight terminal velocity was indicated for all three airplanes.
Bibliography:
Includes bibliographic references (p. 12).
Statement of Responsibility:
by Ralph P. Bielat.
General Note:
"Report no. L-78."
General Note:
"Originally issued August 1945 as Advance Confidential Report L5G31."
General Note:
"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 previously held under a security status but are now unclassified. Some of these reports were not technically edited. All have been reproduced without change in order to expedite general distribution."

Record Information

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University of Florida
Rights Management:
All applicable rights reserved by the source institution and holding location.
Resource Identifier:
aleph - 003638273
oclc - 71826762
sobekcm - AA00006252_00001
System ID:
AA00006252:00001

Full Text
SAcACP N LL-73
ACE No. L5G31


NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS


WAR'RTIMi RI REPORT
ORIGINALLY ISSUED
August 1945 as
Advance Confidential Report L5G31

A SIMPLE METHOD FOR ESTIMATING TERMINAL VELOCITY
INCLUDING EFFECT OF COMPRESSIBILITY ON WRAG
By Ralph P. Bielat

Langley Memorial Aeronautical Laboratory
Langley Field, Va.


NACA


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-
i 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 78


DOCUMENTS DEPARTMENT






































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/simplemethodfore001ang







TIACA ACR lio, L5G41

NATIONAL ADVISORY CCVMITTEE FOR AEt.O"AUTICS


,DVUIICE COi.PTIDEL!TIAL rFP'rT

A SISLE ITSOD FOR .SI..*TI' TG T'RITlL VELOCITY

I!' LTDIT'iG EF EC'T -:F COYi PRL- SIBLr TY -' DR .G

By Ralph P. Bielat




.. gner.lid .iJr'g curve that provides an estimate
for the rag rise jne to c-:p'es sibiity nas been obtained
front an inal,:L s 'o .,inJ-:ur.nnel data ..f several airfoils,
fusela es3, nacellE:s. And ,Ird.idsia"e.l r-" at :;c-ds up to and
above tihe win,.- critics.] need, hd air'fo'ils analyzed had
little or no s.-eei:o-c: an,:' effective aspectt ratios above 6.5.
A chart based on the ,.erneralizd d-: C. curve is presented.
fro., whichh the ter...nal vrlczity :f conv-entional airplane
that employs a w'.n. 3f e: t &..suct ratio and very
little s'.eepback' in vertical (dive ..y 0. rapidly esti-
mated. In -rder t' uce thie ch-rt, che )nly data that
ne.d bE .:novr abcut th.e airlate are a lo.i-'oe.'d drag
coefficient, t.ie v.ing critical seed, an, tlhe wing loading.
The terriinal velocities for three airp'lcinc were computed
in rider r to illustrate th.1 use of the method and chart.
Good agreermnrt between the et.lrriated ttriAinal velocity and
th-e -.easured flight terminal v-locity was indicated for
all three airplanes.


Ii TRO- UCT IO


Several hl,.h-sneec militar:.- air'l,.anes in lives s have
encountered difricultie- that could not oe easily con-
trolled by normal means. These Cdifficult es, whi ch may
consist -of giving znor..ents. .arge charges in trim, large
stick forces, tail buffetinc, -in, the lile, occur in high-
soeed dives '.'when the sr.eed of the airplane exceeds the
critical .speed by a large a:o'..Int. .,or th-.se airpl-.nes for
which i,.aximu-n diving speeds are at or near the critical
speed, little or no trouble occurs. The more recent
fighter airplanes, i:,'ever, have Lerr,-inal r'ach numbers
well in excess of the critical ;.lrch number and, as a
result, often encounter difficulties in dives. Determi-
nation of the terminal velocity of the airplane is there-
fore import-.nt in order that thet probability of encoun-
tering trouble in divts ;iay be esti:;.a-e-.









CONFTDEII'TAL NACA ACR No. L5G51


The terminal velocity is also im-po rtant because it
forms the outer limits of the V-G diagram. Usually the
outer limit of the V-G diagram is established by multi-
plying the maximum level-flight speed of an airplane by
an arbitrary factor somewhat greater than 1.0. The termi-
nal velocity of most recent airplanes, however, generallyy
falls much below this arbitrary Miaximum speed, and these
airplanes are therefore unnecessarily penalized by extra
weight because they are designed for conditions that are
not reached in actual flight.

The present report outlines a si'inle method for
obtaining the terminal velocity of an airplane in a verti-
cal dive and includes an estimate for the drag increase
due to compressibility effects. The wind-tunnel test data
were obtained from model tests conducted in the Langley
2L-inch and 8-foot high-speed tunnels. All the data
presented herein .;ere obtained for zero lift.

The problem of determ'rnirn the terminal velocity for
airplanes for which the terminal velocity is near the
critical sDeed is c:-aratively simple inasmuch as a
constant value of drag coefficient can be assumed. The
diving speeds of most present-day airplanes, however,
occur beyond the critical speed ond the problem is'not so
simple. The following two factors are involved: (1) the
determination of the critical speed and (2) the rate-of '
drag increase at soee-z above the critical speed.

The critical speed used herein was arbitrarily'taken
as the critical speed of the wing-root section. Pressure-
distribution data obtained from wind-tunnel tests-were
used to determine the critical seed, which is defined as
the flight speed at which sonic velocity is reached
locally. If experimental data are not available,. however,-
the methods outlined in references 1 and 2 can be used
for the determination of the critical speed. Selection.
of the critical speed at the wing-root section for use in
terminal-velocity estimation is justified on the. grounds.
that the root section usually has a lower critical s-'eed
than any other component cart of the airplane. -The j.iiag.
root has the lowest critical speed because of .its-high
thickness ratio an contributes a large part of the total
airplane drag because of the large fraction of the wing_
area concentrated at the inboard sections .of tapered wings.

The rate of drag increase at speeds above the. cliti-
cal speed is more difficult to determine in the calculations


C CO- ID 'TIAL








NACA ACR No. L5G351


of the terminal velocity than the critical speed. A
study of the drag of airfoils, fuselages, nacelles, and
windshields has been ma-e fro.n .,ind-tunnel test data in
order to determine che effects of c-mxiressibility on the
drag. Because the rate of draw inc-'e-se at spee-s above
the critical seed is so great, it ..a fond that, within
the accuracy required for terminal-velocity calculations,
an average rate of' drag increase O'.ay be used. .4 curve
indicating the average rate of drag incre se is presented
herein. This ci'.rve wvas derived from an analysis of wind-
tunnel data.

The method described herein for obtaining the termi-
nal velocity of an airplane in a vertical dive has been
in use at the NACA since 1901. Publication of the method,
however, had been delayed pending. the investigation of
constriction corrections to be applied to ..he wind-tunnel
data and the com'plcti'n of high-sprecd five tests made
with several airplanes in order to compare term'Linal
velocities obtained in flight oith terminal velocities
estimated by the simn.le :rethod described herein. This
method is not applicable to airplanes that utilize wing
shapes of low aspect ratio and large sweepback but should
be applied only to airplanes of conventional design that
employ wing shapes of moderate aspect ratio and small
amounts of seaspback due to wing taper rat.io.

S -L BC L3

V velocity
a speed of sound in air

MI Mach numbEr (V/a)
CD drag coefficient

CL lift coefficient

p mass density of air
S wing area
"'V weight of airplane

p atImosoheric pressure at any altitude


COTrFIDENTIAL


C NFIDEI'TIAL








C :'C. ACR No. L5G51


ratio of specific heats (1.40 for air)
t/c ratio of thickness to chord of wing
, scriptst:
or critical (when local sonic valocit- has been
reached on some point of body)
m in minimum
T terminal

DECD.IPTIOT! OF YODELS

Airfoil models.- The airfoil models used herein
represent two classes of airfoils namely, the conven-
tional NACA sections and the more recent low-drag higih-
critical-speed NACA sections. The conventional .'.CA air-
foil sections are characterized by pressure distributions
that have high peak pressures occurring near the leading
edge. The low-a:-g NACA airfoil sections have comoiara-
tively flat pressure distributions iith the peak pressures
occurring at approximately 60 percent of the chord behind
the leading edge.
Airfoils typical of the conventional airfoils are
the NACA 0009-63, 0012-63, 25012, and 23014.7 sections;
a current transport-model airfoil that has an i.iA 2215
section at the root and tapers to an TACA 2212 section at
the tip; and the Davis airfoil with a thickness ratio of
20.15 percent. The low-dr.g airfoils include the following
NACA airfoil sections!
t
16-215 65-type modified, = 0.156
16-50Q 66, 1-115
16-515 67-111 5
1.7-215 67 0-215
65(21,8)-220
The effective as,)."ct ratio of the airfoil models
tested varied from 6.5 to infinity.
Fuselage models.- T'-.e fusel-e models are t-.ical of
fusel7s7X:i ", i se on current airplanes. T.,e various
fusel,'es rec' :-sent bomber, fijhter, and trans-.,;rt air-
planes. Figure 1 shovws the side-view draw"inf n," the
fineness ratio in si-r; elevation of the different fuselage
sha,'-r-s. These fusel';g!-- models were tested in conjunction
with win.Ls (shown as dashed lines in fl.1. 1) and represent
a wide variation in winm-fuselj,. interference.


-': I 'T"IL


CCNFLE'TIAL








HACA ACR No. L5.331 CONFIDENTIAL 5


iacelle and windshield models.- The data for the
various nacelles and windshields were obtained from refer-
ences 5 and 4, respectively. The nacelle and windshield
designations used herein correzoond t) the designations
used in references 5 and 4. All the nacelle models were
tested with the same wing model, which consisted of the
outboard panel of a wing section designed for use on a
bomber airplane. The ,,1ing was a thick low-drag airfoil
that had an .ACA. 65(21E)-221 section at the root and
tapered to an NACa 66(,215)-146 section at the t.p. The
windshields were tested with a w*.ing-fuselage combination.
Drawings of the nacelle and windshield models are shown
in figures 2 and 5, respectivel.7.


RES'TLTS AD DISCl'ZsTljr

Dra; Characteristi cs


Drag analysis.- In order to ottain a correlation of
the rate of drag increase at speec.s abooe the critical
speed, the drag results for the various component parts
of the airplane have been reduced to nondimensional carwrn:-
ete-r--- that is, CD/ Cmin is plotted against M/I r
for each part tested. The use of th-se parameters
represents a convenient method of ina;ring the data non-
dimensional in such a manner that the unknown quantities
are expressed in terms of the known quantities.

The drag results at speeds up to and above the criti-
cal seed for the conventional ULKCA airfoils are presented
in figures 1 and c.. Figures 6 to 9 show the variation of
CD/Cin :ith M/Mcr for the low-drag high-critical-
spseel airfoils. It will be noted that all the airfoils
presented in figures 4, 6, 7, and 8 cxhioited approxi-
nately the sa.ne rate of drag increase at speeds above the
critical speed* for this reason a curve of the average
rate of drag increase at speeds above the critical speed maybe
used. An average increase in drag of approximately 50 per-
cent above the minimum drag was indicated at 1.0;
Mcr
at speeds of only 10 to 15 percent above the critical
speed, however, the drag increased approximately 90
to 200 percent. This rapid increase in drag at speeds
above the critical speed is associated with the formation


CCo:FIDENT IAL








TUACA ACR No. L5G31


of compression shock waves and their effect on the boundary
layer over the surface of the airfoils. The family of
airfoils used in figure 5 showed less percentage of
increase in drag at the critical speed than the NACA 0009,
0012, or the low-drag high-critical-speed airfoil sections.
Both published (reference 5) and unpublished high-speed
data show that the NACA 230-series airfoils differ from
most of the other airfoils in that the critical speed can
be exceeded by as much as 0.15 in I.ach number before any
serious changes in the aerodynamiic characteristics of the
airfoil occur. The critical speed of the NACA 250-series
airfoils is therefore exceeded by approximately 71 percent
before the same percentage of increase in drag occurs as
is shown for the other airfoils. The importance of this
difference in the shape of the ir:-b curve above the criti-
cal speed on the estimation of terminal velocity is dis-
cussed in the section entitled "Terminal-Velocity Calcula-
tion."

Tie rapid increase in drag before the critical speed
is reached, which is shown for the NACA 67-114.5 airfoil
in figure 9, is due to early separation of the flow over
the after portion of the airfoil. This condition also
affects the method for estimating the terminal velocity.
An error in estimating the terminal velocity when the
flow separates will occur only for those airplanes for
which t-.r-ninal velocities are at or near the critical
speed; this separation of flow will not appreciably affect
the determination of the terminal velocity for high-
performance airplanes for which the terminal velocity
occurs at speeds well above the critical speed.

Figure 10 shows the variation of CD/CD.in with

M/Mcr for several fuselage shapes and fineness ratios.
The drag increments for the nacelles and windshields are
presented in figures 11 and 12, respectively. The criti-
cal s:e .s for these bodies were based on the wings with
which the '.-dels were tested and were determined for the
wing-root juncture. TheL effect of compressibility on the
rate of drag increase at sp-.ds above the wing critical
speed for these bodies is similar to that for the air-
foils,

In the correlation of the avera,; drag increases of
the various co- onrenLs of the airplane through-ut the
Mach number range, a generalizedd drag curve was derived


CC "T'DENTIAL


COF IDL'TIAL









HACA ACR o0. L5G51


and is presented in figure 13. The data presented in
figures ., 6, 7, C, 10, 11, and 12 wer-e used to obtain
the generalized drag curve. The generalized drag curve
is an average of the drag data for the airfoils, fuselages,
nacelles, and windshields at speeds up to 10 percent
above the critical speed. Only the average drag of the
airfoils at speeds from 10 to 15 percent above the criti-
cal speed was used. The generalized drag curve was extra-
polated by use of a straight-line extrapolation from 15
to 25 percent above the critical speed. The straight-
line extrapolation is believed to be sufficiently accurate
for estimation of the terminal velocity in this region
where the drag rises rapidly due to compressibility
effects.

Constriction correctjons.- Corrections for constriction
effects have been applied to the data. The constriction
corrections have been determined from pressure measure-
ments obtained in the Langley 24-inch and 8-foot high-
speed tunnels on NhCA 0012 airfoil models of various
sizes. The magnitude of the corrections applied to the
drag coefficients amounted to less uhan one-half of 1 per-
cent of the dyna:,ic pressure q at low speec.s and
increased to approximately 2 percent of q at the criti-
cal speeds and to approximately 5 percent of q at a
value of the Mach number below the choking speed of the
tunnel. The corrections to the Mach numbers amounted to
approximately one-half of these values. The constriction
corrections were such that the coefficients were reduced
and the Mach numbers were increased by the values stated.
The greatest percentage of increase in correction, as
would be expected. occurred for the models that had the
largest ratio of model area to tunnel area.

Comparison with flight data.- Figure 1i shows the
variation of over-all drag coefficient with Miach number
for the XP-51 airplane as measured in flight and the
variation with Mach number of the wing-profile drag at
the rrid-semispan station measured by the wake-survey
method. These flight data are preliminary as corrections
to the data have not been applied. The results obtained
by use of the generalized drag curve in estimating the
drag increases with Mach number are also shown in fig-
ure 14 for comparison with the flight measurements of
over-all drag and wing-profile-drag data of the XP-51 air-
plane. The curves for the wing-profile drag and the over-
all airplane drag in flight begin to rise rather steeply


CONFIDENTIAL


CONFIDENTTIAL









NACA ACR No. L5G31


at about the same Mach number. This fact tends to justify
the assumption that the wing-root critical speed is a
suitable criterion to use in terminal-velocity calcula-
tions. The estimated drag derived from the generalized
drag relation indicates higher drug coefficients at Mach
numbers of a0pro;imately 0.55 to 0.75 than are shown for
both the measured wing-profile drag and the over-all drag
coefficients. Of more importance, however, is the good
agreement that is shown for the values obtained by use of
the gen-ralized drag curve and the measured flight data
at Mach numbers greater than 0.75, which is the region
where the terminal Mach number usually occurs.

Figure 15 shows a comparison of measured flight drag
and estimated drag for the FkFi-2 airplane of reference 6.
An important difference in the drag curves occurs at Miach
numbers around the critical Mach number. The estimated.
drag indicates lower drag coefficients than do the flight
measurements. This difference is believed to be due to a
combination of early shock formation on the cowling and
airplane-wing roi;-ness, which is believed to have caused
some separation of the flow. Good arrerment is indicated
between the fli-ght measurements and the estimated fra., in
the region where the drag coefficients rise steeply, which
is the region that determines the terminal iTch number.


Terminal-Velocity Calculation

The generalized drag curve (fig. 15) may be used as
an approximation in determining the terminal velocity of
an airplane in a vertical dive. 'Te terminal velocity is
reached when the drag of the airplane is equal to the
weight of the airplane. The drag of the airplane in a
dive combines both airplane and propeller characteristics.
Tn the present analysis, however, zero prrpller thrust
is asrunied and the ,ropeller drag or thrust is therefore
neglected. At supercritical speeds the drag or thrust
caused by the propeller is considered to be negliAible as
compared with the c.raZ of the airlane, particularly if
the pilot throttles the engine and adjusts the pri.peller
to a high blade-angle position.

The tmrn.inral velocity for zero-lift conditions is
given by the relation


CON IDEiTIAL


CONPIDE!UTIAL








NACA ACR D LC5G51 CONFIDENTIAL 9


V=V (1)
1 3SP CD

or, in terms of the terminal Mach number Mirp with the
speed of sound equal to ,


VT 2 1 1(2)
a QS p C3

Equation (2) can be rewritten in the f-rm

2. = h 2 1 1 1
S y P CD/Ci-

12
"Y -'min CD/CD.uin


where is constant for each airplane at the
PSCDmin
altitude for which MT is calculated and


CD K (MT '

C^nin \ cr

which is obtainable from the generalized drag curve.
Equation (3) can then be solved for the paramr.eter
V9 -. Figure 16 sho31s soluti,-'ns of equation (3) for
pSCDmin
various values of M-'cr and ,.


In computations of the terminal velocity, the only
data that must be known about the airplane are uhe minimum


CONFIDENTIAL








NACA ACR No. L5G51


drag coefficient at zero (or a-,prox.riately zero) lift
coefficient or a low-sreed drag coefficient whereby the
minimum drag coefficient can be computed by use of the
generalized drag curve, the ,v.ig critical speed, and the
wi.i.; loading. Values of these quantities, all of which
are used in calculations other than those for the terminal
velocity, are easily obtained. With these values known
for a ..drticular airplane, the parameter can be
ZDmin
calculated for different altitudes; then, for given values
W
of Mcr and the terminal Mach number can be

obtained by use of figure 16.

In order to illustrate the method of obtaiirdni the
terminal velocity ..i Lcally, the te-rminal velocities
have been calculate'. i'or the ':2.-.-2, P--3T-l, and P-47 air-
planes. The pertL ent data for these airpha *.--s are given
in the following Lable:


TABLE I

AiF:FLA:TB DATA


CD
Airplane Mcr min


XF2A-2 o0.61 (flight) 0.022 (fli:.ht)
.66 (corrected)

P-39N-17 0.675 (estimated) I0.019 (estimated)

P-7 GO.'!,. (wind tunnel) 0.020 (fliht)
.69 (corrected)


(Ib/sq ft)

26.1


5k.1

h5.o
L._____


By use of these -ita th,- parameter is
Dmin
co,.' -uted for each airplane. The use of figure 16 to
esr.4mate the terminal Mach number is illustrated for the
P-47 airplane at 15,000 feet altitude. The variation of


C O fIT-' :TIAL


CO;PIDEi;TIAL








NACA ACR No. L5G51


terminal Mach number with altitude thus obtained for the
three airplanes is presented in figure 17.

Also included in figure 17, for comparison with the
estimated variation of terminal M'ach number with altitude,
are records of flight data for the XF2A-2, the P-47,
the P-47C-1-RE, and the P-59N-1 airplanes. The flight
record for the P-47C-1-RE airplane -.-;as obtained by the
late Major Perry Ritchie in a terminal-velocity dive made
at Wright Field in July 1943. The points represented by
circles were obtained from a dive of a P-47 airplane made
by a test pilot for the Republic Aviation Corporation.
Unfortunately, a complete dive history is not available
for this dive but it is believed that, had one been
available, it would have followed a path similar to that
obtained by the late Wajor Ritchie for the P-47C-1-RE air-
plane. It is further believed that the test points
obtained at altitudes of 22,000 feet and 10,000 feet
represent entry into and pull-out from the dive, respec-
tively. Data for the XF2A-2 airplane were obtained from
reference 6 and the data for the P-359N-1 were obtained
from dive tests made at Aries Aeronautical Laboratory. The
present method for estimating the terminal Mach number
yields results that compare favorably with the flight
measurements; the difference between the two is no greater
than 0.02 in Mach number. This method for estimating the
terminal Mach number is therefore believed to be suffi-
ciently accurate for usual engineering purposes.

The section entitled "Drag Characteristics" indicates
that the NACA 250-series airfoils and airfoils similar to
the NACA 250-series could exceed the critical speed by
approximately 0.05 to 0.15 in Mach number before any
important changes in the aerodynamic characteristics
occurred. At =-L 1.0, therefore, the NACA 250-series
cr
airfoils and similar airfoils did not show the same
percentage increase in drag as was shown for almost all
the other airfoils and for the generalized drag curve.
Since in the calculation of the terminal velocity the
critical speed of the airplane is based on the critical
speed of the wing, it can be expected that for airplanes
utilizing NACA 250-series airfoil sections or similar
sections the estimation of the terminal velocity will be
in error. If the generalized drag curve is used in the
estimation of the terminal velocity, the indicated wing
1
critical speed must be increased approximately 72 percent


CONFIDENTIAL


CONFIDENTIAL









NACA ACR No. L5G51


for the NACA 250-series sections. This correction was
applied to the critical speeds of the P-47 and E.F2.--2 air-
planes (see table I), since these airplanes have NACA 250-
series sections. The dashed curve on figure 17 for
M*cr = 0.64 is the result obtained if the indicated criti-
cal Uach number is used rather than the effective critical
1
Mach number, which is about T, percent higher.


Langley Memorial Aeronautical Laboratory
national Advisory Committee for Aeronautics
Langley Field, Va.




,R LFE."CES


1. Robinson, Russell G., and ..ight, Ray H.: Estimation
of Critical Speeds of Airfoils and Streamline
Bodies. LACA ACR, March 1940.

2. Heaslet, >.;.. A.: Critical M'ach iuTn.bers of Various
Airfoil Sections. ,iA.CA ACR No. 4G18, 1944.

5. -3ecker, John V.: High-Speed. Tsts of Radial-Engine
"acelles on a Thick Low-Drr..- ing. 1,ACA ACR, May
19LJ2.

4. Delano, Janes B., and Vri.?ht, Ray H.: Investigation
of Drag and Pressure Distribution of ..i'ishields at
Hi9. Speeds. NACA ARR, Jan. 1942.

5. Becker, John V.: High-Speed ..ii L-Tunnel Tests of the
-A.A 27.12 and :-'012-64 Airfoils. -isC? ACR, Feb.
1941.

6. Rhode, Richard V., and Pearson, H. A.: Observations
of C: -or-ssibility Y-hZr',,ena in Fli ht. NtC, e-CR
7.. 3D15, 194).


CO1-'I DENTAL


CONFIDENTIAL






NACA ACR No. L5G31


CONFIDENTIAL




F-iretness ra io, 10.32





F'ne ess raLio, 6.65




Pr'eress raiio 7.30





Fireress rahio 6.14





Firieness r*io 5.60



F~ineress ratio 4.75

NATIONAL ADVISORY
C) COMMITTEE FOI AElRONUtniCS
CONFIDENTIAL
Fineres5 rafio 3.30
Figure / .- 5ide -view drawirqng of various
fuselage shape.


Fig. la-g





NACA ACR No. L5G31


CONFIDENTIAL


Circular cross section




Nacelle I


Elliptical cross section

4^E ^>


Nacelle


i


Circular cross


Nacelle 5


Circular cross Secfion




Nacelle 4


section


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS


CONFIDENTIAL


Figure Z.- Nacelles
( Des 5 ignatiors From


tested.
reference e


Plan-view profile


Nacelle


3.)


Fig. 2






NACA ACR No. L5G31


Windshield
3-1-1










2-0-3


CONFIDENTIAL A

^biWA




K2IL"~hj>


4-0-3


NATIONAL ADliSOAV
COMMITTEE FO AmIONWATICS


X-I


A
Plane A-A of all windshields is located
at Same posii-ion on fJs5elaqe .


Figure 3.- Windshields tested.
(Designoationrs fror reFerence 4.)
CONFIDENTIAL


Fig. 3







NACA ACR No. L5G31


Q
C-)
Q
C)


CONFIDENTIAL



68---------------------------------------------*--



Airfroil sech ons
4/'ACA 0009-62
0009-63
S 0009-64
---- 0009-65
------------
00--------6 -----


.8 - -







-- -- --- --- -- -- ?-
.8 i





!.4






f. NATIONAL ADVISORY
CONFIDENTIAL COnNITTIE For "AIITIrcs


/./ /.Z.


Fi'gure 4. Variab/on of ra/Ao Co/Com,Vn wi/h M/^lcr
for A//4CA 0009 anod 0012 Q0rf-il6.


Fig. 4


.3 .4 .5 .6 .7 .-
A*Adcr







NACA ACR No. L5G31


18



1.6



t-


1.O


1.0


Figure 5. Variaf-ion of rcf io CD/CoD,. w/ih AM/MAcr
For threee Conventinal/ AI/ACA cirO/'r/s.


/.6





Q0
U



1.0


.4 .5


.7 .8 .8 /.0 // /
AM/A4cr


Figure 6. Varia/iori of roaio COCo/C,, wifLh A4M/4cr
for NACIA 16-509 anrd 16-515 Q rfoi/5. TrQisi'/i/on
fixed a' 0.10 chord.


CONFIDENTIAL



Airfoil sec- -ion s
-/ACA 23012
--- 230/4.7
---- 2l(root) -o 22/2 (4)







NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS
........ I l l i


Airfoil sec t'os
NACA /6-509
----- 16-515
/






NATIONAL ADVISORY
CONFIDENTIAL COMMITTEE FOR AERONAUTICS


Figs. 5,6


.7 .8
M/IA--r


.9 10







Fig. 7a,b


18--


16--


14 -


Sis --


NACA ACR No. L5G31


(a) Trancidiion fixed oj 0.10 chord.


Airfoil section
NACA 66, /-//5
47-2/5
67,0-2/5
/6-2/5


16








1.0
to


3 .4 .5 .6 .7 .8

(b) Transi/tion f xed ai 0.40 chord.
Figure 7 .- Vcar/a f/on oAf roa//o Co/Corn
for severe/ NACA oi/rfoi/ls .


1.0 l 12.


With A t/41"cr


NATIONAL ADVISORY
CONFIDENTIAL CObMITTEE Imf AEAU TIKS


*.







NACA ACR No. L5G31


-b I CONFIDENTIAL -
Airfoil sec+iorins
4 NACA 47-215
---- 67,0-2/5
-/6-2/5




2.-0_---------


18-----------


1.6- -- --



/
14----------------------------





10 NATIONAL ADVISORY
S___ CONF IDEN TIAL ____ COMMITTEE FOR AEROMUTICS


.3 .4 .5 .6 .7 .8
A/Amcr


Figure 8 .- Var/atioti of rafio CD/Co,, with A1/A4lcr
for several NACA airroils No "frao ils>ioor).


.9 10


/.1 l/


C

0


Fig. 8







NACA ACR No. L5G31


.c


Q
C)


-I I CONFIDENTIAL




Airfoil sections
NACA 65(218)-220
65-fype raodfied, c =0./ 96
-- ---- D iS oil, t/c =0.20/5
S ---- 67-114.5 ,


2.0 --------------------------------LL





--------------77-


1------------------------ --- t->- --
1.


1 /


1.2-- 1 / 4 1
-- --------__ -/---------

NATIONAL ADVISORY
S CONFIDENTIAL COMMITTEE Iro AEIouAwTICS


.5 .6 .7 .8
AM/MAcr


.9 /.t


1.1 12


FicuLre 9 .- Variafiorw of ratio C/Ci wl Ah M/A4cr
for Davis airfoil and several NACA airfoils.


Fig. 9







NACA ACR No. L5G31


CONFIDENTIAL
2. 2 Fuselage Fireness
(See /ig. I.) ratio
(0) 10.32_
-- ( b) 6.65 /
------ ) 730
(d) 6. 14
------ e) S.60
-------- f) 4.75
12-- ------- -3.30 -









10



NATIONAL. ADVISORY
CONFIDENTIAL COMMITTEE FOl AERONAUTIcS


.3 .4 .5


.6


7 c
M2/A4cr


.9 1O 1/1 12


FigLdure /0. Variafion of ratio C/CD wi/h A^4/Mcr
for several Fuselage .sapes.


C
0

0
U


Fig. 10







NACA ACR No. L5G31


" CONFIDENTIAL

Nacelle (See 2g. 2)
16 -I-2
-~2
-~3
-------4 /
4




1.4-0



NATIONAL ADVISORY
COMMITTEE FOR AEMONaITIC-S


.3 .4 .5 .6 .7 .8
A/-IAcr


.9 1.0 1./ 1.2


Figure // Varic-iori of rafio C/Comi, wHih ,///Vlcr
for several nacell//e sAapes.


.3 .4 .5 .6 .7 .05
MA4f/ Cr


.9 10 i. /Z


Fiqcgre /12 .- Varia-ifon of raf-io CD/Co, wvith /V/A1cr
for several wincidshiel/c shapes.


Q
(~1
Q
0


W/i 'shie/dc (See fg7.3)
3-1-1
---- 3-1-2I
------ 2-0-3
-- -- 4-0-3
X-I




-------7--



-------------------------------------------^^ _--- -

NATIONAL ADVISORY
CONFIDENTIAL COMMITTEE fM AEtOIAUTICS


Figs. 11,12


A








NACA ACR No. L5G31


0
U
0
C-)


CONFIDENTIAL













4I2
5.4-------------------------------------,- -




















24 .
5.0---------------------
2.--------- --- ----------------- ---------- -----i---
4.6---------------------------















.4--------------------------------------------
3.8 -----------













1.---------------------- -----------------------------h----











NEATIONApo ADVISORY
CONFIETIAL COMMITTEE F AERONAUICS


.3 .4 .5 .6 .7 .6 .3
A1lA4cr

Fibwre 13.- Geeeralized dfrcxg clrvoe.


1.0 1/ /2 1.3


Fig. 13







Figs. 14,15


CON
.05-



O .04-



C .03-
*0


U
u

i .201


NACA ACR No. LGI71


Aach ndA mber, Av4

Figure /4.- Varic -oro o F measured y'lht' d Or acid
esfiPma-ed dcrag wi#/A /vlch rmimber for XP-51
Cf/rpfarme .


.0



0
+.
o .0



.0


S.0


.1 .2


.j .4- .- .0
AMach rtcrmber, AM


.7 .0


Figure /5.- Variafioor) of measured y/4A-f drag cOnd
estimated drcq with /tvich riLumber for XF2/t-2
oirp/one .


5

----- Over-all cdroa (qigt/ ) ,-/
4 Estimao+ed drag from /
generoalhed drag c"rve /,









,i- ------ ---

NATIONAL ADVISORY
CONFIDENTIAL CONMITTEE rOI MAEROMWCS
7.. f







NACA ACR No. L5G31 Fig. 16


CONFIDENTIAL %-- A/
/. 25 - - ^ -- 5. 8
5.4
4.6
120 4.2
3.4
3.0
/.15
A, X. VIA


1.4

101





.95
S.90 ---" ,





AY .77
II





.65 / "', .....a.,5000 i
g -7




07. ..






X e 6 .- i r .I .....
ol, I / :II I I

/" ~ A 5L, :::


-rfci A/c nuimbe:., M :
.,wr /6.-' Te mia Mc k num er...-
_ _ _ ^ I I I I I- ll _
_ ^ Z '- ^ : : :_ _
,, I I I I I I I I I 1 1 1 [ 1 1 I ll'.^ ^ 7 _

< O FIE TA /.o - i II 1i ",N^T-- -- /i[-i- C$---

0 .155 70-_- __- ---.-- ^0 85 .9 Z ^_ .95 lO_ __
? ___ _________ __ loc/ ^ -._7^/_ e r __ _c r __ _
F<.re1 .__ _- Termi ^7^l _7kc/ rl__^ner cr_,







NACA ACR No. L5G31


SI oI

-- -- ----
S 0 )
0-\ Q 2-
--% -- ,-- -
--- ^^^T^-TA----- ---------- -- 5


Cz
--- ^8-^- ^-i---
---^^^^-- --- ----S




*1-
-~ ~- ,. ^ -^v- _-- --o -- -





~--5---------r----------k------




--- --t-
___ Z
mmm~dr-^8-E


0 N N 0 00 0
CD 00 CD G Io
Ias .dquwnu, YpokV /O4lous.v/q


N


Pig. 17


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u



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UNIVERSITY OF FLORIDA

3 1262 08105 001 4



U: ,.,E, -T',' OF F..O IDA
DOCUJMErNTS DR.rTENT
120 IIARSTON SCiENCE LIBRARY
-.U. OX 117011
GAINESVILLE, FL 32611-7011 USA






















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