Climb and high-speed tests of a Curtiss no. 714-1C2-12 four-blade propeller on the Republic P-47C airplane

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Material Information

Title:
Climb and high-speed tests of a Curtiss no. 714-1C2-12 four-blade propeller on the Republic P-47C airplane
Series Title:
NACA WR
Alternate Title:
NACA wartime reports
Physical Description:
13 p., 20 leaves : ill. ; 28 cm.
Language:
English
Creator:
Vogeley, A. W
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:
Propellers, Aerial   ( lcsh )
Aerodynamics -- Research   ( lcsh )
Genre:
federal government publication   ( marcgt )
bibliography   ( marcgt )
technical report   ( marcgt )
non-fiction   ( marcgt )

Notes

Summary:
Summary: Flight tests were made of a Curtiss No. 714-1C2-12 four-blade propeller on a Republic P-47C airplane in climb and at high speed. The loss in efficiency when power was increased from normal to military was found to be from 5 to 8 percent in climbs at an indicated airspeed of 165 miles per hour. This loss was attributed primarily to reductions in section lift-drag ratios resulting from increased operating lift coefficients.
Bibliography:
Includes bibliographic reference (p. 11).
Statement of Responsibility:
by A.W. Vogeley.
General Note:
"Originally issued December 1944 as Advance Confidential Report L4L07."
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

Source Institution:
University of Florida
Rights Management:
All applicable rights reserved by the source institution and holding location.
Resource Identifier:
aleph - 003807441
oclc - 126841844
System ID:
AA00009432:00001

Full Text


ACR No. L4L07




NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS





WARTIME REPORT
ORIGINALLY ISSUED
December 1944 as
Advance Confidential Report L4L07

CLIMB AND HIGH-SPEED TESTS OF A CURTISS NO. 714-1C2-12
FOUR-BLADE PROPELLER ON THE REPUBLIC P-47C AIRPLANE
By A. W. Vogeley

Langley Memorial Aeronautical Laboratory
Langley Field, Va.

jUN,,ERSITY OF FLORIDA
DOCLUM ENTS DEPARTMtENTE
2 MIARSTON SCIENCE LIBRARY
P-'. B0\ 117011
-..l SiLLE, FL 32611-7011 USA








WASHINGTON

NACA WARTIME REPORTS are reprints of papers originally issued to provide rapid distribution of
S advance research results to an authorized group requiring them for the war effort. They were pre-
viously held under a security status but are now unclassified. Some of these reports were not tech-
nically edited. All have been reproduced without change in order to expedite general distribution.

L -177



































Digitized by the Internet Archive
in 2011 with funding from
University of Florida, George A. Smathers Libraries with support from LYRASIS and [he Sloan Foundation


http://www.archive.org/details/climbhighspeedteiOilang










:I'.CA AC !r LLT '7

NATIONAL ADVISORY COMMITTEE FtO AERONAUTICS


ADV.L-'T: CONFIDENTIAL REPORT


CTI-'B AI ir HIGH-SPEED TESTS OF A CURTISS NO. 714-1C2-12

FOUR-BLADE PROPELLER ON THE REPUBLIC P-J7C AIRPLANE

By A. W. Vogeley


STIUMMARY


Plight tests were made of a Curtiss No. 714-1C2-12
four-blade propeller on a Republic P-47C airplane in
climb and at high speed. The loss in efficiency when
power vas increased .:,n. normal to military was found to
be from 5 to 8 percent in climbs at an indicated airspeed
of 165 miles per hour. This loss was attributed primarily
to reductions in section lift-drag ratios resulting from
increased operating lift coefficients.

In high-speed flight at military power, losses in
efficiency due to compressibility started at an airplane
Mach number less than 0.4 and increased steadily to 10
or 11 percent at an airplane Mach number of 0,7. These
losses were encountered whenever the propeller-tip Mach
nui.,t:r e-:ceeded 0.88 and the propeller efficiency
decre-9.ed at a rate of about 7 percent for an increase
cf 0.! in tip Mach number. At an airplane Mach number
of 0.7 aid constant propeller rotational speed the
propeller efficiency decreased with a decrease in power
bel:. rililtary power. In comparison with the efficiencies
of l..w-sc~ ed flight tests (a Mach number of approxi-
rmately 0.53) at the same advance-diameter ratio, however,
the c.-rnpressibility loss was relatively independent of
i_. :.' V = r .

The tests indicated that, by suitably increasing the
solid.lit, and reducing the rotational speed, it may be
possi:.le to improve the propeller efficiency in both
cli:rib and high-speed operation.










I"ACA ACR No. L4LO7


INTROHLDUCTTICIT


a part of a program of flight tests of several
nrooellers on the Renublic -l,'7S airplane for the purpose
of detcrmninng climb and high-speed characteristics,
tests have been mede of a Curtiss No. '14-C12-12 four-
blade nroneller. Results of these tests and a brief
analysis are presented herein.

The climb tests consisted of runs at normal rated
nowe.r indicated airspeeds of 160 and 165 miles per hour,
and altitudes from sea level to about 50,000 feet and
runs at Iilitary pow3r, an indicated airspeed of 165 miles
per hInr, and altitudes from sea level to about 25,000 feet.
High-soeed tests consisted of a series of runs covering
a 'acn number rane : .'- J.1 to 0.7 at approximately
constant opwer and robation'l speed and a series of runs
at a :Mach number of 0.7 and constant rotational speed
with varying oover. -n order to determine the effects
of colmressibility, t~c efficiencies iaeasured in the
high-spseed runs :cre co-rpared vwith those measured in runs
made at the same rove2 coefficient and advance-diameter
ratio but au a Tach number of about 0.5.





V true airspeed

n propeller rotational speed, revolutions per
second

D propeller diameter

J advance-diameter ratio (V/nD)

4 section blade angle at 0.75R

9 blade angle at LnyV section

R propeller-tip radius

r pronpller-section radius

b blade-section chord


COTF'DL :' [LAL


CONFTLE? TIAL










HACA ACR :'o. LY LO7 CONFIDENTIAL 5


7h blade-section thickness

r2s radial distance from thrust axis to survey point

X r
_s = s


Po frce-stream static pressure

pTO free-stream total pressure

ApT difference between slilotrearm total pressure and
free-stre~a total pressure

T oro(eller tiihrust

Q propeller torque

CT propeller thrust coefficient

Cp propeller power coefficient

n oropeller efficiency

c ratio of density of free air to density of air
at sea level

p density of free air

Iairnlane Yach number

:t propeller-tip Mach number


PFROPELLER AND TEST EQUIl.T .;T

':eral specifications of the propeller and power
-l:~it -r3 as follows:


:.i.r. f blades .... ... ....... Four
l= signin . Curtiss ITo. 71L4-1C2-12
: sections . . Clark Y
r .11er diameter .. 12 feet, 2 inches
Frn .ell- r gear ratio .. 2:1
n .. .... Pratt & Whitney R-2800-21


COi TFI E .:i AL










D1 CI,:;1r rTIAL ::ACA ACR i c. L)4LO7


.1ilitary-po~er rating of en,Ciz.:
Engine s eed, rp .. .. 2700
Ianifold rssuressure, inches of mercury 52
Horsepo e r . . 2000
Critical altitude, feet ..... (approx.) 27,000

iormal-power rating of engine:
:.i;ine speed, r : . 2550
kanifold pressure, incnes of mercury .
Horsepower . . 1625
Critical altitude, feet (approx.) 29,000

The propeller, as tested, was equip-_ with the standard
production cooling cuffs. Blade-form curves are presented
in figure 1.

Propeller thrust wvs measured "' the slipstream
total-pressure survey method. For this purpose two survey
r.kes, connected to 2di- record.: multiple ne.nonmters,
were mounted Lho-rzc::cjly on either side of the fuselage
at the rear of the engine acoling, as shown in figure 2.
A photograph of eAe :- lano, propeller, ana survey rakes
is presented as f .'. e I.

Propeller btorc:- was measuredd with a standard
Pratt e ..'...tney tcrque r:set.r, to v:hicn vias connected a
stand~ d NA.A pressure recorder. An indicating pressure
gage was mounted in the cockpit for iuse by the pilot.
Standard _.' "A recording instruments wsre used to record
engine speed, i.:-, ct pressure, static pressure, and free-
air temperature. Propeller blade ergle was measured
with a special .ACA soark--'I e blede-angle recorder.





Climb tests.- .ith engine speed, manifold pressure,
and inuicated siospeed adjusted to the desired values,
short records on 1ll instruments were taken at intervals
of 2000 feet as the airplane climbed from sea level to
altitude.

(Climbs were made under the following conditions:

(1) military .-wer at normal climbing ir locatedd airspeed
of 165 miles per hour


CO:' I' 'TIAL










ITACA ACR No. LiL07 CONFIDENTIAL 5


(2) :Irnmal power at indicated airspeed of 160 miles rer
hour

(3) Normal power at indicated airspeed of 165 miles per
hour

The climb at military power was terminated at the
relatively low *1lituie of 23,000 feet because nf
insuff-cient en-; :e cooling indicated by bign cylinder-
head temperature

High-speed tests.- Each high-speed run was made at
values of engine s ped, to :'-, indicated airspeed, and
pressure altitude selected to produce a desired combination
of values of air ane Lach number, propeller advance-
diameter ratio, and power coefficient. Because the air-
plane was usually either climbing or diving during a run,
only engine speed, torque, and airspeed could be fixed.
These values were therefore held constant as the airplane
passed through the desired altitude, when a short record.
was taken.

The low-speed runs (M 7 0.5), used as a basis for
determining the effects of compressibility, were made in
the same manner as the high-speed runs.


REDUCTION 0F DATA


True airspeed, airplane Mach number, and air density
vere obtained by standard reduction methods from the
recorded values of impact pressure, static pressure, and
indicated free-air temperature. Engine speed, torque,
and propeller blade angle were recorded directly.

'ropeller power coefficient was calculated by the
form la

S2wQ
p n2D5


Propeller-tip Mach number was obtained from the
Sequa t ion


Mt = M4 + ( )2
\ / "--


CO\nTIDFPFT Ty









IA"TA ACR !o. L4L07


Propeller thrust coefficient was evaluated from the
measurements of slipstream total r-ressure by the method
described in reference 1, which gives


_/Pc \5/7 (i)
/ A7 p T (1 )
n d(r2)

In order to obtain the nondimensional quantities used in
the present report, equation (1) was reduced as follows:

dT dT T 1

d(xs2) Tr d(r2) 2 pn2D2
The areas under the curves of dCT/d(xs2) against x2
are equal to the thrust coefficients.


RESULTS AND DISCUSSION

Climb tests.- The variations of blade angle,
advance-diameter ratio, power and thrust coefficients,
efficiency, and propeller-tip and airplane a' :, '. numbers
with density altitude for the cli.-.;s are presented in
figures ) to 6. These flight data are also given in
table I.

In each of the climbs, changes in propeller effi-
ciency with altitude eppear to be small. Except for a
slight initial increase, efficiency tends to decrease
with altitude. This decrease is to be expected, since
the operating lift coefficients of the blade sections
increase with increasing altitude and approach the stall
region; the final result is to reduce the section lift-
drag ratios and to lower the efficiency.

Comrrressibility effects become evident in each of
the climbs whenever the propeller-tip T.alh number exceeds
about 0.06. Thrust-grading curves for climb at normal
power and an indicated airspeed of i60 miles per hour
are presented in figure 7 to show these effects of
compressibility at Ihl;'- propeller-tip "ach numbers. The
effects of compressibilit- are not evident in runs
20-1 to 20-11, in which tip :'.;cL numbers are below 0.55
(figs. 7(a) to 7(f)). Th-i. first effects are evident on
CO0::F iDE: TRIAL


COITFIDENTTAL










IACA ACR No. L-L07


the right side of the propeller disk for run 20-12
(fig. 7(g)), in which the tip Mach number has reached 0.86.
These effects continue to increase with tip Mach number.
Little or no evidence of compressibility loss exists on
the left side of the propeller disk, probably because the
left side is less heavily loaded than the right side
owing to inclination of the thrust axis to the air stream.
To tne extent, therefore, that the disk load distribution
is affected, the tip Mach number at which compressibility
effects first become evident is influenced by the airplane
attitude with respect to the flight path.

The tern "cororessibility effects" as used herein
means the effects shown by changes in the general shape
of the thrust-grading curves, for example, the dip in
the curve between xs2 = 0.6 and xs2 = 0.9 as measured
with the right survey rake in run 20-15 (fig. 7(h)). The
term does not include the effect that causes the grading
curves for bcth the right and left surveys to approach
zero at the tip at different values of xs2. This effect
is directly attributable to an unintentional yawed
attitude of the airplane held during the run, which
causes the slipstream to be displaced laterally at the
survey rakes.

Losses in thrust due to compressibility are present
at the hi,.gh;r tip ,ach numbers but no marked decrease in
efficiency attributable to this cause is apparent. With
further increases in altitude that result in higher
section lift coefficients and Mach numbers, however, it is
expected that the losses would extend over an increasing
oart of the disk area and that the effect on efficiency
would become significant. Compressibility losses can be
delayed by reducing the tip blade angles. This reduction
would result in a transfer of load to the inboard sections,
which operate at lower Mach numbers and can therefore
absorb the additional load without serious compressibility
effcctz. The inboard shift of load would also tend to
bring the blade loading into closer agreement with the
theoretically ideal load distribution for a propeller
coDrating at low advance-diameter ratios; thus the
possibility of a reduction in induced losses exists. The
use of this method is suggested only if particular emphasis
is put on climb performance, since large losses in efficiency
at high speed may result.

Since in the range of advance-diameter ratio for
climb the propeller operates at power coefficients greater


CONFIDENTIAL


CONFIDENTIAL










ONAC. ACR No. L4L07


than the values for maxiimim efficiency, it is 6enrerally
recognized that either a reduction in power coefficient
or an increase in advance-diameter ratio is necessary to
increase efficiency. These methods are illustrated by
comparing the efficiency levels (at tip Mach numbers
below 0.86) of the climbs (figs. to 6).

Tn the military-power clino (fig. 4), the propeller
operates at an efficiency of about 70 percent. By
reducing the p-wer coefficj tnt at essentially the same
advance-diameter ratio, as in the normal-power climb at
an indicated airsoeed of 160 miles oer hour (fig. 5), the
propeller efficiency is increased to ar.orximately
80 percent. An additional E-in in efficiency of about
5 percent is achieved by incre-asln, the airplane speed
and thereby increasing the adv'nce-diameter ratio, as in
the climb of figure o. These -,ins in efficiency are
due primarily to reductions in the section lift coef-
ficients that cause the sections to operate at lift-drag
ratios aoproaching the optimum. The climb performance
of the airplane is, of course, not iir~" roved by the
increase in prooeller efficiency because of the large
reduction in Dower required to effect the increase. In
order to improve the airplane climb performance, a
proneller designed to absorb military power at these
higher section lift-drag ratios is necessary; in effect,
an increase in solidity is required.

T', h-hoeed t-est. In order to deter'-ir.e the effects
of co:,.:.rc.sibili' y on prooeller operation at constant
power, two series of runs were made at airplane
Mach numbers ranln- fr ~.i O.1 to 0.7. One series was
made at a power coefficient of about 0.53, which
corres:n-"' s approximately to military-r.ower operation at
critical altitude (27,000 ft). 'y= second series was
made at a r'.ower coefficient of about 0.29, which
corresponds to military oower at an altitude of about
18,O00 feet. if e data obtained in these tests are given
in table II.

The ni.-peller efficiencies measured at hglh seeds
are compared in figure 8 with the efficiencies measured
at low soted (F = 0.5) in runs covering the same
ranges of nower coefficient and advance-diameter ratio.
The low-speed tests are .3.i:,,,arized in figure 9, which
shows the variation of propeller efficiency with -power
coefficient ai-n advance-diameter ratio.


CC;! IDzUiTIAL


CON7 DFT AL










NACA ACR No. LLO7


At. the oroneller speed used in the runs of fi, ire 8,
1sjes in efficiency due to compressibility apparently
betsn 5t an airplane Mach number below 0.L, increase
stetdilly, and reach 10- to 11 percent at an airplane
Nsacl rn.:Toer of U.7. The corresponding propeller-tip
Sach. nu.b.:-rs range from about 0.95 to 1.07.

T-h effect of propeller-tip Mach number on efficiency
is s.h ,vn in figure 10, in which the ratio of high-speed
efficiencY to low-speed efficiency is given as a function
of the high-speed propeller-tip Mach number. Figure 10
shYow, that losses in efficiency begin at Mt = 0.88,
which is in close agreement with the results of the
cli.icr tEsts. The efficiency loss due to compressibility
is shown to increase at the rate of about 7 percent for
an incr-.a.ae of 0.1 in tip Mach number.

Thrist-grading curves of runs at a power coefficient
of 0.75 are presented in figure 11. As in the climb
runs, only the right side of the propeller disk shows
any aprreciable compressibility loss-(fig. 11(a)). As
the Tach number is increased, however, compressibility
1:-ss also become evident on the left. side (fig. 11(b)).
.ith further increase in Mach number, the losses become
larger .-nd extend inboard over a greater portion of the
or:oseller blade.

The thrust-grading curve of a run made at an
airilsj'. ;iach number of about 0.5 and at a reduced
rotational speed is presented in figure 12. The advance-
oia7eter ratio and power coefficient are approximately
the are as those of figure 11(f). .The marked difference
in thie s3ipe .of these grading curves indicates the extent
of the l)oses in the high-speed run of figure 11(f).
Fig.ire 12 may also be c,-,:;:ared with figure ll(b). These
two r''.ns were made at roughly the same power coefficient
and airriclane Mach number. The curves for the two runs
il1,.Lr..te how compressibility losses may be reduced by
cecr,-asin the propeller rotational speed and thereby
reducing the section Mach numbers. By reducing the
rotational speed, the propeller efficiency is increased
about L percent or about one-half the increase to be
expected from the reduction in tip Mach number alone
(fig. 10). This difference indicates that the propeller-
tip :'ach number alone does not determine the magnitude
of the cor.ipressibility losses.


COTiFI DETIAL


COnFIDE,'_TIAL










NACA ACR No. .147


Ti,- effect of loading on the proreller efficiency
at high seed was investigated "- making a series of runs
at an airplane Lsach number of about 0.7 and constant
nrooeller speed with varying power. The results of tiese
tests are compared in figure 13 with the results taken
from figure 9 of low-speed tests at the same advance-
diameter ratio and power coefficients. The extrapolated
point in figure 15 was determined by first extending the
curve for the high-speed tests (Cp z 0.55) in figure 8
to an airplane Mach number of 0.7 and an advance-diameter
ratio of 2.6. The value of efficiency obtained was then
corrected to an advance-diameter ratio of 2.7 by using
the curve for the low-speed tests (Cp = 0.35) of
figure 8. As the power is reduced the propeller
efficiency decreases at both low and high speeds. The
compressibility loss at high speed, as measured by the
difference in high-speed and low-speed efficiency, appears
to be relatively independent of power and is about
10 to 14 percent throughout the rrng- investigated.

The effect of compressibility is to reduce the lift
coefficient for maximum section efficiency as the critical
Mach number is exceeded. A decrease in power would,
consequently, be expected to cause a reduction in
comoressibility loss. In this case, however, some
sections of the propeller are apparently operating at
approximately maximum efficiency and some, at lift coef-
ficients above those for maximum efficiency. Under such
circumstances a reduction in power would result in a
decrease in efficiency of some sections and an improvement
in efficiency in others; the over-all effect would be
only a small c',.iige in comressibility loss. Figure 14
shows that the tip sections are operating at highest
efficiency at hi.h nower, since' as the power is reduced
the tip sections produce a decreasing amount of thrust
in comparison with the inboard sections. Some gain in
high-speed efficiency could probably be obtained by an
adjustment in load distribution.

A comparison of the results of the high-sc-ed and
low-speed tests indicates that, in order to prevent
large losses in efficiency, blade-section Mach numbers
must be limited by reducing the -rotational speed. At the
same time, however, any adverse effect due to the
increase in section lift coefficients necessary to
absorb the same engine power at a lower rotational speed
must be avoided by a proper increase in propeller solidity.


CONFIDENTIAL


CONFIDENTIAL










HIACA ACR No. LiL07


COjiC LUSIONS


Fljht tests of the Curtiss To. 714-1C2-12 four-
blade propeller on a Republic P-47C airplane indicated
tie following conclusions:

1. In climbs at an indicated airspeed of 165 miles
pe-r ihui. from 5 to 8 percent was lost in efficiency by
increase ng fr:m normal to military power, primarily
b-:c- use .f the reductions in section lift-drag ratio that
re'slted from increased operating lift coefficients.

2, Riith military nower, losses in efficiency due
to c:i.mpr!essibility started at an airplane Mach number
less than 0.4, increased steadily, and reached 10
t-' 11 oe cent at an airplane Mach number of 0.7. Compressi-
bility, losses becilne evident whenever the propeller-tip
'acIh inu.'ler exceeded about 0.88, and the propeller
effici-iicy decreased at a rate of about 7 percent for an
increases of 0.1 in ti .Tach number.

5. At an airplane Mach number of 0.7, a reduction
inr engine power below military power resulted in a lower
pr-opell: r efficiency, but the loss in efficiency due to
comrires ability (based on low-speed tests at a corre-
sponding advance-diameter ratio) was relatively independent
of power.

4. By suitably increasing the solidity and reducing
the rotational speed, an improvement in the propeller
efficiency in both climb and high-speed operation may be
c :ssitle.


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


REFERENCE


1. V),aley, A. W.: Flight Measurements of Compressi-
bility Effects on a Three-Blade Thin Clark Y
Propeller Operating at Constant Advance-Diameter
Ratio and Blade Angle. NACA ACR No. 3G12, 1943.


C1OFIDE'TIAL


CO1ITDi DETIAL








fACA ACFE lo. L4LO"


TABLE I
FLIGHT DATA OBTAINED FROM CLIMB TESTS OF
CURTISS NO. 714-1C2-12 FOUR-BLADE PROPELLER


Fig. Run J Cp CT I lM It
S(rps)(deg)


4
4
4






5, 7(a)
5
5, 7(b)
5
5, 7(c)
5
5, 7(d)
5
5, 7(e)
5
5, 7(f)
5, 7(g)
5
5
5, 7(h)
6
6
6
6
6
6
6
6
6
6
6
6
6
6
6
6


29-1
29-2
29-3
29-4
29-5
29-6
29-7
29 -;
29-9
29-10
29-11
20-1
20-2
20-3
20-4
20-5
20-6
20-7
20-8
20-9
20-10
20-11
20-12
20-15
20-14
20-15
18-1
18-2
18-3
18-4
18-5
18-6
18-7
18-8
18-9
18-10
18-11
18-12
18-15
18-14
18-15
18-16


0.94
.96
.99
1.02
1.08
1.12
1.13
1.19
1.22
1.26
.94
.99
1.00
1.06
1.08
1.09
1.15
1.17
1.20
1.26
1.28
1.36
1.41
1.4
1.:.
1.00
1.01
1.07
1.11
1.15
1.16
1.20
1.24
1.28
1.32
1.34
1.37
1.42

1.54
1.56


0.146
.179
.187
.200
.212
.230
.238
.253
.273
.289
.302
.140
.150
.165
.191
.189
.195
.215
.228
.240
.261
.272
.30

.536
.145
.158
.171
.184
.196
.205
.220
.237
.252
.269
.279
.296
.320
.332
.566
.379


0.122
.159
.15
151
154
160
.162
.165
.171
.178
.179
.116
.121
.152
.146
i40
143
.150
.156

.168
.179
.135
.186
.113
.126
.133
157
l15
.149
.153
156
.163
.168
.170
.176
.178
.181
.191
.186


0.783
.750
.764
.768
.761
.750
.758
.74
.749
.749
.779
.798
.799
.810
.798
.802
.8o6
.799
.790



.786
.787
.816
.809
.838
.847
.845
.8533
.819
.822
.822
.819
.811
.791
.808
.801
.764


22.52
22.63
22.62
22..4
22.50
22.44
22.37
22.54
22.49
22.61
22.49
21.21
21.29
21.36
21.20
21.52
21.49
21.50
21.74
21.52
21.45
21.5
21.:1
21.15
21.48
21.25
21 .37
21.36
21.22
21.55
21.50
21.4o
21.45
21.57
21.60
21.48
21.51
21.58
21.40
21.65


.233
.240
.246
.254
.263
.274
.28 4
.290
.306
.5320
.334
.223
.257
.240
.254
.263
.271
.284
.290
.302
.5317
.523
.342
.357
.570
.5388
.237
.241
.254
.264
.274
.283
.294
.505
.317

.350
.368
. 89
.405
.420


0.812
.818
.825
.825
.8*?
.845
.850
.869
.865
.882
.896
.782
.790
.793
.792
.812
.828
.825
.849
.852
.852
.861
.874
.891
.920
.780
.788
.786
.794
.79
.825

.843
.851
.874
.882
.922
.922
.952


0.923
.8582
.805
.756
.709
.662
.629
.592
.553
.512
89
.968

.775
.731
.692
.648
.6108
.566|
.5281
.506
.67


.31
.438
9832
.764
.726
.683
645
.604



.4o 38

.-57
.551


I I I I L S


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS


26.6
28.4
28.9
29.
30.8
31.6
32.4
35.2
_4.3
5.-2
36.2
26.4
27.6
28.6
30.2
30.4
31.1
32.2
52.9
3.8
354.8
35.8
37.0
37..8
39.1
40.0
27.2
28.0
29.0
30.2
30.7
31.6
32.4

35.3
36.2
37.2
38.3
59.2
40.4
41.2








IJACA ACE No. L4L07


TABLE II
FLIGHT DATA OBTAINED FROM HIGH-SPEED TESTS OF


CURTISS NO.


714-1C2-12 FOUR-BLADE PROPELLER


Pig. Run J Cp CT n M Yt a P
(rps) __(deg)


24-6

24-5
24-1
24-2

24-3
24-4
17-1
17-2

17-3
17-4
17-5
18-17
19-18
20-18
21-9
21-10
21-11
21-12
21-13


1.59
1.84
2.08
2.21

2.45
2.47
2.69

2.77
2.70

2.73
2.75
2.68
2.67

2.58
2.68
2.55
2.32
2.14

1.95-


11(a)
11(b)
11(c)
11(d)
11(e)
11(f)

I4(a)




14(b)




14(c)


14(d)







12


0.343
.347

.352
.351
.358
.346
.1144
.151
.164

.176
.204
.216
.221

.256
.282
.290
.292
.291

.295
.330


0.792
.801

.786

.759
.735
.708

.395
.482

.500
.476

.543
.636

.557
.609
.672

.704
.758

.797
.803


22.62
22.47
22.44
22.46
22..30
22.48
22.62
22.08
22.32
22.20
22.03
22.30
22.28

23.31
22.42
22.45
22.47
22.41
22.49


0.171
.151
.134
.121

.107
.099
.021
.026

.030
.031
.040
.051
.046
.060

.071
.080
.096
.108
.122
.109


0.431
.495
.557
.594
.655
.666

.711
.712
.702
.706
.705
.701
.693
.699
.695
.662
.604

.554
.508
.505


0.952
.979
1.009

1.035
i.064

1.077
1.092

1.075
1.079
1.077
1.o69

1.078
1.070
1.100

1.071
1.047
1.015

.984
.960
.803


0.416
.416
.418
.422
.422
.432
.576
.560

.575
.569
.557
.539
.545
.551
.541
.521
.523
.527
.515
.716


39.8
41.4
44.1
45.3
47.5
47.9


47.8
,---


47.8

47.2


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS


.842 17.81


12-1 2.54


- --


---~


-~--
,,,,








IJACA ACE No. L4L07


4' -..- 4 4

...2. .... ...... .:tu .... 1-H H 4 f & t'w t
T 4 4
1 ^^ t Illi|;|i *lffl^ ^lti~ l: l t::1: E~i~i l lf llllll:! ll^;!^^ lf l tl l li ^
11 i~lK iSII ^-l i^ Iffei+


-I


I


figure B/ode -form curves f/r Curhss Vo 7/4-/C2-/Z
fodr- b/ade prope/fer,


I 4


Fig. 1


[


*T-H H






.. .. .t .. ..



1n 1 111 11 1 |l l l 11 |
S.. .... ......i|



!| l l| l | l|! ll


I
|ll~ ~ ~ !l | il 4+ ....ll !1 11!
.| | l ||.| .| |. .1 1| |

... ...........| | | || || |

iiii ^ i 1 1 ^ 1 1 l i inm
....... ..... ....... ..... + M 4 -


4 44,
MT : | i ----


IT



I I .. 4











MM|X


14

- -------I "

+~l + I I




:: m1; ^ i. : I













.7;: is^


++a1 :i =
f 41A


4-+








m









..
t |;g






.
l l~llI~ Ii


114.



m


P
-ii


^It^













V





r1
lilll


-ill n


4-;g








NACA ACR No. L4L7. Fig. 2


Figure 2.- Location of propeller and survey rakes on a Republic P-47C airplane.











NACA ACR No. L4L07


4 ri

' r '



II


SI,




I


ii.




r

I.: r
C:


r;I 1rlaE~
.t r:i
i ,
''
r


Fig. 3


al





CO





Co



zo








dd
0c








-14




0) A









I 4
c)
*1









u-H


'ri
St





M)0
r-l
Q)I-
h 0









NACA ACR No. L4L07


35
B, deg
30

25

/.4


1.2



.8


.30
.21 ---- -- -- -- -- -- -- ---- -- -- -- --



.20---------
.210

./15

.18

16



.12



80 0

70
s I--I__\ .--- I- 0-_ ,-I-.--0

/.0



.9 --7-- -. .


.4

'14 .' _.___,____---_-o-
7 -------- --i-----1-- -

.2
NOIONAL 0D
COMMITI TEE FOR A


0 4 8 /2 /6 20 24 28 x
Density altitude, ft

Figure 4.- Military-power climb of an indicated airspeed of
/65 mi/es per hour. Curhss No. 7/4-/C2-/2 four- blade
propeller on AReputbic P-47C airplane.


,7, percent






,,4


O3


Fig. 4









NACA ACR No. L4L07


J5
Sdeg F


0 -


1.4





3'.0 -- -- -- -



Cp .25 ---

20 .



/B
Cr 16--
/6
Cr rf


30

Opercen.+ O : "

70


[De


.8

.7

4

A4 7


0 4 8 /2
Densd


/6 20
a/fi+ude, ift


NATI NAL A
:OMMITT E FOR


AISORY
:RONAUVCS


24 28 32 x/03


Figure 5.- Norma/-power clmb .at an indlicaed airspeed of
/60 mi/es per hour. Cur iss No. 7/4-/C2-12 four -bade
propeller on Republc P-47C a/ipplane.


2_ -


-Fig. 5









I.ACA ACR tNo. L4L07


,le deg





J






Cp


30---

.35 -




L0 -- -- -- -'--- -- -^s^ -- -- --







.J1
./.5 --- --"- ^" -- -







.14- -

.12 ---
9--


7, -peren/ &o- I ------- ------ ----

70




A1 i I


0

.7


.4-
M 3

.2

/


0 4 8 / /C6 20 4
Density al/fude, ft


Figure 6.- Norma/-power chmb at an indicated airspeed of
/65 m//es per hour. Curfhss No. 7/4-/C2 -12 four-b/ade
propeller on Republic P-47C airplane.


m ,,-FO--A-- -----=-
NAT ONAL A IVSGRY
OMMIT EE FOR I i)


Fig. 6


28 32YlOJ








NACA ACR tlo. L4L07


0








.J



.2




^ .1


(a) Ru n 20-1


(/) Pan 20-3.
Figure 7- Thrust-grading curves for cllmb
at normra/ power. /ndica fed airspeed, /6O
ai/les per hour.


Figs. 7a,b







NACA ACR No. L4L07



.3



---Q- .--- ----- t
+-
1o +- -









.22 .-















zefy
(C) Pan 20-5"






+
0----t--L--.sr_--





2-7









.2 .4 -



(cY) *un 20-7.


Figs. 7c,d


Figure 7. Conf/nued.








NACA ACR No. L4L07


.3




.2

































0
-.











I-












-.1


-+ -R+-9+-.
qs

.+ | \ J =/,O28
-\ _-Cp =.272
I /9~ \ \\ ^ = C ./68















F-gure 7.- Conh-nued.
Figure. 7.- Conhinued.


Figs. 7e,f


(e) Run 20-9.








NACA ACR No, L4L07


"31

























I- i


(g) Run 20-12.


fh) Pun 20-IS.
Figure 7 Concluded.


Figs. 7g,h







NACA ACR No. L4LO7 Fig. 8
















+-------------------------------------------

-.. ----------
-. I '4 4w





i-I-- k -,






( 0--- 0 i Z
L -- I I I -- i -- -- I I I -- -- l L O L ,
S~~I SP O o


4 uoawad '0 4houalolya ^ff/fadw








NACA ACR No. L4L07


I


(O



Qj
Q QQ

04,.















vo
L CZ

0 ^0







In '*"*

^1^,
1 1~~
^ ,a

CO'0\5

u


Fig. 9


do '4LCIaI3Ja0o3 ,aMOc&








NACA ACR No. L4L07 Fig. 10




'I-




-|i --





__ __ __r 0
Ju+ .





-S`
++



__ O/ + O




0- +




/ 3


7-




N ,
I __ __ __ __ ___ ^ ^


'O--W 'pcpagd Mof- /1 I
padjds7 q 61 q pt1









NACA ACR No. L4L07


k
>3


(a) Pun 24-6.


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS


(6) f-un 24-5.


Figure //. Thrust-grading curves for Cp 0.3'-


Figs. lla,b







NACA ACR No. L4L07


.3



.2 I--


I-~


(c) Pun 24-/.


Figs. llc,d


(d) Run 24-2.


Figure /. Contfnued.


-- ---- -2111-- --- -----
J = 2.20
__ -+ + _-a F:/ight surrv C_ = .JS/
S"+/ -...L + C7 =.21
+ F__ _= .359

l l Zef* .srveq / 35

1
---^-




.4 .6 .8 /.O /.2

/ NATI NAL AT I SORRY
___ l) ___ OMMITTI E FOR A :RONAUT







NACA ACR No. L4L07







0 -
CI
'---S' ri

S/






/


(e) Run 24-3.


y
A"
-<5


Figs. lle,f


(f) Pan 24-4.
F/gure //. Concluded.


--- /gh su/v-ey TI = 2.7
-- .I Cp-=.34
S/ +- Cr =.099
S__ _- Ia-- =.708
Lef s1 rv =.666
-\ M= /.,077

I
I .4 .6 .8 /, /2





NATI AL AD ISORY
C)MMITT FOR A ONAUTI .$







NJACA ACR No. L4L07 Fig. 12









I i II II I I




I- -
,___/ _. __ __ _


4-4









\ o

__ s__ A
+'j)


.
'i- --- 40
~ /_ _
N
-- a -. ---- 8


ILJ p







NACA ACR No. L4L07 Fig. 13





I S-




















Q.
-- --L -0---------- __











__
So







tz











-ou
--t --- u

Y --- -- s
K \ h










-^-v^- ^
-^ ^ /> S- <
^ I ---j-- ^ i >
^-ua







NACA ACR No. L4L07




.2



./



FI o _


fa) Run /7-1.


(b) kPun /7-4.

FIqure /4. Thrusu-grading curves for runs at
J,2.70 ond M 7OZ


Figs. 14a,b








NACA ACR No. L4L07


't3


(c) Pun /9-/8.


(d) Pan 21-9.


Figure /4-- Concluded.


Figs. 14c,d
















a





UNIVERSITY OF FLORIDA
II I III II lllll 1II
3 1262 08106 554 1




I ; i ~~CiT' F FL FDA


;- ,+ ,':.'., 1i170 11
,- :.LL.E, FL 32611-7011 USA












-1
7 1