Effect of compressibility on the hovering performance of two 10-foot-diameter helicopter rotors tested in the Langley fu...

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Title:
Effect of compressibility on the hovering performance of two 10-foot-diameter helicopter rotors tested in the Langley full-scale tunnel
Series Title:
NACA RM
Physical Description:
43 p. : ill. ; 28 cm.
Language:
English
Creator:
Jewel, Joseph W
Harrington, Robert D
Langley Aeronautical Laboratory
United States -- National Advisory Committee for Aeronautics
Publisher:
NACA
Place of Publication:
Washington, D.C
Publication Date:

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

Notes

Summary:
Summary: An investigation of the effects of compressibility on the hovering-performance characteristics of two 10-foot-diameter helicopter rotors having solidities of approximately 10 percent has been conducted in the Langley full-scale tunnel. One rotor, having NACA 0012 airfoil sections, a plan-form taper ratio of 3/1, and -8° of twist, was tested to a tip Mach number of 0.95 and a disk loading of 16 pounds per square foot. The other rotor had NACA 64-series airfoil sections, tapering to a 6-cent-thick tip, a plan-form taper ration of 3/1, -16° of twist, and was tested to a tip Mach number of 1.0 and a disk loading of 20 pounds per square foot.
Bibliography:
Includes bibliographic references (p. 15).
Statement of Responsibility:
by Joseph W. Jewel, Jr., and Robert D. Harrington.
General Note:
"Report date April 21, 1958."
General Note:
"Declassified May 29, 1959"

Record Information

Source Institution:
University of Florida
Rights Management:
All applicable rights reserved by the source institution and holding location.
Resource Identifier:
aleph - 003834396
oclc - 150505417
sobekcm - AA00006135_00001
System ID:
AA00006135:00001

Full Text












ARCH MEMORANDUM




OF COMPRESSBILITY ON THE HOVERING PERFORMANCE
'I:.UTWO 10-FOOT-DIAMETER HELICOPTER ROTORS TESTED
IN THE LANGLEY FULL-SCALE TUNNEL
By Joseph W. Jewel, Jr., and Robert D. Harrington
Langley Aeronautical Laboratory
Langley Field, Va.

i. UNIVERSITY OF FLORIDA
DOCUMENTS DEPARTMENT
l120 MARSTON SCIENCE UBRARY
p.O. BOX 117011
GAINESVILLE, FL 32611-7011 USA


NATIONAL ADVISORY COMMITTEE
FOR AERONAUTICS
WASHINGTON

April 21, 1958
Declassified May 29, 1959
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IACA RM L58B19

NATIONAL ADVISORY COMMITTEE FOR AEROIALITICS


RESEARCH I.EM0ORAI[DUM


EFFECT OF COMPRESSIBILITY ON THE HOVERING PERFOR'I4AICE

OF TWO 10-FOOT-DIAMETER HELICOPTER ROTORS TESTED

UIN THE LANGLEY FULL-SCALE TUINHEL

By Joseph W. Jewel, Jr., and Robert D. Harrington


SUMILARY


An investigation of the effects of compressibility on the hovering-
performance characteristics of two 10-foot-diameter helicopter rotors
having solidities of approximately 10 percent has been conducted in the
Langley full-scale tunnel. One rotor, having IJACA 0012 airfoil sections,
a plan-forn taper ratio of 5/1, and -0o of twist, was tested to a tip
Iach number of 0.95 and a disk loading of 16 pounds per square foot.
The other rotor had IIACA 04-series airfoil sections, tapering to a
6-percent-thick tip, a plan-form taper ratio of 5/1, -160 of twist, and
was tested to a tip Mach number of 1.0 and a disk loading of 20 pounds
per square foot.

As the tip Ilach number was increased to O.86, the IIACA 0012 rotor
encountered compressibility losses at progressively lower thrust coef-
ficients. At tip Mach numbers of 0.91 and above, compressibility losses
were evident at the lowest values of thrust obtained in the investigation.
No effects of compressibility were evident at any thrust coefficient
tested on the IIACA 64-series rotor up to tip Mach numbers of 0.91. At
tip Mach numbers of about 0.92 and above, compressibility losses were
evident at the lowest values of thrust measured. The measured power
requirements of both rotors at the lowest tip Mach numbers are sig-
nificantly more than predicted by usually reliable theory. This effect
is believed to be due to the unusually large ratio of hub diameter to
rotor diameter. The absolute values of the measured data are therefore
not believed to be representative of full-scale rotors; however, the
effects of Mach number on either of the rotors or comparisons between
rotors are believed to be valid.

The maximum sound level measured in the tunnel test chamber at a
point 21.5 feet and 140 from the vertical shaft on the NACA 64-series
rotor was 122 decibels at a tip Ilach number of 1.0.






2 NACA EM L58B19


INTRODUCTION


One method of increasing the forward speed and lifting capabilities
of the helicopter and thus its utility is to operate the rotor at high
tip speeds. Accordingly a general research program is currently under
way at the Langley full-scale tunnel and the Langley helicopter test
tower to determine the effect of compressibility on the hovering and
forward-flight performance characteristics of helicopter rotors.

As one part of this program two rotors 10 feet in diameter have
been designed to study the forward-flight performance characteristics
of rotors operating at high tip speeds. Hovering tests of these two
rotors have been made as a preliminary step to aid in the interpretation
of the forward flight results as well as in evaluation of the hovering
performance of rotors designed for high forward speeds. This report pre-
sents the results obtained from these hovering tests.

One set of rotor blades tested had an NACA 0012 airfoil section, a
plan-form taper ratio of 5/1, and -8 of twist. This rotor was tested
to a tip Mach number of 0.95 and a disk loading of 16 pounds per square
foot. The second set of blades had symmetrical NACA 64-series airfoil
sections, tapering in thickness ratio from a 642-015 section at the cen-
ter of rotation to a 64-006 section at the tip. These blades had a plan-
form taper ratio of 5/1 and -160 of twist. They were tested to a tip
Mach number of 1.0 and disk loadings to 20 pounds per square foot. For
this rotor, sound-level measurements were taken at each test condition.


SYMBOLS


a slope of section lift coefficient against section angle of attack
(radian measure)

b number of blades per rotor

c blade section chord, ft


cR
f0cr2dr

Ce equivalent blade chord (on thrust basis), R ft

o r2dr





NACA EM L58B19 5


cd,o section profile-drag coefficient

cl section lift coefficient

6CT
-l mean lift coefficient, -


CQ torque coefficient,
nR2p(QR) 2

Qi
CQi induced-drag torque coefficient,
nR2p(nR)2R

QO
CQo profile-drag torque coefficient, Qo
nR p(QR) R

T
CT thrust coefficient,
aR2p(,R)2

I mass moment of inertia of blade about flapping hinge, slug-ft2

C 5/2
MI rotor figure of merit, 0.707---
CQ

Mt blade-tip Mach number (ratio of blade-tip speed to speed of
sound)

Q rotor-shaft torque, Ib-ft

Qi rotor induced-drag torque, lb-ft

Qo rotor profile-drag torque, lb-ft

r radial distance to blade element, ft

R blade radius, ft

T rotor thrust, lb

v induced inflow velocity at rotor, ft/sec






NACA RM L58B19


ratio of blade-element radius to rotor-blade radius, r/R


blade-element angle of attack, measured from line of
9 p, radians

cepR4
mass constant of rotor blade, ep
I


zero lift,


blade pitch angle, radians

mass density of air, slugs/cu ft

bee
rotor solidity,
iR


local solidity of blade element at spanwise station x,


inflow angle at blade element in plane perpendicular

span axis, a -1 + 1 + radians
16x a'a


bcX
xR


to blade-


Q rotor angular velocity, radians/sec

Subscripts:

t rotor tip

x at station x


APPARATUS AND TESTS


Rotor blades.- The rotor blades tested in this investigation were
constructed from resin-impregnated glass cloth, balsa planking, and spruce
plywood. The main spar of the blades extended to about the 51 percent
chord and was formed in a pressure mould from resin-impregnated glass
cloth. The rear portion of the blades was formed by spruce plywood ribs
covered with 1/8-inch balsa planking. The complete blade was then wrapped
with resin-impregnated glass cloth. Templates indicated that the leading
edge conformed to the true contour of the airfoil sections and that the
blade surfaces rearward of the spar were smooth and fair. Two different
sets of blades were tested in this investigation. One set had an NACA
0012 airfoil section, a plan-form taper ratio of 5/1, and -8 of twist






NACA RM L58B19


(fig. 1). The other set had symmetrical NACA 64-series airfoil sections,
tapering in thickness ratio from a 642-015 section at the center of rota-
tion to a 64-006 section at the tip. These blades had a plan-form taper
ratio of 5/1 and -160 of twist. A more complete description of the phys-
ical characteristics of the rotors is given in table I and in figure 2.

General arrangement.- The general arrangement of the model mounted
in the test section of the Langley full-scale tunnel is shown in figure 3.
The blades were attached to a fully articulated rotor hub. The hub flap-
ping axis was located at the center of rotation and the lag axis at the
10.4 percent radius. For these tests the lag hinges were locked at zero
lag angle. The hub extended to 20.4 percent radius. The relative size
of the hub and blades is shown in figure 4.

The rotor hub was mounted on an extension shaft which was connected
through a flexible coupling to the drive shaft of a 1,000-horsepower
electric motor. The complete assembly was supported by the struts leading
to the tunnel balance.
*
Instrumentation.- The following quantities were measured and recorded:
rotor thrust, torque, coning angle, and shaft speed. Rotor thrust was
measured on the wind-tunnel balance, a complete description of which is
given in reference 1. Torque measurements were obtained from a strain-
gage beam mounted on the rotor drive shaft. Rotor speed was indicated
by an electromagnetic pickup which was also mounted on the drive shaft.
Blade-root coning angles were obtained from a strain-gage beam mounted
at the hub flapping axis. Rotor torque and coning-angle information were
transmitted through a 24-ring slipring assembly mounted on the bottom
motor shaft.

Corrections for down load on the supporting structure have been cal-
culated and applied to the thrust measurements, and blade-removed torque
tares have been subtracted from the torque data. The estimated accuracies
are as follows: thrust, 5 pounds; torque, 1/2 percent at 1,000 pound-
feet; rotor speed, 1/2 percent at 2,100 revolutions per minute; and
angular measurements, 0.05.

Tests.- The rotor having the NACA 0012 airfoil section was tested
over a range of tip Mach numbers from 0.45 to 0.95. The maximum values
of CT obtained varied from 0.011 at Mt = 0.45 to 0.0048 at Mt = 0.95.
The rotor having the NACA 64-series airfoil sections was tested over a
range of tip Mach numbers from 0.45 to 1.0. In this case, the maximum
values of CT obtained varied from 0.0106 at Mt = 0.75 to 0.0068 at
Mt = 1.0. The coning angles for both rotors varied within 10 of 00
coning angle for the complete range of tests.






NACA RM L58B19


Sound-pressure measurements were made over the complete range of
test conditions on the NACA 64-series rotor with the aid of a sound-level
meter. The microphone was located 21.5 feet from the center of the hub
and at an angle of 1400 measured from the vertical shaft and below the
plane of the rotor. The indicated sound levels could be read to
0.5 decibel; however, the absolute sound level is not representative
of free-air conditions since the acoustical characteristics of the full-
scale-tunnel test chamber are not known.


RESULTS AND DISCUSSION


The results of this investigation are discussed in three parts.
The first section presents the analysis of the power-required data
obtained with both the NACA 0012 rotor having -80 of twist and NACA
64-series rotor having -160 of twist. The second section presents a
comparison of the relative static-thrust characteristics of the two
rotors, and the third section discusses the sound measurements obtained
during the NACA 64-series rotor tests.


Power Requirements of NACA 0012 Rotor

Performance measurements.- The measured variation of thrust coef-
ficient CT with torque coefficient Cq over a range of tip Mach num-
bers from 0.45 to 0.95 for the rotor having NACA 0012 airfoil sections
and -8 of twist is shown in figure 5. These data indicate that as the
tip Mach number increases the rotor encounters compressibility losses at
progressively lower thrust coefficients, as indicated by the departure
of the higher Mach number data from the Mt = 0.45 curve. This general
trend is consistent with the results of previous investigations refss. 2
to 4). However, in the present investigation, data were obtained at
somewhat higher tip Mach numbers than in the past. At tip Mach numbers
of 0.91 and above, substantial power increases are indicated even at zero
thrust. Power increases of the type shown for tip Mach numbers between
0.45 and 0.86 could result either from compressibility effects or stall.
However, analysis indicates that the maximum blade-tip angles of attack
at any Mach number obtained during the tests were less than would be
required to produce stall. Therefore, it is probable that the power
increases shown in figure 5 are the result of the rotor reaching com-
binations of Mach number and section angle of attack above the compres-
sibility drag rise. At tip Mach numbers of 0.91 and above, the blade
tip sections are operating well above drag rise even at zero thrust.

Profile-drag power ratio.- The compressibility losses encountered
by a rotor are associated primarily with an increase in profile-drag power.





NACA RI.I L5gB19


Analysis has shown that the induced power coefficient at constant CT
is essentially independent of tip Mach number. A sudden increase in
profile-drag power occurs when a combination of blade-section angle of
attack and Mach number exceeds the values for drag divergence.

A convenient way to show the onset and magnitude of the compres-
sibility losses is by means of a plot of the ratio of the profile-drag
torque coefficient CQ,o at any Mach number to the incompressible
profile-drag torque coefficient as a function of rotor mean-lift coef-
ficient F,. Such a plot is shown in figure 6. The profile-drag torque
coefficients were determined by subtracting a calculated induced-drag
torque coefficient CQi from the corresponding measured total torque
coefficient CQ. It is apparent that as the tip Mach number increases
the rotor encounters compressibility losses at progressively lower mean
rotor lift coefficients. At tip Mach numbers of 0.91 and greater, the
rotor is operating well above the drag rise even near zero mean lift
coefficient.

These same profile-drag torque-coefficient ratios are shown as a
function of the calculated rotor-blade-tip angles of attack in figure 7.
From this plot it is possible to estimate the combinations of tip Mach
number and angle of attack at which drag divergence occurs on the rotor.
A comparison of the drag-divergence Mach numbers estimated from figure 7
with those indicated by two-dimensional airfoil data is shown in fig-
ure 8. In general the rotor drag-divergence Mach numbers are about 0.10
higher than those shown by the two-dimensional airfoil data. This value
compares with values of about 0.02 to 0.06 which have been measured in
previous tests refss. 2, 4, and 5).

Comparison with theory.- The experimental and calculated hovering-
performance curves for this rotor at tip Mach numbers of 0.45, 0.75,
and 0.81 are compared in figure 9. The calculated results are based on
a strip analysis (see ref. 6) in which section lift and profile-drag
coefficients were varied with Mach number along the blade. The varia-
tion of lift and profile-drag coefficient with Mach number used for these
calculations is shown in figure 10. Two different tip loss factors were
applied to the calculated thrust and induced torque. One of these,

B = 1 .T, is based on a triangular inflow distribution, whereas
b
the other, B = 1 1.86 is based on a uniform inflow distri-
b CT

button. The development of these equations is shown in the appendix.
bution. The development of these equations is shown in the appendix.






8 NACA RM L58B19


Because of the twist and taper this rotor probably has an inflow distri-
bution somewhere between triangular and uniform. In addition, at tip
Mach numbers of 0.75 and 0.81 a tip Mach number relief was included in
the calculations. This relief is characterized by an effective reduc-
tion in tip Mach number due to three-dimensional flow at the blade tip.
It was assumed that the effective tip Mach number was 0.05 less than that
calculated from test conditions and that this relief varied linearly to
zero at the 5/4 blade radius. The more usual value of 0.05 was chosen
for these calculations rather than the value of 0.10 indicated in fig-
ure 8. Inspection of figure 9 shows the relative effect of each of these
correction factors on the performance calculations.

The theoretical performance calculations made for this rotor at a
tip Mach number of 0.45 predict an average of 6 to 9 percent more thrust
for a given amount of power than was actually measured depending on the
type of inflow distributions assumed in the calculations. (See fig. 9(a).)
This agreement is considered only fair when compared with the close agree-
ment generally obtained at low blade-tip Mach numbers on larger diameter
rotors. (See, for example, refs. 2, 5, 4, and 7.) An explanation of why
the measured rotor performance is significantly less efficient than the
usually accurate predicted results has not been precisely defined. There
may be several factors present in the test setup, however, which have a
detrimental effect on the measured rotor performance. It is believed
that the primary reason for the discrepancy is a flow disturbance caused
by the unusually large ratio of hub diameter to rotor diameter, which in
these tests was about 0.2.

For this reason it is believed that the measured results are repre-
sentative only of rotors having similar ratios of hub diameter to rotor
diameter. The calculated results are believed to be more nearly repre-
sentative of rotors having smaller ratios of hub diameter to rotor diam-
eter and similar hub configurations. However, it is felt that the
effects of Mach number on either of the rotors or comparisons between
the rotors are valid.

Somewhat better agreement between experimental and calculated per-
formance is indicated at tip Mach numbers of 0.75 and 0.81. (See figs. 9(b)
and 9(c).) It is not known whether this better agreement is fortuitous
or whether the effect of the flow disturbances due to the hub are mini-
mized because of the presence of supercritical flow existing at these
higher Mach numbers. Symbols were used to identify the different methods
of calculating the theoretical performance.


Power Requirements of NACA 64-Series Rotor

Performance measurements.- The measured variation of thrust coef-
ficient CT with torque coefficient CQ over a range of tip Mach numbers





NACA RM L58B19


from 0.45 to 1.0 for the rotor having the NACA 64-series airfoil sections
and -160 of twist is shown in figure 11. For this rotor no compressibility
losses are evident up to and including a tip Mach number of 0.91. At tip
Mach numbers of 0.95 and 1.0 compressibility losses were measured over
the entire range of thrust coefficients covered in the tests. However,
the characteristic spreading of the performance curves is not evident
even at the maximum thrust coefficient obtained CT = 0.0068) at a tip
Mach number of 1.0. This point corresponds to a disk loading of 20 pounds
per square foot. The reduced compressibility effects measured on this
rotor, compared to the NACA 0012 rotor, are attributed to the differences
in thickness ratio and blade twist.

Profile-drag power ratio.- The profile-drag torque ratios for the
NACA 64-series rotor as a function of rotor mean lift coefficient and
calculated tip angle of attack are shown in figures 12 and 15, respec-
tively. The profile-drag torque ratios remain constant with tip Mach
number and rotor mean lift coefficient or tip angle of attack to a tip
Mach number of 0.91. At tip Mach numbers of 0.95 and 1.0 there is an
increase in profile-drag torque ratio. However, the ratio remains rel-
atively constant over the mean lift coefficient and tip angle-of-attack
range which indicates that the profile-drag torque is independent of tip
angle of attack at a given Mach number over the range of the tests.

The combinations of tip Mach number and angle of attack at which
drag divergence occurs on the 64-series rotor have been estimated from
the data of figure 15. Figure 14 presents a comparison of the experi-
mental drag-divergence Mach numbers with those indicated by two-dimensional
airfoil data. The experimental drag-divergence Mach numbers range from
about 0.08 to 0.11 higher than two-dimensional results over the range of
tip angles of attack of the tests.

Comparison with theory.- The experimental and calculated hovering-
performance curves of the NACA 64-series rotor, operating at tip Mach
numbers of 0.45, 0.95, and 1.0, are compared in figure 15. In general,
the agreement between calculation and experiment at a tip Mach number
of 0.45 is similar to that obtained for the NACA-0012 rotor. As indi-
cated in the discussion of the IACA 0012 rotor, the better agreement at
high Mach numbers may be fortuitous. The calculating procedure and cor-
rection factors used in the analysis of the 64-series rotor are the same
as those used for the 0012 rotor.

At tip Mach numbers of 0.95 and 1.0 the outboard blade sections were
operating above the limits of the available two-dimensional airfoil data.
Therefore, it was necessary to extrapolate the airfoil data to a Mach
number of 1.0. This was accomplished by using the results obtained on
similar airfoil sections, tested up to a Mach number of 1.0 in a differ-
ent wind tunnel, as a guide in the extrapolation. The two-dimensional






NACA RM L58B19


lift and profile-drag data used in the performance calculations for the
64-series rotor at tip Mach numbers of 0.95 and 1.0 are shown in fig-
ures 16 and 17.


Comparison of Compressibility Effects on the

NACA 0012 and NACA 64-Series Rotors

Figure of merit.- One method to determine the relative character-
istics of two rotors of the same solidity is to compare the variation
of their figures of merit M with the ratio of thrust coefficient to
solidity CT/a at equal tip Mach numbers. Plots of figure of merit as
a function of CT/a for the NACA 0012 and 64-series rotors are shown
in figures 18 and 19, respectively. Comparison of figures 18 and 19
indicates that at CT/a less than 0.04 and at tip Mach numbers up to
0.81 there are only minor differences in figure of merit between the
two rotors. At CT/a greater than 0.05, the 64-series rotor is superior
at tip Mach numbers of 0.81 and above. At a tip Mach number of 0.86 the
64-series rotor becomes more efficient than the 0012 rotor for all values
of CT/a greater than 0.02. Since the solidity of the two rotors is
almost identical, these differences in efficiency must be attributed to
the use of the thinner low-drag airfoil sections and to the increased
twist. At the lower values of tip Mach number the 0012 rotor would be
expected to have somewhat higher figures of merit because of the better
lift-drag ratios of the blade sections at higher lift coefficients. How-
ever, at the higher tip Mach numbers the thinner tip sections and lower
tip angles of attack resulting from the higher twist tend to delay the
compressibility effects.

Thrust-horsepower ratio.- A dimensional way in which rotors of the
same solidity can be compared is on the basis of the variation of thrust-
horsepower ratio with disk loading. Plots of this type for the NACA 0012
and 64-series rotors are presented in figures 20 and 21, respectively.
In general, as the tip Mach number increases, the maximum thrust-horsepower
ratio is reduced. At the same time the variation of thrust-horsepower
ratio with disk loading is reduced and the optimum value for a given tip
Mach number occurs at a higher disk loading as the tip Mach number is
increased. The improvement in hovering efficiency at high blade tip Mach
numbers brought about by the combined benefits of reduced airfoil thick-
ness ratio and increased blade twist is indicated by comparing figures 20
and 21.





NACA RM L58B19


Sound Measurements

It is well known that the sound emitted by rotors and propellers
increases in intensity with increases in tip Mach number (ref. 8). As
a matter of interest, sound-level measurements were made at each test
condition for the NACA 64-series rotor. The results of these measure-
ments are shown in figure 22 as the variation of sound pressure in deci-
bels with tip Mach number for several values of thrust coefficient. At
a tip Mach number of 0.45 the sound pressure varied from 82 to 94 deci-
bels for the test range of thrust coefficient. At a tip Mach number
of 1.0 the maximum sound pressure was about 122 decibels.

As a matter of interest, some calculations were made of the "Gutin"
or discrete frequency noise from this rotor by means of the method out-
lined in reference 8. The results of these calculations are shown by
the dashed line of figure 22 which is a summation of the first five-blade
passage frequencies for a value of CT of 0.006. It can be seen that
the general trends of the calculated and measured data are similar and
that the sound pressure levels agree within about 4 decibels. It should
be pointed out, however, that the method of reference 8 does not predict
random noise levels. Hence, for conditions where the random noise is
important (full-scale rotors having large blade areas) this method is
probably not adequate.


CONCLUDING REMARKS


An investigation of the effects of compressibility on the hovering-
performance characteristics of two 10-foot-diameter helicopter rotors
having solidities of about 0.10 has been conducted in the Langley full-
scale tunnel. One rotor, having NACA 0012 airfoil sections, a plan-form
taper ratio of 3/1, and -80 of twist, was tested to a tip Mach number
of 0.95 and a disk loading of 16 pounds per square foot. The other rotor
had NACA 64-series airfoil sections, a plan-form taper ratio of 5/1, -16
of twist, and was tested to a tip Mach number of 1.0 and a disk loading
of 20 pounds per square foot. As a result of this investigation the fol-
lowing conclusions were noted:

1. As the tip Mach number was increased up to 0.86, the NACA 0012
rotor having -80 of twist encountered compressibility losses at progres-
sively lower thrust coefficients. At tip Mach numbers of 0.91 and above,
compressibility losses were evident at the lowest values of thrust obtained
in the rotor tests.






NACA RM L58B19


2. No effects of compressibility were evident at any test thrust
coefficient on the NACA 64-series rotor having -160 of twist up to tip
Mach numbers of 0.91. At tip Mach numbers of about 0.92 and above,
compressibility losses were evident at the lowest values of thrust meas-
ured in the rotor tests.

5. The measured power requirements of both rotors at the lowest tip
Mach numbers are significantly more than predicted by usually reliable
theory. This effect is believed to be due to the unusually large ratio
of hub diameter to rotor diameter. The absolute values of the measured
data are therefore not believed to be representative of full-scale rotors;
however, the effects of Mach number on either of the rotors or compari-
sons between rotors are believed to be valid.

4. Calculations indicate that the drag-divergence Mach numbers of
these rotors are about 0.1 higher than would be predicted from two-
dimensional airfoil data.

5. For tip Mach numbers of 0.86 and above, the NACA 64-series rotor
had higher figures of merit than the 0012 rotor for all values of thrust-
coefficient-solidity ratio above 0.02.

6. As the tip Mach number increases, the maximum thrust-horsepower
ratio is reduced. The optimum value for a given tip Mach number occurs
at a higher disk loading as the tip Mach number is increased.

7. The maximum sound level measured in the tunnel test chamber at
a point 21.5 feet and 1400 from the vertical shaft on the NACA 64-series
rotor was 122 decibels at a tip Mach number of 1.0.


Langley Aeronautical Laboratory,
National Advisory Committee for Aeronautics,
Langley Field, Va., February 5, 1958.






NACA RM L58B19 15


APPENDIX


DERIVATION OF BLADE-TIP LOSS FACTORS


The basic equation for tip loss (i.e., effective reduction in blade
radius) due to three-dimensional flow at a propeller blade tip was devel-
oped by Betz and is given in reference 9. In the terminology of the
present report this equation becomes:


S1.3586 xA
B=lX
b F1+ 2

where

=Vt
QR

If a uniform distribution of induced velocity is assumed,



vt = DR



CT
\2




B = 1 1.386 2
b \
CT
1+ -
2


If a triangular distribution of induced velocity is assumed,

vt = 1.5QR T
fE2_





NACA RM L58B19


A = 1.06


Since CT << 1.0 and


1.586 2,


B= 1 -
b


? s CT






NACA RM L58B19


REFERENCES


1. DeFrance, Smith J.: The N.A.C.A. Full-Scale Wind Tunnel. NACA
Rep. 459, 1935.

2. Carpenter, Paul J.: Effects of Compressibility on the Performance
of Two Full-Scale Helicopter Rotors. NACA Rep. 1078, 1952.
(Supersedes NACA TN 2277.)

5. Powell, Robert D., Jr.: Compressibility Effects on a Hovering Heli-
copter Rotor Having an NACA 0018 Root Airfoil Tapering to an NACA
0012 Tip Airfoil. NACA RM L57F26, 1957.

4. Shivers, James P., and Carpenter, Paul J.: Experimental Investiga-
tion on the Langley Helicopter Test Tower of Compressibility Effects
on a Rotor Having NACA 652-015 Airfoil Sections. NACA TN 5850, 1956.

5. Gustafson, F. B.: The Application of Airfoil Studies to Helicopter
Rotor Design. NACA TN 1812, 1949.

6. Gessow, Alfred, and Myers, Garry C., Jr.: Aerodynamics of the Heli-
copter. The Macmillan Co., C.1952, pp. 72-73.

7. Rabbott, John P., Jr.: Static-Thrust Measurements of the Aerodynamic
Loading on a Helicopter Rotor Blade. NACA TN 5688, 1956.

8. Hicks, Chester W., and Hubbard, Harvey H.: Comparison of Sound
Emission From Two-Blade, Four-Blade, and Seven-Blade Propellers.
NACA TN 1554, 1947.

9. Glauert, H.: Airplane Propellers. Propellers of Highest Efficiency.
Vol. IV of Aerodynamic Theory, div. L, ch. VII, sec. 4, W. F. Durand,
ed., Julius Springer (Berlin), 1955, pp. 261-266.






NACA RM L58B19


TABLE I


PHYSICAL CHARACTERISTICS OF NACA 0012 AI'[D NACA 64-SERIES ROTOR BLADES



Iten NACA NACA
0012 rotor 64-series rotor

Weight (blaie alone), lb ...... .12.17 15.95
Fadius, ft . ..... 5.073 5.068
Chordwise center of gravity
(blade alone), percent chord 24.19 25.26
c, in . ... .8.98 8.77
Tip -hord, in . ... ..16. 5.84
Fot chord (extrapolated to center-
linr rotation), in ........ 16.00 15.77
Epanwise center of gravity (blade
alone). percent span .. 42.60 51.90
TwiZt,. . ..... -8 -16
j . 0.095 0.092
7' includingg hub) ... 0.45 0.40
I (including hub). slug-ft2 .... 2.75 2.88
Airfoil section:
Blaie tip . .. N ACA 0012 NACA 64-006
Elidem root (extrapolated to
-tntcr-line rotation) .... NACA 0012 NACA 642-015





NACA RM L58B19


L-92295
Figure 1.- NACA 0012 airfoil rotor blade used in hovering tests.







NACA RM L58B19


ID

m


Ij
I I
I I
/ I
/ I


I I I I I I


Co -d' O 0
1-I -I 1- r u-I


OD D


0




*--P

0




















-rI
0


k
m





CD




a,-
m








0



t0




1U





-P
.rl












-4.
I1)
CM












10
cn
rJ




-,--
r

rl a




-
'*.b
.r


pJoo lua oJad 'oTvIe ssauoiqM ap'eBg


/I
/I





I


I
-d' CM C 0
1-4 r-I I-i ,






NACA iEM L58B19 19











0
O


0





o
















.--
o ... .... ....... ... .... .. ..


















..:








= O


ED
.,. .. .




ma0



to







NACA RM L58B19


Figure 4.- Sketch showing relative


r


size of rotor head and blades.







NACA R4 L58B19 21





in






Soo
0

-1.





"-




o .






-, --( (D..DO l -


o a
0 ON








So0'



8 dH










o, L-
4 <\ 0


-\_ (< -- D-I- -I- -L -
Ll 0























r-
q- o .na









I 0
08









c-
So4-)





-O.-





o 0\
0 \C0













r.4
0L.0
CC i---0 c-C- to -)--------N 0 >








o 00 0 0 6 0
mmrmmS^^mmg
'-^ \- \ o I <







NACA RM L58B19


CO ( COD 1 O
oM C,- ,--f


S$'O = ^ Pe pGOjnsEam o'b)

pajnspaxm oOb.


qH
0
0











I
r*H
4-,





0





-l
C*

*H
P4
0





-4












rO
ar *



E,
. rJ








IO
o Co











0 C0






-c-4
aI









pI
-1
OHr-







NACA RM L58B19


CD CM CO d 0
oN N r4 r-l r


gS'o = W T peianseam o'b,

paJnslam o'?0


bO
0,



" l
C',


0
0
-I
4-,





0






I
4J




(U


01
MO
0c











0

O
0







Ui
cd
4-1
0
0



HZ



4-


Crz
0
O

ai

o

^N

?4







NACA RM L58B19


O O co C C3
U S S S S


tt
U)




!' -


an,'UJOjAjAp pjp joJ jaqumnu Lqouei


H
O
*-

0


cO

O
o










53


I)
Sr-
-rl
!
M0
* h













U
IC








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CJ
u









> ld
a





cd a



















S,- 4
r-.
0 -







1


O M


S+3



-4



dL4
hp i
*o E/







NACA RM L58B19


"' \ _




V' '- r
..

F '\


1: .\-









".')
-i.
_ ^ ^ _ _


_ ^ _ _


'*' <


'4 0 m m t co
1 8 0 0 0
o C, 9 0 C' 0


0 0 0 0
* c: C 0


:

I ,
U1 r

-3 -a
o o
u k
ra
U U


I I I I I I I I I Ii I I I I I T T F I I I


0
4-,



I

0
*d
o
o




II
Ob
c.4
0 -0

OJ





0 0
P4
0 0
O)


o





C-J



0-
o





U)
r-l ,
















nc
RO
C .,










S0












0
IC

ch






E -i








NACA RM L58B19


ITIIIK iiiiTI i


- --


0











o C
C o ra r -
u a
CD







,-z E-

0 rl L
p- 0 -. Q



0~ -
.-' OL OL

C 0I




00 '5c


0 3 c C" co
r- 0 0 0 0
0 -8 0 C' 0 0
* J


o 8 8 0
o O o


0
N
'-j


0
0







o
0
r-4


0
0









0
10













0o
0
0


0
0
0


0

0

0
0








0
0



0
0
0


0
0
0


0
0
0
*i


.



d


c, I






NACA RM L58B19


.011


.010


.009


.008


.007


.006

CT
.005


.004


.003


.002


.001


0 .0001 .0002 .0003 .0004 .0005 .00)06 .0007 .0008 .0009 .0010 .0011
CQ



(c) Mt = 0.81.


Figure 9.- Concluded.


-- -----

-- .- E-'erimental




S No tip Mach relief
S Tip Mach relief 8 = I _
-- -- No tip Mach relief 1 x-p -
o Tip Mach relief B1- i
__ /._ __ __ G TipMac rlie B 1- 2rr _






!8 NACA RM L58B19






a, deg
[1.0 --------------------------
1.0_
8.8





.6





.L6


.4 -- --- --- -- ---- .






0-_ -- __ -- -- -_- -
.6 ^ S- -











00
.2 .3 .4 .5 .6 .7 .8 .9
Mach number


(a) cl.

Figure 10.- Twordimensional NACA 0012 airfoil data used for theoretical
performance calculations.







NACA RM L58B19 29


.14



.13

- -- I -

.12

-----___-I---

.11



.10





.00
--/-- -7- -



- .07--





.06



.05






.03




a, deg
.02
8 0

.01--- -- -
.03 ------_ L _-^-- ----
o -




n __ _ ___ _


.1 .2 .3 .4


.6 .8 .7


.B .9


Mach number



(b) Cd,0



Figure 10.- Concluded.























0-- --- --










q.
---- --------------------------
3 ,,
----------------------------------------------------- ---- ----------------------














0
oooooo






LD S








""b I-



-- -- -
S i i j,
__ __ \ __ __ __ __ __ ^p *a-

0
Ci C 0


0 0 0 0 0 0
0 0 0 0 0 !


NACA HM L58B19


I
C)



O
00





ob
01
H c1






a,A

*H 0


rl ;


0 0

MO





OO
P *
4J



6.

















1
0
0* 0















c l
I"
$4,













0

$4
4-U














O

00




0 l
$4







O
I -I
0)







to
%-q




r.4






NACA RM L58B19


-4I
a,
0


LO


Ix



0


II 01

___














^ll===


9!'O = l, iE peJnspem o0b,

pajnsem o C03


0
o
o

a,




0




4l O

0 1




UO





OO
H\
*r-I



a)













0
-4-)
o

























r-I
Ci:I





c +J


*H
<- d
eM '-
(Ll t-
0 0
u
OP-
til >
*H
H <
1

1-

0l








































































0 (D CM


S"'O = 7I Ip pajnselp 0o'3

pajrnstam 0'0


NACA RM L58B19


O
4-4
0
0







.I



I
4-
0











o
4w-











.rd
P *i















O
E30
0
I
c\








Ci

*cd 0
0U
'd o









P4
H
O
P




a.






NACA RM L58B19


-3-- S














0 o O

-A1 0 0



aou9Sj.AIP SEJP jo Jequmu 40OIEI


c,



r-I
'l

1-1


C'




c -I
Qd



4, *0-
a a





*H
O
04-

id


\D



o
0











'do
1o_


U
II0







d
1 2













00)


co









ai












r-I
to


























4-3
a

























-1I







34 NACA RM L58BI9










a)



I


44





- -, --
0I 0



o o
- ------! -----







O 0

o b
S \ d II







a \ 0 IO o
-------------------- |--- ,











0 O0



0-
SI -

CdO
'. '. 0" |lj













___ m ._:_ g Q

0
C-
0




c0 0
-- ---c- a -
















id



9O
_r. 0 0 0 0T
















\-
LC H
,F
-- -- -- -- -- -- -- -- -- ^ -- -- -- '






NACA RM L58B19


.0005 .0006 .0007 .0008 .0009 .0010


(b) Mt = 0.95.


Figure 15.- Continued.







NACA RM L58B19


.006



.005



.00-- -- -
.007~ ~ ~~~~~~~~~~~~~~~~~ -- -- -- -- -- -- -- -- -- -- -- -
(^

__ __ ^ __ _
.06 - 7 z -/- -


- -- -- --


- Experimental

Calculated
Condition Tip loss factor
O No tip Macn relief
O Tip Mach relief B = 1 -
O No tip Mach relief
F Tip Mach relief Bl1-1 -6 2
b FJ1.CT/2

S1-


0 .0001 .0002 .0003 .0004 .0005 .0006 .0007 .000 .0009 .0010


C(c) M

(c) Mt = 1.O.


Figure 15.- Concluded.


.001








NACA RM L58B19













u .6






2. /
a--------------------------------------------------------,----_





4- /



0--
0-




Extrapolated





.0


.0

a, degrees

a .4 I
0.4




S1.---------------------- ----------------









---.4
.6----------------------------------i--------------i-----






























.1 .2 3 .4 .5 .7 .8 .9 1.0

Mach number



Figure 16.- Two-dimensional NACA 64-006 airfoil data used for theo-
retical performance calculations.
.4
C,
0-


g o -- o ^ =-= -- -- -- -- -- -- -




Mah ume

Fiur i6- -w-di sinl- AA-I----ar -il-aa-se-fr- he-







NACA RM L58B19


.4 .5 .6 .7 .8 .9 1.0


Mach number


Figure 17.- Two-dimensional NACA 64-008 airfoil data used for theo-
retical performance calculations.


-.4
.1





NACA RM L58B19 59





r-I


o




01 0
0







/ o
co_ 0


__ ____ _____+_____ 0












IIO 0%
9 Q .o
i






// / E 0
SIC I
m o









0 Z
U) 0




--- V ^ -v-0 ,















Son U

co (0o



71 'WQJi JO QjnIc2I
I \ I I I






NACA RM L58B19


OD (0
* 0


W '.J9Ta JO ajn.ld


0
0
r%





I

0
4-
0




I






F4)




0
ri









P-H
*1



























(U
0












4)
I






























p0
H
fS


I:
(U










?l





NACA RM L58B19


14

13
Mt
12 00.45
o, 0.63
11 0 0.68
0 V 0.73
V 0.77
10 0_ _- D 0.81
0 0.86
9 0o.91
P 0.95





06 4

5 .

o ti-t_----=- --0-- -- --



2




0 2 4 6 8 10 12 14 16 18 20

Disk Loading, lb t2


Figure 20.- Effect of disk loading and Mach number on power loading of
a 10-foot-diameter rotor having NACA 0012 airfoil sections and -80
of twist. a = 0.095.






NACA RM L58B19


0 2 4 6 8 10 12 14 16 18 20

Disk Loading, lb t2


Figure 21.- Effect of disk loading and Mach number on power loading of
a 10-foot-diameter rotor having NACA 64-series airfoil sections and
-16 of twist. a = 0.092.











































ra

'r-



0






ctO
03




rQ0






0d)
m o


qp 'emansejd punoS


IIACA P11 L58B19


4-.
C~





00


I

I


0



u
,-4
-d
U
u


4-)





0
a,















0U
.





.14-






0
*g -























PI
mrt














cu
cu


E- U)
0 r-
0 0
* 0-
*


I I


NACA Langley Field, Va.


I I
































































































































































2











UNIVE. SITY OF FLORIDA
iAHlllljki1UiMIIII .illlllIII li.
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UNIVERSITY OF FLORIDA
DOCUMENTS DEPARTMENT
120 MARSTON SCIENCE UBRARY
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