Effect of Advanced Rotorcraft Airfoil Sections on the Hover Performance of a Small-scale Rotor Model

34 Pages·2003·7.18 MB·English

https://ntrs.nasa.gov/search.jsp?R=19890014893 2019-04-12T20:34:19+00:00Z
NASA
Technical
Paper
2832
AVSCOM
Technical
Paper
Effect of Advanced
88-B-001
Rotorcraft Airfoil
1988
Sections on the Hover
Performance of a
Small-scaleR otor Model
Susan L. Althoff
Aerostructures Directorate
USAAR TA-AVSCOM
Langley Research Center
Hampton, Virginia
National Aeronautics
and Space Administration
Scientific and Technical
Information Division
Summary the NACA 0012 airfoil section on many helicopter
rotors (ref. 3). The 12 percent thickness is a com-
A hover test was conducted on a small-scale
promise between the thin section desired for good
rotor model for two sets of tapered rotor blades. The
compressibility effects and the thick section required
baseline rotor-blade set used an NACA 0012 airfoil
for high lift. The symmetric section also generates
section, whereas the second rotor-blade set had ad-
little pitching moment and is easy to manufacture.
vanced rotorcraft airfoils RC(4)-10, RC(3)-10, and
With the recent developments in composite man-
RC(3)-08 distributed along the radius. The experi-
ufacturing technology, it has become possible to build
ment was conducted for a range of thrust coefficients
rotor blades with an aerodynamically tailored airfoil
and tip speeds, and the data were compared with
distribution along the radius (refs. 6 to 8). This tech-
the predictions of three analytical methods. The
nology allows the rotor designer the freedom to select
data show the advantage of the advanced airfoils at
the optimum airfoil as a function of radial location
the higher rotor thrust levels; two of the methods
along the blade, thus increasing the performance of
predicted the correct data trends.
the rotor. The benefits to the vehicle from optimizing
the rotor airfoils include a gain in maximum vehicle
Introduction
speed, an increase in gross weight, and more maneu-
The helicopter rotor airfoil is required to oper- verability (ref. 6). With these types of benefits avail-
ate in a complex rotating aerodynamic environment. able, there has been increased interest in designing
In forward flight, the rotational velocity of the ro- rotor airfoils.
tor blades and the free-stream velocity combine to A series of rotorcraft airfoils (the RC series) was
produce periodic loading and unloading of the rotor designed by U.S. Army researchers at the NASA
blades. When the blade is advancing into the flight Langley Research Center as part of an effort to
direction, the local velocity is high and the blade pro- design an advanced rotor-blade geometry with im-
duces high lift. However, when the blade is retreating proved hovering performance and no degradation of
from the flight direction, the relative velocity is low; forward flight performance. The design philosophy
the local velocity may even be reversed with respect is discussed in reference 8; the rotor is optimized
to the airfoil on the rotor. Therefore, the retreat- by using combinations of planform, twist, and air-
ing blade operates at a lower lift than the advancing foil distribution to improve rotor performance as pre-
blade. This imbalance in lift causes a rolling moment dicted by simple computational methods. The hover-
about the rotor hub. In order to balance the lift and prediction method used in this design process was a
“zero” the moments, the advancing blade is usually momentum blade-element analysis based on the the-
operated at lower angles of attack than the retreat- ory outlined in reference 9. The forward flight analy-
ing blade, whereas the retreating blade is operated ses were based on the work in references 10 to 12 and
at angles of attack near the stall limit. on the rotor flight-simulation code (C81) discussed in
These variations in local blade velocity and angle reference 13.
of attack place several conflicting restrictions on ro- The airfoils that were designed for this applica-
tor airfoils. (See refs. 1 to 4.) The outboard portion tion were the RC(3)-12, RC(3)-10, and RC(3)-08. In
of the advancing blade operates in a low-lift, high- the RC(x)-xx format, the “RC” designates a rotor-
velocity region. Therefore, the airfoil is required to craft airfoil, (x) indicates the sequential number, and
have a high drag-divergence Mach number at low lift the “-xx” provides the thickness in percent of chord.
to avoid excessive drag. The data in reference 2 sug- These airfoils were tested in two-dimensional wind-
gest that a symmetric or slightly cambered thin air- tunnel tests, and the results are documented in ref-
foil section would have this characteristic. However, erence 14 by Bingham and Noonan. These airfoils
the inboard section of the retreating blade requires exhibited high drag-divergence Mach numbers and
high-lift capability at low velocity. It is shown in ref- low pitching moments, but they had average high-lift
erences 2 and 5 that a thick, cambered airfoil would characteristics. The RC(4)- 10, designed and tested
provide this capability. The airfoil also should gen- by Noonan, proved to have good high-lift capabil-
erate little or no pitching moment at all operating ity, but the two-dimensional wind-tunnel test results
conditions in order to avoid aeroelastic instabilities have not been published at this time.
and high control loads. Even though a symmetric The RC(3)-12, RC(3)-10, and RC(3)-08 airfoils
airfoil generates little pitching moment, a cambered were used on the wind-tunnel model rotor of ref-
airfoil will produce a moment, as the data in refer- erence 15, and the RC(3)-10 and RC(3)-08 airfoils
ence 5 show. were used during the test of reference 16. The
A compromise among these conflicting require- RC(4)-10, RC(3)-10, and RC(3)-08 airfoils were used
ments for rotor airfoils has resulted in the use of on the model rotor in reference 17. In these test
programs, performance gains were measured for the L rotor lift force, lbf
rotors with the RC-series airfoils. The gains were
Mtip rotor tip Mach number
seen in terms of increased rotor efficiency in hover
and power reductions in forward flight. However, Mz rotor torque, ft-lbf
the effect of only the airfoils on rotor performance
NR€! Reynolds number, pVc/p
could not be determined from these tests because the
advanced airfoil rotor blades also included changes R rotor radius, ft
in planform and twist distribution from the baseline
T radial distance along blade, ft
blade.
+ +
The purpose of the present investigation was to T rotor thrust, (L2 D2 Y2)lI2,
determine the hover-performance benefits that are lbf
gained solely through use of the advanced rotor air-
V local velocity, ft/sec
foils. Therefore, a hover study was conducted using
rotor systems for which the only difference was airfoil rotor tip speed, ft/sec
Kip
shape. This hover test used a small-scale, four-bladed
rotor model with two tapered rotors. The reference x,Y , Cartesian coordinates
rotor used an NACA 0012 airfoil, whereas the sec- Y rotor side force, lbf
ond rotor had advanced rotorcraft airfoils RC(4)-10,
coefficient of viscosity,
RC(3)-10, and RC(3)-08 distributed along the blade P
slugs/ft-sec
span. The rotors were identical in planform and
twist distribution, and they were tested over a range atmospheric density of air,
P
of thrust coefficients and tip speeds. The experi-
slugs/ft3
mental data were compared with the hover perfor-
mance predicted by three analytical methods. One of torque-weighted solidity, 4cq/rR
UQ
the methods used a momentum blade-element analy-
R rotor rotational speed, rad/sec
sis, whereas the other two methods used a free-wake
lifting-surface analysis.
Abbreviations and acronyms:
Symbols
HOVER lifting-surface hover-performance
The data in this report were measured in U.S. code
Customary Units and are referenced to the shaft axis
LSAF Lifting-Surface Aerodynamics
system shown in figure 1.
and Performance Analysis of
rotor profile drag coefficient Rotors in Axial Flight
rotor torque coefficient, ROBIN generic fuselage shell
Mz/prR3( R R)2
RTC rotor test cell
rotor torque coefficient at zero
2MRTS 2-meter rotor test system
thrust
rotor thrust coefficient, Model and Test Description
TlprR 2(
The test program was conducted in the rotor test
local blade chord, ft cell (RTC) at the Langley 14- by 22-Foot Subsonic
Tunnel. The RTC is a high-bay area that is 69 ft
torque-weighted equivalent blade
high by 42 ft wide by 48 ft long with a steel chain
link fence around the walls; it is arranged specifically
for the buildup and testing of powered rotor models
in hover. Two walls of the RTC have louvers that
rotor drag force, lbf
can be opened to alleviate some of the recirculation
rotor figure of merit, of air from the hovering rotor. The ambient dry-bulb
temperature, dew point, and pressure are measured
by instrumentation located in the RTC. The rotor
hub was located 1.74 rotor diameters above the floor
alternate rotor figure of merit,
c of the RTC on a post mount. The model is pictured
c3, /2/ ( Q -cQ,)fi mounted for testing in the RTC in figure 2.
I
I The model system used for the experiment was were constructed using a graphite-epoxy D-spar with
the 2-meter rotor test system (2MRTS) with a tungsten leading-edge weights. The baseline blades
generic fuselage shell (ROBIN). The ROBIN fuselage had a foam trailing edge, whereas the advanced airfoil
I shape is detailed in reference 18, and the 2MRTS is blades had a balsa trailing edge. Both sets of blades
I
I documented in reference 19. A brief description of were very stiff, as there was no attempt to match the
the 2MRTS is repeated here for the convenience of aeroelastic characteristics of full-scale rotor blades.
the reader. The test procedure was as follows: the rotor rota-
The 2MRTS (fig. 2) is a drive system that consists tional speed was established and maintained, a col-
.
of a 29-hp electric motor, a go', two-stage trans- lective blade angle was input, and the shaft angle was
mission with a 4-to-1 gear reduction ratio, and a set to zero. The flapping of the rotor blades was mon-
four-bladed rotor hub. The motor is water cooled itored and was maintained within a tenth of a degree
by a chiller system, and the transmission is cooled of zero during the test. When the model was at a test
and lubricated by an oil pump and vacuum system. point condition, a data point was taken. After the
The motor, transmission, and rotor hub are all sus- data had been collected, the collective angle was in-
pended on a gimbal that allows the system to pitch creased and the procedure was repeated. The upper
and roll; however, both the pitching and rolling mo- limit of the thrust sweep was determined by motor
tions are damped by mechanical dampers in order to power and temperature; i.e., when the drive motor
avoid a ground resonance condition. The rotor hub is exceeded critical temperatures, the thrust sweep was
a four-bladed, fully articulated hub with coincident terminated. Thrust sweeps were made at rotational
flap and lag hinges. The lagging motion of the blades speeds from 2000 to 2500 rpm for a range of thrust co-
is damped by viscous dampers. efficients CT from 0 to approximately 0.010 for each
The collective and cyclic angles of the rotating set of rotor blades.
blades are controlled through a conventional swash The data were acquired through a static data ac-
plate mechanism. The swash plate is positioned quisition system that sampled the data. Four hun-
by three actuators that are remotely controlled at dred data measurements were acquired in 8 sec for
an operator's console. The control of the actuators each data point presented. Each thrust sweep was
is accomplished in a feedback-loop control system repeated three times in order that some measure of
using a minicomputer to calculate the position of the the data accuracy and scatter could be determined.
actuators for the desired rotor-blade pitch angle. One condition was tested on 2 different days in order
Forces and moments are measured separately on to ensure that daily variances in temperature, hu-
the rotor and fuselage by two six-component, strain midity, and pressure were being properly corrected.
gauge balances. Other instrumentation of the system The ambient winds in the RTC were measured daily,
includes three strain gauges on the rotor blades to and data were acquired only when the wind condi-
measure bending moments, potentiometers to mea- tions were in the range from 0 to 1 knot. The preci-
sure flapping and lead-lag motion, a digital rotational sion of the data measurements is estimated from the
speed encoder, and thermocouples to monitor critical repeatability tests to be f2.5 x in thrust co-
temperatures in the motor, transmission, and swash efficient C~,f0.01in rotor figure of merit FM, and
plate. f3.0 x in torque coefficient CQ.
The two sets of rotor blades used in this investiga-
tion are described in figure 3. The blade sets differed Description of Analytical Methods
only by airfoil sections. The baseline set of blades
was constructed using NACA 0012 airfoils. The sec- The selection of the hover analytical methods
ond set of blades had advanced airfoils specifically used for this investigation was based on the expe-
designed for the rotating environment (RC(4)-lO, rience with hover-performance codes in reference 20
RC(3)-10, and RC(3)-08) distributed along the blade and the desire to investigate the capabilities of the
radius, as shown in figure 3. Smooth transitions were relatively new lifting-surface free-wake hover codes.
accomplished between airfoil sections over 5 percent It should be recognized that there are many differ-
of the blade radius. The three airfoil sections, to- ent prediction methods under development; the codes
gether with the NACA 0012 section, are shown in used in this effort are only a representative sample.
figure 4; the coordinates for the advanced airfoils are It was found in reference 20 that a simple momen-
given in tables I, 11, and 111. Both sets of rotor blades tum blade-element analysis was able to predict hover
had a tapered planform (with the 3-to-1 taper ratio data trends. It was also found in that study that the
initiating at 75 percent of the 32.5-in. radius) and a simple free-wake analysis of reference 21 was unable
-13' linear twist distribution. The thrust-weighted to converge on a solution, whereas the prescribed-
solidity of the rotor was 0.0977. The rotor blades wake lifting-line analysis of reference 22 predicted
3
the wrong trends for the hover performance of ta- model discussed in reference 25. The calculations for
pered rotor-blade configurations. Therefore, it was the rotor performance are similar to those used in
decided to use two sophisticated lifting-surface free- HOVER; i.e., the lift is calculated using the circula-
wake analyses as well as the simple momentum blade- tion distribution, and the drag and torque calcula-
element theory for this investigation. tions rely on two-dimensional airfoil data.
The momentum blade-element analysis used for
Presentation of Results
this investigation is a code based on the equations
developed in reference 9. The rotor disk is assumed The experimental data are presented in tabular
to be an infinitely thin actuator disk that allows no and graphical format. The values of thrust coefficient
velocity discontinuities across the disk. The induced CT,t orque coefficient CQ, and figure of merit FM
vertical velocity through the rotor disk is found by for the baseline and advanced airfoil rotor blades
equating the change in the momentum of the air to can be found in tables IV and V, respectively. The
the thrust of the rotor. Then, the rotor blade is con- analytical comparisons to the experimental data are
sidered to be a combination of individual airfoil seg- shown graphically. The presentations are made using
ments, and the forces on each section are calculated. the following figures:
The airfoil section properties are obtained from ta-
Figure
bles of two-dimensional airfoil data. The total ro-
tor forces are determined by integrating the segment Basic aerodynamic characteristics of
forces over the rotor blade. rotors for several Mtip values . . . . . . . 5
The second analytical method used in this in- Comparison of aerodynamic characteristics
vestigation was the lifting-surface hover-performance of baseline and advanced rotors . . . . . . 6
code (HOVER). The method is documented in ref- Power required to hover for baseline and
erence 23. The rotor blades are modeled using a advanced rotors at Mtip = 0.639 . . . . . 7
vortex-lattice panel distribution, whereas the wake Comparison of an alternate figure of
is represented by discrete vortex segments. The ro- merit for the baseline and advanced
tor wake geometry is determined through two sets of airfoil rotors at Mtip = 0.639 . . . . . , . 8
iterations. The first iteration set consists of establish- Comparison of prediction methods with
ing a wake geometry from a set of prescribed-wake experimental data at Mtip = 0.639 . . , . 9
equations and matching the circulation solution of Comparison of three analytical methods
the rotor blades. In the second iteration set, the ro- at Mtip = 0.639 . . . . . . . . . . . . 10
tor wake is allowed to distort as a free wake from Effect of two-dimensional airfoil data
the generalized wake solution in response to the in- tables on analytical results for base-
duced velocities from the rotor blades and from self- line rotor at Mtip = 0.639 . . . . . . . . 11
induced velocities in the wake itself. After a wake ge- Comparison of analytical predictions for
ometry has been determined, HOVER calculates the several values of Mtip . . . . . . . . . . 12
circulation induced by the wake at the rotor-blade Effect of user input for radial position
surface through use of the Biot-Savart Law. Once T/Ro f maximum circulation on
the circulation is known, the lift can be calculated performance calculations by HOVER
from the Kutta-Joukowski Law. The drag is calcu- at Mtip = 0.639 . . . . . . . . . . . . 13
lated by combining the induced drag with the airfoil
Results and Discussion
profile drag from two-dimensional airfoil data. The
rotor torque is calculated from the integrated rotor-
Experimental Results
blade drag. There is no model for stall or separation
in HOVER. Compressibility effects on lift are cal- The results of the experiment are shown in fig-
culated using a Prandtl-Glauert correction, whereas ure 5 as plots of figure of merit FM versus thrust
the effects on drag are assumed to be contained in coefficient C,. Figure of merit is an efficiency term
the two-dimensional airfoil data tables. that expresses the ratio of the ideal power required
The third analytical method used in this in- for hover to the actual power required for hover.
vestigation was the Lifting-Surface Aerodynamics The data for the baseline rotor are shown in fig-
and Performance Analysis of Rotors in Axial Flight ure 5(a) for several tip Mach numbers. The data
(LSAF) which is discussed in reference 24. The rotor show very similar trends below CT = 0.007; however,
blades and wake were represented as vortex boxes, or at the higher thrust coefficients, there is a decreas-
lattices. The version of LSAF that was used for this ing efficiency trend with increasing tip Mach num-
study incorporated a velocity-coupled wake model ber. This trend can be attributed to the decreased
into the program in addition to the prescribed-wake maximum lift coefficient and increase in drag of the
4
baseline NACA 0012 airfoil that is associated with Figure 7 is a plot of CQ against C$I2 for the
an increase in Mach number. baseline and advanced airfoil rotors at a tip Mach
The performance of the advanced airfoil rotor is number of 0.639. The CQ, for the baseline 0012
shown in figure 5(b) for the same three tip Mach rotor was found from the intercept of a second-order
numbers. The effect of Mach number on the ad- least-squares curve fit to be 0.00012, and the CQ,
vanced airfoil rotor is not the same as that observed for the advanced airfoil was 0.00013. In figure 8,
for the 0012 airfoil rotor. This difference in behav- the alternate figure of merit (FM)* is plotted as a
ior may be attributed to the different Mach number function of CT for the baseline and advanced airfoil
characteristics of the advanced airfoil sections. From rotors at the tip Mach number of 0.639. The data are
the two-dimensional airfoil data in references 14 and not plotted for the lower thrust coefficients because
26, it is known that the outboard advanced airfoils the CQ measurement became approximately equal
experience an increase in maximum normal-force co- to CQ,, and (FM)* became meaningless. When
efficient with increased Mach number in the range of compared using this alternate figure of merit, the
Mach numbers tested, whereas the maximum lift of advanced airfoil rotor was found to have a higher
the 0012 airfoil decreases in this Mach number range. (FM)* than the baseline 0012 rotor over this CT
The experimental data for the baseline 0012 ro- range. This effect indicated that the advanced airfoil
tor and the advanced airfoil rotor are compared in rotor had greater efficiency in producing lift than
figure 6. As was evident in the earlier figures, the the baseline rotor, although the advantage of the
advanced airfoil rotor performs more efficiently than advanced airfoil rotor in terms of overall efficiency
the baseline 0012 airfoil rotor at the higher thrust (induced and profile drag effects) occurred at the
coefficients. Torque coefficient CQ is plotted against higher thrust coefficients. (See fig. 6.) From this
thrust coefficient CT to provide an indication of the analysis, it can be concluded that if the profile drag
power required by the rotor for the highest and low- of the advanced airfoil rotor could be reduced, the
est tip Mach numbers. advanced airfoil rotor would have an improvement in
The rotor efficiency term, Le., the figure of merit efficiency (FM) over the baseline rotor for the entire
FM, has been traditionally defined (ref. 9) for hover- thrust range.
ing rotors as
Analytical Results
The prediction of performance by analytical
methods can be used for several different purposes.
This expression relates the lowest possible power re-
For example, a rotor designer needs an analysis that
quirea for hover with an ideal rotor to the actual
will predict the correct trends in performance for a
power required for hover. This definition assumes
systematic parametric study. Another use for the
that the ideal rotor has zero power, or torque, re-
analyses is to predict the level of performance that
quired as a result of the effects of profile drag. Since
the operating rotor actually does experience the ef- would be expected for a given rotor. Figure 9 shows
the ability of several analytical methods to predict
fects of profile drag, to compare only the torque pro-
performance trends, whereas figure 10 compares the
duced by the induced drag requires that the torque
ability of the analytical methods to predict perfor-
generated by the rotor profile drag be removed from
mance levels. In all cases, a tension spline was ap-
the data. An approximation of the torque due to the
profile drag CQ, is obtained by plotting CQ against plied to the analytical predictions to produce the
curves shown in the figures, and airfoil data obtained
C$I2 and extrapolating the data to the zero thrust
at full-scale Reynolds numbers were used to make the
condition (ref. 27). The result is only an approxima-
predictions.
tion of the torque due to profile drag effects because
It can be seen in figure 9 that both the momen-
the rotor is twisted; therefore, even at the zero thrust
tum analysis and the HOVER analysis predicted an
condition, a small portion of the measured torque is
improvement in performance for the advanced airfoil
due to induced effects. However, it is the best ap-
rotor. However, both methods predicted that the
proximation to the true profile torque effects that
improvement occurred at a higher thrust coefficient
can be made from these data. When the torque due
than was measured in the experiment, as expected
to profile drag is removed, an expression for an alter-
because of the use of airfoil data obtained at full-
nate figure of merit that compares only the induced
scale Reynolds numbers. The LSAF method predic-
power effects is obtained:
tion, which is shown in figure 9(c), exhibited signs
of numerical instability at thrust coefficients greater
than 0.0102.
5
In figure 10 it is shown that the HOVER analysis Mach number. Both the momentum analysis and
predicts the figure of merit levels of the experimental the HOVER methods predicted the t,rend that was
rotor model more accurately than the other two measured for the baseline 0012 rotor. However,
methods for both the baseline 0012 rotor and the the momentum analysis predicted a change in rotor
advanced airfoil rotor. In order to investigate the performance for the advanced airfoil rotor that was
effect of using the full-scale airfoil data, a sensitivity not present in the experimental data as a result
study of the effects of the two-dimensional airfoil of changing the tip Mach number. HOVER also
tables was conducted using the momentum analysis predicted some effects of changing the tip Mach
and HOVER methods. number for the advanced rotor, especially at a thrust
The two-dimensional data for the NACA 0012 coefficient of 0.011.
airfoil were obtained from tests in the Langley 6- The values calculated by the HOVER and LSAF
by 28-Inch Transonic Tunnel for two Reynolds num- performance codes were affected greatly by user in-
ber ranges (ref. 26), one representing model-scale put. During this investigation, it was found that the
and the other full-scale helicopter Reynolds numbers. LSAF calculations were sensitive to user inputs such
The lower Reynolds number range was 0.7 x lo6 to as number of wake iterations, increment in thrust
1.5 x lo6, and the higher Reynolds number range level, starting values of thrust and control angles, and
was 3.0 x lo6 to 6.6 x lo6. The advanced airfoil wake azimuth increments. The program was not able
two-dimensional data were available only for repre- to converge on a solution for most of the input vari-
sentative full-scale Reynolds numbers. The experi- ations that were attempted. The tapered planform
mental rotor operated at a tip Reynolds number of of the rotor blades may have been a factor in the nu-
0.39 x lo6 and an inboard maximum Reynolds num- merical instability of the LSAF code; the blade-load
ber at 75-percent radius of 0.89 x lo6. In figure 11, distribution due to the taper may have triggered a
the performance predicted by the momentum analy- computational instability between the wake geome-
sis and HOVER methods is shown for different two- try and the circulation solution.
dimensional airfoil tables for the baseline 0012 ro- This type of problem was also apparent in the
tor at a tip Mach number of 0.639. Calculations HOVER program. In the HOVER input, the user
are shown for both the higher and lower Reynolds was allowed to specify how the radial position of
number airfoil data, as well as calculations performed maximum circulation is determined through an in-
using a modified high Reynolds number data table. put parameter. The effect that this decision may
The modification to the data table consisted of in- have on the values calculated by HOVER is shown
creasing the drag coefficient in the table by the dif- in figure 13. HOVER was executed for three cases.
ference between the Cd, in the data table and the In the first case, the position of maximum circula-
cd, determined from the data of this test (where tion was held at a constant radial position through-
out the prescribed-wake iterations of HOVER, but
cd, = ~CQ,/CTQ)T. his modification was selected
the program was free to move the position when it
based on reference 27 where it was suggested that
started the free-wake geometry perturbations. For
the increment in performance between the model-
the second case, the position of maximum circula-
scale and full-scale values was due to the difference
tion was held at a constant radial position of 75 per-
in rotor profile drag.
cent throughout the entire prescribed-wake and free-
As was expected, the calculations by both meth- wake geometry calculations. The third case allowed
ods indicated that the use of the lower Reynolds num- HOVER to have the freedom to determine the po-
ber data table decreased the predicted performance sition of maximum circulation during each wake ge-
of the rotor. In the case of the momentum analy- ometry iteration. As can be seen in figure 13, for
sis, the predicted FM was closer to the experimental both the baseline 0012 and the advanced airfoil ro-
data, whereas for HOVER the lower Reynolds num- tors at a tip Mach number of 0.639, the calculations
ber predictions were farther from the experimental were closest to the experimental data when the ra-
data than the full-scale Reynolds number predictions. dial position was held at a constant value through-
The HOVER program appears to be able to predict out the prescribed-wake calculations and was then
the correct model-rotor performance level only when allowed to change position during the free-wake iter-
using the airfoil tables obtained at full-scale Reynolds ations. The question remains as to how the constant
numbers. This result is fortunate because the full- radial value should be obtained: in order to be consis-
scale airfoil data are generally the only data available tent, Some criteria should be imposed for determining
to the user for advanced rotorcraft airfoil sections. this value. For this investigation, the radial position
Figure 12 shows the ability of the analyses to of maximum circulation that was calculated in the
predict the variation in rotor performance with tip momentum blade-element analysis was used as the
6
input into HOVER. This combination of the momen- Keys, C. N.: Volume II: Performance Prediction of
tum analysis and HOVER methods provided good Helicopters.
agreement with these experimental data. Further 2. Davenport, Franklyn J.; and Front, John V.: Airfoil Sec-
studies are required to determine whether this combi- tions for Helicopter Rotors-A Reconsideration. Pro-
ceedings of the Twenty-Second Annual National Forum,
nation will prove to be a consistently reliable method.
American Helicopter SOC.,I nc., May 1966, pp. 29-44.
3. Johnson, Wayne: Helicopter Theory. Princeton Univ.
Summary of Results
Press, c.1980.
A hover test was conducted on a small-scale ro- 4. Dadone, Leo: Rotor Airfoil Optimization: An Under-
tor model for two sets of tapered rotor blades. The standing of the Physical Limits. Preprint No. 78-4, 34th
Annual National Forum, American Helicopter SOC.,I nc.,
baseline rotor-blade set used an NACA 0012 air-
May 1978.
foil section, whereas the second rotor-blade set used
5. Abbott, Ira H.; and Von Doenhoff, Albert E.: Theory of
three advanced rotorcraft airfoil sections distributed
Wing Sections. Dover Publ., Inc., c.1959.
, along the radius. Both blade sets had identical plan-
6. Thibert, Jean-Jacques; and Gallot, Jacques: Advanced
form and twist distributions. The experiment was
Research on Helicopter Blade Airfoils. Vertica, vol. 5,
conducted for a range of thrust coefficients and tip no. 3, 1981, pp. 279-300.
speeds, and the data were compared with the predic- 7. McVeigh, Michael A.; and McHugh, Francis J.: Influence
tions of three analytical methods. The methods used of Tip Shape, Chord, Blade Number, and Airfoil on
,
in this investigation were a simple momentum blade- Advanced Rotor Performance. J. American Helicopter
element analysis and two free-wake lifting-surface Soc., vol. 29, no. 4, Oct. 1984, pp. 55-62.
hover-performance analyses (HOVER and LSAF). 8. Bingham, Gene J.: The Aerodynamic Influences of
Rotor Blade Airfoils, Twist, Taper and Solidity on Hover
The results of the investigation and the comparison
and Forward Flight Performance. Proceedings of the 97th
are summarized as follows:
Annual Forum, American Helicopter SOC., May 1981,
1. The experimental data show that for the three pp. 37-50.
speeds investigated at the higher thrust levels (above 9. Gessow, Alfred; and Myers, Garry C., Jr.: Aerodynamics
a thrust coefficient C, of 0.007), the advanced airfoil of the Helicopter. Macmillan Co., c.1952. (Republished
1985 by College Park Press.)
rotor had better hover performance than that of the
10. Gessow, Alfred: Equations and Procedures for Numeri-
baseline NACA 0012 rotor.
cally Calculating the Aerodynamic Characteristics of Lift-
2. The baseline NACA 0012 rotor experienced ing Rotors. NACA TN 3747, 1956.
some detrimental Mach number effects for CT above 11. Gessow, Alfred; and Crim, Almer D.: A Method for
0.007 at the higher tip Mach numbers. The advanced Studying the Dansient Blade-Flapping Behavior of Lift-
airfoil rotor data did not show any effect of the tip ing Rotors at Eztreme Operating Conditions. NACA
Mach number variation. TN 3366, 1955.
12. Gessow, Alfred; and Tapscott, Robert J.: Tables and
3. Of the three analytical methods used for this
Charts for Estimating Stall Effects on Lifting-Rotor Char-
study, the lifting-surface analysis HOVER and the
acteristics. NASA TN D-243, 1960.
momentum analysis predicted the greater efficiency 13. Van Gaaabeek, J. R.: Rotorcraft Flight Simulation,
of the advanced airfoil rotor at the higher thrust Computer Program C81. Volume ZZ-User 's Manual.
coefficients. USARTL-TR-77-54B, U.S. Army, Oct. 1979. (Available
4. The predictions of both lifting-surface analyses from DTIC as AD A079 632.)
14. Bingham, Gene J.; and Noonan, Kevin W.: Two-
used in this study were very sensitive to the user
inputs. It was found that using the output of the Dimensional Aerodynamic Characteristics of Three
Rotorcraft Airfoils at Mach Numbers From 0.95 to 0.90.
momentum analysis as a guide for the input into
NASA TP-2000, AVRADCOM TR 82-B-2, 1982.
HOVER resulted in a prediction that was in general
15. Berry, John D.: Performance Testing of a Main Rotor
agreement with the experimental data. System for a Utility Helicopter at l/4 Scale. NASA
TM-83274, AVRADCOM TR 82-B-3, 1982.
NASA Langley Research Center 16. Kelley, H. L.; and Wilson, J. C.: Aerodynamic Perfor-
Hampton, VA 23665-5225
mance of a 27-Percent-Scale AH-64 Wind-Tunnel Model
August 5, 1988
With Baseline/Advanced Rotor Blades. Proceedings
I
of the /lst Annual Forum, American Helicopter SOC.,
References May 1985, pp. 491-499.
17. Yeager, William T., Jr.; Mantay, Wayne R.; Wilbur,
1. Stepniewski, W. Z.; and Keys, C. N.: Rotary-wing Matthew L.; Cramer, Robert G., Jr.; and Singleton, Jef-
Aerodynamics. Dover Publ., Inc., 1984. frey D.: Wind-nLnnel Evaluation of an Advanced Main-
Stepniewski, W. Z.: Volume Z: Basic Theories of Rotor Rotor Blade Design for a Utility-Class Helicopter. NASA
Aerodynamics (With Application to Helicopters). TM-89129, AVSCOM TM 87-B-8, 1987.
18. Freeman, Carl E.; and Mineck, Raymond E.: Fuselage Lift Aircraft Forum (Aix-en-Province, France), Aug.-
Surface Pressure Measurements of a Helicopter Wind- Sept. 1982, Paper No. 2.8.
nnnel Model With a 3.15-Meter Diameter Single Rotor. 24. Kocurek, J. David; and Berkowitz, Lenard F.:
NASA TM-80051, 1979.
Velocity Coupling-A New Concept for Hover and Ax-
19. Phelps, Arthur E., 111; and Berry, John D.: Description
ial Flow Wake Analysis and Design. Prediction of
of the U.S. Army Small-scale 2-Meter Rotor Test System. Aerodynamic Loads on Rotorcraft, AGARD-CPP-334,
NASA TM-87762, AVSCOM TM 86-B-4, 1987.
Apr. 1982, pp. 8-1-8-13.
20. Phelps, Arthur E., 111; and Althoff, Susan L.: Efects
25. Kocurek, J. David; and Tangler, James L.: A Prescribed
of Planform Geometry on Hover Performance of a 2-
I Wake Lifting Surface Hover Performance Analysis. J.
Meter-Diameter Model of a Four-Bladed Rotor. NASA
American Helicopter SOC., vol. 22, no. 1, Jan. 1977,
TM-87607, AVSCOM TR 85-B-6, 1986.
pp. 24-35.
21. Miller, R. H.: Simplified Free Wake Analyses for
Rotors. ASRL TR 194-3, Dep. of Aeronautics & Astro- 26. Noonan, Kevin W.; and Bingham, Gene J.: Aerodynamic
nautics, Massachusetts Inst. of Technology, Aug. 1981. Characteristics of Three Helicopter Rotor Airfoil Sections
22. Landgrebe, Anton J.: An Analytical and Experimental at Reynolds Numbers From Model Scale to Full Scale
Investigation of Helicopter Rotor Hover Performance and at Mach Numbers From 0.35 to 0.90. NASA TP-1701,
Wake Geometry Characteristics. USAAMRDL Tech. AVRADCOM TR 80-B-5, 1980.
I Rep. 71-24, U.S. Army, June 1971. (Available from 27. Keys, Charles N.; McVeigh, Michael A.; Dadone, Leo;
DTIC as AD 728 835.) and McHugh, Francis J.: Considerations in the Estima-
23. Summa, J. Michael: Advanced Rotor Analysis Meth- tion of Full-scale Rotor Performance From Model Rotor
I
ods for the Aerodynamics of Vortex/Blade Interactions Test Data. Proceedings of the 39th Annual Forum of the
in Hover. Eighth European Rotorcraft and Powered American Helicopter Society, May 1983, pp. 34-43.
8
Table I. Airfoil Coordinates for RC(4)-10
[Stations and ordinates given in fraction of airfoil chord]
Station Lower surface Station Upper surface
.oooooo
1 0.000203 0.000000 -0 .0057 26
.975148 -.003160 .002864 .004313
.950227 -.005728 .009072 .013175
.925298 -.008079 .023543 .025980
.go0374 -.010312 .047036 .038875
.875421 -.012443 .073686 .047953
.850386 -.014420 .lo0188 .053673
.825282 -.016215 .126143 .057324
.800114 - .01781 8 .151842 .059790
.774866 -.019273 .177227 .061579
.749797 - .020610 .202556 .062995
.724840 -.021873 .227760 .064163
.699792 - .023108 .252956 .065143
.674849 -.024327 .303145 .066614
.649997 - .025542 .353142 .067381
.625136 - .026750 .3781 40 .067422
.600297 - .027933 .403297 .067163
.575437 - .029062 .428390 .066543
.550567 - .030090 .453678 .065499
.525714 - .030972 .478891 .064013
.500894 -.031659 .503763 .062129
.476002 - .032163 .528707 .059876
.451451 -.032474 .553618 .057324
.426974 -.032600 .578512 .054510
.402303 -.032553 .60341 7 .051447
.377696 -.032369 .628341 .048144
.352929 -.032090 .653244 .044621
.303398 - .031393 .678157 .040912
.254059 - .030703 .702978 .037093
.229490 - .030430 .727694 .033251
.204930 -.030257 .752502 .029451
.180495 -.030276 .777197 .025808
.156116 -.030611 .801713 .022378
.132051 - .031269 .826309 .019139
.lo8242 -.032011 .850970 .O 16086
.084979 -.032337 .875699 .013211
.061 865 - .031 57 6 .go0509 .010514
.035595 -.028199 .925350 .0080 12
.025184 -.025664 .950185 .005722
.016462 -.022703 .975028 .003652
.oooooo
.014350 - .021823 1 .OO 1785
.004687 -.015907
9