m

NASA Technical Memorandum 4754

ARL Technical Report 1283

ATCOM Technical Report 97-A-006

Performance and Vibratory Loads Data From a

Wind-Tunnel Test of a Model Helicopter MainRotor Blade With a Paddle-Type Tip

William T. Yeager, Jr., Kevin W. Noonan, Jeffrey D. Singleton, Matthew L. Wilbur,

and Paul H. Mirick

May 1997

NASA Technical Memorandum 4754

ARL Technical Report 1283

ATCOM Technical Report 97-A-006

Performance and Vibratory Loads Data From a

Wind-Tunnel Test of a Model Helicopter MainRotor Blade With a Paddle-Type Tip

William T. Yeager, Jr.

Vehicle Technology Center

U.S. Army Research Laboratory

Langley Research Center • Hampton, Virginia

Kevin W. Noonan

Joint Research Program Office

Aeroyqightdynamics Directorate

U.S. Army Aviation and Troop Command

Langley Research Center • Hampton, Virginia

Jeffrey D. Singleton, Matthew L. Wilbur, and Paul H. Mirick

Vehicle Technology Center

U.S. Army Research Laboratory

Langley Research Center • Hampton, Virginia

National Aeronautics and Space Administration

Langley Research Center • Hampton, Virginia 23681-0001

May 1997
Available electronically at the following URL address:

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(301) 621-0390

http://techreports.larc.nasa.gov/ltrs/ltrs.html

National Technical Information Service (NTIS)

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Springfield, VA 22161-2171

(703) 487-4650
Summary

An investigation was conducted in the Langley Transonic Dynamics Tunnel to obtain data to permit evaluation of paddle-type tip technology for possible use in

future U.S. advanced rotor designs. Data were obtained

for both a baseline main-rotor blade and a main-rotor

blade with a paddle-type tip. The main-rotor blade with

the paddle-type tip has the same planform as that developed under the British Experimental Rotor Program

(BERP) but uses different airfoils and is referred to in

this report as a "BERP-type" blade. The intent of using

these two blade sets was to evaluate the effect of the

BERP planform geometry on performance and loads, not

to conduct an exhaustive study of the BERP concept. The

baseline and paddle-type tip blades were compared with

regard to rotor performance, oscillatory pitch-link loads,

and 4-per-rev vertical fixed-system loads. Data were

obtained in hover and forward flight over a nominal

range of advance ratios from 0.15 to 0.425. Results indicate that the paddle-type tip offers no performance

improvements in either hover or forward flight. Pitchlink oscillatory loads for the paddle-type tip are higher

than for the baseline blade, whereas 4-per-rev vertical

fixed-system loads are generally lower.

Introduction

The next generation of rotorcraft, particularly those

used by the U.S. Army, will be required to be faster,

more maneuverable and agile, and to carry higher payloads than current generation rotorcraft. To achieve the

goals of increased speed, higher load factor, and reduced

vibratory loads, designers have continually struggled to

obtain the proper combination of airfoils, twist, and planform geometry. Parameters such as advanced airfoils,

nonrectangular planforms, static twist, and tip sweep

have been investigated (refs. 1 through 5) to determine

their effect on rotor performance and efficiency. Conformable rotor concepts (refs. 6 through 13) have also

investigated the use of dynamic twist to improve rotor

performance and reduce vibratory blade loads.

In 1986, a modified Lynx helicopter equipped with

main-rotor blades incorporating a unique paddle-type tip

shape claimed the Class E-1 (helicopters without payload) speed record (refs. 14 through 17).The rotor blades

used on this Lynx helicopter were developed under the

British Experimental Rotor Program (BERP) and thus

became known as BERP blades, and the paddle-type tip

became known as the BERP tip. Much of the success of

the BERP blades in high-speed forward flight has been

attributed to the tip shape. On the retreating side of the

rotor disk, the BERP tip utilizes vortex flow to delay

flow separation in the tip region to attain higher angles of

attack than for a rectangular tip before the tip region

stalls. The BERP blades also incorporated three

advanced airfoil sections, which undoubtedly contributed

to the favorable performance improvement over the baseline blades used on the Lynx. The baseline Lynx blades

used a RAE 9615 airfoil section in the working region of

the blade and over most of the rest of the blade span. The

RAE 9615 is a moderately cambered, l 1-percent-thick

airfoil and its aerodynamic characteristics are a modest

improvement over those of the symmetric NACA 0012

airfoil. The BERP blade used a 9-percent-thick cambered

airfoil (RAE 9634) in the tip region, an aft-loaded

12-percent-thick airfoil (RAE 9645) in the working

region (65-85 percent radius), and a reflexed 12-percentthick airfoil (RAE 9648) in the inboard region. The

9-percent-thick airfoil provides lower drag at high Mach

numbers than an NACA 0012 airfoil and about the same

maximum lift as the NACA 0012. The aft-loaded airfoil

provides about a 35-percent improvement in maximum

lift over an NACA 0012 airfoil at the expense of high

nose-down pitching moments. The use of a reflexed airfoil with its nose-up pitching moments in the inboard

region was necessary to offset the adverse effects of the

nose-down pitching moments of the aft-loaded airfoil.

No experimental data are currently in the public domain

to quantify the contribution that the paddle-type tip made

to the performance of the BERP rotor. Because of the

success of the BERP blade, it was felt that data should be

obtained to permit an evaluation of the paddle-type tip

technology for possible use in future U.S. advanced rotor

programs.

The data were obtained in the Langley Transonic

Dynamics Tunnel in a DuPont Freon 12 atmosphere

using Mach and aeroelastically scaled model rotor

blades. A set of BERP-type blades were used as well as

rectangular planform baseline blades. The designation

"BERP-type" is used because the model paddle-type tip

blades had the same planform and twist as the full-scale

BERP blades but utilized different airfoils. The intent of

using these two blade sets was to determine the effect on

performance of the BERP planform geometry. Testing

was conducted in hover and forward flight at nominal

advance ratios from 0.15 to 0.425. Model lift and drag

were varied to simulate changes in vehicle gross weight

and propulsive force. The model hover tip Mach number

was maintained at 0.628. The discussion in this report is

intended to provide an overview of the data obtained.

Symbols

The positive directions of forces, angles, and velocities are shown in figure 1.

A balance axial force, lb

a speed of sound, ft/sec
CD Apparatus and Procedures

CL

CQ

D

d

E1

FM

FSO4p

fo

GJ

L

MT

m

N

PL HPP

Q

R

r

V

z

O_ S

0

01

P

ql

D

rotor drag coefficient,

p/_R2(_R) 2

L

rotor lift coefficient, p_R2rf2R_2_J

rotor torque coefficient, Q

p_R3(_-_R) 2

rotor drag, N sin ¢xs + A cos ms, lb

rotor diameter, ft

rotor blade bending stiffness, lb-ft 2

C_/2

rotor figure of merit, 0.707 --

C a

normalized 4-per-rev vertical fixed-system

(balance normal force) load

vehicle equivalent parasite area, ft 2

rotor blade torsional stiffness, lb-ft 2

rotor lift, N cos (xs - A sin ms, lb

_R

rotor hover tip Mach number, -- = 0.628

a

rotor blade section mass per unit length,

slugs/ft

balance normal force, lb

normalized pitch-link half-peak-to-peak load

rotor-shaft torque, measured from balance

yaw moment channel, ft-lb

rotor radius, ft

spanwise distance along blade radius measured from center of rotation, ft

free-stream velocity, ft/sec

distance from wind-tunnel floor to rotor hub,

ft

rotor-shaft angle of attack, deg

rotor blade collective pitch angle at

r

= 0.75, positive nose up, deg

twist angle built into rotor blade, positive nose

up relative to blade root end (station 6.87 in.),

deg (see fig. 4)

V

rotor advance ratio, _R

mass density of test medium, slugs/ft 3

rotor blade azimuth angle, deg

rotor rotational velocity, rad/sec

Abbreviation:

RAE Royal Aircraft Establishment

2

Wind Tunnel

The testing was conducted in the Langley Transonic

Dynamics Tunnel (TI)T). A schematic of the tunnel is

shown in figure 2. The TDT is a continuous-flow tunnel

with a slotted test section and is capable of operation up

to Mach 1.2 at stagnation pressures from near vacuum to

atmospheric. The tunnel test section is 16 fi square with

cropped corners and has a cross-sectional area of 248 ft 2.

Either air or DuPont Freon 12 may be used as a test

medium. For this investigation, Freon 12 at a nominal

density of 0.006 slugs/ft 3 was used as the test medium.

Because of its high molecular weight and low speed of

sound, the use of Freon 12 aids the matching of modelrotor-scale Reynolds number and Mach number to fullscale values. Since the primary purpose of these tests

involved rotor aerodynamic performance, matching fullscale Mach number at Reynolds numbers higher than

those obtainable by testing in air was of particular interest. The use of Freon 12 as a test medium also allows the

easing of some restrictions on model structural design

while still maintaining dynamic similarity. For example,

the heavier test medium permits a heavier, less efficient

structural design to obtain the required stiffness characteristics and thus eases design and fabrication requirements of the model (ref. 18).

Model Description

Two sets of model rotor blades with representative

Mach and aeroelastic scaling were used during these

tests: a set of rectangular planform baseline blades and a

set of BERP-type blades. Both sets of blades have the

same airfoil distribution, twist distribution, and thrustweighted solidity. The planform geometry and twist distribution of each blade set is shown in figures 3 and 4,

respectively. Because the coordinates of the RAE airfoils

used on the BERP rotor blades were not available, two

state-of-the-art U.S. airfoils were selected for use on

these model rotor blades. The BERP blade design philosophy of using an aft-loaded airfoil in the working region

of the blade and a reflexed airfoil from inboard of the aftloaded airfoil to the root end was not incorporated in the

BERP-type model rotor blades. Each blade used

RC(4)-10 and RC(3)-07 airfoils (fig. 3). The 10-percentthick RC(4)-10 airfoil was selected for the inboard region

of the blade (r/R < 0.84) based on its proven performance

(ref. 19) as an inboard airfoil section. The RC(4)-10

airfoil has high maximum lift coefficients at Mach numbers up to 0.5, very low pitching-moment coefficients

over a broad range of Mach numbers and low angles of

attack, and moderately high drag divergence Mach

numbers at low angles of attack. Initially, the 8-percentthick RC(3)-08 airfoil was selected for the tip region
(r/R> 0.866) because of its proven performance (ref. 20)

as a tip airfoil section. The RC(3)-08 airfoil has moderately high maximum lift coefficients at Mach numbers up

to 0.5, very low pitching-moment coefficients over a

broad range of Mach numbers and angles of attack, and

very high drag divergence Mach numbers at low to moderate angles of attack. However, it was discovered that

applying the RC(3)-08 airfoil to this region of the

BERP-type blade would result in an increase in the rotor

blade thickness with increasing r/R, that is, between

r/R = 0.84 and 0.866, rather than the desired constant

thickness or decrease in thickness. Scaling the thickness

distribution of the RC(3)-08 so that the maximum

thickness was 7 percent chord eliminated this problem.

Thus, the RC(3)-07 airfoil was used from r/R = 0.866 to

1.0. For the BERP-type blade, the RC(3)-07 airfoil was

laid out perpendicular to the swept quarter-chord

line for r/R < 0.96 (ref. 17). A smooth transition was

made between these two very different airfoil shapes.

The area solidity, thrust-weighted solidity, and torqueweighted solidity of the baseline rotor were 0.101 and the

same parameters for the BERP-type rotor were 0.096,

0.101, and 0.102, respectively. The structural properties

of each blade set are similar to those of the advanced

blade set described in reference 4. The nominal values of

mass distribution, flapwise stiffness, and torsional stiffness for each blade set are presented in figures 5 through

7. The chordwise stiffness and torsional inertia distributions were not measured for either blade set.

Each blade set was tested by using the aeroelastic

rotor experimental system (ARES) testbed shown in figures 8 and 9. The ARES testbed has a streamlined fuselage shape which encloses the rotor controls and drive

system. The ARES testbed is powered by a variablefrequency synchronous motor rated at 47-hp output at

12 000 rpm. The motor is connected to the rotor shaft

through a belt-driven two-stage speed reduction system.

The ARES testbed rotor blade pitch-control system and

rotor-shaft angle of attack are remotely controlled from

the wind-tunnel control room. The model rotor-shaft

angle of attack is varied by an electrically controlled

hydraulic actuator. Blade collective pitch and lateral and

longitudinal cyclic pitch are input to the rotor through the

swashplate. The swashplate is moved by three hydraulic

actuators. The model rotor hub used in this investigation

was a four-bladed articulated hub with coincident leadlag and flapping hinges. The hub operated with a measured pitch-flap coupling ratio of 0.5 (flap up, pitch

down).

Instrumentation mounted on the ARES model allows

continuous displays of model control settings, rotor

forces and moments, blade loads, and pitch-link loads.

For these tests, one pitch link was instrumented with a

strain gauge to measure pitch-link tension and compression loads. Rotor blade flap and lag motions are measured by rotary potentiometers mounted on the rotor hub.

Rotor-shaft speed is determined by a magnetic sensor.

The rotating blade data are transferred to the fixed system through a 30-channel slip-ring assembly. Rotor

forces and moments are measured by a six-component

strain-gauge balance mounted below the pylon and drive

system. Rotor lift and drag are determined from the measured balance normal and axial forces (fig. 1). Rotor

torque is measured by the balance yawing-moment channel. The balance is fixed with respect to the rotor shaft

and pitches with the fuselage. Fuselage forces and

moments are not transmitted to the balance by design.

Test Procedure

The purpose of this test was to obtain data to evaluate the effect of the BERP planform on rotor performance and loads. Therefore, data for both the baseline

and BERP-type rotors were obtained at the same nominal

test conditions defined by _t, MT, _s, and 0. Data were

obtained in hover and in forward flight from nominal values of IX = 0.15 to 0.425. In hover (kt = 0), data were

obtained at z/d = 0.83. At each test point in forward

flight, the rotor rotational speed and tunnel conditions

were adjusted to give the desired values ofM T and IX.The

ARES testbed was then pitched to the desired t_s. Blade

collective pitch was changed to obtain variations in rotor

lift. To facilitate data acquisition and reduce blade loads,

rotor cyclic pitch was used to remove rotor firstharmonic flapping with respect to the rotor shaft at each

test point. The maximum obtainable value of 0 at each

combination of IXand txs was determined by either blade

or pitch-link loads or ARES testbed drive-system limits.

Model deadweight tares were determined throughout

the range of t_s with each set of blades installed. Aerodynamic rotor-hub tares were determined throughout the

test ranges of t_s and Ix at 640 rpm with the blades

removed and the blade cuffs set at a pitch angle of 8°.

Both deadweight and aerodynamic hub tares have been

removed from the data presented herein. The value of the

sum of the deadweight and aerodynamic tares in coefficient form for _t = 0.40 and t_s = -6 ° are CL = 0.000179,

CD -- -0.000275, and Ca = 0.000020. These values represent the upper limit of the magnitude of the tares

removed from the data. No corrections for tunnel wall

effects have been applied to the data, but, as cited in reference 21, these effects are considered small for the

forward-flight conditions presented herein.

Measurement Uncertainty

The quality of the performance data obtained during

this investigation was addressed. Based on a static calibration of the strain-gauge balance used to measure rotor
lift, drag,andshaftorque,theaccuracyof themeasurementof thesevariablesiwithinthefollowingranges:

CL .......................... +0.000138

CD ......................... +0.000040

C a ......................... +0.000016

During the test, 717 data points were obtained. Out of

this data set, 177 points were repeated at random. The

repeated points included data within any particular run

(increasing then decreasing 0 for constant t_s and _t) as

well as run-to-run (resetting of all test parameters, i.e., 0,

t_s, and bt) repeat points. The average deviation in CL,

CD, and Ca (AC L, AC D, and ACo) was determined

between the originally obtained data points and the

repeat data points. The average deviations for constant

values of kt, t_s, 0, and zero 1-per-rev flapping with

respect to the shaft were determined to be as follows:

Within any particular run:

AC L ............. +0.000031 to-0.000028

AC D ............. +0.000009 to -0.000005

AC a ............. +0.000002 to -0.000001

Run to run:

AC L ............... +0.00013 to -0.00000

AC D ............. +0.000027 to -0.000020

ACQ ............. +0.000016 to -0.000004

Presentation of Results

The hover and forward-flight rotor performance data

obtained during this investigation are presented as combinations of CL, C D, and Ca. The values of CL, CD, and

C a were obtained from the average of 5000 data samples

taken at a rate of 1000 data samples/sec with a filter cutoff of 200 Hz. The value of C a for each forward-flight

rotor task, defined by specific values of CL and CD, was

obtained by plotting CL versus CD and CL versus C o at

each _t and interpolating (from CD versus C a plots) for

the value ofC a at the desired C L and CD. The desired

value of CD was obtained by using a full-scale value of

fD to first determine D as follows:

The value of CD was then determined from D. The

results for C o versus kt are presented for representative

values of CL and full-scalefD at one value ofM T. All the

C a values presented fell within the range of the interpolated data in the plots of CD versus Ca; that is, no extrapolations were made to obtain any of the C a values. The

value of M T is representative of a density corresponding

to a specific combination of geometric altitude and temperature, that is, 4000 ft/95°F.

Loads data presented consist of pitch-link and fixedsystem oscillatory loads. A harmonic analysis is used to

reduce ARES rotor and fixed-system loads to magnitude

and phase components for each harmonic of 8 per rev

and below. Typically, 12 rotor revolutions of data are

processed by using a fast Fourier transform (FFT) with

checks provided on the signal periodicity and rotor

speed. These checks ensure data integrity and allow processing to occur without signal windowing and the associated loss of signal power. Using more than 12

revolutions of data has been shown to produce no

increase in the accuracy of results for harmonics of 8 per

rev and below; therefore, all data presented herein have

been processed by using 12 rotor revolutions of data.

Pitch-link loads data are presented as normalized oscillatory (1/2 peak-to-peak) load versus CL. Fixed-system

loads data are presented as normalized 4-per-rev normal

force versus CL. The pitch-link and fixed-system loads

data are presented for three values of kt at values of t_s

representative of what would be required to produce the

necessary propulsive force to represent the same fullscale equivalent parasite areas as for the performance

data. Thus, the two rotors may not be at the same CD.

This method of presentation was chosen because the

loads data were believed to be too nonlinear to allow

interpolation as with the performance data. All pitch-link

oscillatory loads were normalized to the largest pitchlink oscillatory load measured on either rotor throughout

the speed range tested. All 4-per-rev fixed-system normal forces were normalized to the largest 4-per-rev

fixed-system normal force generated by either rotor

throughout the test speed range. This normalizing procedure was chosen because neither the fixed system nor the

rotating system of the ARES model is a dynamically

scaled representation of any existing helicopter. Therefore, scaling measured model data up to full-scale values

would not be meaningful. Ground vibration tests of the

ARES testhed have shown no dynamic amplification of

any of the strain-gauge balance channels at the rotor 4-

per-rev frequency.

The data are presented in the following order:

Figure

Rotor hover performance ...................... 10

Rotor forward-flight performance ......... 11 and 12

Pitch-link oscillatory loads .................... 13

Fixed-system oscillatory loads ................. 14
Discussion of Results

The primary purpose of this investigation was to

obtain data that would permit an evaluation of any benefits offered by the BERP planform for possible use in

future U.S. advanced rotor programs. The forward-flight

performance of both the baseline and BERP-type rotors

was defined by the Co required at a given rotor task

specified by the parameters CL, fD, tt, and M T. The values of fD chosen were intended to be representative of

both a utility- and scout-type aircraft.

Rotor Performance

Figure 10 presents the hover performance of the

baseline and BERP-type rotors at MT=0.628 and

z/d = 0.83. Figure 10(a) presents the performance in

terms of C L versus Ca; figure 10(b) makes a comparison

in terms of rotor figure of merit versus CL. Two values of

CL are of interest in this figure: 0.0081, which is representative of a UH-60A helicopter at a gross weight of

18500 lb at a density of 4000 ft/95°F, and 0.0086, which

is representative of a scout-type helicopter at a gross

weight of 10500 lb at the same density condition. The

data indicate that at both 0.0081 (UH-60A) and 0.0086

(scout-type helicopter) the BERP-type rotor requires a

higher CQ than the baseline rotor. This trend of a higher

required CQ for the BERP-type rotor for a given CL is

evident at CL from about 0.0054 up to the maximum

value of CL obtained. The figure of merit of the baseline

rotor is clearly higher than that for the BERP-type rotor

for CL greater than about 0.0052. The maximum figure of

merit of the baseline rotor occurs at a higher value of CL

than the corresponding maximum of the BERP-type

rotor. The relative inefficiency of the BERP-type rotor is

probably caused by an increase in both the induced and

profile torque. A higher induced torque for the BERPtype rotor may result from higher inflow angles in the tip

region and additional drag caused by vortex flow around

the swept tip. A higher profile torque for the BERP-type

rotor would be expected because of the increased chord

in the tip region.

Figures 11 and 12 show the forward-flight performance of the baseline and BERP-type rotors, in terms of

C a versus It, for a range of rotor lift coefficients and

three values offD at M T = 0.628. The value offD used in

figure 11 is representative of a UH-60A helicopter, and

the values offD used in figure 12 are representative of

both low-drag and high-drag configurations for a scouttype helicopter. The data in both figures 11 and 12 show

that, for the ranges of C L and fD presented, no improvement in forward-flight performance is offered by the

BERP-type rotor. The reduced forward-flight performance of the BERP-type rotor is probably caused by the

increased profile drag from the paddle-type tip.

Rotor and Fixed-System Loads

Figure 13 shows the effects of C L on pitch-link oscillatory loads (1/2 peak to peak) at three nominal values of

It. The range of rotor-shaft angle of attack presented for

each It brackets the value of propulsive force required to

offset the value offD used in figures 11 and 12. The data

show that the pitch-link oscillatory loads for the BERPtype rotor are generally higher than those for the baseline

rotor as CL is increased, particularly in the C L range of

interest (C L = 0.007 to 0.009). The increase in pitch-link

oscillatory loads with increasing C L is indicative of an

increase in blade torsional activity which is not unexpected because of the large tip area of the BERP planform. Reference 12 has indicated that increased blade

torsional activity can contribute to reduced rotor performance, as was noted in the previous section for the

BERP-type rotor. Changes in ¢xs, representing an

increase infD, show no significant increase in the oscillatory pitch-link loads for either rotor. This same trend was

found for the rotors discussed in reference 4.

Figure 14 shows the effect of CL on the 4-per-rev

fixed-system normal force for the same values of It and

ms as used in figure 13. The data show that generally the

BERP-type rotor produces lower 4-per-rev fixed-system

normal forces than the baseline rotor in the C L range of

interest (C L = 0.007 to 0.009). This trend is different

from that shown for the pitch-link oscillatory loads

(fig. 13) where the BERP-type rotor produced the higher

loads. These trends (figs. 13 and 14) indicate different

torsional loads characteristics and vertical hub shear

characteristics for the baseline and BERP-type rotors.

Increasing the rotor shaft tilt, by varying O_s, generally

showed a reduction in the 4-per-rev fixed-system loads

for both rotor systems at It = 0.25 and 0.35. The 4-per-rev

fixed-system pitching and rolling moments (not shown)

exhibited similar trends as a function of CL for both the

baseline and BERP-type rotors with variations in It and

ms up to It = 0.35. At _t = 0.35, the BERP-type rotor was

found to have higher values of both 4-per-rev fixedsystem pitching and rolling moments as a function of CL.

Conclusions

An investigation has been conducted in the Langley

Transonic Dynamics Tunnel to evaluate differences

between the performance and loads characteristics of a

rotor using blades with a paddle-type tip and a rectangular planform baseline rotor. The rotor blades utilizing the

paddle-type tip are referred to as the "BERP-type" blades

since they have the same planform, but different airfoils,

as the rotor blades used in the British Experimental Rotor

Program (BERP). Data on the BERP-type rotor are of

interest to permit an evaluation of the BERP blade planform for possible use in future advanced rotor programs

5
in the United States. Based on the data obtained for the

test conditions and model configurations investigated,

the following conclusions have been reached:

1. When compared with the baseline rotor, the

BERP-type rotor offers no performance improvements in

either hover or forward flight.

2. Pitch-link oscillatory loads produced by the

BERP-type rotor are higher than those produced by the

baseline rotor at lift coefficients of interest.

3. The 4-per-rev vertical fixed-system loads produced by the BERP-type rotor are generally lower than

those produced by the baseline rotor at lift coefficients of

interest.

NASA Langley Research Center

Hampton, VA 23681-0001

January 2, 1997

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Helicopter Rotor. NASA TP-1926, AVRADCOM TR 81-B-5,

1981.

13. Mantay, Wayne R.; and Yeager, William T., Jr.: Parametric

Tip Effects for Conformable Rotor Applications. NASA

TM-85682, AVRADCOM TR 83-B-4, 1983.

14. Lowson, Martin Vincent; St. David, Barton; Hawkings, David

Leonard; Byham, Geoffrey Malcolm; Perry, Frederick John;

and Denham, Corton: Helicopter Rotor Blades. U.S.

Pat. 4,077,741, Mar. 7, 1978.

15. White, R. W.: Developments in UK Rotor Blade Technology.

AIAA Aircraft Design, Systems and Technology Meeting, Oct.

1983. (Available as AIAA-83-2525.)

16. Wanstail, Brian: BERP Blades--Key to the 200 kn Helicopter.

INTERAVIA, vol. 3, 1986, pp. 322-324.

17. Perry, F. J.: Aerodynamics of the Helicopter World Speed

Record. Proceedings of the 43rd Annual Forum of the American Helicopter Society, May 1987, pp. 3-15.

18. Lee, Charles: Weight Considerations in Dynamically Similar

Model Rotor Design. SAWE Paper No. 659, May 1968.

19. Noonan, Kevin W.: Aerodynamic Characteristics of Two

Rotorcraft Airfoils Designed for Application to the Inboard

Region of a Main Rotor Blade. NASA TP-3009, AVSCOM

TR-90-B-005, 1990.

20. Bingham, Gene J.; and Noonan, Kevin W.: Two-Dimensional

Aerodynamic Characteristics of Three Rotorcraft Airfoils at

Mach Numbers From 0.35 to 0.90. NASA TP-2000, AVRADCOM TR 82-B-2, 1982.

21. Mantay, Wayne R.; Yeager, William T., Jr.; Hamouda,

M-Nabil; Cramer, Robert G., Jr.; and Langston, Chester W.:

Aeroelastic Model Helicopter Rotor Testing in the Langley

TDT. NASA TM-86440, USAAVSCOM TM 85-B-5, 1985.

6
L

V

_g = 180 ° Rotor-shaft axis

N

Figure 1. Notation showing positive directions of forces, angles, and velocities.

7
- Remotely

/ adjustable vanes

Fan blades (47) --x / Fixed r- Streamlined

\¥ / /- vatnes -7 ! stairwell ,4

,. .., .owve. c 

"_ _ _- .... _ _ tube cooler_ //

I "_"_ I 30000-hp / ]l Electrical equipment II_ _ vf/

[ _ _ I motor---/ II buildihg- _ valve_ I_ _ I I

' '/_ _ _"J "_fTest chamber_ _ "'__ _

I // Butterfly valve-.£ _ . . _ I _ _ c,._,, \\ I /f _ ,, ,_=___c _====__? II_V _ _

_ _ 69_ ',:?'.'Vanes-J \\

/I _ . / _di,am. y...E'- \

Observers d°me --/ _ I

Freon equipment building Laboratory building

o' 60

Scale, ft

(a) Top view.

Control

Air

Bypass

valve

Streamlined stairwell

Motor

nacelle

Figure 2.

/ Freon manifold

I I I

0 30

Scale, ft

(b) Cross-sectional view.

The Langley Transonic Dynamics Tunnel (TDT).

Sea level

8

C

il
1_3.0QI_3.87 I

T -I

1.40

L Pitch-link

, attachment point

ql

ql

4.086

47.230

48.690

56.224

1/4 chord

Transition

RC(4)-10 I I RC(3)-07

_84_ Taxww'ist

• It " D

BERP-type rotor blade

Iii

i 3.0q Ti_3.87 1i

__ 1.40 [zz_

L T \

Pitch-link

attachment point

4.454 1/4 chord

t

56.224

Transition

RC(4)-10 .L.I RC(3)-07 I

vr 1 _ i

TaxWWi2t

II

0.84R 0.866R

It

Baseline rotor blade

Figure 3. Rotor blade geometries. Linear dimensions are in inches.
-2

_

01, deg -4

-6

-8

-10 I I I I I

0 .2 .4 .6 .8 1.0

r/R

Figure 4. Twist distribution of baseline and BERP-type blades.

30

20

e_o

10

× 10-3

40

30

20

10

-- Baseline

BERP type

_._1 I I I I I I I I I I

10 15 20 25 30 35 40 45 50 55 60

r, in

I I I I I I I

0 1 2 3 4

r, ft

Figure 5. Rotor blade mass distributions (typical).

10
t"-I

i

_2

x 102

_

L

i

60 x 103

50-

40

t'q

,_ 30

20

10

I

0 10

I i I

-- Baseline

-- --- BERP type

I I I I I

20 30 40 50 60

r, in

i I I I I I i I

0 1 2 3 4 5

r, ft

Figure 6. Rotor blade flapwise stiffness distributions (typical).

4 x_102 60

50

3

40

t"-I

z 2- __.30

20

1

10

0

Figure 7.

x 103

-- Baseline

-- -- -- BERP type

m--- 1

I I I I I I

0 10 20 30 40 50 60

in

I t I r I i I t I t I

0 1 2 3 4 5

Rotor blade torsional stiffness distributions (typical).

11
L-92-10165

Figure8. Aeroelasticrotorexperimental system(ARES)modelinLangleyTransonicDynamicsTunnel.

1.35

I

2.03

I

I

I

7.78

8.37

4.99 -_

_ /-- Model pivot

,_ point

J

f

6.04

Balance

centroid

Figure 9. Schematic of aeroelastic rotor experimental system (ARES) model. All dimensions are in feet.

12
CL

.012

.010

.OO8

.006

.004

.002

.8

E/

/

,,,,,,,, ..........

.0002

/

.0004

......... ,,,,,,,,,

.0006 .0008

CQ

(a) CL versus CQ.

.0010 .0012

O Baseline

[] BERP type

2

/

.7

.6

.5

FM .4

.3

.2

.1

0

/

0 Baseline

[] BERP type

.002 .004 .006 .008 .010 .012

c,

(b) Figure of merit.

Figure 10. Rotor hover performance at M T = 0.628 and zld = 0.83.

13

/
.0012

.0010

.0008

CQ .0006

.0004

.0002

illllttpr tllE,ii_l

0 .1

/

lllr_lll_ _llllrlll 111111111

.2 .3 .4 .5

0 Baseline

[] BERP type

14

(a) CL = 0.007.

.0012

.0010

.0008

CQ .0006

.0004

.0002

i,l_llllr iiiirll_r

0 .1 .2

iil_lll,i

.3

llll_r411 qlllllpll

.4 .5

O Baseline

[] BERP type

(b) CL = 0.0081.

Figure 11. Variation of rotor torque coefficient with advance ratio forfD = 29.94 ft 2 and M T = 0.628.
CQ

.0012

.0010

.0008

.0006

.0004

.0002

iiRl,llll iirqqllql

.1

IIiIIl_lt

.2 .3

g

/

(c) CL = 0.009.

Figure 11. Concluded.

Iiifllr_l

.4 .5

O Baseline

[] BERP type

15
C a

C a

.0012

.0010

.0008

.0006

.0004

.0002

/

/

I I I I I I I i i I r p p p I I p I

.2 .3

g

(a) CL = 0.0086;fD = 12.0 ft 2.

]

P

q_lllllll

.4 .5

O Baseline

[] BERP type

.0012

.0010

.0008

.0006

.0004

.0002

I I I F P I I I I

0 .1

/

r i r r i i _ i i p _ i i i p i p i r i r i i I I r I i i i i q i i i i

.2 .3 .4 .5

g

(b) CL = 0.0095;fD = 12.0 ft 2.

O Baseline

[] BERP type

Figure 12. Variation of rotor torque coefficient with advance ratio for M T = 0.628.

16
C a

C a

.0012

.0010

.0008

.0006

.0004

.0002

0

.0012

.0010

.0008

.0006

.0004

.0002

/

.2 .3 .4

g

(c) CL = 0.0086;f0 = 18.5 ft 2.

ilpplllll

.5

/

iilll,lll i_rlr_llq

.2 .3

g

ii_lrpllr

.4

(d) C L = 0.0095;fD = 18.5 ft 2.

Figure 12. Concluded.

.5

O Baseline

[] BERP type

O Baseline

[] BERP type

17
PLHPP

PLHPP

1.0

.9

.8

.7

.6

.5

.4

.3

.2

.1

0

1.0

.9

.8

.7

.6

.5

.4

.3

.2

.1

OBaseline; Ix=0.16

[] BERPtype;_t=0.16

/

.002 .004 .006 .008

cL

/

iiirllll _tlrlllt

.010 .012 .014

(a) ms = 0°; _t = 0.16.

o Baseline; _t = 0.16

[] BERP type; _t = 0.16

J

.002 .004 .006

/j

J

p,ll_lll lllll,iF I_Jll_lp

.008 .010 .012 .014

cL

(b) ms = -2°; _t = 0.16.

Figure 13. Pitch-link oscillatory loads for baseline- and BERP-type rotors.

18
PLHPP

PLHPP

1.0

.9

.8

.7

.6

.5

.4

.3

.2

.1

iiiiip111

.002 .004 .006 .008 .010 .012 .014

c,

(c) ms = 0°; _t -- 0.25 and 0.26.

/

O Baseline; Ix = 0.25

[] BERP type; Ix = 0.26

O Baseline; Ix = 0.25

[] BERP type; Ix = 0.26

19
PLFIPP

PLHPP

1.0

.9

.8

.7

.6

.5

.4

.3

.2

.1

1.0

.9

.8

.7

.6

.5

.4

.3

.2

.1

/

IIIIIIII rllllrll'rllllqlll ilrrllrl

.002 .004 .006 .008 .010 .012 .014

cL

(e) ms = -2°; _t = 0.35 and 0.36.

O Baseline; _t = 0.35

[] BERP type; _ = 0.36

O Baseline; _ = 0.35

[] BERP type; 1a = 0.35

0

iiiirrll iil_llll, iiiirlll iil_qlll PlPPlIII ipll_rlrlp Ilplllll

.002 .004 .006 .008 .010 .012 .014

cL

(f) ms = -6°; _t = 0.35.

Figure 13. Concluded.

2O
FSO4p

FSO4p

1.0

.9

.8

.7

.6

.5

.4

.3

.2

.1

0

1.0

.9

.8

.7

.6

.5

.4

OBaseline; !1 = 0.16

[] BERP type; _t = 0.16

i_rllql Ivllllllllll_rlll_lltll_lllplltltrl_llll rlllqrlrr Iplllllrll

.002 .004 .006 .008 .010 .012 .014

CL

(a) cxs = 0°; _t = 0.16.

.3

.1

O Baseline; _t = 0.16

[] BERP type; _t = 0.16

iiit_ rrllelllrl_lllllllllllpllllllllrlll illl,l_l Itllll_ll I_lqll_

.002 .004 .006 .008 .010 .012 .014

CL

(b) czs = -2°; _t = 0.16.

Figure 14. The 4-per-rev fixed-system loads for baseline- and BERP-type rotors.

21
FSO4p

FSO4p

1.0

.9

.8

.7

.6

.5

.4

.3

.2

.1

1.0

.9

.8

.7

.6

.5

.4

.3

.2

.1

%

iiil_lll 11111111 Irrlllll Irlllllll, Irtlllll IIIIrlll

0 .002 .004 .006 .008 .010 .012 .014

cL

(c) ms = 0°; _t = 0.25 and 0.26.

o Baseline; _t = 0.25

[] BERP type; _t = 0.26

O Baseline; _t = 0.25

[] BERP type; _t = 0.26

i_lllpll qll_lrl_l Irtll_ll_ iil_rlrrp iiiiiprlp Ii ,_lllll_llllllll

.002 .004 .006 .008 .010 .012 .014

(d) ms = -4°; _t = 0.25 and 0.26.

Figure 14. Continued.

22
FSO4p

1.0

.9

.8

.7

.6

.5

.4

.3

.2

.1

1.0

.9

.8

©

,lllll_l tllllllll ir,lt,ii IBIIFIll Illll_lll Ill,llll

.002 .004 .006 .008 .010 .012 .014

cL

(e) c_s = -2°; g = 0.35 and 0.36.

O Baseline; g = 0.35

[] BERP type; la = 0.36

O Baseline; g = 0.35

[] BERP type; g = 0.35

FSO4p

.7

.6

.5

.4

.3

.2

.1

0

-iiiiillllllll illllrlllt,,I,tg

.002 .004

LplIIplII itllqtlrl Ipllllll

.006 .008 .010 .012

cL

(f) % = -6°; g = 0.35.

Figure 14. Concluded.

23


REPORT DOCUMENTATION PAGE Form Approved

OM B No. 0704-0188

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1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED

May 1997

TITLE AND SUBTITLE

Performance and Vibratory Loads Data From a Wind-lhnnel Test of a Model

Helicopter Main-Rotor Blade With a Paddle-Type Tip

Technical Memorandum

AUTHOR(S)

William T. Yeager, Jr., Kevin W. Noonan, Jeffrey D. Singleton, Matthew L.

Wilbur, and Paul H. Mirick

PERFORMINGORGANIZATIONNAME(S)ANDADDRESS(ES)

Vehicle Technology Center Joint Research Program Office

U.S. Army Research Laboratory Aeroflightdynamics Directorate

NASA Langley Research Center U.S. Army Aviation and Troop Command

Hampton, VA 23681-0001 NASA Langley Research Center

Hampton, VA 23681-0001

9. SPONSORING/MONITORINGAENCYNAME(S)ANDADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546-0001

U.S. Army Research Laboratory

Adelphi, MD 20783-1145

U.S. Army Aviation and Troop Command

St. Louis, MO 63120-1798

5. FUNDING NUMBERS

WU 505-63-36-02

PR A5008

8. PERFORMING ORGANIZATION

REPORT NUMBER

L-17558

!10. SPONSORING/MONITORING

AGENCY REPORTNUMBER

NASA TM-4754

ARL-TR- 1283

ATCOM TR-97-A-006

11. SUPPLEMENTARY NOTES

Yeager, Singleton, Wilbur, and Mirick: Vehicle Technology Center, ARL, Langley Research Center, Hampton, VA;

Noonan: JRPO, Aeroflightdynamics Directorate, ATCOM, Langley Research Center, Hampton, VA.

12a. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified-Unlimited

Subject Category 02

Availability: NASA CASI (301) 621-0390

12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum 200 words)

An investigation was conducted in the Langley Transonic Dynamics Tunnel to obtain data to permit evaluation of

paddle-type tip technology for possible use in future U.S. advanced rotor designs. Data were obtained for both a

baseline main-rotor blade and a main-rotor blade with a paddle-type tip. The baseline and paddle-type tip blades

were compared with regard to rotor performance, oscillatory pitch-link loads, and 4-per-rev vertical fixed-system

loads. Data were obtained in hover and forward flight over a nominal range of advance ratios from 0.15 to 0.425.

Results indicate that the paddle-type tip offers no performance improvements in either hover or forward flight.

Pitch-link oscillatory loads for the paddle-type tip are higher than for the baseline blade, whereas 4-per-rev vertical

fixed-system loads are generally lower.

14. SUBJECT TERMS

Helicopters; Model rotor testing; Helicopter rotor performance; Blade loads;

Blade tip geometry

17. SECURITY CLASSIFICATION

OF REPORT

Unclassified

NSN 7540-01-280-5500

18. SECURITY CLASSIFICATION

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Unclassified

15. NUMBER OF PAGES

24

16. PRICE CODE

A03

19. SECURITY CLASSIFICATION 20. LIMITATION

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