NASA/TM-2000-208785

AFDD/TR-99-A-001

Preliminary Study of a Model Rotor in Descent

K. W. McAlister, C. Tung, D. L. Sharpe, S. Huang, and E. M. Hendley

November 2000

https://ntrs.nasa.gov/search.jsp?R=20010038031 2018-05-15T06:54:20+00:00Z
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NASA/TM-2000-208785

AFDD/TR-99-A-001

Preliminary Study of a Model Rotor in Descent

K. W. McAlister, C. Tung, D. L. Sharpe, S. Huang, and E. M. Hendley

aeroflightdynamics Directorate

U.S. Army Aviation and Missile Command

Ames Research Center, Moffett Field, California

National Aeronautics and

Space Administration

Ames Research Center

Moffett Field, California 94035-1000

November 2000
Available from:

NASA Center for AeroSpace Information

7121 Standard Drive

Hanover, MD 21076-1320

(301) 621-0390

National Technical Information Service

5285 Port Royal Road

Springfield, VA 22161

(703) 487-4650 !
Preliminary Study of a Model Rotor in Descent

K. W. MCALISTER, C. TUNG, D. L. SHARPE, S. HUANG, AND E. M. HENDLEY

Aeroflightdynamics Directorate, U.S. Army Aviation and Missile Command

Ames Research Center

Summary

Within a program designed to develo p experimental

techniques for measuring the trajectory and structure of

vortices trailing from the tips of rotor blades, the present

preliminary study focuses on a method for quantifying the

trajectory of the trailing vortex during descent flight

conditions. This study also presents rotor loads and blade

surface pressures for a range of tip-path plane angles and

Mach numbers. Blade pressures near the leading edge and

along the outer radius are compared with data obtained on

the same model rotor, but in open jet facilities. A triangulation procedure based on two directable laser-light

sheets, each containing an embedded reference, proved

effective in defining the spatial coordinates of the trailing

vortex. When interrogating a cross section of the flow that

contains several trailing vortices, the greatest clarity was

found to result when the flow is uniformly seeded.

Surface pressure responses during blade-vortex interactions appeared equally sensitive near the leading edge

and along the outer portion of the blade, but diminished

rapidly as the distance along the blade chord increased.

The pressure response was virtually independent of

whether the tip-path plane angle was obtained through

shaft tilt or cyclic pitch. Although the shape and frequency of the pressure perturbations on the advancing

blade during blade-vortex interaction are similar to those

obtained in open-jet facilities, the angle of the tip-path

plane may need to be lower than the range covered in

this study.

Nomenclature

a speed of sound

A1 longitudinal cyclic pitch angle

B1 lateral cyclic pitch angle

c blade chord

Cp pressure coefficient, (p - ps)/ql

ACp differential pressure coefficient, (Pu - P_)/q_

CT thrust coefficient, thrust/pn_2R 4

M Mach number, (rf_ + V. sinx_)/a

M,t advancing tip Mach number, (Rg2 + V.)/a

Mh hover tip Mach number, RD./a

p blade pressure (either p_ or Pl)

pl blade pressure on lower surface

ps freestream static pressure

pu blade pressure on upper surface

q freestream dynamic pressure, pV.2/2

ql local blade dynamic pressure,

1/2p(r_ + V_sin_) z

r local radius along blade

R blade radius at tip

Re Reynolds number, eVwv

Vat advancing tip velocity, Rf_ + V_

V_ freestream velocity

x chordwise distance from leading edge of blade

t_tpp angle between freestream and tip-path plane

0 blade pitch angle, 0 = 0o + Alsinv + Blcos_

0o collective pitch angle

_t advance ratio, V../Rf_

v fluid kinematic viscosity

p fluid density

rotor blade azimuth

rotational speed of rotor

Introduction

Although the current civil helicopter has been used

extensively because of its excellent hover and low-speed

forward flight capabilityl the fleet continues to exhibit

noise levels that are too high for widespread community

acceptance. A major source of helicopter noise originates

from the rotor blade cutting through its own wake. This

phenomenon is known as blade-vortex interaction (BVI)

noise and is most prominent when descending for a

landing.
Manyexperimental studieshavebeenperformedover the

last decade to investigate the BVI phenomenon. The

experiments have ranged from full-scale flight tests

(refs. 1 and 2) to model-scale tests (refs. 3-5). At lower

advance ratios (_t < 0.2), it was found that BVI noise can

be scaled (refs. 6 and 7), suggesting that expensive flight

tests can often be simulated by model rotor tests.

BVI events are highly impulsive. Specifically, rotor-blade

acoustics are a function of surface-pressure fluctuations,

which in turn are influenced by the geometry and composition of the trailing vortices within the rotor wake and

with which the blades interact. Analytical methods for

calculating rotor acoustics are dependent on an accurate

description of the vortex wake (refs. 8 and 9). Reliable

and detailed information of this sort is normally derived

from experiments; however, reports of these measurements are rather sparse.

The most dominant feature characterizing the rotor wake

is the tangle of vortices trailing from the tips of the rotor

blades. Attempts to visualize the rotor wake in wind

tunnels typically fall into two groups. One group depends

on density variations in the flow that result from the low

pressure near the center of the vortex. Schlieren and

shadowgraph techniques (refs. 10-15) can image the

res-tiliing variation in air density, and therefore the vortex

geometry. The other grouP relies on the induced flow

around the vortex to entrain the particles that are

deliberately injected into the stream. The particles are

redistributed in a characteristic pattern that reveals the

location of the vortex when illuminated with a laser light

sheet (refs. 16-19). The wake geometry presented in

reference 20, which features the locations of vortex

segments at discrete azimuths, is a good example of wake

data obtained in this fashion.

This present experimental study was undertaken to

establish a technique for defining the wake geometry and

structure of the trailing vortex, and to explore the possibility of complementing the acoustic measurements on

model rotors performed in other facilities. The bridge

between a wake study in one facility and an acoustic

study in another facility would be the correlation of blade

pressures in the two experiments. A model rotor experiment performed in another facility, which focused on

acoustics and included blade pressure but lacked

information about the wake geometry, was reported in

reference 21. That test was conducted at high forward

speed and for several BVI conditions on a model rotor in

an open jet, anechoic test facility. In the present study a

test is performed on the same model rotor, but in a closed

test section and with the angle of the tip-path plane

adjusted to replicate one of the BVI conditions in the

open jet test. Aside from exploring the efficacy of

matching open and closed jet results, the long-range

benefit from this effort is the development of techniques

for measuring model rotor loads, wake geometry, blade

deformation, and trailing vortex structure (core size and

strength) during a BVI (descent) condition. This report

contains the results from the first phase of this study

(identified as LoTuS) and focuses on the rotor loads and

the technique for measuring wake geometry.

Description of Experiment

Facility and Model

The test was conducted in the 7- by 10-Foot Wind Tunnel

at Ames Research Center, using a 1/7-scale model of the

AH-I/OLS rotor. The rotor consists of two blades, with a

chord dimension of 4.09 in. and a span of 37.7 in. measured from the axis of rotation to the blade tip (fig. 1). The

rotor disk was located approximately in the center of the

test section, with the rotor diameter spanning about 63%

of the width of the test section (fig. 2). Beginning at 25%

radius, the twist of each blade was -0.27°/in., so that the

total variation in geometric twist along the span was

7.55 °. Collective pitch settings were referenced to the

geometric pitch angle at 75% radius. Therefore, a

collective pitch setting of zero means that the blade pitch

was about 5.0 ° at the root (held fixed for locations < R/4)

and -2.5 ° at the tip. The rotor hub has a built-in blade

precone of 1.3°.

The blades were instrumented with surface pressure

transducers and safety-of-flight gages. One blade

contained 28 absolute pressure transducers (14 were

located on the upper surface and 14 on the lower surface)

and the other contained 18 differential pressure transducers (fig. 3). Calibrations were performed by applying a

vacuum to each transducer, either individually for the

differential transducers or collectively (using a cylindrical

cuff) for the absolute transducers. Information related to

flight safety was obtained by flap- and chord-bending

gages mounted on the blade (fig. 4). The dragl_n_ used to

adjust the blade sweep was instrumented with a strain

gage that responded to the aerodynamic drag on the rotor,

and as such, offered a redundant measurement of chordwise bending. Depending on the sweep of the blade and

the location of the center of mass of the blade, this gage

could also be loaded by a centrifugal force induced

moment. Strain gages were placed on the pitch horns that

transmitted commands to the rotor blades. Reactions to

these commands resulted in bending loads on these

elements that were proportional to the pitching moments -

on the blades. A transducer (called the teeter pot) for

measuring the angular tilt of the rotor was mounted on top

2
ofthehub.Therotoroperator monitoredthesignalsfrom

eachofthesegages.

Theexcitationandsignalwiresfromthesafetygagesand

pressuretransducers werepasseddownthroughtheshaft

(fig.5)andthejunctionsweresecuredwithtapeand

string(fig.6).Theswashplateandrelatedpitchcontrols

areshowninfigure7.Therotorhubwasmountedonthe

VSB-54forceandmoment balance, andtherotorstand

wasattachedtoaplatetoachievearangeoftip-path

planesbasedonshaftilt(fig.8).Theentirerotorassemblycouldbepositionedinthestreamwisedirection

(fig.9)sothattherotorwakewouldbeinthemost

optimumpositionforviewingthroughthetestsection

windows(locationestablishedduringmodelinstallation).

Basedonmicrophoneandbladepressuremeasurements

collectedinearliertests(refs.3and7),appreciableBVI

noisewasfoundtooccuratseveral descent conditions

(ref.8).Consideringoneofthesetestconditions,

Mh=0.664andla=0.164(Deutch-Niederlandischer

Windkanal (DNW)casenumber10014), andgivena

freestreamtemperature of58°Finthepresent test,the

requiredtunnelandrotorspeedsforthepresent test

werecalculatedtobeV.=122fps(q=0.122psi)and

=2250rpm.Thecorresponding advancingtipMach

andReynoldsnumbersare M,t= 0.77 and Re = 1.88 × 10 6

(advancing tip speed V,t = 862 fps).

Since the center of mass of each rotor blade was located

aft of the feathering axis (aligned with the quarter chord),

it was necessary to sweep the blades forward 1.2 ° to

minimize steady chordwise bending due to centripetal

force. The blades were strobed at 2/rev to verify that the

blades were tracking. A blade sweep of 1.2 ° corresponds

to a 0.19c movement of the tip. Stand accelerometers

indicated any imbalance in the rotor system and lead

weights were added as needed to the blade retention bolt.

Data Acquisition

Blade related instrumentation consisted of gages

measuring load reactions (flap and chord bending) and

gages measuring local surface pressure. These dynamic

signals were conditionally sampled based on an encoder

attached to the transmission and having a resolution of

1024 per revolution. A limited number of pressure

transducers were selected based on their present working

condition, the existence of archived electronic files from

the DNW test, and presentations of their response during

BVI in figures contained in references 8 and 22. These

criteria resulted in the selection of 19 transducers (see

table 1).

ID

IU

3L

4L

3U I

--------4

5U Absolute

8U Absolute

12U I Absolute

9L__ Absolute

10___ Absolute

16L ]Absolute

..........2

1D__ Differential

2D Differential

6D I Differential

15D Differential

9D Differential

10D Differential

16D Differential

17D Differential

Table 1. Blade pressure transducers

x/c

Type

Absolute

Absolute

Absolute

Absolute

Absolute

Wind tunnel related data consisting of freestream

temperature and freestream static and dynamic pressures

were recorded separately and later used in the calculation

of dimensionless quantities such as Mach number, thrust

coefficient, and pressure coefficient. Values for the

longitudinal and lateral cyclic pitch, collective pitch, and

shaft tilt were also recorded as part of the parameter set

identifying the test conditions.

A strain gage bridge mounted on the flex coupling located

between the transmission and the balance measured

torque. The remainder of the rotor loads was obtained

with the temperature compensated VSB-54 force balance.

The allowable load limits are given in table 2.
Table2.Allowableloads

Gage ID

Normal! I

Normal2 2

Side1 3

Side2 4

Axial1 5

Axial2 6

Torque

Max.allowableMax.alternating

500lb 200Ib

500Ib 200lb

500lb 200lb

500lb 200lb

600lb 240lb

1400Ib 560lb

510ft-lb

Toaccount forinteractionsbetweenthevariousbalance

gages, eachforcewascalculatedaccordingto:

F_=(Vi- Z_)CU

WhereF,V,Z,andCrepresent therestiltant gageforce,

loadvoltage,zero-loadvoltage,andcalibrationconstant,

respectively. Eachofthebalancegagesisreferencedby

thesubscript "i" andthematrixofcalibrationconstantsis

givenintable3.

Themostimportant quantitymeasuredbytherotor

balancewastheaxialforce.Withthismeasurement the

thrustcoefficient wascalculatedanddisplayedinreal

timesothatthecollectivepitchcouldbesetoyield

CT = 0.0054 (the condition matching the DNW data). The

procedure used to match the targ_NWv" condition is

outlined in figure 10. Although data were taken at lower

rotor speeds, the DNW matching quantities were fixed at

Mh = 0.664 (M,t = 0.77), la = 0.1 64, and CT = 0.0054.

Given the freestream temperature in the present test, the

necessary rotor rpm and freestream velocity were calculated. Assuming the tip-path plane to be the only variable

to be adjusted to account for the difference between an

open and closed jet test, the angle of the tip-path plane

was arbitrarily Set (but within a few degrees of 1°) and

some value for the collective pitch was selected. If the tippath plane was to be established by shaft tilt (rather than

with cyclic pitch), then the cyclic pitch was adjusted until

the flapping was nulled. The axial force was examined to

determine if the correct thrust coefficient had been

achieved. If not, a new value for the collective pitch was

set. Once the thrust coefficient had been matched, blade

pressure data were recorded and compared with the DNW

results. If the pressure did not match, a new value for the

tip-path plane was set, and the procedure was repeated.

The procedure for quantifying the wake geometry was

based on flow visualization. The technique depended on

the flow being properly seeded and illuminated, and the

image accurately controlled to coincide with particular

rotor azimuths. Liquid particles (primarily propylene

glycol) were released into the flow downstream of the test

section. This location was chosen to take advantage of the

mixing action of the tunnel fan, and the distance before

entering the test section, to establish a relatively uniform

spatial distribution of particles in the flow approaching

the rotor. Given the correct size and composition of the

particles, and the right temperature and humidity of the

air, the core of the vortex was comparatively void of

particles. Thisprovided a means of distinguishing the :

vortex from the surrounding flow in the rotor wake.

Table 3. Calibration matrix for VSB-54 balance

J=l

I = 1 0.00161

2 0.061664

3

4

-1.6E-05

1.64E-05

5 0.000769

6 0.000351

-0.01829

0.061696

3.36E-05

-1.6E-05

-1.6E-06

0.000373

-1.1E-05

-4.1E-05

2.85E-05

-5.8E-05

1.23E-05

0.000141

0.001629 -0.01849 0 -9.9E-06

0.062126 0.062422 0 0.00037

0.00071 0.001902 0 0.19434

0.00036 0.000386 0.1 2.96E-06

4
Anargonlaserbeamwasformedintotwocontinuous,

thinlightsheets. Eachlightsheetwaspartiallynulledand

reflectedoffmirrorsthatdirectedthemdownontothe

rotorwake.Themirrorswerelocated56in.apartand

wereconfiguredtorotateinthesameplane.Theplane

definedbythelightsheetscouldberotatedtoany

azimuthaboutanaxisparalleltotherotor(whenatzero

shafttilt),andcouldbetranslatedinthestreamwise

direction. Eitherlightsheetcouldbeusedtoscantheflow

forafeatureofinterest(inthiscase,thetrailingvortex).

Sincetheflowwasunsteady, theimagewas"frozen"

usinganintensified, gated(about0.0005secinthistest)

camera. Byaimingthenulledregionofeachlightsheetat

thecenterofthevortex,thecoordinates ofthevortex

segment inviewcouldthenbedeterminedby

triangulation.

Aschematicofthesetupforvisualizingtheflowisshown

infigure11.Elementsofthelight-sheet systemappearin

figure12abovethemodelrotor.Adetailedviewofthe

twodirectableightsheets, alongwithembedded

referencelinesthatpreciselyocatethetarget,areshown

infigure13.Thisuniquetargetingsystemisreferredtoas

LLS-ER(laserlight-sheet systemwithembedded

references).

Discussion of Results

Control Settings

Two methods were used for setting the angle of the rotor

tip-path plane, o_,p. The first method involved setting the

shaft tilt of the rotor stand to a given angle and then

adjusting the cyclic pitch to eliminate blade flap. In the

second method, the shaft tilt angle of the rotor stand was

fixed at 0 ° and the cyclic pitch was adjusted to give the

desired angle for the tip-path plane. Balance and blade

pressure data were obtained for a range of rotor speeds

and tip-path plane angles using both methods. The run

numbers corresponding to these different conditions are

given in tables 4 and 5.

Table 4. Run numbers for ottppbased on shaft tilt and zero blade flap

Tip-pathplane, Rotor, RPM

degrees 1200 1500 1750 2000 2250

0.5 162 164 166 168 170

1 62 72 82 92 174

1.5 148 152 156 158 160

2 68 74 84

2.5 108 172 112

3 64 76 86

70

66

78

80

88

90

94 104

114 116

96 106

98

100

Table 5. Run numbers for Otwpbased on blade flap and zero shaft tilt

Tip-path plane, Rotor, RPM

degrees 1200 1500 1750 2000 2250

0 118 130 138 176

0.5 122

1 124

1.5 126

132 140 178

134 142 184

136 144 186
Thecollectiveandcyclicpitchsettingsrequiredinthis

testareshowninfigure14.Bladebendingloadsare

presentedinfigure 15 for the case when the tip-path plane

angle is obtained by tilting the shaft while maintaining

zero blade flap. For 1o < cq,p < 5°, the steady component

of the flap-bending moment increased slightly with RPM

as well as with increases in the angle of the tip-path plane,

but was always less than 25 in-lb in magnitude. The

oscillatory component generally decreased with RPM

from around 85 in-lb to 50 in-lb.

The chordwise locations of the strain gages forming the

bridge for measuring lead-lag bending were incorrectly

selected. As a consequence, the bridge was dominated by

the centrifugal force induced moment and was therefore

disregarded. However, the chord bending moment can be

calculated from the load on the drag link gage, which

indicates a steady increase with RPM from around

100 in-lb to 310 in-lb. The oscillatory components from

both chord bending and drag link gages show a relatively

flat response for RPM _<2000, followed by a 60%

increase at the highest RPM. In general, the chord bending moment decreased with increases in the angle of the

tip-path plane.

Likewise, blade-bending loads are presented in figure 16

for the case when the tip-path plane angle was obtained

by means of blade flap while maintaining zero shaft tilt.

In this Case the steady component of the flap-bending

moment generally decreased with increases in RPM as

well as with decreases in the angle of the tip-path plane.

The oscillatory component decreased with increases in the

angle of the tip-path plane. The response to changes in

RPM was somewhat flatter than in the shaft tilt case. The

steady component of chordwise bending (calculated from

the load on the drag link) shows a steady increase with

increases in RPM. Although generally lower in magnitude

when compared with the shaft tilt case, the steady

component of chordwise bending shows little sensitivity

to the angle of the tip-path plane inboth the shaft tilt and

cyclic pitch cases. The oscillatory component generally

decreases with increases in the angle of the tip-path plane.

Although lower in magnitude than in the shaft tilt case,

the trend with changes in RPM is similar.

Balance Loads

Rotor balance loads are presented in figure 17 for tip-path

plane angles obtained by tilting the shaft while maintaining zero blade flap. In all cases the thrust coefficient

was held fixed at C.r= 0.0054 and p = 0.164. For

1o _<Ott_ < 5% the axial force correctly increases with the

square of rotor RPM. Although more erratic, the torque

exhibits a similar dependence on RPM. The torque

decreases with increases in o-tr, because of the reduction

in blade drag (induced and profile) that accompanies the

decrease in collective pitch (recall fig. 14). The side and

normal forces are due to the integrated effects of induced

and profile drag, which also contribute to the rotor pitch

and roll moments. Balance loads are given in figure 18 for

tip-path plane angles obtained by cyclic pitch with zero

shaft tilt. For the limited number of cases considered, the

trends are similar to those for shaft tilt, except that the

moments are smaller and less erratic.

Blade Pressures

Among the blade pressure transducers listed in table 1,

archived data (either an electronic file or a report figure)

from two absolute pressures (IU and 2U) and two

differential pressures (9D and 15D) were selected to

compare with the present test data. All of the pressure

data presented from the present test have been averaged

over 14 rotor revolutions.

Even when the test parameters are duplicated, blade

pressures may not match because of differences in the

relative size of the model rotor to the wind tunnel jet.

This was demonstrated with data taken from the DNW

and CEPRA 19 facilities for the same test conditions of

la = 0.16, Mat-- 0.77 (or Mh= 0.66), and Ct= 0.0054 and

Using the sarffe-mbdei rotor _-_ii-the present test. Both

facilities are open jet i but they differ in shape and size. To

match the blade pressure response that was obtained in

the DNW for oqpv= 1°, conditions in the CEPRA 19 were

found to require t_p---- 3.5 ° (ref. 22). Although some

differences are apparent (fig. 19), the pressure responses

were considered sufficiently similar with regard to the

resulting acoustic radiation.

Since earlier data sets for a given tip-path plane angle

could have been obtained using either shaft tilt or cyclic

pitch, both approaches were investigated and the signal

from a differential pressure near the leading edge (15D)

was used for comparison. The results for three tip-path

plane angles (0q_ = 0.5 °, 1.0°, and 1.5 °) are presented in

figure 20. Based on the similarity between the pressure

responses, the two approaches were considered

equivalent:

While holding Cr--0.0054 and/a = 0.16 constant, the

pressure response at r/R = 0.80, x/c = 0.03 provides an

indication o_thesensitivity of blade pressure to variations

in Mach number (fig. 21) and tip-path plane angle

(fig. 22). With eqvo fixed at 1° (fig. 21), the primary

consequence of reducing the Mach number (constant

means that f_ and V_ are simultaneously reduced) is a

slight decrease in the magnitude of the pressure peaks

and an advance in the time of occurrence of these peaks

(to lower values of _). With Mh fixed at 0.66 (fig. 22),
adecreaseinthetip-pathplaneangleinducesasimilar

reductionithemagnitudeofthepressurepeaks;

however, thereappearstobenoshiftinthetimeof

occurrenceofthesepeaks.

Pressureresponses ateveral radial(fig.23)andchord

(fig.24)locations, allofwhichareforCx=0.0054,

la=0.16,andMb=0.66,indicatewhatpositionsonthe

surfaceofthebladearemostsensitivetochangesinthe

aerodynamic environment ofherotor.Theresultsin

figure23showthatallofthepressures neartheleading

edgehaveequallylargeresponses andcontainnumerous

distinctivefaturesthatreflectthedifferentinteractions

occurringwiththetrailingvorticesthatarepresent inthe

rotorwake.Incontrast totheseleading-edge responses,

figure24showsadecreaseinthepressurevariation

towardtheaftportionofthebladechord.Althoughthe

datainthesetwofiguresarefrombothdifferentialnd

absolutepressuretransducers, thedistinguishing

differenceintheirresponsetoBVIisdueprimarilyto

theirlocationontheblade.

Thetip-pathplanewasassumedtobetheonlysignificant

parameter requiringadjustment toaccount fordifferences

betweenthepresent closedtestsectionconditionsand

thosefortheDNWtest.Thepressureresponseforthree

valuesoft_p are compared with DNW data obtained for

the same values of CT, P., and Mh. The differential

pressure (15D) results shown in figure 25 indicate that

numerous features are exhibited by each of the _p curves

that are similar to those found in the DNW data for

_.t_, = 1°. The shape, frequency, and amplitude of the

perturbations in the pressure response on the advancing

blade during 20 ° < _ < 130 ° bear the greatest similarity to

the DNW data; however, there are differences in both the

absolute value and phase. Those differences appear to be

insensitive to how the nondimensional groups are formed

(that is, using a instead of V. serves only to emphasize

the advancing side response relative to the retreating

side). The CEPRA data also differed in absolute value on

the advancing blade from the DNW data (recall fig. 19),

but appear to show good agreement in the phase angles of

these events.

Dynamic responses in terms of gage pressure, measured

by an absolute transducer (2U) at x/e = 0.03 and

r/R = 0.955, are compared with DNW gage pressure

data in figure 26 to illustrate the effects of modest

changes in M,_t and oq_. In all cases, the pressure perturbations on the advancing blade are similar and indicate

that small changes in these two parameters have little

effect on the advancing blade pressure. Although the

number of perturbations during this interval is the same as

in the DNW data, the amplitude of the perturbations is

less. Another difference is the higher magnitude of the

pressure in the LoTuS test; however, this is possibly due

to a difference in air density between the two tests.

Rotor Wake

The fact that the rotor blade is always operating in its own

wake is well illustrated in figure 27 by CAMRAD/JA

(Comprehensive Analytical Model of Rotorcraft Aerodynamics and Dynamics/Johnson Associates) calculations

of the tip-vortex trajectories. This particular wake

geometry was calculated for Cr = 0.0054, M.t = 0.77,

ILl= 0.16, and oqpo= 1° at intervals of A_= 15°. Focusing

on the principal source of high-speed BVI noise, the

segment of vortex trajectory that is shed when the blade

moves past an azimuth of_ = 135 ° bears the greatest

interest. Following this segment, marked with a square

symbol in figure 28, as it moves downstream (primarily

by freestream convection) shows that several classical

types of interaction occur. Designating blade #1 as the

source of the shed vortex to be tracked, the first classical

interaction occurs at _ = 285 ° as blade #2 moves past this

vortex segment with a nearly perpendicular intersection.

The second classical interaction occurs at _ = 435 ° as

blade #1 moves past its own vortex segment with an

oblique intersection. The final classical interaction occurs

at _ = 585 ° as blade #2 moves past this vortex segment

with a parallel intersection. This particular interaction is

the most critical in terms of BVI noise.

Tracking of the vortex segment depends on proper

seeding of the flow as well as correct illumination of the

cross section of interest. Introducing the seed material

directly upstream of the rotor (in the settling chamber

about 85 fl ahead of the rotor in this test) resulted in a

dramatic display of the flow pattern induced by a particular trailing vortex (fig. 29a). Concentrated seeding has

the advantage of emphasizing the gross features of the

flow so that the location of the vortex can be readily

identified. The disadvantage of concentrated seeding is

that the particular features that are emphasized in the rotor

wake are biased by the exact cross-stream position where

the particles are introduced. A more uniform distribution

of seed particles in the test section is achieved when the

particles are released into the flow downstream of the

rotor. This location allows more time for diffusion into

the flow and also takes advantage of the mixing action of

the tunnel fan. Increasing the proportion of particles

released downstream, relative to those released directly

upstream, lowers the biasing of the flow visualization and

enables other features that might be present to be identified (fig. 29b). In this case there happens to be three

vortices that are present in the flow. Vortex #2 was the

only apparent vortex in figure 29a, whereas vortex #1

was the segment actually being tracked since its release at
= 135°.Themostuniformdistributionofseedparticles

inthetestsectionisobtainedwhentheyareinjectedinto

theflowonlyatthedownstreamlocation(fig.29c).The

advantageofferedbythisconditionisthatreal-time

opticaltrackingofthevortexcanbeemployedwhenever

meander isaproblem. Thedisadvantage isthathepopulationofseedparticlesreachingthetestsectionmayalso

beuniformlytoosmall.Inthiscasefewerparticleswill

centrifugeawayfromthevortexcoreandthevortexwill

belessdistinguishable fromtheremainder oftheflow

aroundit.Thisdistinguishable changeintheparticle

countisgenerallyinterpretedtobeareasonable estimate

forthesizeofthevortexcore.

Tocorrectlymapthespatialhistoryofagivenvortex

segment requirestaggingthesmalldomainoffluid

comprisingthatsegment andfollowingitssubsequent

excursionaspartoftherotorwake.Anequivalent

experimental procedurewouldrequirebrieflymarkingthe

fluidwithinthecoreofthevortexatthemoment it isshed

fromthe blade. Rat_er th_ marking the fluid and tracking the trajectory of the same segment Of ruled as it moves

through the wake, a hybrid approach was taken. The

calculated streamwise location of the vortex segment for a

given blade azimuth was used to determine the streamwise position of the plane of the light sheets. The reference line contained in each of the light sheets w_s then

used to locate the p0sifion of the vortex within that plane

(fig. 30). For exarnple, if the plane of the light sheets was

oriented perpendicular to the freestream direction, then

aiming the light sheets at the center of the vortex would

yield the crossfiow location. Because of the flexibility in

the present system for positioning the plane of the light

sheets relative to theflow, several paramei_rs rnust be

known in order to determine the spatial location of the

targeted region. These parameters are (1) the streamwise

location of the light-sheet pivot, (2) the angle of the lightsheet plane relative to the flow, and (3) the angles

defining the direction of each light sheet.

The measured trajectories of vortex segments shed at

V = 112 ° and V = 135 ° are shown in figure 31. Tlae _

results show both segments moving inboard and initially

rising above the rotor disk as they are convected downstream. Included in the top view are small figures indicating the azimuthal age of the vortex segment, since

being shed by the rotor blade. The segment associated

with _ = 112 ° was tracked for a greater distance and

appears to be drawn downward over the aft portion of the

rotor disk. The calculated trajectories using CAMRAD/JA

for vortex segments shed when V = 105 °, _ = 120 °, and

= 1350 show the same trends as the measured results,

although the magnitudes of the crossflow and vertical

movements are different.

Conclusions

Blade surface pressure depends on the angle of the tippath plane, but appears to be independent of whether this

angle is obtained using shaft tilt or cyclic pitch. Surface

pressures near the leading edge appear to be equally

sensitive along the entire outer half of the blade, but

exhibit a marked decrease in sensitivity as the distance

along the chord increases.

The shape and number of pressure perturbations caused

by blade-vortex interactions on the advancing blade

are similar to those obtained in an earlier DNW test.

However, the amplitude and phase of the perturbations

were slightly different for the range of tip-path plane

angles in this study.

Downstream injection of seed particles yields the most

uniform seeding distribution in the test section. Comparatively close upstream injection of seed particles reveals

selected gr0sst'eaiures of the flow and is dependent on the

lateral (or cross'Stream) location of th_]njeci_ p_in-t.

The diameter of the core region of the trailing vortex is

approximately equal t-o the thickness of the blade at the

tip. The laser light-sheet system with embedded references proved to be an accurate and simple means for

ascertaining the spatial coordinates of the trailing vortex.

The measured trajectories of vortex segments shed from

the rotor blade at particular azimuths on the advancing

side show an inboard movement toward the rotor hub and

an arched vertical movement that is positive in the second

quadrant and negative in the first quadrant. These trends

were found to be in agreement with CAMRAD/JA _

calculations, but the magnitudes were slightly different.

References

1. Sakowsky, P. C.; and Charles, B. D.: Noise

Measurement Test Results for AH-1G Operational

Loads SUrvey. Vols. I and II, Bell Helicopter Co.

Report 299-099-831, 1976.

Boxwell, D. A.; and Schmitz, F. H.: Full-Scale

Measurements of Blade-Vortex Interaction Noise.

Presented at the 36th Annual National Forum of the

American Helicopter Society, Washington, DC,

1980.

Boxwell, D. A.; Schmitz, F. H.; Splettstoesser,

W. R.; and Schultz, K. J.: Model Helicopter Rotor

High-Speed Impulsive Noise: Measured Acoustics

and Biade Pressures. Presented at the 9th European

Rotorcraft Forum, Paper No. 17, Stressa, Italy, 1983.

.

.

8
4. Dadone, L.U.;Caradonna, F. X.; Ramachandra, K.;

Silva, M. J.; and Poling, D.: The Prediction of Loads

on the Boeing Helicopters Model 360 Rotor.

Presented at the 45th Annual National Forum of

the American Helicopter Society, Boston, MA,

May 1989.

5. Lorber, P. F.: Aerodynamic Results of a PressureInstrumented Model Rotor Test at the DNW.

Presented at the 46th Annual National Forum of the

American Helicopter Society, Washington, DC,

1990.

6. Schmitz, F. H.; Boxwell, D. A.; Lewy, S.; and

Dahan, C.: A Note on the General Scaling of Helicopter Blade-Vortex Interaction Noise. Presented at

38th Annual National Forum of the American

Helicopter Society, Anaheim, CA, 1982.

7. Boxwell, D. A.; Schmitz, F. H.; Splettstoesser,

W. R.; and Schultz, K. J.: Helicopter Model Rotor

Blade Vortex Interaction Impulsive Noise:

Scalability and Parametric Variations. Journal of the

American Helicopter Society, vol. 32, no. 1, 1987.

8. Yu, Y. H.; Tung, C.; Gallman, J.; Splettstoesser,

W. R.; Schultz, K. J.; Spiegel, P.; Rahier, G.; and

Michea, B.: Aerodynamics and Acoustics of Rotor

Blade-Vortex Interactions: Analysis Capability and

its Validation. Paper AIAA 93-4332, presented at

15th AIAA Aeroacoustics Conference, Long Beach,

CA, 1993.

9. Beaumier, P.; Prieur, J.; Rahier, G.; Demargne, A.;

Tung, C.; Gallman, J. M.; Yu, Y. H.; Kube, R.;

Van der Wall, B. G., Schultz, K. J.; Splettstoesser,

W. R.; Brooks, T. F.; Burley, C. L.; and Boyd, D. D.:

Effect of Higher Harmonic Control on Helicopter

Rotor Blade-Vortex Interaction Noise and Initial

Validation. AGARD Meeting, Berlin, Germany,

1994.

10. Tangler, J. L.: Schlieren and Noise Studies of Rotors

in Forward Flight. Proceedings of 33rd Annual

Forum of the American Helicopter Society,

Washington, DC, 1977.

11. Parthasarathy, S. P.; Cheo, Y. I.; and Back, L. H.:

Wide-Field Shadowgraph Flow Visualization of Tip

Vortices Generated by a Helicopter Rotor. AIAA

18th Fluid Dynamics and Plasmadynamics and Laser

Conference, AIAA 85-1557, Cincinnati, OH, 1985.

12. Norman, T. R.; and Light, J. S.: Rotor Tip Vortex

Geometry Measurements Using the Wide-Field

Shadowgraph Technique. Journal of the American

Helicopter Society, vol. 32, no. 2, 1987, pp. 40-50.

13. Bagai, A.; Leishman, J. G.; and Samak, D. K.: A

Study of Rotor Wake Development and Wake/Body

Interaction in Hover. Proceedings of the American

Helicopter Society International Specialists' Meeting

on Rotorcraft Basic Research, Georgia Institute of

Technology, Atlanta, GA, 1991.

14. Light, J. S.; Frerking, A. A.; and Norman, T. R.:

Application of the wide-Field Shadowgraph

Technique to Helicopters in Forward Flight.

Proceedings of the 46th Annual Forum of the

American Helicopter Society, Washington, DC,

1990.

15. Leishman, J. G.; and Bagai, A.: Fundamental Studies

of Rotor Wakes in Low Speed Forward Flight Using

Wide-Field Shadowgraphy. AIAA 9th Applied

Aerodynamics Conference, AIAA 91-3232,

Baltimore, MD, 1991.

16. Brand, A. G.; Komerath, N. M.; and McMahon,

H. M.: Results from Laser Sheet Visualization of a

Periodical Rotor Wake. AIAA 26th Annual Science

Meeting, AIAA 88-0192, Reno, NV, 1988.

17. Leighty, B. D.; Rhodes, D. B.; Franke, J. M.; and

Jones, S. B.: A Synchronous Strobed laser Light

Sheet for Rotor Flow Visualization. NASA

TM-4266, 1991.

18. Elliott, J. W.; Peryea, M. A.; Brand, A. G.; and

Wood, T. L.: Induced Inflow Velocity and Blade

Surface Pressure Measurements for a Helicopter

Model in Forward Flight. NASA TM-104224, 1992.

19. Ghee, T. A.; and Elliott, J. W.: The Wake of a

Small-Scale Rotor in Forward Flight using Flow

Visualization. Journal of the American Helicopter

Society, vol. 40, no. 3, 1995, pp. 52-65.

20. Yu, Y. H.; Gmelin, B.; Heller, H.; Philippe, J. J.;

Mercker, E.; and Preisser, J. S.: HHC Aeroacoustics

Rotor Test at the DNW_The Joint German/French/

U.S. HART Project. Proceedings of the 20th

European Rotorcraft Forum, Amsterdam,

Netherlands, 1994.

21. Splettstoesser, W. R.; Schultz, K. J.; Boxwell, D. B.;

and Schmitz, F. H.: Helicopter Model Rotor-Blade

Vortex Interaction Impulsive Noise: Scalability and

Parametric Variations. NASA TM-86007, 1984.

22. Boxwell, D. A.; Schmitz, F. H.; Splettstoesser,

W. R.; Schultz, K. J.; Lewy, S.; and Caplot, M.:

A Comparison of the Acoustic and Aerodynamic

Measurements of a Model Rotor Tested in Two

Anechoic Wind Tunnels. NASA TM-88364, 1986.
R---37.7" = I

_£o;- ....I.............. _................. I c-4o_.

0.18 R R

t r .....

Max Thickness-O.097 C

Figure 1. Dimensions of model OLS rotor blade (modified BHT 540).

Figure 2. installation of model rotor in 7- by I 0-Foot Wind Tunnel.

10
Blade 2

Differential Transducers

Blade 1

Absolute Transducers

i Surface

i

I

i

l

Surface

Figure 3. Location of absolute and differential pressure transducers.

Figure 4. Rotor safety-of-flight gages.

11
Figure 5. Excitation and signal wires from safety gages and pressure transducers.

_--:--T- ..... 7--

7

Figure 6. Instrumentation wires secured with tape and string.

12
Figure 7. Blade pitch controls.

13
RotorCenterline /

Tunnel Centerline

2.5:1 Transmission

X Inside Pancake Housing

VSB 54

Balance

Housing

Enc0der

Flex Coupling

Housing "x_

Test Section

Pancake

Housing

Tilt Actuator

Tilt Plate _c::_° o o

o o Tunnel Way

Rail

Figure 8. Front view of rotor stand installation in 7- by 10-Foot Wind Tunnel.

14
Rail

Way

Downstream

Direction

©

Tilt Plate

Pancake Housing

Test Section Center

Rotor Stand Center

(! .4" offset)

Tilt Actuator

Rail Clamp

Figure 9. Top view of rotor stand installation in 7- by 10-Foot Wind Tunnel.

15
Mh=0.664

Blade Pressures

la--O.164 Ct=0.0054

Cttpp=1° open tunnel

given

adj,lst

bracket

I

refinelbracket

freestream /

, temperature

sound ....

speed

=- gwen

_ - a M h / R ................ ,,;....... rotor RPM

V._ _R _ given

= $a ...................................... freestream velocity

q -- 0.5 p V_, 2 ........ dynamic pressure

"-1 6*

collective

pitch

cyclic

pitch

_[ null flapping pot

match coefficient thrust

record data

[ compare

[ blade press

not zero

1

¢one =0

I_-- ao - altos ¥ - 1sin

.... .................... _:note: moments _ 0 also

given

] C_t=thrust/p_tf_2R 4

L balance

[................ "..i_.. note cyclic A 1 & B 1 values

0= 00 + Alsin _+ BlCOS _

collective longitudinal cyclic lateral cyclic

Figure 10. Procedure for matching DNW test conditions.

16
Argon

laser

Multi-Mode

Optical Fibers

Path of Vortex

Segment Shed

at Specific Time

Upper

Test-Section

Window

Center of

Test Section

Laser

Light-Sheet

System

Gated/

Intensified

Camera _,

IIIII!111111test section IIII11111111Overhead

Traverse

Seed

Particles

Flow

I

Figure l 1. Setup for flow visualization.

17
Figure 12. Elements of the light-sheet system.

Figure 13. Detailed view of directable light sheets with coincident reference lines.

18
"10

r-

.--

4=d

03

cL)

•t-,. 0

t_

-3

-6

Tip-Path Plane

O =1 °

0 =2 °

O =3 °

==4 °

=5 °

Shaft Tilt with Zero Blade Flap

e=eo+Alsin_+BlCOS_

Collective, 8o

Lateral Cyclic, B1

Longitudinal Cyclic, A 1

1000 1500 2000 2500

RPM

Tip-Path Plane

o =0.0 °

[] =0.5 °

=1.0 o

A ----1.5 o

Blade Flap with Zero Shaft Tilt

e=eo+Alsin_+B_cos_

Collective, 8o

Lateral Cyclic, B_

Longitudinal Cyclic, A 1

I I

000 1500 2000 2500

RPM

Figure 14. Rotor-blade pitch settings required for various tip-path plane angles

(Cr=0.0054, p.=0.164, Mh and M.t varying).

19
75

u_ 50

.O

|

•-- 25

r"

(D 0

E

o

-25

-50

0

U3

J_ -600

"T

c

_-; - 1200

c

(D

E

o . 1800

-2400

5OO

400

!

E

•- 300

E

® 200

E

o

:_ 1oo

o

Flap

-I- Flap Bending

l

, + Chord Bending I

ub

Tip-Path Plane

[] =1.0 °

0=2.0 0

o = 3.0 °

=4.0 °

v =5.0 °

120

9O

6O

3O

Steady

I 1

Lead Lag

0

Not shown - measurement

dominated by centrifugal loads

Steady

400

300

200

100

0

Steady

I I

500

400

_.. 200

100

i ......... 0

1000 1500 2000 2500

Rotor RPM

Drag Link

Chord Bending Gages

Flap Peak-Peak

Flap Bending Gage

..... 4

Lead Lag Peak-Peak

1 I

Drag Link Peak-Peak

o_

v-------v_¢--'---,

I I

1000 1500 2000

Rotor RPM

2500

Figure 15. Bending loads for various O_tppobtained by shaft tilt with zero blade flap.

I

I

2O
75

u} 5O .(3

"T

•- 25

o 0

E

o

:E -26

-50

O

t_

43 -600

.m

'_ - 1200

C

E

o_ -lsoo

2

-2400

500

m 4O0

JO

"7

•- 300

+C

200

E

o

,oo

Tip-Path Plane

= 0.O o

o = 0.5 °

o =1.0 °

Flap

+ Flap Bending

I

I

, + Chord Bending I

_ Chord Bending Gages

--- Flap Bending Gage

Lead Lag

Drag Gage__J Rotor Hub

Not shown - measurement

dominated by centrifugal loads

Drag Link

Steady

i I

[ • _

Steady

Steady

Flap Peak-Peak

120

90

60

30 1

ol

600

400

3OO

20O

100

0

500

4O0

300

200

100

0

2500

J ,, ,_r

Lead Lag Peak-Peak

f

1 I

Drag Link Peak-Peak

0 | ...... I

1000 1500 2000 2500

Rotor RPM

I i

1000 1500 2000

Rotor RPM

Figure 16. Bending loads for various a_vp obtained by cyclic pitch with zero shaft tilt.

21
__._Axial

orce

Torque

Tip-Path Plane

o ==10

6=2 °

0==3 °

_----4 o

V ==5 °

-.--- _ ............. _ FloW

-----''- - -, Normal Fo,c Rolling ........

. Pitching'_:'_ _?:'_

Moment TM _= .blae

Shaft Tilt with Zero Blade Flap Force

250

200 /_

150

1oo

50

0

3O

_ 15

O 0

u_ -15

-3O

30

-- 15

=-

o

LL 0

E

b..

o -15

Z

-30

vJ v

.

3O

e-i

-- 20

to

I I 0

60

I

40

E

E

O 20

E 0

e'-

n

-20

40

I

_- 20

o

E o

o

-20

0

n--

-40

2500

I, , , ,J,

I I i I

1000 1500 2000 1000 1500 2000

RPM RPM

2500

Figure 17. Balance loads for various t_wp obtained by shaft tilt with zero blade flap.

22
25O

,,_ 200

,.Q

15o

2

o

LL

<

100

5O

0

30

15

0

-15

-30

30

o;

(3

i..

o

LL

(D

"10

U3

Axial

orce

Torque

Tip-Path Plane

o = 0.5 °

[] =1.0 °

o = 15 °

_.I i ...............F,ow

Force Momenz ....._ Axis

• .Pitching L__ _.,

Moment ' _lae Force

J:l

-- 15

=-

o

0

LL 0

E

h.

o -15

Z

Blade Flap with Zero Shaft Tilt

3O

ea)

E

o

¢:

J:

fJ

O.

4.a

tG)

E

o

to

nT 20

CT

'- 10

0

60

4O

20

0

-20

40

20

-2O

I I

I t

-30 I I -40 i i

1000 1500 2000 2500 1000 1500 2000 2500

RPM RPM

Figure 18. Balance loads for various tXtpp obtained by cyclic pitch with zero shaft tilt.

23
Upper Surface

Pressure 2U

x/c- O. 030

dR = O. 955

Mat _tpp

DNW 0.77 1.O o

CEPRA 0.77 3.5 °

-12

O

,.A

// ;I ,"

# I _, .-Y t ,'Y_

_ SJ O'

\_., f \-., .. _,Lr nt-_ .,-/ It_

DNW I

o 9o _so =70 380

_, deg

Figure 19. Blade pressures from tests in CEPRA-19 and DNW facilities (ref. 22).

24
O4

o.

O

<1

I

o4

Q.

(J

,<1

!

o4

¢3.

O

I

Differential

Pressure 15D

x/c = O. 030

r/R= O. 800

LoTuS Data

Mat----0.69

0.9

0.6

0.3

0

O.9

0.8

0,3

0

0.9

0.6

0.3

I

Shaft Tilt Run--- 158

I

Cyclic Pitch Run= 186

..,A.,,A

..... Shaft Tilt Run=92

Cyclic Pitch Run- 184 /3

% _% I $

O_tpp__.1.0o

Shaft Tilt Run- 168

Cyclic Pitch Run=178

Ot.tpp---_O,50

I I t

O 90 180 270 360

_, deg

Figure 20. Blade pressure using two methods for obtaining the same tip-path plane.

25
I

OLtpp'--I ° Fixed

Differential

Pressure 15D

26

Used As

Reference

Below

(Dashed)

Q.

2

.el

i

0

/_== 0.16 Fixed M. 0.66 !

(1 = 2260 RPM i

Mh---- 0..60

(1 = 2000 RPM

M," 0.53

(1 = 1750 RPM

Mh= 0.45

(1 = 1500 RPM

(1 = 12OO RPM

I I i I

O 90 180 270 360

_, deg

Figure 21. Periodic pressure response for range of Mach numbers (tx_,p fixed).
O_tpp

Diffe rential _

Pressure 15D

Used As

Reference

Below

(Dashed)

Mh=O.66 Fixed

O_tpp=2.5 °

O_tpp'-2.0 °

¢N

Q.

(.3

<1

I

0.9

0.6

0.3

0

,A, ../ _ _. ,I

" ts, _

•",,;, ;, ;'

t I

0 90

I I I

180 270 360

"_, deg

Ottpp-----1.5 °

O(tpp=l.0 °

O[tppD_0.5 0

Figure 22. Periodic pressure response for range of tip-path plane angles (Mb fixed).

27
_ tp p--'_l ° Fixed

Differential

Pressures

x/c=O.03 Fixed

28

¢N

3;

Q.

!

0.9

0.6

0.3

0

I I

! I

I !

,,

I

I

I

I

I

I

I

0 90 180 270 360

_, deg

Mh----0.66

r/R- O6o

r/R=0.70

Fixed

r/R=0.80

r/R--0.955

Figure 23. Periodic pressure response for range of radial stations (atop and Mb fixed).
I

Ottpp-----1 ° Fixed

Absolute

Pressures

r/R=0.955 Fixed

Mh=0.66 Fixed

• x/c-0.50

0.9

0.6

¢N

!

0.3

I

I

I

I I

I

,,

,,

!

I

!

!

!

!

I I

0 go 180 270 360

_, deg

x/c--0.35

x/c=0.15

x/c--O.03

Figure 24. Periodic pressure response for range of chord stations (tx_vpand Mh fixed).

29
Differential

Pressure 15D

Mat o_tpp

DNW 0.77 1.O°

LoTuS 0.77 2.5°_

1.5° _> TiltShaft

0.5°_

¢N

o.

I

N

o.

_J

!

¢N

o.

!

3O

0.9

0.6

0.3

0

0.9

0.8

0.3

O

0.9

O.8

0.3

...... LoTuS

-- DNW

...... LoTuS

__ DNW

3O

r_

|

r_

!

!

Run----116 _tpp=2.5 °

I

Run=160 (Ztpp==l.5 °

,,

I ss

...... LoTuS

__ DNW

I | O i I I

60 90 120 0 90 180 270

•_, deg "_, deg

Run=170 Ottpp=0.5 °

I

t I

s S !

• t ,,+

360

Figure 25. Blade pressure for several O_wpcompared with DNW test results.

z

.=

m

=
.m

=;

z,.

u}

a)

¢.

Q_

O)

¢0

O

=;

I=.

t,o

t_

n

Q)

¢1D

a.

t_

tO

-12

-9

-6

-3

o

-12

-9

-6

-3

0

-12

-9

-6

-3

Ma t O[tpp

DNW 0.77 1.0 °

Upper Surface

Pressure 2U

i

LoTuS Run=170 Mat Ottp p

0.77 0.5 o

S

• " a/

• j% s S

", ., %, LoTuS ."

DNW

I I I

..f

#

#

#

i'

LoTuS Run=168 Mat (_tpp

0.69 0.5 o

DNW

i I ,, I

LoTuS Run=176 Mat 0.69 eCtp 0.O p°

'_'..., ,- ,, LoTuS

DNW

I I I

O 90 180 270 360

"_, deg

Figure 26. Pressure matching attempts at alternate values of c%p and M.t.

31
_=270 °

Figure 27. Calculated wake geometry for Cx 0.0054, M,t = 0.77, tx 0.16, and o%p

32
_--270 °

lb--180 °

\\ % # l/

x I/'rl • #"11

tt,--285 ° _--300 °

Perpendicular Intersection

%1

%% ii I

_-360 °

I I I !

% #1 %

% %\

_--420 ° _--436 ° li_--460 °

Oblique Intersection

/ tI

I:_:> < ,1

i

¥,-510 °

1_'670° 1_'5860 #--O 0 0 °

Parallel Intersection

Figure 28. Calculated wake geometry showing critical vortex segment interactions.

33
Figure 29. Wake visualizations with different concentrations of flow seeding.

Particles introduced (a) upstream only, (b) both upstream

and downstream, and (c) downstream only.

34
Overhead Window

of Test Section

Separated Light Sheets with

. Embedded Reference Lines

Rotor Hub

Reference

Previously Shed

Trailing Vortex

Targeted Region

of Rotor Wake

Freestream Flow

Figure 30. Laser-light sheet setup for determining coordinates of targeted region of flow.

35
36

¢1

e0

t-

,m

0

0

t-

,rE

0

r,-

U

rU

e"

U

e"

E

0

L,.

q,-

C

:6

E

¢0

i,_

tO

O

8o

-3o

-20

-IDO10-

20-

3O

-BO

Freestream

Flow Direction

Rotor

Rotation

IO

50-

40-

30"

-40

Freestream

Flow Direction

- - - *..... " ........ 7,;,- - --C'. _ ..........

Rotor tip-path plane

Shaft tilt=l ° Precone=l.3 °

w/o blade bending

t i t

-_o -zo -to 6 _b _'o _'o

Downstream distance from hub, inches

4O

Figure 3 l. Trajectories of specific vortex segments within the rotor wake.

Form Approved

REPORT DOCUMENTATION PAGE OM8no.o7o4.o188

Public reporting burden forthiscollection of informationIs estimated to average 1 hourper response,includingthe timefor reviewinginstructions, searchingexisting data sources,

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

November 2000 Technical Mem_)randum

4. TITLE AND SUBTITLE 5. FUNDING NUMBERS

Preliminary Study of a Model Rotor in Descent

6. AUTHOR(S)

K. W. McAlister, C. Tung, D. L. Sharpe, S. Huang, and E. M. Hendley

7. PERFORMINGORGANIZATIONNAME(S)ANDADDRESS(ES)

Army/NASA Rotorcraft Division, Aerot_ghtdynamics Directorate

(AMRDEC), U.S. Army Aviation and Missile Command,

Ames Research Center, Moffett Field, CA 94035

9. SPONSORiNG/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546-0001 and U.S. Army Aviation and Missile

Command, Redstone Arsenal, AL 35898-5000

581-10-12

8. PERFORMING ORGANIZATION

REPORT NUMBER

A-99V0024

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA/TM-2000-208785

AFDD/TR-99-A-001

11. SUPPLEMENTARY NOTES

Point of Contact: K.W. McAlister, Ames Research Center, MS 219-1, Moffett Field, CA 94035-1000

(650) 604-5892

12a. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified -- Unlimited

Subject Category 02

Availability: NASA CASI (301) 621-0390

Distribution: Standard

12b. DISTRIBUTION CODE

;i 3. ABSTRACT (Maximum 200 words)

Within a program designed to develop experimental techniques for measuring the trajectory and structure of

vortices trailing from the tips of rotor blades, the present preliminary study focuses on a method for quantifying the

trajectory of the trailing vortex during descent flight conditions. This study also presents rotor loads and blade surface

pressures for a range of tip-path plane angles and Mach numbers. Blade pressures near the leading edge and along the

outer radius are compared with data obtained on the same model rotor, but in open jet facilities. A triangulation procedure based on two directable laser-light sheets, each containing an embedded reference, proved effective in defining the

spatial coordinates of the trailing vortex. When interrogating a cross section of the flow that contains several trailing

vortices, the greatest clarity was found to result when the flow is uniformly seeded. Surface pressure responses during

blade-vortex interactions appeared equally sensitive near the leading edge and along the outer portion of the blade, but

diminished rapidly as the distance along the blade chord increased. The pressure response was virtually independent of

whether the tip-path plane angle was obtained through shaft tilt or cyclic pitch. Although the shape and frequency of the

pressure perturbations on the advancing blade during blade-vortex interaction are similar to those obtained in open-jet

facilities, the angle of the tip-path plane may need to be lower than the range covered in this study.

14. SUBJECT TERMS

Rotor wake, Trailing vortex, Flow visualization; Model rotor, Helicopter descent

17. SECURITY CLASSIFICATION

OF REPORT

Unclassified

NSN 7540-01-280-5500

18. SECURITY CLASSIFICATION

OF THIS PAGE

Unclassified

19. SECURITY CLASSIFICATION

OF ABSTRACT

15. NUMBER OF PAGES

41

16. PRICE CODE

A03

20. LIMITATION OF ABSTRACT

Standard Form 298 (Rev. 2-89)

Prescribed by ANSI Std. Z39-1B

298-102

=