A/z/jA Z,5" 73

NASA Technical Memorandum 85683 NASA-TM-85683

AVRADCOMTECHNICALREPORT83-B-5 [ct _z_D_OL9P,.O-_

AEROMECHANICALSTABILITYOF A HINGELESS v.-_

ROTORIN HOVERANDFORWARDFLIGHT, '-'_"_,_"_':..._-.-_ ....... :,.._ ' ,. _ r-_..

ANALYSISANDWINDTUNNELTESTS ..... ;..',""-

"gO_...,._. ',

WILLIAMT. YEAGER,JR.,M-NABILH. HAMOUDA,

WAYNER.MANTAY

AUGUST1983 =___::.i_, .....

r r '"2_'T\" ;71 :'_-,

NationalAeronautics and

SpaceAdministration

LangloyRose.archCenter

Hampton, Virginia 23665

AEROMECHANICALSTABILITYOF A HINGELESSROTOR IN HOVER

AND FORWARD FLIGHT: ANALYSISAND WIND TUNNEL TESTS

ABSTRACT

A researcheffort of analysisand testinghas been

conductedin the NASA-LangleyTransonicDynamicsTunnel to

investigatethe ground resonancephenomenonof a soft in-plane

hingelessrotor. Experimentaldata were obtainedusing a 9

ft. (2.74 m) diametermodel rotor in hover and forward flight.

Eight model rotor configurationswere investigated. Configuration parametersincludedpitch-flapcoupling,blade sweep and

droop, and precone of the blade featheringaxis. An analysis

based on a comprehensiveanalyticalmodel of rotorcraftaerodynamicsand dynamicswas used for this study. The movingblock method was used to experimentallydeterminethe regressing lead-lagmode damping. Good agreementhas been obtained

between the analysisand test. Both analysisand experiment

indicatedground resonanceinstabilityin hover. The paper

presentsan outline of the analysis,a descriptionof the

experimentalmodel and procedures,and comparisonof the

analyticaland experimentaldata.

1.0 Introduction

Aeromechanicalstabilityproblemswhich involve

interactionof the rotor and airframeare usually divided into

the categoriesof ground and air resonance. While the

terminologymay imply totally differentphenomena,both are

self-excitedinstabilitiescaused by the coupling between blade

lagging motion and hub motion in the plane of the rotor [I-3].

Althoughaeromechanicalinstabilityis traditionallyassociated

with articulatedrotors, hingelessrotorsare also susceptible

to this problem. Hingelessrotorsare classifiedinto two

types. One is associatedwith a soft inplane system having the

blade inplane frequencyless than the rotor rotationalspeed,

and the other with a stiff inplanesystem where the blade

inplane frequencyis higher than the rotor rotationalspeed.

This paper deals specificallywith ground resonanceof soft

inplanehingelessrotors.

Ground resonancecan occur with helicopterrotor models

just as it can occur with full scale helicopters. The

• resonanceis characterizedby a coalescenceof the blade leadlag regressingfrequency(less than the rotor speed)with a

body vibrationmode. For a full scale helicopter,the body

• softnessis usually related to a landing gear associated

frequencywhen the vehicle is in contact with the ground. For

wind-tunnelhelicoptermodels, the body mode frequencyusually

resultsfrom the model being attachedto either a mounting

strut or a flexibleforce balance. Ground resonanceis now a

1
well understoodphenomenon[4], particularlyfor articulated

rotors. However, hingelessrotors provide substantialstructural and aerodynamiccouplingsthat complicatetheir

aeromechanicalstabilityproblems. A number of analytical °

models [5-7] have been developedto investigatethe aeromechanical stabilityof hingelessrotors. However,these analytical

models have a number of limitationsas indicatedin Reference "

8. Correlationsbetween theoryand experimenthave also been

made [9-11]. These comparisonshave been limited to hover and

the use of small-scalemodels in forward flight.

To investigatethe ground resonanceof soft inplane

hingelessrotors,an analyticaland experimentalstudy was conducted in the NASA Langley TransonicDynamics Tunnel. This

effort was intendedto aid in the identificationof an analysis

that can be used in both the designand testing phases of

hingelessand bearinglessrotor development. Another objective

of the researchwas to developan experimentaltechniquefor

blade excitationand damping measurementsin the rotating

system.

The paper presentsan outline of the analyticalmodel

and describesthe experimentalmodel and procedures. Correlations between the theory and experimentin both hover and

forward flight are presented. Finally, the effects of rotor

couplingparameterson the stabilitycharacteristicsare

discussed. These parametersare blade sweep and droop, precone

of the blade featheringaxis and pitch-flapcoupling.

2.0 AnalyticalModel

The ComprehensiveAnalyticalModel of Rotorcraft

AerodynamicTand Dynamics--(CAMRAD)cTmputerpTogram was used as

the theoreticalt_ol for this investigation. The code was

developedby Johnson [12, 13]. The solutionfor the system

aeroelasticstabilityproceedsas follows: 1) the data input,

2) the trim solution,and 3) the flutteranalysis.

The structuraldynamic model of the rotor input to

CAMRAD includeselastic degrees of freedom in flap bending,

lead-lagbending and torsion,plus a rigid pitch degree of

freedom. The blade is representedby a spanwisedistribution

of mass, bending and torsion stiffnessand moment of inertia.

An estimateof structuraldamping has also been includedin the

rotor data. The aircraft model consists of elasticmotion of

the fuselageand rotor support system in the wind tunnel. The

airframe data input includesgeneralizedmass, structural

damping,frequencyand mode shape of the elastic modes. These

modal characteristicsare set to the measured values. The

rotor blade aerodynamicforces are calculatedusing lifting

line theoryand steady two-dimensionalairfoil characteristics

2
with correctionsfor unsteadyand three-dimensionalflow

effects. The rotor non-uniforminflow is calculatedusing a

vortex-wakemodel. The airfoil characteristicsare constructed

using analyticalexpressions[13]. Fuselageaerodynamicloads

are neglected. The degrees of freedom used in the stability

analysisare the flap and lag motion of the blades, the body

pitch and roll motions and rotor dynamic inflow.

In the trim analysis the controlsare iterateduntil the

requiredoperatingstate is achieved. The trim analysis is

performedfirst for uniform inflow, then for nonuniforminflow

with a prescribedwake, and finally for nonuniforminflowwith

a free wake geometry.

In the flutteranalysis,the dynamicsof the system is

describedby a set of linear differentialequations. These

equationsrepresentthe perturbedmotion of the helicopterfrom

the trim condition. The stabilityof the system is determined

in terms of the eigenvaluesand eigenvectorsof the constant

coefficient(in hover) or periodiccoefficient(in forward

flight)equations.

3.0 ExperimentalModels and Procedures

3.1 Model Description

The rotor model used for this investigationis a soft

in-planehingelessrotor. The principalrotor propertiesare

listed in Table 1. The rotor system is not a dynamically

scaled representationof a specificaircraftbut rather is

representativeof a typical full-scaleaircraft design based on

Mach number, mass ratio and frequencysimulation. The model

blades were fabricatedwith fiberglassspars specificallyfor

Table 1. PrincipalRotor Properties

Parameter Value

Rotor Type ResearchHingelessHub

Number of Blades 4

Rotor Diameter 9 ft. (2.74 m)

• Blade Chord 4.24 in. (1.67 cm)

Solidity O.10

Airfoil Section NACA 0012

• Blade Twist 0 Degrees

Blade Elastic Axis 25% Chord

Blade Pitch Axis 25% Chord

Center of Gravity 25% Chord

3
testing in the Freon-121test medium of the Langley Transonic

DynamicsTunnel.

The rotor hub, shown in Figure 1, consistsof metal

flexuresto accommodateflap and lead-lagmotions and a

mechanicalfeatheringhinge to allow blade pitch motion. The

flap and lead-lagflexuresare each strain-gagedand calibrated •

to measure motion in those directions. The hingelesshub has

the capabilityto independentlyvary blade sweep, droop,and

precone of the blade featheringaxis. These changes are

accomplishedby means of angle blocks as shown in Figure 2.

Two values of blade pitch-flapcouplingare also available. A

list of rotor configurationstested is given in Table 2.

Table 2. Rotor ConfigurationParametricValues

Configuration a3(a) Sweep(b) Droop(c) Pre-cone(d)

(deg) (deg) (deg) (deg)

Baseline 0 0 0 0

1 42.5 0 0 0

2 0 +2 0 0

3 0 0 +2 0

4 42.5 0 +4 +3

5 42.5 0 +2 +3

6 42.5 0 -2 +3

7 42.5 0 +4 +6

a - pitch-flapcouplingangle

b - positiveaft

c - positivedown

d - positiveup

The test-bedused for the experimentis the Langley

AeroelasticRotor ExperimentalSystem (ARES). The ARES, shown

Tn Figure 3,--consiStsof a roto_ drive systemand rotor control

system enclosedby a streamlinedhelicopterfuselageshape.

The ARES is mounted on a six-componentstrain-gagebalance

which is fixed with respect to the rotor shaft and thus pitches

with the ARES. Fuselage forcesand moments are not sensed by

the balance. The entire ARES and balance assemblyare mounted •

on a rigid stand bolted to the floor of the wind tunnel. The

ARES rotor control systemand fuselagepitch attitudeare

1Freon-12is a registeredtrademarkof E.I. duPont de Nemours

and Co., Inc.

4
remotelycontrolledfrom within the wind-tunnelcontrol room.

The swashplateis moved by three hydraulicactuators. Instrumentationon the ARES and in the wind-tunnelcontrol room

• allows continuousdisplaysof model control settings,rotor

forcesand moments,blade loads,and pitch link loads. ARES

pitch attitudeis measuredby an accelerometer,and rotor

" control positionsare measured by linear potentiometers

connectedto the swashplate. Rotating systemdata are transferred into the fixed system througha 30-channelslipring

assembly.

3.2 Wind Tunnel

The test was conductedin the Langley TransonicDynamics

Tunnel. This tunnel is a continuousflow tunnel with a slotted

test section. The tunnel test section is 16 ft. (4.88 m)

square^withcropBed corners and has a cross-sectionalarea of 248 ftz (23.04m ). Either air or Freon-12may be used as

a test medium. For this studyaFreon-12at a nominal density

of 0.0047 slugs/ft3 (2.42 Kg/m_) was used as the test

medium. Because of its high densityand low speed of sound,

the use of Freon-12aids the matchingof model-rotor-scale

Reynolds numberand Mach number to full-scalevalues. The

heavier test medium permitsa simplifiedstructuraldesign to

obtain the requiredstiffnesscharacteristicsfor dynamic

similarity,and thus eases designand fabricationrequirements

of the model (ref. 14).

3.3 Test Procedures

Pre-wind-tunneltests were conductedto determinethe

rotor blade frequenciesas well as the ARES body modes and

damping. A shake test was conductedto determinethe nonrotatingrotor blade natural frequencies.This informationwas

desired to substantiatecalculatedvalues. Good agreementwas

obtainedbetweenmeasured frequenciesand those predictedby

the CAMRAD code. The first flap and lead-lagmode frequencies

were predictedwithin 0.3 Hz. The rotor blade rotatingnatural

frequencyin the lead-lagdirectionwas measured during the

wind-tunneltest as will be discussedlater. The calculated

rotor first lead-lagand flap frequenciesare O.55/REVand

1.14/REV respectivelyfor a nominal rotor speed of 618 rpm.

The measured frequencyand damping values of the ARES as

mounted in the tunnel are given in Table 3. These data were

input for the aeroelasticstabilityanalysis.

D

5
Table 3. Measured ARES Dynamic Properties

Mode Frequency Damping

Hz. %

Critical

Roll 5.4 7.3

Pitch 5.9 5.7

Testing in hover was conductedfor Configuration1 over

a range of rotor speeds from 400 rpm up to a value where rotor

in-planeinstabilitywas encountered. All configurationsshown

in Table 2 were tested in forward flightat a rotor speed of

618 rpm. Advance ratio was variedup to 0.3 with three values

of collectivepitch set at each advance ratio.

At each test point the tunnel speed was adjustedto give

the desired advance ratio. The model was then pitched to the

desired shaft angle of attack and collectivepitch was set. At

each advance ratio, data were taken at shaft angles of attack

of O, -5, and -10 degrees. The correspondingcollectivepitch

at each shaft angle of attack was 4, 8, and 12 degrees. Cyclic

pitch control was used to remove the rotor first harmonic

flappingwith respect to the shaft. Once the test condition

was established,a test techniquewas initiatedto determine

the inherentstabilitylevels for the model. This test

techniqueinvolvedusing the moving-blockmethod describedin

Reference15 as an interactiveprogram to determinethe rotor

in-planefrequencyand damping.

The test techniqueconsistedof two steps. First, the

model was excited in the fixed system by applyinga longitudinal cyclic oscillationto the rotor through the swashplate.

The amplitudeof the swashplateoscillationwas nominally0.75

degrees. The frequencyof the swashplateoscillationwas

initiallyset equal to the fixed system value of the rotor

in-planefrequency (lead-lagregressingmode) predictedby

CAMRAD. The swashplateoscillationwas then adjusted slightly

to obtain the maximum rotor in-planeresponse. Once the rotor

in-planeresponsewas established,the swashplateoscillation

was removedand the moving-blockprocedurewas initiated. A

typical real-timemoving-blockdisplay is shown in Figure 4.

Utilizationof this displaywas as describedin reference15.

The frequencyof interestwas selectedfrom the fast Fourier

transform(FFT) of the lead-lag signaltrace, and the damping °

ratio was computed from the log of the amplitudeof the

filteredlead-lagresponse. It is worth noting that in case of

a ground resonanceinstability,not only was the swashplate

excitationremoved but the rotor speedwas immediatelyreduced.

6
This action was sufficientto eliminatethe rotor disturbance

in the unstable region. By applyingthe moving-block

technique,the rotor in-planedampingwas measured in the

rotatingsystem. These damping valueswere then transferred

into the fixed systemas follows:

t

_F = _R(R-T_-_m)

where:

_F = fixed systemdamping ratio

_R = rotatingsystem damping ratio determinedby the

moving-blockmethod

= rotatingsystem first lead-lagmode regressing

frequencydeterminedby the moving-blockmethod

= rotor rotationalspeed

4.0 Results and Discussion

4.1 Hover

Testing in hover was conductedat 8 degrees collective

pitch over a range of rotor speed from 400 rpm to a value where

inplane instabilitywas encountered. The predictedand

measured stabilityof Configuration1 versus rotor speed is

shown in Figure 5. Figure 5(a) shows a predictedcoalescence

of the body roll mode frequencywith the regressinglag mode

frequencyabove 650 rpm. The predictedand experimentalleadlag frequencyare seen to be in good agreement. Over the range

of rpm tested,differencesbetween the predictedand experimental values of lead-lagfrequencyaverage 0.3 Hz. As can be

seen from Figure 5(b), the regressinglag mode dampingin the

fixed system is also well predicted. In the criticalregion

near the indicatedground resonanceinstability,differences

between the analyticaland experimentaldamping ratio average

0.006. Inplanedamping ratio decreasesgraduallyuntil a

sudden decreaseoccursaround 640 rpm. An unstable region is

indicatednear the regressinglag-rollresonancerotor speed.

" The predicteddamping level has recoveredto a positive value

(stable)at 775 rpm. Due to rotor stress level limitations,

the test could not be carried out for rotor speeds higher than

' 650 rpm. It is concludedthat the analyticalmodel accurately

predicts ground resonanceof hingelessrotors in hover.

7
Furthermore,the excitationtechniqueusing the swashplate

actuatorsproducesexcellentstabilitydata.

4.2 Forward Flight

Figure 6 shows the variationof the predictedand

measured lead-lagmode damping with collectivepitch for the

baselineconfigurationat advance ratio = 0.2, 0.25 and 0.3.

The rotor speed is 618 rpm. Both data and theory show a trend

of generallyincreasingdamping with collectivepitch. The

increasedblade pitch angle generatesaerodynamicand inertial

flap-lagcouplingthat increasesthe lead-lagdamping. A

comparisonof the damping levels for differentadvance ratios

at constantcollectivepitch indicatesthat both the analysis

and test show the lead-lagmode generallybecomesmore stable

as the forward speed increases. Good correlationis shown

between the analyticalpredictionof the dampingand experimental data. It should be noted that for both the baseline

configurationand all subsequentconfigurationsdiscussed,

differencesbetween the analyticaland experimentaldamping

ratio average 0.004.

Results for Configuration1 (63 = +42.5°) are shown in

Figure 7 for the regressinglead-lagdamping versus collective

pitch in forward flightat normal rotor speed (618 rpm). The

agreementbetween the predictedand measureddamping is good.

Comparedto the baseline configuration(no pitch-flapcoupling)

in Figure 6 it can be seen that there is a decreasein damping

due to negativepitch-flapcoupling (positive63). This may

be due to changes in pitch-lagstabilitycaused by 63 as

describedin Reference16.

Figure 8 shows the variationof predictedand measured

lead-lag dampingwith collectivepitch for Configuration2 (2°

aft blade sweep) in forward flight. As can be seen there is a

good agreementbetween the analysisand test. Comparedtothe

baselinezero sweep configurationin Figure6, these data show

that there is a change in dampingwith aft sweep.

The results for Configurations3 and 4 are shown in

Figures 9 and 10, respectively. These two configurations

incorporatechanges in blade droop and pre-cone.The predicted

and measured lead-lagdamping are seen to be in good agreement

for both configurations. Similar correlationswere obtained

between the analysis and test for configurations5, 6 and 7.

Presentedin Figures 11 and 12 are the analyticaland '

experimentallead-lagdamping resultswhich show the effect of

two configurationparameters,namely blade droop and pre-cone

angles on the damping levels. Figure 11 shows the variationof

dampingwith collectivepitch for three values of droop angle

8
in forward flight. For constantcollectivepitch, the lead-lag

damping increasesfor decreasingdroop angles. Figure 12 shows

damping results for two differentpre-coneangles. As can be

seen, the configurationwith a 6 degreepositivepre-conehas

higher lead-lagdamping than the 3 degree positivepre-cone

• configuration.

5.0 Conclusions

Theoreticaland experimentalresults for the aeromechanical stabilityof a soft-inplanehingelessrotor in hover and

forward flight have been presented. The effect on rotor stability of pitch-flapcoupling,blade sweep, droop, and pre-cone

of the featheringaxis has been demonstrated. Based on the

results of this researchthe followingconclusionshave been

reached:

1) Consistentand repeatablemeasurementsof the rotor lead-lag

regressingmode dampingwere made.

2) The analysisaccuratelypredictedthe ground resonance

instabilityfound by experimentin hover.

3) Good agreementwas found between the theoreticaland

experimentalfrequencyand damping values for the rotor

lead-lagregressingmode in hover and forward flight. The

theory predictedthe correct trendswith collectivepitch,

advance ratio, pre-cone,blade droop,and pitch-flap

coupling.

6.0 References

1. Hohenemser,K. H.: HingelessRotorcraftFlight Dynamics.

AGARD-AG-197,Sept. 1974.

2. Donham, R. E.; Cardinale,S. V.; and Sachs, I. B.: Ground

and Air ResonanceCharacteristicsof a Soft Inplane Rigid

Rotor System. Journal of the AmericanHelicopterSociety,

Vol. 14, No. 4, Oct. 1969, pp. 33-41.

3. Miao, Wen-Liu;and Huber, Helmut B.: Rotor Aeroelastic

StabilityCoupled with HelicopterBody Motion. NASA

SP-352, Feb. 1974, pp. 137-146.

4. Coleman, R. P.; and Feingold,A. M.: Theory of SelfExcitedMechanicalOscillationsof HelicopterRotors with

Hinged Blades. NACA Report 1351, 1958.
5. Young, Maurice I.; and Bailey, David J.: Stability and

Control of Hingeless Rotor Helicopter Ground Resonance.

Journal of Aircraft, Vol. II, No. 6, June 1974, pp.

333-339.

6. Ormiston, Robert A.: Aeromechanical Stability of Soft

Inplane Hingeless Rotor Helicopters. Presented at the o

Third European Rotorcraft and Powered Lift Aircraft Forum,

Aix-en-Provence, France, Sept. 1977.

7. Hodges, D. H.: A Theoretical Technique for Analyzing

Aeroelastic Stability of Bearingless Rotors. Proceedings

of the 19th AIAA Structures, Structural Dynamics and

Materials Conference, April 3L-5, 1978, pp. 282-294.

8. Johnson, W.: Comprehensive Helicopter Analyses - A State

of the Art Review. NASATM-78539, 1979.

9. Burkam, John E.; and Miao, Wen-Liu: Exploration of

Aeroelastic Stability Boundaries with a Soft-lnplane

Hingeless Rotor Model. Journal of the American Helicopter

Society, Vol. 17, No. 4, Oct. 1972, pp. 27-35.

I0. Bousman, William G.: An Experimental Investigation of

Hingeless Helicopter Rotor-Body Stability in Hover. NASA

TM-78489, June 1978.

11. Weller, William H.: Correlating Measured and Predicted

Inplane Stability Characteristics for an Advanced

Bearingless Rotor. NASACR-166280, Jan. 1982.

12. Johnson, W.: A Comprehensive Analytical Model of

Rotorcraft Aerodynamics and Dynamics, Part I - Analysis

Development. NASATM-81182, 1980.

13. Johnson, W.: A Comprehensive Analytical Model of

Rotorcraft Aerodynamics and Dynamics, Part II - User's

Manual. NASATM-81183, 1980.

14. Lee, Charles: Weight Considerations in Dynamically Similar

Model Rotor Design. S.A.W.E. Paper No. 659, May 1968.

15. Hammond,C. E.; and Doggett, R. V., Jr.: Determination of

Subcritical Damping by Moving-Block/Randomdec Applications.

NASASP-415, 1976.

16. Gaffey, Troy M.: The Effect of Positive Pitch-Flap

Coupling (Negative a3) on Rotor Blade Motion Stability

and Flapping, Journal of the American Helicopter Society,

Vol. 14, No. 2, April 1969, pp. 49-67.

10
Figure 1. Model rotor hub.

Figure 2. Details of rotorhub root flexures.

11
, i.

Figure3• ARES mountedin TransonicDynamicsTunnel.

•008 - -2 - •

•006- -3

FFT .004- LIA) -4

•002- -5 -

_' L--L.....L..J-I -6 I .. I ]. I

0 20 40 60 FREQ 0 1 2. ,3 4 TIME

AMP OF FFTVS. FREQ PEAKPLOT!

•200 ; ': "

: t

•150

AMPL .100 ," , .

•050 ,_

0 2 4 6 TIME

SIGNALTRACE

Figure4. Samplereal-timedisplayof moving-blockresults•

12
ANALYSIS

8 - 0 EXPERIMENT

BODYPITCH

5-

BO

FREQUENCY, 4

3

2

I- _ _-ROTOR LEAD-LAG

REGRESSINGMODE

0 t I I I I I

300 400 500 600 700 800

ROTORSPEED,rpm

(a)Modalfrequencies

ANALYSIS

0 EXPERIMENT

8-

4- 0

REGRESSINGLAG 0

0 MODEDAMPING 0

RATIO,PERCENT

CRITICAL

0 UNS_BLE

-4 I I I I I 1

300 400 500 600 700 800

ROTORSPEED,rpm

(b)Regressinglagmodedampingratio

Figure5. Comparisonof predictedandmeasuredstabilityas a

functionof rotorspeedin hoverforConfiguration1.

13
5 -

ANALYSIS

0 EXPERIMENT

4

REGRESSINGLAG 3 0

MODEDAMPING 0 RATIO,PERCENT

CRITICAL 2

(

i

0

I I I I I I

0 2 4 6 8 10 12

COLLECTIVEPITCH,deg

(a) Advance ratio = 0.20

5 -

--- ANALYSIS

4 - O EXPERIMENT /

REGRESSING

RAT

MODE

CRI

IO

DA

,

T

PERCENT

ICAL

MP

tAG

ING

2

3 ojo

1

I I I I I 1

O 2 4 6 8 10 12

COLLECTIVEPITCH,deg

(b) Advance ratio = 0.25

5

--ANALYSIS /

4 0 EXPER_o.

REGRESSINGLAG 3

MODE DAMPING

RATIO,PERCENT

CRITICAL 2

1 o

€

I I I I I 1

O 2 4 6 8 I0 12

COLLECTIVEPITCH,deg

(c) Advanceratio= 0.30

Figure 6. Lead-lag damping as a function of collective pitch

in forwardflight for baseline configuration.

14
B

ANALYSI S

_. O EXPERIMENT

3-

REGRESSINGLAG O

MODEDAMPING2 -

RATIO, PERCENT

CRITICAL

1- O

I I I I I I

0 2 4 6 8 I0 12

COLLECTIVPEITCH,deg

(a)Advanceratio= 0.20

ANALYSIS

3 - 0 EXPERIMENT __

REGRESSINGLAG

MODEDAMPING

RATIO,PERCENT2- _O

CRITICAL

II I I I I I

0 2 4 6 8 I0 12

COLLECTIVEPITCH,deg

(b) Advanceratio= 0.30

Figure7. Lead-lagdampingas a functionof collectivepitch

in forwardflightfor ConfigurationI.

15
ANALYSIS

0 EXPERIMENT O

4- 0

REGRESSINGLAG 3 - r

MODE DAMPING

RATIO,PERCENT

CRITICAL 2 -

iI I I I I I

0 2 4 6 8 I0 12

COLLECTIVEPITCH,deg

(a) Advanceratio= 0.20

m

ANALYSIS

O EXPERI

4-

REGRESSINGLAG 3 -

MODEDAMPING

RATIO,PERCENT

CRITICAL 2 -

O

I I I I I I

0 2 4 6 8 10 12

COLLECTIVEPITCH,deg

(b) Advanceratio= 0.30

Figure8. Lead-lagdampingas a functionof collectivepitch

in forwardflightfor Configuration2.

16
ANALYSIS

0 EXPERIMENT

3-

)

REGRESSINGLAG

MODEDAMPING2-

RATIO,PERCENT

CRITICAL

1-

I I I I I I

0 2 4 6 8 I0 12

COLLLECTIVEPITCH,deg

(a)Advanceratio= 0.20

m

ANALYSIS

O EXPERIMENT

3-

REGRESSINGLAG

MODEDAMPING 2 -

RATIO,PERCENT

CRITICAL

B

I I I I I l

0 2 4 6 8 10 12

COLLECTIVEPITCH, deg

(b) Advance ratio = 0.30

Figure 9. Lead-lag damping as a function of collective pitch

in forwardflightfor Configuration3.

17
ANALYSIS

O EXPERIMENT

3-

r

REGRESSINGLAG O

MODEDAMPING 2- O_

RATIO,PERCENT

CRITICAL

1-

I I I I I I :

0 2 4 6 8 10 12

COLLECTIVEPITCH, deg

(a) Advanceratio : 0.20

. ANALYSIS

O EXPERIMENT

3-

REGRESSINLAG G

MODEDAMPING 2-

RATIO,PERCENT

CRITICAL

1-

I I I I I I

0 2 4 6 8 10 12

COLLECTIVEPITCH,deg

(b) Advanceratio= 0.30 f

Figure10. Lead-lagdampingas a functionof collectivepitch

in forwardflightfor Configuration4.

18
EXPERIMENT ANALYSIS DROOP CONFIGURATION

Q 40 4

B 20 5

@ -2o 6

#

) 5- J

/ *

4 - /

I

REGRESSINGLAG t

MODEDAMPING 3 - ,, ,/I

RATIOCR,ITPERCE ICAL NT2- ,! _.._...... ,,.J ......-I", _, ,//,_'"

I I I I I I

0 2 4 6 8 10 12

COLLECTIVEPITCH, deg

Figure11. Effectof bladedroopangleon lead-lagdampingat

advanceratio= 0.30.

EXPERIMENTANALYSIS PRE-CONE CONFIGURATION 5-

• 30 4

• 6o 7

4-

/

/

/

REGRESSINGLAG 3 - //

MODEDAMPING i /_,. _

RATIO,PERCENT

CRITICAL 2 - • _... i

• ._ __- ..

1- _ •

I I I I I I

•) 0 2 4 6 8 10 12

COLLECTIVEPITCH,deg

Figure12. Effectof bladepre-coneangle on lead-lagdamping

at advanceratio= 0.30.

19
l. Report No. TM-85683 2 Government Accession No. 3. Recipient's Catalog No.

AVRADCOMTR 83-B-5

4. Title and Subtitle 5. Report.Date

Aeromechanical Stability of a Hingeless Rotor in Hover August 1983

and Forward Flight: Analysis and Wind Tunnel Tests 6.PerformOinrgg anizationCode

505-42-23-05

7. Author(s) 8. Performing Orgamzation Report No.

William T. Yeager, Jr., M-Nabil H. Hamouda,

Wayne R. Mantay 10. Work Unit No.

9. Performing Organization Name and Address

Structures Laboratory ¢

AVRADCOMResearch and Technology Laboratories 11.ContracortGrantNo.

NASALangley Research Center

Hampton_VA 23665 13. Type of Report and Period Covered

12. Sponsoring Agency Name and Address Technical Memorandum

National Aeronautics and Space Administration

Washington, DC 20546 14ArmyProject No.

and IL262209AH76

U.S. Army Aviation Research and Development Command

St. Louis, MO 63166

15. Supplementary Notes

William T. Yeager, Jr., and WayneR. Mantay, Structures Laboratory, AVRADCOMResearch

and Technology Laboratories. M-Nabil H. Hamouda, Vigyan Research Associates, Inc.

16. Abstract

A research effort of analysis and testing has been conducted in the NASA-Langley

Transonic Dynamics Tunnel to investigate the ground resonance phenomenon of a soft

in-plane hingeless rotor. Experimental data were obtained using a 9 ft. (2.74 m)

diameter model rotor in hover and forward flight. Eight model rotor configurations

were investigated. Configuration parameters included pitch-flap coupling, blade

sweep and droop, and precone of the blade feathering axis. An analysis based on a

comprehensive analytical model of rotorcraft aerodynamics and dynamics was used for

this study. The moving-block method was used to experimentally determine the

regressing lead-lag mode damping. Good agreement has been obtained between the

analysis and test. Both analysis and experiment indicated ground resonance instability in hover. The paper presents an outline of the analysis, a description of the

experimental model and procedures, and comparison of the analytical and experimental

data.

17. Key Words (Sugg_ted by Author(s)) 18. Distribution Statement •

Helicopters Unc!assified_Unlimited

Aeromechanica! stability _

Coupled rotor/airframe Subject Category, 39

Rotor dynamics

Hingeless rotors

19. S_urity Classif. (of this report) 20. Security Classif. (of this _ge) 21. No. of Pages 22. Price"

Unclassified Unclassified 20 A02

For saleby the National Technical Information Service,Springfield, Virginia 22161

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