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SIXTH EUROPEAN ROTOR.CRAFT AND POWERED LIFT AIRCRAFT FORUM .

Paper No. 25

EXPERIMENTAL AND ANALYTICAL STtiDIES OF A ,..:- IE- •

MODEL HELICOPTER. ROTOR IN HOVER "C:

Aeromechanics Laboratory U.S. Army Research and Technology Laboratories

Moffett Field, California 94035 U.S.A.

September 16-19, 1980 Bristol, England

THE UNIVERSITY, BRISTOL, BS8, lHR, ENGLAND

DISTRIBUTION STA

Approved tor publ1e releaM; Distribution Unllmltec!

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EXPERIMENTAL AND ANALYTICAL STUDIES OF W'MODEL HELICOPTER ROTOR IN HOVER

F. X. CaradonnaC. Tung

Aeromechanics LaboratoryU.S. Army Research and Technology Laboratories (AVRADCOM)

Moffett Field, California 94035 U.S.A.

ABSTRACT

The present study is a benchmark test to aid the development of variousrotor performance codes. The study involves simultaneous blade pressure measure-ments and tip vortex surveys. Measurements were made for a wide range of tipMach numbers including the transonic flow regime. The measured tip vortexstrength and geometry permit effective blade loading predictions when used asinput to a prescribed wake lifting surface code. It is also shown that withproper inflow and boundary ayer modeling, the supercritical flow regime canbe accurately predicted.

LIST OF SYMBOLS

A = ratio of vortex circulation to maximum blade-bound circulation

AR = aspect ratio

C= = sectional lift coefficient

d - radial distance from a vortex to a flow-field point

R = radius of the rotor blade

r = radial distance from the rotor center of rotation

V = vortex-induced velocity<iVR = residual velocity in the wake

y = r/R, nondimensional radial coordinate

z = axial distance from rotor

Q=- rotational speed

T = azimuthal angle measured from the point of blade overhead passage

T v = vortex age, the azimuth angle, T, when vortex strikes the probev

1. Introduction

The past two decades have seen a continuing development of methods topredict rotor hover performance with increasing accuracy. These methodsinclude lifting line [1,2,31, lifting surface [4,5,6], and finite difference [7]methods. Practically speaking, none of these methods is self-contained; theyall require the specification of empirically obtained wake data (strength andgeometry) in order to have a correct downwash distribution. Inevitably, thedevelopment of these codes becomes a tuning process in which it is determinedjust how detailed and accurate a wake description must be. This stage of codedevelopment places great reliance on the available body of experimental rotordata.

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The available rotor data include a sizeable number of tests where detailedblade loading is obtained using surface pressure transducers [8,9,10,11], andmore recently by laser doppler velocimetry [12,13). There is also a number oftests in which the rotor wake geometry is defined by flow visualization tech-niques [5,14] for a wide variety of blade configurations. Of the various wakestudies, only Boatwright [15] and Cook [16] made detailed investigations ofthe wake flow structures. Cook's work is especially significant in that hewas able to measure the strength of the tip vortex by a curve-fitting techniqueusing hot-wire data. However, there seem to be no useable data in the literaturein which simultaneous blade load distribution and wake measurements are made.

It is the intention of the present study to help fill this gap in theliterature. This paper will describe the experimental set-up in which steadyblade pressures were obtained using hub-mounted transducers and tip vorticeswere measured using Cook's technique. The data obtained are for unstalled flowranging from the low subsonic to transonic conditions. It is shown herein (usingSumma's lifting surface code [61) that the blade-loading distribution could nothave been predicted using only the classical vortex data of Landgrebe 114) andKocurek [5].

2. The Experiment

The data presented in this paper were gathered in the Army AeromechanicsLaboratory's hover test facility, a large chamber with special ducting designedto eliminate room recirculation. The rotor, situated in the center of thechamber, was mounted on a tall column containing the drive shaft (Fig. 1). Therotor employed two cantilever-mounted, manually adjustable blades. These bladesused an NACA 0012 profile and were untwisted and untapered. An aspect ratioof 6 was chosen in order to maximize Reynolds Number and available instrumenta-tion space. The blades were grooved to accommodate 60 pressure tubes each.These tubes connect to a special cluster of three 48S8 Scanivalves (usingStatham PA 856-15 transducers) driven by one SS5-48 solenoid drive mounted inthe rotor hub. This arrangement permits an ample number of ports for five

measurement locations - three radial locations on each blade, with one locationbeing identical on both blades for comparison purposes. The Scanivalvestepper motor was actuated by a digital data system which acquired the data,computed the centrifugal pressure drops, and displayed the final pressuredistribution. The resulting pressure distriubtions for collective pitchsettings of 50, 80, and 12° are shown in Figs. 2, 3, and 4. It is seen herethat the inboard pressure distributions are only slightly affected by rotorspeed. However, the outboard sections show considerable pressure alterationand shock development as the tip Mach number approaches near sonic values.Overall, however, the spanwise load distribution (obtained by pressure integra-tion) is remarkably little affected by tip Mach number (Fig. 5).

Wake data were acquired with a traverse-mounted DISA 55P01 hot-wire probemounted beneath the rotor. Data from the wire are acquired at various pointsalong the tip vortex trajectories and can give both the tip vortex geometry

and strength. One problem with this approach is that the vortex trajectory isnot steady and the probe location (which is chosen by an on-the-spot decisionas to where the number of vortex core "hits" is maximized) contains some as yetundetermined error. The resulting data stream has considerable variability.However, in order to be certain of the vortex location, the only acceptabledata are those where the vortex core actually hits the probe. In the digitiza-tion process (done off-line at a reduced tape speed), the above-mentioned datasystem was coded to look for and accept only those data which showed the

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characteristic signal dip which indicates a probe-vortex strike. This turnsout to be a very small percentage of the total amount of data actually recorded.A typical hot-wire trace displaying the above-mentioned variability is shown inFig. 6.

3. Hot Wire Data Analysis

The idea of the current data analysis is that a tip vortex should looklike an infinite line vortex to a sufficiently close probe. Unfortunately, theprobe measures not only the velocity induced by the vortex at hand, Vi, but also

that induced by the blade and the remainder of the wake system as well, VR.The problem in analyzing the probe data is, then, how to separate this residualvelocity, VR, from the immediate vortex-induced velocity, Vi. Cook [16] handledthis probem by assuming that the residual velocity was constant and given by thetranslation velocity of the tip vortex. He then was able to find the vortexstrength by a fitting process. This strength was found to be far less than thecomputed maximum blade bound circulation of the single, full-scale blade used inthat test. It was also found that the vortices measured were distinctly non-classical in that they contained a large rotational region outside of the viscouscore. In what follows, we shall use a process very similar to Cook's in analyz-ing wake data.

First consideration is given to the vortex trajectories. Figure 7 showsthe axial and radial components of the vortex trajectories for a pitch settingof 80. This figure gives data for a wide range of rotor speeds, and it isapparent that the trajectory is essentially independent of tip speed - even intothe transonic regime. Figure 7 together with Fig. 5 suggests that the nonlineartransonic flow on the blade has little effect on the far-field induced flow aslong as the local lift is not greatly altered. Also plotted on this figure isthe vortex trajectory given by Kocurek's wake-fitting formula for rotors infree air. Although the axial component of the trajectory compares well withKocurek's formula, there appears to be a greater discrepancy in the contractionthan can be explained by measurement error. The vortex trajectories for pitchsettings ranging from 50 to 120 are given in Fig. 8.

The present aim in analyzing the rotor wake is only to find the vortexstrength and not a complete description of the structure. This strength willbe found by fitting the wake data to the velocities obtained from an appropriatecombination of inviscid, two-dimensional vortices. The velocity from one suchvortex is given by

Vi (C Y)ma__ _ ____ _ 1(1)

SIR 47rAR d/R

where the strength of the vortex is described by A, the ratio of the vortexcirculation to the maximum bound circulation of the blade (determined from thepressure data). To accomplish this fitting, it is first necessary to convertthe spatially dependent Equation (1), into a time-dependent expression, as thevortex data are tme-based. Assuming that A is constant (which seems to betrue within reasonable error bounds), the conversion to a time-dependentfunction is accomplished by expressing d as a function of time using the vortextrajectory data of Fig. 8. The next step is the determination of the residualvelocity, V1 , which must be vectorially added to Vi before a comparison can bemade with te probe data. We have done this in two different ways:

1) The first way to determine VR involves very young vortices (about 50*old). For these it was assumed that VR was given by the vortex trajectories(Fig. 8). The fitting process always commenced when the vortex core hit the

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probe and ended when the following blade passed over; this assured the simplestpossible flow field, as there would be vortices on only one side of the probeand minimal influence of vortex sheets and blade bound vorticity. Figure 9 showssome typical comparisons of probe data with the fitting expression. This figureshows the vortex velocity-time traces for pitch settings of 80 and 120. It isseen here that the fitting curve provides a good match to the data outside ofthe immediate core region. Furthermore, the vortex strength is very close tothe maximum blade bound circulation.

2) A second means to determine VR was required in analyzing older

vortices (about 210* old). The flow is more complex in this case, as the probealways lies between two vortices in the fitting region, and the expression forthe vortex-induced velocity is correspondingly complicated. In fact, Vi, forthis case was determined using three vortices - one outboard of the probe andtwo inboard. Again, the data were fit for the time period between a probe-vortex strike and the subsequent blade passage. It was found that with VRdetermined by the vortex trajectory data, it was not possible to obtain a goodfit of the classical vortex expression to the wake data. Instead we found thata better value for VR was found by use of the minimum measured velocity betweentwo vortices. At this point, the vortex-induced velocity is small, but notzero (due to the differing instantaneous translation velocities of the threevortices). The minimal induced velocity is calculated (assuming some valueof A) and subtracted from the minimum measured inter-vortex velocity toobtain VR. This task was rendered quite simple by the fact that the radialcomponent of these velocities turns out to be very small (this was checked bycalculations and measurements with a second probe). Since the two methodsabove do not give the same value for the residual velocity, it is clear thatVR is not a constant in this case. We assume, however, that it changes suf-ficiently slowly to render the fitting process meaningful. In fact, the resultsthus obtained are consistent with the young vortex data. Figure 10 shows sometypical comparisons of the older vortex data with the fitting expressions. Thiqfigure shows the fittings for pitch settings of 50, 80, and 12. It is seenthat the 8* and 120 cases show vortex strengths which match the maximum blade-bound vorticity very well. At 5* pitch, however, the strength is seen to beconsiderably smaller.

It seems from the above data, which are taken at a low rotor speed, thatthe tip vortex develops its full strength very early in life (mainly before 500).Although there is a fair amount of variability between vortices, it is ratherstriking that very many vortices closely approach a classical Rankine vortexin appearance. Furthermore, the vortices (except for the 50 case) seem to con-tain all of the blade circulation. This vortex strength and structure differsmarkedly from the result obtained by Cook and probably reflects the considerabledifferences in blade geometries. As rotor tip speed increases (Fig. 11), how-ever, there appears to be an increasing departure from the Rankine vortexappearance. Nevertheless, the nondimensional vortex strength seems unaffectedby tip speed.

4. Comparison of Theory and Experiment

In order to integrate the present wake and loading data into a believablewhole, it is necessary to be able to reproduce the blade loading computationally.We have chosen to do this using A.M.I.'s lifting surface code [6]. This is avery flexible, compressible, lifting surface code which can handle eitherprescribed or free wakes.

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Initial efforts to compute the blade loading were done using the Kocurekwake geometry [5]. The resulting computed thrust coefficient was too high byabout 20%. The next step was to compute the loading using the measured vortexlocations and strength. Figure 12 shows a comparison of the measured andcomputed loading using the measured vortex parameters for a collective pitchof 8* (the trajectory is given by Fig. 8 and we choose A - 1.0). The comparisonis now considerably improved and the thrust coefficient is over-predicted byless than 5%. In view of the previously mentioned uncertainties in the vortextrajectory measurements, these computations were also performed with the entirevortex trajectory perturbed such that at T = 1800, the axial and radial perturba-tions were ±0.025R. The results derived from all possible combinations of theseaxial and radial changes fill the shaded area in Fig. 12. That the abovemeasured and computed results are roughly centered on this shaded region indicatesthat for this case the measured trajectories are fairly accurate. However, thebest comparison with the measured loading occurs when the vortex radial location(at T = 1800) is increased (that is, the contraction is decreased) by 0.025R.The identical situation was found to occur in computations of the 12* pitchcases; i.e., the best comparison occurred when the radial vortex location wasincreased by 0.025R over the measured value (Fig. 13). For the 5° collective-pitch case, the situation was a little different in that a reasonable comparisonof computation and loading data could not be made until the vortex strength wasreduced to A = 0.75. In this case, the vorticity which would otherwise havebeen in the tip vortex was now included in the vortex sheet model. (For acomplete description of the assumed vortex sheet model see Ref. [6]). Thisresult is consistent with the measured vortex strength and gives the comparisonshown in Fig. 14.

The previous comparisons have been made at low tip Mach numbers. Thelifting surface code used should be applicable to predict the spanwise and chord-

wise loading up to the onset of supercritical flow. Beyond this point, linearaerodynamics are not applicable on the blade and a more complete flow descriptionis required. As a preliminary evaluation of the high-speed flow data, two-dimensional computations were made of the flow at the 80% radial station. Thiswas done using Holst's full-potential code [17]. In order to perform thiscomputation, an angle of attack is required. Since the region of supersonicflow is localized (i.e., limited to the immediate vicinity of the upper bladesurface), it should be possible to find the angle of attack using the linearlifting surface code. Of course, the lifting surface code requires the measuredvortex location and strength as mentioned previously. With the angle of attackobtained thereby, the Holst code produced the results shown in Fig. 15. Thisfigure shows two computed results - an inviscid result and one with a viscousramp-boundary layer model [18]. It is seen that a shock-boundary layer inter-action model is very necessary and in this case very effective.

5. Concluding Remarks

The present study was intended as a benchmark to aid in the developmentof hover performance codes. The goal was to obtain simultaneous measurementsof blade load distribution and tip vortex geometry and strength using fairlystandard techniques. In spite of some uncertainties (due mainly to wakeunsteadiness), lifting surface computations show that the load and wake measure-ments are generally consistent with each other.

The main conclusions from this study are:

1. The Cook vortex measurement technique seems to be quite effective fortwo-bladed rotors.

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2. At low rotor speeds, an untwisted, untapered, double-bladed rotorproduces tip vortices which can closely resemble a classical Rankine vortex.Except for the lowest pitch settings, this vortex strength closely approachesthe maximum blade bound circulation. At higher tip speeds, the inner vortexstructure appears increasingly nonclassical; however, the strength is unaltered.

3. It is not possible to predict the blade-spanwise-load distributionwithout accurate vortex location and strength data. The present measured vortexlocation data were significantly different (for presently unknown reasons) fromthe classical data in the literature. However, these measurements were indispen-sable to obtaining a reasonable comparison of theory and experiment.

4. For the present rotor and speed range tested, the onset of transonicflow was found to have no effect on the spanwise loading distribution and thevortex trajectories. The chordwise loading is profoundly altered by the transonicflow and can only be simulated by nonlinear aerodynamic techniques which employa shock-boundary layer interaction model.

ACKNOWLEDGMENTS

This work represents the contributions of many excellent people. We wouldlike to extend our thanks to W. D. Vann (Ft. Eustis Directorate, U.S. ArmyApplied Technology Lab) and H. Jones (U. S. Army Research and Technology Labs)who were instrumental in initiating our computational studies. Special acknowl-edgment is due to Georgene Laub who tirelessly and ably assisted in the runningof the test. Thanks also to M. Summa (Applied Mechanics Inc.) who wrote thelifting surface code and assisted us in running it; S. C. Lee (University ofMissouri) and T. L. Holst (NASA Research Center) who provided us with the finitedifference computations.

REFERENCES

1. P. Crimi, Theoretical prediction of the flow in the wake of a hoveringrotor, CAL Report No. BB-1994-S-l and No. BB-1994-S-2, Cornell AeronauticalLaboratory, Inc., Buffalo, New York (September 1965).

2. A. J. Landgrebe, An analytical method for predicting rotor wake geometry,J. American Helicopter Soc., 14, No. 4 (October 1969).

3. A. J. Landgrebe, An analytical and experimental investigation of helicopterrotor hover performance and wake geometry characteristics, USAAMRDLTechnical Report 71-24, Eustis Directorate, US Army Air Mobility Researchand Development Laboratory, Fort Eustis, Virginia (June 1971).

4. W. Johnson, A lifting surface solution for vortex induced airloads andits application to rotary wing airloads calculations, MassachusettsInstitute of Technology, Aeroelastic and Structures Research Laboratory,TR 153-2 (April 1970).

* 5. J. D. Kocurek and J. L. Tangler, A prescribed wake lifting surface hoveri! performance analysis, Presented at the 32nd Annual National Forum of the

American Helicopter Soc., Preprint No. 1001 (May 1976).

6. J. M. Summa and D. R. Clark, A Lifting-surface method for hover/climbloads, Presented at the 35th Annual Forum of the American Helicopter Soc.,Washington, D.C., Preprint No. 79-1 (May 1979).

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7. F. X. Caradonna, The transonic flow on a helicopter rotor, Ph.D. Thesis,Stanford University (March 1978).

8. J. P. Rabbott Jr., Static-thrust measurements of the aerodynamic loadingon a helicopter rotor blade, NACA TN 3688, Langley Aeronautical Laboratory,National Advisory Committee for Aeronautics, Langley Field, Virginia(February 1956).

9. J. Scheiman and H. L. Kelley, Comparison of flight-measured helicopterrotor-blade chordwise pressure distributions with static two-dimensionalairfoil characteristics, NASA TN D-3936 (1967).

10. P. Brotherhood and C. Young, The measurement and interpretation of rotorblade pressures and loads on a puma helicopter in flight, Presented atthe Fifth European Rotorcraft and Powered Lift Aircraft Forum, Amsterdam,The Netherlands (September 1979).

11. R. B. Gray, H. M. McMahon, K. R. Shenoy, and M. L. Hammer, Surface pressuremeasurements at two tips of a model helicopter rotor in hover, NASACP-3281 (May 1980).

12. J. P. Sullivan, Experimental investigation of vortex rings and helicopterrotor wakes using a laser doppler velocimeter, D. S. Dissertation,Massachusetts Institute of Technology (June 1973).

13. J. D. Ballard, K. L. Orloff, and A. B. Luebs, Effect of tip shape on bladeloading characteristics, Presented at the 35th Annual National Forum of theAmerican Helicopter Soc., Washington, D.C., Preprint No. 79-1 (May 1979).

14. A. J. Landgrebe and B. V. Johnson, Measurements of model helicopter rotorflow velocities with a laser doppler velocimeter, Tech. Note, J. AmericanHelicopter Soc., 19, No. 2 (July 1974).

15. D. W. Boatwright, Measurement of velocity component in the wake of a fullscale helicopter rotor in hover, USAAMRDL TR 72-33 (August 1972).

16. C. V. Cook, The structure of rotor blade tip vortex, AGARD CP-ll(September 1972).

17. T. L. Holst, A fast, conservative algorithm for solving the transonicfull-potential equation, AIAA Paper 79-1456 (July 1979).

18. S. C. Lee, Effect of turbulent boundary layer on transonic flows,Preliminary Report, NASA Interchange Number NCA2-OR450-001 (August 1979).

-- ~~~ ~ ' t[ 1dorn

'*i 25-7

krail and/or

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HUB·SCANIVALVE ASSEMBLY

Fig . 1 . The model and experimental set-up.

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rIR -0.18 nHR 0 .9

.2.0---$I -12S~rpm

-2.0 mop, 0.434CT -=0.00213

nIR - .90 -n 1750 rpmMtip - 0.607

CpA. CT -=0.00218

--- nl 22G8rp,=tp 0.794

CT -0.0200

1.

nIH 0.5 OR - O.K8

Fig. 2. Measured pressure distributions. Collective pitch, 0 - 56.

25-,

r/R =0.09 OIR 0.96

----- -1250 rpm

CT - 0.00459

rIR 0.90 fl -1750 rpm

CP Mtip 0.612-1.0 % %CT - 0.00455

-f- 2250 rpm=tp 0.794

0 CT = 0.00462

0ti - 0.977

CT - 0.00473

rIR =0.50

O/R =0.68

Fig. 3. Measured pressure distributions. Collective pitch, e. 80.

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rIR 0.89 r/R 0.96

-- 2= 1250 rpm-2.0 Mtip - 0.433

% ~CT -=0.00796

O/R0.80 - - 12 -1750 rpm

Cpi - 0.6101.0CT - 0.0087

-- 2f = 2279 rpmMtp0=.794

CT = 0.007920

I,.5 1.0

OIR =0.50 O rR =0.68

Fig. 4. Measured pressure distributions. Collective pitch, ec 120.

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.6TEST DATA, , -=80

o 12 - 1260 rpm CT -0.00460

o El -200 rpm, CT=O0.00 461[3 f2-2500 rpm, CT -O.00464

4 REGION OF LOADING VARIATIONFA DUE TO RPM CHANGES

2In

U.

.3

.4 zS0

.2

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~VORTEX TRAJECTORY(TOP VIEW)

PROBE-VORTEX COREINTERSECTION POINT

SECOND BLADEVORTEX SHEET PASSAGE

FIRST BLADEPASSAGE

Fig. 6. Typical wake probe data.

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0

.2

0 El - 50rpm

.3 A a - 250rpmrA+ n - 2540, 2414 rpm N

x0 n - 2260 rpm N

.AL CURVE FIT OF PRESENT DATA0 -- KOCUREK'S FORMULA( C ICT-=0.0046

.8- -

1.0

0 50 100 150 200 250 300 350 400 4W0VORTEX AGE, P's. dug

Fig. 7. Wake geometry measurements for various rotor speeds and comparisonwith classical data. Collective pitch, ec 8.

0

.3--

=C a- 012

.1 N

A.

0 so 100 ISO 200 250 300 300#W do%, VORTEX AGE

Fig. 8.Wake geometry for various pitch settings.

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.8

A = 0.90

VVCR.8 A

A= 1.0

A

1 ' V'v =53.8

VM.44

0 0, 12' 1800

Ov-52.3'

Fig. 9. Typical vortex velocity-time trace and 1/R curve fit for variouspitch settings. Vortex age - 50 ° (nominal). 0 = 1250 rpm.

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.8 V .8

.44

.8 V/fZ

.8 A- 1.0 .4

01800 8C- .36 0

Ov -2150

1300203

Fig. 10. Typical vortex velocity-time trace and 1/R. curve fit for Variouspitch settings. Vortex age -2000 (nominal). nl 1250 rpm.

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A-093

.5.5

Z 120rpim 180Ov= 480

V/nR A -0.92

0 n 1750 rpim 180

-:.0 Iw

Fig. II. Typical velocity-time trace and 1/ft curve fit for various rotorspeeds. Collective pitch, ec M 8% Vortex age, *Va 5O-650.

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* TEST DATA fe a~ IrR -150 mls 1401 fps). CT - 0.0046-MEASURED WAKE GEOMETRY, CT - 0.0048

VORTEX CONTRACTION REDUCED AILFIG

A BY 0.025 R. CT AM 0.0I FA T EORYA - I REGION OF Ce VARIATIN DUE TO WAKE SRAETER

P ..... GEOMETRY CHANGESz

iwo

Z0 .2

0 .2 .4 .. 31.0rIR (RADIAL STATION)

Fig. 12. Effect of vortex position on loading computation.

.6

z.4

3

.20 ETDT e12.n 5 h1911$.C .0

--w...CDUIGMAURDWK EMTY T.01

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' T'

* TEST DATA 0, S nR =150 6ms (401 fVs),CT- 0.0021 I

AMI CODE USING MEASURED WAKE ,

GEOMETRV. CT" 0.0023 IINCOMPLETE TIP-VORTEX ROLL UP, IA-.75 CT 0.00216

//

//

/

___________ ,__ i I p

.2 .4 .6 .8 1.0

Fig. 14. Comparison of measured and computed loading.

0 TEST DATA AT 0.81 MT .8770a = 2.100 (FROM A.M.I. CODE)

-1.0 /F --- FINITE DIFFERENCECODEINVISCID [T. HOLST]

- FINITE DIFFERENCE CODEWITH VISCOUS EFFECTI

IO

01

Cp

0

-1 lI I

0 A .6 .6 1.0

Fig. 15. Comparison of measured and computed chordvise pressure distribution.

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