Copyright ©1996, American Institute of Aeronautics and Astronautics, Inc.

AIAA Meeting Papers on Disc, January 1996A9618117, AIAA Paper 96-0149

Helicopter rotor tip shapes for reduced blade-vortex interaction - An experimental investigration. II

Baxter R. Jr. MullinsTexas, Univ., Arlington

Dudley E. SmithOklahoma, Univ., Norman

Curtis B. RathTexas, Univ., Arlington

Sherry L. ThomasBell Helicopter Textron, Inc., Fort Worth, TX

AIAA, Aerospace Sciences Meeting and Exhibit, 34th, Reno, NV, Jan. 15-18, 1996

The tip vortex structure of four main rotor tip configurations were examined experimentally in a series of transonic wind tunnel tests. The tip geometries used a NACA 0012 airfoil section and included a constant-chord rectangular planform, a tapered planform, a swept planform, and a swept-tapered planform. The tests were conducted at nominal Mach numbers of 0.75 and 0.78 with respective chord Reynolds numbers of 3.2 and 5.6 million. The latter Mach number-Reynolds number pair provide a full-scale operational match for the candidate rotor tip planforms. Total pressure surveys of the wake were conducted at three chordwise stations aft of each model at an equivalent rotor pitch angle of 8 deg. These tests revealed the initial roll-up region and provided a comparison of the effects of Mach number, Reynolds number, and tip shape on the vortex size, shape, location, and minimum stagnation pressure. The swept and swept-tapered planforms produced larger, more diffuse, tip vortices than the rectangular or tapered planforms. (Author)

Page 1

AIAA-96-0149

HELICOPTER ROTOR TIP SHAPES FOR REDUCED BLADE-VORTEXINTERACTION - AN EXPERIMENTAL INVESTIGATION, PART II

by

Baxter R. Mullins, Jr.*The University of Texas at Arlington

Arlington, Texas

Dudley E. Smith*The University of Oklahoma

Norman, Oklahoma

Curtis B. Rath*The University of Texas at Arlington

Arlington, Texas

and

Sherry L. Thomas5

Bell Helicopter-Textron, Inc.Fort Worth, Texas

Abstract

The tip vortex structure of four main rotor tipconfigurations were examined experimentally in a seriesof transonic wind tunnel tests. The tip geometries useda NACA 0012 airfoil section and included a constant-chord, rectangular planform, a tapered planform, aswept planform, and a swept-tapered planform. Thetests were conducted at nominal Mach numbers of 0.75and 0.78 with respective chord Reynolds numbers of3.2 and 5.6 million. The latter Mach number-Reynoldsnumber pair provide full-scale operational match for thecandidate rotor tip planforms. Total pressure surveys ofthe wake were conducted at three chordwise stations aftof each model at an equivalent rotor pitch angle of 8degrees. These tests provided a picture of the initialroll-up region and provide a comparison of the effectsof Mach number, Reynolds number, and tip shape onthe vortex size, shape, location, and minimumstagnation pressure. The swept and swept-taperedplanforms produced larger, more diffuse tip vorticesthan the rectangular or tapered planforms.

* Adjunct Professor, Mechanical and AerospaceEngineering, Senior Member AIAA.Associate Professor, Aerospace and Mechanical

Engineering, Member AIAA.Graduate Student, Student Member AIAAAssociate Engineer, Member AIAA

Introduction

Blade Vortex Interaction (BVI) is a major area ofstudy with regards to understanding and reducingrotorcraft noise. The interaction noise is a result of atrailing tip vortex from a leading blade passing near theregion of another blade. Rapid fluctuations of thepressure distribution along a blade occurs as a result ofthe relatively high velocities of the vortex passing nearthe blade, generating an intense BVI noise spectrum.Understanding the effect of blade tip shape on theformation of the tip vortex structure is important to theoverall problem of rotorcraft noise reduction.

Three basic approaches exist to minimize BVInoise: (1) avoid operating aircraft in the regime inwhich BVI occurs, (2) incorporate airfoils in the rotorsystem that are designed to minimize the interactioneffects, and (3) reduce the tip vortex interactionvelocities by rapid diffusion of the tip vortex.Operational considerations, such as the 6 degree IFR(Instrument Flight Rules) approach and performancedegradation, may limit many of these options. However,tip shapes may be modified without overbearingrestrictions being applied. Tip shape variations providea mechanism for vortex diffusion by varying local spanloading of the rotor.

The first phase of the research was an analyticalstudy to define candidate planforms that could be usedto reduce BVI noise. This analysis used the VSAEROcomputer code to examine the development of the tipvortex structure and its core development downstream1.

Copyright © 1995 by B. R. Mullins, Jr. Published by theAmerican Institute of Aeronautics and Astronautics, Inc. 1with permission

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Geometric variations evaluated included thicknesstaper, planform taper, sweep, and combinations of theseparameters. A constant chord, rectangular shape with aconstant NACA 0012 airfoil section blade was chosenas a baseline. The baseline blade tip planform waspragmatically varied and the relative effects of thevariations examined.

After candidate blade tips were chosen, each wasevaluated experimentally in a full-scale, low-speed windtunnel test to confirm and examine the vortex diffusioncapacity of the various geometries. These testsevaluated the following planforms: (1) rectangularbaseline, (2) taper, (3) sweep, (4) sweep with taper, (5)hyperbolic, (6) stubwing, (7) BERP, and (8) NASAStar2. In addition to wake velocity measurements, thesetests evaluated various surface flow visualizationtechniques including surface sublimation flowphotography and videos, smoke flows, and chiral-neumatic liquid crystal surface flow. The results ofthese tests were reported in Ref. 3. Of the eight tipsstudied in the low-speed wind tunnel test, only the firstfour were selected to be evaluated in a transonic windtunnel test.

Fig. 1 shows line drawings of the tip planforms.The objectives of this research were: (1) the generationof an experimental database for use in correlatingprediction methodologies, (2) the development ofanalytic design tools, and (3) the comparison of low-speed, full-scale tests to high-speed, model-scale tests.This was accomplished by constructing and testing four,10% scale, tip planforms at high Reynolds numbers andMach numbers.

Wind Tunnel Test

Transonic Wind Tunnel Facilities

The tests were conducted in the AerospaceResearch Center's High Reynolds number Transonicwind tunnel (HIRT) located on the campus of TheUniversity of Texas at Arlington (UTA). The UnitedStates Air Force Arnold Engineering DevelopmentCenter (AEDC) originally developed the Ludwieg-tubetunnel as a prototype concept for the National TransonicFacility. In 1978, the wind-tunnel was donated to UTAand in 1984 it became operational. A completedescription of the facility can be found in Ref. 4. A linedrawing of the tunnel is shown in Fig. 2 and alongitudinal-section drawing of the nozzle, test section,diffuser and sliding-sleeve starting value in Fig. 3.

The HIRT tunnel has exceptional flow quality andachieves high Reynolds numbers by operating at highpressure . The operational Reynolds number iscontrolled by the Mach number and the tunnel pressureas shown in Fig. 4. HIRT has design Reynolds numbers

:

,

M.t8

I

,[3 8

I

, ———————— ,.„„ ———————— .

« ——— I.»h ———

\\Jw

V^'^

\1T

(a) Rectangular Ttp (b) Swept Tip

—>| [*-»» ——•) |*-J»n

(c) Tapered Tip (d) Swept-Tapered Tip

Fig. 1. Line drawings of the tips tested in the UTAHIRT wind tunnel facility.capability of up to 7.5 million per inch for a Machnumber range from approximately 0.5 to 1.2. Nominalcharge tube pressure for this test program was 112 psi,which yields a Reynolds number of 5.6 million based ona 2.0 inch chord.

The test section is 7.34 inches high by 9.16 incheswide and 25.4 inches long with a conventional AEDCporous wall designed to control shock and expansionwave reflections as well as to provide boundary layercontrol. The porous walls vent to a plenum cavity thatexhausts through a ball valve which provided theprimary Mach number control.

Ludwieg-Tube Tunnel Operations

Ludwieg-tube wind tunnels employ an unsteadyexpansion wave for acceleration of high-pressure airstored in the charge tube to transonic Mach numbers.The entire tunnel is pressurized to the desired chargetube pressure level with the starting valve in the closedposition. The starting valve is then opened by means ofa pneumatic actuator, which generates an unsteadyexpansion wave that propagates upstream through thediffuser, test section, nozzle, and into the charge tube.This wave accelerates the primary flow from the chargetube into the nozzle and test section. A secondary flowthrough the porous walls in the test section is initiatedby cutting a diaphragm in the secondary flow system.This generates a second unsteady expansion wave thatpropagates into the test section plenum cavity. Once thetwo waves clear the test section, steady flow is

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33.8m

HIGH PRESSUREAIR CHARGINGSYSTEM

NOZZLE

r THRUST COLLAR

I CHARGE TUBE1 (33.3 cm dia)

• 2.1 m • BUILDINGWALL

£[TEST SECTION LINERPLENUM CHAMBER

EJECTOR FLAP CYLINDER

MODEL SUPPORT SECTION-SLIDING SLEEVESTARTING VALVE

EXHAUSTSILENCER

PLENUM EXHAUST

\_ THRUST STAND rEXHAUST SPHERE

Fig. 2. General line drawing of The University of Texas at Arlington's HIRT facility.

12-IN SLID ING SLEEVESONIC PLENUM PLENUM EXHAUST MODEL SUPPORT START VALVE

CHARGE NOZZLE SHELL (10 PLACES TYP) STRUT (SHOWN CLOSED)TUBE " — .. ^~13.94\

(NOM)X

a / '/EJECTOR//!

TEST SECTION WALL FLAPS INSTRUMENTATION LEADSSUPPORT STRUCTURE AND PRESSURE TUBE BUNDLE

Fig. 3.facility.

Sectional view of the nozzle, test section, diffuser and sliding-sleeve starting valve of the HIRT

maintained for the duration required for the expansionwaves to travel the length of the charge tube, reflect,and return to the test section.

The unsteady starting process takes between 40 to60 msec to stabilize, followed by about 120 msec of asteady, high-quality flow in the test section. Thereturning expansion wave causes a drop in pressureproducing a second shorter period of lower Machnumber steady flow. The Reynolds number of thissecond steady-flow period is about half that of theinitial flow period. As a result, two test points can begenerated during a single tunnel operation cycle.

Instrumentation/Data Acquisition System

The stagnation pressures downstream of the modelare measured with a wake rake mounted on a vertical

traversing mechanism. The rake consists of 13 totalhead tubes spaced 0.125 inches apart. Pressure leadsfrom the rake, the plenum cavity, and the charge tubestatic and total ports are connected to differentialpressure transducers. The plenum cavity transducer hasa range of 100 psig, the others have a range of 250 psig.The transducers, in turn, are connected to a 48 channeldata acquisition system that conforms to the ComputerAutomated Measurement and Control (CAMAC)standard. The transducer signal conditioning isprovided by programmable amplifiers. The analogsignals are concurrently converted to 12-bit digital dataand stored in a buffered memory. A microcomputer,connected to the data acquisition system via a cable,controls the data acquisition sequence. For thisexperiment the system was configured to record all ofthe pressure signals 1024 times at 0.5 msec intervals.

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0.2 1.0 1.20.4 0.6 0.8MACH NUMBER

Fig. 4. Operational capability of the Ludwieg tubetransonic wind tunnel.

Test Setup

Each model was attached to a base plate, Fig. 5,and then mounted on the side wall of the test section atthe prescribed distance ahead of the wake rake. Thebase plate had a wedge shaped leading edge, and wascurved so that it presented the same geometry to theoncoming flow, regardless of the angle of attack. Fig. 6shows the test setup of the model and the wake rake inthe tunnel test section. Angle of attack was set byselecting the appropriate mounting holes in the modelbase plate. The angle of attack was set at 8 degrees forall the test runs.

For each tip shape, the vortex is mapped at threedownstream locations behind the model. Since each testrun provides two periods of steady flow, the data isacquired for two Mach numbers and Reynolds numbers.The downstream locations for the test are 1 chordlength, 3 chords lengths, and 6 chord lengths behind thetrailing edge of the test article. The model was mountedas closely as possible to the wake rake withoutphysically interfering with the model for the 1 chordlength tests. The Mach numbers tested were nominally0.78 and 0.75. The test matrix is summarized in thetable.

Test Procedures

Vortex maps were constructed from the stagnationpressure measurements on grid points of a rectangularregion perpendicular to the freestream flow at presetchordwise positions downstream of the test article.Typically, 7 or 8 runs were required to map the test areafor a given configuration. Additional runs were requiredto locate the core of the vortex, and several more runs

Fig. 5. Base plate for airfoil model tests.

MODEL

HAKE HAKE OF 13TOTAL HEAD TUBES

Fig. 6. Pictorial view of the wind tunnel setup.

were made to insure repeatability of the data. Theseruns are referred to collectively as a sweep, since therake was moved from run-to-run across the vortex. Eachsweep produced two vortex maps, one per steady statetest interval.

Each run was performed according to standardprocedures. Immediately after each run, the charge tubepressure history was graphically displayed to the controlcomputer to verify that the data were obtained. Thecharge tube pressure history, an example is shown inFig. 7, allowed the operator to immediately determinethat the tunnel operated properly and that the dataacquisition system collected the data over the properinterval. This check was done because the tunnel bleedsystem occasionally did not work properly or the dataacquisition system would inadvertently triggerprematurely or not trigger at all. These lost runsconstituted about 5 percent of all the test runsconducted..

Each sweep was performed by setting the rake atthe top or bottom of the mapping region, thenperforming a data run. Next the rake was moved 0.125

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inch, up or down, to the next run position and a data setacquired. This process was repeated until the sweepwas completed. The rake was consistently moved inonly one direction to minimize systematic error inpositioning.

EXPERIMENT TEST MATRIX

150

Tip Shape

Rectangular

Swept

Tapered

Swept-Tapered

Mach

0.750.780.750.780.750.780.750.78

AOA(deg)

88888888

DownstreamLocation

<lc 3c 6cXX

XXXX

X

X

XX

XXXX

X

XXX

XX

X

X

Since the testing occurred over a period of a month,transducer drift was a concern. To minimize drift, thetransducers were calibrated prior to the start of eachdays testing. The calibration procedure, detailed in Ref.6, consists of monitoring a master transducer. Themaster transducer is a very stable and accuratetransducer and is used to measure the charge tube totalpressure. This master transducer is connected to adigital volt meter (DVM) and is monitored duringtunnel pressurization. It is this transducer that serves asthe standard for calibration of the other transducers.The calibration curve for each transducer is constructedby measuring the charge tube pressure at several stablepoints as the tube is being pressurized. At eachcalibration point, the master transducer and the othertransducers were read. Calibration curves for eachtransducer was then constructed.

Data Reduction

Each vortex map was constructed from thestagnation pressures measured over a grid downstreamof the wing tip. These measurements were made in asequence of test runs and recorded as time histories.The geometric parameters were nondimensionalized bythe root chord. The origin of the vortex map coordinatesystem, as shown in Fig. 8, is located at the blade-tip ofthe model at the root quarter-chord intercept. Thedistance downstream, d, is measured from the wing roottrailing edge. Fig. 9 shows the approximate size andshape of the area mapped by the wake surveys.

Next, the steady-state operation intervals of eachsweep were determined using the test section freestream

25D char jttubt mite

A mte raid port 13O chirp tubt stagnation

O pltnum cavtty

100 200 300 400TIME - (msec)

500 600

Fig. 7. A typical charge tube pressure history.

stagnation pressure for each run sweep as a function oftime (Fig. 10). From these graphs the first and secondsteady state intervals were determined. In general, thefirst intervals were generally about 80 to 100 msec long.The second interval was very short, with steady stateintervals on the order of 20 msec.

After identifying the steady-state intervals, thepertinent information was extracted and saved asindividual sweep pressure time histories. The pressureswere normalized based on freestream stagnationpressure. Time averaged normalized pressures over thespecified intervals were constructed so that the vortexmaps could be generated. Compensation for systemlags in the tube pressure and wake-rake systems wereprocessed using a standard correction procedure asdescribed in Ref. 6. The test procedure and data

Top View

Fig. 8. Definition of coordinate system used indeveloping vortex maps.

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- SIDE WALL PLATE

TEST AREA

• TIP SHAPE

7.34" x 9.16 " TEST SECTION WALL

Fig. 9. Looking upstream at the wind tunnel testsection showing the test area used in mapping thetip-shape vortex.

100 200 300 400TIME - (msec)

600

Fig. 10. Typical wake rake pressure measurementsused to select steady state intervals.

reduction methodology are described in Ref. 7 and Ref.

Vortex Maps

Vortex maps were constructed using 91 or 104pressure points evenly distributed over a rectangularregion normal to the flow (see Fig. 9) at distances of 1,3 and 6 chord lengths behind the trailing edge of themodel. The vortex core was defined by only 9 to 12pressure points in the test field. Figure 11 shows atypical vortex map produced from the wake rakepressure data. The vortex map is of the baseline,rectangular, tip-shape at one chord downstream of thetrailing-edge of the model. The "X" symbols note thelocation of the pressure ports relative to the rectangular

field, while the quarter chord of the tip shape is shownas a dashed line. The projected image of the deflected(8° angle of attack) tip-shape is shown as the solidoutline extending from the left hand side of the graph.Lines of equal pressure are in 0.01 increments, while agray-scale progressing from white to black is thepressure ratio gradient of the vortex map as shown bythe bar at the top of the vortex map. The numericalvalues above the gradient bar are the pressure ratios ofthe wake to the free stream value. The strength of thevortex can be determined by examining the minimumpressure ratio of the vortex and by calculating thepressure gradient across the vortex.

1.00 0.98 0.96 0.94 0.92 0.90 0.88 0.86

-0.4 -0.3 -0.2 -0.1 0.0 0.1 0.2 0.3X/C •

Fig. 11. Typical vortex map generated from thepressure measurements downstream of a tip shape.

Results

Test results are provided as four sets of vortexmaps, one set for each tip shape. Figure 12 shows thevortex maps for the rectangular tip, Figure 13 the swepttip, Figure 14 the tapered tip, and Figure 15 the tapered-swept tip. Each figure has three vortex maps showingfrom top to bottom the pressure ratio distribution at 1,3, and 6 chord lengths downstream of the wing trailingedge, except for Figure 13. Figure 13 has only twovortex maps calculated for 3 and 6 chord lengthpositions downstream of the swept tip. The pressuremeasurements could not be taken at the one chordposition as the swept tip interfered with the wake probe.

Rectangular Tip Shape

The three vortex maps for the rectangular tip areshown in Figure 12. The vortex map at one chord lengthdownstream (Fig. 12a) clearly shows the wake along thetrailing edge of the wing with a well defined tip vortex.The pressure ratio at the core center is less than 0.93and the pressure gradient is well defined. As the coremoves downstream to three chord lengths (Fig. 12b),the vortex core pressure ratio decreases to less than 0.86and the vortex core moves inward and down with

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respect to the tip. At six chord lengths downstream(Fig. 12c) the vortex core is well formed with a welldefined pressure gradient. The path of the vortex core atthis position is down and out toward the tip, followingan S-curve path with respect to the initial point.

Swept Tip Shape

Only two vortex maps were constructed for theswept-tip model. These are shown in Figure 13. Data atthe one chord length downstream position was not takenas the swept tip interfered with the wake rake. At thethree chord lengths position (Fig. 13a), the vortex corepressure ratio is less than 0.84 and the vortex core isabout 10% chord inboard of the tip. At six chord

1.00 0.98 0.96 0.94 0.92 0.90 0.88 0.86

•0.4 -0.3 -0.2 -0.1 0.0 0.1 0.2 0.3X/C

(a) 1 chord length downstream

-0.4 -0.3 -0.2 -0.1 0.0 0.1 0.2X/C

(b) 3 chord lengths downstream

0.3

-0.4 -btr-S2 -6!.i 676 6!i 672x/c

(c) 6 chord lengths downstream

Fig. 12. Rectangular tip vortex maps.

0.3

lengths downstream (Fig. 13c) the vortex core is welldefined with a well defined pressure gradient. The pathof the vortex from the 3 chord position to the 6 chordposition is nearly straight down. The pressure gradientis not as steep as that of the baseline tip shape.

1.00 0.98 0.96 0.94 0.92 0.90 0.88 0.86 0.84niiiiii

-0.4 -0.3 -0.2 -0.1 0.0 0.1 0.2 0.3X/C

(a) 3 chord lengths downstream

°-1 iiiiiii,..~,,,.,,™_,

-0.4 -0.3 -0.2 -0.1 0.0 0.1 0.2 0.3X/C

(b) 6 chord lengths downstream

Fig. 13. Swept tip vortex maps.

Tapered Tip Shape

Figures 14a through 14c are the vortex mapsgenerated for the tapered-tip. Figure 14a is the vortexmap constructed at the one chord length downstream ofthe wing tailing edge. This vortex map shows thedownwash along wing trailing edge and the initial tipvortex roll-up. The downwash from the wing appears tobe swept out beyond the wing tip before being caught inthe vortex roll-up as shown in Figure 14b. Although theinitial roll-up is more diffuse than that of the rectangulartip, by the time the vortex has traveled six chord lengthsdownstream, the tip vortex has rolled-up into a tightcore with a steep pressure gradient.

On the left-hand side of Figure 14a, the vortexmaps shows a large pressure ratio deficit. This deficitmay well be part of the vortex roll-up from the corner atthe transition of the straight leading-edge to the sweptleading-edge. Unfortunately, a full set of data along the

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wing span was not taken and the significance of thepressure deficit data can not be realized.

1.00 0.98 0.96 0.94 0.92 0.90 0.88 0.86

1.00 0.98 0.96 0.94 0.92 0.90 0.88

-0.4 -0.3 -0.2 -0.1 0.0 0.1 0.2 0.3X/C

(a) 1 chord length downstream

-0.4 -0.3 -0.2 -0.1 0.0 0.1 0.2 0.3X/C

(b) 3 chord lengths downstream

-0.4 -0.3 -0.2 -0.1 0.0 0.1 0.2 0.3X/C

(c) 6 chord lengths downstream

Fig. 14. Tapered tip vortex maps.

Swept-Tapered Tip Shape

Figure 15a through 15c show the vortex maps forthe swept-tapered tip shape. Figure 15a shows the wingdownwash and a small pressure deficit near the tip ofthe wing at one chord downstream displacement. Thetip vortex is not well defined. At three chord lengthsbehind the wing trailing-edge, a large and diffuse vortexis beginning to form. Even at six chord lengthsdownstream, the vortex is not well defined. The

-0.4 -0.3 -0.2 -0.1 0.0 0.1 0.2X/C

(a) 1 chord length downstream

i T^^^^TvT?^?^0.11 : :'-™^

0.3

-0.4 -0.3 -0.2 -0.1 0.0 0.1 0.2X/C

(b) 3 chord lengths downstream

-0.4 -0.3 -0.2 -0.1 0.0 0.1 0.2 0.3X/C

(c) 6 chord lengths downstream

Fig. 15. Swept-Tapered tip vortex maps.

pressure deficit in the core is the lowest value of thefour tips studied, with the minimum pressure ratio of0.87.

Vortex Core Pressure Deficits

Crosscuts through the vortex core of the taperedand swept-tapered tip shapes are provided in Figure 16..This graph shows the pressure ratio gradient and thepressure deficit for the two cases. The tapered-tip shapehad the sharpest pressure gradient and the largestpressure deficit of the four shapes tested. The swept-tapered tip had the most shallow pressure gradient andleast core pressure ratio deficit of the four tip shapes

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tested. As a result, the swept-tapered tip had the bestperformance of the four tip shapes. The swept-taperedtip shape produced a more diffuse pressure gradient andthe least loss in pressure in the core of the vortex. Thiswould result in reduced BVI and thus a better acousticperformance. The worst performing tip-shape was thetapered-tip shape. The tapered-tip shape had the largestpressure gradient in the tip vortex and it also had thegreatest core pressure loss.

1.00ogCC 0.95

ou_UlQUlCC

0.90

0.85UJCCQ.

0.80

TAPERED TIP

SWEPT-TAPERED TIP

-0.4 -0.2 0.0 0.2 0.4SPANWISE POSITION, X/C

Fig. 16. Comparison of the pressure gradient andpressure deficit of two tip shapes at six chordlengths downstream of the wing trailing edge.

Mach Number and Reynolds Number Effects

Each test run of the HIRT wind tunnel provides twoperiods in which steady-state data can be acquired. Thiswas done for each test run at nominal Mach numbers of0.75 and 0.78 and chord Reynolds numbers of 3.2 and5.6 million, respectively. Since the difference in Machnumber between the two steady-state regions was only0.03, the effect of Mach number could not bepractically evaluated. On the other hand, Reynoldsnumber variations were quite large allowing for acomparison of the results. The primary effect of anincreasing Reynolds number is to increase the pressuredeficit while the overall core diameter remains nearlyunchanged. Figure 17 shows the effect of Reynoldsnumber on the tip vortex of the tapered tip. The otherthree tip shapes have similar results with respect toReynolds number variations.

1.00

CC 0.95

Ou_UJQUJCCID(/)COUJCCCL

0.90

0.85

0.80

TAPERED TIPReynolds No.

5,700,000

3,200,000

-0.4 -0.2 0.0 0.2 0.4SPANWISE POSITION, X/C

Fig. 17. The effect of Reynolds number on thepressure ratio deficit of the tip vortex for theTapered Tip Shape.

Conclusions

This experiment provided data to be used in thestudy of Blade Vortex Interaction noise. The overallgoal of the test was to examine the tip vortex strength offour tip shapes by determining the tip vortex size andpressure deficit at high Mach and Reynolds numbers.The four tip shapes studied included (1) a baselinerectangular tip, (2) a swept-tip, (3) a tapered tip, and (4)a swept-tapered tip.

Of the tip shapes tested in the UTA's HIRT facility,the swept-tapered tip had the best aerodynamicperformance. This planform provided:

(1) the most diffused tip vortex core at six chordlengths downstream of the wing trailing-edge,

(2) the smallest tip-vortex core pressure deficit, and(3) the most shallow tip-vortex pressure ratio

gradient.The tapered tip had the worst overall performance. Thetip vortex rolled up to the tightest core with the largestpressure ratio gradient and the deepest pressure ratiodeficit. The order of performance from best to worstwas (1) the swept-tapered tip, (2) the swept-tip, (3) therectangular tip, and (4) the tapered tip.

Acknowledgment

The authors gratefully acknowledge the financialsupport of Bell Helicopter-Textron, Inc.

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References

1.Maskew, B., and Roo, B. M., "Unsteady Analysis ofRotor Blade Tip Flow " NASA CR3868, May 1985.

2. Sigl, D. G. and Smith, D. E., "Flow Over and BehindVarious Rotor Blade Tip Shapes: Wind Tunnel TestData Report," NASA CR182000, January, 1990.

3. Smith, D., and Sigl, D., "Helicopter Rotor Tip Shapesfor Reduced Blade Vortex Interaction: anExperimental Investigation," AIAA 95-0192, 33rdAerospace Sciences Meeting and Exhibit, Reno, NV,January 9-12, 1995.

4. Wilson, D. R. and Chou, S. Y., "Development of theUTA High Reynolds Number Transonic WindTunnel," AIAA Paper 85-0135, January 1985.

S.Starr, R. F., and Schueler, C. J., "ExperimentalStudies of a Ludwieg Tube High Reynolds NumberTransonic Tunnel, AEDC-TR-168, December 1973.

6. Kalkhoran, I. M., "An Experimental Investigation ofthe Perpendicular Vortex-Airfoil Interaction atTransonic Speeds," Ph.D. Dissertation, TheUniversity of Texas at Arlington, 1987.

7. Rath, C. B., "Wing Tip Vortex Stagnation PressureSurvey of Various Tip Shapes," MSAE Thesis, TheUniversity of Texas at Arlington, 1991.

S.Thomas, S. L., and Wilson, D. R., 'Total PressureSurveys and Force Balance Data for Rotor TipShapes," ARC Report 94-01, Mechanical &Aerospace Engineering Dept., The University ofTexas at Arlington, April 5, 1994.

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Copyright ©1996, American Institute of Aeronautics and Astronautics, Inc. Fig. 12

Copyright ©1996, American Institute of Aeronautics and Astronautics, Inc. Fig. 13

Copyright ©1996, American Institute of Aeronautics and Astronautics, Inc. Fig. 14

Copyright ©1996, American Institute of Aeronautics and Astronautics, Inc. Fig. 15