1

Insights into Airframe Aerodynamics and Rotor-on-Wing Interactions from a 0.25-Scale Tiltrotor Wind Tunnel Model

L.A. YoungD. Lillie

M. McCluerG.K. Yamauchi

Army/NASA Rotorcraft DivisionNASA Ames Research Center

Moffett Field, CA 94035

M.R. DerbyAerospace Computing, Inc.

Mountain View, CA

Abstract

A recent experimental investigation into tiltrotor aerodynamics and acoustics has resulted in the acquisitionof a set of data related to tiltrotor airframe aerodynamics and rotor and wing interactional aerodynamics.This work was conducted in the National Full-scale Aerodynamics Complex’s (NFAC) 40-by-80 FootWind Tunnel, at NASA Ames Research Center, on the Full-Span Tilt Rotor Aeroacoustic Model (TRAM).The full-span TRAM wind tunnel test stand is nominally based on a quarter-scale representation of the V-22 aircraft. The data acquired will enable the refinement of analytical tools for the prediction of tiltrotoraeromechanics and aeroacoustics.

Nomenclature

c Wing reference chord length, 2.0833 ftCD Airframe drag coefficientCL Airframe (including wing) lift

coefficientCN Wing section normal force coefficientCP Rotor power coefficientCT Rotor thrust coefficientDL Hover download, lbFM Rotor hover figure of meritiN Nacelle incidence angle, deg.MTip Rotor tip Mach numberpu, p Wing upper and lower wing static

pressures, psigr/R ‘Radial’ (measured from rotor hub)

spanwise location of wing staticpressure taps

Presented at the AHS InternationalAerodynamics, Acoustics, and Test andEvaluation Specialists’ Conference, SanFrancisco, CA, January 23-25, 2002. Copyright© 2002 by the American Helicopter SocietyInternational, Inc. All rights reserved.

R Rotor radius, 4.75 ftS Wing reference area, 23.9 ft2

T Rotor Thrust, lbV Wind tunnel test section velocity, fpsx/c Wing static pressure tap chordwise

locationαF Fuselage angle of attack, deg.αs Rotor shaft angle, deg, shaft vertical at

zero degrees angle, positive aftδf Flaperon angles, deg.δe Elevator angle, deg.µ Advance ratio, V/ΩRΩ Rotor rotational speed, radians/sec

Introduction

The development of the Tilt Rotor AeroacousticModel (TRAM) has been a joint effort of NASAAmes and Langley Research Centers and theU.S. Army Aeroflightdynamics Directorate(AFDD). Wind tunnel test results were acquiredin support of the NASA Short Haul CivilTiltrotor (SHCT) and Rotorcraft R&T Baseprograms. Two wind tunnel test campaigns have

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4. TITLE AND SUBTITLE Insights into Airframe Aerodynamics and Rotor-on-Wing Interactionsfrom a 0.25- Scale Tiltrotor Wind Tunnel Model

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7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES) Army/NASA Rotorcraft Division,Army Aviation and MissileCommand,Aeroflightdynamics Directorate (AMRDEC), Ames ResearchCenter,Moffett Field,CA,94035

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13. SUPPLEMENTARY NOTES Presented at the AHS International Aerodynamics, Acoustics, and Test and Evaluation Specialists’Conference, San Francisco, CA, January 23-25, 2002

14. ABSTRACT A recent experimental investigation into tiltrotor aerodynamics and acoustics has resulted in theacquisition of a set of data related to tiltrotor airframe aerodynamics and rotor and wing interactionalaerodynamics. This work was conducted in the National Full-scale Aerodynamics Complex’s (NFAC)40-by-80 Foot Wind Tunnel, at NASA Ames Research Center, on the Full-Span Tilt Rotor AeroacousticModel (TRAM). The full-span TRAM wind tunnel test stand is nominally based on a quarter-scalerepresentation of the V-22 aircraft. The data acquired will enable the refinement of analytical tools for theprediction of tiltrotor aeromechanics and aeroacoustics.

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been conducted to date with the TRAM teststand: an isolated rotor test in the Duits-Nederlandse Windtunnel (DNW) in December1997 and April-May 1998; a full-span, dual-rotor, complete-airframe wind tunnel model testin NASA Ames’ National Full-scale Complex(NFAC) 40-by-80 Foot Wind Tunnel inNovember-December 2000. In addition toacquiring fundamental aeromechanics andaeroacoustic data during these tests, wing staticpressure measurements were madecharacterizing tiltrotor airframe aerodynamicsand rotor-on-wing interactional aerodynamics.

This paper discusses the wing pressures andmodel balance measurements acquired with andwithout the rotors installed on the full-spanTRAM. Results from several different modelconfigurations and test conditions are presented.

Description of TRAM Test Stand

TRAM can be configured as either an isolatedrotor test stand or a full-span, dual rotor aircraftmodel. Figures 1 and 2 show the isolated rotorand full-span TRAM configurations,respectively. The rotors and test model arenominally a quarter-scale representation of theV-22 Osprey tiltrotor aircraft (Refs. 1-3).

Fig. 1 -- Isolated rotor TRAM configuration(Installed in DNW Open-Jet Test Section)

Fig. 2 -- Full-span TRAM configuration (Prior toNFAC Tunnel Entry)

The ability to configure the TRAM test stand aseither an isolated rotor or dual-rotor full-spanaircraft model makes it uniquely suited foridentifying and studying tiltrotor aerodynamicsand acoustics in a rigorous experimental manner.The drive train, nacelle assembly, hub and rotorcomponents from the isolated rotor model arealso used to make up the right-hand rotor andnacelle assembly on the full-span model. Eachrotor is powered by an electric motor designed todeliver up to 300 hp at 1588 rpm (through an11.3:1 gear reduction in the model drive train).An interconnect shaft insures that both rotors canbe driven by one motor, if necessary.

The right-hand rotor has a set of pressure-instrumented blades and a set of strain-gaugedblades; the left-hand rotor blades have only straingauge instrumentation. Only the strain-gagedblades were tested in the 40-by-80 Foot WindTunnel test. A rotating amplifier system wasused to condition signals from the heavilyinstrumented right-hand rotor. The TRAMblades and hub retention system are fabricatedwith composite materials and are structurallytailored to match fundamental frequencies of thefull-scale aircraft. The hub and control systemare kinematically similar to the full-scale aircraft.Both left-hand and right-hand rotor forces andmoments are measured with strain-gaugedbalances located in each pylon assembly. Theoverall aerodynamic loads on the full-spanmodel are measured with a strain-gauged balancelocated in the fuselage. A more detaileddescription of the TRAM test stands can befound in Refs. 1-2. A more comprehensivediscussion of the full-span TRAM 40-by-80 FootWind Tunnel test can be found in Ref. 4.

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Left-Hand-Wing and Flap Static PressureMeasurements

Wing pressure measurements were acquiredfrom 185 static pressure taps installed on the left-hand wing and flaperons of the Full-SpanTRAM. These measurements are used to studywing lift and normal forces in hover, helicoptermode, and airplane mode test conditions.

The wing static pressure tap locations, theexpected pressure range, and the maximumexpected dynamic pressure were defined, in part,based on results and insights from previoustiltrotor semispan wing testing (Ref. 5-6). Anelectronic pressure scanning system, made byPSI, was used to acquire data. The 185 taps aredivided amongst 4 modules, each with a range of±2.5 psid. Table 1 lists the locations of eachmodule.

Table 1 -- PSI Module Locations on Left Wing

Module Description Location48-port, 90 deg tube Cove/spar cap48-port, 90 deg tube Outboard flap64-port, 90 deg tube Inboard flap64-port, 90 deg tube Wing leading edge

The left wing’s spanwise pressure stations areshown in Fig. 3. Pressure taps are located alongthe span of the left wing and are identified as afunction of the left rotor radius, r/R (origin at theLH rotor shaft axis). Keeping in mind that theleft rotor turns clockwise when viewed fromabove, the definitions for the r/R station aredetermined when the retreating blade is parallelwith the wing’s forward swept leading edge.Spanwise (r/R) wing pressure stations areparallel with tunnel flow, and not wing chord.Note that fuselage taps were not installed for thefull-span TRAM testing. Fig. 4 shows a wingcross section at on spanwise station.

Fig. 3 – Planform View of TRAM Wing andSpanwise Location of Static Pressure Taps

Fig. 4 – Chordwise Wing Pressure Taps (OneSpanwise Station)

Table 2 summarizes the x/c chordwise locationsfor the wing pressure taps for the five spanwisewing r/R stations.

Table 2 – Left Hand Wing Pressure TapLocations

0.25 r/R 0.48 0.75 0.85 0.972UpperWing

.03,.06,.09,

.12,.15,.2,

.25,.3,.4,.5,

.6,.65,.68

.03,.06,.09,

.12,.15,.2,

.25,.3,.4,.5,

.588,.65,

.68

.03,.06,.09,

.12,.15,.2,

.254,.3,.4,

.5,.6,.65,

.68

.03,.06,.09,

.12,.15,.2,

.25,.3,.4,.5,

.6,.65,.68

.03,.06,.09,

.12,.15,.2,

.25,.3,.4,.5,

.6,.65,.68

UpperFlap.

.686,.696,

.71,.74,.77,

.8,.83,.86,

.9,.94,.98

.686,.696,

.71,.74,.77,

.8,.83,.86,

.9,.94,.98

.686,.696,

.71,.74,.77,

.8,.83,.86,

.9,.94,.98

.686,.696,

.71,.74,.77,

.8,.83,.86,

.9,.94,.98

.686,.696,

.71,.74,.77,

.8,.83,.86,

.9,.94,.98LowerWing

.001,.03,

.05,.09,.15,

.3,.68

.001,.03,

.05,.09,.15,

.25,.68

.001,.03,

.05,.09,.2,

.6,.68

.001,.03,

.05,.09,.12,

.15,.68

.001,.03,

.05,.09,.12,

.15,.68LowerFlap.

.71,.74,.77,

.83,.9,.98.71,.74,.77,.83,.9,.98

.71,.74,.77,

.83,.9,.98.71,.74,.77,.83,.9,.98

.71,.74,.77,

.83,.9,.98

Heater boxes were not be used with the PSImodules during the full-span TRAM test. Since

4

data drift caused by thermal effects wasanticipated, a frequent calibration of the PSIsystem was performed. The PSI system wascalibrated by taking 1024 samples at 10 differentpressure settings (ranging from ±2.5 psid). Thecalibration and test data were low-pass filtered at~4Hz using 1024 point FFT. The calibrationcoefficients were computed using a 2nd orderleast squares fit on the averages from eachpressure setting. The reference pressure for themodules was the tunnel barometric pressure. Thelast port of each module measured the tunnelbarometric pressure. Hence, the last port of eachmodule provided an indication of the amount ofsignal drift during a test run, since, ideally, thisport should provide a zero reading with zerodrift. The results from each calibration werewritten to a computer file. Raw pressure dataand reduced engineering unit measurements weresaved to computer files for each data point.

Experimental Results

This paper discusses three categories ofexperimental observations from the full-spanTRAM 40-by-80 Foot Wind Tunnel test. First,rotor and wing interactions in hover will bediscussed, with emphasis on wing download.Second, rotor-on-wing interactions in helicopter-mode forward-flight will be discussed,presenting rotor and model loads and wingpressure distributions for rotor-on and rotor-offtest points. Finally, balance loads and wingpressures (airframe only, no rotors) in airplane-mode forward-flight will be presented.

Rotor and Wing Interactions in Hover

Figure 5 shows the installation of the full-spanTRAM test stand in the National Full-scaleAerodynamics Complex (NFAC) 40-by-80 FootWind Tunnel. One of the key concernsunderlying any wind tunnel investigation is theimpact of the tunnel test section on the testmodel (and rotor) aerodynamics andperformance, and the determination ofinstallation configurations and/or tunnelcorrection methodologies that minimize thatinfluence. This is a particularly crucial issue forhover performance measurements in an enclosedwind tunnel environment. Wall effects,including flow recirculation, can be considerableif the rotor diameter is too large, or the diskloading too high, for the test section size.

Fig. 5 – Full-Span TRAM Installation in theNFAC 40-by-80 Foot Wind Tunnel

The NFAC 40-by-80 foot test section is anelliptical test section that has ceiling ‘clam shell’doors that open up to expedite the installation oftest models using an overhead high-capacitycrane. The TRAM rotor hubs – at a fuselageangle of attack of zero degrees – are 21.8 feetabove the tunnel floor, when the model is inhelicopter-mode (iN = 90 deg.). To minimizeadverse wall effects on the full-span TRAM rotorhover performance measurements, the tunnelceiling clam shells were left fully open, themodel was pitched nose down to a fuselage angleof attack of –9 deg., and the tunnel (downstream)air exchangers were open to 50%. Figure 6shows for a relatively low thrust coefficient of0.0043 that that there is minimal observedinfluence, or variation, of the wind tunnel walleffects on rotor performance for the full-spanTRAM installation (as represented by theminimum variation in the rotor powercoefficient). As can also be seen in Fig. 6, theremay be a slight ‘ground effect’ due to the rotorwake interaction with the tunnel floor(manifesting as a slightly reduced power level).Sweeping the fuselage angle of attack to -9 deg.appears to yield slightly high power levels.

5

0

0.0001

0.0002

0.0003

0.0004

0.0005

0.0006

-15 -10 -5 0 5 10 15 20Fuselage Angle of Attack (Deg.)

CT = 0.0043(Average)

Fig. 6 – Influence of Wind Tunnel Test SectionInteractions in Hover

Figure 7 compares the hover figure of meritperformance data from the full-span TRAM test(iN=90 deg. and αF=-9 deg.) against isolatedrotor TRAM data from the DNW (airplane-modeconfiguration). Both primary and secondary, leftand right, rotor balance loads are included in thefull-span figure of merit data. The blade, rotor,and nacelle assembly designs used for theisolated rotor and full-span TRAM tests areidentical; in many cases the same hardware wasused for the two tests (the full-span TRAM right-hand hub and nacelle assembly were used in theDNW isolated rotor test). Figure 7 shows asignificant interactional (rotor-on-rotor andwing-on-rotor) aerodynamic effect in the full-span performance data. This result ispreliminary but suggests that furtherinvestigation of interactional aerodynamicphenomena for tiltrotor hover performance iswarranted. Though interactional aerodynamicinvestigations have been performed for tiltrotoraircraft previously (notably, Ref. 7), TRAMtesting represents an opportunity tosystematically explore these phenomena using aversatile modular test stand. It is anticipated – ashas already begun with the isolated rotor TRAMdata (Ref. 8) – that full-span TRAM data willenable the development/refinement ofcomprehensive design and analysis tools fortiltrotor aircraft.

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1

0 0.005 0.01 0.015 0.02Thrust Coefficient

TRAM DNWIsolated RotorFS TRAM 40-by-80 Ft WT

Fig. 7 – Influence of Wing/Airframe on RotorHover Performance

Two representative chordwise static pressuredistribution plots are shown in Fig. 8, forspanwise locations 0.25 and 0.85 r/R. Thechordwise location, x/c, is given in terms of thestatic pressure taps ‘undeflected’ (with respect toflaperon angle) positions. The rotor tip Machnumber is 0.629, the left-hand rotor thrust is 378lb (CT = 0.0046), the nacelle incidence angle is90 degrees, and the flaperon angles areuniformly set at 70 degrees.

Spanwise Station r/R = 0.25

-0.2

-0.15

-0.1

-0.05

0

0.05

0.1

0 0.2 0.4 0.6 0.8 1 1.2

Chordwise Station, x/c

Win

g Pr

essu

re, p

sig

Wing Upper Surface Lower Surface

(a)

Spanwise Station r/R = 0.85

-0.2

-0.15

-0.1

-0.05

0

0.05

0.1

0 0.2 0.4 0.6 0.8 1 1.2

Chordwise Station, x/c

Win

g Pr

essu

re, p

sig

Wing Upper Surface Lower Surface

(b)Fig. 8a-b– Sample of Wing Chordwise Pressure

Distributions (CT = 0.0046)

6

Flaperon cove seals are implemented on the wingby thin, pre-loaded, fiberglass strips. Ideally, noflow (from lower to upper surface) should occuras a consequence of the fiberglass cove seals.The flaperons were remotely controlled in twodiscrete angle ranges: –10 to +35 deg. (positivetrailing edge down) for airplane-mode operationand +35 to +100 degrees for hover andhelicopter-mode forward flight. The flaperons’mechanical linkages must be manually adjustedto switch between the two discrete operatingranges. For each set of flaperons on the left andright side of the wing there was a pair of inboardand outboard flaps. Though each flaperon (ofthe four total) can be remotely controlled, allflaperons were set at the same angle for all testconditions. A separate set of fiberglass coveseals was used for the two different flaperonranges.

A large amount of hover wing pressure data wasacquired during the full-span TRAM test. Thisincluded data at several different rotor thrustconditions, flaperon angles, and nacelleincidence angles. Wing pressure data can beintegrated (from r/R = 0, at the wing tip, tor/R=1.21, at the fuselage plane of symmetry) toprovide estimates of the wing sectional normalforce (i.e. download for hover test conditions), asshown in Fig. 9. Figure 9 is derived from thesame test point and wing pressure data set thatFig. 8a-b are samples plots of.

0

0.05

0.1

0.15

0.2

0 0.2 0.4 0.6 0.8 1 1.2

Wing Span, r/R

Fig. 9 – Representative Wing IntegratedSectional Download Results (CT=0.0046)

Spanwise integration (applying the wingsymmetry condition) of the Fig. 9 results yieldsan estimate for the TRAM download-to-thrustratio of DL/T=0.105 for this test condition. Thisestimate of download-to-thrust ratio is typical ofresults previously noted in the literature (Refs. 5-7, 9).

Both chordwise and spanwise integration of thewing pressure data was accomplished by simpletrapezoidal rule numerical integration. Theintegration limit values for the leading andtrailing edge of the flaperons were assumed to bethe mean value of the nearest adjacent upper andlower surface pressure tap measurements, forchordwise integration. A leading edge pressuretap was provided for the main wing section. Forspanwise integration, the sectional normal forceloading at the wing tip (r/R=0) was assumed tobe zero and to be constant/uniform loading fromthe last measurement station, r/R=0.972, tor/R=1.21 (fuselage plane of symmetry).Symmetry (left- versus right-hand wing) wasalso assumed.

Figure 10 is a qualitative comparison of hoverdownload results between two tiltrotortests/configurations. Hover download estimatesfrom the 0.25-scale full-span TRAM integratedwing pressures are (iN = 90 deg., MTip = 0.63,and δf= 70 deg.) compared to results from a0.658-scale JVX proprotor, semispan wing, andimage plane, Ref. 5 (iN = 85 deg., MTip = 0.71,and δf= 67 deg.). Two observations readily standout from this qualitative comparison. First,despite the nontrivial model configuration andoperating condition differences between the twotests, the download results appear to be in goodagreement with each other for thrust coefficientsgreater than 0.005. Second, the download-to-thrust ratio trend, though seemingly linear and/orasymptotic at thrust coefficients representative ofthe nominal operating conditions forconventional tiltrotor aircraft, appears to steeplyrise for very low thrust coefficients.

0

0.05

0.1

0.15

0.2

0 0.005 0.01 0.015 0.02

Thrust Coefficient

DL

/T

FS TRAM - Integrated WingPressures0.658-Scale JVX w. SemispanWing (NFAC 40-by-80), Ref. 6

Fig. 10 – Qualitative Data Comparison BetweenTiltrotor Hover Download Tests/Configurations

7

Figure 11 is a comparison of the full-spanTRAM hover download test results derived fromintegrated wing pressures versus the measureddownload from the model balances. Thedownload-to-thrust ratio from the modelbalances was derived as DL/T = 1 - TF/(TR + TL),for hover testing at nacelle incidence angles ofninety degrees. TR and TL are rotor thrustmeasurements from the right and left rotorbalances, and TF is the resolved model ‘thrust’measurement from the fuselage balance, whichmeasures the total model loads – including rotorloads. Figure 11 also demonstrates that thedownload contributions of the fuselage – inaddition to the wing download component – is asignificant fraction of an overall ‘airframe’ (wingand fuselage) hover download. This result isconsistent with the experimental observations ofRef. 10 which estimated the fuselagecontribution to total download to be between 36to 42%. Still, there is considerable data scatter inthe fuselage balance download results andadditional testing in the future will be required tofully characterize the fuselage contribution toaircraft download.

0

0.05

0.1

0.15

0.2

0.25

0 0.002 0.004 0.006 0.008 0.01

Thrust Coefficient

DL

/T

Integrated Wing Pressures

Balance Measurements

Fig. 11 – Hover Download as a Function ofThrust Coefficient

Further insight can be gained regarding tiltrotordownload-to-thrust ratio trends by examining thewing sectional download spanwise distribution.Integrated wing pressure data in Fig. 12 allowstwo observations to be made. First, most of thewing download is generated near the wing tip(close to the rotor axis). Second, as rotor thrustis decreased, proportionally more and more ofthe wing download is carried at the wing tip.This is a consequence of highly twisted tiltrotorblades only carrying positive thrust on the

inboard blade sections (and, thereby, producing adownwash) while the outboard blade is unloadedor negatively-loaded (resulting in very littledownwash or, perhaps, even an upwash) at lowrotor collectives.

-0.05

0

0.05

0.1

0.15

0.2

0.25

0.3

0.35

0 0.2 0.4 0.6 0.8 1 1.2

Wing Spanwise Station, r/R

Rat

io o

f Win

g Se

ctio

nal N

orm

al F

orce

to

Rot

or T

hrus

t

CT=0.0046CT=0.0044CT=0.0060CT=0.0097CT=0.0032CT=0.0025CT=0.0018

Fig. 12– Wing Sectional Download Distributionas a Function of Rotor Thrust Coefficient

Another qualitative comparison can be madebetween the full-span TRAM integrated wingpressure hover download data and the Ref. 60.658-scale JVX data. Figure 13 shows thedependence of tiltrotor hover download on wingflaperon angle. The qualitative agreementbetween the Ref. 6 data and the full-span TRAMresults is quite good. A flaperon angle of δf ~70deg. results in a minimum wing download, forboth cases.

0.000

0.020

0.040

0.060

0.080

0.100

0.120

40 50 60 70 80 90 100

Flaperon Angles, deg.

DL

/T

0.658-Scale JVX & Semispan Wing& Image Plane, Ref. 6Full-Span TRAM Integrated WingPressures

Fig. 13 – Tiltrotor Download as a Function ofFlaperon Angles

8

Insight into the Fig. 13 results can be arrived byconsidering the wing sectional normal-force-to-thrust ratio spanwise distributions for severalflaperon angles, Fig. 14, and selected wingchordwise pressure profiles, Fig. 15a-b. Thegeneral character of the spanwise distributions inFig. 14 does not to appear to change withflaperon angle other than a shift in overallmagnitude of the sectional normal-force-to-thrustratio curves. (The two fitted curves (at δf=70and δf=90 deg.) highlight the minimum andmaximum observed download measurements.)

0

0.02

0.04

0.06

0.08

0.1

0.12

0.14

0.16

0.18

0 0.2 0.4 0.6 0.8 1 1.2

Wing Spanwise Station, r/R

Rat

io o

f Win

g Se

ctio

nal N

orm

al F

orce

to

Rot

or T

hrus

t

Flap = 65 Deg.Flap = 60 Deg.Flap = 70 Deg.Flap = 76 Deg.Flap = 80 Deg.Flap = 86 Deg.Flap = 90 Deg.

Fig. 14—Wing Download Distribution as aFunction of Wing Flaperon Angle (CT=0.008)

The wing chordwise pressure distributions ofFig. 15a-b, for flaperon angles of 60 and 70degrees (at spanwise stations 0.25 and 0.85 r/R),reveal subtle upper surface pressure profiledifferences, particularly at the leading andtrailing edge of the wing. The lower surfacechordwise pressures are indistinguishable fromeach other.

-0.25

-0.2

-0.15

-0.1

-0.05

0

0.05

0.1

0.15

0 0.2 0.4 0.6 0.8 1 1.2

Chordwise Station (Undeflected Coordinates), x/c

Pres

sure

, psi

g

Upper Surface, Flap = 60 Deg.Lower, 60 Deg.Upper Surface, Flap = 70 Deg.Lower, 70 Deg.

(a) r/R = 0.25

-0.25

-0.2

-0.15

-0.1

-0.05

0

0.05

0.1

0.15

0 0.2 0.4 0.6 0.8 1 1.2

Chordwise Station (Undeflected Coordinates), x/c

Pres

sure

, psi

g

Upper Surface, Flap = 60 Deg.Lower, 60 Deg.Upper Surface, Flap = 70 Deg.Lower, 70 Deg.

(b) r/R = 0.85

Fig 15 – Wing Chordwise Pressure Profiles forDifferent Flaperon Angles

The full-span TRAM hover download trends areconsistent with results obtained with tiltrotorsemispan wing and image plane test models,such as Refs. 5 and 6. It is interesting to notethat for wing download studies, tiltrotor modelsusing semispan wings and image planes may besufficient for hover wing download studies. Thishas been a point of considerable debate withinthe tiltrotor research community for severalyears. The same can not necessarily be said forrotor performance measurements; the literaturedoes not reveal anywhere near the level of wing-on-rotor and/or rotor-on-rotor (or image-plane onrotor) interactions on rotor performance forsemispan test models as has been observed forthe full-span TRAM hover results (Fig. 7).

9

Rotor and Wing Interactions in Helicopter-ModeForward-Flight

Figure 16 shows the influence of the wing andairframe (and the second rotor) on rotorperformance in helicopter-mode forward flight.The rotor thrust is directly measured from theTRAM rotor balances. The rotor thrust andpower coefficient data is shown for an advanceratio, µ, of 0.15. The isolated rotor performancedata is from the TRAM DNW test, Ref. 3.Figure 16 can only be considered a preliminaryexamination of the influence of wing and/orairframe interactions on helicopter-mode forwardflight rotor performance. No wind tunnel wallcorrections have been applied to the data shownin Fig. 16. As noted in Ref. 8, wind tunnelcorrections can be very important for correlatingdata between tests and between experiment andanalysis. In particular, Ref. 8 identifiedsatisfactory corrections for the isolated rotorTRAM DNW data.

0

0.0001

0.0002

0.0003

0.0004

0.0005

0.0006

0.0007

0.0008

0 0.005 0.01 0.015 0.02Thrust Coefficient

Pow

er C

oeff

icie

nt

FS TRAM NFAC 40-by-80 FtWTTRAM DNW Isolated Rotor

Fig. 16 -- Wing/Airframe on Rotor Interactionsin Helicopter Mode Forward Flight (MTip = 0.63,

µ=0.15, αF = +11 deg., and αS = +6 deg.)

Hub weight and aerodynamic tares have beenapplied in Fig. 16 to the rotor power and thrustcoefficients for both the isolated rotor and full-span TRAM data sets. In both cases the samemethodology (and hub configuration) was usedto derive and apply the hub tares. Figure 17shows the hub tare configurations for the isolatedrotor and full-span configurations.

(a) isolated rotor

(b) full-span TRAM

Fig. 17 – Hub Tare Configurations

The influence of the rotor wakes on tiltrotorairframe aerodynamics can be seen in Fig. 18 forδf = 43 deg., δe = 0 deg., and iN = 85 deg.(nacelles pitched five degrees forward of verticalwhen the fuselage angle of attack is zerodegrees). The tunnel velocity is 62 knots. Therotor thrust coefficient for the rotor-onconfiguration is CT=0.008. Due to problemswith the left hand rotor balance during thehelicopter-mode forward flight testing,symmetry has been assumed between the loadsfor the left and right hand rotor balances and so(two-times) the right hand balance thrust resultswere subtracted from the fuselage balance loads.The wing is ‘clean’ – no vortex generators orfences are installed; no forebody strake isrepresented on the model.

10

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

-20 -10 0 10 20

Fuselage Angle of Attack (Deg.)

Air

fram

e A

erod

ynam

ic

Coe

ffic

ient

s

Airframe Only LiftCoeff.

Airframe OnlyDrag Coeff.

Airframe LiftCoeff. With RotorInteractionsAirframe DragCoeff. WithInteractions

Fig. 18 – Influence of Rotor on Wing/AirframeInteractions on Airframe Lift and Drag

The rotor-on-wing interactions seen in Fig. 18are shown further in Figs. 19 and 20. In thiscase, the influence of varying rotor thrust is evenmore clearly seen on both the airframe lift anddrag results. Figure 19 and 20 are results for afuselage angle of attack of +11 deg. (and a rotorshaft angle of –5 deg.). Unfortunately, thrustsweeps for other fuselage angles of attack werenot acquired during the full-span TRAM test.

0.5

1

1.5

2

0 0.005 0.01 0.015 0.02

Rotor Thrust Coefficient

Air

fram

e Li

ft C

oeffi

cien

t

Airframe Only (Baseline Value)Lift Coeff. With Rotor Interactions

Fig. 19 – Influence of Rotor Thrust on TiltrotorAirframe Lift

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0 0.005 0.01 0.015 0.02

Rotor Thrust Coefficient

Air

fram

e D

rag

Coe

ffic

ient

Airframe Only (Baseline Value)Drag Coeff. With Rotor Interactions

Fig. 20 -- Influence of Rotor Thrust on TiltrotorAirframe Drag

Figure 21 shows the influence of rotor-on-winginteractions on wing normal force coefficients, asa function of fuselage angle of attack, as derivedfrom integrated wing pressures. Figure 21results are for the same model configuration andtest conditions as Fig. 18 results – i.e. CT=0.008,δf = 43 deg., δe = 0 deg., iN = 85 deg, andµ=0.15. The rotor and fuselage balancemeasurements (Fig. 18) and the integrated wingpressure data (Fig. 21) are both in generalagreement as to the effect of rotor-on-winginteractions observed during helicopter-modeforward-flight testing of the full-span TRAM. Itappears that the observed effect of the rotor-on-wing interactions is primarily one of influencingthe onset and severity of wing stall. This muststem from the rotor wake’s unsteady flowadversely interacting with the wingaerodynamics. On the other hand, the meaninduced vertical velocities of the rotor wake (at aCT=0.008) do not appear to be of sufficientmagnitude to significantly reduce the wingangle-of-attack. More wind tunnel testing willbe required to fully understand wing-on-rotorinteractions.

11

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

-15 -10 -5 0 5 10 15 20

Fuselage Angle of Attack (Deg.)

Win

g N

orm

al F

orce

Coe

ffic

ient

Airframe OnlyWing-on-Rotor Interactions

Fig. 21 – Effect of Rotor-on-Wing Interactionson Wing Normal Force Coefficients

Aerodynamics of Airplane-Mode Forward-Flight

Figure 22 shows the TRAM model in itsairplane-mode (rotor-off) forward flightconfiguration. The TRAM model was tested upto 291 knots in this configuration. Fuselagebalance and wing pressure data was acquired.

Fig. 22 – Full-Span TRAM in Airplane-Mode(Rotors Off)

Figures 23 and 24 show the lift curve and lift-drag polars for the full-span TRAM in airplane-mode forward flight. Lift and drag data arepresented for two different wing Reynoldsnumbers: 3.3x106 (V = 150 knots) and 6.4x106

(V = 291 knots). The airplane-mode

configuration is δf = 0 deg., δe = –17 deg., and iN

= 0 deg. (δe = –17 deg. was primarily set tominimize the pitching moment load carried onthe fuselage balance). The lift curve results forFig. 23 compare reasonably well with Ref. 11computational and measured results for the V-22tiltrotor aircraft.

-1

-0.5

0

0.5

1

1.5

2

-15 -10 -5 0 5 10 15 20

Fuselage Angle of Attack (Deg.)A

irfr

ame

(Onl

y) L

ift C

oeffi

cien

t

Wing Reynolds # = 3.3E+6

Reynolds # = 6.4E+6

Fig. 23 – TRAM Airframe-Only Airplane-ModeLift Curves

0

0.05

0.1

0.15

0.2

0.25

0.3

0.35

0.4

-1 -0.5 0 0.5 1 1.5 2

Lift Coefficient

Air

fram

e (O

nly)

Dra

g C

oeff

icie

nt

Wing Reynolds # = 3.3E+6

Reynolds # = 6.4E+6

Fig. 24 – TRAM Airframe-Only Airplane-ModeLift/Drag Polar

Tare and interference (T&I) corrections for themodel mounting strut have not been made in thelift and drag data presented in Figs. 23 and 24.Attempting to acquire T&I corrections in the 40-by-80 Foot Wind Tunnel with the full-spanTRAM installation would be an arduous task thatmay or may not be feasible given the large sizeof the facility and test model. As such this

12

airframe lift and drag data derived from themodel fuselage balance can not be fullyconsidered representative of the ‘free air’aerodynamics of a tiltrotor aircraft.

Figure 25 are representative wing chordwisepressure distributions for the TRAM wing inairplane-mode forward-flight (rotors-off) for r/R= 0.25 and 0.85, respectively. The fuselageangle of attack is zero degrees and δf = δe = 0deg. The tunnel dynamic pressure is 13.5 psf(tunnel velocity is 63 knots).

Spanwise Station r/R = 0.25

-0.2

-0.15

-0.1

-0.05

0

0.05

0.1

0 0.2 0.4 0.6 0.8 1 1.2

Chordwise Station, x/c

Win

g Pr

essu

re, p

sig

Wing Upper Surface Lower Surface

(a) r/R = 0.25

-0.2

-0.15

-0.1

-0.05

0

0.05

0.1

0 0.2 0.4 0.6 0.8 1 1.2

Chordwise Station, x/c

Win

g Pr

essu

re, p

sig

Wing Upper Surface Lower Surface

(b) r/R = 0.85

Fig. 25a-b – Representative Airplane-ModeForward-Flight (Rotors-Off) Wing Pressure

Distributions

Tunnel velocity and fuselage angle-of-attacksweeps were performed for the rotor-offairplane-mode testing. Integrated sectionalnormal force coefficient spanwise distributionestimates can be made from the wing pressuredata. Figure 26 shows the (left-hand wing)

normal force coefficient distribution for the sametest condition as the Fig. 25 data.

0

0.1

0 .2

0 .3

0 .4

0 .5

0 .6

0 0.2 0 .4 0 .6 0 .8 1 1.2W ing S pan, r/R

Sect

iona

l Win

g N

orm

al F

orce

Coe

ffici

ent

Fig. 26 – Representative Sectional Normal ForceCoefficient Spanwise Distribution

The relative contribution of the wing versus theairframe (wing plus fuselage) normal forcecoefficient can be seen in Fig. 27. The twocurves of Fig. 27 were derived from fuselagebalance measurements for the airframe normalforce coefficient curve and integrated wingpressure data for the wing normal force curve.Two observations can be made from Fig. 27.First, integration of the wing pressure dataappears to yield satisfactory integratedaerodynamic coefficients. Second, the fuselagedoes appear to be a significant contributor to theairframe normal force for large fuselage anglesof attack. Future work can focus on comparingfull-span TRAM wing pressure data andsectional normal force coefficient properties withexisting airfoil, wing, and unpowered forcemodel for the V-22 tiltrotor aircraft (Refs. 11 and12, for example).

-0.6

-0.4

-0.2

0

0.2

0.4

0.6

0.8

1

1.2

1.4

-15 -10 -5 0 5 10 15

Fuselage Angle of Attack (Deg.)

Nor

mal

For

ce C

oeffi

cien

t (B

ody

Axi

s)

Airframe (Wing &Fuselage) - Balance

Wing - WingPressures

Fig. 27 – Airframe Versus Wing Normal ForceCoefficient Contribution (iN = 0 deg., δf = 0 deg.,

δe = -17 deg., V = 104 knots, Re = 2.3x106)

13

The wing sectional normal force coefficientspanwise distributions for the Fig. 27 data areshown in Fig. 28. Figure 28 is based onintegrated wing pressure data. Over all, giventhe results of Figs. 27 and 28, reasonable wingsectional normal force values can be derivedform the full-span TRAM wing pressure data.This data will be a valuable aid in the refinementof comprehensive rotorcraft aeromechanicsanalysis tools and computational fluid dynamicsanalyses.

-0.6

-0.4

-0.2

0

0.2

0.4

0.6

0.8

1

1.2

1.4

0 0.2 0.4 0.6 0.8 1 1.2

Wing Spanwise Station, r/R

Win

g Se

ctio

nal N

orm

al F

orce

C

oeffi

cien

t ALPHA_F = -9 Deg.ALPHA_F = -3 Deg.ALPHA_F = 0 Deg.ALPHA_F = 6 Deg.ALPHA_F = 9 Deg.ALPHA_F = 12 Deg.

Fig. 28 – Influence of Fuselage Angle of Attackon Wing Spanwise Distributions

Concluding Remarks

The paper discusses in detail several observedtiltrotor aerodynamic characteristics asevidenced by the TRAM wing static pressureand model balance load measurements. Theseairframe aerodynamics and rotor and winginteraction observations encompass threeprimary operating regimes for tiltrotor aircraft:hover, helicopter-mode forward flight, andairplane-mode cruise forward flight.Comparisons are made between isolated rotorTRAM data acquired in the DNW wind tunneland data from the recent full-span TRAM test inthe Ames Research Center NFAC 40-by-80 FootWind Tunnel. Hover results are comparedbetween the 0.25-scale full-span TRAM andpreviously reported 0.658-scale JVX proprotorand semispan wing tests results acquired in theNFAC. The results from this paper, thoughpreliminary in nature, point to several important

interactional aerodynamic phenomena fortiltrotor aircraft.

The Tilt Rotor Aeroacoustic Model projectcontinues to explore fundamental tiltrotoraeromechanics and aeroacoustics. Planning forfollow-on tests are currently underway. Thismajor research investment on the part of NASAand the U.S. Army will enable the developmentof a new generation of tiltrotor analytical designtools and advanced tiltrotor technologies.

Acknowledgements

The experimental results in this paper werederived from research performed under theauspices of the NASA Short Haul Civil Tiltrotor(SHCT) and the Rotorcraft R&T Base programs.The success of the full-span TRAM developmentand initial wind tunnel testing was due to theoutstanding support provided by members of theAmes Research Center Army/NASA RotorcraftDivision, the U.S. Army AeroflightdynamicsDirectorate, and the Wind Tunnel OperationsDivision. Finally, the authors would like toextend special acknowledgement of theprogrammatic and technical contributions ofWilliam J. Snyder to civilian tiltrotor technology,and to the TRAM project in particular. Thanks,Bill, for never losing the faith.

References

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14

Meeting on Advanced RotorcraftTechnology and Disaster Relief, Gifu,Japan, April 1998.

2. Johnson, J.L. and Young, L.A., “Tilt RotorAeroacoustic Model Project,”Confederation of European AerospaceSocieties (CEAS) Forum on Aeroacousticsof Rotors and Propellers, Rome, Italy, June9-11, 1999.

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Annual Forum of the American HelicopterSociety, Montreal, Canada, May 25-27,1999.

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6. Felker, F.F., Shinoda, P.R., Heffernan,R.M., and Sheehy, H.F., “Wing Force andSurface Pressure Data form a Hover Test ofa 0.658-Scale V-22 Rotor and Wing,”NASA TM 102244, February 1990.

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9. Liu, J., McVeigh, M.A., Mayer, R.J., andSnider, R.W., “Model and Full-ScaleTiltrotor Hover Download Tests,” 55th

Annual Forum of the American Helicopter

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