flow environment study near the empennage of a 15-percent...
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March 2000
NASA/TP-2000-210085AFDD/TR-00-A-004
Flow Environment Study Near theEmpennage of a 15-Percent Scale HelicopterModel
Susan Althoff GortonLangley Research CenterHampton, Virginia
John D. BerryDirectorate of Aviation EngineeringUS Army Aviation and Missile CommandRedstone Arsenal, Huntsville, Alabama
W. Todd Hodges and Deane G. ReisAeroflightdynamics DirectorateUS Army Aviation and Missile CommandLangley Research Center, Hampton, Virginia
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March 2000
NASA/TP-2000-210085AFDD/TR-00-A-004
Flow Environment Study Near theEmpennage of a 15-Percent Scale HelicopterModel
Susan Althoff GortonLangley Research CenterHampton, Virginia
John D. BerryDirectorate of Aviation EngineeringUS Army Aviation and Missile CommandRedstone Arsenal, Huntsville, Alabama
W. Todd Hodges and Deane G. ReisAeroflightdynamics DirectorateUS Army Aviation and Missile CommandLangley Research Center, Hampton, Virginia
Available from:
NASA Center for AeroSpace Information (CASI) National Technical Information Service (NTIS)7121 Standard Drive 5285 Port Royal RoadHanover, MD 21076-1320 Springfield, VA 22161-2171(301) 621-0390 (703) 605-6000
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Summary
Development of advanced rotorcraft configurations has highlighted a needfor high-quality experimental data to support the development of flexible andaccurate analytical design tools. To provide this type of data, a test programwas conducted in the NASA Langley 14- by 22-Foot Subsonic Tunnel to measurethe flow near the empennage of a 15-percent scale powered helicopter modelwith an operating tail fan. Three-component velocity profiles were measuredforward of the horizontal tail for four advance ratios using laser velocimetry(LV) to evaluate the effect of the rotor wake impingement on the horizontal tailangle of attack. These velocity data indicate the horizontal tail can experienceunsteady angle of attack variations of over 30° due to the rotor wake influence.The horizontal tail is most affected by the rotor wake above advance ratios of0.10. Velocity measurements of the flow on the inlet side of the tail fan werealso made for a low-speed flight condition using LV techniques. The velocitydata show an accelerated flow near the tail fan duct, and vorticity calculationstrack the passage of main rotor wake vortices through the measurement plane.
Introduction
As rotor and fuselage designs become moreintegrated, compact, and complex, rotor-wake-fuselage aerodynamic interactions are anincreasingly important part of the overallperformance characteristics of rotorcraft.Reference 1 attributes the importance ofinteractional effects for modern helicopters toincreased disk loading, more compact designs,low-level flight requirements, and the increasedrequirement for directional trim after the loss ofthe tail rotor that results in larger vertical tailsurfaces. These effects are especially importantin the design and placement of the anti-torquesystem, such as a tail rotor, and the horizontaland vertical stabilizers as documented inreferences 2-3.
Much work has already been doneexperimentally and analytically to define theinteraction effects between the rotor and thefuselage (refs. 4-15). A more limited amount ofexperimental data is available for analyzing themain rotor/anti-torque interactions (refs. 16-22).
Advanced configurations such as the RAH-66are designed and manufactured withsophisticated and new anti-torque devices, andthere is a need for high-quality experimentaldata to support the development of more flexibleanalytical models capable of treating these typesof configurations (refs. 23 and 24). Reference25 specifically cites the difficulty in predictingunsteady empennage loads at speeds below 40knots. Reference 26 provides experimentalpressure data at model scale for a generic T-tailempennage, and reference 27 discusses thetremendous amount of testing involved in theLight Helicopter (LH) design process.However, there does not appear to be specificinformation in the literature on the velocities inthe flowfield of a lifting rotor near an operatingtail fan.
In order to investigate the rotor wake-fuselage-empennage interactions near theempennage of a powered small-scale helicopterwith an operating tail fan and a T-tail, the U. S.Army Joint Research Program Office,Aeroflightdynamics Directorate, in cooperationwith the NASA Langley Research Center,
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conducted a wind tunnel test program in the 14-by 22-Foot Subsonic Tunnel. Velocity datawere acquired forward of the horizontal tail forfour flight conditions to document the unsteadydownwash near the horizontal tail. Velocitydata were also obtained on the inlet side of thetail fan for one flight condition, providinginformation about the inflow into the tail fan.
Symbols and Abbreviations
CT Rotor thrust coefficient, T
ρπR2(ΩR)2
R Rotor radius, ft
T Rotor thrust, lb
u Streamwise component of velocity,ft/sec
U∞ Freestream velocity, ft/sec
vf Induced velocity in forward flight, ft/sec
vh Induced velocity in hover, ft/sec
v Lateral component of velocity, ft/sec
w Vertical component of velocity, ft/sec
x, y, z Cartesian coordinates (see fig. 3), in.
α Rotor shaft angle, positive nose up, deg
µ Main rotor advance ratio, U∞
ΩR
Χ Wake skew angle, deg,
tan−1 U∞cos α( )vf − U∞sin α( )
Ω Main rotor rotational speed, rad/sec
ρ Density of air, slugs/ft3
BL Baseline reference configuration:fuselage, tail fan covered with plates
FDP Frequency Domain Processor
LV Laser Velocimetry
MR Main rotor configuration: fuselage, tailfan free-wheeling, main rotor operating
MRTF Main rotor/tail fan configuration:fuselage, tail fan operating, main rotoroperating
OR Order Ratio, frequency spectrum ofvelocity signal in multiples of rotorfundamental frequency
TF Tail Fan configuration: fuselage, tailfan operating
2MRTS 2-Meter Rotor Test System
Model and Instrumentation
The test program was conducted in theLangley 14- by 22-Foot Subsonic Tunnel usingthe Army’s 2-Meter Rotor Test System(2MRTS) with a four-bladed, 15-percent scalerotor, a fuselage model representative of theRAH-66, and the tunnel’s three-component laservelocimetry (LV) system.
The 14- by 22-Foot Subsonic Tunnel is aclosed-circuit, atmospheric wind tunnel designedfor the low-speed testing of powered and high-lift configurations (ref. 28). In the open testsection configuration, the walls and ceiling arelifted out of the flow, leaving a solid floor underthe model. In this configuration, the tunnel canachieve a maximum dynamic pressure of about92 lb/ft2. This investigation was conducted withthe tunnel in the open test section configurationto allow complete optical access to the rotorflowfield. For this test program, the test section
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floor was lowered two feet to install the LVoptics. A false floor with a window, flush withthe rest of the tunnel, was placed over the LVoptics.
Figure 1 shows the 2MRTS ready for testingin the tunnel. The LV system is also visible inthe photograph. The rotor system, which wasinstalled on the 2MRTS, had a 4-bladed,articulated hub with blades that closely matchedthe planform, twist, and airfoils of the RAH-66blades. No attempt was made to dynamicallyscale these blades. Because the only hubavailable for testing was a 4-bladed hub, therewere some deviations from scale from an actualmodel of the 5-bladed RAH-66. The radius ofthe blades when installed on the 4-bladed hubwas reduced by one inch from a true 15-percentscale RAH-66. In addition, the use of only fourblades reduced the rotor solidity and resulted inhigher blade loads for any given thrustcoefficient. The blades and hub are described inmore detail in table 1 and a planform sketch ofthe blades is shown in figure 2. The 2MRTS isdescribed in further detail in reference 29.
The fuselage was a 15-percent scale model ofthe RAH-66 and was instrumented with over200 surface pressure ports and 4 dynamicpressure gages. Forces and moments on therotor and fuselage were measured separately bytwo six-component, strain-gage balances. Thefuselage is shown in detail in figure 3. Fuselagesurface pressure data were acquired during thistest program, and samples of the pressure dataare used in reference 30 for computational fluiddynamics (CFD) code calibration.
The anti-torque device of the configurationwas modeled by an air-powered, tip-driven, 8-in.diameter, 22-bladed fan mounted in the tail fanduct. The fan configuration is shown in figure 4.As can be seen in the photograph, the fan ductsection was painted black to minimize theoptical reflections from the surface. Althoughthe fan configuration did not match the physical8-bladed tail fan assembly of the full-scaleconfiguration, the model fan was used to
simulate the general flow physics and trimconditions for the model by using the fan rpm tocontrol the fan thrust. For this study, a generalsimulation of the thrusting fan environment wasthought to be sufficient to yield informationregarding the flowfield around the fuselage.Obviously, a detailed assessment of the full-scale fan design could not be made using thistype of simulation.
Laser Velocimeter System
The LV system was a three-componentsystem operated in the backscatter mode tominimize alignment difficulties between thetransmit and receive optics packages. Mostcomponents of the system are described inreferences 31-32; this paper presents the firstdata obtained with the upgraded three-component system. The streamwise and verticalcomponents of velocity are measured by opticslocated on the side of the tunnel, out of the flow;the lateral crossflow component of velocity ismeasured by optics which are located beneaththe tunnel floor. The traversing mechanisms ofthe three components are computer-controlled toensure the sample volumes of the three sets ofbeams are positioned at a single location. Ascan be seen in figure 1, the third componentbeams originating beneath the floor were angledat 33° relative to the vertical. This angle wasnecessary to optically access the inflow area ofthe tail fan due to the cant of the tail fan duct.Corrections for this rotation in the lateralvelocity component were applied to the dataduring post-processing.
Except for its long focal length and zoomlens assembly, the system was a standard fringe-based LV system. Polystyrene particles (1.7micron) suspended in an alcohol and watermixture were used to seed the flow. Thevelocity data were acquired using frequencydomain processors (FDP’s) to maximize thesignal to noise ratio of the data. The LV dataacquisition system was designed to allowacquisition of rotor azimuth position in additionto the velocity measurements so that an
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“azimuthal history” of the velocity could bereconstructed in post-processing.
Measurement Locations and TestProcedures
The laser velocimeter measurement locationsare described briefly below and are shown infigure 5. The operating conditions for eachconfiguration are documented in table 2.
Horizontal Tail
LV data were obtained for ten points in avertical line one chord forward and one chord(mid-semi-span) to the right of center of thehorizontal tail with both the main rotor and tailfan operating (MRTF) for main rotor advanceratios of 0.055, 0.076, 0.102, and 0.150. Therotor thrust coefficient was 0.007, and the rotorshaft angle was held at a constant -0.65°.
Tail Fan
In order to investigate non-linear interferenceeffects between the main rotor wake and the tailfan, velocity data were acquired on the inlet sideof the tail fan for several combinations ofunpowered and powered main rotor and tail fanconditions.
Baseline Data were acquired for a baselinereference condition (BL), which consisted ofonly the fuselage (no main rotor installed) andthe tail fan covered with plates to prevent flowthrough the tail duct. This established thereference flow conditions at the measurementplane due to just the presence of the fuselage inthe freestream. The tunnel speed was 55 ft/sec,which was the speed for a main rotor advanceratio of 0.076 if the main rotor had beeninstalled and operating. The velocity wasmeasured with LV in three locations near thecovered tail fan. At each location, the velocitywas very close to the freestream value,indicating little interference due to the fuselagealone at these locations.
TF For the tail fan operating alonecondition (TF), the main rotor was not installed,and the tail fan was operated at an rpm whichwas known to generate about 340 in-lbs of anti-torque. This was the amount of anti-torque thatwas predicted before the test program to berequired to trim the configuration in yaw. Thetunnel speed was again set to 55 ft/sec.Measurements of velocity with the LV weremade at a limited number of locations asreflections from the tail fan spinner and the ductmade the measurements difficult to acquire withthe LV system operating in the backscattermode.
MR LV measurements were obtained inthe same measurement plane for the main rotoroperating alone condition (MR). Thisconfiguration had the main rotor installed andoperating, and the tail fan was uncovered andunpowered, but it did free-wheel during the testcondition. The tail fan was uncovered tominimize the optical reflections from the cover.As a result, the tail fan was free-wheeling at anominal rpm of about 150. This wasapproximately 3% of the operating rpm for thefan; however, during static thrust sweeps of thefan, it required 1200 rpm to generate 0.3 lbs offan thrust. Therefore, the free-wheeling fan rpmof 150 was considered too low in magnitude togenerate any significant thrust or create anymeasurable flow effects. In addition, thefuselage yawing moment did not vary greatlybetween this configuration and the BLconfiguration, so it was not expected that thischange from covered fan to free-wheeling fanwould appreciably affect the data comparisons.
The operating conditions were a main rotoradvance ratio of 0.076, a thrust coefficient of0.0051, and a shaft angle of -0.60°. The mainrotor was trimmed to zero longitudinal andlateral first harmonic flapping. Due to therelatively small size of the model and the lowthrust coefficient, the calculated wind tunnelwall effects corrections were insignificant, andno wall corrections were applied to these data.Measurements of the rotor torque averaged 330
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in-lbs, indicating the TF anti-torque setting wasvery close to that required for trimming thesystem for the MR configuration.Measurements of velocity were made with LV atmore locations near the tail fan for thisconfiguration as it was assumed that the rotorwake would cause more non-uniformdistribution of velocity than either the BL or theTF configuration.
MRTF The final configuration was forboth the main rotor and tail fan operating(MRTF). In this configuration, the main rotorwas trimmed in the same manner as for the MRconfiguration with the advance ratio = 0.076, thethrust coefficient = 0.0051, and the shaft angle =-0.63°. The tail fan was operated at the samerpm and pressure as for the TF configuration.However, although the fan rpm wasapproximately the same as for the TFconfiguration, the anti-torque produced was 640in-pounds, almost twice as much as needed totrim the model to zero yawing moment. Thisincreased performance by the tail fan may bedue to the favorable interference effects betweenthe rotor wake and the tail fan. This favorableinterference resulted in a model testconfiguration that was out of trim in yaw whencompared to a flight test condition.Unfortunately, there was not enough testing timeto acquire a second MRTF test condition forbetter trim matching.
It was noted during the test program that forthis operating condition, the model wouldoccasionally shake or “twitch” in yaw; thisobserved phenomenon indicated the fan mighthave been experiencing some type of inlet stallphenomenon. The main concentration of LVmeasurements was for this MRTF configuration.
LV Data Acquisition and Reduction
The LV data acquisition process consisted ofplacing the sample volume at the measurementlocation and acquiring data for a period of nine
minutes or until 4096 velocity measurementswere made in each of the longitudinal, vertical,and lateral components of velocity. The LVmeasurements were not made in coincidence,which would have required that each componentof velocity be measured at the same time fromthe same particle. Instead, the flow wasassumed to be periodic with rotor blade passage,and each component was allowed to bemeasured individually; this dramaticallyreduced the time required to obtain the LV data.During this process, as was mentioned earlier,conditional sampling techniques were employedto associate each measured velocity with theazimuth of the rotor blades at the time when themeasurement was made. At the conclusion ofthe process, the measurement location waschanged, and the acquisition process wasrepeated.
For each measurement location, the raw datawere reviewed, and the histograms of thevelocities in each of the three components wereprocessed to improve the signal-to-noise ratio.The data were “binned” into 128 bins (2.8°azimuth each), and the mean velocity for thelocation was calculated from the mean of all theazimuth bins. Since the data were associatedwith a rotor position, it was possible to sort thedata by azimuthal position, therebyreconstructing a time history of velocity at eachmeasurement location that represented oneaverage rotor revolution.
The largest contributors to the uncertainty inthe LV measurements are the measurement ofthe crossbeam angle and the particle lag. Usingthe error estimation techniques described inreferences 32-34, the LV system error for thevelocity measurements in this paper is estimatedat 1.3% measured velocity.
Discussion of Results
Horizontal Tail Velocity Measurements
The average downwash angle, as measured
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using LV at a location one chord forward of thehorizontal tail, is shown in figure 6 for severaladvance ratios. In each of these cases, both themain rotor and the tail fan are operating at theconditions indicated in table 2. As expected, thedownwash angle decreases with increasingadvance ratio. Similarly, the average sidewashangle is shown in figure 7. The sidewash angledecreases with increasing advance ratio; that is,the average lateral flow tends more to thestarboard side of the model with increasingforward speed.
Large variations in the unsteady downwashand sidewash angles were also measured usingthe LV system. Typical plots of the unsteadyflow angles calculated from the unsteadyvelocity data are shown in figure 8 for a heightone-half inch below the horizontal tail for eachof the advance ratios tested. The results indicateover 30° of unsteady fluctuation are encounterednear the horizontal tail at the blade passagefrequency with the most unsteadiness occurringat an advance ratio of 0.10. Carpet plots of theunsteady angles for all the advance ratios thatwere tested are presented in figure 9. Theseplots show the variation in unsteady angle withheight above the tail section at each advanceratio.
From the unsteady data, an experimentaldetermination of the position of the rotor wakerelative to the horizontal tail can be made byanalyzing the 4/rev content of the velocity. Thevertical component of velocity was used tocalculate the 4/rev RMS content of the rotorwake, and the results are shown in figure 10.The strong 4/rev content indicates that the rotorwake is the dominant flow feature. Figure 10also shows the position of the rotor wakerelative to the horizontal tail for the advanceratios tested. In figure 10, the tail is immersedin the wake above advance ratios of 0.10 asshown by the high 4/rev content at allmeasurement locations for these advance ratios.These results generally agree with those inreference 26, considering the different geometryand flight conditions of the two test
configurations.
A theoretical evaluation of the wake skewangle can be made based on the equations inreference 35. Generally accepted practice inapplying the equations of reference 35 tocalculate wake skew angle is to use the hoverinduced velocity, vh, in the equation for wakeskew angle. Since vh is defined as:
vh = T
2ρπR2
this is a fairly straightforward calculation.However, in determining the wake skew anglefor a forward flight configuration, the inducedvelocity in forward flight, vf, can be shown to bea function of thrust, tip speed, and forwardspeed. By solving the equation and assumingthat the rotor shaft angle is small, vf is given by:
vf = 1
2CT
2 ΩR( )4 + U∞4[ ]−U∞
2
The skew angle is then calculated by:
Χ = tan−1 U∞ cosα( )vf −U∞ sin α( )
This formulation for Χ using the forward-flight induced velocity results in a significantlydifferent skew angle calculation than if the hovervalue of induced velocity is used. The results ofthe skew angle calculation are plotted in figure11 along with the geometry of the configuration.In figure 11, the calculated wake skew angle,which theoretically defines the edge of the rotorwake, is shown to approach the position of thehorizontal tail at advance ratios of 0.10 andabove. This correlates well with theexperimental determination of rotor wakeposition shown in figure 10.
Figure 12 illustrates an interesting feature ofthe unsteady vertical flow near the horizontaltail. At the lower advance ratios of the test
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program, represented by the sample plot infigure 12a, the flow is dominated by the 4/revblade-passage frequency and its multiples. Atan advance ratio of 0.15, a significant 2/revcontent becomes present in the flow as shown infigure 12b. Reference 26 also reports a strong8/rev in the flow near the horizontal tail; the datain the present investigation show periodiccontent at several multiples of the 4/revfrequency, as well as frequencies between themultiples of the 4/rev. During the inflow studiesof references 36-38, 2/rev frequencies werenoted in several instances; it is possible thesefrequencies are generated by vorticity shed fromthe hub and pylon that moves into themeasurement area at the increased advanceratios.
Tail Fan Velocity Measurements
Figure 13 presents contour plots in the fansystem coordinates of the average streamwisevelocity, u, the lateral (perpendicular to the fan)velocity, v, and the vertical (parallel to the fan)velocity, w, on the inlet side of the duct. Theseare presented for the MRTF configurationoperating at the conditions listed in table 2.Note the accelerated flow at the forward sectionof the duct. The photographs in figure 14 showsurface flow visualization of the empennage forone of the runs. The flow visualization,supported by the velocity data, indicate the flowis separated along the upper half of the upstreamlip of the tail fan duct. The photographs alsoshow a large region of separation on the aft partof the tail fan shroud. There are also severalseparation lines on the vertical tail and thejunction between the vertical tail and the tail fanduct. This occurs on both the right and left handside of the empennage.
The purpose of acquiring data for severaldifferent model configurations was to allow thedetermination of the non-linear interferenceeffects between the main rotor wake and the tailfan flow. This was determined by subtractingthe combination of the MR and TF velocitiesfrom the MRTF velocities. The results for four
locations in the measurement plane are given intable 3. There are limited results for this part ofthe investigation due to the small number ofmeasurements made for the TF configuration.From a percentage standpoint, the non-lineareffects are most significant in the lateral velocity(v) component, which is influenced the most bythe tail fan flow.
For each measurement location, the unsteady,azimuthally-dependent velocity was measuredby LV in each velocity component. For a givenazimuth, the velocity at each measurement pointcan be extracted and plotted on a contour plot togive an effective velocity “snapshot” of theentire measurement grid. As these data wereprocessed at azimuth intervals of 2.8°, therewere 128 snapshots of velocity in each of thethree components.
From each snapshot of velocity, the vorticitycomponent normal to the measurement gridplane was calculated for the MRTFconfiguration. By examining each azimuthal“snapshot”, it became evident that areas ofconcentrated vorticity were convecting throughthe measurement plane. Figure 15 showsexamples of two azimuth angles. Both positive(into the plane) and negative vorticity arepresent in the plots. The phenomenon occurs 4times per revolution as would be expected due tothe 4-bladed rotor configuration. This periodiccontent in the data indicates that the main rotorblade wake vortices are passing through themeasurement plane at the blade passagefrequency. The rotor wake contains bothpositive and negative vorticity due to thegeometric structure of the blade tip vortices.Whether the vortices are seen as positive ornegative is dependent on the wake age and thevortex origination point. Wake flowvisualization of this phenomenon can be foundin reference 39.
The convection velocities, on average, werecalculated to be 45 ft/sec in the downstreamdirection and 52 ft/sec in the vertical direction.This equates to an experimental skew angle of
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41°. Using the equations in the above sectionfor skew angle, the calculated momentum skewangle is determined to be 62°. Although thecalculation and the data have a 21° difference inskew angle, this can be explained by thedifference in the inflow velocity data betweenmomentum theory and experimentalmeasurement. It is known from references 36-40 that the experimental inflow for this type ofrotor contains more downwash on the aft portionof the disk than is predicted by momentumtheory. Therefore, it is expected that theexperimental skew angle would be less inmagnitude than the skew angle calculated frommomentum theory.
Conclusions
In order to investigate the rotor wake-fuselage-empennage interactions near theempennage of a powered small-scale helicopterwith an operating tail fan and a T-tail, the U. S.Army Joint Research Program Office,Aeroflightdynamics Directorate, in cooperationwith the NASA Langley Research Center,conducted a wind tunnel test program in the 14-by 22-Foot Subsonic Tunnel. Velocity datawere acquired forward of the horizontal tail forfour flight conditions, documenting the unsteadydownwash near the horizontal tail. Velocitydata were also obtained on the inlet side of thetail fan for one flight condition, providinginformation about the inflow into the tail fan.The major conclusions from this study are:
1. The horizontal tail surface experiences largechanges (over 30°) in the unsteady sidewash anddownwash angles due to the influence of therotor wake. The horizontal tail is most affectedby the rotor wake above advance ratios of 0.10.
2. The wake skew angle, calculated throughsolving the momentum equation for inducedvelocity, can be significantly different than theexperimental value.
3. There is an accelerated flow pattern near theoperating tail fan. Flow visualization, as well as
the measured velocity data, indicate the flow isseparated on part of the forward duct lip and atthe base of the vertical tail for an advance ratioof 0.07 and a main rotor thrust coefficient of0.005.
4. Velocity measurements show the passage ofvorticity at the inlet to the tail fan. Both positiveand negative vorticity are measured in theflowfield, and the vorticity is convected at anangle and rate consistent with passage of therotor wake through the tail fan inlet area.
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24. Philippe, J.: ONERA Makes Progress In RotorAerodynamics, Aeroelasticity, and Acoustics.Vertiflite, Sept/Oct. 1992, pp. 48-53.
25. Torok, M. S.; and Ream, D. T.: Investigation ofEmpennage Airloads Induced by a HelicopterMain Rotor Wake. Proceedings of the 49thAnnual Forum, American Helicopter Soc., 1993.
26. Moedersheim, E.; and Leishman, J. G.:Investigation of Aerodynamic InteractionsBetween a Rotor and a T-Tail Empennage.Proceedings of the Aeromechanics SpecialistMeeting, American Helicopter Soc., 1995.
27. Keys, C.; Sheffler, M.; Weiner, S.; andHeminway, R.: LH Wind Tunnel Testing: Key
10
to Advanced Aerodynamic Design. Proceedingsof the 47th Annual Forum, American HelicopterSoc., 1991.
28. Gentry, G. L.; Quinto, P. F.; Gatlin, G. M.; andApplin, Z. T.: The Langley 14- by 22-FootSubsonic Tunnel: Description, FlowCharacteristics, And Guide For Users. NASATP 3008, 1990.
29. Phelps, A. E.; and, Berry, J. D.: Description ofthe Army’s 2-Meter Rotor Test System. NASATM-87762, AVSCOM TM-86-B-4, 1987.
30. Duque, E. P. N.; Berry, J. D.; Budge, A. M.; andDimanlig, A. C. B.: A Comparison of Computedand Experimental Flowfields of the RAH-66Helicopter. Proceedings of the AeromechanicsSpecialist Meeting, American Helicopter Soc.,1995.
31. Mace, W. D. Jr.; Elliott, J. W.; Blancha, B.; andMurphy, J.: Comparison Of Frequency DomainAnd Time Domain Laser Velocimeter SignalProcessors. 14th International Congress onInstrumentation in Aerospace SimulationFacilities, 1991.
32. Gorton, S. A.; Poling, D. R.; and Dadone, L.:Investigation Of Blade-Vortex Interaction UsingLaser Velocimetry And Pressure-InstrumentedRotor Blades. Volume 1—Advance Ratio of 0.2,Rotor Lift Coefficient Normalized by Solidity of0.07, and Shaft Angle of 0 Degrees. NASA TM4570, ATCOM-TR-94-A-003, 1995.
33. Young, W. H., Jr.; Meyers, J. F.; and Hepner, T.E.: Laser Velocimeter Systems Analysis Appliedto a Flow Survey Above a Stalled Wing. NASATN D-8408, 1977.
34. Dring, R. P.: Sizing Criteria for LaserAnemometry Particles. J. Fluids Eng., vol 104,Mar. 1982, pp. 15-17.
35. Stepniewski, W. Z.; and Keys, C. N.: Rotary-Wing Aerodynamics. Dover Publications, Inc.,New York, 1984, p. 64.
36. Elliott, J. W.; Althoff, S. L.; and Sailey, R. H.:Inflow Measurements Made with a LaserVelocimeter on a Helicopter Model in ForwardFlight— Volume I: Rectangular Planform at anAdvance Ratio of 0.15. NASA TM 100541,AVSCOM TM 88-B-004, 1988.
37. Elliott, J. W.; Althoff, S. L.; and Sailey, R. H.:Inflow Measurements Made with a LaserVelocimeter on a Helicopter Model in ForwardFlight—Volume II: Rectangular Planform at anAdvance Ratio of 0.23. NASA TM 100542,AVSCOM TM 88-B-005, 1988.
38. Elliott, J. W.; Althoff, S. L.; and Sailey, R. H.:Inflow Measurements Made with a LaserVelocimeter on a Helicopter Model in ForwardFlight—Volume III: Rectangular Planform at anAdvance Ratio of 0.30. NASA TM 100543,AVSCOM TM 88-B-006, 1988.
39. Ghee, T. A.; Berry, J. D.; and Zori, L. A. J.:Wake Geometry Measurements and AnalyticalCalculations on a Small-Scale Rotor Model.NASA TP-3584, 1996.
40. Hoad, D. R.: Rotor Induced-Inflow-RatioMeasurements and CAMRAD Calculations.NASA TP-2946, AVSCOM-TM-89-B-010, 1990.
11
Tables
Table 1. Description of rotor blades.
Airfoil sections 23.7-percent radius . . . . . . . . . . . . . . . . . . . . VR-12 84.6-percent radius . . . . . . . . . . . . . . . . . . . . VR-12 91.8-percent radius . . . . . . . . . . . . . . . . . . SSC-A09 100-percent radius . . . . . . . . . . . . . . . . . . . SSC-A09Chord, in. 23.7-percent radius . . . . . . . . . . . . . . . . . . . . . . 2.25 74.3-percent radius . . . . . . . . . . . . . . . . . . . . . . 2.25 91.8-percent radius . . . . . . . . . . . . . . . . . . . . . . 2.25 100-percent radius . . . . . . . . . . . . . . . . . . . . . . . 1.35Cutout, in. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8.2Flapping hinge offset, in. . . . . . . . . . . . . . . . . . . . . 2.0Lag hinge offset, in. . . . . . . . . . . . . . . . . . . . . . . . . 2.0
Number of blades . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4Pitch axis, percent of chord . . . . . . . . . . . . . . . . . . . 25Radius, in. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 34.55Solidity, thrust-weighted . . . . . . . . . . . . . . . . . 0.07866Tip sweep angle (of 1/4 chord), deg . . . . . . . . . . . . . 30Tip sweep begins, in. . . . . . . . . . . . . . . . . . . . . . . . 31.7Twist, deg 0-percent radius . . . . . . . . . . . . . . . . . . . . . . . . . . . 0 23.7-percent radius . . . . . . . . . . . . . . . . . . . . . . . . . 0 74.3-percent radius . . . . . . . . . . . . . . . . . . . . . . . -6.6 84.6-percent radius . . . . . . . . . . . . . . . . . . . . . . . -7.6 91.8-percent radius . . . . . . . . . . . . . . . . . . . . . . . -9.5 100-percent radius . . . . . . . . . . . . . . . . . . . . . . . -9.5
Table 2. Test Conditions.
Variable BL TF MR MRTF Horizontal TailAdvance ratio . . . . . . . . . — — 0.076 0.076 0.055 0.076 0.102 0.15Collective, deg . . . . . . . . — — 7.1 7.4 11.0 10.1 8.9 7.5Density, slug/ft3 . . . . . . . . .00249 .00243 .00242 .00241 .00236 .00237 .00236 .00235Fuselage angle of attack, deg . 4.3 4.4 4.3 4.3 4.1 4.0 4.1 4.1Freestream velocity, ft/sec . . 54.9 55.0 55.2 55.2 40.0 54.7 73.8 108.9Freestream velocity, knots . . 32.6 32.6 32.7 32.7 23.7 32.4 43.7 64.5Fuselage yaw moment, in-lb . -25.7 -342.8 -75.0 -641.9 -473.4 -631.3 -815.5 -781.4Lateral cyclic, deg . . . . . . — — 1.1 1.1 1.2 1.3 1.9 2.5Longitudinal cyclic, deg . . . — — -3.0 -3.0 -2.8 -3.0 -3.2 -2.7Rotor drag, lb . . . . . . . . — — 2.1 1.7 0.6 2.8 2.6 3.3Rotor lift, lbs . . . . . . . . . — — 170.0 169.0 230.5 228.9 229.4 228.8Rotor rpm . . . . . . . . . . — — 2401 2400 2400 2399 2402 2401Rotor shaft angle, deg . . . . — — -0.60 -0.63 -0.66 -0.61 -0.61 -0.61Rotor thrust coefficient . . . . — — .00512 .00512 .00714 .00706 .00707 .00709Rotor yawing moment, in-lb . — — 330.3 324.4 540.3 497.2 421.8 310.0Tail fan rpm . . . . . . . . . — 5007 — 5394 4860 5197 6262 5435
12
Table 3. Non-linear Interference Velocities.
Location (measured fromcenter of fan) MRTF - (MR + TF), ft/sec Velocity, % MRTF, ft/sec
u v w u v w
1 inch upstream, 0.7 inch up -3.5 2.9 -9.1 -4.0 -18.8 11.6
1 inch downstream, 0.7 inch up -6.7 0.9 -7.8 -8.3 -4.6 10.0
1 inch upstream, 2.7 inch up -3.9 6.9 -6.2 -4.7 -34.8 7.5
1 inch downstream, 2.7 inch up -6.9 5.2 -15.9 -9.3 -23.3 19.9
13
L-95-1069
Figure 1. Model and LV system installed in tunnel.
34.55
VR-12
LinearTransition SSC-A09
30 deg
1.35
Center of rotation
2.25
Airfoil Sections
Figure 2. Description of blade planform. All dimensions in inches.
14
69.10
34.55
17.95
6.70
12.05
28.65
69.15
78.00
8.00
7.65
20.05
z
y
z
x
Figure 3. Description of fuselage. All dimensions in inches.
L-95-1071
Figure 4. Tail fan configuration.
15
ZX
Y
5
4
U¥
w
u
(a) Side view of horizontal tail measurementlocations.
ZX
Y
U¥
(c) Side view of fan measurementlocations.
w
v
(b) Rear view of horizontal tailmeasurement locations.
(d) Rear view of fan measurementlocations.
Figure 5. Velocity measurement locations. All dimensions in inches.
16
U¥ = 40 fps
(a) µ = 0.05. Average downwash angle is57 degrees.
U¥ = 74 fps
(c) µ = 0.10. Average downwash angle is35 degrees.
U¥ = 55 fps
(b) µ = 0.07. Average downwash angle is47 degrees.
U¥ = 109 fps
(d) µ = 0.15. Average downwash angle is23 degrees.
Figure 6. Average downwash angle forward of the horizontal tail for MRTF.
17
Reference Vector = 40 fps
(a) µ = 0.05. Average sidewash angle is 11degrees.
Reference Vector = 74 fps
(c) µ = 0.10. Average sidewash angle is 4degrees.
Reference Vector = 55 fps
(b) µ = 0.07. Average sidewash angle is 7degrees.
Reference Vector = 109 fps
(d) µ = 0.15. Average sidewash angle is -3degrees.
Figure 7. Average sidewash angle forward of the horizontal tail for MRTF.
18
0 60 120 180 240 300 3600
10
20
30
40
50
60
70
Azimuth, deg
m = 0.05
m = 0.07
m = 0.10
m = 0.15
Do
wn
wash
,d
eg
(a) Downwash angle.
0 60 120 180 240 300 360-20
-10
0
10
20
30
40
50
Azimuth, deg
m = 0.05
m = 0.10
m = 0.15
m = 0.07
Sid
ew
ash
,d
eg
(b) Sidewash angle.
Figure 8. Unsteady angles for a location 0.5 inches below the horizontal tail.
19
090
180270
360
20
30
40
50
60
70
-2.5
-1.5
-0.5
0.5
1.5
2.5
Height
Azimuth, deg
Downw
ash
angle
, deg
090
180270
360
20
10
0
10
20
30
-2.5
-1.5
-0.5
0.5
1.5
2.5
Height
Azimuth, deg
Sidewas
h an
gle, d
eg
(a) µ = 0.05.
090
180270
360
20
30
40
50
60
70
-2.5
-1.5
-0.5
0.5
1.5
2.5
Height
Azimuth, deg
Downw
ash
angle
, deg
090
180270
360
20
10
0
10
20
30
-2.5
-1.5
-0.5
0.5
1.5
2.5
Height
Azimuth, deg
Sidewas
h an
gle, d
eg
(b) µ = 0.07.
Figure 9. Unsteady downwash and sidewash angles for MRTF.
20
090
180270
360
20
30
40
50
60
70
-2.5
-1.5
-0.5
0.5
1.5
2.5
Height
Azimuth, deg
Downw
ash
angle
, deg
090
180270
360
20
10
0
10
20
30
-2.5
-1.5
-0.5
0.5
1.5
2.5
Height
Azimuth, deg
Sidewas
h an
gle, d
eg
(c) µ = 0.10.
0
90180
270360
0
10
20
30
40
50
-2.5
-1.5
-0.5
0.5
1.5
2.5
Height
Azimuth, deg
Downw
ash
angle
, deg
090
180270
360
20
10
0
10
20
30
-2.5
-1.5
-0.5
0.5
1.5
2.5
Height
Azimuth, deg
Sidewas
h an
gle, d
eg
(d) µ = 0.15.
Figure 9. Concluded.
21
Figure 10. Unsteady vertical wake impingement at horizontal tail for MRTF.
Figure 11. Wake skew angle calculation for CT = 0.007 projected on model geometry.
22
0 60 120 180 240 300 360-40
-20
0
20
40
Azimuth, deg
w,fp
s
(a) µ = 0.10.
0 60 120 180 240 300 360-40
-20
0
20
40
Azimuth, deg
w,fp
s
0 4 8 12 16 20 240
3
6
9
Order Ratio
RM
So
fw
,fp
s
2/rev
4/rev
(b) µ = 0.15.
Figure 12. Velocity and order ratio analysis for location 2 inches below centerlineof horizontal tail, MRTF.
23
5
9
7
875
8
5
6
ZX
Y
U¥ = 55 ft/sec
A 95
9 90
8 85
7 80
6 75
5 70
4 65
3 60
2 55
1 50
u, fps
(a) Streamwise velocity, u.
1
13
31
1
36
56
34
4
4
356
2
ZX
Y
U¥ = 55 ft/sec
A -54
9 -57
8 -60
7 -63
6 -66
5 -69
4 -72
3 -75
2 -78
1 -81
w, fps
(b) Vertical velocity, w.
31
554
89A
2
657
2
56
2
ZX
Y
U¥ = 55 ft/sec
A 7
9 4
8 1
7 -2
6 -5
5 -8
4 -11
3 -14
2 -17
1 -20
v, fps
(c) Lateral velocity (inflow), v.
Figure 13. Contour plots of LV average velocity data for MRTF configuration.
24
(a) Inlet side of tail fan.
(b) Outflow side of tail fan.
Figure 14. Surface flow visualization of empennage for µ = 0.07 and thrust coefficient of 0.005, MRTFconfiguration.
25
-202-2
-1
0
1
2
3
4
5
5
5
5
6
6
6
6
6
6
6
6
7
7
7
7
7
7
8
9
Inch
es
up
fro
mta
ilfa
ncen
terl
ine
Inches downstream from tail fan centerlineForwardAft
Level
B 200
A 160
9 120
8 80
7 40
6 0
5 -40
4 -80
3 -120
2 -160
1 -200
Vorticity intoplane, sec
-1LV measurement locations
(a) Azimuth = 150 degrees.
-202-2
-1
0
1
2
3
4
5
2
3
3
4
4
4
5
5
5
5
5
6
66
6
6
6
6
6
7
7
7
7
7
7
8
88
8
8
Inch
es
up
fro
mta
ilfa
ncen
terl
ine
Inches downstream from tail fan centerlineForwardAft
Level
B 200
A 160
9 120
8 80
7 40
6 0
5 -40
4 -80
3 -120
2 -160
1 -200
-
6 6 6 6 6
LV measurement locationsVorticity intoplane, sec
-1
(b) Azimuth = 220 degrees.
Figure 15. Vorticity contours calculated from unsteady LV data for MRTF configuration.
REPORT DOCUMENTATION PAGE Form ApprovedOMB No. 0704-0188
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1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE
March 20003. REPORT TYPE AND DATES COVERED
Technical Publication4. TITLE AND SUBTITLE
Flow Environment Study Near the Empennage of a 15-Percent ScaleHelicopter Model
5. FUNDING NUMBERS
WU 581-10-11-01
6. AUTHOR(S)
Susan Althoff Gorton, John D. Berry, W. Todd Hodges, Deane G. Reis
7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)NASA Langley Research Center U.S. Army Aviation and Missile CommandHampton, VA 23681-2199 Aeroflightdynamics Directorate
Joint Research Programs Office NASA Langley Research Center Hampton, VA 23681-2199
8. PERFORMING ORGANIZATIONREPORT NUMBER
L-17940
9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)
National Aeronautics and Space AdministrationWashington, DC 20546-0001
and U.S. Army Aviation and Missile Command Moffett Field, CA 94035-1000
10. SPONSORING/MONITORINGAGENCY REPORT NUMBER
NASA/TP-2000-210085AFDD/TR-00-A-004
11. SUPPLEMENTARY NOTES
Berry: U.S. Army Aviation and Missile Command, Directorate of Aviation Engineering, Redstone Arsenal,Huntsville, AL 35898
12a. DISTRIBUTION/AVAILABILITY STATEMENT
Unclassified-UnlimitedSubject Category 03 Distribution: StandardAvailability: NASA CASI (301) 621-0390
12b. DISTRIBUTION CODE
13. ABSTRACT (Maximum 200 words)
Development of advanced rotorcraft configurations has highlighted a need for high-quality experimental data tosupport the development of flexible and accurate analytical design tools. To provide this type of data, a testprogram was conducted in the Langley 14- by 22-Foot Subsonic Tunnel to measure the flow near the empennageof a 15-percent scale powered helicopter model with an operating tail fan. Three-component velocity profileswere measured with laser velocimetry (LV) one chord forward of the horizontal tail for four advance ratios toevaluate the effect of the rotor wake impingement on the horizontal tail angle of attack. These velocity dataindicate the horizontal tail can experience unsteady angle of attack variations of over 30 degrees due to the rotorwake influence. The horizontal tail is most affected by the rotor wake above advance ratios of 0.10. Velocitymeasurements of the flow on the inlet side of the tail fan were made for a low-speed flight condition usingconventional LV techniques. The velocity data show an accelerated flow near the tail fan duct, and vorticitycalculations track the passage of main rotor wake vortices through the measurement plane.
14. SUBJECT TERMS
helicopter, rotor, wake, laser velocimetry, empennage, tail fan, fenestron,15. NUMBER OF PAGES
30horizontal tail 16. PRICE CODE
A0317. SECURITY CLASSIFICATION
OF REPORT
Unclassified
18. SECURITY CLASSIFICATIONOF THIS PAGE
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