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Overview of the Testing of a Small-Scale Proprotor 


Larry A. Young 

NASA Ames Research Center 

Moffett Field, CA 


Gloria K. Yamauchi 
NASA Ames Research Center 
Moffett Field, CA 


Earl R. Booth, Jr. 

NASA Langley Research Center 
Hampton, VA 


Gavin Botha 

NASA Ames Research Center 
Moffett Field, CA 


Seth Dawson 

The Boeing Company 

Mesa, AZ 


Summary 

This paper presents an overview of results from the wind tunnel test of a 1/4-scale V-22 proprotor in the 
Duits-Nederlandse Windtunnel (DNW) in The Netherlands. The small-scale proprotor was tested on the 
isolated rotor configuration of the Tilt Rotor Aeroacoustic Model (TRAM). The test was conducted by a 
joint team from NASA Ames, NASA Langley, U.S. Army Aeroflightdynamics Directorate, and The 
Boeing Company. The objective of the test was to acquire a benchmark database for validating 
aeroacoustic analyses. Representative examples of airloads, acoustics, structural loads, and performance 
data are provided and discussed. 


Nomenclature 

C P Rotor power coefficient 

C T Rotor thrust coefficient 

FM Rotor hover figure of merit 

Mtip Rotor tip Mach number 

Presented at the American Helicopter Society 
55th Annual Forum, Montreal, Canada, May 25- 
27, 1999. Copyright © 1999 by the American 
Helicopter Society, Inc. All rights reserved. 


R Rotor radius 

x/R non-dimensional tunnel longitudinal 

coordinate, origin at hub, positive 
downstream 

y/R non-dimensional tunnel lateral 

coordinate, origin at hub, positive on 
rotor advancing side 

z/R non-dimensional tunnel vertical 

coordinate, origin at hub, positive up 

V Wind tunnel test section velocity 

a s Rotor shaft angle, deg, shaft vertical at 

zero degrees angle, positive aft 



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Overview of the Testing of a Small-Scale Proprotor 

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Army/NASA Rotorcraft Division,Army Aviation and Missile 

Command,Aeroflightdynamics Directorate (AMRDEC), Ames Research 
Center,Moffett Field,CA,94035 

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Presented at the American Helicopter Society 55th Annual Forum, Montreal, Canada, May 25-27,1999 

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Standard Form 298 (Rev. 8-98) 

Prescribed by ANSI Std Z39-18 








Advance ratio, V/QR 


r| Proprotor efficiency, C T |i/C P 

\\) Rotor azimuth, deg. 

Q Rotor rotational speed, rad/s 


Introduction 

The successful introduction of civil tiltrotor 
aircraft is dependent in part on identifying and 
reducing, or suppressing, the noise generation 
mechanisms of tiltrotor aircraft proprotors. To 
accomplish these goals, a series of wind tunnel 
tests with a new generation of tiltrotor models is 
required. The purpose of the Tilt Rotor 
Aeroacoustic Model (TRAM) experimental 
program is to provide data necessary to validate 
performance and aeroacoustic prediction 
methodologies and to investigate and 
demonstrate advanced civil tiltrotor technologies. 
The TRAM project is a key part of the NASA 
Short Haul (Civil Tiltrotor) (SH(CT)) program. 
The SH(CT) program is an element of the 
Aviation Systems Capacity Initiative within 
NASA. Reference 1 summarizes the goals and 
objectives and the overall scope of the SH(CT) 
program. 

The current scope of TRAM experimental 
investigations is focused on the following: 

1. Acquisition and documentation of a 

comprehensive isolated proprotor 

aeroacoustic database, including rotor 
airloads. 

2. Acquisition and documentation of a 

comprehensive full-span tiltrotor 

aeroacoustic database, including rotor 
airloads, to enable assessment of key 
interactional aerodynamic and aeroacoustic 
effects by correlating isolated rotor and full- 
span TRAM wind tunnel data sets with 

advanced analyses. 

3. An advanced technology demonstrator test 
platform for low-noise proprotors. 


The first wind tunnel test of the TRAM project 
was an isolated rotor test in the Duits- 
Nederlandse Windtunnel (DNW) in The 
Netherlands (Fig. 1). This isolated rotor test was 
the first comprehensive aeroacoustic test for a 
tiltrotor proprotor, including not only noise and 
performance data, but airload and wake 
measurements as well. The TRAM isolated rotor 
test stand was installed and tested in the DNW 
open-jet test section during two tunnel entries, in 
December 1997 and April-May 1998. 



Fig. 1. TRAM Isolated Rotor Configuration in 
the Duits-Nederlandse Windtunnel (DNW) in 
The Netherlands 


This paper provides an overview of the data 
acquired during the TRAM DNW test. Follow- 
on plans for the full-span (dual-rotor, complete 
airframe) TRAM will be briefly discussed. 


Description of Model and Wind Tunnel 
Facility 


A general description of the TRAM isolated 
rotor configuration is found in Reference 2. A 
description of the DNW and its rotary-wing test 
capability is found in Reference 3. The model 
proprotor tested on the TRAM isolated rotor test 
stand was a 1/4-scale (9.5 ft diameter) V-22 
rotor. The rotor was counterclockwise rotating 




(planform view over the rotor); i.e., it was a 
right-hand side or starboard rotor. The isolated 
proprotor was tested in the 8x6m open-jet test 
section of the DNW. The 1/4-scale V-22 
proprotor was tested at reduced tip speed of 0.63 
tip Mach number because of operational 
considerations (the nominal design tip speed of 
the V-22 Osprey aircraft is M t i p = 0.71). All 
airplane-mode proprotor data were acquired at 
0.59 tip Mach number (equivalent to that of the 
V-22 aircraft). 

The TRAM isolated rotor test stand was 
comprised of two major elements: the rotor and 
nacelle assembly and the motor mount assembly. 
The rotor and nacelle assembly was attached to 
the acoustically treated isolated rotor test stand at 
a mechanical pivot or ‘conversion axis’ (fig. 4). 
This conversion axis allowed the nacelle to be 
manually rotated (in between tunnel runs) in 5 
degree increments from airplane to helicopter 
modes. An electric motor provided power to the 
rotor via a super-critical driveshaft. Rotor shaft 
angle changes were accomplished in flight with 
the DNW sting, which automatically maintained 
the hub on tunnel centerline. All rotating data 
channels were amplified by a Nationaal Lucht-en 
Ruimtevaartlaboratorium (The Netherlands 
National Aerospace Laboratory, NLR) developed 
Rotating Amplifier System (RAS) to enhance 
transducer signal to noise ratios before entering 
the slipring (reference 4). 



Figure 4 — Isolated Rotor Configuration 
(Planform View With Nacelle Assembly 
Positioned in Airplane-mode) 


A more detailed summary of the characteristics 
of the TRAM isolated rotor test stand and the 
1/4-scale V-22 rotor is noted below: 


• TRAM Isolated Rotor Configuration 
Description (Fig. 4) 

Test stand is wind tunnel sting- 
mounted; compatible with DNW and 
National Full-scale Aerodynamics 
Complex (NFAC) stings 
Drive train designed for nominal 300 
HP and 18,000 RPM motor; employs 




one right-angle 1:1 gearbox and a 
11.3:1 gear reduction transmission 
Nacelle tilt/incidence angle (about the 
‘conversion axis’) is ground adjustable 
Six-component rotor balance and 
instrumented torque coupling (with 
primary and secondary measurements) 
300-ring slip ring and rotating amplifier 
system 

Three electromechanical actuators and a 
rise-and-fall swashplate; rotating and 
nonrotating scissor sub-assembly design 
allows full proprotor collective/cyclic 
ranges without changing hardware 
rotor control system and console 
designed to minimize re-rigging 
between different operating regimes 
Self-contained model utilities within 
nacelle and motor-mount assemblies 
Motor-mount acoustically treated with 
foam panels 

Nacelle assembly not acoustically 
treated but geometrically scaled for V- 
22 aircraft outer mold 
Model capable of being tested to full V- 
22 operating envelope 
Rotor shaft interface hardware designed 
to easily install and test advanced 
proprotors on TRAM test stands 
Isolated rotor configuration is hardware 
compatible with the full-span model; 
hardware is shared between the two test 
stands 


Rotor Characteristics and Instrumentation 
(Fig. 5) 

Gimballed rotor hub with constant 
velocity joint (spherical bearing and 
elastomeric torque links) 

Rotor hub is dynamically and 
kinematically similar to V-22 aircraft 
hubs 

1/4-scale V-22 rotor set with both 
strain-gauged and pressure- 
instrumented blades 
Rotor pressure-instrumentation consists 
of 150 transducers (three different types 
of Kulite transducers) distributed over 
two rotor blades 

High fidelity scaling with respect to the 
V-22 rotor for blade/airfoil contours 


First elastic modes (flapwise, 
chordwise, and torsional) of blades 
dynamically scaled to V-22 frequencies 
Both strain-gauged and pressure- 
instrumented blades have nominally 
identical mass distributions and CG 
locations 

Adequate instrumentation provided to 
acquire a good blade/hub structural load 
data set for analytical correlation. 



(a) 



(b) 

Fig. 5a-b TRAM 1/4-scale V-22 Rotor (a. Blade 
and b. Hub) 



Test Description 

Two tunnel entries were conducted in the DNW 
wind tunnel with the TRAM isolated rotor test 
stand and the 1/4-scale V-22 rotor. The first 
tunnel entry was in December 1997 and it was 
focused on TRAM test stand risk reduction and 
envelope expansion for the 1/4-scale V-22 rotor. 
The second entry in April-May 1998 was 
devoted to acquiring a high quality isolated rotor 
aeroacoustic database for the SH(CT) program. 

The German Dutch Wind Tunnel (DNW) is 
located in Emmeloord in The Netherlands. It is a 
world-class acoustic wind tunnel facility that has 
been used for several important international 
rotorcraft acoustic test campaigns since it 
became operational in the late 1970’s. NASA 
and the U.S. Army have previously conducted 
joint helicopter acoustic tests in the DNW. Like 
many of these previous tests, the U.S. Army 
enabled access to the DNW facility for the 
TRAM isolated rotor acoustic test. 

The DNW test focused mostly on low-speed 
helicopter-mode test conditions. The test 
objective priorities were in order of importance: 
detailed acoustic survey of Blade Vortex 
Interaction (BVI) phenomena in helicopter-mode 
descent; broadband noise in hover and low-speed 
helicopter-mode flight; parametric trend data (as 
a function of a s , q, M t i P , and C T ) on helicopter¬ 
mode acoustics; airplane-mode performance and 
acoustic measurements; transition flight 
performance and loads measurements. Because 
of time and load limit constraints, transition 
flight (-15 < a s < -75 degrees) measurements 
were not made. Data was acquired to meet all 
other test objectives. 

The TRAM isolated rotor test stand, as earlier 
noted, was mounted on the DNW sting. The 
DNW sting is articulated to allow for not only 
angle of attack sweeps but beta/yaw angle and 
vertical sting translation sweeps as well. This 
sting articulation capability proved to be very 
useful during the successful execution of the test 
program. The TRAM test stand was positioned 
inside the open-jet test section of the DNW. The 
outer (outside the tunnel flow) containment of 
the test section was acoustically treated with 
foam fairings. A DNW-provided acoustic 


traverse was positioned underneath the model for 
acoustic measurements during the test. The 
DNW also provided laser and associated 
equipment to make laser-light-sheet flow 
visualization and particle image velocimetry 
measurements during the TRAM test. 


Proprotor Performance 


Rotor performance measurements were made 
during the DNW test with the TRAM six- 
component rotor balance and an instrumented 
(torque and residual thrust) flex-coupling (see 
figure 2). 



0 0.005 0.01 0.015 0.02 

CT 


Fig. 2 Hover Figure of Merit (M t i P = 0.63) 

The 1/4-scale V-22 TRAM figure-of-merit data 
compares reasonably well with data from a large- 
scale proprotor hover test at NASA Ames 
Research (Reference 5). No Reynolds number 
corrections have been made for the above data 
points. The two sets of data (airplane- and 
helicopter-mode) reflect the two configurations 
for which hover data was taken. (Helicopter¬ 
mode is when the nacelle incidence angle with 
respect to the motor-mount assembly and DNW 
sting axis is 75 degrees (near perpendicular) and 
airplane-mode is when the nacelle incidence 
angle is zero degrees.) Because of body 
interference effects from the TRAM test stand 
motor-mount sub-assembly, it is not surprising to 
note a fairly substantial impact on hover figure 




of merit for the ‘helicopter-mode’ versus the 
‘airplane-mode’ configurations. 


i 

0.95 


<l) 


t- 0.8 

5 0,5 

2 0.7 

o 

0.65 
0.6 
0.55 
0.5 


Cl 

o 



♦ Mu = 0.325 
□ Mu = 0.350 
O Mu = 0.375 


0 0.001 0.002 0.003 0.004 0.005 0.006 

CT 


Figure 3 - Airplane-Mode, Low-Speed Cruise, 
Proprotor Efficiency (M t i P = 0.59) 


Figure 4 shows a power polar in helicopter-mode 
forward-flight for an advance ratio of |i=0.15 and 
shaft angles ranging from a s = -10 to +12 
degrees. This is only a small subset of the data 
acquired during the DNW test. Hub/spinner 
aerodynamic tares were applied to the figure 4 
performance data. The 1/4-scale V-22 data is 
consistent with full-scale XV-15 helicopter¬ 
mode isolated rotor test results at NASA Ames 
Research Center (Reference 7); power polar 
trends are in general agreement, given the 
differences between rotor solidity and scale 
between the two tests. However, the 1/4-scale 
V-22 performance data is far more 
comprehensive in terms of the C T and rotor shaft 
angle sweeps performed. 


Figure 3 summarizes low-speed airplane-mode 
proprotor efficiency data acquired during the 
test. A clean spinner fairing aerodynamic tare 
was applied to the data in figure 3. This 
performance data was acquired at the V-22 
aircraft’s airplane-mode tip mach number 
(M ti p=0.59) and for q=0.325, 0.35, and 0.375. 
The data was acquired for the maximum 
practical open-jet tunnel velocity, where the test 
section flow field was still reasonably steady. 
Both primary and secondary rotor balance 
measurements are shown in the figure. The 
proprotor efficiency trends measured are 
comparable to performance data from previous 
tests (reference 6). The 1/4-scale V-22 TRAM 
DNW test results, though, will greatly augment 
this extremely limited airplane-mode cruise 
performance data set in the literature. 


0.0014 

H0ASHAFT= -10 deg. 

♦ -6 deg. 



6> 

0.0012 

0 0 deg. 

A+6 deg. 

□ ASHAFT= +12 deg. 


O 

♦ 

0.001 

& 

♦ 

o 


O 

♦ 


0.0008 

£> ♦ 

♦ 

o 


m 

0.0006 

o «> 

4* 

▲ 

mn 

0.0004 

A A 

m ° 

□ 

□ 


0.0002 





o — 

0.008 0.01 0.012 0.014 

CT 

Fig. 4 Helicopter-Mode Power Polar (M t i P =0.63) 


Rotor Structural Loads and Trimmed 
Control Settings 


The TRAM rotor had a strain-gaged blade for 
safety-of-flight monitoring and blade structural 
load measurement. Rotor trim (for zero gimbal 
angle) control settings as a function of rotor shaft 
angles and tunnel speed were also measured. 

Figure 5 shows representative data for the rotor 
control trim settings for cyclic pitch for the 1/4- 
scale V-22 rotor in helicopter-mode flight. The 
rotor was trimmed for zero one-per-rev gimbal 
angle. The cyclic pitch measurements were 
derived from the nonrotating control actuator 
stroke positions, given the kinematic control 
system equations. The general trend for the 1/4- 
scale cyclic pitch trim settings is consistent with 
previous test results for proprotors in low-speed 
helicopter-mode forward-flight (reference 7). 




6 


C7) 

<D 

Q 


CL 

U 

U 

CJ 


ijfl □ 




R D 
D □ 


□ D 


-15 


15 


♦ ♦ 


♦♦Ml 


• ♦ ♦ ♦ ♦ 


♦ Lateral Cyclic Pitch '* 
□ Longitudinal Cyclic Pitch 


Rotor Shaft Angle, Deg. 


Figure 5 - Rotor Cyclic Pitch Trim Settings 
(ji=0.15 and C T =0.009) 


Figure 6a-d is a sample set of contour plots for 
the rotor flap bending loads. The redistribution 
of the flap bending moments across the rotor 
disk can be seen as the tunnel advance ratio 
increases from = 0.125 (Fig. 6a) to 0.20 (Fig. 
6d). Other structural load data were acquired 
during the DNW test, including blade chord 
bending moment and torsional loads and 
pitchlink, pitch-case, and flexbeam loads. The 
data will be used to validate a new generation of 
comprehensive aeromechanics analysis codes. 



(a) |i=0.125 




(b) q=0.15 



(c) |i=0.175 


(d) \i=020 


Fig. 6a-d — Flap Bending Moments for 
Helicopter-Mode Flight (C T =0.013 and a s = -2 
deg.) 


Acoustics 


Among the most crucial information acquired 
during the TRAM DNW isolated rotor test was 
the acoustic data from the 1/4-scale V-22 rotor. 
The TRAM test stand and DNW acoustic 
traverse are shown in figure 7. 









Fig. 7 TRAM Isolated Rotor Configuration and 
the DNW Acoustic Traverse 


Acoustic data were acquired using a combination 
of in-flow traversing and out-of-flow fixed 
microphones. Thirteen microphones were 
equally spaced from -1.86 y/R to 1.86 y/R on a 
traversing microphone wing (also shown in 
figure 7). In addition, two microphones were 
placed outside the test section flow, one above 
the model and another located adjacent to the 
hub on the advancing side of the rotor. Piston- 
phone calibrations, background noise 
measurements, and installed model reflection 
tests were performed. 

For each rotor test condition, the rotor hub height 
was maintained constant while the hub 
longitudinal location was allowed to change with 
shaft angle. The microphone wing was traversed 
along a plane 1.73 z/R beneath the center of the 
rotor hub from -2.76 x/R to 2.76 x/R, centered on 
the actual rotor hub x-location for the test 
condition. At seventeen equally spaced locations, 
the traversing microphone wing motion was 
stopped and data was acquired. 

Figure 8 is a representative contour plot of the 
Blade Vortex Interaction Sound Pressure Level 
(BVISPL) acoustic survey measurements in 
helicopter-mode operation. Prominent in figure 
8 is the BVI ‘hot spot’ on the advancing-side of 
the rotor. 



31 - 1 - 1 - 1 - 1 ■ 

" 2-10 1 2 Quieter 

y/R 


Fig. 8 — Acoustic Survey of Proprotor in 
Helicopter-Mode (BVI Descent Condition) 


The general characteristic of the 1/4-scale V-22 
acoustic survey is similar to 0.15-scale JVX rotor 
results reported in reference 8 and full-scale XV- 
15 results in reference 9. However, the 1/4-scale 
V-22 data from the DNW test is unique in its 
scope (in terms of test envelope and acoustic 
parametric trends measured), its quality (with 
respective to the tunnel flow and acoustic 
characteristics), and its comprehensiveness (with 
respect to the number and types of aeroacoustic, 
aeromechanic, and rotor wake measurements 
made). The acoustic results from the 1/4-scale 
V-22 DNW test will improve the understanding 
and reduction of tiltrotor BVI noise which is 
important for civil tiltrotor passenger and 
community acceptance. 

Acoustic data were acquired for a range of 
advance ratios (p,=0.125, 0.15, 0.175, 0.2), rotor 
shaft angles (a s = -14 to +12 degrees), and thrust 
sweeps (C t = 0.009 to 0.014) were investigated. 
A detailed and comprehensive discussion of the 
1/4-scale V-22 TRAM acoustic results will be 
found in reference 10. One example of the 
BVISPL acoustic trend with increasing q, for 
constant a s and C T , is presented in figure 9. As 
would be expected, the maximum noise levels 
increase with increasing advance ratio. 


















r/R — 33 % 



^ 2-10 1 2 

y/R 

(a.) a = 4.4°, 
|i = 0.15 



ILL = 0.175 p = 0.20 


Figure 9 — BVISPL directivity trend with 
increasing p for C T = 0.013. 


Airloads 


Airloads data for proprotors is extremely limited. 
Reference 11 reported on experimental 
measurements for a generic tiltrotor 
configuration in hover. Several CFD studies 
have been conducted for hovering tiltrotors (for 
example, reference 12) to understand the flow 
mechanisms underlying high thrust conditions. 
Prior to the TRAM DNW test, there existed no 
airloads data for proprotor forward-flight 
operating conditions (either in helicopter- or 
airplane-modes). 

Figure lOa-b is a sample set of sectional pressure 
coefficient data for the 1/4-scale V-22 rotor in 
low-speed helicopter-mode forward-flight. The 
difference between advancing and retreating side 
rotor pressure coefficient distributions is 
demonstrated by comparing figures 10a and 10b. 
The pressure coefficients are 

nondimensionalized by the local velocity. The 
individual pressure coefficient distributions in 
figure lOa-b are all scaled the same. 





0.0 0.2 0.4 0.6 0.8 1.0 

x/c (percent chord) 

r/R = 82% 




0.0 0.2 0.4 0.6 0.8 1.C 

x/c (percent chord) 

r/R = 96% 



0.0 0.2 0.4 0.6 0.8 1.C 

x/c (percent chord) 

r/R = 98% 



0.0 0.2 0.4 0.6 0.8 1.C 

x/c (percent chord) 


Fig. 10a — Helicopter-Mode Forward-Flight 
Sectional Pressure Coefficient Distributions 
(x|i=90 deg., p=0.15, C t = 0.009, & a s = +2 deg.) 





























r/R_= 50% 






r/R = 98% 



Figure lOa-b represents a very limited sample of 
the literally gigabytes of airloads data acquired 
during the 1/4-scale V-22 isolated rotor test at 
the DNW. The airloads data will enable new 
insights into tiltrotor noise mechanisms, will be 
an important validation data set for a new 
generation of aeroacoustic prediction tools, and 
will hopefully inspire new noise reduction 
strategies for tiltrotor aircraft. 


Wake Flow Visualization and 
Measurements 


Both laser-light-sheet (LLS) flow visualization 
and vortex trajectory measurements were made 
for the TRAM 1/4-scale V-22 rotor, as well as 
particle image velocimetry (PIV) vortex velocity 
measurements. Figure 11 is a representative 
laser-light sheet flow visualization picture 
acquired during the DNW test. The rotor blade 
seen in figure 11 is at ip=45 degrees and the 
visible vortices in the figure are on the 
advancing-side of the rotor. 



Fig. 11 — Laser Light Sheet Flow Visualization 
of Trailed Tip Vortices 


Fig. 10b — Helicopter-Mode Forward-Flight 
Sectional Pressure Coefficient Distributions 
(x|i=270 deg., q=0.15, C T =0.009, & a s = +2 deg.) 

Figure 10a is the radial pressure coefficient 
distribution for the rotor advancing-side (xp=90 
deg.) and figure 10b is the distribution for the 
rotor retreating side (i|i=270 deg.). The 
retreating side pressure coefficients are 
significantly greater in magnitude than the 
advancing side coefficients - as would be 
expected for a trimmed rotor in forward-flight. 


Acquisition of a series of laser-light sheet 
pictures enabled the definition of the three- 
dimensional vortex filament trajectories for the 
1/4-scale V-22 rotor on the advancing-side of the 
rotor (fig. 12). Both clockwise and 
counterclockwise vortices were observed in the 
laser-light sheet pictures. The rotation direction 
of the observed advancing-side vortices is noted 
in figure 12. 
















Fig. 12 - Planform View of Vortex Filament 
Trajectories (Low Thrust, Moderate Positive 
Rotor Shaft Angle) 


Reference 13 presents more detailed findings 
from the DNW test with respect to the LLS 
images, the vortex trajectory, and PIV results. 
One important observation made during the 
DNW test was the successful imaging of dual 
rotor vortices being trailed on the advancing side 
of the proprotor in BVI descent conditions. 


Future Plans 

Preparations are underway to conduct a test of 
1/4-scale V-22 proprotors on the Full-Span 
TRAM test stand (figure 13) in the National 
Full-scale Aerodynamics Complex (NFAC) 40- 
by-80 Foot Wind Tunnel at NASA Ames 
Research Center. 



Comparison of 1/4-scale V-22 test results from 
the DNW isolated rotor test and the upcoming 
NFAC full-span TRAM test will enable 
assessment of interactional aerodynamic and 
acoustic effects. 


Conclusion 

NASA and the U.S. Army have made a major 
infrastructure investment in tiltrotor test 
technology through the continuing development 
of the TRAM. This investment has begun to 
payoff through acquisition of fundamental 
aeroacoustic and aeromechanics data from a 1/4- 
scale V-22 isolated rotor tested in the DNW on 
the TRAM isolated rotor configuration. The 
DNW data will enable substantial improvements 
in the predictive capability for tiltrotor aircraft. 
This paper has presented an overview of the 
scope of the data acquired during this 
experimental investigation. Continuation of in- 
depth tiltrotor experimental investigations will 
proceed with tests on a full-span (dual-rotor and 
complete airframe) TRAM configuration. 


Acknowledgements 

The experimental results in this paper were 
derived from research performed under the 
auspices of the Tilt Rotor Aeroacoustic Model 
(TRAM) project and the NASA Short Haul Civil 
Tiltrotor program (SH(CT)). The TRAM and 
SH(CT) programs are led at NASA Ames 
Research Center by the Army/NASA Rotorcraft 
Division and Advanced Tiltrotor Technology 
Project Office, respectively. Other major funding 
partners and research participants in the 
experimental research effort were the U.S. Army 
Aeroflightdynamics Directorate (AFFD) located 
at Ames, NASA Langley Research Center Fluid 
Mechanics and Acoustics Division, and The 
Boeing Company (Mesa, Arizona). In addition, 
the outstanding support provided by the Duits- 
Nederlandse Windtunnel staff during the 
execution of the wind tunnel test was critical to 
the success of the test. 


Fig. 13 - Full-Span TRAM 









The TRAM DNW test team also extends their 
thanks to the staff of the Dutch Nationaal Lucht- 
en Ruimtevaartlaboratorium (NLR) and U.S. 
Army European Research Office (ERO) for their 
technical and programmatic contributions to the 
TRAM development and the DNW wind tunnel 
access, respectively. 


References 

1. Marcolini, M.A., Burley, C.L., Conner, 
D.A., and Acree, Jr., C.W., “Overview of 
Noise Reduction Technology in the NASA 
Short Haul (Civil Tiltrotor) Program,” SAE 
Technical Paper #962273, SAE International 
Powered Lift Conference, Jupiter, Florida, 
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