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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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,
November 1996.
2. Young, L.A., “Tilt Rotor Aeroacoustic
Model (TRAM): A New Rotorcraft
Research Facility,” American Helicopter
Society (AHS) International Specialist’s
Meeting on Advanced Rotorcraft
Technology and Disaster Relief, Gifu,
Japan, April 1998.
3. Van Ditshuizen, J.C.A., “Helicopter Model
Noise Testing at DNW — Status and
Prospects,” Thirteenth European Rotorcraft
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