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Low-Speed Wind Tunnel Investigation of a Full-Scale UH-60 Rotor System
Low-Speed Wind Tunnel Investigation of a Full-Scale UH-60 Rotor System
Thomas R. Norman
Army/NASA Rotorcraft Division
tnorman@mail.arc.nasa.gov
Patrick M. Shinoda
Army/NASA Rotorcraft Division
US Army Aeroflightdynamics Directorate (AMCOM)
pshinoda@mail.arc.nasa.gov
Cahit Kitaplioglu, Stephen A. Jacklin, and Alex Sheikman
Army/NASA Rotorcraft Division
NASA Ames Research Center, Moffett Field, California
ABSTRACT
An experimental program to test a full-scale UH-60 rotor system in the NASA Ames 80- by 120-Foot Wind Tunnel was
completed. The rotor system was installed and tested using a new test stand/facility, the Large Rotor Test Apparatus (LRTA).
The experimental program had both operational and research objectives, including 1) demonstration of LRTA capabilities, 2)
evaluation of an Individual Blade Control system to reduce vibration and noise, 3) acquisition of low-speed performance and
load data for comparison with flight test results and analyses, and 4) validation of a new flow measurement technique. In this
paper, the specific objectives and approach for the wind tunnel test are presented along with examples of the research results.
Particular attention is placed on describing the experimental program, including the new testing capabilities available with
the LRTA.
NOTATION
C
L
Rotor lift coefficient
C
P
Rotor power coefficient
C
T
Rotor thrust coefficient
C
X
Rotor propulsive force coefficient
F
M
Figure of merit
M
TIP
Rotor tip Mach number
r
Radial station
R
Rotor radius
s
Rotor shaft angle measured from vertical, positive
aft, deg
µ
Advance ratio Rotor solidity
INTRODUCTION
Testing of a full-scale UH-60 rotor system in the NASA
Ames 80- by 120-Foot Wind Tunnel was recently
completed. The motivation for this testing was based on the
objectives of two major programs. The first is a cooperative
program between NASA, ZF Luftfahrttechnik GmbH (ZFL),
Sikorsky Aircraft Corporation, and the U.S. Army

Presented at the American Helicopter Society 58
th
Annual
Forum, Montreal, Canada, June 11-13, 2002. Copyright ©
2002 by the American Helicopter Society, Inc. All rights
reserved.
Aeroflightdynamics Directorate to evaluate Individual Blade
Control (IBC) technology on noise, vibration and
performance. The second is a NASA/Army program to
obtain comprehensive measurements of a modern
technology rotor for analysis validation and physical
understanding.
The ultimate objective of the cooperative IBC program is to
demonstrate the technology in flight on a Sikorsky UH-60
helicopter. As a part of the risk reduction effort, it was
decided to first test the UH-60 IBC system in both test
sections of the National Full-scale Aerodynamic Complex
(NFAC) at NASA Ames Research Center. The NFAC
includes both the 80- by 120-Foot Wind Tunnel and the 40-
by 80-Foot Wind Tunnel. The objective of the 80- by 120-
entry was to verify the functionality of the IBC system
(integrated with the wind tunnel systems) and to assess the
effect of IBC on low-speed noise and vibration. Reference 1
describes (and this paper summarizes) the results of this first
test entry. The objective of the planned 40- by 80- entry is to
evaluate the ability of IBC to control noise and vibration at
all airspeeds and to improve rotor performance.
The objective of the second major program is to obtain
comprehensive measurements (including detailed blade
pressure measurements) on a current technology rotor (Ref.
2). In particular, the goal is to acquire data in flight and in
the wind tunnel as well as equivalent data for a model-scale rotor. Testing has been completed for a 1:5.73-scaled UH-
60A model rotor in hover (Ref. 3) and in the wind tunnel
(Ref. 4), as well as for a full-scale rotor in flight (Ref. 5).
The final piece of this program is full-scale wind tunnel
testing of the UH-60A in the NFAC. Although the current
test in the 80- by 120-Foot Wind Tunnel does not fully meet
the goals of the original program (no detailed blade pressure
measurements, no high-speed results), it does provide a
wealth of data for experimental and analytical comparisons.
In this paper, the UH-60 test program in the 80- by 120-Foot
Wind Tunnel is described in detail. Information on test
hardware, instrumentation, and data systems is provided.
Since this was the first test program to use a new rotor test
stand, the Large Rotor Test Apparatus (LRTA), particular
emphasis is given to describing its capabilities. Test
objectives and approaches are then presented, followed by
sample results and discussion.
DESCRIPTION OF THE EXPERIMENT
The test program was conducted in the NASA Ames 80- by
120-Foot Wind Tunnel using a Sikorsky Aircraft UH-60
rotor system mounted on the LRTA. Figure 1 shows the
model installed in the wind tunnel. In the following sections,
detailed information is provided describing the experiment,
including test hardware, instrumentation, data systems, pre-
test activities, standard test procedures, and data validation
activities.
Figure 1. UH-60 Rotor System installed on the Large
Rotor Test Apparatus in the Ames 80-by 120-Foot Wind
Tunnel.
Hardware
The wind tunnel, test stand, and all rotating hardware are
described in this section. Since this was the first test to use
the LRTA, specific emphasis is given to describing its
functionality and capability.
NASA Ames 80- by 120-Foot Wind Tunnel
The 80- by 120-Foot Wind Tunnel is part of the NFAC
located at NASA Ames Research Center (Fig. 2). The tunnel
has an open circuit with a rectangular test section that is 79
ft high, 119 ft wide, and 193 ft long. The maximum test
section velocity is approximately 100 knots. The tunnel
walls are treated with 6 in of acoustically absorbent material
to reduce reflections that can contaminate the noise field.
This material provides an absorptivity of greater than 90%
down to a frequency of approximately 250 Hz. To reduce
contamination by hard surfaces on the test hardware,
additional absorptive material was added to selected hard
spots on the test section floor.
Figure 2. 80-by 120-Foot Wind Tunnel at the National
Full-Scale Aerodynamics Complex (NFAC).
LRTA Test Stand
The LRTA (Figs. 1 and 3) is a special-purpose drive and
support system designed to test helicopters and tilt rotors in
the NFAC. Developed for NASA and the U.S. Army by
Dynamic Engineering, Inc., the LRTA is capable of testing
rotors at thrust levels up to 52,000 lb. Its primary design
features include 1) a drive system powered by two 3000 HP
motors, 2) a five-component rotor balance to measure steady
and unsteady rotor hub loads, along with an instrumented
flex-coupling to measure rotor torque, 3) a six-component
fuselage load-cell system to measure steady fuselage loads,
4) a complete rotor control system (including console) with
primary and higher harmonic control, and 5) an output shaft assembly with a replaceable upper shaft for mating with
different rotor systems.
The current maximum capabilities of the LRTA (and rotor
balance) are provided in Table 1. Note that the maximum
resultant hub moment is dependent upon the installed hub
height above the balance moment center. Also note that
these maximum capabilities represent the limits of the
primary LRTA structure, with the exception of the
replaceable upper shaft. For most rotor test programs, this
upper shaft will have significantly less load-carrying
capability than the rest of the LRTA.
Detailed LRTA Description. The LRTA main support
chassis (Fig. 3) is designed to mount in the wind tunnel on
three struts, two forward and one aft. The length of the aft
strut can be adjusted to vary the pitch angle of the apparatus
(from 30 deg nose down to 15 deg nose up). The chassis also
provides the base to which all LRTA components are
attached, including the fuselage fairing, electric drive
motors, and transmission.
Table 1. LRTA Capabilities
Parameter
Value
Normal Force
52,000 lb
Shear Force (Resultant)
15,000 lb
Moment (Resultant)
125,000 ft-lb *
Torque
165,000 ft-lb
Rotational Speed
320 RPM
Power
6000 HP
Actuator Loads
5000 ± 6000 lb
* at the balance moment center
The fuselage fairing is basically a body of revolution with an
overall length of 480 in and a maximum diameter of 100 in.
The fairing structure is connected to the chassis by a
statically determinant arrangement of six load cells. These
load cells allow determination of the steady aerodynamic
forces and moments on the fairing