DTIC ADA629356: Wind-Tunnel Evaluation of the Effect of Blade Nonstructural Mass Distribution on Helicopter Fixed-System Loads

252 Pages·0.97 MB·English

by  Defense Technical Information Center

NASA/TM-1998-206281
ARL-TR-1401
Wind-Tunnel Evaluation of the Effect of
Blade Nonstructural Mass Distribution on
Helicopter Fixed-System Loads
Matthew L. Wilbur, William T. Yeager, Jr., Jeffrey D. Singleton, Paul H. Mirick,
andW.Keats Wilkie
Vehicle Technology Center
U.S. Army Research Laboratory
Langley Research Center, Hampton, Virginia
January 1998
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NASA/TM-1998-206281
ARL-TR-1401
Wind-Tunnel Evaluation of the Effect of
Blade Nonstructural Mass Distribution on
Helicopter Fixed-System Loads
Matthew L. Wilbur, William T. Yeager, Jr., Jeffrey D. Singleton, Paul H. Mirick,
andW.Keats Wilkie
Vehicle Technology Center
U.S. Army Research Laboratory
Langley Research Center, Hampton, Virginia
National Aeronautics and
Space Administration
Langley Research Center
Hampton, Virginia 23681-2199
January 1998
Available from the following:
NASA Center for AeroSpace Information (CASI) National Technical Information Service (NTIS)
800 Elkridge Landing Road 5285 Port Royal Road
Linthicum Heights, MD 21090-2934 Springfield, VA 22161-2171
[phone] [phone]
Summary Numerous analytical investigations have been con-
ducted to evaluate the effectiveness of aeroelastic tailor-
This report provides data obtained during a wind-
ing in reducing helicopter vibrations. However, few have
tunnel test conducted to investigate parametrically the
included experimental results to verify the conclusions.
effect of blade nonstructural mass on helicopter fixed-
References 2 and 3 offer substantial data sets to validate
system vibratory loads. The data were obtained with
the aeroelastic tailoring techniques and the analyses
aeroelastically scaled model rotor blades that allowed for
used; however, no detailed information is provided about
the addition of concentrated nonstructural masses at mul-
the blade designs, which effectively disallows any fur-
tiple locations along the blade radius. Testing was con-
ther research outside the respective organizations. It is
ducted for advance ratios ranging from 0.10 to 0.35 for
apparent that a substantial need exists for additional
10 blade-mass configurations. Three thrust levels were
experimental data to support the validation of aeroelasti-
obtained at representative full-scale shaft angles for each
cally tailored rotor design techniques.
blade-mass configuration. This report provides the fixed-
system forces and moments measured during testing. The The goal of this research program is to provide a
comprehensive database obtained is well suited for use in detailed experimental database to be used for comparison
correlation and development of advanced rotorcraft and validation of aeroelastic tailoring optimization tech-
analyses. niques, as well as for correlation with and to aid in the
development of new rotor dynamics analyses. The
Introduction approach taken has been to vary parametrically the mass
distribution of an aeroelastically scaled model rotor dur-
Significant research and development efforts have ing testing at the Langley Transonic Dynamics Tunnel
been devoted to the reduction of helicopter vibratory (TDT) and to measure the effect of the distributed mass
loads throughout the years. A survey documenting many on the fixed-system loads and response. This testing has
of the vibration-reduction techniques developed has been resulted in a database that relates one aeroelastic tailoring
presented by Reichert in reference 1. Although tech- parameter (mass distribution) to the resulting dynamic
niques are available to isolate or detune the helicopter response. A summary of these test results appears in ref-
fuselage from forcing frequencies present in the com- erence 4. This paper is intended to provide the reader
plete system, the most obvious course of action is to with the results of the testing in sufficient detail for fur-
reduce the major vibratory forcing functions. For the ther, independent evaluation and analysis.
case of the helicopter, the main rotor generates the bulk
of these vibratory loads and transmits them to the fuse- Symbols
lage primarily through the main-rotor shaft. Therefore,
the main-rotor dynamic response is of primary concern Forces and moments are measured in U.S. customary
when attempting to minimize helicopter vibrations. Both units. The positive directions of forces, moments, angles,
passive- (e.g., pendulum absorbers) and active-control and velocities are shown in figure 1.
(e.g., higher harmonic control) techniques have been suc-
A balance axial force, lb
cessfully used to reduce the main-rotor dynamic response
A amplitude ofnth harmonic, lb or in-lb
(ref. 1). However, these techniques have the disadvan- n
tage of significant weight penalty, added parts, increased a speed of sound, ft/sec
maintenance requirements, and possible degradation of
c blade chord, ft
rotor performance. A better approach is required.
L
It is possible to achieve vibration reduction by C rotor lift coefficient, ------------------------------
aeroelastic tailoring of the rotor. This approach, which L p R2r W( R)2
has the advantage of addressing the vibration problem in
D
the design phase, has become a viable option in recent C rotor drag coefficient, ------------------------------
years because of the increased use of composite materials D p R2r W( R)2
in rotor blade construction. These materials allow the
Q
rotor designer more freedom when selecting blade stiff- C rotor torque coefficient, ------------------------------
ness, mass, torsional inertia, local center of gravity, and Q p R3r W( R)2
elastic-coupling characteristics. Using computer analy-
D rotor drag,N sina +A cosa , lb
ses, the designer may apply these parameters to tune the s s
rotor response in an effort to minimize the rotor loads L rotor lift,N cosa - A sina , lb
s s
that are transmitted to the fuselage. The underlying
assumption is, however, that the analyses used are reli- W R
M rotor-tip Mach number in hover, --------
able and well understood. T a
N balance normal force, lb Finally, the high-density environment increases the
Reynolds number throughout the test envelope, which
n index of harmonic
allows more accurate modeling of the full-scale aerody-
nP frequency,n times rotor rotational frequency
namics of the system. A more detailed discussion of the
W
advantages of the heavy-gas test medium and further dis-
Q rotor torque, in-lb cussions on rotorcraft testing in the TDT are presented in
reference 5.
R rotor radius, ft
r blade radial location measured from center of
Model Description
rotation, ft
T rotor thrust, lb Testbed. Th e ARES-1 (Aeroelastic Rotor
Experimental System 1) generic rotorcraft testbed shown
T thrust required to simulate 1g flight condition
1g in figures 3 and 4 was used for the wind-tunnel testing.
(285 lb for model scale and 18500 lb for full
The ARES-1 is the most basic in a series of three rotor-
scale)
craft testbeds, which are more fully described in refer-
V free-stream velocity, ft/sec ence 5. It is a single-degree-of-freedom system (allowing
x blade chordwise measurement from leading for pitching motions of the shaft) that is soft-mounted to
edge, ft the floor to permit limited flexibility.
a s rotor-shaft angle of attack, deg The ARES-1 model is powered by a variable-
d hub pitch-flap coupling ratio frequency synchronous motor rated at 47-hp output at
3
12000 rpm. The motor is connected to the rotor shaft
q collective pitch at 0.75R, deg
0 through a belt-driven, two-stage speed-reduction system.
Vcosa Control of rotor systems on the ARES-1 testbed is
m rotor advance ratio, ------W-----R---------s achieved through variable shaft angle of attack and a
standard rise-and-fall swashplate. All control is achieved
r density, slugs/ft3 hydraulically with a fly-by-wire control system, with the
f phase angle referenced in direction of rotor shaft angle of attack actuated by one and the swashplate
n
rotation from 0(cid:176) over model tail, deg by three independent hydraulic actuators.
y rotor-blade azimuth angle referenced in direc- Instrumentation on the ARES-1 model allows con-
tion of rotor rotation from 0(cid:176) over model tail, tinuous display of model control settings, rotor speed,
deg rotor forces and moments, blade loads and position, and
W rotor rotational frequency, rad/sec pitch-link loads. All rotating-system data are transferred
through a 30-channel slip-ring assembly to the model
fixed-system. Rotor forces and moments are measured
Apparatus and Procedures
by a six-component strain-gauge balance placed in the
fixed system 21.0 in. below the rotor hub. The balance
Wind Tunnel
supports the rotor pylon and drive system, pitches with
The Langley TDT, a schematic of which is shown in the model shaft, and measures all forces and moments
figure 2, is a continuous-flow pressure tunnel capable of generated by the rotor model. A streamlined fuselage
speeds up to Mach 1.2 at stagnation pressures up to shape encloses the rotor controls and drive system; how-
1atm. The TDT has a 16-ft square slotted test section ever, fuselage forces and moments are not sensed by the
that has cropped corners and a cross-sectional area of balance. A photograph of the ARES-1 testbed with the
248ft2. Either air or refrigerant-12 (R-12), a heavy gas, rotor hardware installed is shown in the TDT test section
may be used as the test medium. For this research, R-12 in figure 4.
was used at a nominal density of 0.006 slug/ft3. The TDT
is particularly suited for rotorcraft aeroelastic testing pri- Rotor.A summary of the model rotor system charac-
marily because of three advantages associated with the teristics is presented in table 1. The rotor used a four-
heavy gas. First, the high density of the test medium bladed, articulated hub with coincident lead-lag and flap
allows model rotor components to be heavier, thereby hinges placed 3.0 in. (0.0534R) from the center of rota-
more easily meeting structural design requirements while tion and pitch-bearings placed directly outboard of the
maintaining dynamic scaling. Second, the low speed of hinges. Rotary potentiometers placed on the hub and
sound of R-12 allows much lower rotor rotational speeds geared to the blade cuffs permitted measurement of blade
and forward flight velocities to match full-scale aerody- lead-lag, flap, and pitch angles. Lead-lag dampers on the
namic parameters (e.g., M and m, among others). hub provided an effective damping output of
T
2
980in-lb-sec/rad. For this test the hub was configured 8.7 percent and 3.5 percent, respectively, of total blade
with a measured pitch-flap coupling ratio of 0.5 mass, if one neglects the mass of the root-end connecting
(d =26.6(cid:176) , flap up-pitch down) with no pitch-link cant. hardware. For the baseline configuration tested, no
3
masses were inserted in the blades. For each of the subse-
The blades tested were a 1/6-size, aeroelastically
quent configurations, a single mass was inserted in each
scaled representation of a U.S. Army candidate design
blade. Mounting a steel or tungsten mass in the blades
for the "growth" version of the UH-60 Black Hawk
produced no change in the chordwise center of gravity
utility-class helicopter. The blades, therefore, were simi-
and negligible changes to the blade torsional inertia and
lar to the advanced blade design in references 6 and 7.
stiffness distributions.
The current blades, however, have provisions for the
addition of insertable nonstructural masses at 13 loca-
Blade structural properties.Model scale structural
tions along the blade span to permit parametric variation
properties are provided in table 2. The flapwise stiffness
of the blade-mass distribution.
and mass distributions are measured values. All other
properties are estimates based on computer analysis of
Aerodynamic design.The blade planform and airfoil
the blade design. The increased mass and stiffness
distributions are shown in figure 5. The blades have a
caused by the root-end hardware are not included in the
wide root chord with a tapered, unswept tip. Taper initia-
data. The chordwise center-of-gravity distribution and
tion is at 0.80R with a 3:1 taper ratio extending to the tip.
The blades have a- 16(cid:176) (nose-down) linear twist distribu- the elastic axis are coincident and placed at the blade
quarter-chord.
tion. The airfoils used in the design are the RC(4)-10,
RC(3)-10, and the RC(3)-08, with blends between the
different airfoil shapes. For each of the airfoils, the aero- Rotating blade modes.To provide confidence in the
dynamic center is considered to be at the quarter-chord. blade construction and structural properties, a rotating
Further information regarding the airfoils used may be shake test was performed to identify the rotating blade
found in references 8 and 9. Information pertaining to the modes and compare them with analysis. The modes and
techniques used to develop the aerodynamic design of frequencies identified are shown in tables 3 and 4 and are
the blades may be found in reference 10. described in the “Results” section of the paper.
Blade construction.In order to evaluate the effects
Blade-mass configurations.Ten blade-mass config-
of blade-mass distribution on fixed-system vibratory
urations were tested. Table 5 provides a listing of the
loads, the blade set was fabricated with a facility for the
configurations and the notation used to designate these
addition of nonstructural mass at discrete locations along
configurations throughout the paper.
the blade radius. An illustration of the concept is shown
in figure 6, with a cross-sectional view of the blade
shown in figure 7. The blade consists of an airfoil glove Test Procedures
and an insertable steel spar. The airfoil glove maintains
The purpose of this test was to measure vibratory
the aerodynamic shape and provides the majority of the
forces and moments in the fixed system while varying
blade stiffness. As shown in figures 6 and 7, the airfoil
parametrically the location and magnitude of discrete
glove has an internal channel centered about the quarter-
nonstructural masses installed in the rotor blades.
chord that is made to accept the steel spar. The spar may
Because frequency response and dynamic response of
be inserted in the channel at the root end of the blade
strain-gauge balances are of utmost concern during test-
glove and locked in place for testing with keys and
ing of this nature, load comparisons were made only
shoulder bolts fastened near the root. The primary func-
when the rotor was operating at a specified frequency.
tion of the steel spar is to provide a mounting area for the
For this test, a rotor speed of 662 rpm was established as
insertable nonstructural masses and to provide the
the nominal frequency and used for all data presented in
mechanical connection of the blade to the rotor hub. Cut-
this paper. Because of changing wind-tunnel conditions,
out areas of the spar, placed every 5 percent of blade
this method had the drawback of not being able to
radius beginning inboard at 0.30R and extending out-
exactly match the hover tip Mach number M (which
board to 0.90R, permit the mounting of either tungsten or T
ranged from 0.656 to 0.680) for each configuration.
steel masses to effect the desired mass-distribution
However, it was determined that it was more important
modifications.
to remove the possibility that balance dynamic response
A single tungsten mass and its associated mounting might contaminate the data than it was to exactly match
hardware provided a mass of 0.00838 slug (0.27 lb). A the aerodynamic environment. Therefore, a dynamic cal-
single steel mass and mounting hardware were 0.00342 ibration of the balance is not required to observe trends in
slug (0.11 lb). These two masses represent approximately the fixed-system loads.
3
Testing was conducted in a heavy-gas test medium at Results
a nominal density of 0.006 slug/ft3, which provided a
nominal Reynolds number of 7.5 · 106/ft at a nominal Blade Frequencies and Modal Moments
hover tip Mach number (M = 0.65). Test data were Four of the 10 configurations were chosen for exper-
T
acquired for nine advance ratios that ranged from 0.10 imental verification of the blade flap and chord frequen-
to0.35. Rotor blade limit loads established an upper cies and the flapwise bending modal moments. The
speed limit ofm= 0.35. At each advance ratio, data were chosen configurations were the baseline configuration
acquired for the rotor-shaft angles of attack shown in and the configurations with a tungsten mass at
table 6, with the rotor collective pitch trimmed to pro- 0.30R(T30), 0.60R (T60), and 0.80R (T80). Blade exci-
duce a 1.0g rotor thrust level. With the shaft angle of tation was insufficient to verify the torsional frequencies
attack held constant, the collective pitch was then re- in the same manner. Tables 3 and 4 provide the experi-
trimmed to acquire data for 0.75g and 1.25g thrust. For mental frequencies obtained and calculated frequencies
all test conditions, rotor cyclic pitch was used to mini- for the elastic blade modes of the four configurations.
mize (<0.1(cid:176) ) the first-harmonic components of blade Figure 8 compares the experimental and calculated
flapping with respect to the shaft. The conditions pro- modal moments for the first elastic flapwise mode. Fig-
vided in table 6 represent the results obtained for the ure 9 compares the experimental and calculated modal
baseline configuration only. Slight variations were noted moments for the second elastic flapwise mode. The cal-
in trim conditions for the other configurations tested. culations were performed using the CAMRAD (ref. 11)
comprehensive rotor analysis.
For each of the test points, a high-speed dynamic Fixed-System Loads
data acquisition system acquired 5 sec of data for each
Tables 7–66 present the results of the fixed-system
channel at a rate of 1000 samples/sec. Concurrently,
loads measured at the ARES-1 fixed-system balance.
wind-tunnel conditions and ARES-1 control settings
Data are presented for the mean load A , one-half peak-
0
were recorded. Fixed-system loads were then reduced off to-peak load, magnitude A , and phase f for the first
n n
line with a Fourier analysis program. This analysis pro-
eight harmonics of rotor speed. Phase angles are pre-
vided the mean, one-half peak-to-peak load, and magni-
sented in degrees referenced to the direction of rotor rota-
tude and phase information for the first eight harmonics tion from 0(cid:176) over the tail of the model. The total load at
of rotor rotational frequency. For these calculations,
any azimuthal location may be reconstructed from the
52rotor revolutions were used.
tabulated harmonics by using the equation
8
During a separate set of tests on a hover test stand in (cid:229)
Load = A + A sin (ny +f )
an air environment, data were acquired to determine the 0 n n
rotating blade bending frequencies and modal moments. n=0
These data were obtained by operating the ARES-1 in The data are grouped according to measured fixed-
hover and allowing rotor wake recirculation to excite the system forces and moments for each rotor configuration
blade modes. Data for rotor speeds of 200 rpm to and thrust condition. Each set of data is presented in
700rpm were obtained. The frequencies and modal order of ascending advance ratio for ease in assessing
moments were reduced from blade strain-gauge data with data repeatability and for comparison with other configu-
a Fast Fourier Transform that provided approximately rations and loads. The data are presented in the following
0.25-Hz resolution for frequencies ranging to 100 Hz. order:
Table for—
Configuration Normal Axial Side
force force force Pitch Roll Yaw
Baseline 7 17 27 37 47 57
T30 8 18 28 38 48 58
T40 9 19 29 39 49 59
T50 10 20 30 40 50 60
T60 11 21 31 41 51 61
T70 12 22 32 42 52 62
T75 13 23 33 43 53 63
T80 14 24 34 44 54 64
T85 15 25 35 45 55 65
S80 16 26 36 46 56 66
4
The 4P fixed-system normal force and pitching and 2. Weller, William H.; and Davis, Mark W.: Wind Tunnel Tests of
rolling moments are presented in bar graph format, Helicopter Blade Designs Optimized for Minimum Vibration.
which permits a more direct comparison of the effects of J. Am. Heli. Soc., vol. 34, no. 3, July 1989, pp. 40–50.
each blade-mass configuration. To generate the bar
3. Young, Darrell K.; and Tarzanin, Frank J., Jr.: Structural
graphs, all data were first plotted as shown in figure 10,
Optimization and Mach Scale Test Validation of a Low
where the 4P fixed-system normal-force loads for the
Vibration Rotor.Proceedings of the 47th American Helicopter
baseline configuration at a 1.0g thrust condition are pre-
Society Annual Forum, AHS, May 1991, pp. 955–968.
sented versus advance ratio. The data for the bar charts
were then developed based on the best-fit curves for each
4. Wilbur, Matthew L.: Experimental Investigation of Helicopter
load-versus-advance-ratio plot. Figures 11–19 present
Vibration Reduction Using Rotor Blade Aeroelastic Tailoring.
the resulting bar graphs for the 4P normal, pitch, and roll Proceedings of the 47th American Helicopter Society Annual
balance loads at each thrust condition and advance ratio Forum, AHS, May 1991, pp. 969–983.
tested.
5. Yeager, William T., Jr.; Mirick, Paul H.; Hamouda,
M-NabilH.; Wilbur, Matthew L.; Singleton, Jeffrey D.; and
Concluding Remarks
Wilkie, W. Keats: Rotorcraft Aeroelastic Testing in the
Langley Transonic Dynamics Tunnel. J. Am. Heli. Soc.,
A database has been obtained in the Langley
vol.38, no. 3, July 1993, pp. 73–82.
Transonic Dynamics Tunnel (TDT) which relates rotor-
craft fixed-system vibratory loads to main rotor blade- 6. Yeager, William T., Jr.; Mantay, Wayne R.; Wilbur, Matthew
mass distribution. The database, obtained with 10 differ- L.; Cramer, Robert G., Jr.; and Singleton, Jeffrey D.: Wind-
ent blade-mass configurations, is particularly suited to Tunnel Evaluation of an Advanced Main-Rotor Blade Design
the validation and development of rotor dynamics analy- for a Utility-Class Helicopter. NASA TM-89129, 1987.
ses and aeroelastic tailoring techniques. The data set,
which was obtained with an aeroelastically scaled model 7. Singleton, Jeffrey D.; Yeager, William T., Jr.; and Wilbur,
rotor system with an advanced aerodynamic blade Matthew L.: Performance Data From a Wind-Tunnel Test of
Two Main-Rotor Blade Designs for a Utility-Class Helicopter.
design, was acquired on the Aeroelastic Rotor
NASA TM-4183, 1990.
Experimental System 1 (ARES-1) testbed. The testing
was performed at full-scale tip Mach numbers and
8. Noonan, Kevin W.: Aerodynamic Characteristics of Two
Reynolds numbers sufficient to provide a realistic aero-
Rotorcraft Airfoils Designed for Application to the Inboard
dynamic environment. This report presents the fixed-
Region of a Main Rotor Blade. NASA TP-3009, 1990.
system force and moment results in tabular and graphical
formats. 9. Bingham, G. J.; and Noonan, K. W.: Two-Dimensional
Aerodynamic Characteristics of Three Rotorcraft Airfoils at
Mach Numbers From 0.35 to 0.90. NASA TP-2000, 1982.
NASA Langley Research Center
10. Bingham, G. J.: The Aerodynamic Influences of Rotor Blade
Hampton, VA 23681-2199
Airfoils, Twist, Taper and Solidity on Hover and Forward
April 15, 1997
Flight Performance. Proceedings of the 37th American
Helicopter Society Annual Forum, AHS, May 1981,
References pp. 37–50.
1. Reichert, G.: Helicopter Vibration Control—A Survey. 11. Johnson, W.:A Comprehensive Analytical Model of Rotorcraft
Vertica, vol. 5, no. 1, 1981, pp. 1–20. Aerodynamics and Dynamics. NASA TM-81182, 1980.
5
Table 1. Model Rotor System Description
Number of blades . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4
Radius, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.685
Root cutout,r/R. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .0.213
Root chord, ft. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .0.450
Tip chord, ft. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .0.150
Taper initiation,r/R. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .0.800
Taper ratio . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3:1
Solidity:
Area weighted. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .0.114
Thrust weighted. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .0.101
Torque weighted . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .0.0956
Twist, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .- 16 (linear)
Pitch axis,x/c. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .0.25
Elastic axis,x/c . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.25 (straight)
Chordwise center of gravity,x/c . . . . . . . . . . . . . . . . . 0.25 (straight)
Aerodynamic center,x/c . . . . . . . . . . . . . . . . . . . . . . . 0.25 (straight)
Flap hinge,r/R. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .0.0534
Lag hinge,r/R . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .0.0534
Pitch horn offset,x/R. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .0.0249
Pitch-link radial attachment,r/R . . . . . . . . . . . . . . . . . . . . . . .0.0400
6