The sensitivity of blade tracking in hover to variations in root pitch was examined for two rotor configurations. Tests were conducted using a four bladed articulated rotor mounted on the NASA-Army aeroelastic rotor experimental system (ARES). Two rotor configurations were tested: one consisting of a blade set with flexible fiberglass spars and one with stiffer (by a factor of five in flapwise and torsional stiffnesses) aluminum spars. Both blade sets were identical in planform and airfoil distribution and were untwisted. The two configurations were ballasted to the same Lock number so that a direct comparison of the tracking sensitivity to a gross change in blade stiffness could be made. Experimental results show no large differences between the two sets of blades in the sensitivity of the blade tracking to root pitch adjustments. However, a measurable reduction in intrack coning of the fiberglass spar blades with respect to the aluminum blades is noted at higher rotor thrust conditions

Langston, Chester W.

Mirick, Paul H.

Singleton, Jeffrey D.

Wilbur, Matthew L.

Wilkie, W. Keats

Yeager, William T., Jr.

English

NASA Technical Reports Server

NASA
Technical Memorandum 4313
AVSCOM
Technical Report 91-B-017
An Experimental Study of the
Sensitivity of Helicopter Rotor
Blade Tracking to Root Pitch
Adjustment in Hover
W. Keats Wilkie, Chester W. Langston,
Paul H. Mirick, Jeffrey D. Singleton,
Matthew L. Wilbur, and William T. Yeager, Jr.
Aerostructures Directorate
U.S. Army-AVSCOM
Langley Research Center
Hampton, Virginia
National Aeronautics and
Space Administration
Office of Management
Scientific and Technical
Information Program
1991
https://ntrs.nasa.gov/search.jsp?R=19920003771 2020-03-17T14:01:45+00:00Z

SuInnlary
The sensitivity of blade tracking in hover to w_ria-
tions in root pitch was examined for two rotor config-
urations. Tests were conducted using a four-bladed
articulated rotor mounted on the NASA/U.S. Army
aeroclastic rotor experimental system (ARES) at tile
Langley Research Center. Two rotor configurations
were tested: one consisting of a blade set with flexi-
ble fiberglass spars and one with stiffer (by a factor
of five in flapwise and torsional stiffnesses) aluminuin
spars. Both blade sets were identical in planform and
airfoil distribution and were untwisted. The two con-
figurations were ballasted to the same Lock number
so that a direct comparison of the tracking sensitivity
to a gross change in blade stiffness couht be made.
Experimental results show no large differences
between the two sets of blades in the sensitivity of the
blade tracking to root pitch a(t.iustments. However,
a measurable reduction ill in-track coning of the
fiberglass spar blades with respect to the aluminum
blades is noted at. higher rotor thrust conditions.
Introduction
A major concern in developing any new rotor
design is tile relative sensitivity of that design to
track adjustnmnt (or nlisadjustment). An out-of-
track blade (;all often cause unacceptable once-per-
revohltion vibrations on a helicopter, and correct-
ing such a track problem can be a laborious and
time-consunfing t)rocess. Usual methods employed
to correct a track problem involve adjusting blade
root pitch, deflecting trailing edge tabs, or adding
balance weights to lower the once-per-revolution vi-
brations caused by the errant blade to a.cceI)table
levels. Such methods are for the most part trial-and-
error in nature and unresponsive blades oft(,u must
be discarded altogether. Past studies (refs. 1 and 2)
have suggested that torsionally %oft" rotor blades
are especially sensitive to parametric track adjust-
merits. The rather large sensitivity, and relatively
unpredictable tracking response, of such blades to
small tracking adjustnmnts was recognized as a po-
tential problenl for full-scale rotors. As a first step
to understanding out-of-track blade phenonlena, this
study attempts to experimentally quantify the track-
ing sensitivities in hover of two rotor configurations
with comparatively large differences in blade struc-
tural stiffness. The results presented in this paper ad-
dress only tracking response (tile to discrete changes
in root pitch, and only for an articulated rotor.
Symbols
a blade section two-dimensional lift-curve
slope, per radian
b number of blades
c blade chord, fl
CT rotor thrust coefficient, T/[prcR2(t_R) '2]
I b rotor blade flapping mass moment of invrtia
about flapping axis, slug-ft 2
I o rotor blade torsional mass moment of
inertia, per refit length about blade elastic
axis, lb-scc ')
/i' rotor radius, ft
r spanwise distance along t)lade radius
measured from center of rotation, ft.
T rotor thrust force, measured fl'om balance
normal-force channel, lb
/_ blade coning angle, (leg
A/'_ induce(t blade out-of-tra.ck with respect to
reference blade, (tog
")' blade Lock number, cpoI?l/Ib
0 rotor blade collective pitch angle, (leg
AO offset increment of root t)itch at)t)lie(t to
perturb blade, deg
[i nlass density of test Illediulll, slug/ft :_
o- rotor solidity, bc/rcl?
[_ rotor rotational velocity, rad/scc
ca natural frequency of rotating blade mode,
rad/sec
Apparatus and Procedures
Test Facility
Testing was conducted in tile Langley Helicopter
Hover Facility' (HHF), shown in figure 1. The HHF ix
a high-t/ay facility enclose(t by a 30-ft x 30-ft x 20-ft
coarse-mesh screen and is used for h(/ver testing
and rotoreraft model buildup and checkout prior to
testing in the Langley Transonic Dynanlics Tunnel
(TDT). Models are mounted on tile test stand such
that the rotor plane of rotation is high enough above
the floor to avoid ground effect (15 ft, or approxi-
mately 1.6 times the rotor diameter). All hover text-
ing in the HHF is performed at. sea level atmospheric
conditi(ms (nominal p = 0.002378 slug/ft.'_).
Model Description
A four-bladed articulated hul) with coincident
lead-lag and flapping hinges was used in this investi-
gation. Two sets of rotor blades were used: one set
with flexiblefiberglassparsandonesetwith stiffer
aluminumspars.Thestructuralandinertialproper-
tiesof both bladesetsarelistedin tablesI andII.
Rotatingnatural frequencieswerecomputedusing
the ComprehensiveAnalyticalModelof Rotorcraft
AerodynamicsandDynamics(CAMRAD)computer
code(refs.3 and4). Thefiberglassbladesetwasde-
signedfor usein theR-12("Freon-12")testenviron-
mentof theTDT andhasscaledaeroelastieproper-
tiesin R-12similarto thoseofafull-scaleutility-class
helicopter.The aluminumbladeset wasdesigned
for Math-number-sealedtestingin air. Dueto rotor
speedlimitationsof thetest-bedmodel,full-scaletip
Machnumbervalueswerenot possiblefor this test.
Bothbladesetshaveidenticalrectangularplanforms,
areuntwisted,andusea NACA0012airfoil. Blade
geometryfor bothsetsis shownin figure2. Theso-
lidity for both rotorconfigurationswas0.0982.The
massandinertialcharacteristicsofthebladesmaybe
variedby alteringthedistributionsof tungstenand
aluminumweightslocatedin twospanwisechannels
alongeachblade.Locknumbersof bothbladesets
werematched(_ = 4.35)by appropriatedistribu-
tionsof theseballastweights,thuseliminatingtrack
effectscausedbyvariationsin bladeflappinginertia.
Theballastweightsdonot alter the stiffnessdistri-
butionsof theblades.
The test-bedusedfor this experimentwasthe
NASA/U.S.Armyaeroelasticrotorexperimentalsys-
tem (ARES)model,shownin figure3. TheARES
modelhasa streamlinedfuselageshellthat encloses
the rotor controlsanddrive system.The fuselage
shellisnot usuallyinstalledwhentestingtheARES
modelin theHHFandwasomittedduringthis test.
Themodelrotor is poweredbya variable-frequency
synchronouselectricmotor(ratedat 47 hp output
at 12000rpm) that is connectedto therotor shaft
througha belt-driventwo-stagespeedreductionsys-
tem. Bladecollectivepitch inputs andlateraland
longitudinalcyclicpitchinputsareprovidedthrough
a conventionalswashplatearrangement,with the
swashplatepositionedbythreeelectricallycontrolled
hydraulicactuators.Rootpitchadjustmentsto in-
dividual bladesare madeusinga motorizedpitch
link system.Therotor controlsystemis remotely
operatedfrom the HHF controlroom,with instru-
mentationmountedon the ARESmodelproviding
a continuousdisplayof model control settings, ro-
tor forces and moments, blade loads, and pitch link
loads. Rotary potentiometers mounted at the blade
cuffs are used to measure flapping and lagging on
two of tile four blades. These potentiometer sig-
nals are transferred to the fixed system through a
30-channel slip-ring assembly. Rotor forces and too-
ments are measured in the nonrotating system by a
six-component strain-gauge balance mounted below
the drive system.
Test Procedure
This test was designed to provide a direct com-
parison of the coning and tracking characteristics of
the two blade configurations. Baseline data were ob-
tained for each set of blades, with all blades tracked
with respect to the reference blade. At each test
condition the rotor speed was set to 650 rpm. Blade
collective pitch was swept fl'om 0 ° to 16 °, with data
being collected at every 1° increment. Repeat mea-
surements were made from 16 ° collective pitch to 0°
collective pitch in increments of 2° . The nominal tip
Mach number was 0.27 for the entire test. Blade
track was visually monitored by means of a strobo-
scopic light system. Subsequent runs were performed
similarly with the blade opposite the reference blade
forced out of track by providing an offset in the root
pitch of that blade (A0 = -2 °, -1 °, +1 °, +2°). At
each test point 5 seconds of data were obtained (cor-
responding to approximately 54 rotor revolutions).
Blade position data and rotor force and moment data
were sampled at a rate of 1000 samples per second per
channel, averaged, and stored digitally. Rotor thrust
and torque were calculated from tile balance normal
force and yawing moment channels, with balance in-
teractions removed during off-line data reduction.
Presentation of Results
Plots are presented in figures 4 and 5 of induced
blade out-of-track (A_) versus the ratio of thrust co-
etticient to solidity (CT/c_) for each rotor configura-
tion. Figure 6 is a comparison of blade out-of-track
versus blade root pitch offset for both sets of blades
at various fixed thrust levels. Coning response ver-
sus rotor thrust for the baseline in-track cases of each
rotor arc compared in figure 7.
Because of an error in setting the root pitch dur-
ing one run, measurements at the desired A0 of +2 °
were not made for the aluminum blades. '/'he A0
for the run shown is estimated to be approximately
+1.5 °. Repeatability for CT/o" has been estimated to
be within -t-0.0025. Accuracy of the angular measure-
ments is estimated to be within 4-0.1 °. Repeatability
for the angular measurements is within :t:0.05 ° .
Discussion of Results
The results presented in figures 4 and 5 show that
both rotors exhibit essentially the same trends in
induced blade out-of-track due to the incremental
adjustments made in root pitch. Cross-plots of A_
versus A0 at fixed values of CT/Cr (fig. 6) indicate
that the fiberglass blades are perhaps slightly less re-
sponsive to the root pitch inputs than tile ahuninum
blades. It should be emphaMzed that tile differences
in sensitivity shown here are on the limits of the res-
olution of tim data.
Figure 7 shows the in-track coning angle (/_) ver-
sus the ratio of thrust coefficient to solidity (Cr/cr)
for both configurations. These data show a slight but
measurable difference in the coning behavior of the
two rotor configurations at higher values of CT/a;
specifically, the fiberglass spar blades exhibit less
coning (approximately 0.2 °) than the atuminuln spar
blades. This behavior could be the result of an effec-
tive nose-down twist caused by "propeller nloment"
(ref. 5) acting on the blades at the higher collective
pitch settings. As the fiberglass blades are signifi-
cantly less stiff ill torsion than the ahuninum blades,
they presumedly would be more susceptible to pro-
peller moment-induced twist.
Concluding Remarks
The sensitivity of blade tracking in hover to vari-
ations in root. pitch was determined experimentally
for two rotor configurations possessing large differ-
ences in flapping and torsional stiffnesses. The data
obtained here indicate that both rotors exhibit, for
the most part, a similar sensitivity to root pitch ad-
justment in hover. However, slight differen(:es ill
coning response between the two configurations are
observed at higher thrust levels. Although no analy-
sis is presented in this paper, the experimental data
shown here may eventually prove useful to the ro-
tor blade designer in diagnosing or preventing sonic
blade tracking difficulties, and to the analyst in val-
idating future rotor aeroela_stic theories.
NASA Langley Research Center
Hampton, VA 23665-5225
November 5, 19!)1
References
1. Mantay, \Vayne R.; and Yeager, \Villiam T., .Jr.: Para-
metric Tip Effects for CoT_formable Rotor Applications.
NASA TM-85682, AVRADCOM Ttt-S3-B-4, 1983.
2. Mantay, Wayne R.; and Yeager, William T., ,Jr.: Acro-
dastic Considerations for TorsioTudly S@ Rotors. NASA
TM-87687, USAAVSCOM TR-86-B-1, 1986.
3. Johnson, Wayne: A Comprehensive A_alytical Mod_l of
Rotorcraft Aerodynamics a_zd Dynamics. Part I: Analg-
sis Development. NASA TM-81182, USAAVRADCOM
TR-80-A-5, 1980.
4. ,Johnson, Wayne: ,4 Comprehensive Analytical Model of
Rotorcrafl Aerodgrtamics and D?lnamic._, Part ll: User's
Manual. NASA TM-81183, USAAVSCOM TR-80-A-6,
1980.
5. Johnson, Wayne: Hclicopttr TheoEq. Princeton Univ.
Press, c. 198{}.
TableI. Propertiesof FiberglassModelBlade
(a) Structuralproperties
Inboard
section,
r/R
0.055
.125
.161
.193
.227
.280
.284
.325
.432
.514
.927
.959
.964
.982
.986
Stiffness, lb-ft 2
Section
Inass;
slugs
5.11 × 10 2
Flap
3.47 × 104
Chord
3.47 × 101
Torsion
6.94 x 103
4.57 x 10 -3
2.22 × 10 -3
8.82 x 10 -3
1.31 x 10 -2
2.55 x 10 -4
3.48 x 10 -3
8.65 x 10 a
6.43 x 10 -a
3.18 x 10 -2
2.45 x 10 -a
3.67 x 10 -4
2.45 × 10 3
4.72 x 10 -1
2.57 × 10 -4
3.47 x 103
2.78 x 102
2.69 x 102
2.69 x 102
1.73 × 102
1.73 x 102
1.60 x 102
1.26 × 102
1.05 x 102
1.05 x 102
1.11 x 102
3.47 x 102
2.81 x 102
3.47 x 101
1.04 x 104
1.74 x 103
1.86 x 10 a
1.86 x 103
1.75 x 10 a
1.75 x 103
2.11 x 103
1.83 x 10 a
1.70 x 10 a
1.70 x 10 a
1.70 x 10 a
3.47 x 103
2.78 x 10 a
3.47 x 102
3.47 x 103
2.78 x 102
3.04 x 102
3.04 x 102
2.36 x 102
2.36 x 102
1.63 x 102
1.37 x 102
1.17 x 102
1.17 x 102
1.22 x 102
3.47 x 102
2.78 x 102
3.47 x 101
/07
lb-sec 2
5.70 × 10 -4
1.14 x 10 4
6.49 x 10 -5
1.46 x 10 4
1.56 x 10 4
8.41 x 10 -5
8.98 x 10 -5
8.68 x 10 -5
8.49 × 10 .5
8.35 x 10 -5
8.35 x 10 -5
8.52 × 10 -5
1.03 x 10 -4
9.23 x 10 .5
1.14 x 10 -5
(b) Blade rotating natural frequencies
/vlodal identity *w/_
Flap 2.61
Chord 4.18
Flap 4.83
Torsion t5.61
*f_ = 68.07 rad/sec
TableII. Propertiesof AhmfinumModelBlade
(a) Structuralproperties
Inboard
section,
,/R
0.055
.125
.161
.193
.227
.280
.284
.325
.400
.432
.514
.918
.927
.959
.964
.982
.986
Stiffness, lb-ft 2
Section
mass,
slugs
5.11 x 10 -2
Flap
3.47 x 104
Chord
3.47 x 104
Torsion
6.94 x 103
4.57 x 10 -3
2.27 x 10 .3
8.59 x 10 .3
1.39 x 10 2
3.27 x 10 4
2.95 x 10 3
4.67 x 10 .3
2.63 x 10 .3
6.37 x 10 3
2.98 x 10 -2
8.61 x 10 .4
3.13 x 10 .3
4.72 x 10 -4
2.62 x 10 .3
5.48 × 10 .4
2.56 x 10 -4
3.47 x 103
1.01 x 103
9.72 x 102
9.72 x 102
1.01 × 103
8.47 x 102
7.47 x 102
7.47 x 102
6.20 x 102
5.22 x 102
5.22 x 102
5.22 x 102
6.25 x 102
9.72 x 102
8.33 x 102
3.47 x 101
1.04 x 104
9.38 x 103
8.68 x 103
8.68 x 103
9.38 x 103
8.68 x 103
8.47 x 103
8.47 x 103
8.33 x 103
7.92 x 103
7.92 x lO 3
7.92 x 103
8.33 x 103
8.33 x 103
7.64 x 103
3.47 x 102
3.47 x 103
1.39 x 103
1.04 x 103
1.04 x 103
1.39 × 103
1.15 x 103
1.02 x 103
1.02 x 103
7.76 x 102
6.47 x 102
6.47 x 102
6.47 x 102
6.94 x 10 `)
6.94 x 102
6.25 x 102
3.47 x 101
0_
lb-sec 2
5.70 x 10 -'i
1.14 x 10 -1
4.70 × 10 -5
7.43 x 10 -5
1.15 x 10 .4
8.87 × 10 5
7.95 x 10 5
7.49 x 10 -5
9.84 x 10 5
9.38 x 10 -r,
9.06 x 10 -5
1.17 x 10 ,l
1.17 x 10 -4
1.21 x 10 -4
1.11 x 10 -1
9.23 x l0 -5
1.14 x 10 -r,
(b) Blade rotating natural frequencies
Modal identity *w/ft
Flap 3.28
Flap 7.78
Chord 7.99
Torsion 15.50
*f_ = 68.07 rad/sec
5
ORIGINAL PA_'_E
BLACK AND WHITE Pr-iO1-OGRAPI',
Figure 1. Helicopter Hover Facility (HHF).
6
NACA 0012 airfoil
--- 4.24
"_---- 12.5
Center of rotation
Figure 2.
55.0
Rotor blade geometry. Dimensions in inches.
l-
.!
Pitch axis
(at c/4)
7
.)"'_!'I,_.ILPAGE
_I.ACt_ &NO Wi-IIIL PtiOt-OGRAPH
Figure 3. ARES model mounted in HHF.
L-86-11,726
8
Ap,
deg
1.0 --
\
.5
<
-.5
-1.0 I I I
0 .04 .08 .12
CT/_
Figure 4. Tracking response to root pitch adjustment for fiberglass spar blades.
A0, deg
O o
[] -2
A -1
+2
1.0 --
.5
0
deg
-.5
-1.0
YP"C;_V_7 V'_' VVV V V _Y V V
_000@ 0<_ <_ 0 @ 0 0 0 0
Z)[O0_ 0 (_ 5) O 0 0 0 0 0
_.AA& A_ AAA AA AA A
_i"l_ i-i_ r-l_l I:El ri Eli r-i []
I I I
0 .04 .08 .12
CT/_
A@,deg
O o
1"-1-2
A-1
<}+1
V +1 5
Figure. 5. Tracking response to root pitch adjustment for ahlminuui spar blades.
9



An experimental study of the sensitivity of helicopter rotor blade tracking to root pitch adjustment in hover

Abstract

Similar works

Full text

Available Versions

NASA Technical Reports Server