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NASA Technical Memorandum 101 040
Hub and Blade Structural Loads Measurements of an SA349/2 Helicopter
Gloria K. Yamauchi Ruth M. Heffernan, Ames Research Center, Moffett Field, California Michel Gaubert, Aerospatiale Division Helicopteres, Marignane, France
December 1988
National Aeronautics and Space Administration
Ames Research Center Moffett Field, California 94035
https://ntrs.nasa.gov/search.jsp?R=19890010753 2020-03-12T02:03:51+00:00Z
NOMENCLATURE
a
A
A,
An
b
BL
Bn
C
COLL
CT
CZM
EIf
E11
ftx
ftY
ftz
FS
FX
FY
FZ
g
GJ
IAS
speed of sound (m/s)
rotor area (m2)
mean coefficient (harmonic analysis)
cosine coefficient (harmonic analysis); n = harmonic number
number of blades = 3
buttline (m)
sine coefficient (harmonic analysis); n = harmonic number
chord (m)
collective pitch, referenced to 75% radial station (deg)
thrust coefficient
thrust parameter
blade flapwise bendlng stiffness (N-m2)
blade edgewise bending stiffness (N-m2)
total rotating in-plane force at hub in xs-direction (N)
total rotating in-plane force at hub in ys-direction (N)
total rotating out-of-plane hub force (N)
fuselage station (m)
total nonrotating in-plane force in Xs-direction (N)
total nonrotating in-plane force in Ys-direction (N)
total nonrotating out-of-plane force (N) (Fz is equivalent to ftz)
gravitational acceleration (m/s2) = 9.8
torsional stiffness (N-m2)
indicated air speed (m/s)
... 111
PRECEDING PAGE BUNK NOT FILMED
blade flap inertia (kg-rn2)
blade lag inertia (kg-m2)
blade polar moments of inertia (kg-m2) = jj (Y2 + 22) dm, I./ (22 - Y2) dm, respectively
sectional polar moments of inertia (kg-mz/m) = jj (Yt2 + Zt2) dm, 11 (Zt2 - Yt2) dm, respectively
aircraft principal moments of inertia (kg-m2) = X2 dm, !I Y2 dm, !I 2 2 dm, respectively I
aircraft products of inertia (kg-m2) = !! X Y dm, I! XZ dm, YZ dm, respectively 1
lag damper displacement (mm)
vertical distance from shaft strain gage at A to rotor hub center (m) = 0.394
vertical distance from shaft strain gage at B to rotor hub center (m) = 0.149
distance between lag hinge and lag damper line of action (m)
lateral blade cyclic pitch, positive roll to the right (deg)
longitudinal blade cyclic pitch, positive nose up (deg)
helicopter mass (kg)
reduced helicopter mass (kg)
total rotating in-plane moment at shaft section A in xs-direction (N-m)
total rotating in-plane moment at shaft section A in ys-direction (N-m)
total rotating in-plane moment at shaft section B in xs-direction (N-m)
total rotating in-plane moment at shaft section B in ys-direction (N-m)
total rotating in-plane hub moment in xs-direction (N-m)
total rotating in-plane hub moment in ys-direction (N-m)
total rotating out-of-plane hub moment (N-m)
blade sectional mass (kg/m)
main rotor
I
total nonrotating in-plane hub moment in Xs-direction (N-m)
iv
MY
MZ
n
N
. Po
Ps
r 4
R
STAT
TAS
TO
TS
TR
total nonrotating in-plane hub moment in Ys-direction (N-m)
total nonrotating out-of-plane hub moment (N-m) (Mz is equivalent to mtz)
harmonic number
load factor
standard atmospheric pressure (mbar) = 1013.25
static flight pressure (mbar)
blade radial location (m)
rotor radius (m)
static
true air speed ( d s )
standard atmospheric temperature ( O K ) = 288.16
static flight temperature ("C)
tail rotor
velocity ( d s )
water line (m)
aircraft c.g. axis system (origin at aircraft c.g., fig. la)
blade reference axis system (origin at one-quarter chord, fig. 5 )
rotating-shaft reference axis system (origin at hub center, fig. 14)
wind axis system (origin at aircraft c.g., fig. IC)
rotor axis system (origin at hub center, fig. lb)
blade principal inertia axes (origin at chordwise c.g., fig. 5 )
nonrotating-shaft reference axis system (origin at hub center, fig. 15)
YG
YN
ZN
chordwise blade c.g. location in xb,yb,zb axis system (m)
Yb-axis location in xb,yb,zb axis system (neutral axis) (m)
Zb-aXiS location in xb,yb,zb axis system (m)
V
a
w R
angle of attack (deg)
specific heat ratio for air = 1.4
lag angle (deg)
pitch angle (deg)
advance ratio
air density (kg/m3)
standard atmospheric density (kg/m3) = 1.225
solidity
density ratio
angle between Yb,Zb axes and yb,zb axes (deg)
azimuth angle, measured from tail boom toward advancing side of rotor disk (deg)
rotor rotational speed (rads)
I
vi
SUMMARY
Data from 23 flight conditions, including level flights ranging from p. = 0.14 to 0.37 and steady turning flights from p = 0.26 to 0.35, are presented for an Aerospatiale SA349/2 Gazelle helicopter. The data include hub loads data (for 6 of the 23 conditions), blade structural data at eleven different blade radial stations, and fuselage structural data. All dynamic data are presented as harmonic analysis coefficients (ten harmonics per rotor revolution). This report also documents the data acquisition and reduction proce- dures. Blade structural and inertial properties are provided in addition to control system geometry and properties.
INTRODUCTION
Measuring and analyzing helicopter oscillatory hub loads is an important step toward the ultimate goal of minimizing vibration. One important purpose of the 1987 flight test of the SA349/2 helicopter was to measure rotor hub loads. This was accomplished by instrumenting the rotor shaft at two locations and determining the resulting rotating and nonrotating loads at the hub. Additional flight measurements acquired at two different blade chordwise c.g. locations, included blade moments at eleven radial stations, pitch link loads, blade flap and lag angles, shaft angle, seat accelerations, servo loads, gearbox strut loads, and lag damper force. Data for 23 flight conditions covering a wide range of thrust levels and speeds are documented in this report.
This flight test of the Aerospatiale SA349/2 helicopter initiated the second phase of joint work between the National Aeronautics and Space Administration and the French Ministry of Defense. The objective of this program is to gain a better understanding of rotor aerodynamic and dynamic phenomena by correlating flight test data from the SA349/2 with French and U.S. rotor analyses. The first phase of cooperative work also consisted of an extensive flight-test program, which was conducted in 1984 in France. The 1984 flight test produced blade airloads and structural loads data (ref. 1). The data documented in this report complement the data given in reference 1 since many of the 1984 flight conditions were duplicated during the 1987 flight program (also conducted in France).
AIRCRAFT AND ROTOR DESCRIPTION
Properties and characteristics of the aircraft and rotor are given in reference 1, but some are repeated . here for convenience.
w SA349/2 Aircraft
The SA349/2 aircraft consists of a three-bladed, fully-articulated main rotor and a Fenestron tail rotor. Basic aircraft information is given in table 1. Reference axes and sign conventions for aircraft loads are shown in figures 1 and 2, respectively.
1
Research Rotor
I
The rotor consists of the Non Articule en Trainee (NAT) hub and three, constant-chord Grande Vitesse (GV) blades. The GV blades are rectangular tipped and have OA209 airfoil profiles. Rotor characteristics and control system geometry are provided in table 2 and figure 3. Figure 4 shows the GV blade planform and twist distribution. Reference axes for the blade sectional properties are shown in figure 5. The sec- tional characteristics of the blade are tabulated in tables 3 and 4 for chordwise c.g. locations of 24.7% and 26.7%, respectively. The blade chordwise c.g. location is varied by removing or inserting weights in a tube that is part of the blade structure. The tube extends along the length of the blade and is located at the 25% chordwise station. The rotating frequencies of the GV blade for the two different chordwise c.g locations are shown in figure 6a. The calculations were made using CAMRAD (ref. 2). Similar calcula- tions were made using an Aerospatiale analysis, IAK40 (described in ref. I), as shown in figure 6b. The CAMRAD results were computed in vacuo, while the IAK40 results include the effects of aerodynamics. Figures 7 and 8 show the IAK40 calculated blade damping ratio and modal frequencies for both c.g. loca- tions. Note that figure 7 shows a calculated instability at the operating rpm for the 26.7% c.g. location, which did not occur during the flight test. This calculated instability may be the result of neglecting unsteady aerodynamics in the IAK40 analysis.
,
INSTRUMENTATION
The aircraft instrumentation and data acquisition processes are described in this section.
Flight Condition Parameters
The environmental conditions of the flight test were fully documented. Flight condition parameters recorded for each flight test included: helicopter mass; altitude; indicated airspeed; ground pressure and temperature; static flight pressure and temperature; collective-pitch stick position; lateral and longitudinal cyclic pitch stick position; tail rotor pedal position; aircraft pitch and roll angles; aircraft pitch, roll, and yaw rates; load factor; rotor rotational speed; and engine power. All stick positions were translated into the equivalent pitch angles. The collective pitch was referenced to the 75% radial station value.
Rotating Frame Instrumentation
The GV blades were instrumented with 31 strain gages to measure structural loads at eleven radial sta- tions (fig. 9). Flapwise, edgewise, and torsion moments were measured at 9, 10, and 6 radial stations, respectively. Two of the blades were instrumented to measure blade flap angle and lag damper displace- ment. Figure 10 shows locations of the rotating NAT hub instrumentation. The flap angle was measured using an angular sensor located about the flap hinge. The sensor is composed of two components: one component is fixed relative to the blade flap motion and the other component rotates as the blade flaps about the hinge. The sensor delivers an induced sinusoidal tension. The linear part of this sinusoidal sig- nal is proportional to the rotation angle of the sensor rotating component. The blade lag displacement was measured using a potentiometer. Properties of the lag damper are given in reference .1. All three pitch links were instrumented.
2
The rotor shaft instrumentation produced four measurements: shaft bending moments at two locations, shaft torque, and shaft vertical force. Figure 11 shows the gage locations. FlSHAFT and F3SHAFT each represent a gage couple (two gages located diametrically opposite each other). The direction of the bending moment is perpendicular to the diameter formed by the gage couple.
Fixed Frame Instrumentation
Fixed frame instrumentation included strain gages on two of the three primary control system servos and strain gages on each of the four main gearbox struts. In addition, pilot- and copilot-seat vertical vibration were measured using accelerometers. Locations of all nonrotating instrumentation are shown in figure 12.
On-Board Data Acquisition
A schematic of the data acquisition process is shown in figure 13. The flight condition parameters were digitized by an onboard modular acquisition unit (UAM). The resulting digitized data stream was recorded on magnetic tape. The flight parameter values presented in this report represent data averaged over one second (approximately seven consecutive rotor revolutions). The rotating frame signals were transferred to the nonrotating frame using 45 channels of a 50-channel slipring. The rotating and nonro- tating frame signals were then relayed to seven frequency domain multiplexers. All multiplexing was per- formed according to the Inter Range Instrumentation Group standard.
DATA REDUCTION
The calculation of additional flight condition parameters and the reduction of strain gage data are presented. Also, the procedure for determining hub loads from shaft loads measurements is described.
Calculated Rotor Flight Parameters
In addition to the parameters discussed in the Flight Condition Parameters section, the following values were calculated:
A = X R ~
CT = bchR
P =Po00
3
p =TAS/RR
CT/O = Nmg/(pA(RR)20)
Strain Gage and Potentiometer Signals
Strain gage signals were recorded over many rotor revolutions; however, only seven consecutive revolutions were reduced. These seven revolutions correspond temporally to the seven revolutions used to acquire the flight condition parameters. A harmonic analysis of the gage signals was performed for each of the seven revolutions. The results were then averaged. The strain gage signals were converted to harmonic analysis coefficients, ten per rotor revolution, which were then converted into physical units. Although the strain gage data were acquired from all three blades, all data were referenced to the blade 1 azimuth position, starting at 0" azimuth (rear blade position). The lag angle was calculated from the lag displacement (potentiometer signal) using the following relationship.
I
Time histories of the strain gage and Potentiometer signals can be reconstructed using the following formula.
10
n X = A0 + 2 [An cos (nY) + Bn sin (nY)]
where Y = 0" corresponds to the rear blade position.
Determination of Hub Loads from Shaft Loads
The hub loads were calculated from measured shaft bending moments. Bending moments in the x- direction, mtAx and mtBx, were measured at distances 1~ and fg from the rotor center (fig. 14). The rotating hub forces and moments are represented in figure 14 by (ftx, fty) and (mtx, mty), respectively. The shaft is assumed to be a beam isolated from other external forces and moments. The relationships between the loads at the hub and the loads at sections A and B are:
I 4
fty = (ma, - mtAx)/(lA - 1 ~ ) (20)
The moments mtAx and mtBx are measured by the gage couples FlSHAJT and F3SHAFT, respectively. Through symmetry, lmtAx I = lmtAy I and ImtBx I = ImtBy I. Also, the phases of mtAy and mtBy lag mtAx and mtBx by 90’. Therefore, all terms required to solve for the total rotating forces and moments at the hub are known. For a rotor with b equally spaced blades, only the nb f 1 harmonics exist for the rotating frame in-plane hub loads (ftx, fty, mtx, mty), and only the nb harmonics exist for the out-of-plane loads (ftz, mtz). A coordinate transformation is performed to obtain the fixed frame hub loads, Fx, F y Mx, and M y (fig. 15). These loads can be expressed in terms of the rotating frame loads by
c
9
FX = ftx cos@!) + fty sin(Y) (21)
F y = -ftx sin(Y) + fty cos(Y) (22)
M x = mtx cos(”) + mtY sin(”)
My = -mtx sin(Y) + mtY cos(Y)
(23)
(24)
Mz =mtz
Through phase cancellation, only the nb harmonics exist for the fixed frame loads.
FLIGHT CONDITIONS
The flight envelope included advance ratios from 0.0078 to 0.371 and values of CT/O from 0.062 to 0.133 (fig. 16). A complete list of the 23 flight conditions is presented in Appendix A. This flight test attempted to duplicate selected flight conditions from the first test documented in reference 1. This was done by directing the pilot to an altitude such that a desired reduced-mass value was obtained. Once at the desired altitude, the aircraft forward speed was adjusted to obtain a desired true air speed (TAS).
DATA PRESENTATION ..
As shown in Appendix A, flight V31 consisted of 6 conditions, flight V32 consisted of 7 conditions, and flight V33 consisted of 10 conditions for a total of 23 flight conditions. A li’st of measured parameters and their sign conventions are provided in Appendix B. Flight V3 1 and V32 were performed with the blade chordwise c.g. located at 24.7%. The chordwise c.g. for flight V33 was located at 26.7%. Note that Appendix B also indicates which measurements were inoperable for the three flights.
5
Appendix C contains the following for each of the 23 flight conditions: a) flight test parameters, b) blade bending moments, and c) airframe loads.
Appendix D provides rotating and nonrotating hub loads derived from the measured shaft forces and moments. Only values for flight V31 are provided, since F3SHAFT and the shaft torque gage, CZSHAFT, were inoperable during flights V32 and V33.
The data were checked for consistency where possible. Redundant blade strain gage instrumentation provided blade-to-blade comparisons at 12%, 29%, and 87% radial stations. Also, pitch link loads, blade flap angle, and blade lag angle were redundantly measured. Discrepancies in mean values were noticed for the bending moments, pitch link loads, and lag angle for some of the flight conditions. However, the data are presented here without alteration.
.
Flight Test Parameters
A list of the environmental conditions and aircraft and rotor trim values precede the tabulated data for each of the flight conditions in Appendix C (for example, see Appendix C, page Cl). The average, mini- mum, and maximum values over seven consecutive rotor revolutions are presented for each parameter. Also, the standard deviation (STD.DEV.) is provided.
Rotor and Aircraft Measurements
The blade structural moments, flap and lag angles, lag damper force, pitch link loads, servo loads, shaft loads, gearbox strut loads, and seat accelerations are all presented in harmonic coefficient form. The coefficients represent the average value over seven consecutive rotor revolutions (10 harmonics per revo- lution). Note that redundant instrumentation is available for several of the rotor measurements (for exam- ple, blade flap angle).
Hub Loads Data
Appendix D presents hub force and hub moment data. Hub load data were measured for flight V31 only. The mean value of the vertical force, Fz, was not accurate because of signal drift. FZ is identical to the parameter FZSHAFT in Appendix C. For convenience, the 3/rev and 6/rev values of FZSHAFI' are repeated in Appendix D as Fz.
Appendix D provides the l/rev, 2/rev, .?./rev, Yrev, and 7/rev harmonic coefficient values for f b (total, rotating in-plane force). Recall from the section Determination of Hub Loads from Shaft Loads that I ftx I = I fty I ; therefore, only values for ftx are given. In addition, the hub forces in the fixed frame are given in terms of the mean, 3/rev and 6/rev for F x and Fy (in-plane forces) and in terms of 3/rev and 6/rev for Fz. As discussed earlier, ftz = Fz.
Similar information is provided for the hub moments in Appendix D. M i is identical to the shaft torque parameter, CZSHAFT, presented in Appendix C. Again, the mean, 3/rev, and 6/rev values of CZSHAFT are repeated in Appendix D as MZ for convenience.
6
~ REFERENCES
1. Heffernan, R.; and Gaubert, M.: Structural and Aerodynamic Loads and Performance Measurements of an SA349/2 Helicopter With an Advanced Geometry Rotor. NASA TM-88370, 1986.
I 2. Johnson, W.: A Comprehensive Analytical Model of Rotorcraft Aerodynamics and Dynamics. Part 1: ~
Analysis Development. NASA TM-81182, 1980.
8
. .
7
TABLE 1.- SA349/02 HELICOPTER DATA
Design aircraft gross weight (kg): 2000.0 387.0
59 19.0 Design main-rotor rotational speed (rpm): Design tail-rotor rotational speed (rpm):
Estimated aircraft moments of inertia (kg - m2): (for reference axis: see fig. 1)
Ixx = 800.0 SWASHPLATE FORWARD
LEAD ANGLE \ Iyy = 4200.0 Izz = 3600.0
Ixz = 680.0 / I'
FIXED SWASHPLATE
Ixy = 0.0
Iyz = 0.0
Shaft angle of attack (degrees): (Positive rearward) Main rotor -4.0 Tail rotor 0.0
Shaft cant angle (degrees): Main rotor 0.0 Tail rotor 0.0
Horizontal tail cant angle (degrees): - 1 .O (Relative to floor reference, positive
rearward, see fig. 1)
Vertical tail cant angle (degrees): 0.0
Aircraft CG location (m): (for reference axis: see fig. 1) FS (Fuselage station) X = -0.09 BL (Butt line) Y = 0.0 WL (Waterline) 2 = -1.32
Tail rotor hub location (m): (for reference axis: see fig. 1) FS X = 5.86 BL Y = 0.0 WL 2 = -0.83
Horizontal tail aerodynamic center (m): FS X = 4.82 BL Y = 0.0 WL 2 = -1.10
(for reference axis: see fig. 1)
Vertical tail center of action (m): (for reference axis: see fig. 1) FS X = 5.89 BL Y = 0.0 WL Z. = -0.13
Fixed swashplate azimuth lead angle (degrees): 34.0
I '
8
TABLE 2.- ADVANCED GEOMETRY ROTOR DATA
Number of blades Airfoil Chord (m) Radius (m) Rotor solidity Design tip speed (dsec) Rotation direction, viewed from top Blade mass (kg) Blade flap mass (mass outboard of flap hinge) (kg) Blade mass moment (m - kg) (Reference: flapping axis) Blade polar inertia (m2 kg) (Reference: flapping axis)
= Ij(y2+z2)dm - - I,
Blade flap inertia (m2 - kg) (Reference: flapping axis) Lock number Flap hinge offset (m) Lag hinge offset (m) Distance between lag hinge and lag damper line of action (m) Control system stiffness (N/m) (not measured) Pitch bearing radial location (m) Pitch horn arm length (m) Pitch link length (m) Pitch horn cant angle (degrees) Pitch link cant angle (degrees) Precone angle (degrees) Droop angle (degrees) Sweep angle (degrees) Feathering axis droop angle (degrees) Feathering axis sweep angle (degrees)
3 0a209
0.35 5.25 0.06366
212.76 clockwise
39.64 74.67
112.69
0.367
0.286
375.59 4.5417 0.1 1 0.475 0.177 3.0e4 0.250 0.170 0.326 2.0 7.05 0.0 0.0 0.0 0.0 0.0
9
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W n
z 0 E t, w m
n r
c
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m F m o O m d O F o 0 + o o o + m m m + E 0 0 0 0 0 0 0 0 . ? ~ S S S S ~ S S 5 0 0 0 0 0 0 0 d 0 0 0 0 0 0 0 0 0 0 . 0 0 0 0
10
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9 d
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12
Radial station, m
0.445 0.475- 0.475+
0.505+ 0.525-
F 0.525+ 0.545- 0.545+ 0.565- 0.565+
0.585+
0.625+ 0.665 0.705
0.745+
0.795+ 0.815 0.885 0.955 1.025 1.095
1.165+
1.305+ 1.375 1.515 1.605 1.635
3.276+
3.396+
3.590+' 3.634- 3.634+ 3.754-
0.505-
0.585-
0.625-
0.745-
0.795-
1.165-
1.235 1.305-
3.276- *
3.396-
3.590-
TABLE 4.- SECTIONAL BLADE PROPERTIES CG = 26.7%
M, kg/m
25. 25. 25. 25. 15. 15.
105. 105. 100. 100. 38.648 37.848 16.148 13.648 8.548 8.148 7.748 7.598
10.098 9.648 7.148 6.802 6.247 6.819 6.726 6.747 6.886 7.238 6.954 6.466 6.149 5.721 5.686 5.241 5.146 5.146 6.67 1 6.67 1 4.895
. 4.895 5'.0 15 5.015
' 6.675 6.675
Itp, kg-m2/m
0.479 0.4336 0.4336 0.3572 0.1952 0.1228 0.2828 0.295 1 0.293 1 0.2753 0.1745 0.1679 0.2347 0.1986 0.0141 0.0107 0.0085 0.0079 0.0079 0.0077 0.0077 0.0086 0.0099 0.0161 0.0270 0.0432 0.0600 0.0612 0.0723 0.0738 0.0708 0.0656 0.0594 0.0454 0.0385 0.0385 0.0450 0.0450 0.0377 0.0377 0.0382
' 0.0382 0.0456 0.0456
13
Itpp, kg-m2/m
0.3 104 0.285 0.285 0.2526 0.1626 0.0751 0.1311 0.1393 0.1313 0.1 186 0.1169 0.1115 0.1985 0.1725 0.0072 0.0044 0.0026 0.0023 0.0023 0.0028 0.0028 0.0033 0.0053 0.01 18 0.0238 0.0409 0.0583 0.0595 0.07 10 0.0727 0.0697 0.0646 0.0585 0.0445 0.0376 0.0376 0.0440 0.0440 0.0367 0.0367 0.0373 0.0373 '
0.0445 0.0445
YG, m
0. 0. 0. 0. 0. 0. 0. 0. 0. 0. 0. 0. 0. 0. 0. 0. 0. 0. 0. 0. 0.
-0.0017 -0.0021 -0.0070 -0.0148 -0.0245 -0.03 14 -0.0298 -0.0357 -0.0408 -0.0386 -0.03 66 -0.0264 -0.0133 -0.0070 -0.0070 0.0102 0.0102
-0.0056 -0.0056 -0.0038 -0.0038 0.01 12 0.01 12
Radial station, m
3.754+ 4.173- 4.173+
4.350+
4.390+
4.410+
4.440+
4.490+
4.610+
4.623+
4.640+
4.665+
4.705+
4.770+
4.800+
4.855+
4.875+
4.890+
4.950+
4.990+
5.030+
5.070+
5.150+ 5.250
4.350-
4.390-
4.410-
4.440-
4.490-
4.610-
4.623-
4.640-
4.665-
4.705-
4.770-
4.800-
4.855-
4.875-
4.890-
4.950-
4.990-
5.030-
. 5.070-
5.150-
M, kg/m
5.176 5.176 6.185 6.185 6.955 6.955 7.272 7.186 7.751 7.68 1 8.236 8.126
13.664 13.664 12.961 12.961 13.516 13.516 18.763 18.763 19.859 19.859 19.686 19.686 17.742 17.742 19.596 19.596 6.108 6.108
11.547 11.547 5.328 5.328 5.48 1 5.48 1 6.053 6.053 4.199 4.199 3.843 3.843 3.690 3.690
TABLE 4.- Concluded.
Itp, kg-m2/m
0.039 1 0.039 1 0.0454 0.0454 0.0478 0.0478 0.0549 0.0550 0.0568 0.0567 0.0666 0.0665 0.0740 0.0740 0.0640 0.0640 0.0740 0.0740 0.0801 0.0801 0.07 1 1 0.07 11 0.08 16 0.08 16 0.0864 0.0864 0.1385 0.1385 0.0582 0.0582 0.06 1 1 0.061 1 0.0539 0.0539 0.0539 0.0539 0.0541 0.054 1 0.0507 0.0507 0.0457 0.0457 0.0457 0.0457
14
Itpp, kg-m2/m
0.0381 0.038 1 0.0444 0.0444 0.0467 0.0467 0.0537 0.0539 0.0555 0.0554 0.065 1 0.065 1 0.07 19 0.07 19 0.0621 0.0621 0.07 19 0.07 19 0.0777 0.0777 0.0693 0.0693 0.0799 0.0799 0.0840 0.0840 0.1362 0.1362 0.0573 0.0573 0.0601 0.0601 0.0530 0.0530 0.0530 0.0530 0.0532 0.0532 0.0499 0.0499 0.0449 0.0449 0.0449 0.0449
YG, m
-0.0048 -0.0048 0.0093 0.0093 0.0145 0.0145 0.0082 0.0082 0.0055 0.0052
-0.0010 -0.0008 0.0022 0.0022 0.0073 0.0073 0.0035 0.0035
-0.0074 -0.0074 -0.0015 -0.0015 -0.0070 -0.0070 -0.0172 -0.0172 -0.02 15 -0.0215 0.0054 0.0054 0.0026 0.0026
-0.0 188 -0.01 88 -0.0160 -0.0 160 -0.0 150 -0.0150 -0.0 141 -0.0141 -0.0088 '
-0.0088 -0.0 125 -0.0125
-* -
FLOOR REFERENCE
\
Figure 1.- Aircraft reference axis systems. (a) Aircraft c.g. axis system. (b) Rotor axis system. (c) Wind axis system.
15
LIFT LATERAL FORCE A
Figure 2.- Aircraft force and moment sign conventions.
16
TOP VIEW-HUB PLANE A
1 -
I
ROTAT I ON FLAP PITCH LAG AX IS HINGE BEARING HINGE
BLADE 1/4- I I I CHORD AXIS
I I 0 . 1 1 0 y I
I A I PIT(
I HOF 3N 2 . 4 I
I
0 . 2 5 0 4 i I
I 0.475 ; I J I i
A
AS VIEWED FROM A' - A'
I I L+ 0.210 A, I CILHUC I ' 0.170
I I- PITCH HORN 7 I I I f I -
HUE PLANE
I h ' I r I - . . - - . -. I I 2mo- 1/4 CHORD STATION (COL L EC :TIVE =
I I I I
DIMENSIONS IN METERS
0" 1
Figure 3.- Control system geometry.
17
E E ln ln d 0 m 2 r
/ p_p____-___----p
FEATHERING AXIS (25% CHORD)
-i E ln
/ CONSTANT OA209 AIRFOIL I
E cq d
2
REMOVABLE TIP
A i REMOVABLE TIP - -7.392~~1,
-1.28' T- - \ \ *
I E ' " E I E
21 L 5
bl
I
I-
1 I I I I I -7 I
100 40 50 60 70 75 80 90 10 20 25 30 RADIAL LOCATION, % (REF: ROTOR SHAFT)
Figure 4.- Grande Vitesse blade planform and twist distribution.
YN > 0 ZN > 0 YG > 0
FORWARD OF 25% CHORD TOWARD UPPER SURFACE FORWARD OF 25% CHORD
+--- - I I \ \
.. v w i\ \ 25%CHORD
' 'N . REFERENCE AXES
Figure 5.- Grande Vitesse blade sectional axis systems.
18
8
50
fl 40 >- 0
3 0 w u.
E 30
20 40
2 30 >- 0 z w 3 U W a 20 u.
10
(a)
CHORDWISE C.G. LOCATION, % GIREV
5lREV
4lREV
1
r 7IREV I GIREV v
20
2 >- 3 10 0
3 0 W a u.
0 100 200 300 400 ROTOR ROTATIONAL SPEED, rpm
(a) Rotating blade modal frequency vs rpm: CAMFWD calculated.
Figure 6.- Grande Vitesse rotating blade properties.
19
7lREV 6IREV
40
35
30
25
2 5 5 20
8 w K L
15
10
5
0
I
CHORDWISE C.G. LOCATION, % A 24.7 0 26.7
.
100 200 300 400 ROTOR ROTATIONAL SPEED, rpm
(b) Rotating blade modal frequency vs rpm: IAK40 calculated.
Figure 6.- Concluded.
20
c
.45
-40
.35
.30
.25 w
0
a (3 .20 Z n
0
- 2
5 -
.15
-
-
-
-
-
-
-
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.05
0 -
-.05
K W n 01
-
-
I I I 1 I I I I I I J
1F
2F
1L 1T
2L
3F
Figure 7.- Blade damping ratio vs rprn: IAK40 calculated.
21
40
35
30
25
2 >- 5 20
E
0
3
K LL
15
10
5
0
CHORDWISE C.G. LOCATION, % A 24.7 0 26.7
rn - 0 2L
A d
3F
A B 1F I
Q 1L
I I I I I 1 I 1
2 4 6 8 10 12 14 16 BLADE PITCH, deg
Figure 8.- Blade modal frequency vs blade pitch: IAK40 calculated (at 387 rpm).
22
F: FLAP BENDING L: LAG BENDING T: TORSION
.12 .20 -29 -37 .46 .54 .63 .71 .80 .85.87
CONSTANT OA209 AI R FOI L
1 I I I I I I I I I
I I
I
0 .1 .2 .3 .4 .5 .6 r/R
Figure 9.- Blade instrumentation.
.7 .8 .9 1 .o
23
U 0
v)
w a
H z L -
24
HUB
F3SHAFT I I t
t I
I
I FlSHAFT, FSSHAFT: FZSHAFT: CZSHAFT:
I -
V I
I I I 0.314 m t l l I
.331
CZSHAFT FZSHAFT
0.394
I
m
BENDING THRUST TORQUE
25
26
z u 0 L
-+ I
I t
L
1
O N l l V l O U D N l l V l O U DN l lV1OU -NON -NON
27
.
Figure 14.- Hub and shaft loads reference system.
28
$ = 180"
mtx
tll = 0"
NONROTATING LOADS: Fx, Fy, Mx, My ROTATING LOADS: fw, fty. mtx' mty
Figure 15.- Relationship between rotating and nonrotating frame hub loads.
29
.14
.13
-12
.11
.09
.08
.07
FLIGHT C.G., % 0 V31 24.7 0 V32 24.7 A v33 26.7
A
a
0
A 0
a O h
nh @ I I 1 1 1 1 1 1 1
-08 .16 .24 .32 .40 .06 0
ADVANCE RATIO
Figure 16.- Flight envelope.
30
APPENDIX A
FLIGHT CONDITIONS
The following list summarizes the flight test conditions presented in this report.
Flight
V3101 V3103 V3105 V3106 V3 109 V3111
V3202 V3204 V3207 V3208 V3209 V3211 V3218
V330 1 V3303 V3304 V3305 V3308 V3310 V3312 V33 13 V3314 V33 15
Condition
LEVEL FLIGHT LEVEL FLIGHT LEVEL FLIGHT LEVEL FLIGHT STEADY TURN STEADY TURN
STEADY TURN STEADY TURN LEVEL FLIGHT LEVEL FLIGHT LEVEL FLIGHT STEADY TURN HOVER O.G.E.
LEVEL FLIGHT LEVEL FLIGHT LEVEL FLIGHT LEVEL FLIGHT STEADY TURN STEADY TURN STEADY TURN STEADY TURN STEADY TURN HOVER O.G.E.
TAS ( d s )
30 56 74 79 51 55
74 75 51 69 72 70
2
32 59 73 76 51 53 71 75 70
3
Load Factor
1.012 1.023 0.964 1.006 1.474 1.937
1.761 1.533 0.99 1 1.013 0.989 1.480 1.007
1.014 0.978 0.986 0.999 1.524 2.1 15 1.475 1.587 1.855 1.016
cT/O
0.06615 0.06622 0.06277 0.06552 0.09619 0.125 10
0.1 1390 0.10020 0.06470 0.08708 0.085 13 0.12400 0.06182
0.66280 0.063 18 0.06426 0.06592 0.096 17 0.13300 0.09355 0.10320 0.1 1300 0.06225
c1
0.1422 0.2647 0.3498 0.3712 0.2418 0.2602
0.3459 0.3509 0.2406 0.3250 0.3395 0.33 10 0.0078
0.1498 0.2752 0.3424 0.3614 0.2403 0.2479 0.3354 0.3543 0.33 10 0.01 19
31
APPENDIX B
MEASURED PARAMETERS AND SIGN CONVENTIONS
Appendix B lists all measured flight test parameters for the 23 flight conditions. Given for each parameter is the parameter name as used in Appendix C, the physical description, the units, and the sign convention.
Parameter Name Description
FLAP BEND flap bending moment 12% R; blades 1,3 20% R; blade 2 29% R; blades 1,3 37% R blade 2 54% R blade 2 63% R blade 1 a 7 1 % R; blade 2 80% R; blade 1 b 85% R; blade 2
EDGE BEND edgewise bending moment 12% R; blades 1,3 20% R; blade 2 29% R; blade 1 37% R; blade 2 46% R blade 1 54% R blade 2 63% R blade 1 71% R; blade 2 80% R blade 1 85% R; blade 2
TORSION
FLAP
LAG *
torsion moment 12% R blades 1,3 20% R blade 2 29% R blades 1,3 54% R blade 2 80% R; blade 2 87% R; blades 1,3
flap angle blade 1 blade 2
lag angle blade' 1 blade 2
33 PRECEDING PAGE BLANK NOT FILMED
Units
N-m
N-m
N-m
deg
deg
Sign Convention
positive moment created by downward force on blade at tip
positive moment created by force at blade tip directed from trailing to leading edge
positive moment created by blade nose-up deflection at tip
positive angle created by blade flapping upwards
positive, deflection in the direction against rotation .
FLAGDAMP
PTCH LNK LD
SERVO
GEAR BOX STRUT
VERT ACCEL
FlSHAFT
F2SHAFT
F3SHAFTc
FZSHAFT
CZSHAFTd
lag damper force N
pitch link load blade 1 blade 2 blade 3
servos left right
N
N
gear box strut load N rear left rear right front left front right
seat acceleration g front right seat front left seat
longitudinal shaft bending N-m moment
lateral shaft bending moment N-m
longitudinal shaft bending N-m moment
shaft vertical force N
shaft torque N-m
positive, lag damper in tension
positive, pitch link in tension
positive, upwards force
positive, strut compressed positive, strut compressed positive, strut in tension positive, strut in tension
positive, bending toward front of aircraft
positive, bending toward left side of aircraft (facing forward)
positive, bending toward front of aircraft
positive, upwards force
a not functioning for conditions V3101, V3103 b not functioning for V3109, V32 (all conditions), and V33 (all conditions) C not functioning for V32 (all conditions) and V33 (all conditions) d not functioning for V32 (all conditions) and V33 (all conditions)
positive, counterclockwise as viewed from above
34
APPENDIX C
SA349/2 HELICOPTER FLIGHT TEST DATA
Flight test data for each of the 23 flight conditions are listed. A list of flight parameters precedes the data for each condition.
35
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APPENDIX D
FLIGHT TEST DATA - HUB LOADS
The measured hub forces and moments for six flight conditions, V3101, V3103, V3105, V3106, V3109, and V3111 are presented for both the rotating and nonrotating frame.
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I .
1 Report No I 2 Government Accession No
im 3 Recipient's Catalog No
Report Documentation Page
7. Author(sl 8. Performing Organization Report No.
NASA "l-101040 5. Report Date 4. Title and Subtitle
Gloria K. Yamauchi, Ruth M. Heffeman, and Michel Gaubert (Aerospatiale Division Helicopteres, Marignane, France)
9. Performing Organization Name and Address
Ames Research Center Moffett Field, CA 94035
12. Sponsoring Agency Name and Address
National Aeronautics and Space Administration Washington, DC 20546-0001
Hub and Blade Structural Loads Measurements of an SA349/2 Helicopter
A-88300 10. Work Unit No.
505-61 -5 1 11. Contract or Grant No.
13. Type of Report and Period Covered
Technical Memorandum 14. Sponsoring Agency Code
December 1988
19. Security Classif. (of this report) 120. Security Classif. (of this page) 21. No. of pages 22. Price
Unclassified
18. Distribution Statement 17. Key Words (Suggested by Authorls))
Unclassified 133 A07
Helicopter flight-test data Rotor dynamics Hub loads
Unclassified-Unlimited
Subject Category - 05