A flight flutter testing technique is described in which the rotor controls are oscillated by series actuators to excite the rotor and airframe modes of interest, which are then allowed to decay. The moving block technique is then used to determine the damped frequency and damping variation with rotor speed. The method proved useful for tracking the stability of relatively well damped modes. The results of recently completed flight tests of an experimental soft-in-plane rotor are used to illustrate the technique. Included is a discussion of the application of this technique to investigation of the propeller whirl flutter stability characteristics of the NASA/Army XV-15 VTOL tilt rotor research aircraft

Matthys, C. G.

Viswanathan, S.

Yen, J. G.

English

NASA Technical Reports Server

FLIGHT FLUTTER TESTING OF ROTARY WING AIRCRAFT 
USING A CONTROL SYSTEM OSCILLATION TECHNIQUE 
Jing G. Yen, Sathy Viswanathan, 
and Carl G. Matthys 
Bell Helicopter Company 
SUMMARY 
This paper describes a flight flutter testing technique in which the 
rotor controls are oscillated by series actuators to excite the rotor and air- 
frame modes of interest, which are then allowed to decay. The moving block 
technique (see ref. 1) is then used to determine the damped frequency and 
damping variation with rotor speed. The method has proved useful for tracking 
the stability of relatively well damped modes. 
pleted flight tests of an experimental soft-in-plane rotor are used to illus- 
trate the technique. This technique will also be used for flight flutter 
tests of the NASA/Army XV-15 Tilt Rotor Research Aircraft, to investigate its 
propeller whirl flutter stability characteristics, and this application is also 
discussed briefly . 
The results of recently com- 
INTRODUCTION 
A soft-in-plane rotor has recently been built and flight tested by Bell 
Helicopter Company. This type of rotor system has the potential for ground and 
air resonance instability. The potential mode of instability, the lead-lag 
motion of the blades coupled with the fuselage rigid body roll mode, has a fre- 
quency of 0.75 per rev (3.55 Hz) in the rotating system. 
to be well damped, since sufficient elastomeric lead-lag damping was employed. 
However, since it was the first rotor of this type developed by Bell Helicopter 
Company, an extensive ground/air resonance test program was conducted to inves- 
tigate the ground/air resonance phenomena. 
The mode was predicted 
The measurement of the stability characteristics of rotary wing aircraft is 
complicated by the need to excite modes in a rotating system and by the fact 
that both the rotor and the airframe are subject to steady state harmonic load- 
ing. 
damped mode. 
helicopters or tilt-rotor aircraft hence becomes a formidable task, both to 
cause the proper excitation and to reduce the data. A flight flutter testing 
technique has been developed, as described in this paper, to oscillate the rotor 
controls by means of the series actuators in a sense and a magnitude prescribed 
by the user. 
This loading tends to mask the transient response of any relatively well 
The'development of an on-line flight flutter testing technique for 
The authors acknowledge the assistance of Ken Builta of Electronics 
Development in the design of the rotor excitation system, and Al 'Eubanks of 
Scientific and Technical Computing in development of the analysis. 
https://ntrs.nasa.gov/search.jsp?R=19770014097 2020-03-22T11:20:39+00:00Z
SYMBOLS 
F(w) F o u r i e r  t ransform of f ( t )  
SCAS s t a b i l i t y  and c o n t r o l  augmentation system 
x(w) 
f ( t )  e x c i t a t i o n  s i g n a l  
Four i e r  t ransform of a response 
t t i m e  
n r o t a t i o n a l  speed of  r o t o r  
w frequency of a mode i n  a dynamics system 
WI frequency of t h e  mode of i n t e r e s t  i n  a dynamics system 
b lade  lead-lag n a t u r a l  f requency 5 w 
ROTOR EXCITATION SYSTEM 
For a so f t - in -p l ane  r o t o r  t es t  program, i t  i s  d e s i r a b l e  t o  excite t h e  r o t o r  
a t  a r e l a t i v e l y  low ampli tude and a t  a p resc r ibed  frequency.  The system 
developed t o  perform the  e x c i t a t i o n  f o r  f l i g h t  f l u t t e r  t e s t i n g  of  t h e  Bell Model 
609 so f t - in -p l ane  r o t o r  u s e s  t h e  SCAS a c t u a t o r s ,  d r i v i n g  them from an  HP203A 
dua l  s i g n a l  gene ra to r .  T h i s  r o t o r  exciter can supply e i t h e r  a cont inuous s i g n a l  
of a s i n g l e  f requency,  o r  a p u l s e  s p e c t r u m  which can be  s e l e c t e d  t o  excite only  
one n a t u r a l  frequency of  several which may be present. A f u n c t i o n a l  diagram of 
t h e  system i s  shown i n  f i g u r e  1. The upper p o r t i o n  shows the  system which d r i v e :  
t h e  r o t o r ,  and the lower p a r t  shows t h e  n u l l i n g  system. Th i s  n u l l i n g  system i s  
an op t ion  which al lows t h e  u s e r  t o  cancel t h e  s teady  s t a t e  one-per-rev and two- 
per - rev  s i g n a l s b e f o r e  t h e  i n p u t  spectrum i s  app l i ed ,  t hus  i s o l a t i n g  t h e  response 
t o  the  i n p u t  s i g n a l  only,  F igure  2 shows a diagram of the r o t o r  exciter. P o i n t  
A has  a p u l s e  output  t o  the computer t o  i n d i c a t e  when t h e  inpu t  has been com- 
p l e t e d .  P o i n t  B i s  the  demodulated one-per-rev t o  t h e  n u l l i n g  system. 
The under ly ing  p r i n c i p l e  behind t h e  single-mode e x c i t a t i o n  technique can be  
explained i n  terms of a l i n e a r  system w i t h  many degrees  of  freedom. The F o u r i e r  
t ransform of t h e  response v e c t o r  is  obta ined  by pos t -mul t ip ly ing  the matrix of 
t h e  frequency response f u n c t i o n  by the column of t h e  F o u r i e r  t ransform of the 
fo rc ing  func t ion .  (See, f o r  example, r e f .  2)  The frequency response  f u n c t i o n s  
e x h i b i t  peaks corresponding t o  n a t u r a l  f r equenc ie s  of t h e  system. 
The shape of t he  F o u r i e r  t ransform of the f o r c i n g  f u n c t i o n  i s  governed by 
the shape of t h e  fo rc ing  func t ion  i n  the time domain. For example, the 
Four ie r  t r ans fo rm of the f o r c i n g  f u n c t i o n  depic ted  i n  f i g u r e  3a is  shown i n  
f i g u r e  3b. F(y) has  a l a r g e  magnitude corresponding t o  frequency w1, and a 
r a t h e r  low magnitude elsewhere.  S ince  the Four i e r  t ransform of t h e  response,  
X(w), i s  ob ta ined  as a matrix product  of frequency response f u n c t i o n  and F(w), 
the response motions w i l l  b e  motions predominately a t  t h e  frequency w1. (This ,  
however, does n o t  mean that t h e  mode of in terest  a l o n e  i s  exc i t ed  and the  o t h e r  
502 
modes are properly suppressed.) 
chosen to produce a desired spectrum F(W), 
Various types of forcing function f(t) can be 
ON-LINE STABILITY ANALYSIS 
The Bell experimental Model 609 flex-hinge rotor, a four-bladed, soft-in- 
plane design, is mounted on a Bell 1501 fuselage. The elastomeric lead-lag 
hinge offset is 1 3 . 2  percent of the rotor radius which is 752 cm (296 in.). 
The blade first flapping frequency is 5.08 Hz ( 1 . 0 7  per rev). There are two 
potentially unstable modes for this experimental soft-in-plane rotor system. 
One, a coupled roll-yaw mode of the fuselage on the landing gear, has a 
frequency of 2 . 3  Hz. 
of 3 . 5 5  Hz at low strain conditions. 
The other, the blade first lead-lag mode, has a frequency 
Ground Resonance Test 
The roll mode stability was monitored by an accelerometer at the top of the 
mast. The test oscillated the SCAS actuators in a sense equivalent to moving 
the cyclic stick in a counter-clockwise direction at a frequency near 2 . 3  Hz. 
The aerodynamic hub moments and hub shears then excited the fuselage roll mode 
at its natural frequency. The high damping, the presence of high one-per-rev 
response, and some unknown noise made the data reduction difficult. Figure 4a 
shows the roll mode response at 260 rpm when a 7-Hz filter was used. 
block analysis of this response indicated the modal damping to be at 1 9 . 7  per- 
cent critical. Figure 4b shows the result of filtering the same raw data 
through a 3-Hz analog filter. In this case, the moving block technique showed ’ 
the modal damping to be at 20.8  percent critical. Since the 3-Hz filter brings 
out the highly damped roll mode more clearly, and since the calculated damping 
is approximately the same as in the other case, the decay plots of the roll mode 
in other ground run conditions were all filtered through 3-Hz filters before 
undergoing the moving block analysis. 
The moving 
The blade lead-lag motion was sensed by a strain gage on the grip damper 
arm. For the blade lead-lag stability test, the SCAS actuators were cycled as 
if the cyclic stick were being moved in a counter-clockwise direction at a fre- 
quency of one-per-rev minus the blade lead-lag frequency ( f 2 - 0 ~ ) .  This is equiv- 
alent to exciting the lead-lag mode aerodynamically at its natural frequency in 
the blade rotating system. Figure 5a shows the response of the lead-lag mode to 
a SCAS input at high collective (immediately before lift-off) and a rotor speed 
of 280 rpm. The modal damping obtained from the moving block technique was 4 . 3  
percent critical in the rotating system. The data were then passed through a 
4-Hz filter with the results shown in figure 5b. The modal damping determined 
by the moving block analysis was 4 . 2  percent critical in the rotating system. 
The filtering process again highlighted the modal decay and did not affect the 
moving block result. 
A summary of damping variations measured at various rotor speeds for the 
lead-lag mode and fuselage roll mode in the fixed reference system is shown in 
figure 6 .  
uted to the destabilizing Coriolis force and the characteristic of an elasto- 
meric damper, a phenomenon predicted by the ground resonance analysis. 
The decrease in damping with increase in blade collective is attrib- 
, 
503 
Air Resonance Analysis 
The blade lead-lag mode in hover was also excited with the SCAS actuators. 
Figure 7 shows the lead-lag decay at the design rotor speed of 285 rpm filtered 
through a 4-Hz filter, The moving block technique determined the modal damping 
to be 5.4 percent critical. 
XV-15 VTOL FLIGHT FLUTTER TEST PLANS 
- 
in September 1976. Bell Helicopter Company, the prime contractor, is building 
two of these aircraft which will have a design gross weight of 57.8 kN(13000 lb: 
and a maximum speed of 364 knots. Of primaiy interest from the standpoint of 
flutter are the stability of the coupled rotor/airframe system and flutter of 
the empennage. 
The NASAlArmy XV-15 Tilt Rotor Research Aircraft will enter flight testing 
Coupled Rotor/Airframe Stability 
The XV-15 has a nacelle on each wingtip. Each nacelle houses a T-53 engine 
and a transmission, and each transmission drives a 25-foot, gimbaled, stiff-in- 
plane three-bladed rotor. The nacelles are oriented with their shafts vertical 
for takeoff, landing, and flight in the helicopter mode, and are mechanically 
tilted 90 degrees for flight in the airplane mode. 
stability of the coupled rotor/airframe system, the wing is designed to be very 
stiff in torsion and in bending (it is 23% thick with spars at 5% and 50% chord, 
fully monocoque) and the nacelle is attached to the wing at the front and rear 
spar to make the attachment stiff in pitch and yaw. 
To prevent aeroelastic in- 
The calculated coupled rotor/airframe stability characteristics in the air- 
plane mode of flight indicate that instability occurs first in the wing chord- 
wise bending mode and, at higher speeds, in wing beamwise bending and in torsion 
These are shown in figure 8. The instability is similar in nature to propeller 
nacelle whirl flutter; but involves elastic bending of the blade’ and elastic 
deflections of the blade pitch control system in addition to the precession of 
the rotor disc. There is considerable confidence in the predicted stability 
characteristics, since the coupled rotor/airframe stability analysis has shown 
excellent agreement with flutter model tests and with tests of a full-scale semi 
span wing in the NASA Ames 40 x 80 foot wind tunnel (ref. 3). 
Empennage Flutter 
The XV-15 has an H-tail, a configuration that gives it good high-speed 
directional stability characteristics. Although flutter was of concern during 
the design, the empennage was designed to avoid resonance of the empennage modes 
with rotor excitation frequencies. For good frequency placement, the horizontal 
stabilizer is very stiff in bending and torsion. The elevator and rudder are 
powered by irreversible hydraulic actuators (dual for the elevator) with lock 
and load mechanisms to make them irreversible in the event of a hydraulic system 
failure. As a result, the empennage has a large flutter margin that has been 
confirmed by flutter model tests in the 16-foot transonic tunnel at NASA Langley 
(ref. 4 ) .  
504 
F1 ight Flutter Testing 
For tests to determine the frequency and damping of the coupled rotor/ 
The copilot or 
airframe modes, the research XV-15 will have series actuators in the wing 
flaperon and rotor blade collective pitch control linkages. 
flight test engineer will control the amplitude and frequency of the actuators, 
which have limited authority so that a hardover cannot cause excessive stresses 
or aircraft responses that the pilot cannot easily control. 
quency response is flat to well above the frequency of the highest coupled mode 
of interest. 
The actuator fre- 
Frequency and damping of each mode of interest will be determined from the 
decay of that mode. The test procedure will be to select either the flaperon or 
collective actuator and tune its excitation frequency to the modal frequency, 
then turn off (the actuators automatically center) and record the decay. This 
procedure was used in the full-scale test in the 40 x 80 foot wind tunnel, and 
gave good results. 
flight safety during flight enveiopi expansion. 
The decays will be monitored and analyzed on the ground for 
The excitation system will also be used to generate transfer functions by 
slow-sweeping excitation frequency. These transfer functions will be used for 
an additional check on the validity of the coupled rotor/airframe stability 
analysis. 
Tests to evaluate the empennage flutter characteristics will excite the 
empennage modes with doublet inputs to the elevator and rudders. 
generated through the series SCAS actuators. 
These will be 
CONCLUDING REMARKS 
1. For an elastomeric damper such as the one used on the Model 609 blade 
lead-lag hinge, the characteristics of the material depend on its strain. Since 
the lead-lag displacement (hence strain) varies with flight conditions, the lead- 
lag frequency varies throughout the flight test. It was learned from this study 
that, in the moving block analysis, a good estimate of the assumed frequency of 
computation could help the convergence. 
2. The dynamic and aerodynamic environment of a rotor change as the flight 
condition changes; hence the steady state one-per-rev and two-per-rev harmonic 
loads also vary. Therefore it is suggested that the nulling system be retuned 
whenever the flight condition changes. 
the Model 609 test because-the rotor synchro was not operational. 
However this option was not used during 
3. Because of the limitations on SCAS authority for the Model 609 testing, 
the signal-to-noise ratio for the modes of interest was relatively low. 
fore, a number of different analog filters were used to clean up the data. Low- 
pass filters from 3 Hz to 12 Hz, however, made no appreciable difference in 
damping calculations. 
There- 
505 
4. The single-mode excitation technique, using the SCAS actuators, produces 
excellent stability data since the input signal is well under the user's control. 
Depending on the input magnitude phasing and limitations of the SCAS authority, 
any mode in a rotary wing aircraft can be excited by the oscillation of the rotor 
controls in a prescribed manner. But whether the initial condition (mode shape) 
of the aeroelastic mode of interest is excited properly remains to be seen. The 
moving block analysis in most cases can be used in conjunction with the single- 
mode excitation technique to assess the stability information from ground run or 
flight test with real time computation. This testing and data reduction package 
is useful for on-line flight flutter testing of rotary wing aircraft. 
REFERENCES 
1. Anderson, William D.: Investigation of Reactionless Mode Stability Charac- 
teristics of a Stiff Inplane Hingeless Rotor System. 
the 29th Annual National Forum of the American Helicopter Society, May 
1973. 
Preprint No. 734, 
2 .  Lin, Y. K.: Probabilistic Theory - of Structural Dynamics. McGraw-Hill Book 
Company, 1967. 
3. Edenborough, Kipling H., Gaffey, Troy M., and Weiberg, James A.: Analysis 
and Tests Confirm Design of Proprotor Aircraft. 
August 1972. 
AIAA Paper No. 72-803, 
4 .  Marr, Roger L., and Neal, Gordon T.: Assessment of Model Testing of a Tilt- 
Proprotor VTOL Aircraft. The American Helicopter Society Symposium on 
Status of Testing and Modeling Techniques for V/STOL Aircraft, October 
1972. 
506 
F/A PILOT INPUT - - -9 
I 
I 
-- 
TRANSDUCER 
LATERAL PILOT INPUT- 
* T h i s  demodulator i s  physical ly  located i n  the rotor exciter. 
BUFFER 
, 1/mv 4 AMPLIiE'ATION 
PHASE AMPLIFIER 
* ADJUSTMENTS MAIN ROTOR 
SYNCHRO 4 DEMODU- - 
( l /REV SINUSOID) LATOR 
FREQUENCY 
DOUBLER To ADJUSTMENTS RECORDER 
Figure 1. Functional block diagram of rotor excitation system. 
I 
I 
I 
L W T  AND 
SYNCHRO EXCITATION 
I 
I 
I 
c 
218 VDC 
Figure 2. Diagram of rotor exciter. 
507 
(a) Plot of f(t). 
W / W l  
0 
(b)  Four i e r  transform of f(t). 
F i g u r e  3 .  Forc ing  f u n c t i o n  f ( t ) .  
508 
7-Hz FILTER 
AMPLITUDE 
TIME (SEC) 
(a) With 7-Hz f i l t e r .  
I 
0 .5 1.0 I. 5 2.0 2.5 3 . 0  3.5 
TIME (SEC) 
(b)  With 3-Hz f i l t e r .  
F igu re  4 .  Fuselage r o l l  mode decay i n  ground run. 52 = 260 rpm; 
l o w  c o l l e c t i v e .  
509 
AMPLITUDE 
AMPLITUDE 
NO FILTER 
.5 1.0 1.5 2.0 2.5 
TIME (SEC) 
(a) With no f i l t e r .  
.5 1.0 1.5 2.0 2.5 
TIME (SEC) 
( b )  With 4 - H x  f i l t e r .  
F i g u r e  5. Blade l ead - l ag  decay i n  ground run. 
high c o l l e c t i v e .  
fi = 280 rpm; 
510 
LOW COLLECTIVE 
HIGH COLLECTIVE 
0 
50 
40 
DAMPING 30 
(% CRITICAL) 0 
20 
10 
0 
2 40 260 280 30 0 320 
ROTOR SPEED ( R P M ) !  
(a) Blade lead-lag damping. 
50 
40 
DAMPING 
(% CRITICAL) 30 
20 
m 0 10 
0 
a LOW COLLECTIVE 
HIGH COLLECTIVE 
240 260 2 80 300 3 20 
ROTOR SPEED (WM) 
(b) Fuselage roll mode damping. 
Figure 6 ,  Damping in fixed system. 
511 
0 .5 1.0 1.5 2.0 2.5 
10 + 
TIME (SEC) 
- WING BEAM 
- -- - WING TORSION 
-.-.- W I N G  CHORD 
- - PYLON YAW 
F i g u r e  7 .  F i l t e r e d  blade lead- lag  decay i n  hover,  52 = 285 rpm. 
(% CRITICAL) 
0 
STABLE ---':.,.--. - Y - - - -  . '. ~- c --- -. \ -.- '. - -- 
I I I \  I i 
\ 
'"I 5 FREQUENCY (Hd 
O J  I 1 I I I 
0 100 200 300 400 500 600 700 
TRUE AIRSPEED (KNOTS) 
Figure  8. Calcula ted  s t a b i l i t y  for XV-15 a t  d e n s i t y  a l t i t u d e  
6096 m. (20000 f t ) .  
512 NASA-Langley, 1976 L-105 67 


Flight Flutter Testing of Rotary Wing Aircraft Using a Control System Oscillation Technique

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