Ludi, L. H.

English

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SUMMARY OF ROTOR-BIADE VIBRATORY-LOAD STUDIES
By LeRoy H. Ludl
Langley Research Center
The rotor-powered aircraft is directly associated with fatigue
since the rotor is subjected to alternating aerodynamic loadings during
all flight regimes. These alternating loadings on the rotor can pro-
duce periodic loads on the various components of the helicopter vhich
could severely limit the service life of these components because of
fatigue. In order to enable prediction of a satisfactory service life
for these components, it is necessary to know which conditions result
in the most severe periodic loads so that they may be investigated
during prototype testing. Examples of conditions which require investi-
gation are as follows:
(1) Level flight throughout speed range
(2) Retreating-blade stall
(3) Landing approaches
(4) Partial-power descents
(9) Droop-stop pounding on the ground
(6) Atmospheric turbulence
(7) Moderate maneuvers
(8) Transition
(9) High-speed level turns
(lO) Pull-outs from autorotatlon
(ll) Autorotation at high forward speed with high rotor speed
The loads in these conditions in combination with the amount of time
spent in the conditions are the basic factors in a rational determina-
tion of the fatigue life of a structure. In order to determine experi-
mentally the information on the relative severity of the periodic moments
encountered by a rotor blade during the conditions listed previously, a
program utilizing the helicopter shown in figure l, equipped with strain
gages at 14 percent and 40 percent of the radius on one of the blades,
was undertaken. These locations were chosen to give a maximum amount of
information on the blade moments and were considered adequate for this
investigation even though stress surveys usually involve more locations.
The results are of general interest in that the flight conditions, which
resulted in the most severe rotor-blade loads are defined. The condi-
tions numbered 1 to 5 in the previous list were found to produce the
more severe blade loads and will be discussed in more detail. Additional
information on all the conditions listed previously can be obtained from
references 1 to 4.
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Before going into the results of the rotor-blade vibratory-load
studies, it is interesting to see how much time a helicopter actually
spends in its various operating airspeed ranges under continued opera-
tional use as shown in figure 2. The information in this figure is a
result of a continuing study of operating experiences being obtained
from both civil and military operations. Published information on
surveys of helicopter operating conditions is contained in refer-
ences 5 to 8. Figure 2 shows the percent of total time spent in the
various speed regimes as a function of the ratio of forward velocity V
to the maximum forward velocity VMAX" The maximum forward velocity is
determined from the pilot's handbook for the particular configuration.
The distribution represents the complete flight profile of climb,
en route, and descent for a helicopter in both a civil application as
shown by the solid line and a military application as shown by the
dashed line. As can be seen from the figure, the airmail operation
tends to concentrate the major percentage of total time at speeds
beyond 65 percent of VMA X. The military operation, in this case
pilot training, tends to shift the major percentage of total time
toward the lower speeds. In fact, the military operation did not
spend any time above 87 percent of the maximum speed. These results
show that the entire speed spectrum must be checked for large moments,
particularly when the helicopter has a dual mission. The high-speed
end of the spectrum, where the large percentage of time spent and the
expected increased moments are combined, is extremely important in
fatigue-life calculations.
In order to illustrate how the rotor-blade vibratory moments are
affected by forward speed, figure _ shows the variation of measured
vibratory moments as a function of the ratio of forward velocity to
maximum forward velocity. The measured vibratory moments at any
speed M are divided by the measured vibratory moments at cruise
speed MCRUISE. Cruise speed for this helicopter is 60 percent of its
maximum velocity, and the vlbratorymoments at cruise speed are used
as a reference for most of the following figures. The moments shown
in figure 3 are flapwise bending moments at 40 percent and 14 percent
of the blade radius and torsional moments at 14 percent of the blade
radius. The torsional moments on the blade, while not as critical
as bending moments in determining the fatigue life of the blade itself,
are of prime importance in control-system design. In figures 3 to 8,
amplitude and revolution or cycle are defined as shown in the small
inset in figure 3. It can be seen that the trends for all three
moments are similar. The buildup in moment value at 25 percent of
the maximum forward speed is caused by the transition region, which is
the region between hovering and the speed for minimum power and is
characterized by a change in airflow through the rotor. As the speed
increases above 70 percent of VMAX, retreating-blade stall becomes
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more responsible for the increased moments. At 95 percent of VMAX,
the highest speed achieved during this investigation, the flapwise
bending momentsat 40 percent of the blade radius had increased to
twice the cruise value while the torsional momentsat l_ percent of
the blade radius had increased to 3_ times the cruise value. The trends
indicate that the use of turbine engines and the resultant higher ratio
of cruise speed to maximumforward speed will produce very high moments
that must be accounted for during fatigue-life calculations.
An additional problem which arises as a result of high-speed
flight is increased periodic control loads in high-performance proto-
type helicopters. The increased periodic control loads, a result of
increased torsional momentscaused by retreating-blade stall, have
seriously restricted the normal operating limits of these prototypes.
The effect of stall on the torsional momentsis illustrated in fig-
ure 4. In this figure, the torsional momentsat the various forward
speeds M are divided by the torsional momentat the speed which
first produced a retreating tip angle of 12° MSTALL and plotted
against increasing forward speed. The open symbols represent unstalled
conditions, and the solid symbols represent stalled conditions. It
can be seen that the momentsincrease at a fairly shallow rate until
the retreating tip angle of attack exceeds the stall angle; then, they
increase rapidly. Moderate penetration into stall results in ratios
that are still fairly small while extreme penetration results in ratios
of almost 3. These increased vibratory torsional momentsillustrate
the fact that careful consideration must also be g_ven to control-
system loads during high-speed flight in order to avoid operating
limitations.
At the other end of the speed spectrum, there are flight condi-
tions which were found to produce severe periodic rotor-blade moments
that would be of interest in the design of the various rotorcraft com-
ponents. Landing approaches and partial-power descents at zero or low
forward speeds resulted in momentsthat were the highest encountered
i during the investigation. The maximumvibratory momentsexperienced
by the rotor blade during a landing approach are shownin figure _. In
_i this figur , the inb ard blade oments M are divided by the cruise
moment MCRUISEand plotted as a function of the forward speedratio
during the approach. Since the results obtained during the landing-
approach tests are not readily repeatable, the momentsin this figure
represent the maximumvalues obtained during the investigation. Since
there is scatter in this type of maneuver, the amount of scatter in the
area of most interest is indicated by the shaded area and is discussed
" in detail later. In general, figure 5 shows that flapwise bending
momentsas high as 5.75 times the cruise momentsare encountered during
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the approach. These increased moments in the landing approach are
believed to be caused by a change in the airflow similar to that which
occurs during the transition region.
The effect of partial-power descents on the maximum vibratory moments
encountered by a rotor blade is shown in figure 6. In this figure, the
ratio of the inboard flapwlse moments is shown as a function of the rate
of descent, for two values of forward speed (0 and 28 knots). Since the
moments at a particular rate of descent also are not readily repe_atable,
the moments plotted in the figure represent the maximum moments encoun-
tered in the various descents. As in figure 5, the amount of scatter
during the partial-power vertical descents is sho_nby the shaded area.
Figure 6 shows that inboard moments 6.8 times the cruise moments are
obtained during partial-power descents at zero forward speed. Because
of the random character of the flow during the vertical descents, con-
trol.was generally poor and required large movements of the controls to
maintain steady conditions. As forward speed is increased to 28 knots,
the moments peak at a lower rate of descent and reach a lower numerical
value than those encountered in vertical descents. At the higher for-
ward speed, control is generally improved and smaller control movements
are needed to hold steady conditions. The partial-power vertical descents
were found to produce the highest vibratory moments of all the conditions
investigated.
Since the moments encountered during partlal-power descents have a
random character, the degree of conservatism involved in using the maxi-
mum moments during a partial-power vertical descent, with the assumption
that the maximum moments occurred during the entire time for a fatigue-
llfe determination, was considered. An indication of the random nature
of the moments encountered during a partlal-power vertical descent is
given by the distribution shown in figure 7. In this figure, the ratio
of the inboard bending moments is plotted as a function of the percent
of total cycles in the various bending-moment ranges. Figure 7 shows
that the maximum moments shown in figure 6 occur only during 1 per-
cent of the cycles and, therefore, might be too conservative if used
exclusively. Furthermore, other conditions such as hlgh-speed flight,
where the maximum moments might be above the endurance limit and occur
during a larger percentage of time than the maximums in partlal=power
descents, must not be neglected in any rotor-blade finite llfe
assessment.
In addition to the flight conditions investigated, static droop-
stop pounding, where the blade impinges on a mechanical stop due to
flapping during rotor operation on the ground, also produced some addi-
tional moments which were of interest. Static droop-stop po_n_ding, even
though artificially produced for these tests, can inadvertently occur
during rotor operation on the ground in strong, gusty winds or in ground
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taxiing. In fact, some cases of droop-stop pounding in the air have
been reported. The potential ability to produce droop-stop pounding
moments that are higher than maximum flight moments, if the pounding is
allowed to progress, is illustrated in figure 8. Here the inboard flap-
wise moments during droop-stop pounding M are divided by the maximum
flight moments MFLIGRT,MAX (in this case, those encountered in partial-
power vertical descent) and plotted as a function of the change in longi-
tudinal cyclic pitch &B1, s. The cyclic pitch at which droop-stop
pounding first occurs is used as a reference. In general, the figure
shows that the moments increase linearly with cyclic control; therefore,
the moments would also increase with increasing flapping angle. This
trend indicates that if the design and the operating circumstances should
produce large down flapping, such as fuselage clearance often allows,
very high moments will result. Even though the moments during droop-
stop pounding occur at a low frequency, they can become very important
from a fatigue standpoint if the magnitude of these moments follows the
trend shown in the figure. The moments in the figure are felt to be an
indication of the rate of increase of moment with control displacement
if droop-stop pounding is inadvertently permitted, at least for a blade
with uniform spar and uniform mass.
Utilization of the moments illustrated in the figures and the
moments encountered during the other flight conditions investigated, in
conjunction with the results of helicopter operating surveys which Illus-
trate the amount of time spent in various flight conditions, permitted
rough sample fatigue-life calculations to be made for the rotor blade.
One calculation, utilizing Just two of the flight conditions that pro-
duced the highest moments, was made. The fact thet these moments occur
during the entire time in this flight condition was assumed. The blade
fatigue life under these conditions was approximately 775 hot_s. A sec-
ond calculation, utilizing all the flight conditions that produced
moments above the endurance limit and more accurate distributions of the
times spent during these flight conditions, was made. The blade fatigue
life under these conditions was approximately 1,510 hours. These resul_
illustrate the necessity for an accurate evaluation of the number of
flight conditions which result in the most severe periodic loads and of
the time spent in each condition to be used in a fatigue-life determina-
tion for the blade in order to provide a safe blade life without undue
penalties.
The results of the rotor-blade vibratory-load study indicate that
- severe rotor-blade loads can be expected to occur during high-speed
level flight, landing approaches, partial-power descents, and droop-stop
pounding during rotor operation on the ground. In addition, retreating-
blade stall can cause increased torsional moments which cause increased
control loads. In the future design of rotorcraft_ these severe rotor-
blade loads must be accurately accounted for in component fatigue-life
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determinations so that undue limitations will not seriously restrict the
operational usage of the rotorcraft.
REFERENCES
1. Ludl, LeRoy H.: Flight Investigation of Effects of Atmospheric
Turbulence and Moderate Maneuvers on Bending and Torsional Moments
Encountered by a Helicopter Rotor Blade. NACA TN 4203, 1958.
2. Ludl, LeRoy H.: Flight Investigation of Effects of Retreating-Blade
Stall on Bending and Torsional Moments Encountered by a Helicopter
Rotor Blade. NACA TN 4254, 1958.
3- Ludi, LaRoy H.: Flight Investigation of Effects of Transition,
Landing Approaches, Partial-Power Vertical Descents, and Droop-Stop
Pounding on the Bending and Torsional Moments Encountered by a
Helicopter Rotor Blade. NASA M_NO _-7-59L, 1959.
_. Ludi, LeRoy H.: Flight Investigation of Effects of Additional
Selected Operating Conditions on the Bending and Torsional Moments
Encountered by a Helicopter Rotor Blade. (Prospective NASA Paper.)
5- Crim, Almer D., and Hazen, Marlin E.: Normal Accelerations and
Operating Conditions Encountered by a Helicopter in Air-Mall Opera-
tions. NACA TN 2714, 1952.
6. Hazen, Marlin E.: A Study of Normal Accelerations and Operating Con-
ditions Experienced by Helicopters in Commercial and Military
Operations. NACA TN 3_34, 195}.
7. Connor, Andrew B., and Ludi_ LeRoy H.: A Summary of Operating Con-
ditions Experienced by Two Helicopters in a Commercial and a
Military Operation. NASA TN D-2_l, 1960.
8. Connor, Andrew B.: A Summary of Operating Conditions Experienced by
Three Military Helicopters and a Mountain-Based Commercial Heli-
copter. NASA TN D-_32, 1960.
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TEST HELICOPTER
Figure i
DISTRIBUTION OF OPERATING AIRSPEED
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% 40
OFTOTALTIME
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V/VMAx
---MILITARY
--AIRMAIL
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Figure 2
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EFFECT OF FORWARD SPEED ON VIBRATORY MOMENTS
M
MCRUISE
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AMPLITUDE _ l /7 • '
2- ,--., /,/_______
/ -.., /A,""
• I I L... I I I - I I I I
0 .l .2 .3 .4 .5 .6 .7 .8 .9 I.O
VIVMAX
TORSION
---- FLAPWISE
BENDI NO
Fi_e3
EFFECT OF RETREATING-BLADE STALL ON TORSIONAL MOMENTS
4
0 UNSTALLED
• STALLED
INCREASING V _-
Fi_e_
i • INBOARD FLAPWlSE MOMENTS DURING LANDING APPROACH
4
M 3
MCRUIS E
11
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0 .1 .2 .3 .4
V/VMA x
1
Figure 5
INBOARD FLAPWISE MOMENTS DURING PARTIAL-POWER DESCENTS
M
MCRUISE 3
2
v: <is
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400 800 1,200 1,60(I
RATE OF DESCENT. FT/MIN
Figure 6
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DISTRIBUTION OF INBOARD FLAPWlSE MOMENTS DURING
PARTIAL-POWER VERTICAL DESCENT
6.87
5.13_ _
M 4.26 _ _': _
2.52
0 2 4 6 8 10 12 14 16
CYCLES,% OFTOTAL
MCRUISE
Figure 7
INBOARD FLAPWlSE MOMENTS DURING DROOP-STOP POUNDING
1.4
1.2
M
MFUGHT,MAX i
0 1 2 3 4 5
ABI,s,DEG
F igure 8


Summary of rotor-blade vibratory-load studies

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