Croman Corporation 

801 Ave C White City, OR 97503  (541) 826-4455  (541) 826-4730 FAX 

June 24, 2013 

Jeffery Lee 

Aviation Safety Engineer 

Boston Aircraft Certification Office 

12 New England Executive Park 

Burlington, MA 01803 

INTRODUCTION 

Croman Corp is an operator of S-61 helicopters. We have a history of 34 years of operation of the 

S-61 for logging operations involving REL cycles. In the years of operation, no catastrophic 

structural failures have occurred. We are supportive of any FAA mandates which enable us to 

maintain our safety record, but we take exception to the proposed life limits in the AD, as they will 

impose an unacceptable punitive cost (the current cost of a new MRS is $81K, not the $44K cited in 

the NPRM) which we strongly believe will result in the scrapping of serviceable MRS assemblies 

and is not justified. 

We believe the proposed retirement time of 30,000 REL cycles or 5000 hr TIS, whichever occurs 

first, for MRS which have been modified to the steel planetary plate and 200 ksi bolts is unrealistic 

and not supported by data for the following reasons: 

• Based on flawed interpretation of test data 

• Based on flawed assumptions of use by the operators 

• Ignores the restorative effects of maintenance practices recommend by the OEM 

• Ignores the service history of modified MRS which have accrued over 14 years of service 

without cracking without a restriction on REL cycles 

• Is inconsistent with the proposed retirement time of 30,000 REL cycles or 1500 hr TIS for 

MRS with original titanium planetary plates and 160 ksi bolts. 

SUBSTANTIATION 

The following is a summary of the reasons we believe the proposed retirement limits are 

unreasonable and not supported by data. Our comments are based on the results of the laboratory 
Jeffery Lee 

Aviation Safety Engineer 

Boston Aircraft Certification Office 

Page 2 of 9 

Croman Corporation 

801 Ave C White City, OR 97503  (541) 826-4455  (541) 826-4730 FAX 

testing presented at the 1998 S-61 operators conference, Attachment 1. In the attachment, Sikorsky 

presented the SN curve generated by fatigue testing two specimens under repetitive torque cycles, 

along with its then recommended life limit. We believe the conclusions drawn by Sikorsky 

regarding revised life limits are erroneous for the following reasons. 

• Flawed Interpretation of the Laboratory Test Data. The proposed life limit of 30,000 

REL cycles is based on a -3σ ”Non-Chafing” working SN curve generated from the test 

failure points. The premise for using a “non-chafing” SN curve instead of a “chafing” SN 

curve is that it leads to a lower number of cycles than the chafing SN curve. Using the nonchafing SN curve in lieu of the chafing SN curve is inconsistent with the fretting fatigue 

failures observed on the test specimens. Per Sikorsky report SER 61364, Methods of 

Service Life Analysis of Structural Components Substantiated by Laboratory Testing for 

FAA Certification of S-61L/S,S61R (CH-3) Helicopter, a characteristic SN curve for 

chafing is to be used for generating a component SN curve from test data if the failure mode 

is fretting. Attachment 1 from the 1998 operator’s conference shows that the test failures 

were indeed due to fretting. 

Normally, the use of a SN curve generated from test data using the characteristic SN curve 

shape for chafing results in a curve which predicts a lower allowable number of cycles than 

obtained using a non-chafing curve. This is illustrated by the notional SN curves for chafing 

and non-chafing shown below. 

High Cycle

Fatigue

Low Cycle

Fatigue

No Chafing

Chafing

N

S

Figure 1 Notional SN Curve 
Jeffery Lee 

Aviation Safety Engineer 

Boston Aircraft Certification Office 

Page 3 of 9 

Croman Corporation 

801 Ave C White City, OR 97503  (541) 826-4455  (541) 826-4730 FAX 

In the notional SN curve, the low cycle fatigue portion of the chafing and non-chafing 

curves are identical, as chafing has negligible effect on low cycle fatigue life. In the high 

cycle fatigue portion (generally defined as N>10,000 cycles), the chafing SN curve is lower 

than the non-chafing curve. The endurance limit for a part subjected to chafing should 

always be lower than the endurance limit for a part without chafing. If sufficient test data 

points exist, the chafing SN curve generated from the high cycle fatigue data points would 

invariably lead to a lower retirement time than the non-chafing SN curve for the same 

loading. The fact that a lower life is obtained from the non-chafing curve is indicative of 

insufficient data. 

Figure 2 shows mean and working SN curves for chafing and non-chafing generated from 

the test failure of specimen 1, which failed at 200,000 cycles at a vibratory torque amplitude 

of 29,515 ft-lb, which corresponds to REL cycle between 0%Q and 103%Q. The curves 

were generated using the method of SER 61364 , see Attachment 2. 

0

5000

10000

15000

20000

25000

30000

35000

40000

45000

1.00E+04 1.00E+05 1.00E+06 1.00E+07 1.00E+08

N~cycles

Vibratory Torque~ft-lb

Working-No 

Chafing

WorkingChafing

Test Failure Point Mean--No 

Chafing

Mean--Chafing

Figure 2. SN Curves Generated from Single Test Failure Point. 

Figure 2 shows that at a vibratory torque value of 29,515 ft-lb, the -3σ working life if 

fretting is assumed is close to 200,000 cycles while the working life if fretting does not 

occur is 35,000 cycles. From a structural mechanics point of view, the life with chafing 

should always be lower than the life without chafing, due to the role of fretting scars in 
Jeffery Lee 

Aviation Safety Engineer 

Boston Aircraft Certification Office 

Page 4 of 9 

Croman Corporation 

801 Ave C White City, OR 97503  (541) 826-4455  (541) 826-4730 FAX 

promoting fatigue cracking. This apparent anomaly is the result of the characteristic SN 

curve shapes of the Sikorsky methodology and of insufficient test data points. The SN 

curves of Figure 2 are generated from a single data point to illustrate the shortfalls of using 

insufficient data points. For multiple test failure points, SN curves as shown in Figure 2 

must be generated for each failure point. The mean SN curve is then obtained by averaging 

the imputed endurance limit for each curve. The working curve is then obtained, in the case 

of steel by regenerating a SN curve using a working endurance limit which is 70% of the 

mean endurance limit. If sufficient data points are available, the averaged SN curves for 

chafing and non-chafing will follow the trending of Figure 1, i.e., the chafing curve will 

always predict a lower life than the non-chafing curve for high cycle fatigue. 

Since only two data points were available, we believe that using a non-chafing SN curve for 

determining life simply because it led to a lower life is physically incorrect and a departure 

from the FAA approved fatigue methodology of SER 61364. SER 61364 specifies the use 

of a characteristic chafing curve for generating a SN curve from laboratory test data if the 

failure mode is fatigue abetted by fretting. We believe that the non-chafing SN curve should 

be disregarded, since its use is physically incorrect. Based on the chafing SN curve, we 

believe that a retirement time in excess of 100,000 REL cycles is appropriate for a REL 

cycle between 0%Q and 103%Q. 

Another flaw in the interpretation of the test data is the assumption that the MRS is 

subjected to continuous operation without scheduled maintenance to periodically check the 

MRS for damage and to remove damage before returning the MRS to service. The MRS is 

inspected every 1100 hr. as part of the overhaul of the gearbox. At that time, the gearbox is 

disassembled, the MRS is disassembled from the planetary plate and the flange holes are 

subjected to dimensional checks and to NDI inspection for cracks. Any fretting damage on 

the faying surface between the flange and the planetary plate is blended out. If any hole 

exhibits fretting damage or is worn beyond allowable limits, the damage is cleaned out by 

reaming the hole and installing an oversize dowel. See Attachment 4 – Sikorsky Overhaul 

Manual pages 464-476D. Reaming out the hole removes any fretting scars which can serve 

as initiation sites for fatigue cracking and also remove any incipient fatigue damage not 

detectible by NDI. Thus, in effect, at each overhaul cycle, the MRS plate dowel holes and 

the flange faying surface are restored to a “like new” condition. Thus, the life limits inferred 

by the SN curve constructed from the test data can be viewed as representing the maximum 

number of cycles permissible between replacements of the dowels with oversized dowels, 

rather than the service life for the MRS. 

• Flawed Assumptions about Operator Usage 

We believe that the proposed life limits for both the unmodified and modified MRS result 

from overly simplified assumptions about the S-61 duty cycle which ignore the statistical 

variability of loads and operator fleet management practices. As an example, the maximum 

rotor torque which can be achieved during a high torque event such as climb varies 

significantly. The variability of loads stems from differences in pilot techniques and also 

from physical limitations on the amount of torque which can be attained. The physical 
Jeffery Lee 

Aviation Safety Engineer 

Boston Aircraft Certification Office 

Page 5 of 9 

Croman Corporation 

801 Ave C White City, OR 97503  (541) 826-4455  (541) 826-4730 FAX 

limitations on torque stem from the air density reduction with increased pressure altitude and 

temperature. As an example, in summer or during hot temperatures, the main rotor torque is 

usually limited by the engine. In winter or during cold temperatures, the main rotor torque 

is usually limited by pilot technique. Another factor which influences the maximum torque 

is whether the logs are ferried up a slope or down a slope. The highest torque would occur 

during an uphill logging operation, where the base camp is located uphill from the log 

pickup site. The preferred practice is to locate the base camp below the logging site to 

minimize the power demands on the engines. In addition, further variability in the max 

torque results from differences in pilot technique and in the payload for the REL cycle.

The statistical variability of loads was recognized by Thompson and Adams of Sikorsky in a 

paper presented at the 1990 AHS Forum (A Computation Method for the Determination of 

the Structural Reliability of Dynamic Components). In the paper, Thompson and Adams 

argued that the industry practice of using top-of-scatter values to define a load cycle resulted 

in a reliability rate which was higher than the .999(3σ ) required by the FAA for type 

certification of dynamic system components. Thompson and Adams also argued that the 

variability in loading could be characterized by Weibull analysis of flight test data. Figure 3 

shows a Weibull plot of the cumulative rate of occurrences versus shaft bending moment for 

the UH-60A MRS for hover. The plot shows that the recorded shaft bending moments varied 

between 175,000 in-lb and 360,000 in-lb. The plot shows that bending moments between 

342,000 in-lb and 360,000 in-lb (95% to 100% max bending moment) comprise 

approximately 5% of the occurrences. The shaft bending moment was used by Sikorsky as 

the characterizing parameter for setting the life limit of MRS, as the hub moment was 

considered the primary loading defining the life of the shaft. The S-61 MRS was also 

substantiated using the shaft bending moment as the substantiating parameter, as the original 

testing for type certification was performed before REL cycles from high torque events 

became recognized as causing a different mode of failure than shaft bending and hence 

required separate testing and substantiation 
Jeffery Lee 

Aviation Safety Engineer 

Boston Aircraft Certification Office 

Page 6 of 9 

Croman Corporation 

801 Ave C White City, OR 97503  (541) 826-4455  (541) 826-4730 FAX 

Figure 3. Weibull Cummulative Rate of Occurrences for MRS Bending Moment During 

Hover. 

By accounting for the statistical variability, rather than using top of scatter loading to define the 

fatigue loading, Thompson and Adams showed that the imputed reliability was 0.999999, which 

was 3 orders of magnitude higher than the -3σ (0.999) reliability required for dynamic system 

components by FAA order 8110.9(Handbook for Vibration Substantiation and Fatigue 

Evaluation of Helicopter and Other Power Transmission Systems). While this result was 

obtained for MR shaft bending, a similar conclusion can be reached if MR torque is used as the 

substantiating parameter, because of the aforementioned reasons for variability. Thus, the 

assumption of 103%Q as the unvarying maximum for all REL cycles is unrealistic and 

excessive. 103% Q can occur, as indicated by Sikorsky surveys. We do not dispute that. 

However, as acknowledged by the FAA, 103%Q is a value which has been recorded in logging 

surveys, albeit only infrequently. Thus, it should not be used for establishing retirement times, 

unless credit is given for the undamaging or less damaging lower magnitude cycles. 

With this variability in mind, the test data can be reinterpreted. For example, specimen 2 

survived 1x106 cycles at 87%Q with no indication of cracking, followed by an additional 

1.41x106 cycles at 96% Q with no indication of cracking , followed by 200,000 cycles at 

103%Q, at which point the specimen was disassembled and fretting cracking found. From a 

statistical point of view, the testing is equivalent to testing of the specimen to a spectrum in 
Jeffery Lee 

Aviation Safety Engineer 

Boston Aircraft Certification Office 

Page 7 of 9 

Croman Corporation 

801 Ave C White City, OR 97503  (541) 826-4455  (541) 826-4730 FAX 

which 38% of the cycles occur at a max of 87%Q, 54% of the cycles occur at a max of 97%Q 

and 8% of the cycles occur at a max of 103%Q. If the specimen cracked at 200,000 cycles, the 

total number of cycles applied to the specimen is 2.6x106 cycles, rather than 200,000 cycles. 

Thus the interpretation of the test data failed to account for the fact that only the 103%Q loading 

was damaging, and that the torque cycles at lower maximum values were not damaging, but 

should have been included in the total cycle count. 

In addition to failing to account for the variability of loading due to physical limitations and 

pilot technique, the assumption of REL cycles between 0%Q and 103%Q is incorrect. 0%Q 

only occurs during autorotation or during a high speed flared approach to hover, neither of 

which is appropriate for a logging mission. A more realistic minimum for the REL cycle is the 

torque during the Partial Power Descent (PPD) portion of a logging turn. In addition, frequent 

excursions down to 0%Q will result in premature damage to the free wheeling clutch and 

require removal of the gearbox well before its scheduled TBO. The susceptibility of the S-61 

free-wheeling unit to damage from autorotation cycles is well known in the helicopter industry. 

For this reason, as a practical matter, most, if not all, S-61 operators require pilots to maintain a 

minimum torque. At Croman, for example, pilots are prohibited from going below 15% torque. 

So critical is staying above 15%Q that Croman modified its fleet of S-61’s to insure that pilots 

do not violate this mandate. A torque meter is mounted on the left hand side of the pilot by the 

collective so that as he is descending and looking out the bubble canopy, he can monitor the 

torque without looking up. 

Accounting for a realistic minimum torque can have a major impact on the life. As an example, 

consider the difference between a REL cycle between 0%Q and 103%Q torque and a REL cycle 

between 15%Q and 103%Q torque. As shown in Attachment 3, the use of 15% Q for the 

minimum for the REL cycle results in a reduction in the vibratory amplitude of 15%. This 

reduction, in turn, leads to a doubling of the retirement time using both the chafing and nonchafing SN curves. 

• Not consistent with Service History 

No cracks have been reported in 14 year period since the publication of AD 98-26-02. In 

particular no cracks have been found in oversized dowel holes repaired in accordance with the 

Sikorsky ORI. Additionally, with the incorporation of Customer Service Notice 6135-10A and 

ASB 611335-53 (installation of the steel planetary plates and 200 ksi fasteners) none of the 

shafts in service on our aircraft have required the dowel hole repair. 

REL Operations – Under the new proposed AD a new main shaft that is modified with 

Customer Service Notice 6135-10A and ASB 611335-53 will be life limited to 5,000 hours or 

30,000 cycles. With REL operations of 30 cycles per hour the main shaft will cycle out in 1,000 

hours. So this proposed AD will reduce the current 2,200 hour life imposed by AD 98-26-02 to 

approximately 1,000 hours for operators that are consistently involved in helicopter logging 

operations. AD98-26-03 Imposed a life limit of 2,200 hours with unlimited cycles for main 

shafts that are is modified with Customer Service Notice 6135-10A and ASB 611335-53.
Jeffery Lee 

Aviation Safety Engineer 

Boston Aircraft Certification Office 

Page 8 of 9 

Croman Corporation 

801 Ave C White City, OR 97503  (541) 826-4455  (541) 826-4730 FAX 

The proposed AD increases the life limit for shafts with the incorporation of Customer Service 

Notice 6135-10A and ASB 611335-53 to 5,000 hours but imposes the same cycle limit for 

modified and unmodified shafts. While 30,000 REL cycles may be an appropriate limit for 

unmodified MRS with titanium planetary plates to protect the plate from fatigue damage, it is 

not appropriate for MRS with the steel plates and 200 ksi bolts. The steel plates represent a 

substantial design improvement over the titanium plates due to the higher fatigue strength of the 

steel, the anti-fretting coating applied to the faying surface, and the greater resistance to wear of 

the dowel holes in both the shaft flange and in the plate. The greater resistance is attained by 

the higher yield strength of the steel and also by the higher elastic modulus of the steel. The 

higher elastic modulus is important as it provides a more rigid “socket” support for the dowels, 

minimizing the relative slip movement between the shaft flange and the plate from the 

overturning moment acting on the dowels as torque is transferred from the shaft to the plate. In 

addition, the use of 200 ksi bolts in lieu of 160 ksi bolts increases the permissible bolt 

installation torque. This, in turn, increases the contact pressure between the shaft and the plate, 

leading to a higher percent of the torque which can be transferred by friction and thus relieving 

the stresses in the shaft dowel holes. Thus, the steel plate with 200 ksi bolts reduces the fretting 

fatigue damage to the shaft, leading to a life increase in the shaft, as well as in the plate. 

Furthermore, maintaining the 30,000 REL restriction while increasing the maximum hours from 

1500 hours to 5000 hours is inconsistent. If the number of cycle is increased by the same ratio 

as the maximum TIS hours permitted, the number of REL cycles for MRS with steel plate 

should be 100,000 cycles. 

Cost of Compliance 

The Proposed AD estimates the cost of compliance at $44,753 for each helicopter. The current 

price for Part Number S6137-23040-001 Main Rotor Shaft is $81,215.10. The costs of the new 

shaft plus the NDI bring the compliance cost of the AD to $81,452.10. For operators involved in 

helicopter logging operations these costs would be incurred every 1,000 flight hours. 

Required Actions 

Paragraph (f) (1) (ii) directs the operator to substitute the helicopter’s Time in Service for the 

MRS time in service if no records of the Time in Service exist for the main rotor shaft. Few, if 

any, of the affected helicopters have less than 5,000 hours time in services and the vast majority 

of the helicopters have in excess of 13,000 hours time in service. A better measure of the 

potential shaft time in services would be the transmission time in service. 

SUMMARY 

Because of the reasons above, we believe that the proposed limit of 30,000 REL cycles for MRS 

with steel plates is unreasonable and is contrary to the OEM’s FAA approved fatigue evaluation 

practices. Further, the limit of 30,000 REL cycles for modified MRS ignores the crack-free 

service history for the modified MRS, ignores an operator’s prudent fleet management practices, 
Jeffery Lee 

Aviation Safety Engineer 

Boston Aircraft Certification Office 

Page 9 of 9 

Croman Corporation 

801 Ave C White City, OR 97503  (541) 826-4455  (541) 826-4730 FAX 

ignores variability in loads due to pilot technique and differences in payload, ignores the 

physical limits on torque due to pressure altitude or temperature, and ignores the benefits of 

scheduled maintenance performed per the OEM’s MM and service bulletins. We believe that 

in view of the arguments presented, that the fatigue test data supports, at a minimum, a 

retirement time of 100,000 REL cycles or 5000 hr TIS, which ever occurs first, for shafts 

which have been modified for steel plates and 200 ksi bolts. This retirement time is: 

• Consistent with the proposed limit of 30,000 REL cycles for the unmodified shaft as 

this value is simply the 30,000 REL limit multiplied by the increase in hr. from 1500 

to 5000 hr with steel plate. 

• Consistent with OEM FAA approved fatigue life prediction methodology 

• Consistent with statistical variability of maximum torques 

• Consistent with the manner in which the helicopter is flown and maintained 

• Consistent with service history 

Furthermore, this recommendation of 100,000 REL cycles is conservative. An even higher 

number of REL cycles can be justified if all factors are taken into account. As shown above in 

the previous section, this extension in the proposed retirement cycles from the proposed 30,000 

REL cycles can be justified just by using a chafing SN curve, instead of a non-chafing SN 

curve, and assuming an ultraconservative duty cycle for each REL cycle between 0 and 103% 

Q for determining life. The life increases attainable if the variability of usage and the 

prohibition against going below 15%Q are accounted for, and if restorative maintenance 

practices are accounted for, add further support for treating the recommended 100,000 REL 

cycle life limit as being conservative and justified. 

Sincerely, 

Greg Williams 

Director of Maintenance 

Croman Corporation