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 yea rs 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 th e scrapping of serviceable MRS assemblies and is not justified.

We believe the proposed retirement time of 30,000 REL cycl es or 5000 hr TIS, whichever occurs first, for MRS which have been modified to the steel p lanetary 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 practic es recommend by the OEM

• Ignores the service history of modified MRS which have ac crued 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 b olts.

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 bas ed on the results of the laboratoryJeffery 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, Atta chment 1. In the attachment, Sikorsky presented the SN curve generated by fatigue testing two speci mens under repetitive torque cycles, along with its then recommended life limit. We believ e the conclusions drawn by Sikorsky regarding revised life limits are erroneous for the foll owing reasons.

• Flawed Interpretation of the Laboratory Test Data. The proposed life limit of 30,000 REL cycles is based on a -3 σ ”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 th an the chafing SN curve. Using the nonchafing SN curve in lieu of the chafing SN curve is incons istent with the fretting fatigue failures observed on the test specimens. Per Sikorsky r eport SER 61364, Methods of Service Life Analysis of Structural Components Substanti ated by Laboratory Testing for FAA Certification of S-61L/S,S61R (CH-3) Helicopter, a c haracteristic SN curve for chafing is to be used for generating a component SN curve fr om test data if the failure mode is fretting. Attachment 1 from the 1998 operator’s confere nce 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 th e 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 CurveJeffery 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 with out 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 t he non-chafing curve is indicative of

insufficient data.

Figure 2 shows mean and working SN curves for chafing and n on-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 an d 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~cyclesVibratory Torque~ft-lb

Working-No Chafing WorkingChafing

Test Failure Point Mean--No Chafing

Mean--Chafing

Figure 2. SN Curves Generated from Single Test Failure P oint.

Figure 2 shows that at a vibratory torque value of 29,515 ft-lb , the -3 σ working life if fretting is assumed is close to 200,000 cycles while the wo rking 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 inJeffery 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 r esult of the characteristic SN curve shapes of the Sikorsky methodology and of insuffic ient test data points. The SN curves of Figure 2 are generated from a single data point t o illustrate the shortfalls of using insufficient data points. For multiple test failure p oints, SN curves as shown in Figure 2 must be generated for each failure point. The mean SN c urve is then obtained by averaging the imputed endurance limit for each curve. The working curv e 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 t hat 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 s pecifies the use of a characteristic chafing curve for generating a SN cur ve from laboratory test data if the failure mode is fatigue abetted by fretting. We believe t hat the non-chafing SN curve should be disregarded, since its use is physically incorrect. B ased on the chafing SN curve, we believe that a retirement time in excess of 100,000 REL c ycles is appropriate for a REL cycle between 0%Q and 103%Q.

Another flaw in the interpretation of the test data i s the assumption that the MRS is subjected to continuous operation without scheduled maintena nce 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 gea rbox. 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 c racks. Any fretting damage on the faying surface between the flange and the planetary pl ate is blended out. If any hole exhibits fretting damage or is worn beyond allowable limit s, the damage is cleaned out by reaming the hole and installing an oversize dowel. See A ttachment 4 – Sikorsky Overhaul Manual pages 464-476D. Reaming out the hole removes any fretting scars which ca n serve as initiation sites for fatigue cracking and also remove any incipient fatigue damage not detectible by NDI. Thus, in effect, at each overhaul cy cle, the MRS plate dowel holes and the flange faying surface are restored to a “like new” condi tion. Thus, the life limits inferred by the SN curve constructed from the test data can be view ed 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 c ycle which ignore the statistical variability of loads and operator fleet management practi ces. As an example, the maximum rotor torque which can be achieved during a high torque event such a s climb varies significantly. The variability of loads stems from di fferences in pilot techniques and also from physical limitations on the amount of torque which ca n be attained. The physicalJeffery 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 tem peratures, the main rotor torque is usually limited by the engine. In winter or during cold tem peratures, the main rotor torque is usually limited by pilot technique. Another factor which i nfluences the maximum torque is whether the logs are ferried up a slope or down a slo pe. The highest torque would occur during an uphill logging operation, where the base camp is lo cated 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 Th ompson and Adams of Sikorsky in a paper presented at the 1990 AHS Forum (A Computation Method fo r the Determination of the Structural Reliability of Dynamic Components). In t he paper, Thompson and Adams argued that the industry practice of using top-of-scatter va lues to define a load cycle resulted in a reliability rate which was higher than the .999(3 σ ) required by the FAA for type certification of dynamic system components. Thompson a nd Adams also argued that the variability in loading could be characterized by Weibull a nalysis of flight test data. Figure 3 shows a Weibull plot of the cumulative rate of occur rences versus shaft bending moment for the UH-60A MRS for hover. The plot shows that the rec orded shaft bending moments varied between 175,000 in-lb and 360,000 in-lb. The plot shows that bend ing moments between 342,000 in-lb and 360,000 in-lb (95% to 100% max bending moment) compris e approximately 5% of the occurrences. The shaft bending mom ent 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 shaf t. The S-61 MRS was also substantiated using the shaft bending moment as the substanti ating parameter, as the original testing for type certification was performed before RE L cycles from high torque events became recognized as causing a different mode of failure th an shaft bending and hence required separate testing and substantiationJeffery 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 t han 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 -3 σ (0.999) reliability required for dynamic system components by FAA order 8110.9(Handbook for Vibration Substan tiation 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 r easons for variability. Thus, the assumption of 103%Q as the unvarying maximum for all REL cyc les is unrealistic and excessive. 103% Q can occur, as indicated by Sikorsky survey s. We do not dispute that. However, as acknowledged by the FAA, 103%Q is a value which h as been recorded in logging surveys, albeit only infrequently. Thus, it should not be use d for establishing retirement times, unless credit is given for the undamaging or less damaging l ower magnitude cycles.

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

survived 1x10 6 cycles at 87%Q with no indication of cracking, follo wed by an additional

1.41x10 6 cycles at 96% Q with no indication of cracking , fol lowed by 200,000 cycles at 103%Q, at which point the specimen was disassembled and fret ting cracking found. From a statistical point of view, the testing is equivalent to testing of the specimen to a spectrum inJeffery 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 speci men cracked at 200,000 cycles, the total number of cycles applied to the specimen is 2.6x10 6 cycles, rather than 200,000 cycles. Thus the interpretation of the test data failed to acco unt 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 a pproach to hover, neither of which is appropriate for a logging mission. A more reali stic 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 th e free wheeling clutch and require removal of the gearbox well before its schedul ed 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 pro hibited from going below 15% torque. So critical is staying above 15%Q that Croman modified it s fleet of S-61’s to insure that pilots do not violate this mandate. A torque meter is mounted on th e 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 majo r 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 t he vibratory amplitude of 15%. This reduction, in turn, leads to a doubling of the retirement ti me using both the chafing and nonchafing SN curves.

• Not consistent with Service History No cracks have been reported in 14 year period since the publi cation of AD 98-26-02. In particular no cracks have been found in oversized dowel h oles repaired in accordance with the Sikorsky ORI. Additionally, with the incorporation of Cus tomer Service Notice 6135-10A and ASB 611335-53 (installation of the steel planetary plates an d 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 s haft that is modified with

Customer Service Notice 6135-10A and ASB 611335-53 will be life limit ed 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 l ife 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 unl imited 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 w ith the incorporation of Customer Service Notice 6135-10A and ASB 611335-53 to 5,000 hours but imposes the same cycl e limit for modified and unmodified shafts. While 30,000 REL cycles may b e 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 b olts. 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 sur face, and the greater resistance to wear of the dowel holes in both the shaft flange and in the pl ate. The greater resistance is attained by the higher yield strength of the steel and also by the high er 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 shaf t flange and the plate from the overturning moment acting on the dowels as torque is tra nsferred from the shaft to the plate. In addition, the use of 200 ksi bolts in lieu of 160 ksi bolts in creases 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 trans ferred by friction and thus relieving the stresses in the shaft dowel holes. Thus, the ste el 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 inc reasing the maximum hours from 1500 hours to 5000 hours is inconsistent. If the number of cycl e is increased by the same ratio as the maximum TIS hours permitted, the number of REL c ycles 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 Servi ce 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 transmissi on time in service.

SUMMARY

Because of the reasons above, we believe that the propo sed limit of 30,000 REL cycles for MRS with steel plates is unreasonable and is contrary to t he OEM’s FAA approved fatigue evaluation practices. Further, the limit of 30,000 REL cycles for mo dified MRS ignores the crack-free service history for the modified MRS, ignores an operat or’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 diff erences 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 servic e 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 t he 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 m ethodology

• 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 co nservative. 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 reti rement cycles from the proposed 30,000 REL cycles can be justified just by using a chafing SN c urve, 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 i f 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