[Federal Register Volume 63, Number 127 (Thursday, July 2, 1998)]
[Rules and Regulations]
[Pages 36158-36161]
From the Federal Register Online via the Government Publishing Office [www.gpo.gov]
[FR Doc No: 98-17523]
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DEPARTMENT OF TRANSPORTATION
Federal Aviation Administration
14 CFR Part 39
[Docket No. 98-NM-121-AD; Amendment 39-10642; AD 98-14-09]
RIN 2120-AA64
Airworthiness Directives; Boeing Model 737-100, -200, -200C
Series Airplanes
AGENCY: Federal Aviation Administration, DOT.
ACTION: Final rule; request for comments.
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SUMMARY: This amendment adopts a new airworthiness directive (AD) that
is applicable to certain Boeing Model 737-100, -200, and -200C series
airplanes. This action requires repetitive inspections to detect
fatigue cracking and certain discrepancies of the forward engine mount
support (FEMS) fitting and its attachments, and repair, if necessary.
This amendment is prompted by reports of fatigue cracks on the lower
flange of the FEMS fitting, broken bolts and bolts with loose or
detached nuts on the upper inboard attachment of the FEMS fitting, and
cracked or severed lugs at the outboard support link attachment of the
FEMS fitting. The actions specified in this AD are intended to detect
and correct fatigue cracking and certain discrepancies of the FEMS
fitting and its attachments, which could result in an in-flight
separation of an engine.
DATES: Effective July 17, 1998.
The incorporation by reference of certain publications listed in
the regulations is approved by the Director of the Federal Register as
of July 17, 1998.
Comments for inclusion in the Rules Docket must be received on or
before August 31, 1998.
ADDRESSES: Submit comments in triplicate to the Federal Aviation
Administration (FAA), Transport Airplane Directorate, ANM-114,
Attention: Rules Docket No. 98-NM-121-AD, 1601 Lind Avenue, SW.,
Renton, Washington 98055-4056.
The service information referenced in this AD may be obtained from
Boeing Commercial Airplane Group, P.O. Box 3707, Seattle, Washington
98124-2207. This information may be examined at the FAA, Transport
Airplane Directorate, 1601 Lind Avenue, SW., Renton, Washington; or at
the Office of the Federal Register, 800 North Capitol Street, NW.,
suite 700, Washington, DC.
FOR FURTHER INFORMATION CONTACT: Gregory L. Schneider, Aerospace
Engineer, Airframe Branch, ANM-120S, FAA, Transport Airplane
Directorate, Seattle Aircraft Certification Office, 1601 Lind Avenue,
SW., Renton, Washington 98055-4056; telephone (425) 227-2028; fax (425)
227-1181.
SUPPLEMENTARY INFORMATION: The FAA has received reports of certain
problems affecting the forward engine mount support (FEMS) fitting on
certain Boeing Model 737 series airplanes. This support fitting is one
of the primary structural elements that attach the engine to the wing.
The reports indicate that three critical elements of the FEMS fitting
have proved to be susceptible to fatigue damage or other problems as
summarized below:
Lower Flange of the FEMS Fitting:
The FAA has received 17 reports of cracks of the lower flange ``I''
section of the FEMS fitting. Analysis indicates that the cracks were
initiated by fatigue. A FEMS fitting that has a cracked lower flange
may not be capable of withstanding certain limit load conditions.
Upper Inboard Attachment Bolt:
There have been 13 cases of the upper inboard attachment bolt
fracturing in service due to fatigue, and 4 cases of the nut being
broken, loose, or detached. Investigation revealed that the original
production bolt installation was subject to relative motion between the
bushing and the attachment bolt. As a result, the production nut (which
has no secondary locking features) tended to come loose in service. A
later configuration change that was intended to correct this problem
consisted of installing a stronger bolt and nut, and a new bushing.
This change, which has subsequently been adopted by almost the entire
fleet of affected airplanes, requires the nut to be torqued to a higher
value than is appropriate for the bolt and nut installation.
Specifically, the torque applied to the new nut is applicable to a
``non-lubricated'' thread condition, whereas the nut material tends to
act as a ``dry'' lubricant. Consequently, the higher torque applied to
the new bolt and nut configuration induces an excessive pre-load on the
bolt threads. This excessive pre-load, in conjunction with certain
operational loads, causes an overload condition on the bolt threads,
which in turn leads to premature fatigue cracking of the bolt.
Additionally, results of an analysis indicate that the FEMS fitting
cannot react certain limit load conditions with a fractured or detached
bolt at this location.
Upper Outboard Lug of the FEMS Fitting:
The upper outboard lug of the FEMS fitting contains a bearing that
has proved susceptible to excessive wearing. This lug is designed to
secure the outboard end of the FEMS fitting to the wing. A severely
worn bearing could drastically reduce the fatigue life of the lug. This
condition has been observed on six airplanes to date; on three of those
airplanes the lug was found to be completely fractured. Analysis has
revealed that the FEMS fitting cannot react certain limit load
conditions with a severed lug.
Explanation of the Unsafe Condition
The fatigue cracking problems that affect the three areas of the
FEMS fitting are examples of ``multiple element damage.'' The existence
of any one of these conditions could result in an engine separation
under certain limit load conditions. The simultaneous existence of any
two conditions could result in an immediate engine loss at loads that
are much lower than the design limit loads. These problems, if not
corrected, could result in an in-flight separation of an engine.
[[Page 36159]]
Explanation of Relevant Service Information
The FAA has reviewed and approved the following three service
bulletins:
Boeing Service Bulletin 737-54A1012, Revision 4, dated
March 26, 1998, addresses fatigue cracking of the lower flange of the
FEMS fitting. The service bulletin notes that the fatigue cracking
problem affects only ``older-type'' FEMS fittings that have a lower
flange thickness of 0.32 inches (nominal). Therefore, the service
bulletin describes procedures for performing repetitive detailed visual
inspections of the lower flange of the ``older-type'' FEMS fitting to
detect fatigue cracking, and corrective action, if necessary. The
corrective action includes replacement of the ``older-type'' FEMS
fitting with a ``newer-type'' FEMS fitting, which would eliminate the
need for the repetitive detailed visual inspections. These inspections
are not required on ``newer-type'' FEMS fittings [i.e., those FEMS
fittings having lower flanges that are 0.40 inches (nominal) thick],
since there have been no reports of fatigue cracking of the lower
flange of these parts.
Boeing Service Bulletin 737-54-1007, Revision 1, dated
March 26, 1998, describes procedures for performing repetitive detailed
visual inspections of the upper inboard attachment of the FEMS fitting
to detect bolt deformation or fatigue damage. Additionally, the service
bulletin recommends that operators perform a torque check during each
inspection to ensure that the nut and bolt installation has retained
its integrity. The service bulletin also describes procedures for an
initial and two follow-on ultrasonic inspections of the bolt to detect
fatigue cracking, and replacement of any discrepant part.
The service bulletin recommends that, if the three successive
ultrasonic inspections (i.e., the initial and the two follow-on
inspections) reveal that the bolt is undamaged, the need for further
ultrasonic inspections would be eliminated. In addition, the service
bulletin describes procedures for replacement of the bolt and nut
installation with a new Nickel Alloy 718 bolt and associated nut, which
would eliminate the need for the repetitive detailed visual inspections
and torque checks.
Boeing Service Bulletin 737-54-1009, Revision 1, dated
March 26, 1998, describes procedures for repetitive detailed visual
inspections of the lug of the outboard support link attachment of the
FEMS fitting to detect cracked or severed lugs; and corrective action,
if necessary. The service bulletin notes that some of the lug structure
will not be visible during the detailed visual inspection. If a crack
is detected, the corrective action is to replace the cracked FEMS
fitting with a ``newer-type'' FEMS fitting and to install a new
bearing. The service bulletin also describes procedures for an optional
preventive modification, which entails removing the engine, installing
a new bearing, and re-installing the existing fitting (provided that a
magnetic particle inspection shows that the lug of the existing FEMS
fitting is free of cracks).
Explanation of the Requirements of the Rule
Since an unsafe condition has been identified that is likely to
exist or develop on other airplanes of the same type design, this AD is
being issued to detect and correct fatigue cracking and certain
discrepancies of the FEMS fitting and its attachments, which could
result in an in-flight separation of an engine. This AD requires
accomplishment of the actions specified in the service bulletins
described previously, except as discussed below. This AD also requires
that operators report any adverse (negative) inspection findings to the
FAA.
Differences Between the AD and the Service Bulletins
Boeing Service Bulletin 737-54A1012, Revision 4, specifies that if
cracking of the lower flange of the FEMS fitting is found, the cracked
FEMS fitting should be replaced with a ``newer-type'' FEMS fitting.
Such installation of a ``newer-type'' FEMS fitting would constitute
terminating action for the repetitive detailed visual inspection
requirements of this AD. However, since sufficient parts may not be
available for all of the affected airplanes, this AD allows operators
to install either an ``older-type'' FEMS fitting that is
``serviceable,'' or a ``newer-type'' FEMS fitting. The installation of
a ``serviceable'' FEMS fitting instead of a ``newer-type'' FEMS fitting
would not terminate the repetitive detailed visual inspections required
by this AD. Rather, these inspections would continue until a ``newer-
type'' FEMS fitting is installed. For the purposes of this AD, a
``serviceable'' FEMS fitting is defined as an ``older-type'' FEMS
fitting that has been shown to be free of cracks by means of a magnetic
particle inspection. This AD also requires operators to perform the
magnetic particle inspection in accordance with a method approved by
the FAA.
Although Boeing Service Bulletin 737-54-1007, Revision 1, advises
operators to examine the nut of the FEMS fitting inboard attachment for
looseness, it does not provide procedures for determining if the nut is
too tight. This AD requires operators to examine the nut for both
looseness and excessive tightness. This AD also requires that, if the
nut is found to be too loose or too tight, the nut is to be re-torqued
to a value of 440 to 650 pound-inches, provided that a run-on torque
value of at least 18 pound-inches can be achieved. If the run-on torque
value cannot be achieved, the nut is to be replaced with a new nut.
This run-on torque check is to be accomplished by loosening the nut
sufficiently to demonstrate that a minimum run-on torque value of 18
pound-inches can be achieved. Finally, this AD requires operators to
perform this same run-on torque check on any new nut that is installed
on the bolt. If a new nut should fail the 18 pound-inches minimum
requirement, then this would imply that the bolt thread was defective.
Therefore, if this were to occur, this AD requires the operator to
replace the existing bolt installation with a stronger bolt
installation in accordance with the service bulletin.
Boeing Service Bulletin 737-54-1009, Revision 1, specifies that the
manufacturer may be contacted for disposition of certain repair
conditions (i.e., for a repair of a cracked lug). However, this AD
requires that the repair of those conditions be accomplished in
accordance with a method approved by the FAA.
Previously Modified Airplanes
Each of the three Boeing service bulletins specified in this AD
contains the following statement: ``If an airplane has a non-Boeing
modification or repair that affects a component or system affected by
this service bulletin, the operator is responsible for obtaining
appropriate regulatory agency approval before incorporating this
service bulletin.''
The FAA is aware that a certain proportion of the airplanes listed
in the effectivity sections of the three service bulletins have already
been modified by certain non-Boeing engine hush-kit supplemental type
certificates (STC). The FAA has determined that the following hush-kit
STC's are compatible with the service bulletins; therefore, operators
of airplanes modified with the following STC's need not seek prior FAA
approval before accomplishing the requirements of this AD.
SA5730NM, issued June 26, 1992; amended October 2, 1992.
ST00131SE, issued November 8, 1994; amended January 26,
1995; May
[[Page 36160]]
13, 1996; September 13, 1996; and February 20, 1997.
ST223CH, issued July 7, 1994; amended August 11, 1994;
December 19, 1994; May 30, 1995; and October 14, 1997.
Interim Action
This is considered to be interim action. The FAA is currently
considering requiring replacement of the attachment bolt installation
and the bearing with new and improved replacement parts. However, the
planned compliance time for installation of new and improved parts is
sufficiently long that notice and opportunity for prior public comment
will be practicable.
Determination of Rule's Effective Date
Since a situation exists that requires the immediate adoption of
this regulation, it is found that notice and opportunity for prior
public comment hereon are impracticable, and that good cause exists for
making this amendment effective in less than 30 days.
Comments Invited
Although this action is in the form of a final rule that involves
requirements affecting flight safety and, thus, was not preceded by
notice and an opportunity for public comment, comments are invited on
this rule. Interested persons are invited to comment on this rule by
submitting such written data, views, or arguments as they may desire.
Communications shall identify the Rules Docket number and be submitted
in triplicate to the address specified under the caption ``ADDRESSES.''
All communications received on or before the closing date for comments
will be considered, and this rule may be amended in light of the
comments received. Factual information that supports the commenter's
ideas and suggestions is extremely helpful in evaluating the
effectiveness of the AD action and determining whether additional
rulemaking action would be needed.
Comments are specifically invited on the overall regulatory,
economic, environmental, and energy aspects of the rule that might
suggest a need to modify the rule. All comments submitted will be
available, both before and after the closing date for comments, in the
Rules Docket for examination by interested persons. A report that
summarizes each FAA-public contact concerned with the substance of this
AD will be filed in the Rules Docket.
Commenters wishing the FAA to acknowledge receipt of their comments
submitted in response to this rule must submit a self-addressed,
stamped postcard on which the following statement is made: ``Comments
to Docket Number 98-NM-121-AD.'' The postcard will be date stamped and
returned to the commenter.
Regulatory Impact
The regulations adopted herein will not have substantial direct
effects on the States, on the relationship between the national
government and the States, or on the distribution of power and
responsibilities among the various levels of government. Therefore, in
accordance with Executive Order 12612, it is determined that this final
rule does not have sufficient federalism implications to warrant the
preparation of a Federalism Assessment.
The FAA has determined that this regulation is an emergency
regulation that must be issued immediately to correct an unsafe
condition in aircraft, and that it is not a ``significant regulatory
action'' under Executive Order 12866. It has been determined further
that this action involves an emergency regulation under DOT Regulatory
Policies and Procedures (44 FR 11034, February 26, 1979). If it is
determined that this emergency regulation otherwise would be
significant under DOT Regulatory Policies and Procedures, a final
regulatory evaluation will be prepared and placed in the Rules Docket.
A copy of it, if filed, may be obtained from the Rules Docket at the
location provided under the caption ADDRESSES.
List of Subjects in 14 CFR Part 39
Air transportation, Aircraft, Aviation safety, Incorporation by
reference, Safety.
Adoption of the Amendment
Accordingly, pursuant to the authority delegated to me by the
Administrator, the Federal Aviation Administration amends part 39 of
the Federal Aviation Regulations (14 CFR part 39) as follows:
PART 39--AIRWORTHINESS DIRECTIVES
1. The authority citation for part 39 continues to read as follows:
Authority: 49 U.S.C. 106(g), 40113, 44701.
Sec. 39.13 [Amended]
2. Section 39.13 is amended by adding the following new
airworthiness directive:
98-14-09 Boeing: Amendment 39-10642. Docket 98-NM-121-AD.
Applicability: Model 737-100, -200, -200C series airplanes,
manufacturer's line positions 001 through 1585 inclusive;
certificated in any category.
Note 1: This AD applies to each airplane identified in the
preceding applicability provision, regardless of whether it has been
modified, altered, or repaired in the area subject to the
requirements of this AD. For airplanes that have been modified,
altered, or repaired so that the performance of the requirements of
this AD is affected, the owner/operator must request approval for an
alternative method of compliance in accordance with paragraph (e) of
this AD. The request should include an assessment of the effect of
the modification, alteration, or repair on the unsafe condition
addressed by this AD; and, if the unsafe condition has not been
eliminated, the request should include specific proposed actions to
address it.
Note 2: The performance of the requirements of this AD is not
affected by modifications in accordance with the following
supplemental type certificates (STC's).
SA5730NM, issued June 26, 1992; amended October 2,
1992.
ST00131SE, issued November 8, 1994; amended January 26,
1995; May 13, 1996; September 13, 1996; and February 20, 1997.
ST223CH, issued July 7, 1994; amended August 11, 1994;
December 19, 1994; May 30, 1995; and October 14, 1997.
Compliance: Required as indicated, unless accomplished
previously.
To detect and correct fatigue cracking and certain discrepancies
of the forward engine mount support (FEMS) fitting and its
attachments, which could result in an in-flight separation of an
engine, accomplish the following:
(a) For airplanes on which a ``newer-type'' FEMS fitting having
part number (P/N) 65-46850-9/-10 or 65-46850-13/-14 has not been
installed: Within 90 days or 700 flight cycles after the effective
date of this AD, whichever occurs later, perform a detailed visual
inspection to detect fatigue cracking of the lower flange of the
FEMS fitting, in accordance with the Accomplishment Instructions of
Boeing Service Bulletin 737-54A1012, Revision 4, dated March 26,
1998.
(1) If no fatigue cracking of the lower flange of the FEMS
fitting is found, or if a ``serviceable'' FEMS fitting is installed
in lieu of a ``newer-type'' FEMS fitting, repeat the inspection
thereafter at intervals not to exceed 700 flight cycles in
accordance with the service bulletin.
Note 3: For the purposes of this AD, a ``serviceable'' FEMS
fitting is defined as an ``older-type'' FEMS fitting that is free of
cracking, as shown by a magnetic particle inspection performed in
accordance with a method approved by the Manager, Seattle Aircraft
Certification Office (ACO), FAA, Transport Airplane Directorate.
(2) If any cracking of the lower flange of the FEMS fitting is
found, prior to further flight, replace the FEMS fitting with a
``serviceable'' or a ``newer-type'' FEMS fitting in accordance with
the service bulletin. Replacement of this part with a ``newer-type''
FEMS fitting constitutes terminating action for the repetitive
inspection requirements of paragraph (a)(1) of this AD.
[[Page 36161]]
(b) Within 90 days or 700 flight cycles after the effective date
of this AD, whichever occurs later, perform a detailed visual
inspection to detect deformation or fatigue damage of the bolt at
the upper inboard attachment of the FEMS fitting; perform a torque
check to detect any bolt that is under-or over-torqued; and perform
an ultrasonic inspection to detect any cracking of the bolt; in
accordance with the Accomplishment Instructions of Boeing Service
Bulletin 737-54-1007, Revision 1, dated March 26, 1998.
(1) If no bolt deformation or fatigue damage, under- or over-
torqued nut, or fatigue cracking is found: Thereafter, repeat the
detailed visual inspection and torque check required by paragraph
(b) of this AD at intervals not to exceed 700 flight cycles.
Additionally, repeat the ultrasonic inspection two more times at
intervals not to exceed 700 flight cycles, but no earlier than 600
flight cycles.
(2) If any deformation, fatigue damage, or fatigue cracking of
the inboard attachment bolt is found during any inspection required
by this paragraph: Prior to further flight, replace the inboard
attachment bolt and nut with a new Nickel Alloy 718 bolt and
associated nut in accordance with the service bulletin. Replacement
of the inboard attachment bolt and nut in accordance with the
service bulletin constitutes terminating action for the repetitive
inspection requirements of paragraphs (b)(1), (b)(2), and (b)(3) of
this AD.
(3) If the torque check shows that a nut is torqued to any value
outside the limits of 440 to 650 pound-inches, prior to further
flight, accomplish paragraphs (b)(3)(i) and (b)(3)(ii) of this AD.
(i) Loosen the affected nut enough to demonstrate that a minimum
run-on torque value of 18 pound-inches can be achieved. If this
value cannot be achieved, install a new nut in accordance with the
service bulletin, and repeat the run-on torque check prior to
tightening the nut to 440-650 inch pounds. If a run-on torque value
of 18 pound-inches still cannot be achieved, prior to further
flight, replace the inboard attachment bolt and nut with a new
Nickel Alloy 718 bolt and associated nut in accordance with the
service bulletin.
(ii) Tighten the affected nut to 440-650 pound-inches in
accordance with the service bulletin.
(c) Within 90 days or 700 flight cycles after the effective date
of this AD, whichever occurs later, perform a detailed visual
inspection to detect any cracked or severed lug of the outboard
support link attachment of the FEMS fitting, in accordance with the
Accomplishment Instructions of Boeing Service Bulletin 737-54-1009,
Revision 1, dated March 26, 1998.
(1) If no cracked or severed lug is detected: Repeat the
detailed visual inspection required by paragraph (c) thereafter at
intervals not to exceed 700 flight cycles, or perform the optional
terminating modification, in accordance with Part II of the
Accomplishment Instructions of the service bulletin. Where the
service bulletin specifies that the manufacturer may be contacted
for disposition of certain repair conditions, repair in accordance
with a method approved by the Manager, Seattle ACO. Accomplishment
of this modification constitutes terminating action for the
repetitive inspection requirements of paragraph (c) of this AD.
(2) If any cracked or severed lug is found, prior to further
flight, accomplish the requirements of paragraphs (c)(2)(i) and
(c)(2)(ii) of this AD.
(i) Replace the FEMS fitting with a ``serviceable'' or a
``newer-type'' FEMS fitting in accordance with Accomplishment
Instructions of Boeing Service Bulletin 737-54A1012, Revision 4,
dated March 26, 1998. Replacement of the FEMS fitting with a
``newer-type'' FEMS fitting in accordance with the service bulletin
constitutes terminating action for the repetitive inspection
requirements of paragraph (a) of this AD.
(ii) Install a new bearing, which is inserted into the lug of
the replacement FEMS fitting, in accordance with the Accomplishment
Instructions of Boeing Service Bulletin 737-54-1009, Revision 1,
dated March 26, 1998. Replacement of the existing bearing with an
improved bearing constitutes terminating action for the repetitive
inspection requirements of the lug that are specified in paragraph
(c) of this AD.
(d) Within 20 days after accomplishing the initial inspections
required by paragraphs (a), (b), and (c) of this AD, or within 20
days after the effective date of this AD, whichever occurs later,
submit a report of the inspection results (adverse findings only) to
the Manager, Seattle ACO, FAA, Transport Airplane Directorate, 1601
Lind Avenue, SW., Renton, Washington 98055-4056; fax (425) 227-1181.
Required information for each report must include the following: A
description of the adverse finding, airplane serial number and total
flight cycles and flight hours accumulated, number of flight cycles
and flight hours accumulated since the last engine change, and the
number of flight cycles and flight hours accumulated since the last
inspection of the affected part. Information collection requirements
contained in this regulation have been approved by the Office of
Management and Budget (OMB) under the provisions of the Paperwork
Reduction Act of 1980 (44 U.S.C. 3501 et seq.) and have been
assigned OMB Control Number 2120-0056.
(e) An alternative method of compliance or adjustment of the
compliance time that provides an acceptable level of safety may be
used if approved by the Manager, Seattle ACO. Operators shall submit
their requests through an appropriate FAA Principal Maintenance
Inspector, who may add comments and then send it to the Manager,
Seattle ACO.
Note 4: Information concerning the existence of approved
alternative methods of compliance with this AD, if any, may be
obtained from the Seattle ACO.
(f) Special flight permits may be issued in accordance with
sections 21.197 and 21.199 of the Federal Aviation Regulations (14
CFR 21.197 and 21.199) to operate the airplane to a location where
the requirements of this AD can be accomplished.
(g) Except as provided in paragraph (c)(1) of this AD, the
actions shall be done in accordance with Boeing Service Bulletin
737-54A1012, Revision 4, dated March 26, 1998; Boeing Service
Bulletin 737-54-1007, Revision 1, dated March 26, 1998; and Boeing
Service Bulletin 737-54-1009, Revision 1, dated March 26, 1998. This
incorporation by reference was approved by the Director of the
Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR part
51. Copies may be obtained from Boeing Commercial Airplane Group,
P.O. Box 3707, Seattle, Washington 98124-2207. Copies may be
inspected at the FAA, Transport Airplane Directorate, 1601 Lind
Avenue, SW., Renton, Washington; or at the Office of the Federal
Register, 800 North Capitol Street, NW., suite 700, Washington, DC.
(h) This amendment becomes effective on July 17, 1998.
Issued in Renton, Washington, on June 25, 1998.
John J. Hickey,
Acting Manager, Transport Airplane Directorate, Aircraft Certification
Service.
[FR Doc. 98-17523 Filed 7-1-98; 8:45 am]
BILLING CODE 4910-13-P
