[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:
    <bullet> 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.
    <bullet> 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.
    <bullet> 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:
    <bullet> 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.
    <bullet> 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.
    <bullet> 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.
    <bullet> SA5730NM, issued June 26, 1992; amended October 2, 1992.
    <bullet> ST00131SE, issued November 8, 1994; amended January 26, 
1995; May

[[Page 36160]]

13, 1996; September 13, 1996; and February 20, 1997.
    <bullet> 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).

    <bullet> SA5730NM, issued June 26, 1992; amended October 2, 
1992.
    <bullet> ST00131SE, issued November 8, 1994; amended January 26, 
1995; May 13, 1996; September 13, 1996; and February 20, 1997.
    <bullet> 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