36158 Federal Register / Vol. 63, No. 127 /Thursday, July 2, 1998 /Rules and Regulations

63 FR 20298, in the third column,

§ 1980.108, introductory paragraph

(a)(1)(vi) and the first sentence of

paragraph (a)(1)(vi)(A), should be

corrected to read as follows:

§ 1980.108 General provisions.

(a) * * *

(1) * * *

(vi) The Agency may subordinate

direct loan basic security under

paragraph (a)(1)(v)(D) of this section

only when both of the following

additional conditions are met:

(A) The total unpaid principal and

interest balance of all of the borrower’s

direct loans secured by the property

being subordinated is less than or equal

to 75 percent of the value of all of the

basic security for the direct loan,

excluding the value of growing crops or

planned production, on the date the

Agency approves the subordination.

* * *

Signed at Washington, DC, on June 22,

1998.

August Schumacher Jr.,

Under Secretary for Farm and Foreign

Agricultural Services.

Dated: June 10, 1998.

Jill Long Thompson,

Under Secretary for Rural Development.

[FR Doc. 98–17562 Filed 7–1–98; 8:45 am]

BILLING CODE 3410–05–P

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.

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 inflight separation of an engine.
Federal Register / Vol. 63, No. 127 /Thursday, July 2, 1998 /Rules and Regulations 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 ‘‘oldertype’’ 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 ‘‘oldertype’’ 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 ‘‘newertype’’ 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 ‘‘newertype’’ 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 runon 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
36160 Federal Register / Vol. 63, No. 127 /Thursday, July 2, 1998 /Rules and Regulations

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.

§ 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 inflight 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.
Federal Register / Vol. 63, No. 127 /Thursday, July 2, 1998 /Rules and Regulations 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 underor 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

DEPARTMENT OF TRANSPORTATION

Federal Aviation Administration

14 CFR Part 71

[Airspace Docket No. 98–ASO–6]

Amendment of Class E Airspace;

Daytona Beach, FL; Correction

AGENCY: Federal Aviation

Administration (FAA), DOT.

ACTION: Final rule; correction.

SUMMARY: This action corrects an error

in the geographic position coordinates

of a final rule that was published in the

Federal Register on June 19, 1998, (63

FR 33544) Airspace Docket No. 98–

ASO–6. The final rule modified Class E

airspace at Daytona Beach, FL.

EFFECTIVE DATE: 0901 UTC, August 13,

1998.

FOR FURTHER INFORMATION CONTACT:

Nancy B. Shelton, Manager, Airspace

Branch, Air Traffic Division, Federal

Aviation Administration, P.O. Box

20636, Atlanta, Georgia 30320;

telephone (404) 305–5586.

SUPPLEMENTARY INFORMATION:

History

Federal Register Document 98–16355,

Airspace Docket No. 98–ASO–6,

published on June 19, 1998 (63 FR

33544), amended the Class E surface

area airspace at Daytona Beach, FL. A

VHF Omnidirectional Range (VOR) or

Global Positioning System (GPS)

Runway (RWY) 17 Standard Instrument

Approach Procedure (SIAP) has been

developed for Ormond Beach Municipal