64514 Federal Register / Vol. 62, No. 235 / Monday, December 8, 1997 / Rules and Regulations

Docket Number 97–ANE–21–AD.’’ The

postcard will be date stamped and

returned to the commenter.

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 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:

97–25–08 General Electric Company:

Amendment 39–10232. Docket 97–ANE–

21–AD.

Applicability: General Electric Company

(GE) CJ610 series turbojet and CF700 series

turbofan engines, with turbine torque rings

and compressor drive shafts identified in GE

CF700 Service Bulletin (SB) No. A72–155,

dated May 22, 1997, and GE CJ610 SB No.

A72–147, dated May 22, 1997. These engines

are installed on but not limited to the

following aircraft: Learjet 20 series, Israel

Aircraft Industries Westwind series, Hansa

Jet, Aero Commander Jet Commander,

Dassault Falcon 20 series, Sabreliner 265

series.

Note 1: This airworthiness directive (AD)

applies to each engine 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 engines 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 (d)

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.

Compliance: Required as indicated, unless

accomplished previously.

To prevent turbine torque ring or

compressor drive shaft failure due to a

manufacturing defect, which could result in

an uncontained engine failure, accomplish

the following:

(a) For GE CF700 series turbofan engines,

accomplish the following in accordance with

GE CF700 SB No. A72–155, dated May 22,

1997:

(1) Remove from service affected turbine

torque rings, listed by serial number (S/N) in

paragraph 1. A.(3a) of GE CF700 SB No. A72–

155, dated May 22, 1997, and replace with

serviceable parts, within 50 hours time in

service (TIS), or 60 days after the effective

date of this AD, whichever occurs first.

(2) Remove from service affected turbine

torque rings and compressor drive shafts,

listed by S/N in GE paragraph 1.A.(3b) of GE

CF700 SB No. A72–155, dated May 22, 1997,

and replace with serviceable parts, within

300 hours TIS, or 12 months after the

effective date of this AD, whichever occurs

first.

(b) For GE CJ610 series turbojet engines,

accomplish the following in accordance with

GE CJ610 SB No. A72–147, dated May 22,

1997:

(1) Remove from service affected turbine

torque rings, listed by S/N in paragraph

1.A.(3a) of GE CJ610 SB No. A72–147, dated

May 22, 1997, and replace with serviceable

parts, within 50 hours TIS, or 60 days after

the effective date of this AD, whichever

occurs first.

(2) Remove from service affected turbine

torque rings and compressor drive shafts,

listed by S/N in paragraph 1.A.(3b) of GE

CJ610 SB No. A72–147, dated May 22, 1997,

and replace with serviceable parts, within

300 hours TIS, or 12 months after the

effective date of this AD, whichever occurs

first.

(c) After the effective date of this AD,

installation of uninstalled affected parts

identified by S/N in the SBs referenced in

paragraphs (a) and (b) of this AD is

prohibited.

(d) 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, Engine

Certification Office. Operators shall submit

their requests through an appropriate FAA

Principal Maintenance Inspector, who may

add comments and then send it to the

Manager, Engine Certification Office.

Note 2: Information concerning the

existence of approved alternative methods of

compliance with this airworthiness directive,

if any, may be obtained from the Engine

Certification Office.

(e) 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 aircraft to

a location where the requirements of this AD

can be accomplished.

(f) The actions required by this AD shall be

done in accordance with the following GE

service documents:

Document No. Pages Date

CF700 SB No. A72–155 1–9 May 22,

1997.

Total Pages: 9.

CJ610 SB No. A72–147 1–9 May 22,

1997.

Total Pages: 9.

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 GE Aircraft Engines,

1000 Western Ave., Lynn, MA 01910;

telephone (781) 594–3140, fax (781)

594–4805. Copies may be inspected at

the FAA, New England Region, Office of

the Assistant Chief Counsel, 12 New

England Executive Park, Burlington,

MA; or at the Office of the Federal

Register, 800 North Capitol Street, NW.,

suite 700, Washington, DC.

(g) This amendment becomes effective

on January 2, 1998.

Issued in Burlington, Massachusetts, on

November 26, 1997.

Jay J. Pardee,

Manager, Engine and Propeller Directorate,

Aircraft Certification Service.

[FR Doc. 97–31862 Filed 12–5–97; 8:45 am]

BILLING CODE 4910–13–U

DEPARTMENT OF TRANSPORTATION

Federal Aviation Administration

14 CFR Part 39

[Docket No. 97–ANE–04; Amendment 39–

10234; AD 97–25–10]

RIN 2120–AA64

Airworthiness Directives; Pratt &

Whitney JT9D Series Turbofan Engines

AGENCY: Federal Aviation

Administration, DOT.

ACTION: Final rule.
Federal Register / Vol. 62, No. 235 / Monday, December 8, 1997 / Rules and Regulations 64515

SUMMARY: This amendment supersedes

an existing airworthiness directive (AD),

applicable to certain Pratt & Whitney

JT9D series turbofan engines, that

currently requires initial and repetitive

eddy current inspection (ECI) or

fluorescent penetrant inspection (FPI)

for cracks in first stage high pressure

turbine (HPT) disk cooling air holes.

This amendment requires initial and

repetitive FPI for cracks in cooling air

holes of additional first stage HPT disks,

and replacement with serviceable parts.

In addition, this amendment requires

initial and repetitive FPI for cracks in tie

bolt holes of certain other affected

second stage HPT disks installed in PW

JT9D series turbofan engines. This

amendment is prompted by reports of a

cracked cooling air hole on one first

stage HPT disk, and a cracked tie bolt

hole on one second stage HPT disk. The

actions specified by this AD are

intended to prevent turbine disk failure

due to cooling air hole or tie bolt hole

cracking, which could result in an

uncontained engine failure and damage

to the aircraft.

DATES: Effective January 12, 1998.

ADDRESSES: The service information

referenced in this AD may be obtained

from Pratt & Whitney, 400 Main St., East

Hartford, CT 06108; telephone (860)

565–6600, fax (860) 565–4503. This

information may be examined at the

FAA, New England Region, Office of the

Assistant Chief Counsel, 12 New

England Executive Park, Burlington,

MA. This information may be examined

at the Federal Aviation Administration

(FAA), New England Region, Office of

the Assistant Chief Counsel, 12 New

England Executive Park, Burlington, MA

01803–5299; or at the Office of the

Federal Register, 800 North Capitol

Street, NW., suite 700, Washington, DC.

FOR FURTHER INFORMATION CONTACT: Tara

Goodman, Aerospace Engineer, Engine

Certification Office, FAA, Engine and

Propeller Directorate, 12 New England

Executive Park, Burlington, MA 01803–

5299; telephone (781) 238–7130, fax

(781) 238–7199.

SUPPLEMENTARY INFORMATION: A

proposal to amend part 39 of the Federal

Aviation Regulations (14 CFR part 39)

by superseding airworthiness directive

(AD) 91–04–10, Amendment 39–6859

(56 FR 5343, February 11, 1991),

applicable to Pratt & Whitney (PW) JT9D

series turbofan engines, was published

in the Federal Register on March 19,

1997 (62 FR 12979). That action

proposed to require initial and

repetitive fluorescent penetrant

inspections (FPI) for cracks in cooling

air holes of affected first stage high

pressure turbine (HPT) disks, and, if

necessary, replacement with serviceable

parts. In addition, the action proposed

to require initial and repetitive FPI for

cracks in tie bolt holes of all affected

second stage HPT disks. Finally, that

action proposed to require reporting

findings of cracked turbine disks.

Interested persons have been afforded

an opportunity to participate in the

making of this amendment. Due

consideration has been given to the

comments received.

One commenter states that since the

root cause of the crack which was found

on the cooling hole of the improved

disk, part number (P/N) 840301, was

caused by improper tooling application

(use of reamer instead of a carbide

insert) at the specified supplier, only the

suspect supplier and lots should be

affected by the AD. The commenter

maintains that if the FAA suspects PW’s

qualified carbide insert machining

process, every maintenance process

requiring the manufacturer’s

qualification should always require

FAA qualification in future. The Federal

Aviation Administration (FAA) does not

concur. The improved disk, P/N 840301,

is manufactured with the new reamer

tooling method; however, a disk of this

P/N was found with cooling hole

cracking. The reamer method was

introduced by manufacturers to

preclude abusive machining that was

found using the carbide insert method.

Investigation has shown that both the

carbide insert method and the reamer

method are subject to the same abusive

machining phenomenon. The FAA

approves manufacturing and

maintenance processes, which are

updated as necessary.

The same commenter requests that the

FAA and PW develop an inspection

procedure that can detect not only a

crack but detect whether a severely

worked layer of material exists or not,

so that unnecessary repetitive

inspections can be minimized by

removing those disks. The FAA does not

concur. It is not possible to detect 100%

of the possible cracking conditions in

the field using current inspection

methods. Therefore, repetitive

inspections are necessary for disks that

are in service because cracking can

propagate in fatigue from a layer of

severely worked material resulting from

the manufacturing process.

The same commenter requests the

FAA extend the initial inspection

requirement for disks that have been

previously inspected, noting that the AD

as proposed would allow 1,500 cycles in

service (CIS) before initial inspection for

disks never inspected while only 250

CIS for disks that have been inspected

but that have accumulated more than

3,500 CIS since last inspection. The

FAA does not concur. The current AD,

effective in 1991, requires repetitive

inspections of disks installed in JT9D–

59A, –70A, –7Q, and –7Q3 engines at

intervals not to exceed 3,500 CIS.

Therefore, the example cited by the

commenter should not occur, as disks

installed in JT9D–59A or –7Q engines

should never exceed 3,500 CIS since last

inspection. Only if a disk installed in a

JT9D–7R4 engine had exceeded 3,500

CIS since last inspection would the

requirement to inspect no later than 250

CIS after the effective date of this AD

apply. This inspection requirement was

considered in the risk analysis and the

FAA has determined that it is necessary.

The commenter does not indicate how

many JT9D–7R4 engines might be

affected by the 250 CIS initial

inspection requirement. Individual

operators can apply for an adjustment to

that compliance time under paragraph

(d) of the AD.

One commenter states that the

mandated use of FPI does not provide

all possible assurance that defective

HPT disks will be removed from service.

The commenter believes that eddy

current inspection, or a combination of

the two methods, would clearly provide

a greater probability of crack detection.

The FAA does not concur. The first

stage turbine cooling air holes and

second stage tie bolt holes have low

aspect ratios. The FAA has determined

that FPI of low aspect ratio holes is

adequate for detecting cracks in these

locations.

One commenter states that there

appears to be anomalies in the

requirements for disks that have been in

service for over 6,000 cycles since new

(CSN), as stated in paragraph (b)(1)(i) of

the compliance section. The commenter

maintains that as the paragraph reads in

the proposed rule, this inspection will

always occur later than accumulating

8,000 CSN if CSN is greater than 6,000

on the effective date of the AD. The

FAA concurs and has changed this

paragraph to require inspection within

2,000 CIS if a disk has over 6,000 CSN

on the effective date of the AD.

The same commenter suggests that at

next engine shop visit would be

sufficient as an interval for cooling hole

and tie bolt hole inspections. The FAA

does not concur. Since the timing of

engine shop visits varies widely

between operators, the use of shop visits

to define inspection intervals in ADs

does not provide adequate objectivity on

which to assess the effectiveness of the

required actions in addressing the

unsafe condition. The FAA has

determined that a maximum of 2,000

CIS interval is required.
64516 Federal Register / Vol. 62, No. 235 / Monday, December 8, 1997 / Rules and Regulations

Subsequent to the publication of the

proposed rule, two JT9D–7R4D/E

operators indicated that the 6,000 CIS

re-inspection interval for second stage

turbine hub tie bolt hole mismachining

would require them to remove engines

prematurely due to their high cycle

utilization rate. The FAA has

determined through a review of risk

analysis that the additional risk

involved in extending the re-inspection

interval from 6,000 CIS to 8,000 CIS is

sufficiently low and has changed the reinspection interval accordingly.

In addition, the manufacturer has

recommended that Special Process

Operation Procedure (SPOP) 70 be used

in lieu of SPOP 84 in order to permit the

inspection of the second stage turbine

hub tie bolt holes when the second stage

turbine rotor is removed from the HPT

module assembly without necessitating

the removal of the second stage turbine

blades. The FAA concurs and has

changed the reference to the inspection

procedure accordingly.

After careful review of the available

data, including the comments noted

above, the FAA has determined that air

safety and the public interest require the

adoption of the rule with the changes

described previously. The FAA has

determined that these changes will

neither increase the economic burden

on any operator nor increase the scope

of the AD.

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.

For the reasons discussed above, I

certify that this action (1) Is not a

‘‘significant regulatory action’’ under

Executive Order 12866; (2) is not a

‘‘significant rule’’ under DOT

Regulatory Policies and Procedures (44

FR 11034, February 26, 1979); and (3)

will not have a significant economic

impact, positive or negative, on a

substantial number of small entities

under the criteria of the Regulatory

Flexibility Act. A final evaluation has

been prepared for this action and it is

contained in the Rules Docket. A copy

of it 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

removing amendment 39–6859 (56 FR

5343, February 11, 1991) and by adding

a new airworthiness directive,

Amendment 39–10234, to read as

follows:

97–25–10 Pratt & Whitney: Amendment 39–

10234. Docket 97–ANE–04 Supersedes

airworthiness directive (AD) 91–04–10,

Amendment 39–6859.

Applicability: Pratt & Whitney (PW) JT9D–

59A, –70A, –7Q, –7Q3, –7R4D, –7R4D1,

–7R4E, and –7R4E1 (AI–500) series turbofan

engines, installed on but not limited to

Airbus Industrie A300 and A310, Boeing 747

and 767, and McDonnell Douglas DC–10

series aircraft.

Note 1: This AD applies to each engine

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

engines 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 (d) 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.

Compliance: Required as indicated, unless

accomplished previously.

To prevent turbine disk failure due to

cooling hole or tie bolt hole cracking, which

could result in an uncontained engine failure

and damage to the aircraft, accomplish the

following:

(a) For first stage high pressure turbine

(HPT) disks, part numbers (P/Ns) 768001,

792701, 812901, 819801, 840501, 840401,

840701, 840601, and 840301, installed in PW

JT9D–59A, –70A, –7Q, and –7Q3 engines,

accomplish the following:

(1) Disks that have not been fluorescent

penetrant inspected or eddy current

inspected since introduction into service,

perform an initial fluorescent penetrant

inspection (FPI) for cracks in all 40 cooling

air holes in accordance with PW Turbojet

Engine Standard Practices Manual, P/N

585005, Chapter/Section 70–33, Special

Process Operation Procedure (SPOP) 84, as

follows:

(i) Disks with 3,500 cycles since new (CSN)

or more on the effective date of this AD,

inspect prior to accumulating 5,000 CSN, or

within 1,500 cycles in service (CIS) after the

effective date of this AD, whichever occurs

later.

(ii) Disks with less than 3,500 CSN on the

effective date of this AD, inspect prior to

accumulating 5,000 CSN.

(2) Disks that have been reoperated in

accordance with PW Service Bulletin (SB)

No. 5815, Revision 2, dated July 31, 1992, or

prior revisions, that have not been

fluorescent penetrant inspected or eddy

current inspected since reoperation, perform

an initial FPI for cracks in all 40 cooling air

holes in accordance with PW Turbojet Engine

Standard Practices Manual, P/N 585005,

Chapter/Section 70–33, SPOP 84, as follows:

(i) Disks with 3,500 CIS or more since

reoperation on the effective date of this AD,

inspect prior to accumulating 5,000 CIS since

reoperation, or within 1,500 CIS after the

effective date of this AD, whichever occurs

later.

(ii) Disks with less than 3,500 CIS since

reoperation on the effective date of this AD,

inspect prior to accumulating 5,000 CIS since

reoperation.

(3) Disks that have been fluorescent

penetrant inspected, or eddy current

inspected, since introduction into service or

since reoperation, in accordance with PW SB

No. 5744, Revision 3, dated March 31, 1993,

or prior revisions, or PW JT9D–7Q, –7Q3

Engine Manual, P/N 777210, 72–51–00,

Inspection –03, or PW JT9D–59A, –70A

Engine Manual, P/N 754459, 72–51–00,

Heavy Maintenance Check –03, perform an

FPI for cracks in all 40 cooling air holes,

prior to accumulating 3,500 CIS since last FPI

or ECI, or within 250 CIS after the effective

date of this AD, whichever occurs later, in

accordance with PW Turbojet Engine

Standard Practices Manual, P/N 585005,

Chapter/Section 70–33, SPOP 84.

(4) Thereafter, perform FPI for cracks in all

40 cooling air holes at intervals not to exceed

3,500 CIS since last FPI, in accordance with

PW Turbojet Engine Standard Practices

Manual, P/N 585005, Chapter/Section 70–33,

SPOP 84.

(5) Prior to further flight, remove from

service cracked disks, and replace with

serviceable parts.

(b) For second stage HPT disks, P/N

5001802–01, installed in PW JT9D–7R4D,

–7R4D1, –7R4E, and –7R4E1 (AI–500)

engines, accomplish the following:

(1) Disks that have not been fluorescent

penetrant inspected since introduction into

service, perform an initial FPI for cracks in

all 30 tie bolt holes in accordance with PW

Turbojet Engine Standard Practices Manual,

P/N 585005, Chapter/Section 70–33, SPOP

70, as follows:

(i) Disks with 6,000 CSN or more on the

effective date of this AD, inspect within

2,000 CIS after the effective date of this AD.
Federal Register / Vol. 62, No. 235 / Monday, December 8, 1997 / Rules and Regulations 64517

(ii) Disks with less than 6,000 CSN on the

effective date of this AD, inspect prior to

accumulating 8,000 CSN.

(2) Disks that have been fluorescent

penetrant inspected since introduction into

service, perform an FPI for cracks in all 30

tie bolt holes, prior to accumulating 8,000

CIS since last FPI, or within 250 CIS after the

effective date of this AD, whichever occurs

later, in accordance with PW Turbojet Engine

Standard Practices Manual, P/N 585005,

Chapter/Section 70–33, SPOP 70.

(3) Thereafter, perform FPI for cracks in all

30 tie bolt holes at intervals not to exceed

8,000 CIS since last FPI, in accordance with

PW Turbojet Engine Standard Practices

Manual, P/N 585005, Chapter/Section 70–33,

SPOP 70.

(4) Prior to further flight, remove from

service cracked disks, and replace with

serviceable parts.

(c) Report findings of cracked turbine disks

within 48 hours after inspection to Tara

Goodman, Aerospace Engineer, Engine

Certification Office, FAA, Engine and

Propeller Directorate, 12 New England

Executive Park, Burlington, MA 01803–5299;

telephone (781) 238–7130, fax (781) 238–

7199, Internet: ‘‘Tara.Goodman@faa.dot.gov’’.

Reporting requirements have been approved

by the Office of Management and Budget and

assigned OMB control number 2120–0056.

(d) 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, Engine

Certification Office. Operators shall submit

their requests through an appropriate FAA

Principal Maintenance Inspector, who may

add comments and then send it to the

Manager, Engine Certification Office.

Note 2: Information concerning the

existence of approved alternative method of

compliance with this AD, if any, may be

obtained from the Engine Certification Office.

(e) 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 aircraft to

a location where the requirements of this AD

can be accomplished.

(f) This amendment becomes effective on

January 12, 1998.

Issued in Burlington, Massachusetts, on

November 28, 1997.

Jay J. Pardee,

Manager, Engine and Propeller Directorate,

Aircraft Certification Service.

[FR Doc. 97–31965 Filed 12–5–97; 8:45 am]

BILLING CODE 4910–13–U

DEPARTMENT OF TRANSPORTATION

Federal Aviation Administration

14 CFR Part 39

[Docket No. 97–ANE–31–AD; Amendment

39–10233; AD 97–25–09]

RIN 2120–AA64

Airworthiness Directives; Allison

Engine Company Model 250–C40B

Turboshaft Engines

AGENCY: Federal Aviation

Administration, DOT.

ACTION: Final rule; request for

comments.

SUMMARY: This amendment adopts a

new airworthiness directive (AD) that is

applicable to Allison Engine Company

Model 250–C40B turboshaft engines.

This action requires installation of a

placard requiring pilots to record torque

level and time in service operating

above 86% engine torque until the

defective parts have been replaced, no

later than December 31, 2000, or when

certain maintenance actions are

accomplished, or when certain

operational restrictions are exceeded,

whichever occurs earliest. This

amendment is prompted by a report

from Allison Engine Company of a

manufacturing defect in certain helical

power takeoff gearshaft assemblies,

identified by serial numbers. The

actions specified in this AD are

intended to prevent fatigue failure of the

helical power takeoff gearshaft

assembly, which could result in a loss

of engine power and inflight engine

shutdown.

DATES: Effective December 23, 1997.

The incorporation by reference of

certain publications listed in the

regulations is approved by the Director

of the Federal Register as of December

23, 1997.

Comments for inclusion in the Rules

Docket must be received on or before

February 6, 1998.

ADDRESSES: Submit comments in

triplicate to the Federal Aviation

Administration (FAA), New England

Region, Office of the Regional Counsel,

Attention: Rules Docket No. 97–ANE–

31–AD, 12 New England Executive Park,

Burlington, MA 01803–5299. Comments

may also be sent via the Internet using

the following address: ‘‘9-adengineprop@faa.dot.gov’’. Comments

sent via the Internet must contain the

docket number in the subject line.

The service information referenced in

this AD may be obtained from Allison

Engine Company, P.O. Box 420, Speed

Code U–15, Indianapolis, IN 46206–

0420; telephone: (317) 230–6674. This

information may be examined at the

FAA, New England Region, Office of the

Regional Counsel, 12 New England

Executive Park, Burlington, MA; or at

the Office of the Federal Register, 800

North Capitol Street, NW., suite 700,

Washington, DC.

FOR FURTHER INFORMATION CONTACT: John

Tallarovic, Aerospace Engineer, Chicago

Aircraft Certification Office, FAA, Small

Airplane Directorate, ACE–118C, 2300

East Devon Avenue, Des Plaines, IL

60018; telephone (847) 294–8180, fax

(847) 294–7834.

SUPPLEMENTARY INFORMATION: The

Federal Aviation Administration (FAA)

received a report from Allison Engine

Company of a manufacturing defect

discovered in certain helical power

takeoff gearshaft assemblies, part

number (P/N) 23056617, installed on

Model 250–C40B turboshaft engines.

The manufacturing defect was

discovered while measuring the depth

of the case hardening of the gear. The

manufacturing defect was caused by

excessive removal of case hardened

material from the gear during

manufacturing. This condition, if not

corrected, could result in fatigue failure

of the helical power takeoff gearshaft

assembly, which could result in a loss

of engine power and inflight engine

shutdown.

The FAA has reviewed and approved

the technical contents of Allison Alert

Commercial Engine Bulletin (CEB) No.

A–72–5009, dated May 21, 1997, that

lists by serial number (S/N) 49 affected

engines, gearboxes, and gears. This CEB

also describes procedures for

replacement of affected helical power

takeoff gearshaft assemblies with

serviceable parts.

Since an unsafe condition has been

identified that is likely to exist or

develop on other engines of the same

type design, this AD is being issued to

prevent a loss of engine power and

inflight engine shutdown. This AD

requires installation of a placard

requiring pilots to record torque and

time in service operating above 86%

engine torque until replacement of

defective helical power takeoff gearshaft

assemblies with serviceable parts, and

then the placard can be removed. The

compliance times were determined

based upon an analysis of the effect of

gearbox assembly torque loading on

component life. The actions are required

to be accomplished in accordance with

the CEB described previously.

The operational limitations imposed

by this AD have been coordinated with

the Rotorcraft Directorate.