Federal Register / Vol. 60, No. 244 / Wednesday, December 20, 1995 / Rules and Regulations 65519

Identify Operational Usage Cycles-Per-Hour:

1–6 Operational

Cycles-Per-Hour lllllllllll

7–19 Operational

Cycles-Per-Hour lllllllllll

20–Above Operational

Cycles-Per-Hour lllllllllll

Next Inspection Date (Estimated): lllll

and Flight Hours (Estimated): lllllll

Magnetic Particle Inspection (MPI) Results

(this inspection): Passed lll

Failed lll

If a crack is found, indicate the approximate

location on the part and the length of the

crack in inches: llllllll

Total Time-In-Service (TIS) (Hours):

Estimated lllllllllllllll

Actual llllllllllllllll

Unknown lllllllllllllll

At Retirement lllllllllllll

Inspection results at retirement (if known):

MPI Passed lll Failed lll

Visual Passed lllFailed lll

Log Book Entry for Part No. llllll,

Serial No. llllll, is (date) llll,

at Retirement Hours llllll. This

part’s Serial No. has been marked

unairworthy and unfit for further service on

(date) llllll, 199 ll.

Issued in Fort Worth, Texas, on December

13, 1995.

Daniel P. Salvano,

Manager, Rotorcraft Directorate, Aircraft

Certification Service.

[FR Doc. 95–30771 Filed 12–19–95; 8:45 am]

BILLING CODE 4910–13–U

14 CFR Part 39

[Docket No. 95–SW–21–AD; Amendment

39–9466; AD 95–26–08]

Airworthiness Directives; Sikorsky

Aircraft Model S–58A, S–58B, S–58C,

S–58D, S–58E, S–58F, S–58G, S–58H,

S–58J, S–58BT, S–58DT, S–58ET, S–

58FT, S–58HT, and S–58JT Helicopters

AGENCY: Federal Aviation

Administration, DOT.

ACTION: Final rule; request for

comments.

SUMMARY: This amendment adopts a

new airworthiness directive (AD) that is

applicable to Sikorsky Aircraft Model

S–58A, S–58B, S–58C, S–58D, S–58E,

S–58F, S–58G, S–58H, S–58J, S–58BT,

S–58DT, S–58ET, S–58FT, S–58HT, and

S–58JT helicopters. This action requires

initial and repetitive magnetic particle

inspections of the main rotor shaft

(shaft) for cracks, and defines power

limitations for certain helicopter

operations. This amendment is

prompted by a recent accident in which

a shaft failed, resulting in loss of power.

Subsequent inspections on other aircraft

of the same type revealed cracks in four

additional shafts. The actions specified

in this AD are intended to prevent

failure of the shaft, loss of power to the

rotor system, and subsequent loss of

control of the helicopter.

DATES: Effective January 4, 1996.

Comments for inclusion in the Rules

Docket must be received on or before

February 20, 1996.

ADDRESSES: Submit comments in

triplicate to the Federal Aviation

Administration (FAA), Office of the

Assistant Chief Counsel, Attention:

Rules Docket No. 95–SW–21–AD, 2601

Meacham Blvd., Room 663, Fort Worth,

Texas 76137.

FOR FURTHER INFORMATION CONTACT: Mr.

Francis X. Walsh, Aerospace Engineer,

FAA, Boston Aircraft Certification

Office, 12 New England Executive Park,

Burlington, Massachusetts 01803–5299,

telephone (617) 238–7158, fax (617)

238–7199.

SUPPLEMENTARY INFORMATION: This

amendment adopts a new AD that is

applicable to Sikorsky Aircraft Model

S–58A, S–58B, S–58C, S–58D, S–58E,

S–58F, S–58G, S–58H, S–58J, S–58BT,

S–58DT, S–58ET, S–58FT, S–58HT, and

S–58JT helicopters with shaft assembly,

part number (P/N) S1635–20059–2,

installed. This AD is prompted by an

accident in which the failure of a shaft

resulted in the crash of a helicopter.

Since that accident, inspections have

revealed cracks in four additional shafts.

The shaft transmits power to the main

rotor system to provide lift for the

helicopter. Failure of this shaft results

in loss of power to the main rotor

system and subsequent loss of control of

the helicopter. Due to the criticality of

the shaft, this AD must be issued

immediately to correct an unsafe

condition in the affected helicopters.

Since an unsafe condition has been

identified that is likely to exist or

develop on other Sikorsky Aircraft

Model S–58A, S–58B, S–58C, S–58D, S–

58E, S–58F, S–58G, S–58H, S–58J, S–

58BT, S–58DT, S–58ET, S–58FT, S–

58HT, and S–58JT helicopters of the

same type design, this AD is being

issued to prevent failure of the shaft,

loss of power to the rotor system, and

subsequent loss of control of the

helicopter. This AD requires

determining the operational cycles-perhour on the helicopters, removing the

shaft assembly from the main gear box,

and inspecting the shaft for cracks using

a magnetic particle inspection method

within the next 50 hours time-in-service

(TIS). Following this initial inspection,

repetitive magnetic particle inspections

are required. Additionally, this AD

prescribes operating limitations for

certain helicopter operations.

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 should 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 No. 95–SW–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 that it is not a ‘‘significant
65520 Federal Register / Vol. 60, No. 244 / Wednesday, December 20, 1995 / Rules and Regulations

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, 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 USC 106(g), 40101, 40113,

44701.

§ 39.13 [Amended]

2. Section 39.13 is amended by

adding a new airworthiness directive to

read as follows:

AD 95–26–08 Sikorsky Aircraft: Amendment

39–9466. Docket No. 95–SW–21–AD.

Applicability: Model S–58A, S–58B, S–

58C, S–58D, S–58E, S–58F, S–58G, S–58H,

S–58J, S–58BT, S–58DT, S–58ET, S–58FT, S–

58HT, and S–58JT helicopters with main

rotor shaft assembly (shaft assembly), part

number (P/N) S1635–20059–2, installed,

certificated in any category.

Note 1: The shaft assembly consists of a

main rotor shaft, P/N S1635–20059; an upper

end plug, P/N S1635–20153; and a lower end

plug, P/N S1635–20154. The shaft assembly

P/N (S1635–20059–2) is marked on the edge

of the main rotor shaft lower flange.

Note 2: This AD applies to each helicopter

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

helicopters that have been modified, altered,

or repaired so that the performance of the

requirements of this AD is affected, the

owner/operator must use the authority

provided in paragraph (j) to request approval

from the FAA. This approval may address

either no action, if the current configuration

eliminates the unsafe condition, or different

actions necessary to address the unsafe

condition described in this AD. Such a

request should include an assessment of the

effect of the changed configuration on the

unsafe condition addressed by this AD. In no

case does the presence of any modification,

alteration, or repair remove any helicopter

from the applicability of this AD.

Compliance: Required as indicated, unless

accomplished previously.

To prevent failure of the main rotor shaft

(shaft) loss of power to the rotor system, and

subsequent loss of control of the helicopter,

accomplish the following:

(a) From available helicopter records,

determine the maximum number of actual

operational cycles-per-hour of the current

shaft assembly since installation. An

operational cycle is defined as one

turnaround (external lift cycle) for external

load operations, and as one takeoff and one

landing for internal load operations. A

turnaround is defined as picking up an

external load, transporting that load to a

drop-off point, releasing the load, and flying

to the next load pickup point. If the

maximum number of actual operational

cycles-per-hour cannot be determined, use

25-operational cycles-per-hour as the

maximum operational cycles-per-hour for

purposes of this AD. Record the determined

number of operational usage cycles-per-hour

of the shaft assembly in the appropriate

aircraft maintenance records.

(1) If the maximum operational cycles-perhour has ever equaled or exceeded 20 cyclesper-hour, inspect in accordance with

paragraph (b) of this AD within the next 50

hours time-in-service (TIS), unless previously

accomplished within the last 200 hours TIS.

(2) If the maximum operational cycles-perhour has never exceeded 19 cycles-per-hour,

inspect the shaft in accordance with

paragraph (b) of this AD within the next 50

hours TIS, unless previously accomplished.

(b) Remove the shaft assembly, P/N S1635–

20059–2, from the main gear box. Remove the

upper end plug, P/N S1635–20153 and lower

end plug, P/N S1635–20154, from the shaft

assembly, and conduct a magnetic particle

inspection (MPI) of the shaft for cracks in

accordance with MIL–STD–1949 or ASTM E–

1444. Pay particular attention to the inside

diameter of the 0.7515–0.7510-inch diameter

dowel pin holes in the flange and adjacent

flange surfaces.

Note 3: Section 2D of Sikorsky Aircraft

Alert Service Bulletin 58B35–34, dated June

9, 1995, contains a procedure for conducting

a MPI of the shaft (in agreement with MIL–

STD–1949 or ASTM E–1444).

(c) Conduct repetitive MPI’s of the shaft for

cracks as follows:

(1) If the maximum operational cycles-perhour has ever equaled or exceeded 20 cyclesper-hour, repeat the MPI at intervals not to

exceed 250 hours TIS from the date of the

last inspection.

(2) If the maximum operational cycles-perhour exceeds 6 cycles-per-hour, but has

always been less than 20 cycles-per-hour,

repeat the MPI at 1,250 hours TIS, and

thereafter at intervals not to exceed 250 hours

TIS from the date of the last inspection. If the

last inspection was accomplished between

1,000 hours TIS and 1,250 hours TIS, begin

the repetitive inspections within 250 hours

TIS from the date of the last inspection

instead of at 1,250 hours TIS.

(3) If the maximum operational cycles-perhour has never exceeded 6 cycles-per-hour,

repeat the MPI at 1,250 hours TIS. If the last

inspection was accomplished between 1,000

hours TIS and 1,250 hours TIS, repeat the

MPI within 250 hours TIS from the date of

the last inspection instead of at 1,250 hours

TIS.

(d) Report all inspection results to the

Manager, Boston Aircraft Certification Office,

using the Attachment provided later in this

AD. Reporting requirements have been

approved by the Office of Management and

Budget and assigned OMB control number

2120–0056.

(e) If no crack is discovered, replace the

upper and lower end plugs and reinstall the

shaft assembly into the main gearbox.

(f) If any crack is discovered or on or before

the shaft assembly reaches 2,500 hours TIS,

replace the shaft assembly with an airworthy

shaft assembly, P/N S1635–20059–2. If the

replacement shaft has previously been in

service, determine the maximum operational

cycles-per-hour in accordance with

paragraph (a) and inspect in accordance with

this AD.

Note 4: In accordance with the applicable

maintenance manual, 2,500 hours TIS is the

mandatory retirement life for the shaft

assembly, P/N S1635–20059–2.

(g) If the main rotor shaft assembly

installed on the helicopter has ever equaled

or exceeded 20 or more operational cyclesper-hour, insert the following restrictions

into the Limitations section of the Rotorcraft

Flight Manual:

(1) For turbine engine installations: ‘‘The

main rotor shaft assembly installed on this

helicopter has been operated at 20 or more

cycles-per-hour. Engine power is restricted to

maximum continuous power at 93%Nf.

Takeoff power operations are prohibited.’’

(2) For reciprocating engine installations:

‘‘The main rotor shaft assembly installed on

this helicopter has been operated at 20 or

more cycles-per-hour. Engine power is

restricted to maximum continuous power at

2,500 RPM. Takeoff power operations are

prohibited.’’

(h) If the main rotor shaft assembly

installed on the helicopter has ever equaled

or exceeded 20 or more operational cyclesper-hour, install on the instrument panel,

adjacent to the pilot’s engine (Nf or RPM)

tachometer, torquemeter, or manifold

pressure gauges, a placard made of material

that is not easily erased, disfigured, or

obscured that contains the following

statement in lettering of 0.2 inch minimum

height and stated in one or two lines:

(1) For turbine engine installations: ‘‘MAX

PWR: 101% Q AT 93% Nf’’

(2) For reciprocating engine installations:

‘‘MAX PWR: 47.5 IN. HG at 2,500 RPM’’

(i) Continue to record operational cyclesper-hour of the shaft assembly in the

appropriate maintenance records. If

operational cycles-per-hour increases on an

affected shaft assembly to the extent that it

places the shaft assembly into a higher

cycles-per-hour usage group, the applicable

requirements and limitations contained in

this AD for the higher usage group apply to

that shaft assembly. A replacement shaft

assembly must comply with all requirements
Federal Register / Vol. 60, No. 244 / Wednesday, December 20, 1995 / Rules and Regulations 65521

and limitations of this AD as applicable. If

the number of operational cycles-per-hour

determined for a replacement shaft assembly

does not equal or exceed 20 cycles-per-hour,

the Rotorcraft Flight Manual limitation

specified in paragraph (g) and the placard

specified in paragraph (h) may be removed.

(j) An alternative method of compliance or

adjustment of the compliance time that

provides an acceptable level of safety may be

used when approved by the Manager, Boston

Aircraft Certification Office, FAA, New

England Region. Operators shall submit their

requests through an FAA Principal

Maintenance Inspector, who may concur or

comment and then send it to the Manager,

Boston Aircraft Certification Office.

Note 5: Information concerning the

existence of approved alternative methods of

compliance with this AD, if any, may be

obtained from the Boston Aircraft

Certification Office.

(k) 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 helicopter

to a location where the requirements of this

AD can be accomplished.

(l) This amendment becomes effective on

January 4, 1996.

Attachment—Inspection Results Report

The following information must be

reported as soon as possible, but no later

than 7 days after inspection, to:

Manager, Boston Aircraft Certification

Office, Engine and Propeller Directorate,

Aircraft Certification Service, Federal

Aviation Administration, 12 New

England Executive Park, Burlington, MA

01803–5299, FAX: (617) 238–7199.

Operator/Repair Station lllll

Aircraft Model No. lllll

Aircraft Serial No. lllll

Date of Inspection lllll

Main Rotor Part No. lllll

Main Rotor Serial No. lllll

Type of Aircraft Utilization:

Passenger Carry lllll

Utility/Construction lllll

Firefighting lllll

Logging lllll

Other lllll

Identify Operational Usage Cycles-PerHour:

1–6 Operational Cycles-Per-Hour

lllll

7–19 Operational Cycles-Per-Hour

lllll

20–Above Operational Cycles-PerHour lllll

Next Inspection Date (Estimated):

lllll

and Flight Hours (Estimated):

lllll

Magnetic Particle Inspection (MPI)

Results (this inspection):

Passed lllll Failed lllll

If a crack is found, indicate the

approximate location on the part

and the length of the crack in

inches: llllll

Total Time-In-Service (TIS) (Hours):

Estimated lllll

Actual lllll

Unknown lllll

At Retirement lllll

Inspection results at retirement (if

known):

MPI Passed lllll

Failed lllll

Visual Passed lllll

Failed lllll

Log Book Entry for Part No. lllll,

Serial No. lllll, is (date)

lllll, at Retirement Hours

lllll. This part’s Serial No.

has been marked unairworthy and

unfit for further service on (date)

lllll, 199 ll.

Issued in Fort Worth, Texas, on December

13, 1995.

Daniel P. Salvano,

Manager, Rotorcraft Directorate, Aircraft

Certification Service.

[FR Doc. 95–30772 Filed 12–19–95; 8:45 am]

BILLING CODE 4910–13–U

14 CFR Part 39

[Docket No. 95–NM–238–AD; Amendment

39–9465; AD 95–26–07]

Airworthiness Directives; Bombardier

Model CL–600–2B19 (Regional Jet

Series 100) 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 Bombardier Model

CL–600–2B19 series airplanes. This

action requires revising the Limitations

Section of the Airplane Flight Manual to

provide the flight crew with procedures

to check the travel range of the aileron.

This action also requires inspection for

damage of the shear pins of the aileron

flutter damper and aileron hinge

fittings, and various follow-on actions.

This amendment is prompted by reports

of failure of shear pins in the aileron

flutter damper. The actions specified in

this AD are intended to prevent damage

to the aileron hinge fittings due to failed

shear pins, which subsequently could

cause reduced controllability of the

airplane.

DATES: Effective January 4, 1996.

The incorporation by reference of

certain publications listed in the

regulations is approved by the Director

of the Federal Register as of January 4,

1996.

Comments for inclusion in the Rules

Docket must be received on or before

February 20, 1996.

ADDRESSES: Submit comments in

triplicate to the Federal Aviation

Administration (FAA), Transport

Airplane Directorate, ANM–103,

Attention: Rules Docket No. 95–NM–

238–AD, 1601 Lind Avenue, SW.,

Renton, Washington 98055–4056.

The service information referenced in

this AD may be obtained from

Bombardier, Inc., Canadair, Aerospace

Group, P.O. Box 6087, Station Centreville, Quebec H3C 3G9, Canada. This

information may be examined at the

FAA, Transport Airplane Directorate,

1601 Lind Avenue, SW., Renton,

Washington; or at the FAA, New York

Aircraft Certification Office, Engine and

Propeller Directorate, 10 Fifth Street,

Third Floor, Valley Stream, New York;

or at the Office of the Federal Register,

800 North Capitol Street, NW., suite

700, Washington, DC.

FOR FURTHER INFORMATION CONTACT:

Franco Pieri, Aerospace Engineer,

Airframe Branch, ANE–172, FAA,

Engine and Propeller Directorate, New

York Aircraft Certification Office, 10

Fifth Street, Third Floor, Valley Stream,

New York 11581; telephone (516) 256–

7526; fax (516) 568–2716.

SUPPLEMENTARY INFORMATION: Transport

Canada Aviation, which is the

airworthiness authority for Canada,

recently notified the FAA that an unsafe

condition may exist on certain

Bombardier Model CL–600–2B19

(Regional Jet Series 100) series

airplanes. Transport Canada Aviation

advises that it has received reports

indicating that the shear pins of the

aileron flutter damper had failed.

Investigation revealed that the shear

pins had sheared off and migrated out,

which subsequently damaged the

aileron hinge fittings. This condition, if

not corrected, could result in reduced

controllability of the airplane.

Bombardier has issued Canadair

Regional Jet Alert Service Bulletin S.B.

A601R–27–058, Revision ‘‘A,’’ dated

September 8, 1995, which describes

procedures for:

1. A visual inspection to detect

damage of the shear link, the shear pin,

and the aileron attachment fitting;

2. Repair of the aileron attachment

fitting, if necessary;

3. For airplanes on which any

damaged shear pin is found, removal of

the aileron flutter dampers, the shear

links, the pivots, and the attaching

hardware;

4. For certain airplanes on which no

damaged shear pin is found, repetitive

visual inspections to detect damage of

the shear link, the shear pin, and the

aileron attachment fitting until the

aileron flutter dampers are removed.