70092 Federal Register / Vol. 75, No. 221 / Wednesday, November 17, 2010 / Rules and Regulations 

It must be demonstrated by test, or 

combination of test and analysis, that the 

airplane can continue safe flight and landing 

with inoperative normal engine- and APUgenerated electrical power (for example, 

without electrical power from any source, 

except for the battery and any other standby 

electrical sources). The airplane operation 

should be considered at the critical phase of 

flight and include the ability to restart the 

engines and maintain flight for the maximum 

diversion time capability being certified. 

Issued in Renton, Washington on 

November 5, 2010. 

Jeffrey Duven, 

Acting Manager, Transport Airplane 

Directorate, Aircraft Certification Service. 

[FR Doc. 2010–28998 Filed 11–16–10; 8:45 am] 

BILLING CODE 4910–13–P 

DEPARTMENT OF TRANSPORTATION 

Federal Aviation Administration 

14 CFR Part 25 

[Docket No. NM434; Notice No. 25–412–SC] 

Special Conditions: Bombardier Inc. 

Model CL–600–2E25 Airplane, 

Interaction of Systems and Structures 

AGENCY: Federal Aviation 

Administration (FAA), DOT. 

ACTION: Final special conditions; request 

for comments. 

SUMMARY: These special conditions are 

issued for the Bombardier Inc. Model 

CL–600–2E25 airplane. This airplane 

will have a novel or unusual design 

feature associated with the ruddertraveler limiting system controlling the 

command-by-wire (CBW) rudder. This 

system can serve to alleviate loads in 

the airframe but, in a failure state, can 

create loads in the airframe. The 

applicable airworthiness regulations do 

not contain adequate or appropriate 

safety standards for this design feature. 

These special conditions contain the 

additional safety standards that the 

Administrator considers necessary to 

establish a level of safety equivalent to 

that established by the existing 

airworthiness standards. 

DATES: The effective date of these 

special conditions is November 5, 2010. 

We must receive your comments by 

January 3, 2011. 

ADDRESSES: You must mail two copies 

of your comments to: Federal Aviation 

Administration, Transport Airplane 

Directorate, Attn: Rules Docket (ANM– 

113), Docket No. NM434, 1601 Lind 

Avenue, SW., Renton, Washington 

98057–3356. You may deliver two 

copies to the Transport Airplane 

Directorate at the above address. You 

must mark your comments: Docket No. 

NM434. You can inspect comments in 

the Rules Docket weekdays, except 

Federal holidays, between 7:30 a.m. and 

4 p.m. 

FOR FURTHER INFORMATION CONTACT: 

Todd Martin, FAA, ANM–115, 

Transport Airplane Directorate, Aircraft 

Certification Service, 1601 Lind 

Avenue, SW., Renton, Washington 

98057–3356; telephone (425) 227–1178; 

facsimile (425) 227–1149. 

SUPPLEMENTARY INFORMATION: The FAA 

has determined that notice of, and 

opportunity for prior public comment 

on, these special conditions are 

impracticable because these procedures 

would significantly delay issuance of 

the design approval and thus delivery of 

the affected aircraft. In addition, the 

substance of these special conditions 

has been subject to the public-comment 

process in several prior instances with 

no substantive comments received. The 

FAA therefore finds that good cause 

exists for making these special 

conditions effective upon issuance. 

Comments Invited 

We invite interested people to take 

part in this rulemaking by sending 

written comments, data, or views. The 

most helpful comments reference a 

specific portion of the special 

conditions, explain the reason for any 

recommended change, and include 

supporting data. We ask that you send 

us two copies of written comments. 

We will file in the docket all 

comments we receive, as well as a 

report summarizing each substantive 

public contact with FAA personnel 

concerning these special conditions. 

You can inspect the docket before and 

after the comment closing date. If you 

wish to review the docket in person, go 

to the address in the ADDRESSES section 

of this preamble between 7:30 a.m. and 

4 p.m., Monday through Friday, except 

Federal holidays. 

We will consider all comments we 

receive on or before the closing date for 

comments. We will consider comments 

filed late if it is possible to do so 

without incurring expense or delay. We 

may change these special conditions 

based on the comments we receive. 

If you want us to acknowledge receipt 

of your comments on these special 

conditions, include with your 

comments a self-addressed, stamped 

postcard on which you have written the 

docket number. We will stamp the date 

on the postcard and mail it back to you. 

Background 

On February 28, 2007, Bombardier 

Inc. applied for an amendment to Type 

Certificate No. A21EA, through 

Transport Canada, to include the new 

Model CL–600–2E25 airplane. The CL– 

600–2E25, which is a derivative of the 

CL–600–2D24 currently approved under 

Type Certificate No. A21EA, is to be 

certified for a maximum occupancy of 

110 people, including 5 crewmembers. 

The CL–600–2E25 has increased gross 

weight, extended wing tip, and 

increased fuselage length to 

accommodate the additional passengers 

as compared to the CL–600–2D24. 

The CL–600–2E25 will have a CBW 

rudder-control system that will affect 

the structural performance of the 

airplane. The airplane will use CBW 

Rudder Electronic Control Unit (ECU) 

software as a replacement for the 

Rudder Travel Limiter to limit rudder 

commands. The CBW Rudder ECU 

controls the rudder, trim, and yaw 

damping as well. This system can serve 

to alleviate loads in the airframe but, in 

a failure state, can create loads in the 

airframe. The current rules do not 

adequately account for the effects of this 

system and its failures on structural 

performance. The special conditions 

defined herein provide the criteria to be 

used in assessing the effects of this 

system on structures. 

Type Certification Basis 

Under the provisions of § 21.101, 

Bombardier Inc. must show that the 

Model CL–600–2E25 airplane meets the 

applicable provisions of Title 14, Code 

of Federal Regulations (14 CFR) part 25, 

as amended by Amendments 25–1 

through 25–119, except for earlier 

amendments as agreed upon by the 

FAA. These regulations will be 

incorporated into Type Certificate No. 

A21EA after type-certification approval 

of the Model CL–600–2E25. The 

regulations incorporated by reference in 

the type certificate are commonly 

referred to as the ‘‘original typecertification basis.’’ The regulations 

incorporated by reference in Type 

Certificate No. A21EA are as follows: 

The original type-certification basis 

for the Model CL–600–2D24 (CRJ 900), 

shown on TCDS A21EA, Revision 25, 

and reprinted below. 

Model CL–600–2D15/CL–600–2D24 

Part 25, including Amendments 25–1 

through 25–86, Amendments 25–88 

through Amendments 25–90, and 

Amendments 25–92 through 25–98 with 

the following exceptions: 

• Section 25.783(f) at Amendment 

25–23 shall replace § 25.783(f) at 

Amendment 25–88 for the Aft Cargo 

Compartment and Main Avionics Bay 

Doors only (common doors with CL– 

600–2C10 (CRJ–700); 


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• Section 25.807(d)(6) at Amendment 

25–72 shall replace § 25.807(h) at 

Amendment 25–94; 

• Sections 25.365, 25.831(a), and 

25.1447(c) at Amendment 25–87. Part 

25, Amendment 25–91, is not included 

in the type-certification basis. 

Additional FAA Requirements for 

Model CL–600–2D15/CL–600–2D24 

1. 14 CFR part 36, effective September 

10, 1990, and including all 

amendments effective on the date of 

type certification. 

2. 14 CFR part 34, effective September 

10, 1990, and including all 

amendments effective on the date of 

type certification. 

3. Special Conditions: 

(a) High Intensity Radiated Fields, No. 

25–ANM–109, dated October 31, 

1995. 

(b) Go-around Performance Credit for 

Use of Automatic Power Reserve 

(APR), No. 25–167–SC, dated 

October 24, 2000 (same as CL–600– 

2C10). 

(c) Sudden Engine Stoppage, No. 25– 

217–SC, dated October 04, 2002. 

(d) Passenger Seats with Nontraditional, Large, Non-metallic 

Panels, No. 25–384–SC, dated 

August 12, 2009. 

4. Exemptions: Exemption No. 7447, 

hydraulic-systems testing per 14 

CFR 25.1435(b)(1). Equivalent 

safety has been established for the 

following requirements: 

CL–600–2D15/CL–600–2D24 

1. Section 25.103 and others, Reduced 

Minimum Operating Speed Factors. 

2. Section 25.811(d)(2), Main Door Exit 

Marking Sign. 

3. Section 25.813(c)(2)(i), Emergency 

Exit Access. 

4. Section 25.904, Performance Credit 

for Use of APR During Reduced 

Thrust Takeoff. 

5. Section 25.933(a)(1)(ii), Thrust 

Reverser System. 

6. Section 25, appendix I, § 25.5(b)(4), 

Lack of On/Off Switch for 

Automatic Takeoff Thrust Control 

System (ATTCS). 

7. Section 25.841(b)(6), High Altitude 

Takeoff and Landing Operations 

documented in Transport Airplane 

Directorate ELOS Memo 

AT2587NY–T, dated January 31, 

2007. 

In addition, the certification basis 

includes other regulations, special 

conditions, and exemptions that are not 

relevant to these special conditions. 

Type Certificate No. A21EA will be 

updated to include a complete 

description of the certification basis for 

this airplane model. 

If the Administrator finds that the 

applicable airworthiness regulations 

(i.e., 14 CFR part 25) do not contain 

adequate or appropriate safety standards 

for the CL–600–2E25 because of a novel 

or unusual design feature, special 

conditions are prescribed under the 

provisions of 14 CFR 21.16. 

In addition to the applicable 

airworthiness regulations and special 

conditions, the CL–600–2E25 must 

comply with the fuel-vent and exhaustemission requirements of 14 CFR part 

34, and the noise-certification 

requirements of 14 CFR part 36. 

The FAA issues special conditions, as 

defined in 14 CFR 11.19, in accordance 

with § 11.38, and they become part of 

the type-certification basis under 14 

CFR 21.101. 

Special conditions are initially 

applicable to the model for which they 

are issued. Should the type certificate 

for that model be amended later to 

include any other model that 

incorporates the same or similar novel 

or unusual design feature, or should any 

other model already included on the 

same type certificate be modified to 

incorporate the same or similar novel or 

unusual design feature, the special 

conditions would also apply to the other 

model under § 21.101. 

Novel or Unusual Design Features 

The Bombardier Model CL–600–2E25 

airplane will incorporate the following 

novel or unusual design features: 

The CL–600–2E25 airplane will have 

a CBW rudder-control system that will 

affect the structural performance of the 

airplane. The airplane will use a CBW 

Rudder ECU software as a replacement 

for the rudder-travel limiter to limit 

rudder commands. The CBW Rudder 

ECU controls the rudder, trim, and yaw 

damping as well. 

Discussion 

This CBW system can affect the 

airplane’s structural performance, either 

directly or as a result of failure or 

malfunction. That is, the CBW system 

affects how the airplane responds in 

maneuver and gust conditions, and 

thereby affects the airplane’s structural 

capability. Such systems represent a 

novel and unusual feature when 

compared to the technology envisioned 

in the current airworthiness standards. 

Special conditions are needed to require 

consideration of the effects of the 

system on the structural capability and 

aeroelastic stability of the airplane, both 

in the normal and in the failed state. 

These special conditions require that 

the airplane meet the structural 

requirements of subparts C and D of 14 

CFR part 25 when the airplane systems 

are fully operative. These special 

conditions also require that the airplane 

meet these requirements considering 

failure conditions. In some cases, these 

special conditions allow reduced 

margins (in terms of speed margins and 

factors of safety) for failure conditions, 

as a function of system reliability. 

The Administrator considers these 

special conditions necessary to establish 

a level of safety equivalent to that 

established by the existing 

airworthiness standards. 

Applicability 

As discussed above, these special 

conditions are applicable to the Model 

CL–600–2E25. Should Bombardier Inc. 

apply at a later date for a change to the 

type certificate to include another 

airplane model incorporating the same 

novel or unusual design feature, these 

special conditions would apply to that 

model as well. 

Conclusion 

This action affects only certain novel 

or unusual design features on one model 

of airplane. It is not a rule of general 

applicability. 

Under standard practice, the effective 

date of final special conditions would 

be 30 days after the date of publication 

in the Federal Register. However, as the 

certification date for the Model CL–600– 

2E25 is imminent, the FAA finds that 

good cause exists to make these special 

conditions effective upon issuance. 

List of Subjects in 14 CFR Part 25 

Aircraft, Aviation safety, Reporting 

and recordkeeping requirements. 

■ The authority citation for these 

special conditions is as follows: 

Authority: 49 U.S.C. 106(g), 40113, 44701, 

44702, 44704. 

The Special Conditions 

■ Accordingly, pursuant to the authority 

delegated to me by the Administrator, 

the following special conditions are 

issued as part of the type-certification 

basis for Bombardier Model CL–600– 

2E25 airplanes modified according to 

DCA 0145–000–00020–2008/FAA (latest 

revision approved by the FAA). 

1. CWB Rudder-Control-System Special 

Conditions 

The Bombardier Model CL–600–2E25 

airplane is equipped with systems that 

affect the airplane’s structural 

performance either directly or as a result 

of failure or malfunction. The influence 

of these systems and their failure 

conditions must be taken into account 

when showing compliance with 

requirements of 14 CFR part 25, 

subparts C and D. The following criteria 


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must be used for showing compliance 

with these special conditions for 

airplanes equipped with flight-control 

systems, autopilots, stabilityaugmentation systems, load-alleviation 

systems, flutter-control systems, fuelmanagement systems, and other systems 

that either directly, or as a result of 

failure or malfunction, affect structural 

performance. If these special conditions 

are used for other systems, it may be 

necessary to adapt the criteria to the 

specific systems. 

(a) The criteria defined here address 

only direct structural consequences of 

system responses and performances. 

They cannot be considered in isolation 

but should be included in the overall 

safety evaluation of the airplane. They 

may, in some instances, duplicate 

standards already established for this 

evaluation. These criteria are only 

applicable to structure the failure of 

which could prevent continued safe 

flight and landing. Specific criteria 

defining acceptable limits on handling 

characteristics or stability requirements, 

when operating in the system-degraded 

or inoperative mode, are not provided in 

these special conditions. 

(b) Depending on the specific 

characteristics of the airplane, 

additional studies may be required, 

which go beyond the criteria provided 

in these special conditions, to 

demonstrate the capability of the 

airplane to meet other realistic 

conditions such as alternative gust 

conditions or maneuvers for an airplane 

equipped with a load-alleviation 

system. 

(c) The following definitions are 

applicable to these special conditions: 

(1) Structural performance: The 

capability of the airplane to meet the 

structural requirements of part 25. 

(2) Flight limitations: Limitations that 

can be applied to the airplane flight 

conditions following an in-flight failure 

occurrence, and that are included in the 

flight manual (speed limitations or 

avoidance of severe weather conditions, 

for example). 

(3) Operational limitations: 

Limitations, including flight limitations, 

that can be applied to the airplane 

operating conditions before dispatch, 

and which include, for example, fuel, 

payload, and master minimumequipment-list limitations. 

(4) Probabilistic terms: Terms, 

including probable, improbable, and 

extremely improbable, used in these 

special conditions and which are the 

same as those probabilistic terms used 

in § 25.1309. 

(5) Failure condition: The same term 

as used in § 25.1309. However, in these 

special conditions, the term ‘‘failure 

condition’’ applies only to systemfailure conditions that affect structural 

performance of the airplane. Examples 

are system-failure conditions that 

induce loads, change the response of the 

airplane to inputs such as gusts or pilot 

actions, or lower flutter margins. 

Note: Although failure-annunciationsystem reliability must be included in 

probability calculations for paragraph (d)(2) 

of these special conditions, there is no 

specific reliability requirement for the 

annunciation system required in paragraph 

(e) of these special conditions. 

(d) General. The following criteria 

will be used in determining the 

influence of a system and its failure 

conditions on the airplane structure: 

(1) System fully operative. With the 

system fully operative, the following 

apply: 

(i) Limit loads must be derived in all 

normal operating configurations of the 

system from all the limit conditions 

specified in subpart C of 14 CFR part 25 

(or used in lieu of those specified in 

subpart C), taking into account any 

special behavior of such a system or 

associated functions, or any effect on 

the structural performance of the 

airplane that may occur up to the limit 

loads. In particular, any significant 

degree of nonlinearity in rate of 

displacement of control surface or 

thresholds, or any other system 

nonlinearities, must be accounted for in 

a realistic or conservative way when 

deriving limit loads from limit 

conditions. 

(ii) The airplane must meet the 

strength requirements of part 25 for 

static strength and residual strength, 

using the specified factors to derive 

ultimate loads from the limit loads 

defined above. The effect of 

nonlinearities must be investigated 

beyond limit conditions to ensure the 

behavior of the system presents no 

anomaly compared to the behavior 

below limit conditions. However, 

conditions beyond limit conditions 

need not be considered if the applicant 

demonstrates that the airplane has 

design features that will not allow it to 

exceed those limit conditions. 

(iii) The airplane must meet the 

aeroelastic stability requirements of 

§ 25.629. 

(2) System in the failure condition. 

For any system failure condition not 

shown to be extremely improbable, the 

following apply: 

(i) Establishing loads at the time of 

occurrence. Starting from 1g level flight 

conditions, a realistic scenario 

including pilot corrective actions must 

be established to determine loads 

occurring at the time of failure and 

immediately after failure. 

(A) For static-strength substantiation, 

these loads, multiplied by an 

appropriate factor of safety related to 

probability of occurrence of the failure, 

are ultimate loads to be considered for 

design. The factor of safety (FS) is 

defined in Figure 1. 


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(B) For residual-strength 

substantiation, the airplane must be able 

to withstand two-thirds of the ultimate 

loads defined in paragraph (d)(2)(i)(A) 

of these special conditions. For 

pressurized cabins, these loads must be 

combined with the normal operating 

differential pressure. 

(C) Freedom from aeroelastic 

instability must be shown up to the 

speeds defined in § 25.629(b)(2). For 

failure conditions that result in speeds 

beyond design cruise speed or design 

cruise mach number (Vc/Mc), freedom 

from aeroelastic instability must be 

shown to increased speeds, so that the 

margins intended by § 25.629(b)(2) are 

maintained. 

(D) Failures of the system that result 

in forced structural vibrations 

(oscillatory failures) must not produce 

loads that could result in detrimental 

deformation of primary structure. 

(3) Establishing loads in the systemfailed state for the continuation of the 

flight. For airplane-flight continuation 

in the system-failed state, and 

considering any appropriate 

reconfiguration and flight limitations, 

the following apply: 

(i) Loads derived from the following 

conditions (or used in lieu of the 

following conditions) at speeds up to 

Vc/Mc, or the speed limitation 

prescribed for the remainder of the 

flight, must be determined: 

(A) The limit symmetricalmaneuvering conditions specified in 

§§ 25.331 and 25.345. 

(B) The limit gust-and-turbulence 

conditions specified in §§ 25.341 and 

25.345. 

(C) The limit rolling conditions 

specified in § 25.349 and the limit 

unsymmetrical conditions specified in 

§§ 25.367 and 25.427(b) and (c). 

(D) The limit yaw-maneuvering 

conditions specified in § 25.351. 

(E) The limit ground-loading 

conditions specified in §§ 25.473 and 

25.491. 

(ii) For static-strength substantiation, 

each part of the structure must be able 

to withstand the loads in paragraph 

(d)(3)(i) of these special conditions, 

multiplied by a FS depending on the 

probability of being in this failure state. 

The FS is defined in Figure 2. 

Qj = (Tj)(Pj) 

Where: 

Tj = Average time spent in failure condition 

j (in hours) 

Pj = Probability of occurrence of failure mode 

j (per hour) 

Note: If Pj is greater than 10¥3 per flight 

hour, then a 1.5 FS must be applied to all 

limit-load conditions specified in part 25, 

subpart C. 

(iii) For residual-strength 

substantiation, the airplane must be able 

to withstand two-thirds of the ultimate 

loads defined in paragraph (d)(3)(ii) of 

these special conditions. For 

pressurized cabins, these loads must be 

combined with the normal operating 

differential pressure. 

(iv) If the loads induced by the failure 

condition have a significant effect on 

fatigue or damage tolerance, then the 

effects of these loads must be taken into 

account. 

(v) Freedom from aeroelastic 

instability must be shown up to a speed 

determined from Figure 3. Flutterclearance speeds V′ and V′′ may be 

based on the speed limitation specified 

for the remainder of the flight using the 

margins defined by § 25.629(b). 

V′ = Clearance speed as defined by 

§ 25.629(b)(2) 

V′′ = Clearance speed as defined by 

§ 25.629(b)(1) 

Qj = (Tj)(Pj) 

Where: 


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70096 Federal Register / Vol. 75, No. 221 / Wednesday, November 17, 2010 / Rules and Regulations 

Tj = Average time spent in failure condition 

j (in hours) 

Pj = Probability of occurrence of failure mode 

j (per hour) 

Note: If Pj is greater than 10¥3 per flight 

hour, then the flutter-clearance speed must 

not be less than V′′. 

(vi) Freedom from aeroelastic 

instability must also be shown up to V′ 

in Figure 3 above, for any probable 

system-failure condition, combined 

with any damage, required or selected 

for investigation by § 25.571(b). 

(4) Consideration of certain failure 

conditions may be required by other 

sections of part 25 regardless of 

calculated system reliability. Where 

analysis shows the probability of these 

failure conditions to be less than 10¥9, 

criteria other than those specified in this 

paragraph may be used for structural 

substantiation to show continued safe 

flight and landing. 

(e) Failure indications. For system 

failure detection and indication, the 

following apply: 

(1) The system must be checked for 

failure conditions, not extremely 

improbable, that degrade the structural 

capability of the airplane below the 

level required by part 25 or significantly 

reduce the reliability of the remaining 

system. As far as reasonably practicable, 

the flightcrew must be made aware of 

these failures before flight. Certain 

elements of the control system, such as 

mechanical and hydraulic components, 

may use special periodic inspections, 

and electronic components may use 

daily checks, instead of detection and 

indication systems to achieve the 

objective of this requirement. Such 

certification-maintenance inspections or 

daily checks must be limited to 

components on which faults are not 

readily detectable by normal detection 

and indication systems, and where 

service history shows that inspections 

will provide an adequate level of safety. 

(2) The existence of any failure 

condition, not extremely improbable 

during flight, that could significantly 

affect the structural capability of the 

airplane and for which the associated 

reduction in airworthiness can be 

minimized by suitable flight limitations, 

must be signaled to the flightcrew. For 

example, failure conditions that result 

in a FS between the airplane strength 

and the loads of part 25, subpart C, 

below 1.25, or flutter margins below V′′, 

must be signaled to the crewmembers 

during flight. 

(f) Dispatch with known failure 

conditions. If the airplane is to be 

dispatched in a known system-failure 

condition that affects structural 

performance, or affects the reliability of 

the remaining system to maintain 

structural performance, then the 

provisions of these special conditions 

must be met, including the provisions of 

paragraph (d)(1) of these special 

conditions for the dispatched condition, 

and paragraph (d)(2) of these special 

conditions for subsequent failures. 

Expected operational limitations may be 

taken into account in establishing Pj as 

the probability of failure occurrence for 

determining the safety margin in Figure 

1. Flight limitations and expected 

operational limitations may be taken 

into account in establishing Qj as the 

combined probability of being in the 

dispatched failure condition and the 

subsequent failure condition for the 

safety margins in Figures 2 and 3. These 

limitations must be such that the 

probability of being in this combined 

failure state, and then subsequently 

encountering limit load conditions, is 

extremely improbable. No reduction in 

these safety margins is allowed if the 

subsequent system-failure rate is greater 

than 10¥3 per hour. 

Issued in Renton, Washington, on 

November 5, 2010. 

Jeffrey Duven, 

Acting Manager, Transport Airplane 

Directorate, Aircraft Certification Service. 

[FR Doc. 2010–28999 Filed 11–16–10; 8:45 am] 

BILLING CODE 4910–13–P 

DEPARTMENT OF TRANSPORTATION 

Federal Aviation Administration 

14 CFR Part 39 

[Docket No. FAA–2010–0732; Directorate 

Identifier 2010–NE–04–AD; Amendment 39– 

16509; AD 2010–23–20] 

RIN 2120–AA64 

Airworthiness Directives; General 

Electric Company (GE) CT7–9C and 

–9C3 Turboprop Engines 

AGENCY: Federal Aviation 

Administration (FAA), DOT. 

ACTION: Final rule. 

SUMMARY: We are adopting a new 

airworthiness directive (AD) for the 

products listed above. This AD requires 

inspecting certain serial number (S/N) 

gas generator turbine (GGT) shafts for 

nonconforming land balance-cuts, and if 

found, removing the shaft from service. 

This AD was prompted by reports of a 

manufacturing quality problem. We are 

issuing this AD to detect nonconforming 

GGT shaft land balance-cuts, which 

could result in the shaft failing before its 

published life limit, and which could 

result in an uncontained engine failure 

and damage to the airplane. 

DATES: This AD is effective December 

22, 2010. 

The Director of the Federal Register 

approved the incorporation by reference 

of a certain publication listed in the AD 

as of December 22, 2010. 

ADDRESSES: For service information 

identified in this AD, contact General 

Electric Company, GE–Aviation, Room 

285, 1 Neumann Way, Cincinnati, Ohio 

45215; e-mail geae.aoc@ge.com; 

telephone (513) 552–3272; fax (513) 

552–3329. You may review copies of the 

referenced service information at the 

FAA, Engine & Propeller Directorate, 12 

New England Executive Park, 

Burlington, MA. For information on the 

availability of this material at the FAA, 

call (781) 238–7125. 

Examining the AD Docket 

You may examine the AD docket on 

the Internet at http:// 

www.regulations.gov; or in person at the 

Docket Management Facility between 

9 a.m. and 5 p.m., Monday through 

Friday, except Federal holidays. The AD 

docket contains this AD, the regulatory 

evaluation, any comments received, and 

other information. The address for the 

Docket Office (phone: 800–647–5527) is 

Document Management Facility, U.S. 

Department of Transportation, Docket 

Operations, M–30, West Building 

Ground Floor, Room W12–140, 1200 

New Jersey Avenue, SE., Washington, 

DC 20590. 

FOR FURTHER INFORMATION CONTACT: 

Walter Meibaum, Aerospace Engineer, 

Engine Certification Office, FAA, Engine 

& Propeller Directorate, 12 New England 

Executive Park, Burlington, MA 01803; 

telephone (781) 238–7119; fax (781) 

238–7199; e-mail: 

walter.meibaum@faa.gov. 

SUPPLEMENTARY INFORMATION: 

Discussion 

We issued a notice of proposed 

rulemaking (NPRM) to amend 14 CFR 

part 39 to include an airworthiness 

directive (AD) that would apply to the 

specified products. That NPRM 

published in the Federal Register on 

July 23, 2010 (75 FR 43099). That NPRM 

proposed to require inspecting certain 

S/N GGT shafts, P/N 6068T44P02, for 

nonconforming land balance-cuts, and if 

found, replacing the shaft. 

Comments 

We gave the public the opportunity to 

participate in developing this AD. We 

received no comments on the NPRM or 

on the determination of the cost to the 

public. 


erowe on DSK5CLS3C1PROD with RULES