Federal Register / Vol. 62, No. 68 / Wednesday, April 9, 1997 / Proposed Rules 17117

IV. Solicitation of Comments

The OTS invites public comment on

all aspects of the proposal.

V. Executive Order 12866

The Director of the OTS has

determined that this proposed rule does

not constitute a ‘‘significant regulatory

action’’ for the purposes of Executive

Order 12866.

VI. Regulatory Flexibility Act Analysis

Pursuant to Section 605(b) of the

Regulatory Flexibility Act, the OTS

certifies that this proposal, which will

reduce regulatory burdens, will not have

a significant economic impact on a

substantial number of small entities,

because small entities utilizing the

regulation may be able to reduce the

number of applications they must file in

order to convert to a federal charter.

Accordingly, an initial Regulatory

Flexibility Analysis is not required.

VII. Unfunded Mandates Act of 1995

Section 202 of the Unfunded

Mandates Reform Act of 1995, Pub. L.

104–4 (Unfunded Mandates Act),

requires that an agency prepare a

budgetary impact statement before

promulgating a rule that includes a

federal mandate that may result in

expenditures by state, local, and tribal

governments, in the aggregate, or by the

private sector, or $100 million or more

in any one year. If a budgetary impact

statement is required, Section 205 of the

Unfunded Mandates Act also requires

an agency to identify and consider a

reasonable number of regulatory

alternatives before promulgating a rule.

The OTS has determined that the

proposed rule will not result in

expenditures by state, local, or tribal

governments or by the private sector of

$100 million or more. Accordingly, this

rulemaking is not subject to Section 202

of the Unfunded Mandates Act.

List of Subjects in 12 CFR Part 543

Conversions, Reporting and

recordkeeping requirements, Savings

associations.

Accordingly, the Office of Thrift

Supervision proposes to amend chapter

V, title 12, Code of Federal Regulations,

as set forth below.

PART 543—INCORPORATION,

ORGANIZATION, AND CONVERSION

OF FEDERAL MUTUAL

ASSOCIATIONS

1. The authority citation for part 543

continues to read as follows:

Authority: 12 U.S.C. 1462, 1462a, 1463,

1464, 1467a, 2901 et seq.

2. Section 543.8 is amended by

revising the heading and paragraph (a)

to read as follows:

§ 543.8 Conversion of depository

institutions to Federal mutual charter.

(a) With the approval of the OTS, any

depository institution, as defined in

§ 552.13 of this chapter, that is in

mutual form may convert into a Federal

mutual savings association, provided

that:

(1) The depository institution, upon

conversion, will have deposits insured

by the Federal Deposit Insurance

Corporation;

(2) The depository institution, in

accomplishing the conversion, complies

with all applicable state and federal

statutes and regulations, and OTS

policies, and obtains all necessary

member approvals; and

(3) The resulting Federal mutual

association conforms within the time

prescribed by the OTS to the

requirements of section 5(c) of the Home

Owners’ Loan Act.

* * * * *

3. Section 543.9 is amended by

revising paragraph (a) and the

introductory text of paragraph (c) to

read as follows:

§ 543.9 Application for conversion to

Federal mutual charter.

(a) Filing. Any depository institution

that proposes to convert to a Federal

mutual association as provided in

§ 543.8 shall, after approval by its board

of directors, file in accordance with

§ 516.1 of this chapter an application on

forms obtained from the OTS. The

applicant shall submit any financial

statements or other information the OTS

may require.

* * * * *

(c) Action on application. The OTS

will consider such application and any

information submitted therewith, and

may approve the application in

accordance with section 5(e) of the

Home Owners’ Loan Act and

§ 543.2(g)(1). Converting depository

institutions that have been in existence

less than three years will be subject to

all approval criteria and other

requirements applicable to de novo

Federal associations. Approval of an

application and issuance by the OTS of

a charter will be subject to:

* * * * *

Dated: April 2, 1997.

By the Office of Thrift Supervision.

Nicolas P. Retsinas,

Director.

[FR Doc. 97–8952 Filed 4–8–97; 8:45 am]

BILLING CODE 6720–01–P

DEPARTMENT OF TRANSPORTATION

Federal Aviation Administration

14 CFR Part 25

[Docket No. NM–139, Notice No. SC–97–2–

NM]

Special Conditions: Ilyushin Aviation

Complex Model Il–96T Airplane

AGENCY: Federal Aviation

Administration (FAA), DOT.

ACTION: Notice of proposed special

conditions.

SUMMARY: This notice proposes special

conditions for the Ilyushin Aviation

Complex Model Il–96T airplane. This

airplane will have novel and unusual

design features when compared to the

state of technology envisioned in the

airworthiness standards of part 25 of the

Federal Aviation Regulations (FAR).

This document contains the additional

safety standards that the Administrator

considers necessary to establish a level

of safety equivalent to that provided by

the airworthiness standards of part 25.

DATES: Comments must be received on

or before May 27, 1997.

ADDRESSES: Comments may be mailed

in duplicate to: Federal Aviation

Administration, Office of the Assistant

Chief Counsel, Attn: Rules Docket

(ANM–7), Docket No. NM–139, 1601

Lind Avenue SW., Renton, Washington,

98055–4056; or delivered in duplicate to

the Office of the Assistant Chief Counsel

at the above address. Comments must be

marked: Docket No. NM–139.

Comments may be inspected in the

Rules Docket weekdays, except Federal

holidays, between 7:30 a.m. and 4:00

p.m.

FOR FURTHER INFORMATION CONTACT:

John Hickey, FAA, International

Program Manager, ANM–102, Transport

Airplane Directorate, Aircraft

Certification Service, 1601 Lind Avenue

SW., Renton, WA 98055–4056;

telephone (206) 227–2128.

SUPPLEMENTARY INFORMATION:

Comments Invited

Interested persons are invited to

participate in the making of these

proposed special conditions by

submitting such written data, views, or

arguments as they may desire.

Communications should identify the

regulatory docket or notice number and

be submitted in duplicate to the address

specified above. All communications

received on or before the closing date

for comments will be considered by the

Administrator before further rulemaking

action on this proposal is taken. The
17118 Federal Register / Vol. 62, No. 68 / Wednesday, April 9, 1997 / Proposed Rules

proposals contained in this notice may

be changed in light of the comments

received. All comments received will be

available in the Rules Docket, both

before and after the closing date for

comments, for examination by

interested parties. A report summarizing

each substantive public contact with

FAA personnel concerning this

rulemaking will be filed in the docket.

Commenters wishing the FAA to

acknowledge receipt of their comments

submitted in response to this notice

must include a self-addressed, stamped

postcard on which the following

statement is made: ‘‘Comments to

Docket No. NM–139.’’ The postcard will

be date stamped and returned to the

commenter.

Background

Ilyushin Aviation Complex, 45

Leningradsky Prospect, Moscow,

125190, Russia, has applied for Russian

type certification of their Model Il–96T

airplane by the Aviation Register (AR) of

the Interstate Aviation Committee in

accordance with existing AR standards.

The AR is authorized to perform

airworthiness certification functions on

behalf of the Commonwealth of

Independent States, including the

Russian government. In addition,

Ilyushin applied for U.S. type

certification of the Model Il–96T on

February 16, 1993.

Section 21.29 of 14 CFR part 21 of the

Federal Aviation Regulations (FAR)

prescribes a reciprocal bilateral

agreement between the U.S. and

exporting country as a requirement for

consideration of U.S. design or

airworthiness approval of an imported

aeronautical product. Such agreements

are known as bilateral aviation safety

agreements (BASA). Although the U.S.

does not presently have a BASA with

Russia providing reciprocal acceptance

of transport category airplanes, the FAA

is working with the AR and Russian

government officials to conclude an

agreement of this nature. FAA Advisory

Circular (AC) 21–23, Airworthiness

Certification of Civil Aircraft, Engines,

Propellers, and Related Products

Imported to the United States, provides

further guidance in this regard.

A BASA with Russia may be

concluded following successful

completion of an assessment by the

FAA and the AR of each other’s

technical competence and regulatory

capability for performing airworthiness

certification functions. The scope of the

agreement is defined by each authority

in Implementation Procedures. FAA

type certification of the Model Il–96T

transport airplane is therefore

conditional upon successful

implementation of a BASA with Russia,

providing acceptance of transport

category airplanes.

One of the key elements of any BASA

assessment program is the shadow

certification program. Under the

Russian shadow certification program,

FAA specialists are ‘‘shadowing’’ their

AR counterpart specialists during AR

certification of an example of the

aeronautical product that the BASA is

intended to cover. This program is

intended to provide FAA assessment

specialists with ample opportunity to

evaluate the AR certification process

and the AR specialists’ technical

competencies to support the

airworthiness authority responsibilities

inherent in a bilateral agreement. The

Ilyushin Model Il–96T was selected as

the product for this shadow certification

which, if successful, would lead to a

U.S.-Russia BASA. Conclusion of the

BASA and related implementation

procedures would, in turn, be followed

by issuance of a U.S. type certificate for

that model.

Under the anticipated provisions of

the future BASA, the AR has elected to

certify that the Model Il–96T complies

with the AP–25 type certification

standards, plus any additional

requirements identified by the FAA to

ensure an equivalent level of safety to

that provided by the U.S. type

certification standards. The AP–25

airworthiness standards, which were

developed as the successor to the

NLGS–3 standards of the former Soviet

Union, were approved by the AR in

November 1993 and implemented in

Russia in July 1994. These standards

have also been accepted by many of the

other Commonwealth of Independent

States for type certification of transport

category airplanes. They were

established after extensive

harmonization with part 25 of the FAR

and the European Joint Airworthiness

Requirements (JAR)–25. The AP–25

standards are similar to part 25 of the

FAR; however, there are certain

specified differences in the

requirements of the two documents.

Based on the application date of

February 16, 1993, the U.S. type

certification standards are part 25 of the

FAR, as amended by Amendments 25–

1 through 25–77, and the special

conditions proposed in this notice. In

addition, the type certification basis

includes the sections of part 25, as

amended by Amendment 25–80,

pertaining to lightning protection.

Compliance with those sections is

required under the provisions of

§ 21.17(a)(1)(ii).

Because the AR has elected to certify

that the Model Il–96T complies with the

Russian type certification standards, the

FAA will make a comparison of the

Russian type certification basis and the

U.S. type certification standards

described above. Based on this

comparison, the FAA will prescribe any

additional requirements that are

necessary to ensure that the Model Il–

96T meets a level of safety equivalent to

that provided by the U.S. type

certification standards. For U.S.

clarification of the Model Il–96T, the

FAA will therefore accept the Russian

type certification basis, plus any

additional requirements, and the special

conditions proposed in this notice. As

the program progresses, other features of

the Model Il–96T may be determined to

be novel or unusual. The equivalent

certification basis may therefore include

other special conditions or exemptions

not pertinent to the special conditions

proposed in this notice.

Since noise certification and emission

requirements are beyond the scope of

the possible future bilateral agreement,

the FAA will make findings of

compliance with the applicable U.S.

noise, fuel venting, and exhaust

emission requirements. The U.S. noise

certification basis for the Model Il–96T

is 14 CFR part 36 of the FAR, as

amended by Amendments 36–1 through

36–21, and any subsequent amendments

that are applicable on the date on which

the U.S. type certificate is issued. In

addition to compliance with part 36, the

statutory provisions of Public Law 92–

574, ‘‘Noise Control Act of 1972,’’

require that the FAA issue a finding of

regulatory adequacy pursuant to Section

611 of that Act. The Model Il–96T must

also comply with the fuel venting and

exhaust emission requirements of 14

CFR part 34 of the FAR, as amended by

Amendment 34–1, and any subsequent

amendments that are applicable on the

date the type certificate is issued.

Special conditions are prescribed

under the provisions of § 21.16 of the

FAR when the applicable regulations for

type certification do not contain

adequate or appropriate standards

because of novel or unusual design

features. As discussed below, the new

Ilyushin Model Il–96T airplane

incorporates a number of such design

features.

Il–96T Design Features

General

The Model Il–96T airplane presented

for U.S. type certification is a long

range, four engine, transport category

cargo airplane powered by four (4) Pratt

& Whitney PW2337 engines with 37,500

lbs. thrust ratings and incorporating

Rockwell/Collins avionics. It is
Federal Register / Vol. 62, No. 68 / Wednesday, April 9, 1997 / Proposed Rules 17119

designed to be flown by a two-man

crew; however, it incorporates seats for

2 additional crewmembers. The airplane

is intended for cargo operation only and

is designed to carry cargo on main and

lower decks. The aircraft cargo loading

system includes a large main deck cargo

door (15.91 feet × 9.43 feet) and two

lower deck cargo doors (8.69 feet × 5.74

feet). The main cargo compartment on

the upper deck has a volume of 20,480

cubic feet and can accommodate 25 P–

6 pallets. The two cargo compartments

on the lower deck have a total volume

of 6,900 cubic feet, and can

accommodate a total of 32 LD–3

containers or 9 P–6 pallets. The Il–96T

has a maximum takeoff weight of

595,240 lbs. and a maximum landing

weight of 485,000 lbs. The maximum

cruise altitude is 43,000 feet.

The structure of the Il–96T is

generally of conventional design and

construction. The landing gear system

employs a center landing gear for use

during ground handling conditions with

heavy airplane weights. The structural

design also makes use of an electronic

flight control system which provides the

potential for a wide range of structural

and system interactions.

The Model Il–96T flight control

system is an electro-hydromechanical

system utilizing both fly-by-wire (FBW)

and conventional mechanical (cables

and push-pull rods) linkages between

pilot control column and control surface

hydraulic actuators in two

simultaneously operated and

synchronized channels. The

conventional mechanical channel, in

normal operation, functions as a passive

redundancy of the FBW channel and

provides feedback to the pilots via the

Automatic Feel Load System.

Hydraulic power to the flight control

system is simultaneously provided by

four independent hydraulic systems.

Functions are shared among these

systems in order to ensure airplane

control in the event of loss of one, two,

or three systems. The four systems are

pressurized by variable displacement

pumps driven by the engine accessory

gearbox. In addition, the systems can be

powered by electrically driven pumps.

A ram air turbine (RAT)-driven pump is

available as an emergency hydraulic

power source.

Normal electrical power is supplied

by four constant frequency generators,

one on each engine. An auxiliary power

unit (APU) providing electrical and

hydraulic supply is available for ground

use only and is not used in flight. Five

batteries provide an alternative source

of electrical power for loads required to

continue safe flight and landing in the

case of failure of four generators.

The engine control system consists of

a dual-channel electronic engine control

(EEC) mounted on the fan case of each

engine. Each EEC interfaces with

various airplane computer systems. The

EEC provides gas generator control,

engine limit protection, power

management, thrust reverser control,

and engine parameter inputs for the

flight deck displays. The engine EEC

and associated airplane related systems

form the complete propulsion control

system.

Pitch and roll control inputs are made

through conventional flight deck central

control columns. The flight instruments

are displayed on six cathode ray tube

(CRT) displays. Two CRT’s are mounted

directly in front of both the pilot and

copilot and display primary flight

instruments and navigational

information. The other two CRT’s are

located in the center of the instrument

panel and display engine parameters,

warnings, and system diagnostics.

The proposed type design of the

Model Il–96T contains novel or unusual

design features not envisioned by the

applicable part 25 airworthiness

standards and therefore special

conditions are considered necessary in

the following areas:

Airframe

1. Center Landing Gear

The Ilyushin Il–96T landing gear

arrangement includes a center braking

landing gear under the fuselage. The

center main landing gear does not differ

from that of the right or left main

landing gear in construction and

performs the same functions. The

current landing gear design criteria are

applicable to conventional landing gear

arrangements. Therefore, to provide

additional taxi, takeoff, and landing

criteria for this arrangement, Special

Condition No. 1 is proposed.

2. Design Maneuver Requirements

In a conventional airplane with a

hydro-mechanical flight control system,

pilot inputs directly affect control

surface movement (both rate and

displacement) for a given flight

condition. In the Il–96T, the pilot’s

control and the flight control surfaces

are connected through the electronic

flight control system, which introduces

additional surface movements based on

its design control laws. The control

surface movement during maneuvers

differs from the pilot control

displacements in terms of both rate and

displacement. The additional effects of

the electronic flight control system are

not reflected in the current FAR.

Therefore, Special Condition No. 2 is

proposed.

3. Interaction of Systems and Structure

The Ilyushin Model Il–96T is

equipped with an electrical flight

control system and a load alleviation

system that effects both gust and

maneuver loads. These systems can

directly, or as a result of failure or

malfunction, affect structural

performance. This degree of system and

structures interaction was not

envisioned in the structural design

regulations of part 25 of the FAR for

transport airplanes. To provide

comprehensive criteria in which the

structural design safety margins are

dependent on systems reliability,

Special Condition No. 3 is proposed.

Systems

4. Protection from Unwanted Effects of

High Intensity Radiated Fields (HIRF)

The use of fly-by-wire designs to

command and control engines and flight

control surfaces increases the airplane’s

susceptibility to HIRF sources external

to the airplane. The airworthiness

regulations do not provide adequate

requirements for protection from

unwanted effects of HIRF.

High intensity radiated fields have the

potential to cause adverse and

potentially hazardous effects on fly-bywire systems if design measures are not

taken to ensure the immunity of such

systems. This is particularly true with

the trend toward increased power levels

from ground based transmitters and the

advent of space and satellite

communications.

The Model Il–96T is being designed

with electrical interfaces between crew

inputs and (1) the flight control

surfaces, and (2) the engines. The

interfaces, and the interconnection

among the electronic subsystems

controlling these functions, can be

susceptible to disruption of both

command/response signals and the

operational mode logic as a result of

electrical and magnetic interference.

Traditional airplane designs have

utilized mechanical means to connect

the primary flight controls and the

engine to the flight deck. This

traditional design results in control

paths that are substantially immune to

the effects of HIRF. A special condition

is required to ensure that critical and

essential systems be designed and

installed to preclude component

damage and system upset or

malfunction due to the unwanted effects

of HIRF. Therefore, Special Condition

No. 4 is proposed.
17120 Federal Register / Vol. 62, No. 68 / Wednesday, April 9, 1997 / Proposed Rules

Special conditions may be issued and

amended, as necessary, as part of the

type certification basis if the

Administrator finds that the

airworthiness standards designated in

accordance with § 21.17(a)(1) do not

contain adequate or appropriate safety

standards because of novel or unusual

design features of an airplane. Special

conditions, as appropriate, are issued in

accordance with § 11.49 after public

notice, as required by §§ 11.28 and

11.29(b), effective October 14, 1980, and

become part of the type certification

basis in accordance with § 21.17(a)(2).

These special conditions would be

applicable initially to the Ilyushin

Model Il–96T airplane. Should Ilyushin

Aviation Complex apply at a later date

for a change to the type certificate to

include another model incorporating the

same novel or unusual design features,

the special conditions would apply to

that model as well under the provisions

of § 21.101(a)(1).

Conclusion

This action affects only certain

unusual or novel design features on one

model series of airplanes. It is not a rule

of general applicability and affects only

the manufacturer who applied to the

FAA for approval of these features on

the airplane.

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 Proposed Special Conditions

Accordingly, the Federal Aviation

Administration (FAA) proposes the

following special conditions as part of

the type certification basis for the

Ilyushin Aviation Complex Model Il–

96T series airplanes.

1. Center Landing Gear

Notwithstanding § 25.477 of the FAR,

the requirements of §§ 25.473 and

25.479 through § 25.485 apply, except as

noted:

(a) In addition to the requirements of

§ 25.473, landing should be considered

on a level runway and on a runway

having a convex upward shape that may

be approximated by a slope of 1.5

percent with the horizontal at main

landing gear stations. The maximum

loads determined from these two

conditions must be applied to each

main landing gear and to the center

landing gear.

(b) The requirements of § 25.483

apply and, in addition, the condition

represented by the following figure also

applies:

BILLING CODE 4910–13–M

BILLING CODE 4910–13–C

(c) In lieu of the requirements of

§ 25.485, the following apply:

(1) The airplane is considered to be in

the level attitude with only the main

and central wheels contacting the

ground.

(2) Vertical reactions of one-half of the

maximum vertical reaction obtained at

each main and center gear in the level

landing conditions should be

considered. The vertical loads must be

combined with side loads that for the

main gear are 0.8 of the vertical reaction

(on one side) acting inward and 0.6 of

the vertical reaction (on the other side)

acting outward, and for the center gear

are 0.7 of the vertical reaction acting in

the same direction as main gear side

loads. (Drag load=0)

(d) In addition to the requirements of

§ 25.489, ‘‘Ground handling

conditions,’’ the following applies: The

airplane should be considered to be on

a level runway and on a runway having

a convex upward shape that may be

approximated by a slope of 1.5 percent

with the horizontal at main landing gear

stations. The ground reactions must be

distributed to the individual landing

gear units in a rational or conservative

manner (zero life, shock struts in the

static position).

(e) In lieu of the requirements of

§ 15.503, the following apply:

(1) The airplane is assumed to pivot

about one of the outer main gears with

brakes locked on the selected gear. The

limit vertical load factor must be 1.0 and

the coefficient of friction must be 0.8.

(2) The airplane is assumed to be in

static equilibrium, with the loads being

applied at the ground contact points.

(3) All of the main gear units must be

designed for the scrubbing or torsion

loads, or both, induced by pivoting

during ground maneuvers produced by:

(i) Towing at the nose gear, no brakes

applied; and

(ii) Application of symmetrical or

unsymmetrical forward thrust to aid

pivoting and with or without breaking

on the outside main gear closest to the

pivot center.
Federal Register / Vol. 62, No. 68 / Wednesday, April 9, 1997 / Proposed Rules 17121

(f) The following applies to the center

landing gear in lieu of § 25.723, ‘‘Shock,

absorption tests’’:

(1) The center landing gear should not

fail in a test demonstrating its reserve

energy absorption capacity at design

landing weight, assuming airplane lift

no greater than the airplane weight

acting during an impact simulating:

(i) A center landing gear descent

velocity of 120 percent of the maximum

aircraft descent velocity at the time of

center landing gear ground contact; or

(ii) A 12 fps airplane landing impact

taking into account both the main and

center landing gears acting during the

impact, whichever is more critical.

2. Design Maneuver Requirements

(a) Maximum elevator displacement

at VA. In lieu of compliance with

§ 25.331(c)(1) of the FAR, the airplane is

assumed to be flying in steady level

flight (point A1 within the maneuvering

envelope of § 25.333(b)) and, except as

limited by pilot efforts specified in

§ 25.397 concerning pilot effort forces,

the cockpit pitching control device is

suddenly moved to obtain extreme

positive pitching acceleration (nose up).

In defining the tail load condition, the

response of the airplane must be taken

into account. Airplane loads which

occur subsequent to the point at which

the normal acceleration at the center of

gravity exceeds the maximum positive

limit maneuvering factor, n, need not be

considered.

(b) Pitch maneuvering loads induced

by the system. In addition to the

requirements of § 25.331(c) of the FAR,

it must be established that pitch

maneuver loads induced by the system

itself (e.g. abrupt changes in orders

made possible by electrical rather than

mechanical combination of different

inputs) are acceptably accounted for.

(c) Roll maneuver loads. In lieu of

compliance with § 25.349(a) of the FAR,

the following conditions, speeds, spoiler

and aileron deflections (except as the

deflections may be limited by pilot

effort) must be considered in

combination with an airplane load

factor of zero and of two-thirds of the

positive maneuvering factor used in

design. In determining the required

aileron and spoiler deflections, the

torsional flexibility of the wing must be

considered in accordance with

§ 25.301(b).

(1) Conditions corresponding to

steady rolling velocities must be

investigated. In addition, conditions

corresponding to maximum angular

acceleration must be investigated. For

the angular acceleration conditions, zero

rolling velocity may be assumed in the

absence of a rational time history

investigation of the maneuver.

(2) At VA, sudden deflection of the

cockpit roll control up to the limit is

assumed. The position of the cockpit

roll control must be maintained until a

steady roll rate is achieved and then

must be returned suddenly to the

neutral position.

(3) At VC, the cockpit roll control

must be moved suddenly and

maintained so as to achieve a rate of roll

not less than that obtained in paragraph

(2).

(4) At VD, the cockpit roll control

must be moved suddenly and

maintained so as to achieve a rate of roll

not less than one third of that obtained

in paragraph (2) of this paragraph.

(5) It must also be established that roll

maneuver loads induced by the system

itself (i.e., abrupt changes in orders

made possible rather than mechanical

combination of different inputs) are

acceptably accounted for.

(d) Yaw maneuver loads. In lieu of

compliance with § 25.351 of the FAR,

the airplane must be designed for loads

resulting from the conditions specified

in subparagraphs (a) and (b) of this

paragraph. Unbalanced aerodynamic

moments about the center of gravity

must be reacted in a rational or

conservative manner, considering the

principal masses furnishing the reacting

inertia forces. Physical limitations of the

airplane from the cockpit yaw control

device to the control surface deflection,

such as control stop position, maximum

power and displacement rate of the

servo controls, and control yaw limiters

may be taken into account.

(1) Maneuvering. At speeds from VMC

to VD, the following maneuvers must be

considered. In computing the tail loads,

the yawing velocity may be assumed to

be zero:

(i) With the airplane in unaccelerated

flight at zero yaw, it is assumed that the

cockpit yaw control device (pedal) is

suddenly displaced (with critical rate)

to the maximum deflection, as limited

by the stops.

(ii) With the cockpit yaw control

device (pedal) deflected as specified in

subparagraph (1) of this paragraph, it is

assumed that the airplane yaws to the

resulting sideslip angle (beyond the

static sideslip angle).

(iii) With the airplane yawed to the

static sideslip angle with the cockpit

yaw control device deflected as

specified in sub-paragraph (1) of this

paragraph, it is assumed that the cockpit

yaw control device is returned to

neutral.

3. Interaction of Systems and Structure

(a) General. For an airplane equipped

with flight control systems, load

alleviation systems, or flutter control

systems that directly, or as a result of a

failure or malfunction, affect its

structural performance, the influence of

these systems and their failure

conditions shall be taken into account

in showing compliance with subparts C

and D of part 25 of the FAR.

(b) System fully operative. With the

system fully operative, the following

apply:

(1) Limit loads must be derived in all

normal operating configurations of the

systems from all the deterministic limit

conditions specified in subpart C, taking

into account any special behavior of

such systems 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

nonlinearity (rate of displacement of

control surface, thresholds, or any other

system nonlinearities) must be

accounted for in a realistic or

conservative way when deriving limit

loads from limit conditions.

(2) The airplane must meet the

strength requirements of part 25 (static

strength, 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 systems

presents no anomaly compared to the

behavior below limit conditions.

However, conditions beyond limit

conditions need not be considered when

it can be shown that the airplane has

design features that make it impossible

to exceed those limit conditions.

(3) The airplane must meet the

aeroelastic stability requirements of

§ 25.629.

(c) System in the failure condition.

For any system failure condition not

shown to be extremely improbable, the

following apply:

(1) At the time of occurrence. Starting

from 1 g level flight conditions, a

realistic scenario, including pilot

corrective actions, must be established

to determine the loads occurring at the

time of failure and immediately after

failure. The airplane must be able to

withstand these loads, multiplied by an

appropriate factor of safety, related to

the probability of occurrence of the

failure. These loads should be

considered as ultimate loads for this

evaluation. The factor of safety is

defined as follows:

BILLING CODE 4910–13–M
17122 Federal Register / Vol. 62, No. 68 / Wednesday, April 9, 1997 / Proposed Rules

(i) The loads must also be used in the

damage tolerance evaluation required in

§ 25.571(b), if the failure condition is

probable. The loads may be considered

as ultimate loads for the damage tolerant

evaluation.

(ii) Freedom from flutter and

divergence must be shown at speeds up

to VD or 1.15 VC, whichever is greater.

However, at altitudes where the speed is

limited by Mach number, compliance

need be shown only up to MD, as

defined in § 25.335(d). For failure

conditions that result in speed increases

beyond VC/MC, freedom from flutter and

divergence must be shown at increased

speeds, so that the above margins are

maintained.

(iii) Notwithstanding subparagraph (1)

of this paragraph, failures of the system

that result in forced structural vibrations

(oscillatory failures) must not produce

peak loads that could result in

permanent deformation of primary

structure.

(2) For the continuation of the flight.

For the airplane, in the failed

configuration and considering any

appropriate flight limitations, the

following apply:

(i) Static and residual strength must

be determined for loads induced by the

failure condition, if the loads could

continue to the end of the flight. These

loads must be combined with the

deterministic limit load conditions

specified in subpart C.

(ii) For static strength substantiation,

each part of the structure must be able

to withstand the loads specified in

subparagraph (2)(i) of this paragraph

multiplied by a safety factor depending

on the probability of being in this failure

state. The factor of safety is defined as

follows:

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 factor of safety must be used.

(iii) For residual strength

substantiation as defined in § 25.571(b),

for structures also affected by failure of

the system and with damage in

combination with the system failure, a

reduction factor may be applied to the

residual strength loads of § 25.571(b).

However, the residual strength level

must not be less than the 1g flight load,

combined with the loads introduced by

the failure condition plus two-thirds of

the load increments of the conditions

specified in § 25.571(b) in both positive

and negative directions (if appropriate).

The reduction factor is defined as

follows:
Federal Register / Vol. 62, No. 68 / Wednesday, April 9, 1997 / Proposed Rules 17123

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 residual strength factor of 1.0

must be used.

(iv) Freedom from flutter and

divergence must be shown up to a speed

determined by the following figure:

BILLING CODE 4910–13–C

V1=Clearance speed as defined in

§ 25.629(b)(2).

V2=Clearance speed as defined in

§ 25.629(b)(1).

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 the flutter clearance speed must

not be less than V2.

(v) Freedom from flutter and

divergence must also be shown up to V1

in the above figure for any probable

system failure condition combined with

any damage required or selected for

investigation § 25.571(b).

(vi) If the time likely to be spent in the

failure condition is not small compared

to the damage propagation period, or if

the loads induced by the failure

condition may have a significant

influence on the damage propagation,

then the effects of the particular failure

condition must be addressed and the

corresponding inspection intervals

adjusted to adequately cover this

situation.

(vii) If the mission analysis method is

used to account for continuous

turbulence, all the systems failure

conditions associated with their

probability must be accounted for in a

rational or conservative manner in order

to ensure that the probability of

exceeding the limit load is not higher

than the prescribed value of the current

requirement.

(d) Warning considerations. For

system failure detection and warning,

the following apply:

(1) Before flight, the system must be

checked for failure conditions, not

shown to be extremely improbable, that

degrade the structural capability of the

airplane below the level intended in

paragraph (b) of this special condition.

The crew must be made aware of these

failures, if they exist, before flight.

(2) An evaluation must be made of the

necessity to signal, during the flight, the

existence of any failure condition 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. The

assessment of the need for such signals

must be carried out in a manner

consistent with the approved general

warning philosophy for the airplane.

(3) During flight, any failure condition

not shown to be extremely improbable,

in which the safety factor existing

between the airplane strength capability

and loads induced by the deterministic

limit conditions of subpart C of part 25

is reduced to 1.3 or less, must be

signaled to the crew if appropriate

procedures and limitations can be

provided so that the crew can take

action to minimize the associated

reduction in airworthiness during the

remainder of the flight.

(e) Dispatch with failure conditions. If

the airplane is to be knowingly

dispatched in a system failure condition

that reduces the structural performance

of the airplane, then operational

limitations must be provided whose

effects, combined with those of the

failure condition, allow the airplane to
17124 Federal Register / Vol. 62, No. 68 / Wednesday, April 9, 1997 / Proposed Rules

meet the structural requirements

described in paragraph (b) of this

special condition. Subsequent system

failures must also be considered.

Discussion: This special condition is

intended to be applicable to flight controls,

load alleviation systems, and flutter control

systems. The criteria provided by the special

condition only address the direct structural

consequences of the systems responses and

performances and therefore cannot be

considered in isolation but should be

included in the overall safety evaluation of

the airplane. The presentation of these

criteria may, in some instances, duplicate

standards already established for this

evaluation. The criteria are applicable to

structure, the failure of which could prevent

continued safe flight and landing.

The following definitions are applicable to

this special condition:

Structural performance: Capability of the

airplane to meet the structural requirements

of part 25.

Flight limitations: Limitations that can be

applied to the airplane flight conditions

following an inflight occurrence and which

are included in the flight manual (e.g., speed

limitations, avoidance of severe weather

conditions, etc.).

Operational limitations: Limitations,

including flight limitations, that can be

applied to the airplane operating conditions

before dispatch (e.g., payload limitations).

Probabilistic terms: The probabilistic terms

(probable, improbable, extremely

improbable) used in this special condition

should be understood as defined in AC

25.1309–1.

Failure condition: The term failure

condition is defined in AC 25.1309–1;

however, this special condition applies only

to system failure conditions that have a direct

impact on the structural performance of the

airplane (e.g., failure conditions that induce

loads or change the response of the airplane

to inputs such as gusts or pilot actions).

4. Protection From Unwanted Effects of

High Intensity Radiated Fields (HIRF)

In the absence of specific

requirements for protection from the

unwanted effects of HIRF, the following

apply:

Each airplane system that performs

critical functions must be designed and

installed to ensure that the operation

and operational capabilities of these

systems to perform critical functions are

not adversely affected when the airplane

is exposed to high intensity radiated

fields.

Discussion: The Ilyushin Model II–96T will

utilize electrical and electronic systems that

perform critical functions. These systems

include the electronic displays, intergrated

avionics computer, electronic engine

controls, etc. The existing airworthiness

regulations do not contain adequate or

appropriate safety standards for the

protection of these systems from the effects

of HIRF which are external to the airplane.

Airplane designs that utilize metal skins

and mechanical command and control means

have traditionally been shown to be immune

from the effects of HIRF energy from groundbased and airborne transmitters. With the

trend toward increased power levels from

these sources, plus the advent of space and

satellite communications, the immunity of

the airplane to HIRF energy must be

established. No universally accepted

guidance to define the maximum energy level

in which civilian airplane system

installations must be capable of operating

safely has been established.

For the purposes of this special condition,

the following definition applies:

Critical Functions: Functions whose failure

would contribute to or cause a failure

condition that would prevent the continued

safe flight and landing of the airplane. At this

time the FAA and other airworthiness

authorities are unable to precisely define or

control the HIRF energy level to which the

airplane will be exposed in service.

Therefore, the FAA hereby defines two

acceptable interim methods for complying

with the requirement for protection of

systems that perform critical functions.

(1) The applicant may demonstrate

that the critical systems, as installed in

the airplane, are protected from the

external HIRF threat environment

defined in the following table:

Frequency Field peak

(V/M)

Strength average

(V/M)

10 KHz–500

KHz ................ 60 60

500 KHz–2 MHz 80 80

2 MHz–30 MHz 200 200

30 MHz–100

MHz ............... 33 33

100 MHz–200

MHz ............... 150 33

200 MHz–400

MHz ............... 56 33

400 MHz–1 GHz 4,020 935

1 GHz–2 GHz ... 7,850 1,750

2 GHz–4 GHz ... 6,000 1,150

4 GHz–6 GHz ... 6,800 310

6 GHz–8 GHz ... 3,600 666

8 GHz–12GHz ... 5,100 1,270

12 GHz–18 GHz 3,500 551

18 GHz–40 GHz 2,400 750

or,

(2) The applicant may demonstrate by

laboratory test that the critical systems

elements and their associated wiring

harnesses can withstand a peak

electromagnetic field strength of 100

volts per meter, without the benefit of

airplane structural shielding, in the

frequency range of 10 KHz to 18 GHz.

Compliance Method:

This paragraph describes an

acceptable method of showing

compliance with the HIRF energy

protection requirements.

(1) Compliance Plan: The applicant

should present a plan for Aviation

Register approval, outlining how

compliance with the HIRF energy

protection requirements will be

attained. This plan should also propose

pass/fail criteria for the operation of

critical systems in the HIRF

environment.

(2) System Criticality: A hazard

analysis should be performed by the

applicant for approval by Aviation

Register to identify electrical and/or

electronic systems which perform

critical functions. These systems are

candidates for the application of HIRF

energy protection requirements.

(3) Compliance Verification:

Compliance with the HIRF energy

protection requirements may be

demonstrated by tests, analysis, models,

similarity with existing systems, or a

combination thereof as acceptable to

Aviation Register. Service experience

alone is not acceptable since such

experience in normal flight operations

may not include an exposure to the

HIRF environmental condition.

(4) Pass/Fail Criteria: Acceptable

system performance is attained by

demonstrating that the system under

consideration continues to perform its

intended function during and after

exposure to the required

electromagnetic fields. Deviations from

system specification may be acceptable

depending on an independent

assessment of the deviations for each

application.

(5) Test Methods and Procedures:

RTCA document DO–160C, Section 20,

provides information on acceptable test

procedures. In addition, the following

information on modulation is presented

to supplement that found in DO–160C.

Equipment and subsystem radiated

susceptibility qualification tests should

be conducted by slowly scanning the

entire frequency spectrum with an

unmodulated signal which produces the

required average electric field strength

at the equipment under test (EUT) and

its wiring. A peak level detector should

be used to monitor the peak values of

the signal and these values should be

recorded at each test point. The EUT

should not be damaged by this test and

should operate normally for frequencies

under 400 MHz. Deviations from normal

operation for test frequencies above 400

MHz should be recorded. The test

should be repeated with an appropriate

modulation applied to the test signal. At

each test point, the amplitude of the RF

test signal should be adjusted to the

peak values recorded during the

unmodulated test. The modulation

should be selected as the signal most

likely to disrupt operation of the

equipment under test based on its

design characteristics. For example,
Federal Register / Vol. 62, No. 68 / Wednesday, April 9, 1997 / Proposed Rules 17125

flight control systems might be

susceptible to 3 Hz. square wave

modulation while the video signals for

CRT displays may be susceptible to 400

Hz sinusoidal modulation. If the worst

case modulation is unknown or cannot

be determined, default modulations can

be used. Suggested default values are 1

KHz sine wave with 80% depth of

modulation in the frequency range from

10 KHz to 400 MHz and 1 KHz square

wave with greater than 90% depth of

modulation from 400 MHz to 18 GHz.

For frequencies where the unmodulated

signal caused deviations from normal

operation of the EUT, several different

modulating signals with various waveforms and frequencies should be

applied. Modern laboratory equipment

may not be able to continuously scan

the spectrum in the manner of analog

equipment. These units will only

generate discrete frequencies. For such

equipment, the number of test points

and the dwell time at each test point

must be specified. For each decade of

the frequency test spectrum (a ten times

increase in frequency i.e. 10 Kz to 100

KHz) there should be at least 25 test

points, and for the decades from 10

MHz to 100 MHz, and 100 MHz to 1

GHz there should be a minimum of 180

test points each. The dwell time at each

test point should be at least 0.5 second.

(6) Data Submittal: An

accomplishment report should be

submitted to the Aviation Register

showing fulfillment of the HIRF energy

protection requirements. This report

should contain test results, analysis and

other pertinent data.

(7) Maintenance Requirements: The

applicant (manufacturer) must provide

maintenance requirements to assure the

continued airworthiness of the installed

system(s).

Issued in Renton, Washington, on March

18, 1997.

Ronald T. Wojnar,

Manager, Transport Airplane Directorate,

Aircraft Certification Service, ANM–100.

[FR Doc. 97–9143 Filed 4–8–97; 8:45 am]

BILLING CODE 4910–13–M

DEPARTMENT OF TRANSPORTATION

Federal Aviation Administration

14 CFR Part 39

[Docket No. 97–CE–15–AD]

RIN 2120–AA64

Airworthiness Directives; SOCATAGroupe AEROSPATIALE Model TBM

700 Airplanes

AGENCY: Federal Aviation

Administration, DOT.

ACTION: Notice of proposed rulemaking

(NPRM).

SUMMARY: This document proposes to

adopt a new airworthiness directive

(AD) that would apply to certain

SOCATA-Groupe AEROSPATIALE

(Socata) Model TBM 700 airplanes that

do not have MOD 70–065–32

incorporated. This proposed AD would

require removing the MLG inboard

doors and the door locking control

mechanism (MOD 70–065–32). This AD

is the result of an incident on one of the

affected airplanes where the MLG

inboard door locking hooks (hinges)

corroded, caused the doors to jam, and

prevented the MLG from extending.

Analysis has shown that removing the

MLG inboard doors will not cause any

airplane safety or performance

problems. The actions specified by the

proposed AD are intended to prevent

failure of the MLG to extend because of

corroded MLG inboard locking hinges,

which could result in loss of control of

the airplane during landing operations.

DATES: Comments must be received on

or before June 6, 1997.

ADDRESSES: Submit comments in

triplicate to the Federal Aviation

Administration (FAA), Central Region,

Office of the Assistant Chief Counsel,

Attention: Rules Docket No. 97–CE–15–

AD, Room 1558, 601 E. 12th Street,

Kansas City, Missouri 64106. Comments

may be inspected at this location

between 8 a.m. and 4 p.m., Monday

through Friday, holidays excepted.

Service information that applies to the

proposed AD may be obtained from the

SOCATA-Groupe AEROSPATIALE,

Socata Product Support, Aeroport

Tarbes-Ossun-Lourdes, B P 930, 65009

Tarbes Cedex, France; telephone

62.41.74.26; facsimile 62.41.74.32; or

the Product Support Manager, U.S.

AEROSPATIALE, 2701 Forum Drive,

Grand Prairie, Texas 75053; telephone

(214) 641–3614; facsimile (214) 641–

3527. This information also may be

examined at the Rules Docket at the

address below.

FOR FURTHER INFORMATION CONTACT: Mr.

William J. Timberlake, Program Officer,

Brussels Aircraft Certification Division,

FAA, Europe, Africa, and Middle East

Office, c/o American Embassy, B–1000

Brussels, Belgium; telephone (32 2)

513.38.30; facsimile (32 2) 230.68.99; or

Mr. Karl Schletzbaum, Aerospace

Engineer, FAA, Small Airplane

Directorate, 1201 Walnut Street, suite

900, Kansas City, Missouri 64106;

telephone (816) 426–6934; facsimile

(816) 426–2169.

SUPPLEMENTARY INFORMATION:

Comments Invited

Interested persons are invited to

participate in the making of the

proposed 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 above. All

communications received on or before

the closing date for comments, specified

above, will be considered before taking

action on the proposed rule. The

proposals contained in this notice may

be changed in light of the comments

received.

Comments are specifically invited on

the overall regulatory, economic,

environmental, and energy aspects of

the proposed 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

proposal will be filed in the Rules

Docket.

Commenters wishing the FAA to

acknowledge receipt of their comments

submitted in response to this notice

must submit a self-addressed, stamped

postcard on which the following

statement is made: ‘‘Comments to

Docket No. 97–CE–15–AD.’’ The

postcard will be date stamped and

returned to the commenter.

Availability of NPRMs

Any person may obtain a copy of this

NPRM by submitting a request to the

FAA, Central Region, Office of the

Assistant Chief Counsel, Attention:

Rules Docket No. 97–CE–15–AD, Room

1558, 601 E. 12th Street, Kansas City,

Missouri 64106.

Discussion

The Direction Ge ́ne ́rale de l’Aviation

Civile (DGAC), which is the

airworthiness authority for France,

recently notified the FAA that an unsafe

condition may exist on certain Socata

Model TBM 700 airplanes. The DGAC