[Federal Register Volume 63, Number 13 (Wednesday, January 21, 1998)]
[Rules and Regulations]
[Pages 3023-3030]
From the Federal Register Online via the Government Publishing Office [www.gpo.gov]
[FR Doc No: 98-865]


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DEPARTMENT OF TRANSPORTATION

Federal Aviation Administration

14 CFR Part 25

[Docket No. NM-139, Special Conditions No. 25-ANM-135]


Special Conditions: Ilyushin Aviation Complex Model Il-96T 
Airplane

AGENCY: Federal Aviation Administration (FAA), DOT.

ACTION: Final special conditions.

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SUMMARY: These special conditions are issued 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). These special conditions contain 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.

EFFECTIVE DATE: February 20, 1998.

FOR FURTHER INFORMATION CONTACT: Norm Martenson, FAA, International 
Office, ANM-116, Transport Airplane Directorate, Aircraft Certification 
Service, 1601 Lind Avenue SW., Renton, WA 98055-4056; telephone (425) 
227-2196.

SUPPLEMENTARY INFORMATION:

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

[[Page 3024]]

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.-Russian 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 these special conditions. 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 
Sec. 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. 
certification of the Model Il-96T, the FAA will therefore accept the 
Russian type certification basis, plus any additional requirements, and 
these special conditions. 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 these special conditions.
    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 
Sec. 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 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  x  
9.43 feet) and two lower deck cargo doors (8.69 feet  x  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

[[Page 3025]]

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 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. Special Condition No. 1 provides additional taxi, 
takeoff, and landing criteria for this arrangement.
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 controls 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 provided.
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. Special 
Condition 3 provides comprehensive criteria in which the structural 
design safety margins are dependent on systems reliability.

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-by-wire 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. These 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 provided.
    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 Sec. 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 
Sec. 11.49 after public notice, as required by Secs. 11.28 and 
11.29(b), effective October 14, 1980, and become part of the type 
certification basis in accordance with Sec. 21.17(a)(2).

Discussion of Comments

    Notice of proposed special conditions No. SC-97-2-NM was published 
in the Federal Register on April 9, 1997 (62 FR 17117). No comments 
were received, and the special conditions are adopted as proposed.

Applicability

    These special conditions are 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 Sec. 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 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 the Ilyushin Aviation Complex Model 
Il-96T series airplanes.
    1. Center Landing Gear. Notwithstanding Sec. 25.477 of the FAR, the 
requirements of Secs. 25.473 and 25.479 through Sec. 25.485 apply, 
except as noted:
    (a) In addition to the requirements of Sec. 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 Sec. 25.483 apply and, in addition, the 
condition represented by the following figure also applies:

BILLING CODE 4910-13-P

[[Page 3026]]

[GRAPHIC] [TIFF OMITTED] TR21JA98.010



BILLING CODE 4910-13-C

    (c) In lieu of the requirements of Sec. 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 Sec. 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 lift, shock struts in the static position).
    (e) In lieu of the requirements of Sec. 25.503, the following 
apply:
    (1) The airplane is assumed to pivot about one of the outer main 
gears with the 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 braking on the outside main gear 
closest to the pivot center.
    (f) The following applies to the center landing gear in lieu of 
Sec. 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 Sec. 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 Sec. 25.333(b)) and, except as 
limited by pilot effort as specified in Sec. 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 Sec. 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 Sec. 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 
Sec. 25.301(b).
    (1) Conditions corresponding to steady rolling velocities must be 
investigated. In addition, conditions corresponding to maximum angular

[[Page 3027]]

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 Sec. 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 law 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 Sec. 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 1g 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-P
[GRAPHIC] [TIFF OMITTED] TR21JA98.011



[[Page 3028]]


BILLING CODE 4910-13-C

    (i) The loads must also be used in the damage tolerance evaluation 
required in Sec. 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 Sec. 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:

BILLING CODE 4910-13-P
[GRAPHIC] [TIFF OMITTED] TR21JA98.012


BILLING CODE 4910-13-C

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 
Sec. 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 Sec. 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 Sec. 25.571(b) in both positive and negative directions 
(if appropriate). The reduction factor is defined as follows:

BILLING CODE 4910-13-P
[GRAPHIC] [TIFF OMITTED] TR21JA98.013


BILLING CODE 4910-13-C

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 4910-13-P

[[Page 3029]]

[GRAPHIC] [TIFF OMITTED] TR21JA98.014



BILLING CODE 4910-13-C

V1=Clearance speed as defined in Sec. 25.629(b)(2).
V2=Clearance speed as defined in Sec. 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 in Sec. 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 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 Il-96T will utilize electrical and 
electronic systems that perform critical functions. These systems 
include the electronic

[[Page 3030]]

displays, integrated 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 ground-based 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:

------------------------------------------------------------------------
                                                      Field     Strength
                     Frequency                      peak  (V/  average V/
                                                        M)         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-12 GHz......................................      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 as 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, 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 wave-forms 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 December 16, 1997.
Gilbert L. Thompson,
Assistant Manager, Transport Airplane Directorate, Aircraft 
Certification Service, ANM-101.
[FR Doc. 98-865 Filed 1-20-98; 8:45 am]
BILLING CODE 4910-13-P