[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
