[Federal Register: February 3, 2003 (Volume 68, Number 22)]
[Proposed Rules]
[Page 5241-5246]
From the Federal Register Online via GPO Access [wais.access.gpo.gov]
[DOCID:fr03fe03-11]
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DEPARTMENT OF TRANSPORTATION
Federal Aviation Administration
14 CFR Part 25
[Docket No. NM242; Notice No. 25-03-01-SC]
Special Conditions: Embraer Model 170-100 and 170-200 Airplanes;
Sudden Engine Stoppage; Operation Without Normal Electrical Power;
Interaction of Systems and Structures
AGENCY: Federal Aviation Administration (FAA), DOT.
ACTION: Notice of proposed special conditions.
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SUMMARY: This notice proposes special conditions for the Embraer Model
170-100 and 170-200 airplanes. These airplanes will have novel or
unusual design features when compared to the state of technology
envisioned in the airworthiness standards for transport category
airplanes. These design features are associated with (1) engine size
and torque load which affect sudden engine stoppage, (2) electrical and
electronic flight control systems which perform critical functions, and
[[Page 5242]]
(3) systems which affect the structural performance of the airplane.
The applicable airworthiness regulations do not contain adequate or
appropriate safety standards for these design features. These proposed
special conditions contain the additional safety standards that the
Administrator considers necessary to establish a level of safety
equivalent to that established by the existing airworthiness standards.
Additional special conditions will be issued for other novel or unusual
design features of the Embraer Model 170-100 and 170-200 airplanes.
DATES: Comments must be received on or before March 20, 2003.
ADDRESSES: Comments on this proposal may be mailed in duplicate to:
Federal Aviation Administration, Transport Airplane Directorate,
Attention: Rules Docket (ANM-113), Docket No. NM242, 1601 Lind Avenue
SW., Renton, Washington 98055-4056; or delivered in duplicate to the
Transport Airplane Directorate at the above address. All comments must
be marked: Docket No. NM242. Comments may be inspected in the Rules
Docket weekdays, except Federal holidays, between 7:30 a.m. and 4 p.m.
FOR FURTHER INFORMATION CONTACT: Tom Groves, FAA, International Branch,
ANM-116, Transport Airplane Directorate, Aircraft Certification
Service, 1601 Lind Avenue SW., Renton, Washington 98055-4056; telephone
(425) 227-1503; facsimile (425) 227-1149.
SUPPLEMENTARY INFORMATION:
Comments Invited
The FAA invites interested persons to participate in this
rulemaking by submitting written comments, data, or views. The most
helpful comments reference a specific portion of the special
conditions, explain the reason for any recommended change, and include
supporting data. We ask that you send us two copies of written
comments.
We will file in the docket all comments we receive, as well as a
report summarizing each substantive public contact with FAA personnel
concerning these proposed special conditions. The docket is available
for public inspection before and after the comment closing date. If you
wish to review the docket in person, go to the address in the ADDRESSES
section of this notice between 7:30 a.m. and 4 p.m., Monday through
Friday, except Federal holidays.
We will consider all comments we receive on or before the closing
date for comments. We will consider comments filed late if it is
possible to do so without incurring expense or delay. We may change the
proposed special conditions in light of the comments we receive.
If you want the FAA to acknowledge receipt of your comments on this
proposal, include with your comments a pre-addressed, stamped postcard
on which the docket number appears. We will stamp the date on the
postcard and mail it back to you.
Background
On May 20, 1999, Embraer applied for a type certificate for its new
Model 170 airplane. Two basic versions of the Model 170 are included in
the application. The Model 170-100 airplane is a 69-78 passenger twin-
engine regional jet with a maximum takeoff weight of 81,240 pounds. The
Model 170-200 is a lengthened fuselage derivative of the 170-100.
Passenger capacity for the Model 170-200 is increased to 86, and
maximum takeoff weight is increased to 85,960 pounds.
Type Certification Basis
Under the provisions of 14 CFR 21.17, Embraer must show that the
Model 170-100 and 170-200 airplanes meet the applicable provisions of
14 CFR part 25, as amended by Amendments 25-1 through 25-98.
If the Administrator finds that the applicable airworthiness
regulations (i.e., part 25, as amended) do not contain adequate or
appropriate safety standards for the Embraer Model 170-100 and 170-200
airplanes because of novel or unusual design features, special
conditions are prescribed under the provisions of Sec. 21.16.
In addition to the applicable airworthiness regulations and special
conditions, the Embraer Model 170-100 and 170-200 airplanes must comply
with the fuel vent and exhaust emission requirements of 14 CFR part 34
and the noise certification requirements of 14 CFR part 36, and the FAA
must issue a finding of regulatory adequacy pursuant to section 611 of
Public Law 93-574, the ``Noise Control Act of 1972.''
Special conditions, as defined in 14 CFR 11.19, are issued in
accordance with Sec. 11.38 and become part of the type certification
basis in accordance with Sec. 21.17(a)(2), Amendment 21-69, effective
September 16, 1991.
Special conditions are initially applicable to the model for which
they are issued. Should the type certificate for that model be amended
later to include any other model that incorporates the same novel or
unusual design feature, or should any other model already included on
the same type certificate be modified to incorporate the same novel or
unusual design features, the special conditions would also apply to the
other model under the provisions of Sec. 21.101(a)(1), Amendment 21-
69, effective September 16, 1991.
Novel or Unusual Design Features
The Embraer Model 170-100 and 170-200 airplanes will incorporate
the following novel or unusual design features:
Engine Size and Torque Load
Since 1957 the limit engine torque load which is posed by sudden
engine stoppage due to malfunction or structural failure--such as
compressor jamming--has been a specific requirement for transport
category airplanes. Design torque loads associated with typical failure
scenarios were estimated by the engine manufacturer and provided to the
airframe manufacturer as limit loads. These limit loads were considered
simple, pure torque static loads. However, the size, configuration, and
failure modes of jet engines have changed considerably from those
envisioned when the engine seizure requirement of Sec. 25.361(b) was
first adopted. Current engines are much larger and are now designed
with large bypass fans capable of producing much larger torque, if they
become jammed.
Relative to the engine configurations that existed when the rule
was developed in 1957, the present generation of engines are
sufficiently different and novel to justify issuance of special
conditions to establish appropriate design standards. The latest
generation of jet engines are capable of producing, during failure,
transient loads that are significantly higher and more complex than the
generation of engines that were present when the existing standard was
developed. Therefore, the FAA has determined that special conditions
are needed for the Embraer Model 170-100 and 170-200 airplanes.
Electrical and Electronic Systems Which Perform Critical Functions
The Embraer Model 170-100 and 170-200 airplanes will have an
electronic flight control system which performs critical functions. The
current airworthiness standards of part 25 do not contain adequate or
appropriate standards for the protection of this system from the
adverse effects of operations without normal electrical power.
Accordingly, this system is considered to be a novel or unusual design
feature. Since the loss of normal
[[Page 5243]]
electrical power may be catastrophic to the airplane, special
conditions are proposed to retain the level of safety envisioned by 14
CFR 25.1351(d).
Interactions of Systems and Structures
The Embraer Model 170-100 and 170-200 airplanes will have systems
that affect the structural performance of the airplane, either directly
or as a result of a failure or malfunction. These novel or unusual
design features are systems that can alleviate loads in the airframe
and, when in a failure state, can create loads in the airframe. The
current regulations do not adequately account for the effects of these
systems and their failures on structural performance.
Discussion
Engine Size and Torque Loads
In order to maintain the level of safety envisioned in 14 CFR
25.361(b), a more comprehensive criterion is needed for the new
generation of high bypass engines. The proposed special conditions
would distinguish between the more common seizure events and those
rarer seizure events resulting from structural failures. For the rare
but severe seizure events, the proposed criteria allow some deformation
in the engine supporting structure (ultimate load design) in order to
absorb the higher energy associated with the high bypass engines, while
at the same time protecting the adjacent primary structure in the wing
and fuselage by providing a higher safety factor. The criteria for the
more severe events would no longer be a pure static torque load
condition, but would account for the full spectrum of transient dynamic
loads developed from the engine failure condition.
Electrical and Electronic Systems Which Perform Critical Functions
The Embraer Model 170-100 and 170-200 airplanes will require a
continuous source of electrical power for the electronic flight control
systems. Section Sec. 25.1351(d), ``Operation without normal
electrical power,'' requires safe operation in visual flight rule (VFR)
conditions for a period of not less than five minutes with inoperative
normal power. This rule was structured around a traditional design
utilizing mechanical connections between the flight control surfaces
and the pilot controls. Such traditional designs enable the flightcrew
to maintain control of the airplane while taking the time to sort out
the electrical failure, start engines if necessary, and re-establish
some of the electrical power generation capability.
The Embraer Model 170-100 and 170-200 airplanes will utilize an
electronic flight control system for the pitch and yaw control
(elevator, stabilizer, and rudder). There is no mechanical linkage
between the pilot controls and these flight control surfaces. Pilot
control inputs are converted to electrical signals which are processed
and then transmitted via wires to the control surface actuators. At the
control surface actuators, the electrical signals are converted to an
actuator command, which moves the control surface.
In order to maintain the same level of safety as an airplane with
conventional flight controls, an airplane with electronic flight
controls, such as the Embraer Model 170, must not be time limited in
its operation, including being without the normal source of electrical
power generated by the engine or the Auxiliary Power Unit (APU)
generators.
Service experience has shown that the loss of all electrical power
generated by the airplane's engine generators or APU is not extremely
improbable. Thus, it must be demonstrated that the airplane can
continue safe flight and landing (including steering and braking on
ground) after total loss of the normal electrical power with only the
use of its emergency electrical power systems. These emergency
electrical power systems must be able to power loads that are essential
for continued safe flight and landing. The emergency electrical power
system must be designed to supply electrical power for the following:
[sbull] Immediate safety, without the need for crew action,
following the loss of the normal engine generator electrical power
system (which includes APU power), and
[sbull] Continued safe flight and landing, and
[sbull] Restarting the engines.
For compliance purposes, a test of the loss of normal engine
generator power must be conducted to demonstrate that when the failure
condition occurs during night Instrument Meteorological Conditions
(IMC), at the most critical phase of the flight relative to the
electrical power system design and distribution of equipment loads on
the system, the following conditions are met:
1. After the unrestorable loss of normal engine and APU generator
power, the airplane engine restart capability must be provided and
operations continued in IMC.
2. The airplane is demonstrated to be capable of continued safe
flight and landing. The length of time must be computed based on the
maximum diversion time capability for which the airplane is being
certified. Consideration for speed reductions resulting from the
associated failure must be made.
3. The availability of APU operation should not be considered in
establishing emergency power system adequacy.
Interaction of Systems and Structure
The Embraer Model 170 has systems that affect the structural
performance of the airplane. These systems can serve to alleviate loads
in the airframe and, when in a failure state, can create loads in the
airframe. This degree of system and structures interaction was not
envisioned in the structural design regulations of 14 CFR part 25. This
proposed special condition provides comprehensive structural design
safety margins as a function of systems reliability.
Applicability
As discussed above, these special conditions are applicable to the
Embraer Model 170-100 and 170-200 airplanes. Should Embraer apply at a
later date for a change to the type certificate to include another
model incorporating the same novel or unusual design features, these
special conditions would apply to that model as well under the
provisions of Sec. 21.101(a)(1), Amendment 21-69, effective September
16, 1991.
Conclusion
This action affects only certain novel or unusual design features
on the Embraer Model 170-100 and 170-200 airplanes. It is not a rule of
general applicability, and it affects only the applicant 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 Embraer Model 170-100 and 170-200 airplanes.
Sudden Engine Stoppage. In lieu of compliance with 14 CFR
25.361(b), the following special conditions apply:
1. For turbine engine installations: The engine mounts, pylons and
adjacent supporting airframe structure must be designed to withstand 1g
level flight loads acting simultaneously with the
[[Page 5244]]
maximum limit torque loads imposed by each of the following:
a. Sudden engine deceleration due to a malfunction which could
result in a temporary loss of power or thrust.
b. The maximum acceleration of the engine.
2. For auxiliary power unit installations: The power unit mounts
and adjacent supporting airframe structure must be designed to
withstand 1g level flight loads acting simultaneously with the maximum
limit torque loads imposed by each of the following:
a. Sudden auxiliary power unit deceleration due to malfunction or
structural failure.
b. The maximum acceleration of the auxiliary power unit.
3. For an engine supporting structure: An ultimate loading
condition must be considered that combines 1g flight loads with the
transient dynamic loads resulting from each of the following:
a. The loss of any fan, compressor, or turbine blade.
b. Where applicable to a specific engine design, and separately
from the conditions specified in paragraph 3.a., any other engine
structural failure that results in higher loads.
4. The ultimate loads developed from the conditions specified in
paragraphs 3.a. and 3.b. above must be multiplied by a factor of 1.0
when applied to engine mounts and pylons and multiplied by a factor of
1.25 when applied to adjacent supporting airframe structure.
Operation Without Normal Electrical Power. In lieu of compliance
with 14 CFR 25.1351(d), the following special conditions apply:
It must be demonstrated by test or by a combination of test and
analysis, that the airplane can continue safe flight and landing with
inoperative normal engine and APU generator electrical power (in other
words, without electrical power from any source, except for the battery
and any other standby electrical sources). The airplane operation
should be considered at the critical phase of flight and include the
ability to restart the engines and maintain flight for the maximum
diversion time capability being certified.
Interaction of Systems and Structures: In lieu of compliance with
14 CFR 25.1351(d), the following special conditions apply:
1. General: For airplanes equipped with systems that affect
structural performance, either directly or as a result of a failure or
malfunction, the influence of these systems and their failure
conditions must be taken into account when showing compliance with the
requirements of 14 CFR part 25, subparts C and D. The following
criteria must be used for showing compliance with these special
conditions for airplanes equipped with flight control systems,
autopilots, stability augmentation systems, load alleviation systems,
flutter control systems, and fuel management systems. If these special
conditions are used for other systems, it may be necessary to adapt the
criteria to the specific system.
(a) The criteria defined herein address only the direct structural
consequences of the system responses and performances and cannot be
considered in isolation but should be included in the overall safety
evaluation of the airplane. These criteria may in some instances
duplicate standards already established for this evaluation. These
criteria are only applicable to structures whose failure could prevent
continued safe flight and landing. Specific criteria that define
acceptable limits on handling characteristics or stability requirements
when operating in the system degraded or inoperative modes are not
provided in these special conditions.
(b) Depending upon the specific characteristics of the airplane,
additional studies that go beyond the criteria provided in these
special conditions may be required in order to demonstrate the
capability of the airplane to meet other realistic conditions, such as
alternative gust or maneuver descriptions, for an airplane equipped
with a load alleviation system.
(c) The following definitions are applicable to these special
conditions.
Structural performance: Capability of the airplane to meet the
structural requirements of 14 CFR part 25.
Flight limitations: Limitations that can be applied to the airplane
flight conditions following an in-flight occurrence and that 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., fuel, payload, and Master Minimum Equipment List
limitations).
Probabilistic terms: The probabilistic terms (probable, improbable,
extremely improbable) used in these special conditions are the same as
those used in Sec. 25.1309.
Failure condition: The term failure condition is the same as that
used in Sec. 25.1309; however, these special conditions apply only to
system failure conditions that affect the structural performance of the
airplane (e.g., system failure conditions that induce loads, lower
flutter margins, or change the response of the airplane to inputs such
as gusts or pilot actions).
2. Effects of Systems on Structures. The following criteria will be
used in determining the influence of a system and its failure
conditions on the airplane structure.
(a) System fully operative. With the system fully operative, the
following apply:
(1) Limit loads must be derived in all normal operating
configurations of the system from all the limit conditions specified in
subpart C, taking into account any special behavior of such a system or
associated functions, or any effect on the structural performance of
the airplane that may occur up to the limit loads. In particular, any
significant 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 system 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 will not allow it to exceed those limit
conditions.
(3) The airplane must meet the aeroelastic stability requirements
of Sec. 25.629.
(b) 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.
(i) For static strength substantiation, these loads multiplied by
an appropriate factor of safety that is related to the probability of
occurrence of the failure are ultimate loads to be considered for
design. The factor of safety (FS) is defined in Figure 1.
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[[Page 5245]]
[GRAPHIC] [TIFF OMITTED] TP03FE03.014
(ii) For residual strength substantiation, the airplane must be
able to withstand two-thirds of the ultimate loads defined in paragraph
2.(b)(1)(i) above.
(iii) Freedom from aeroelastic instability must be shown up to the
speeds defined in Sec. 25.629(b)(2). For failure conditions that
result in speed increases beyond Vc/Mc, freedom from aeroelastic
instability must be shown to increased speeds, so that the margins
intended by Sec. 25.629(b)(2) are maintained.
(iv) Failures of the system that result in forced structural
vibrations (oscillatory failures) must not produce loads that could
result in detrimental deformation of primary structure.
(2) For the continuation of the flight. For the airplane in the
system failed state and considering any appropriate reconfiguration and
flight limitations, the following apply:
(i) The loads derived from the following conditions at speeds up to
Vc, or the speed limitation prescribed for the remainder of the flight,
must be determined:
(A) The limit symmetrical maneuvering conditions specified in
Sec. Sec. 25.331 and 25.345.
(B) The limit gust and turbulence conditions specified in
Sec. Sec. 25.341 and 25.345.
(C) The limit rolling conditions specified in Sec. 25.349, and the
limit unsymmetrical conditions specified in Sec. 25.367 and Sec.
25.427(b) and (c).
(D) The limit yaw maneuvering conditions specified in Sec. 25.351.
(E) The limit ground loading conditions specified in Sec. Sec.
25.473 and 25.491.
(ii) For static strength substantiation, each part of the structure
must be able to withstand the loads defined in paragraph 2.(b)(2)(i)
above, multiplied by a factor of safety depending on the probability of
being in this failure state. The factor of safety is defined in Figure
2.
[GRAPHIC] [TIFF OMITTED] TP03FE03.015
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 applied to all
limit load conditions specified in subpart C.
(iii) For residual strength substantiation, the airplane must be
able to withstand two thirds of the ultimate loads defined in paragraph
2.(b)(2)(ii) above.
(iv) If the loads induced by the failure condition have a
significant effect on fatigue or damage tolerance, then their effects
must be taken into account.
(v) Freedom from aeroelastic instability must be shown up to a
speed determined from Figure 3. Flutter clearance speeds VI
and VII may be based on the speed limitation specified for
the remainder of the flight using the margins defined by Sec.
25.629(b).
[[Page 5246]]
[GRAPHIC] [TIFF OMITTED] TP03FE03.016
VI = Clearance speed as defined by Sec. 25.629(b)(2).
VII = Clearance speed as defined by 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
VII.
(vi) Freedom from aeroelastic instability must also be shown up to
VI in Figure 3 above for any probable system failure
condition combined with any damage required or selected for
investigation by Sec. 25.571(b).
(3) Consideration of certain failure conditions may be required by
other sections of 14 CFR part 25, regardless of calculated system
reliability. Where analysis shows the probability of these failure
conditions to be less than 10-9, criteria other than those
specified in this paragraph may be used for structural substantiation
to show continued safe flight and landing.
(c) Warning considerations. For system failure detection and
warning, the following apply:
(1) The system must be checked for failure conditions, not
extremely improbable, that degrade the structural capability below the
level required by 14 CFR part 25, or significantly reduce the
reliability of the remaining system. The flightcrew must be made aware
of these failures before flight. Certain elements of the control
system, such as mechanical and hydraulic components, may use special
periodic inspections, and electronic components may use daily checks,
in lieu of warning systems, to achieve the objective of this
requirement. These certification maintenance requirements must be
limited to components that are not readily detectable by normal warning
systems and where service history shows that inspections will provide
an adequate level of safety.
(2) The existence of any failure condition, not extremely
improbable, during flight that could significantly affect the
structural capability of the airplane, and for which the associated
reduction in airworthiness can be minimized by suitable flight
limitations, must be signaled to the flightcrew. For example, failure
conditions that result in a factor of safety between the airplane
strength and the loads of 14 CFR part 25, subpart C below 1.25, or
flutter margins below VII, must be signaled to the crew
during flight.
(d) Dispatch with known failure conditions. If the airplane is to
be dispatched in a known system failure condition that affects
structural performance, or affects the reliability of the remaining
system to maintain structural performance, then the provisions of these
special conditions must be met for the dispatched condition and for
subsequent failures. Flight limitations and expected operational
limitations may be taken into account in establishing Qj as the
combined probability of being in the dispatched failure condition and
the subsequent failure condition for the safety margins in Figures 2
and 3. These limitations must be such that the probability of being in
this combined failure state and then subsequently encountering limit
load conditions is extremely improbable. No reduction in these safety
margins is allowed if the subsequent system failure rate is greater
than 10-3 per hour.
Issued in Renton, Washington, on January 9, 2003.
Ali Bahrami,
Acting Manager, Transport Airplane Directorate, Aircraft Certification
Service.
[FR Doc. 03-2423 Filed 1-31-03; 8:45 am]
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