[Federal Register Volume 68, Number 78 (Wednesday, April 23, 2003)]
[Rules and Regulations]
[Pages 19933-19937]
From the Federal Register Online via the Government Publishing Office [www.gpo.gov]
[FR Doc No: 03-10044]



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  Federal Register / Vol. 68, No. 78 / Wednesday, April 23, 2003 / 
Rules and Regulations  

[[Page 19933]]



DEPARTMENT OF TRANSPORTATION

Federal Aviation Administration

14 CFR Part 25

[Docket No. NM246; Special Conditions No. 25-231-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: Final special conditions.

-----------------------------------------------------------------------

SUMMARY: These special conditions are issued 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 engine size and torque load which 
affect sudden engine stoppage, electrical and electronic flight control 
systems which perform critical functions, and 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 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.

EFFECTIVE DATE: April 10, 2003.

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:

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 Sec.  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 is 
sufficiently different and novel to justify issuance of special 
conditions to establish appropriate design standards. The latest 
generation of jet engines is 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

[[Page 19934]]

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 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. These 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 specified 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 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. 
These special conditions provide comprehensive structural design safety 
margins as a function of systems reliability.

Discussion of Comments

    Notice of proposed special conditions No. NM246 for the Embraer 
Model 170-100 and 170-200 airplanes was published in the Federal 
Register dated February 3, 2003 (68 FR 5241). No comments were 
received, and the special conditions are adopted as proposed.

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.
    Under standard practice, the effective date of final special 
conditions would be 30 days after the date of publication in the 
Federal Register; however, as the certification date for the Embraer 
Model 170-100 and -200 is imminent, the FAA finds good cause to make 
these special conditions effective upon issuance.

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.

[[Page 19935]]

List of Subjects in 14 CFR Part 25

    Aircraft, Aviation safety, Reporting and recordkeeping 
requirements.


0
The authority citation for 14 CFR part 25, for these special 
conditions, is as follows:


    Authority: 49 U.S.C. 106(g), 40113, 44701, 44702, 44704.

PART 25--AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES

The Special Conditions

0
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 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 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 the 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.

[[Page 19936]]

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.
[GRAPHIC] [TIFF OMITTED] TR23AP03.000

    (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] TR23AP03.001

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

[[Page 19937]]

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).
[GRAPHIC] [TIFF OMITTED] TR23AP03.002

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 April 10, 2003.
Ali Bahrami,
Acting Manager, Transport Airplane Directorate, Aircraft Certification 
Service.
[FR Doc. 03-10044 Filed 4-22-03; 8:45 am]
BILLING CODE 4910-13-P