[Federal Register Volume 60, Number 65 (Wednesday, April 5, 1995)]
[Rules and Regulations]
[Pages 17194-17196]
From the Federal Register Online via the Government Publishing Office [www.gpo.gov]
[FR Doc No: 95-8371]



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DEPARTMENT OF TRANSPORTATION

Federal Aviation Administration

14 CFR Part 25

[Docket No. NM-105, Special Conditions No. 25-ANM-97]


Special Conditions: Saab Aircraft AB Model Saab 2000 Series 
Airplanes

AGENCY: Federal Aviation Administration, DOT.

ACTION: Final special conditions.

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SUMMARY: These special conditions are for the Saab Aircraft AB Model 
Saab 2000 airplane. This airplane will have novel and unusual design 
features, relating to its electronic flight control system, 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 which the 
Administrator considers necessary to establish a level of safety 
equivalent to that provided by the airworthiness standards of part 25.

EFFECTIVE DATE: March 29, 1995.

FOR FURTHER INFORMATION CONTACT:
Mark I. Quam, FAA, Standardization Branch, ANM-113, Transport Standards 
Staff, Transport Airplane Directorate, Aircraft Certification Service, 
1601 Line Avenue SW, Renton, Washington 98055-4056; telephone (206) 
227-2145, facsimile (206) 277-1320.

SUPPLEMENTARY INFORMATION

Background

    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. The new Saab 2000 incorporates a 
number of such design features.
    The Saab 2000, certificated on April 29, 1994, is a twin-engined, 
low-wing, pressurized turboprop aircraft that is configured for 
approximately 50 passengers. The airplane has two Allison Engine 
Company AE 2100A engines rated at 3650 shp. The propeller is a 6 bladed 
Dowty Rotol swept shaped propeller. A single lever controls each prop/
engine combination. An Auxiliary Power Unit (APU) will be installed in 
the tail. The airplane has provisions for two pilots, an observer, two 
flight attendants, overhead bins, a toilet, and provisions for the 
installation of a galley. There is a forward and aft stowage 
compartment and an aft cargo compartment. The airplane has a maximum 
operating altitude of 31,000 feet.
    The Saab 2000 has a fully hydraulically powered electronically 
controlled rudder and will have fully hydraulically powered 
electronically controlled elevators as a follow-on design modification. 
The Powered Elevator Control System (PECS) provides control and power 
actuation of the left and right elevator surfaces. The PECS also 
provides aircraft stability augmentation and trim functions.
    The proposed elevator system is in many respects similar to the 
rudder design and is comprised of a mix of analog and digital circuitry 
and has no mechanical backup. Control columns are connected to Linear 
Variable Differential Transducers (LVDT), stick damper(s), auto pilot 
servo, linear springs with break-outs and are interconnected with an 
electronic disconnect unit.
    The position transducers (LVDT), connected to the control columns, 
provide signals to two Powered Elevator Control Units (PECU). Each PECU 
controls two Elevator Servo Actuators (ESA) through two separate Servo 
Actuator Channels (SAC). Each SAC is subdivided into a primary control 
lane and a monitor lane. Two of the four ESAs, controlled by one PECU, 
positions one elevator side.
    The ESAs have two modes of operation, active and damped. The active 
mode will result when mode control current from the PECU and hydraulic 
pressure are available. One active servo actuator is sufficient to 
operate the elevator surface.
    Elevator Servo Actuators valve and actuator ram position feedback 
are provided by position transducers (LVDT). The PECUs are connected to 
one Flight Control Computer via the trim relay and two Digital Air Data 
Computers. The flight control computer also provides a signal to the 
auto pilot servo.
    Stick to elevator gearing is a function of Indicated Airspeed 
(IAS). Trim and stability augmentation are based on IAS, vertical 
acceleration and flap position. Stick, trim and elevator position and 
status information are fed to the Engine Indicating and Crew Alerting 
System (EICAS).
    Each PECU has built in Automatic Preflight Built in test (PBIT) and 
Continuous Built In Test (CBIT) circuitry and utilizing cross channel 
nomitoring.
    The elevator's actuators are supplied by three hydraulic circuits 
that are physically separated, isolated, fused and located to minimize 
common cause failures. The Number 1 hydraulic circuit is powered by the 
left engine and a backup DC pump and accumulators. The Number 2 
hydraulic circuit is powered by the right engine and a backup AC pump 
and accumulators. THe Number 3 hydraulic circuit is powered by an AC 
drive pump.
    The Number 1 hydraulic circuit powers the left hand (LH) and right 
hand (RH) outboard servo actuators. The Number 2 hydraulic circuit 
powers the RH inboard servo actuator. the Number 3 hydraulic circuit 
powers the LH inboard servo actuator.
    Hydraulic warnings and cautions in the event of hydraulic supply 
failure are provided by the EICAS.
    The elevator system is electrically supported by two system sides, 
a LH and a RH side. The electrical system is normally powered by two AC 
generators, each driven by a propeller gear box. An APU equipped with a 
standby generator is installed. When only one of the three generators 
is working, it supplies power to both LH and RH sides.
    Each LH and RH AC system side is connected via a Transformer 
Rectifier Unit (TRU) to a LH and RH DC system made up of a network of 
DC buses. A third center TRU is connected to a center circuit. The LH, 
RH and center buses can be supplied from batteries or from the TRUs. 
The center TRU will replace a failed RH or LH TRU. When only one TRU 
unit is working, the LH and RH buses are tied together with power being 
received from the remaining TRU.
    Two DC feeders in addition to two AC feeders provide power aft of 
the debris zone. The LH side is routed through the ceiling and the RH 
side is routed through the floor.

Type Certification Basis

    The applicable requirements for U.S. type certification must be 
established in accordance with Secs. 21.16, 21.17, 21.19, 21.29, and 
21.101 of the FAR. Accordingly, based on the application date of June 
9, 1989, and Saab Aircraft AB volunteering for certain later 
regulations, the TC basis for the Saab 2000 airplane is as 
follows: [[Page 17195]] 
    Part 25 as amended by Amendments 25-1 through 25-71, except:
Sec. 25.361  Engine torque, as amended by Amendment 25-72.
Sec. 25.365  Pressurized compartment loads, as amended by Amendment 25-
72.
Sec. 25.571  Damage tolerance and fatigue evaluation of structure, as 
amended by Amendment 25-72.
Sec. 25.772  Pilot compartment doors, as amended by Amendment 25-72.
Sec. 25.773  Pilot compartment view, as amended by Amendment 25-72.
Sec. 25.783(g) Doors, as amended by Amendment 25-72.
Sec. 25.905(d) Propellers, as amended by Amendment 25-72.
Sec. 25.933  Reversing systems, as amended by Amendment 25-72.
Secs. 25.903 and 25.951 as amended by Amendment 25-73.
Secs. 25.851 and 25.854 as amended by Amendment 25-74.
Sec. 25.729 as amended by Amendment 25-75.
Sec. 25.813 as amended by Amendment 25-76.

    Part 34, as amended on the date of issuance of the type 
certificate.
    Part 36, as amended on the date of issuance of the type 
certificate.
    Special Conditions No. 25-ANM-66, dated 1/12/93, for Lightning and 
HIRF Protection.
    Special Conditions No. 25-ANM-82, dated 3/11/94, for Interaction of 
Systems and Structure.
    Special conditions, as appropriate, are issued in accordance with 
Sec. 11.49 of the FAR after public notice, as required by Secs. 11.28 
and 11.29(b), and become part of the type certification basis in 
accordance with Sec. 21.101(b)(2).

Discussion

    Special Conditions No. 25-ANM-82 were written for the rudder and in 
anticipation of the installation of the powered elevator. However, as 
the Saab 2000 could be flown without rudder control during certain 
failure conditions, and the elevator system was not installed for 
initial certification, Special Conditions No. 25-ANM-82 were limited to 
requirements common to both the rudder and follow-on elevator. The Saab 
2000, however, requires control and power to the elevator all the time 
for safe flight and landing. Therefore, these special conditions 
supplement Special Conditions No. 25-ANM-82 for the powered elevator. 
The proposed type design of the Saab 2000 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:

Systems

    1. Operation Without Normal Electrical Power. In the Saab 2000, a 
source of electrical power is required by the elevator electronic 
flight control system. Service experience with traditional airplane 
designs has shown that the loss of electrical power generated by the 
airplane's engines is not extremely improbable. The electrical power 
system of the Saab 2000 must therefore be designed with standby or 
emergency electrical sources of sufficient reliability and capacity to 
power essential loads in the event of the loss of normally generated 
electrical power. The need for electrical power for electronic flight 
controls was not envisioned by part 25 since in traditional designs, 
cables and hydraulics are utilized for the flight control system. 
Therefore, Special Condition No. 1 is adopted as proposed.
    2. Command Signal Integrity. Command and control of the control 
surfaces will be achieved by fly-by-wire systems that will utilize 
electronic (AC, DC, or digital) interfaces. These interfaces involve 
not only the commands to the control surfaces, but all the control 
feedback and sensor input signals as well. These signal paths, as well 
as the electronic equipment that manages them, can be susceptible to 
damage that may cause unacceptable or unwanted control responses. The 
damage may originate from electrical equipment failures, mechanical 
equipment failures or external damage. Therefore, special designs are 
needed to maintain the integrity of the fly-by-wire interfaces to an 
immunity level equivalent to that of traditional hydro-mechanical 
designs. Similar to the conventional steel cable controls, positioning 
of the electrical control equipment and routing of wire bundles must 
provide separation and redundancy to ensure maximum protection from 
damage due to a common cause. Therefore, Special Condition No. 2 is 
adopted as proposed.
    3. Design Maneuver Requirements. In a conventional airplane, pilot 
inputs directly affect control surface movement (both rate and 
displacement) for a given flight condition. In the Saab 2000, the pilot 
provides only one of several inputs to the control surfaces, and it is 
possible that the pilot control displacements specified in 
Secs. 25.331(c)(1), 25.349(a), and 25.351 of the FAR may not result in 
the maximum displacement and rates of displacement of the elevator. The 
intent of these noted rules may not be satisfied if literally applied. 
Therefore, Special Condition No. 3 is adopted as proposed.

Discussion of Comments

    Notice of Proposed Special Conditions No. SC-95-1-NM for the Saab 
Aircraft AB Model Saab 2000 Series Airplanes was published in the 
Federal Register on February 2, 1995 (60 FR 6456). No comments were 
received.
    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 will become part of the type certification basis in 
accordance with Sec. 21.17(a)(2).
    Under standard practice, the effective date of these special 
conditions would be 30 days after publication in the Federal Register. 
As the intended U.S. type certification date for the Saab 2000 is April 
1, 1995, the FAA finds that good cause exists to make these special 
conditions effective upon issuance.

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 airplanes.

List of Subjects in 14 CFR Part 25

    Air transportation, Aircraft, Aviation safety, Safety.

    The authority citation for these proposed special conditions is as 
follows:

    Authority: 49 U.S.C. 1344, 1348(c), 1352, 1354(a), 1355, 1421 
through 1431, 1502, 1651(b)(2); 42 U.S.C. 1857f-10, 4321 et seq.; 
E.O. 11514; and 49 U.S.C. 106(g).

The Special Conditions

    Accordingly, the following special conditions are issued as part of 
the type certification basis for the Saab Aircraft AB Model Saab 2000 
series airplanes.

    1. Operations Without Normal Electrical Power. In lieu of 
compliance with Sec. 25.1351(d), it must be demonstrated by test, or 
combination of test and analysis, that the airplane can continue 
safe flight and landing with inoperative normal engine generated 
electrical power (electrical power sources excluding 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 [[Page 17196]] engines and maintain flight 
for the maximum diversion time capability being certified.
    Discussion: The Electronic Flight Control System installations 
establish the criticality of the electrical power generation and 
distribution systems, since the loss of all electrical power may be 
catastrophic to the aircraft.
    The Saab 2000 fly-by-wire control system requires a continuous 
source of electrical power in order to maintain the flight control 
system. The current Sec. 25.1351(d), ``Operation Without Normal 
Electrical Power,'' requires safe operation in visual flight rules 
(VFR) conditions for at least five minutes with inoperative normal 
power. This rule was structured around a traditional design 
utilizing mechanical control cables for flight control while the 
crew took time to sort out the electrical failure and was able to 
re-establish some of the electrical power generation capability.
    In order to maintain the same level of safety associated with 
traditional designs, the Saab 2000 design must not be time limited 
in its operation without the normal source of engine generated 
electrical power. It should be noted that service experience has 
shown that the loss of all electrical power which is generated by 
the airplane's engines is not extremely improbable. Thus, it must be 
demonstrated that the airplane can continue safe flight and landing 
with the use of its emergency electrical power systems (batteries, 
auxiliary power unit, etc.). This emergency electrical power system 
must be able to power loads that are essential for continued safe 
flight and landing. Also, the availability of emergency electrical 
power sources, including any credit taken for APU start reliability, 
must be validated in a manner acceptable to the FAA.
    The emergency electrical power system must be designed to 
supply:
--electrical power required for immediate safety, which must 
continue to operate without the need for crew action following the 
loss of the normal electrical power system;
--electrical power required for continued safe-flight and landing;
--electrical power required to restart the engines.
    For compliance purposes:
    1. A test demonstration of the loss of normal engine generated 
power is to be established such that:
    a. The failure condition should be assumed to occur during night 
instrument meteorological conditions (IMC) at the most critical 
phase of flight relative to the electrical power system design and 
distribution of equipment loads on the system.
    b. After the unrestorable loss of the source of normal 
electrical power, the airplane engines must be capable of being 
restarted and operations continued in IMC until visual 
meteorological conditions (VMC) can be reached. (A reasonable 
assumption can be made that turbine engine driven transport category 
airplanes will not have to remain in IMC for more than 30 minutes 
after experiencing the loss of normal electrical power).
    c. After 30 minutes of operation in IMC, the airplane should be 
demonstrated to be capable of continuous safe flight and landing in 
VMC conditions. The length of time in VMC conditions must be 
computed based on the maximum flight duration capability for which 
the airplane is being certified. Consideration for speed reductions 
resulting from the associated failure must be made.
    2. Since the availability of the emergency electrical power 
system operation is necessary for safe-flight, this system must be 
available before each flight.
    3. The emergency electrical power system must be shown to be 
satisfactorily operational in all flight regimes.
    2. Command Signal Integrity. In addition to compliance with 
Sec. 25.671 of the FAR, it must be shown that for the elevator 
Electronic Flight Control System (EFCS):
    (a) Signals cannot be altered unintentionally, or that the 
altered signal characteristics are such that the control authority 
characteristics will not be degraded to a level that will prevent 
continued safe-flight and landing; and
    (b) Routing of wire EFCS wires and wire bundles must provide 
separation and redundancy to ensure maximum protection from damage 
due to common cause.
    Discussion: The Saab 2000 will be using fly-by-wire (FBW) as a 
means to command and control the elevator surface actuators. In the 
FBW design being presented, command and control of the control 
surfaces will be achieved by electronic (AC, DC, or digital) 
interfaces. These interfaces involve not only the direct commands to 
the elevator control surfaces, but feedback and sensor signals as 
well.
    Malfunctions could cause system instabilities, loss of functions 
or freeze-up of the control actuator. It is imperative that after 
failure at least one path of the command signal, that is capable of 
providing safe flight and landing, remains continuous and unaltered.
    The current regulations, which primarily address hydro-
mechanical flight control systems, Secs. 25.671 and 25.672, make no 
specific or implied reference that command and control signals 
remain unaltered from external interferences. Present designs 
feature steel cables and pushrods as a means to control hydraulic 
surface actuators. These designs are easily identifiable relative to 
the understanding that they are necessary for safe flight and 
landing and thus should be protected and continually inspected. 
However, the FBW designs are not easily discernible from non-
essential electronics where placement of equipment and wire runs is 
not critical. Therefore, FBW requires additional attention when 
locating the equipment and wire runs.
    It should be noted that:

--The wording ``signals cannot be altered unintentionally'' is used 
in the Special Condition to emphasize the need for design measures 
to protect the FBW control system from the effects of the 
fluctuations in electrical power, accidental damage, environmental 
factors such as temperature, local fires, exposure to reactive 
fluids, etc. and any disruptions that may affect the command signals 
as they are being transmitted from their source of origin to the 
Power Control Actuators.
    3. Design Maneuver Requirements
    (a) 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 in accordance with Sec. 25.397(b), 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) In addition to the requirements of Sec. 25.331(c), 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.

    Issued in Renton, Washington, on March 29, 1995.
Darrell M. Pederson,
Acting Manager, Transport Airplane Directorate, Aircraft Certification 
Service, ANM-100.
[FR Doc. 95-8371 Filed 4-4-95; 8:45 am]
BILLING CODE 4910-13-M