[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
