[Federal Register Volume 64, Number 95 (Tuesday, May 18, 1999)]
[Proposed Rules]
[Pages 26900-26922]
From the Federal Register Online via the Government Publishing Office [www.gpo.gov]
[FR Doc No: 99-12361]
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DEPARTMENT OF TRANSPORTATION
Federal Aviation Administration
14 CFR Part 25
[Docket No. NM156, Notice No. 25-99-04-SC]
Special Conditions: McDonnell Douglas Corporation (MDC) Model MD-
17 Series Airplanes
AGENCY: Federal Aviation Administration, DOT.
ACTION: Notice of proposed special conditions.
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SUMMARY: The FAA proposes to issue special conditions for the McDonnell
Douglas Corporation Model MD-17 airplane. This airplane will have novel
and unusual design features, including the use of power-augmented-lift
from externally blown flaps, for which the applicable airworthiness
standards for transport category airplanes do not contain adequate or
appropriate safety standards. This document contains the additional
safety standards that the Administrator considers necessary to
establish a level of safety equivalent to that provided by the existing
airworthiness standards.
DATES: Comments must be received on or before July 2, 1999.
ADDRESSES: Comments on this document may be mailed in duplicate to:
Federal Aviation Administration, Transport Airplane Directorate,
Program Management Branch, Attention: Rules Docket (ANM-114), Docket
No. NM156, 1601 Lind Avenue SW., Renton, WA 98055-4056; or delivered in
duplicate to the Transport Airplane Directorate at the above address.
Comments delivered must be marked Docket No. NM156. Comments may be
examined in the Rules Docket weekdays, except Federal holidays, between
7:30 a.m. and 4:30 p.m.
FOR FURTHER INFORMATION CONTACT: Gerry Lakin, Project Officer, FAA
Transport Airplane Directorate, Standardization Branch, ANM-113, 1601
Lind Avenue SW., Renton, WA 98055-4056; telephone (425) 227-1187;
facsimile (425) 227-1149; Email: [email protected].
SUPPLEMENTARY INFORMATION:
Comments Invited
Interested persons are invited to participate in the making of
these proposed special conditions by submitting such written data,
views, or arguments as they may desire. Communications should identify
the regulatory docket or notice number and be submitted in duplicate to
the Rules Docket address specified above. All communications received
on or before the closing date for comments will be considered by the
Administrator. The proposals described in this notice may be changed in
light of the comments received. All comments received will be available
in the Rules Docket for examination by interested persons, both before
and after the closing date for comments. A report summarizing each
substantive public contact with FAA personnel concerning this
rulemaking will be filed in the docket. Persons wishing the FAA to
acknowledge receipt of their comments must submit with those comments a
self-addressed, stamped postcard on which the following statement is
made: ``Comments to Docket No. NM156.'' The postcard will be date
stamped and returned to the commenter.
Background
On July 7, 1996, McDonnell Douglas Corporation, 2401 E. Wardlow
Rd., Long Beach, CA 90807-5309, a wholly owned subsidiary of The Boeing
Company, submitted an application for type certification of a
commercial version of the Model C-17 military airplane, designated as
the MDC Model MD-17. The MD-17 is a long range, transport category
airplane powered by four Pratt & Whitney F-117-PW-100 engines, which
are a military version of the PW2040 engines used on other civil
transport category airplane types. The airplane will be offered in a
cargo configuration only and is designed for carriage of outsized cargo
into short runways.
The MD-17 airplane will be certified as a part 25 transport
category airplane and, as such, pilots and flight instructors who
operate it will have a standard airplane multiengine rating.
Type Certification Basis
Under the provisions of Sec. 21.17, McDonnell Douglas must show
that the MD-17 complies with the applicable provisions of 14 CFR part
25, as amended by Amendments 25-1 through 25-87. In addition, the
certification basis includes part 36, as amended at the time of
certification; part 34, as amended at the time of certification; any
subsequent amendments to part 25 that are required for operation under
part 121; and the special conditions resulting from the proposals
specified in this notice.
If the Administrator finds that the applicable airworthiness
regulations (i.e., part 25) do not contain adequate or appropriate
safety standards for the MD-17 because of a novel or unusual design
feature, special conditions are prescribed under the provisions of
Sec. 21.16.
In addition to the applicable airworthiness regulations and special
conditions, the MD-17 must comply with the fuel vent and exhaust
emission requirements of part 34 and the noise certification
requirements of part 36, and the FAA must issue a finding of regulatory
adequacy pursuant to Sec. 611 of Pub. L. 92-574, the ``Noise Control
Act of 1972.''
[[Page 26901]]
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), and become part of the type certification basis in accordance
with Sec. 21.17(a)(2).
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, the special conditions would also apply to the
other model under the provisions of Sec. 21.101(a)(1).
MD-17 Design Features
The MD-17 has novel and unusual design features to support the
operation of a large transport category sized airplane at airports with
very short runways. The MD-17 has externally blown flaps (EBF), which
are fixed-vane double slotted flaps that deflect directly into the
engine exhaust stream. The MD-17 integrated EBF design includes
positioning the engines to provide engine exhaust blowing on the flaps,
and flap slots sized to provide engine exhaust flow over both the upper
and lower flap and vane surfaces. The resulting flap/exhaust stream
interaction provides power-augmented-lift relative to conventional
transport category airplane designs. The total lift produced by the EBF
is made up of three components: (1) conventional aerodynamic lift
produced by the wing and flap; (2) lift due to thrust deflection (the
vertical component of the thrust force); and (3) the powered
circulation lift (the additional aerodynamic lift resulting from the
interaction of the engine exhaust stream on the wing flaps).
To distinguish the new and novel power-augmented-lift design
feature of the MD-17 from conventional transport category airplanes,
the following definition has been established: Power-augmented-lift
means a heavier-than-air airplane capable of operation in regimes of
short field takeoff and short field landing, and low speed flight. The
airplane depends upon the propulsion system for a significant portion
of lift and control during these flight regimes, but relies primarily
on conventional wing lift when in the en route configuration.
The MD-17 features Direct Lift Control (DLC), which uses spoilers
to provide rapid control of the flight path angle in the down direction
for large flight path adjustments without throttle movement. DLC is
actuated via push button switches placed on both sides of the thrust
levers. Another feature of the MD-17 design that differs from
conventional transport category airplanes is that the spoilers are
biased to a non-flush position when the flaps are extended. When in
this configuration, separate from the DLC function, the spoilers are
linked to the thrust levers to provide airplane response equivalent to
instantaneous engine response to thrust lever movement.
The MD-17 Primary Flight Control System (PFCS) provides three-axis
control and envelope protection using conventional cockpit controls and
control surfaces, and a full authority fly-by-wire Electronic Flight
Control System (EFCS) with single-strand mechanical backup. The PFCS
provides stability and command augmentation to improve basic airplane
characteristics and also integrates the trim and high lift controls.
Pitch and roll control inputs are made through a one-handed center
stick controller centrally mounted to the floor in front of each pilot
station. In addition to four electronic displays, the cockpit display
system incorporates pilot and co-pilot full-time head up displays that
can be used as primary flight displays.
The MD-17 will utilize electrical and electronic systems that
perform critical functions. Examples of these systems include the
electronic displays and electronic engine controls.
As the proposed type design of the MD-17 contains novel or unusual
design features not envisioned by the applicable part 25 airworthiness
standards, special conditions are considered necessary in the following
areas:
Power-Augmented-Lift
1. Stall Speeds and Minimum Operating Speeds
The primary purpose of the EBF design feature on the MD-17 is to
reduce the takeoff and landing speeds, and hence the required takeoff
and landing distances. The benefits provided by this novel design
feature are not adequately addressed by the current part 25 stall speed
and minimum operating speeds requirements. A special condition is
needed to fully address the benefits of the MD-17 design features on
stall speeds and minimum operating speeds, and to provide appropriate
safety standards to ensure equivalent safety with current part 25
requirements.
The part 25 minimum allowable operating speeds are derived from
power-off (i.e., zero thrust or power) stall speeds (VS),
except in those instances where the operating speeds are limited by
some other constraint. Appropriate multiplying factors are applied to
these power-off stall speeds to provide adequate safety in the one-
engine-inoperative power-on condition. The beneficial effects of power-
on available lift due to both circulation effects and thrust
inclination were well known at the time the airworthiness requirements
were developed. Evidence for this point is provided by the requirements
associated with the minimum takeoff safety speed, V2MIN, in
Sec. 25.107(b). For airplanes without ``significant'' power-augmented-
lift effects in the one-engine-inoperative condition, V2MIN
must not be less than 1.20 VS, or 1.13 VS if the
1-g stall speed is used. However, for airplanes that realize a
significant reduction in stall speed in the one-engine-inoperative
power-on condition, the multiplying factor is reduced to 1.15.
According to the explanatory information associated with this
requirement that is provided in Civil Aeronautics Manual 4b, ``The
difference in the required factors * * * provides approximately the
same margin over the actual stalling speed under the power conditions
which are obtained after the loss of an engine. * * *''
The MD-17 power-augmented-lift design, however, achieves
significantly more lift from power than would be taken into account by
the part 25 requirements. At the conditions applicable to the
determination of the takeoff safety speed, V2, the MD-17
achieves a 15 percent reduction in power-on stall speed. The four
percent reduction in V2 speed permitted by Sec. 25.107(b)(2)
for ``turbojet powered airplanes with provisions for obtaining a
significant reduction in the one-engine-inoperative power-on stalling
speed'' would therefore not provide ``approximately the same margin
over the actual stalling speed as conventional transport category
airplanes in the one-engine-inoperative power-on condition.'' A further
reduction in V2 speed could be made while maintaining the
same margin over the one-engine-inoperative power-on stall speed.
At approach thrust, the MD-17 achieves over a 50 percent increase
in lift due to power-augmented-lift effects. In the maximum landing
flap configuration, the thrust used for a stable approach results in a
stall speed reduction of approximately 20 percent relative to the zero
thrust stall speed. There are no provisions in part 25, however, for
allowing the landing approach speed to be reduced to account for the
beneficial effects of power-augmented-lift on stall speeds. For a
conventional transport category airplane, thrust or power may vary
[[Page 26902]]
considerably during the landing approach, including reductions to idle
thrust or power. During the landing flare for a conventional transport
category airplane, thrust is typically reduced to idle.
The MD-17 power-augmented-lift design, however, requires a
significant thrust level to be maintained during the approach to remain
on the desired approach flight path. Unlike conventional transport
category airplanes, only minor thrust modulation may be necessary
during the approach to maintain or recover the desired flight path. The
MD-17 design features and operational procedures will discourage use of
thrust reductions to make flight path adjustments during approach.
Adjustments in speed are obtained through changes in airplane pitch
attitude during approach. In addition, the MD-17 is designed to provide
very stable controllability characteristics to allow very slow approach
speeds using a backside control technique, which is explained later in
this preamble. With the backside control technique, airplane pitch
attitude is used to control airspeed and thrust is used to control
flight path angle.
As stated earlier, the MD-17 incorporates a DLC feature, which uses
the spoilers to provide rapid control of the flight path angle in the
down direction for large flight path adjustments without throttle
movement. DLC is actuated via push button switches placed on both sides
of the thrust levers. Separate from the DLC function, the spoilers are
biased to a non-flush position in the flaps extended configurations. In
this configuration, the spoilers are linked to the thrust levers to
provide an airplane response equivalent to instantaneous engine
response to thrust lever movement. This feature provides a high level
of control feedback and further minimizes the need for thrust
adjustments. Because of the unique characteristics of the MD-17 power-
augmented-lift design, thrust reduction is not used to reduce the rate
of descent at touchdown. Instead, a slight thrust increase may
sometimes be used to accomplish this task when desired.
To establish a level of safety equivalent to that established in
the regulations, the MD-17 minimum operating speeds should provide
approximately the same margin over the stall speed as conventional
transport category airplanes under the power conditions that are
obtained after the loss of an engine. In a power-augmented-lift
airplane like the MD-17, significant increases in lift capability can
be achieved not only by increasing angle of attack, but also by
increasing thrust. During the takeoff phase of flight, there is no
capability to add lift due to power because operation is already based
on the use of the maximum thrust available. For approach and landing,
however, the lift reserve due to thrust is much greater than that
available on conventional transport category airplanes. A rapid lift
increase due to increasing thrust is achievable on the MD-17 because it
uses not only a higher approach power setting than conventional
transport category airplanes, but also spoiler modulation to compensate
for engine spool-up time. The higher approach power setting is
necessary to compensate for the high induced drag from the power-
augmented-lift effects, and to compensate for the relatively high
profile drag of the approach and landing configurations, which include
spoilers that are biased in the up direction. Advancing the thrust
levers modulates the spoilers such that engine spool-up time is
compensated for and a rapid increase in lift is achieved.
In addition, the MD-17 design incorporates a feature in which the
deployed spoilers will be retracted should the airplane exceed a
predetermined angle-of-attack that is less than the stall angle-of-
attack. The stall speeds are defined assuming that the spoilers are
flush to the wing at the point of stall. McDonnell Douglas must
demonstrate to the FAA that the probability of the failure of any
system that could change the calculated stall speeds by one-half knot
or more is improbable.
Because there is no regulatory requirement to determine one-engine-
inoperative power-on stall speeds, there is only limited data available
to the FAA for assessing the margins attained under these conditions by
the current fleet of conventional transport category airplanes. Based
on the limited data that are available, and on the precedent
established by Civil Air Regulations part 4b and part 25 for powered-
lift credit, on average, conventional transport category airplanes
without provisions for obtaining significant lift from power obtain
approximately a 4-5 percent reduction in stall speed in the one-engine-
inoperative power-on condition. This 4-5 percent reduction in stall
speed applies to both the takeoff configuration at takeoff power and
the landing configuration at the power for a 3-degree glideslope.
To retain equivalent safety, the MD-17 minimum operating speed in
the takeoff configuration, V2, should retain the additional
4-5 percent safety margin in the one-engine-inoperative power-on stall
speed currently obtained on conventional transport category airplanes.
To use one-engine-inoperative power-on stall speeds to determine
V2MIN for the MD-17, the multiplying factor used to derive
V2MIN from power-off stall speeds for conventional transport
category airplanes should therefore be increased by not less than 4
percent (i.e., V2MIN must be 1.18 times the power-on 1-g
stall speed, rather than 1.13 times the power-off 1-g stall speed). In
determining the thrust effects on stall speeds for V2MIN
determination, the thrust or power on the operating engines should be
no greater than the minimum power that may exist at any point in the
takeoff flight path. This means that the takeoff (or derated takeoff)
power or thrust for the minimum engine would normally be determined at
a height of 1500 feet above the runway surface at the appropriate
takeoff power setting for the conditions existing at the time of
takeoff. However, if the effect of altitude on takeoff thrust or power
up to 1500 feet above the runway surface has a negligible impact on
power-on stall speed used for V2MIN determination, thrust or
power at the runway height may be used. McDonnell Douglas has provided
the FAA with data which show, for the MD-17 power-augmented-lift
design, that the effect of altitude on takeoff thrust up to 1500 feet
above the runway surface has a negligible (less than 0.5 knots) impact
on MD-17 power-on stall speeds used for V2MIN determination.
As noted above, the MD-17 incorporates several design features and
operating characteristics that result in significant fundamental
differences from the way conventional transport category airplanes are
flown in the approach and landing phase of flight. During approach to
landing, the MD-17's power-augmented-lift allows it to fly at speeds
that are less than the speed at which total airplane drag is a minimum.
Therefore, the MD-17 will be operating on the ``backside'' of the drag
(or power) curve, which means that drag increases as speed is reduced
and drag is reduced as speed increases. This variation of drag with
speed is in the opposite sense to that normally encountered on
conventional transport category airplanes operating at higher approach
speeds.
A significant consequence of operating on the backside of the drag
curve is that MD-17 pilots will use a different technique for
controlling airspeed and flight path than is used on conventional
transport category airplanes. In the MD-17, the thrust levers
(including the DLC switches) are
[[Page 26903]]
the primary means for controlling flight path for approach and landing.
Thrust is increased to reduce descent angle. To increase descent angle,
the MD-17 pilot will use small reductions in thrust to make small down
flight path adjustments, and will use the DLC thumb switch on the
thrust lever to make large down flight path corrections. In effect, the
MD-17 pilot uses the throttles in a similar manner to the way a
helicopter pilot uses the collective pitch lever. In contrast, the
pilot of a conventional transport category airplane primarily uses the
pitch control device for flight path control. For airspeed control, the
MD-17 pilot uses pitch, while the pilot of a conventional transport
category airplane primarily uses thrust.
Another significant characteristic of the power-augmented-lift MD-
17 design is that, while operating on the backside of the drag curve,
there is not much cross-coupling between pitch and thrust controls.
This means that changes in thrust result primarily in changes to the
flight path with very little effect on airspeed. Similarly, changes in
pitch affect primarily airspeed with little change to the flight path.
In combination with a full-authority three-axis fly-by-wire stability
and control augmentation system, this characteristic ensures accurate
airspeed control during manipulation of the thrust levers to control
the flight path descent angle. On a conventional transport category
airplane, manipulation of the pitch control to change the flight path
will result in unwanted airspeed excursions. For example, a one degree
change in flight path takes four seconds in a conventional transport
category airplane and is accompanied by a seven knot speed change,
while the same change in flight path for a powered-lift airplane takes
one second and does not result in a speed change.
Analysis of C-17 flight test and piloted simulator data support a
conclusion that airspeed can be controlled to a much higher degree of
precision during an approach with this airplane than with a
conventional transport category airplane. The analysis shows that the
standard deviation in speed due to maneuvering varied from 1 to 1.3
knots, while the speed excursions due to horizontal gusts ranged from
1.6 to 5.3 knots for light to severe turbulence levels. (The 5.3 knot
deviation corresponded with severe turbulence, including a 30-knot
crosswind and 33-knot headwind at a height of 50 feet above the
runway.) The standard deviation for the flight test approaches for
reported crosswinds of 13 to 31 knots, including both steep and normal
path approaches, was about 3.5 knots.
The unique MD-17 design features and operating characteristics
discussed above support a reevaluation of the minimum operating speed
for the approach and landing phase of flight. These design features and
operating characteristics provide the capability for rapid increases in
lift from thrust in the approach and landing configurations. Unlike
conventional transport category airplanes, there is no need to reduce
thrust to idle at any point in the approach or landing (until after
touchdown) for controlling either the flight path or rate of sink at
touchdown. Also, airspeed can be controlled very accurately even when
flight path changes are being made. Since large thrust decreases will
not be necessary nor will thrust be reduced to idle during the
approach, and rapid lift increases are available through the use of the
thrust levers, the FAA considers the use of one-engine-inoperative
power-on stall speeds in determining the reference landing speed,
VREF, for the MD-17 to provide equivalent safety to
conventional transport category airplanes. In addition, due to the
capability for more accurate airspeed control during the approach, the
FAA considers it appropriate to reduce the multiplying factor applied
to the reference stall speed in determining VREF. For the
MD-17, VREF may not be less than 1.20 times the one-engine-
inoperative power-on stall speed.
However, until more operational experience is gained with power-
augmented-lift airplanes, the FAA will not allow an applicant to
establish operating speeds for transport category airplanes lower than
the power-off stall speed. To provide some margin between the operating
speeds and the power-off stall speed, the MD-17's minimum operating
speeds must provide at least a 3 percent speed margin above the power-
off stall speed.
In addition to the speed margin obtained by applying factors to the
one-engine-inoperative power-on stall speeds, other constraints on the
minimum operating speeds must be considered due to the unique
characteristics of power-augmented-lift airplanes. For conventional
transport category airplanes, providing an airspeed margin between the
operating speed and the stall speed provides an adequate angle-of-
attack margin to stall. For a power-augmented-lift airplane like the
MD-17, however, separate airspeed, angle-of-attack, and thrust margins
must be considered. Maneuvering capability may also be more of a
concern on a power-augmented-lift airplane because of the difference in
thrust effects for a maneuver at a constant airspeed compared to a
slowdown maneuver.
Thrust Margin
On the MD-17, variations in thrust at a constant airspeed result in
variations in the stall speed margin. While this characteristic
provides the capability to increase lift (and hence stall speed margin)
simply by increasing thrust, there is also a potential for reductions
in stall speed margin following a thrust reduction. On a conventional
transport category airplane, where thrust is used primarily to control
airspeed, thrust reductions to idle can and do occur. On the MD-17,
thrust is used to control flight path rather than airspeed. The DLC
feature removes the need for large thrust reductions, and loss of stall
margin due to transient thrust reductions can be recovered quickly.
Additionally, because VREF is based on the one-engine-
inoperative power-on stall speed, additional margin is present in the
normal all-engines-operating condition. For the MD-17, the proposed
VREF would result in a speed approximately 1.27 times the
power-on stall speed with all-engines-operating at the thrust required
to maintain the reference approach flight path angle. At maximum
thrust, the proposed VREF would be 1.30 times greater than
the resulting power-on stall speed.
Another type of thrust variation would be a steady-state thrust
reduction that may, for example, be caused by a steady or increasing
tailwind, or a decreasing headwind. In this type of situation,
attempting to maintain a steady approach path with respect to the
ground would result in a steeper descent path angle, which would most
likely be attained by a lower thrust setting rather than through use of
the DLC. For an approach at the limiting tailwind condition, the
steeper approach flight path angle relative to the air mass reduces the
MD-17 airspeed margin to stall by less than one knot for normal and
steep approaches.
Based on the information presented above, an additional airspeed
margin to allow for thrust variation is not considered necessary. The
thrust or power on the operating engines used in the stall speed
determination for VREF should be the power or thrust used to
maintain the steady-state reference flight path angle at
VREF. For the MD-17, the reference flight path angle is
defined as -3 degrees for a normal approach, and the shallower of -5
degrees or the flight path angle associated with a descent rate of 1000
feet per minute for a steep approach.
[[Page 26904]]
Maneuvering Capability
During a banked turn, a portion of the lift generated by the wings
provides a force to help turn the airplane. To remain at the same
altitude, the airplane must produce additional lift. Therefore, banking
the airplane (at a constant speed and altitude) reduces the stall
margin, which is the difference between the lift required for the
maneuver and the maximum lift capability of the wing. As the bank angle
increases, the stall margin is reduced proportionately. Ignoring Mach
effects, this bank angle effect on the stall margin can be determined
analytically for conventional airplanes, and the multiplying factors
applied to the stall speed to determine the minimum operating speeds
are intended to ensure that an adequate stall margin is maintained.
For the MD-17, however, the effect of power-augmented-lift on stall
speeds differs between a slowdown maneuver (i.e., a wings level
deceleration) and a banked turning maneuver at a constant airspeed. The
speed reduction during a slowdown maneuver results in a larger
contribution of lift from thrust than is provided in a constant speed
maneuver. Therefore, for a power-augmented-lift airplane like the MD-
17, the stall CL would be lower in a constant speed turning
maneuver than in a slowdown maneuver. To ensure an equivalent level of
safety, the MD-17 minimum operating speeds should provide a maneuver
margin equivalent to conventional transport category airplanes.
The existing part 25 regulations do not prescribe specific
maneuvering margin requirements. However, as part of the proposed 1-g
stall amendment to part 25, maneuvering margin requirements are
proposed in Notice 95-17 (61 FR 1260, January 18, 1996). These proposed
maneuvering margin requirements represent the minimum maneuvering
margin to stall warning (or other characteristic that might interfere
with normal maneuvering) expected for the current fleet of transport
category airplanes. To provide equivalent maneuvering capability within
the operational flight envelope, the MD-17 must comply with maneuvering
margin requirements equivalent to those proposed in Notice 95-17,
except that the thrust used for the maneuvering capability at
VREF may be adjusted as necessary during the maneuver to
maintain the reference approach flight path angle. This change is
considered appropriate for the backside control technique that will be
used on the MD-17, where thrust, rather than pitch, is used as the
primary parameter to control flight path.
Angle-of-attack Margin
Another characteristic of power-augmented-lift airplanes like the
MD-17 is that the stall angle-of-attack during a slowdown maneuver can
be higher than the stall angle-of-attack achieved at higher speeds.
Again, this characteristic results from the variation of the effect of
power-on lift as speed varies. At higher airspeeds, the contribution of
power-augmented-lift can be less than at lower airspeeds. From an
operational standpoint, this characteristic can be critical during the
approach to landing phase of flight, where a sharp-edged vertical gust
could induce a large change in the angle-of-attack at approach speed.
For a conventional transport category airplane, where the angle-of-
attack margin is generally directly related to airspeed, vertical gust
margins are assured by the speed multiples applied to stall speeds when
determining the minimum allowable operating speeds. For power-
augmented-lift airplanes, this may not be true; therefore, the vertical
gust margin must be evaluated independently.
For conventional transport category airplanes, it has been
determined that approximately 20 knots of vertical gust margin is
provided at the minimum landing approach speed. (Reference: Report No.
FAA-RD-76-100, ``Progress Toward Development of Civil Airworthiness
Criteria for Powered-Lift Aircraft,'' May 1976, a copy of which is
included in the official docket for these special conditions.) To
provide equivalent safety, a vertical gust margin of 20 knots will be
included as a constraint on VREF for the MD-17 with all
engines operating. To ensure safety in the event of an engine failure,
the vertical gust margin in the one-engine-inoperative condition must
also be considered. Considering the short time period for operation in
this failure condition, the FAA has concluded that a vertical gust
margin of 15 knots will be required.
Proposed Special Condition No. 1 for MD-17 stall speeds and minimum
operating speeds takes into account power-augmented-lift effects for
configurations with flaps extended. Additionally, the FAA has
determined that the MD-17 stall speeds will be based on 1-g stall
criteria consistent with those proposed in Notice 95-17.
Systems
2. Head Up Display (HUD) Used as Primary Flight Display (PFD)
The MD-17 flight deck is equipped with two monochrome head up
displays (HUD), one at each pilot station. They are centrally located
in front of each pilot, above the glareshield at the pilot's eye level,
and between the pilot and the forward window. The MD-17 dual HUD
functions as the Primary Flight Display (PFD) for all regimes of normal
and abnormal operation and performs the functions of certain primary
flight instruments required for transport category airplanes by
Sec. 25.1303. The information is electronically projected on a
transparent surface with monochrome strokes. It may be used as the only
visible display, without any alternative flight instrument indications
displayed at the pilot station.
Until recently, HUD certification did not require a special
condition because conventional, certified primary flight instruments
were also provided at each pilot station and were always visible. The
MD-17 dual-HUD installation has the novel and unique feature of being
used when it is the only visible display of primary flight information,
which is not fully addressed by the current regulations. Therefore,
special conditions are proposed for the MD-17 dual HUD installation in
the following areas.
Arrangement and Visibility
Section 25.1321(b) states that the ``flight instruments required by
Sec. 25.1303 must be grouped on the instrument panel. * * *'' Because
of the MD-17 HUD location and its function as the only visible display
of primary flight information, Sec. 25.1303 does not adequately address
the MD-17 HUD's novel and unique features.
As described above, the HUD is not in the same visual field as the
instrument displays on the instrument panel. The electronically
displayed information is projected on a transparent surface and focused
at a distance (i.e., optical infinity). Unlike instrument scanning
between displays on the instrument panel, when scanning between the HUD
and the instrument panel the pilot's eyes must substantially change
viewing angle (about 15 degrees), light adaptation, and focus (from
infinity to 2 feet). Furthermore, information displayed on the
instrument panel cannot as easily be viewed in the pilot's peripheral
vision while simultaneously viewing the HUD, when compared to viewing
the suite of conventional flight instruments.
[[Page 26905]]
Therefore, in addition to compliance with Sec. 25.1321(b), a
special condition is proposed to require that the HUD provide all
information necessary for rapid pilot evaluation of the airplane's
flight state and position, during all phases of flight, for manual
control of the airplane, and for pilot monitoring of the performance of
the automatic flight control system. The HUD must provide equivalent
situational awareness of critical information that is normally
displayed near but not on the conventional PFD.
Pilot Compartment View and HUD Optical Characteristics
Section 25.1321(a) requires that ``[e]ach flight, navigation, and
powerplant instrument for use by any pilot must be plainly visible to
him from his station with the minimum practicable deviation from his
normal position and line of vision when he is looking forward along the
flight path.'' When the pilot is viewing conventional flight
instruments, the variations of pilot seating positions are not
significant in the pilot's ability to view the flight instruments.
However, with the HUD, the optical characteristics require that the
pilot's eyes be located within a very small volume to view all of the
required information, which is not adequately addressed by
Sec. 25.1321(a). There is much less tolerance for changes in eye
position and viewing angles when viewing the HUD. Hence, the proposed
special condition ensures that primary flight information remains
visible to the pilot without inadvertent lapses. In addition to
compliance with Sec. 25.1321(a), the proposed special condition ensures
that the HUD information is fully visible from the cockpit design eye
position, at which the required angular dimensions of the external
field of view, visibility of other cockpit instruments, and access to
cockpit controls are simultaneously realized. Furthermore, the special
condition ensures that pilot viewing of the HUD does not unduly
restrict pilot head movement, cause unacceptable fatigue or discomfort,
or interfere with other required pilot duties.
Also, unlike conventional flight displays, the HUD displays certain
flight information symbols conformally (i.e., graphically with angular
position and movement corresponding to the external view and in the
same angular scale). Mispositioning of conformal symbolic information
can be more hazardous than mispositioning the same information on
conventional displays. There is no specific rule that addresses the use
of conformal symbolic information as primary flight information.
Therefore, the proposed special condition does not permit the display
of electronic or optical misalignment of conformal symbology that would
be hazardously misleading.
Compatibility With Other Cockpit Displays
The existing regulations did not anticipate and do not address the
monochrome limitations associated with the MD-17 HUD. The HUD
electronically displays information with monochrome strokes, while on
conventional displays color is used to highlight and distinguish
different types of information. On color displays, the warning and
caution indications follow the same color scheme, red and amber,
respectively, as described in Sec. 25.1322 for warning, caution, and
advisory lights. This use of red and amber is consistent across the
cockpit and serves to give unmistakable meaning to the indications. The
MD-17 HUD must have an equivalent means to unmistakably highlight and
distinguish the same information.
The monochrome HUD must also have certain display design features
to make other essential flight information conspicuous, distinct, and
meaningful to compensate for the lack of multiple colors. For example,
the conventional primary attitude indication distinguishes angles on
the pitch scale above the horizon (sky) and angles below the horizon
(earth) with different colors, such as blue and brown, respectively. To
perform its intended function as the primary attitude indicator, and to
ensure satisfactory pilot recognition of unusual attitudes, the HUD
must provide clear visual distinction between positive and negative
pitch angles by means other than color.
In summary, the display format of the HUD can differ from the
format of other cockpit displays of the same information due to
differences in their capabilities and limitations. These differences
must be regulated to ensure that one format is not so unlike the other
that the pilot can misinterpret the information hazardously, or that
excessive time and attention is required for the pilot to interpret the
information. During critical high workload or emergency conditions, the
pilot may need to quickly make a transition from the HUD to other
flight instruments to continue safe flight. The existing rules do not
adequately address the compatibility of different display formats in
the MD-17 cockpit. This special condition is required to avoid
potentially hazardous workload and pilot confusion due to display
incompatibility.
To address the above identified inadequacies in current regulations
as related to the acceptability of the HUD as the primary source of
flight information, Special Condition No. 2 is proposed as an
appropriate set of requirements.
Additional Recommendations or Supporting Data
In addition to the special condition for the HUD system, there are
other regulations and advisory material that, although adequate,
warrant special attention due to the unique features of the MD-17 HUD
installation. The following discussion of applicable regulations is
provided for information in the context of this special condition.
Regulations
Section 25.771(e): ``Vibration and noise
characteristics of cockpit equipment may not interfere with safe
operation of the airplane.'' Attention should be paid to the visual
effects resulting from vibration of the cockpit and the optical
components of the HUD, including vibration associated with engine
imbalance resulting from fan blade failure.
Section 25.773(a)(1): ``Each pilot compartment must be
arranged to give the pilots a sufficiently extensive, clear, and
undistorted view, to enable them to safely perform any maneuvers
within the operating limitations of the airplane, including taxiing,
takeoff, approach, and landing.'' Special attention should be paid
to this requirement because of the unique location of the HUD
combiner, between the pilot's eyes and the forward windshield,
compared to conventional displays. The potential of each combiner
structure to obstruct the outside view of both pilots (on-side and
off-side) should be considered.
Section 25.773(a)(2): ``Each pilot compartment must be
free of glare and reflection that could interfere with the normal
duties of the minimum flight crew (established under Sec. 25.1523).
This must be shown in day and night flight tests under non-
precipitation conditions.'' Special attention should be paid to this
requirement because the unique HUD optical system and the location
of the combiner, between the pilot's eyes and the forward
windshield, can be especially susceptible to and be the cause of a
variety of glare and reflections in the cockpit.
Section 25.785(k): ``Each projecting object that would
injure persons seated or moving about the airplane in normal flight
must be padded.'' Typical installations of HUD's include components
that project into the space near the pilot's head. Attention should
be paid to head contact with these components during all expected
operations and pilot activities, especially during turbulence.
Section 25.1301(a): ``Each item of installed equipment
must be of a kind and design appropriate to its intended function.''
Previously, HUD's for transport category airplanes have been
certified with a fully
[[Page 26906]]
certificated set of primary flight instruments/displays visible on a
full-time basis; therefore, the HUD was not required to meet all of
the requirements for primary flight instruments. However, the MD-17
HUD's are a primary source of flight information and must comply
with those requirements, because alternate instrument flight
displays that comply are not in full-time use. Therefore,
consideration should be given to the functionality of the MD-17 HUD
under all foreseeable operating conditions. For example, looking
directly at the sun through the HUD combiner can be painful or
harmful to the pilot's eyes; therefore, an alternate display of
primary flight information, which complies with the applicable
regulatory requirements, must be available on demand. The MD-17 is
capable of displaying primary flight information on any of its four
multi-function displays (MFD's). To comply with Sec. 25.1321, the
two MFD's centered in front of each pilot must be available to
display instrument flight information on demand, and the other two
center displays must be able to simultaneously display other
essential information, such as navigation and engine indications.
Selectable display functionality needs special attention in
determining compliance with Sec. 25.1301 for the MD-17 suite of
displays, including HUD's and MFD's.
The installation of the HUD system must not interfere with or
restrict the use of other installed equipment such as emergency
oxygen masks, headsets, or microphones. HUD installations typically
result in the placement of protruding equipment (e.g., projector,
combiner) in the vicinity of the pilot's head and thereby provide
the potential for compromising the intended function of the
equipment identified above.
The HUD is capable of presenting a large amount of static and
dynamic symbology, numbers, and text that can appear cluttered,
difficult to interpret, and difficult to see through. Special
attention should be given to the potential effects of display
clutter, such as interference between moving symbols, other symbols,
and alphanumeric information on display functionality, flightcrew
task performance, and workload (Sec. 25.1523; Appendix D).
``Declutter'' modes can selectively remove certain data from the
display, so special attention should be given to ensuring that
essential data cannot be removed, when needed to continue safe
flight and landing.
Section 25.1381a(2)(ii): ``Instrument lights must be
installed so that no objectionable reflections are visible to the
pilot.'' Attention should be paid both to reflections from other
sources on the HUD and those from the HUD on to windows and other
displays.
Advisory Material
Advisory Circular (AC) 25-11, ``Transport Category Airplane
Electronic Display Systems,'' provides guidance and policy information
regarding means to demonstrate the acceptability of electronic
displays, including HUD's. All portions of AC 25-11 are applicable to
demonstrate compliance for the special conditions, except for the color
unique criteria of paragraph 5. However, note that the fundamental
principles specified in subparagraph 5b, Color Perception vs. Workload,
do apply and should be followed with non-color means such as size,
shape, and location. Although the HUD does not have color, criteria for
evaluation of clutter, workload, and display perception, considering
distinctive symbology features such as size, shape, and location, are
applicable. Also note that, for HUD's, excessive clutter affects not
only the workload and readability of the presentation, but also the
pilot's ability to see the outside view and visually detect operational
hazards. Also, in spite of its title, the luminance criteria of
subparagraph 6b, Chromaticity and Luminance, applies to evaluation of
the HUD display luminance. Unique HUD requirements for HUD brightness
capability and control are specified in Special Condition No. 2(b)(2).
3. Protection From Unwanted Effects of High Intensity Radiated Fields
(HIRF)
The MD-17 uses electrical and electronic systems that perform
critical and essential functions. These systems include electronic
displays, electronic engine controls, fly-by-wire flight controls, and
others. There is no specific regulation that addresses protection
requirements for these systems from HIRF. Increased power levels from
ground based radio transmitters and the growing use of sensitive
electrical and electronic systems to command and control airplanes have
made it necessary to provide adequate protection.
Changes in technology have given rise to advanced electrical and
electronic airplane systems, use of composite materials in airplane
structures, and higher energy levels from radio, television, and radar
transmitters. The combined effect of these developments has been an
increased susceptibility of electrical and electronic systems to
electromagnetic fields.
Many advanced digital systems are prone to upsets and/or damage at
energy levels lower than analog systems. Digital systems also allow the
location of more complex functions in fewer components. These functions
were previously performed manually, electromechanically, or
hydraulically. The implementation of such advanced systems has found
rapid acceptance since they lower cost, crew workload, and maintenance
requirements, while airplane performance and fuel efficiency are
enhanced.
Propelled by the need to attain higher efficiency, industry has
also proceeded to adopt composite materials for use in airplane
structures, thus reducing or replacing the use of aluminum. Due to
their low conductivity properties, composite materials afford poor
shielding effectiveness, further exposing electrical and electronic
systems to the electromagnetic environment.
At this time, the FAA and other airworthiness authorities are
unable to precisely define or control the HIRF energy level to which
the airplane will be exposed in service. Therefore, to ensure that a
level of safety is achieved equivalent to that intended by the current
regulations, Special Condition No. 3 is proposed. This special
condition would require that new electrical and electronic systems that
perform critical functions be designed and installed to preclude
component damage and interruption of function due to both the direct
and indirect effects of HIRF.
Airframe
4. Interaction of Systems and Structures
The MD-17 airplane utilizes a full-time electronic flight control
system (EFCS). Pilot control commands are sent to flight control
computers which condition the input signals, combine them with other
sensor data indicating airplane configuration and flight condition, and
apply servo position commands to the actuation systems of the control
surfaces. In this way, the EFCS affects control surface actuation and
therefore the airplane flight loads. Failures that occur in the EFCS
may further affect flight loads, both at the time of the event and
thereafter.
The current part 25 airworthiness standards were intended to
account for control laws for which control surface deflection is
proportional to control device deflection. They do not address any
nonlinearities or other effects on control surface actuation that may
be caused by the EFCS, whether fully operative or in a failure mode.
Since the EFCS may affect flight loads, and therefore the structural
capability of the airplane, specific regulations are needed to address
these effects. Thus, Special Condition 4 is proposed.
If a failure occurs within the EFCS, the airplane may still be
capable of operating within a reduced structural envelope. That is, the
airplane may be able to meet the strength and flutter requirements of
part 25, but at reduced factors of safety or airspeed, as applicable.
This reduced structural envelope is considered acceptable provided that
it is based on failure probabilities within the EFCS. Special Condition
4 provides specific structural load and aeroelastic stability
[[Page 26907]]
requirements with reduced factors of safety and/or airspeeds based on
the probability of failure. These requirements ensure that the airplane
structural design safety margins will be dependent on system
reliability. The requirements proposed in Special Condition 4 also
ensure that any influence of the EFCS on airplane flight loads will be
accounted for when the system is fully operative.
5. Design Maneuvering Requirements for Fly-by-Wire
The MD-17 airplane utilizes a full-time electronic flight control
system (EFCS). Pilot control commands are sent to flight control
computers, which condition the input signals, combine them with other
sensor data indicating airplane configuration and flight condition, and
apply servo position commands to the actuation systems of the control
surfaces. In this way, the EFCS affects control surface actuation and
therefore the airplane flight loads.
The current part 25 airworthiness standards were intended to
account for control laws for which control surface deflection is
proportional to control device deflection. They do not address
nonlinearities or other effects on control surface actuation that may
be caused by the EFCS. Since the EFCS may affect flight loads, and
therefore the structural capability of the airplane, specific
regulations are needed to address these effects. Thus, Special
Condition 5 is proposed.
Special Condition 5 differs from current requirements in that it
requires that certain maneuvers be performed by actuation of the
cockpit control device as opposed to the corresponding control surface.
In addition, the special condition requires consideration of loads
induced by the EFCS itself. These requirements ensure that any
influence of the EFCS on airplane flight loads will be accounted for.
6. Limit Engine Torque Loads for Sudden Engine Stoppage
McDonnell Douglas proposes to treat the rare sudden engine stoppage
condition resulting from structural failure as an ultimate load
condition. Section 25.361(b)(1) specifically defines the seizure torque
load, resulting from structural failure, as a limit load condition.
The limit engine torque load imposed by sudden engine stoppage due
to malfunction or structural failure (such as compressor jamming) has
been a specific requirement for transport category airplanes since
1957. The size, configuration, and failure modes of jet engines has
changed considerably from those envisioned by Sec. 25.361(b) when the
engine seizure requirement was first adopted. Engines are much larger
and are now designed with large bypass fans capable of producing much
larger torque loads if they become jammed. It is evident from service
history that the frequency of occurrence of the most severe sudden
engine stoppage events, resulting from structural failures, are rare.
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 a special
condition to establish appropriate design standards. The latest
generation of jet engines are capable of producing engine seizure
torque loads that are significantly higher than previous generations of
engines.
The FAA is developing a new regulation and a new AC that will
provide more comprehensive criteria for treating engine torque loads
resulting from sudden engine stoppage. In the meantime, a special
condition is needed to establish appropriate criteria for the MD-17
type design.
In order to maintain the level of safety envisioned by
Sec. 25.361(b), more comprehensive criteria are needed for the new
generation of high-bypass engines. The proposed special condition would
distinguish between the more common seizure events and those rare
seizure events resulting from structural failures. For these more rare
but severe seizure events, the proposed criteria would allow
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 an additional safety
factor.
To provide appropriate structural design criteria for the engine
torque on the MD-17, Special Condition No. 6 is proposed.
Flight Characteristics
7. Flight Characteristics Compliance via Handling Qualities Rating
Method
The MD-17 will have an Electronic Flight Control System (EFCS).
This system will provide an electronic interface between the pilot's
flight controls and the flight control surfaces (for both normal and
failure states), generating the actual surface commands that provide
for stability augmentation and control about all three airplane axes.
Because EFCS technology has outpaced existing regulations (written
essentially for unaugmented airplanes, with provision for limited ON/
OFF augmentation), a suitable special condition is needed to aid in the
certification of flight characteristics.
In addition, service history and certification experience have
shown that EFCS-type airplanes and others may be susceptible to
airplane-pilot coupling (A-PC) tendencies. Pilot induced oscillations
can be considered a subset of A-PC problems. An example of these
problems are control systems that are rate or position limited during
some pilot commands in which the pilot has no feedback through the
controller.
The proposed special condition provides a means by which flight
characteristics (``satisfactory,'' ``safe flight and landing,'' etc.)
can be evaluated and compliance found. The Handling Qualities Rating
System (HQRS) was developed for airplanes with control systems having
similar functions and is employed to aid in the evaluation of the
following:
For all EFCS/airplane failure states not shown to be
extremely improbable, and where the envelope (task) and atmospheric
disturbance probabilities are each 1.
For all combinations of failures, atmospheric
disturbance level, and flight envelope that yield flight conditions
expected to occur more frequently than extremely improbable.
For any other flight condition or characteristic where
part 25 proves to be inadequate for proper assessment of unique MD-
17 flight characteristics.
The HQRS provides a systematic approach to handling qualities
assessment. It is not intended to dictate program size or need for a
fixed number of pilots to achieve multiple opinions. The airplane
design itself and success in defining critical failure combinations
from the many reviewed in systems safety assessments would dictate the
scope of any HQRS application.
Handling qualities terms, principles, and relationships familiar to
the aviation community have been used to formulate the HQRS. For
example, similarity has been established between the well-known Cooper-
Harper rating scale and the proposed FAA three-part rating system. This
approach is derived, in part, from work on flying qualities of highly
augmented/relaxed static stability airplanes, namely regulatory and
flight test guide requirements.
In addition, experience has shown that compliance with only the
qualitative, open-loop (pilot-out of-the-loop) requirements does not
guarantee that the required levels of flying qualities are achieved.
There must be an evaluation by certification pilots conducting high
gain (wide band width) closed-loop (pilot-in-the-loop) tasks, to
[[Page 26908]]
ensure that the airplane demonstrates the flying qualities required by
Secs. 25.143(a) and (b) and to minimize the hazards associated with
encountering adverse A-PC tendencies in service.
For the most part, these tasks must be performed in actual flight.
For conditions that are considered too dangerous to attempt in actual
flight (i.e., certain flight conditions outside of the operational
flight envelope, flight in severe atmospheric disturbances, flight with
certain failure states, etc.), the closed loop evaluation tasks may be
performed on a validated high fidelity simulator.
Special Condition No. 7 is proposed for the MD-17 to aid in the
certification of flight characteristics. An acceptable means of
compliance with this special condition is provided in AC 25-7A,
``Flight Test Guide for the Certification of Transport Category
Airplanes.''
8. Static Longitudinal Stability
Like other airplanes with similar highly augmented electronic
flight control systems, the MD-17 does not literally comply with the
requirements prescribed by Sec. 25.173 for static longitudinal
stability. In one control mode of the electronic flight control system,
no control force is needed to maintain a speed change from the trimmed
condition. Although this operating system mode provides quick, accurate
pitch response with minimal pilot effort, it does not comply with the
literal requirements for static longitudinal stability.
Static longitudinal stability has been required in accordance with
part 25 for the following reasons:
Provides additional speed change cues to the pilot
through control force changes.
Ensures that short periods of unattended operation do
not result in any significant changes in attitude, airspeed, or load
factor.
Provides predictable pitch response.
Provides acceptable level of pilot attention (workload)
to attain and maintain trim speed and altitude.
Provides gust stability.
In order to achieve an equivalent level of safety with part 25, the
MD-17 should meet the intent of these principles, even though it may
not comply with the literal terms of Sec. 25.173. Special Condition No.
8 is proposed to ensure that the MD-17 has suitable static longitudinal
stability in any condition normally encountered in service. The HQRS
prescribed by Special Condition No. 7 may be used to make this
assessment.
9. Static Lateral-Directional Stability
Because of the MD-17 roll axis design feature in which the
commanded roll rate is proportional to roll stick position, aileron
control movements and forces do not comply with Sec. 25.177 as they are
not proportional to angle of sideslip. This feature is active during
all flight phases and conditions, except when the flap/slat handle is
at or greater than the \1/2\ detent setting, or during a rudder pedal
input.
Dihedral effect (as indicated by aileron forces proportional to the
angle of sideslip) has been required in accordance with Sec. 25.177 for
the following reasons:
In the event that primary lateral control is lost, roll
can be produced by use of the rudder.
In an airplane with positive dihedral effect, the bank
angle and the lateral control forces required to hold heading
provide positive indication of an inadvertent sideslip.
It can have a beneficial effect on spiral stability.
In the event of an engine failure, the roll due to the
asymmetric yawing moment contributes to the ease of identifying the
failed engine.
In order to achieve an equivalent level of safety with part 25, the
MD-17 should meet the intent of these principles even though it may not
comply with the literal terms of Sec. 25.177.
In lieu of showing compliance with Sec. 25.177, Special Condition
No. 9 is proposed to ensure that the MD-17 has suitable static lateral-
directional stability in any condition normally encountered in service.
The HQRS prescribed by Special Condition No. 7 may be used to make this
assessment.
10. Control Surface Awareness
With an electronic flight control system and no direct coupling
from cockpit controller to control surface, the pilot may not be aware
of the actual surface position utilized to fulfill the requested
demand. Some unusual flight condition, arising from atmospheric
conditions and/or airplane or engine failures, may result in full, or
near full, surface deflection. Unless the flightcrew is made aware of
excessive deflection or impending control surface limiting, piloted or
auto-flight system control of the airplane might be inadvertently
continued in such a manner as to cause airplane loss of control or
other unsafe stability or performance characteristics.
In airplanes with electronic flight control systems, there may not
always be a direct correlation between pilot control position and the
associated airplane control surface position. Under certain
circumstances, a commanded maneuver that may not involve a large
control input may nevertheless require a large control surface
movement, possibly encroaching on a control surface or actuation system
limit without the flightcrew's knowledge. This situation can arise in
both manually piloted and autopilot flight, and may be further
exacerbated on airplanes where the pilot controls are not back-driven
during autopilot system operation.
As a result of these concerns, a special condition is proposed for
the MD-17. Special Condition No. 10 proposes a requirement that
suitable flight control position annunciation be provided to the
flightcrew when a flight condition exists in which near full surface
authority (not crew-commanded) is being utilized. Suitability of such a
display or alerting must take into account that some pilot-demanded
maneuvers are necessarily associated with intended full performance,
which may saturate the surface. Therefore, simple alerting systems,
which would function in both intended or unexpected control-limiting
situations, must be properly balanced between needed crew awareness and
nuisance factors. A monitoring system that compares airplane motion,
surface deflection, and pilot demand could be useful for eliminating
nuisance alerting.
Approach and Landing Limitations
11. Steep Approach Air Distance
The MD-17 has a number of design features to support steep approach
flight path capability with precision landing. McDonnell Douglas
proposes to certify MD-17 landing performance for both conventional 3-
degree approach glideslope operation and steep approach operation.
Novel and unique features on the MD-17 provide for increased
touchdown dispersion accuracy during steep approach operations relative
to conventional transport category airplanes. McDonnell Douglas has
proposed an alternative method for defining the airborne portion of the
landing distance in lieu of the demonstrated distance from a 50-foot
height to touchdown. A special condition is proposed to redefine the
air distance portion of the MD-17 landing distance for steep approach
operations conducted under a proposed Special Federal Aviation
Regulation (SFAR), ``Requirements for operational approval of special
approaches to short field landings for the McDonnell Douglas Model MD-
17 power-augmented-lift airplane,'' currently being developed by the
FAA.
Steep approach operations are intended to minimize the air run to
help achieve short field performance. Steep
[[Page 26909]]
approach for the MD-17 is defined as an approach flight path angle not
to exceed 5 degrees, with an approach rate of descent not to exceed
1,000 feet per minute. For the landing reference speeds used by the MD-
17, almost all operations are limited by the 1,000 feet per minute
constraint, which yields approach flight path angles predominantly in
the range from 4 to 4.8 degrees.
Several design features on the MD-17 are intended to enable the
airplane to safely fly steep approaches. First, the landing gear is
designed to withstand touchdown rates of descent of up to 12.5 feet per
second for weights up to 435,800 pounds and 11 feet per second for
weights up to the maximum MD-17 landing weight of 491,900 pounds.
Second, the high lift system with externally blown flaps allows
operation at relatively low landing reference speeds which, when
combined with the MD-17 lift/drag characteristics, allows this airplane
to be flown using a backside control technique. Third, a spoiler
function linking spoilers and throttle movement provides much more
precise throttle control. Fourth, the MD-17 is equipped with a HUD,
which displays the airspeed and the flight path vector, and a pilot-
selectable flight path marker to indicate the desired flight path. The
HUD assists the pilot in precisely controlling the airplane flight path
to an aim point on the runway. With no pitch flare needed, the aim
point is very close to the actual touchdown point. Considered together,
these MD-17 features allow pilots to fly steep approaches and accurate
touchdowns near the aim point, while maintaining control over speed and
the rate of descent at touchdown.
The backside control technique mentioned above uses thrust changes
to primarily affect flight path angle, and pitch changes to primarily
affect airspeed. As with all airplanes, there is some control coupling
such that any control input will affect both flight path angle and
airspeed, but the coupling is minimized for the low speed backside
operation used by the MD-17. Reduced control coupling leads to greater
precision in airspeed and flight path control. The backside control
technique allows throttle inputs to be used to control vertical speed
all the way to touchdown instead of the ``pitch flare'' maneuver used
on other airplanes.
The throttle-spoiler interconnect feature of the MD-17 design
allows spoiler motion to simulate the effect of immediate engine
response to throttle movement. The spoilers are nominally biased in the
up direction during steady-state operation. When the throttles are
moved, the spoilers move in the direction necessary to provide
essentially the same airplane response as an immediate thrust change.
As the engine responds, the spoilers, over time, return to their
original (biased) positions. This feature eliminates the lag often
associated with thrust control.
Over 175 steep approach landings were performed during C-17 testing
to demonstrate the precision landing characteristics. All of these runs
were made using an operational technique performed by pilots with only
three practice runs to gain familiarity with the technique. These
approaches were conducted to establish that no exceptional piloting
skill or training was required to achieve the tested performance
levels. During the demonstrations, only a limited portion of the flight
manual allowable wind and temperature conditions were accounted for.
The purpose of the testing was to demonstrate that the precision
approach accuracy could yield touchdowns with a 2 standard
deviation () band of less than 500 feet relative to the mean
touchdown point, while also maintaining an acceptable rate of descent
at touchdown.
The FAA notes that McDonnell Douglas proposes two distinct types of
landing operations for the MD-17: (1) conventional landings that will
be conducted in accordance with existing part 25 and 121 regulations;
and (2) special approaches to short field landings that will be
conducted in accordance with a proposed SFAR (to be published at a
later date) and associated special conditions. The proposed SFAR would
address additional equipment, training, and operating requirements
associated with conducting special approaches to short field landings.
McDonnell Douglas intends to provide steep approach capability
(allowing operators to seek steep approach approval) for both types of
landing operations.
For conventional landings, the steep approach air distance would be
determined by using the existing applicable type certification and
operating requirements. This proposed special condition for steep
approach air distance would only apply to special approaches to short
field landings conducted in accordance with the proposed SFAR and
Special Condition No. 12, ``Landing Distances for Special Approaches to
Short Field Landings.'' It addresses only the determination of landing
distance to be used in conjunction with those operations and does not
imply approval to conduct steep approach operations.
For MD-17 steep approach operations conducted under the proposed
SFAR, Special Condition No. 11 is proposed in conjunction with proposed
Special Condition No. 12, in lieu of Sec. 25.125(a).
12. Landing Distances for Special Approaches to Short Field Landings
As noted in the discussion of Special Condition No. 11, McDonnell
Douglas proposes two distinct types of landing operations for the MD-
17: (1) conventional landings that will be conducted in accordance with
existing part 25 and 121 regulations, and (2) special approaches to
short field landings that will be conducted in accordance with a
proposed SFAR and associated special conditions.
The operational landing distance margin provided by part 121 takes
into account steady-state variables that are not included in the part
25 landing distances, differences in operational procedures and
techniques from those used in determining the part 25 landing
distances, non steady-state variables, and differences in the
conditions forecast at dispatch and those existing at the time of
landing. Examples of each of these categories include:
----------------------------------------------------------------------------------------------------------------
Non steady-state Operations vs. Flight Actual vs. Forecast
Steady-state variables variables Test conditions
----------------------------------------------------------------------------------------------------------------
Runway slope......................... Wind gusts/turbulence.. Flare technique........ Runway or direction
(affecting slope).
Temperature.......................... Flight path deviations. Time to activate Airplane weight.
deceleration devices.
Runway surface condition (dry, wet, ....................... Flight path angle...... Approach speed.
icy, texture).
Brake/tire condition................. ....................... Rate of descent at Environmental
touchdown. conditions (e.g.,
temperature, wind,
pressure altitude).
Speed additives...................... ....................... Approach/touchdown Engine failure.
speed.
[[Page 26910]]
Crosswinds........................... ....................... Height at
thresholdSpeed control
.
----------------------------------------------------------------------------------------------------------------
Note: This is not intended to be an exhaustive list of variables to be considered.
In order to allow the part 121 operational landing distance margins
to be reduced as proposed in the SFAR for special approaches to short
field landings, additional type certification requirements are needed.
In addition to what is currently required by Sec. 25.125, the landing
distances to be used under the proposed SFAR would be required to
include the effects of runway slope and ambient temperature. Landing
distances on a wet runway would also have to be determined in a manner
acceptable to the FAA. In addition, during the flight testing to
determine the landing distances, the average touchdown rate of descent
and the approach flight path angle would be limited to no greater than
4 feet per second and -3 degrees, respectively.
The applicant would be required to establish operating procedures
for use in service that are consistent with those used to establish the
performance data and can be executed by crews of average skill. The
applicant would be required to include, as applicable, procedures
associated with speed additives for turbulence and gusts for approaches
with all engines operating and with an engine failure on final
approach, and the use of thrust reversers on all operative engines
during the landing rollout.
The operational landing distance margins applicable to the MD-17,
and additional operational considerations associated with the use of
these reduced margins (e.g., runway markings, meteorological
conditions, and flightcrew procedures and training), are covered in the
proposed SFAR.
Although this special condition will explicitly take into account
many of the variables currently accounted for by the part 121
operational landing distance margins, some operational landing distance
margin is still necessary to account for variables that remain. For
example, because Sec. 121.195(d) specifies the maximum takeoff weight
for the conditions forecast at the time of landing (including
environmental conditions such as temperature and pressure altitude,
airport conditions such as runway and direction, and airplane
conditions such as fuel burnoff and approach speed), potential
differences in the forecast and actual conditions should be taken into
account. Other operational issues that should be considered in the
operational landing distance margins include piloting technique and
time to activate deceleration means, unsteady winds and crosswinds, and
airspeed and flight path deviations. Therefore, the proposed SFAR will
still contain operational landing distance margins, although reduced
from those margins currently required by Secs. 121.195 and 121.197,
that would be applied to the landing distance determined in accordance
with this proposed special condition.
The proposed Special Condition No. 12 provides the additional
requirements noted above that the FAA considers necessary to allow
operational use of the landing distance margins prescribed in the
proposed SFAR. Note that the determination of landing distances in
accordance with this proposed special condition does not constitute
operational approval to use landing distance margins reduced from those
specified in part 121. Operational approval to use the reduced landing
distance margins must be obtained in accordance with the proposed SFAR.
13. Thrust for Landing Climb
Section 25.119(a) states that the airplane must achieve a 3.2
percent climb gradient after initiating a thrust increase from the
minimum flight idle position. The thrust allowed is that thrust
attained within eight seconds of engine spool-up time from the
initiation of thrust lever movement. Because of the power-augmented-
lift design, the MD-17 thrust required for a stabilized approach is
significantly above a conventional turbojet minimum flight idle
setting, and thrust would not be reduced to idle during the approach.
Section 25.119(a) was written to assure that the flightcrew would
have sufficient airplane performance to safely transition to a climb
during a go-around in the landing configuration. The rule assumes that
the approach power setting may be as low as the flight idle position.
The MD-17 power-augmented-lift design requires a significant approach
thrust level during the approach to maintain the approach flight path.
Unlike conventional transport category airplanes, thrust reductions
during the approach are not necessary to maintain or recover the flight
path. The MD-17 operational procedures will discourage use of thrust
reduction to make down flight path adjustments during approach. The
direct lift control (DLC) feature provides a down path angle control
for large flight path adjustments without throttle movement.
To improve the control response to throttle movement, the MD-17
uses a spoiler function where the spoilers are linked with the
throttles to simulate the effect of instantaneous engine response to
throttle movement. The throttle-spoiler function is a short-term
response; as the engine responds to throttle movement, the spoilers
return to their original positions. The approach is flown with a non-
zero spoiler bias to allow spoilers to react upward or downward in
response to throttle movement. This function provides instantaneous
response to control input and allows throttle movement to be minimized.
During the segment from 50 feet to touchdown, the MD-17 uses a
backside control technique that does not require either thrust to be
reduced to an idle power setting or the use of a pitch-up flare
maneuver prior to touchdown. With the backside control technique,
airplane pitch attitude is used to control airspeed, and thrust is used
to control flight path angle.
In lieu of compliance with Sec. 25.119(a), Special Condition No. 13
is proposed. The thrust for a stabilized approach, including an
appropriate margin for operational safety, would be used as a basis for
determining the thrust available for the landing climb requirement. In
the proposed special condition, the initial thrust level at the start
of the 8-second spool-up time would be the thrust for a stabilized
approach at a flight path angle 2 degrees steeper than the desired
flight path angle. This thrust level would account for thrust
variations during the approach and conservatively represent the initial
thrust level.
This proposed special condition would be applicable only when the
following design features are present:
At no time in the landing configuration should the
thrust be reduced to idle.
A backside control technique must be used such that a
thrust reduction is not used to reduce the rate of descent at
touchdown.
Procedures must be provided in the Airplane Flight
Manual to define the proper technique for flight path angle
adjustments during approach and landing.
The airplane must have DLC spoilers or other
aerodynamic means of making down path angle adjustments without
thrust reduction.
[[Page 26911]]
Throttle movement should activate a short-term
aerodynamic surface motion in order to provide a high level of
control feedback and to avoid excessive throttle adjustments.
The airplane and engine state (e.g., airplane weight
and engine bleed configuration) and operating conditions (e.g.,
pressure altitude and temperature) should be the most critical
combination relative to the thrust level used to show compliance
with this special condition.
Applicability
As discussed above, these special conditions are applicable to the
McDonnell Douglas Model MD-17 series airplanes. Should McDonnell
Douglas apply at a later date for a change to the type certificate to
include another model incorporating the same novel or unusual design
features, the special conditions would apply to that model as well
under the provisions of Sec. 21.101(a)(1).
Conclusion
This action affects only certain novel or unusual design features
on one model series of airplanes. It is not a rule of general
applicability and affects only the applicant who applied to the FAA for
approval to use 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 McDonnell Douglas Model MD-17 series airplanes:
1. Stall Speeds and Minimum Operating Speeds
(a) In addition to the general definitions, abbreviations, and
symbols provided in Secs. 1.1 and 1.2, this special condition relies on
the following additional definitions, abbreviations, and symbols:
`` Reference flight path angle means -3 degrees for a normal
approach, and the shallower of -5 degrees or the flight path angle
resulting from a 1000 feet per minute rate of descent for a steep
approach.''
``VSR means reference stall speed.''
``VSRPWR means power-on reference stall speed.''
``VSRO means reference stall speed in the landing
configuration.''
``VSROPWR means power-on reference stall speed in the
landing configuration.''
``VSR1 means reference stall speed in a specific
configuration.''
``VSR1PWR means power-on reference stall speed in a
specific configuration.''
``VREF means reference landing speed.''
``VFTO means final takeoff speed.''
``VSW means speed at which onset of natural or
artificial stall warning occurs.''
(b) In lieu of compliance with Sec. 25.103, the following applies:
(1) The reference stall speed, VSR, is a calibrated
airspeed as defined in paragraph (3) below. VSR is
determined with--
(i) Engines idling, or, if that resultant thrust causes an
appreciable decrease in stalling speed, not more than zero thrust at
the stall speed;
(ii) The airplane in other respects (such as flaps and landing
gear) in the condition existing in the test in which VSR is
being used;
(iii) The weight used when VSR is being used as a factor
to determine compliance with a required performance standard;
(iv) The center of gravity position that results in the highest
value of reference stall speed; and
(v) The airplane trimmed for straight flight at a speed selected by
the applicant, but not less than 1.13 VSR and not greater
than 1.30 VSR.
(2) Starting from the stabilized trim condition, apply elevator
control to decelerate the airplane so that the speed reduction does not
exceed one knot per second.
(3) The reference stall speed, VSR, may not be less than
a 1-g stall speed, which is a calibrated airspeed determined in the
stalling maneuver and expressed as:
[GRAPHIC] [TIFF OMITTED] TP18MY99.010
Where:
VCLMAX = Speed occurring when lift coefficient is first a
maximum; and
nZW = Flight path normal load factor (not greater than 1.0)
at VCLMAX.
(4) The power-on reference stall speed, VSRPWR, is a
calibrated airspeed as defined in paragraph (6) below.
VSRPWR is determined with--
(i) The critical engine inoperative and the power or thrust setting
on the remaining engines at the minimum power or thrust level
appropriate for the flight condition used to show compliance with a
required performance standard;
(ii) The airplane in other respects (such as flaps and landing
gear) in the condition existing in the test in which VSRPWR
is being used;
(iii) The weight used when VSRPWR is being used as a
factor to determine compliance with a required performance standard;
(iv) The center of gravity position that results in the highest
value of the power-on reference stall speed; and
(v) The airplane trimmed for straight flight at a speed selected by
the applicant, but not less than 1.18 VSRPWR and not greater
than 1.36 VSRPWR.
(5) Starting from the stabilized trim condition, apply elevator
control to decelerate the airplane so that the speed reduction does not
exceed one knot per second.
(6) The power-on reference stall speed, VSRPWR, may not
be less than a 1-g power-on stall speed, which is a calibrated airspeed
determined in the stalling maneuver and expressed as:
[GRAPHIC] [TIFF OMITTED] TP18MY99.011
Where:
VCLMAX = Speed occurring when lift coefficient is first a
maximum; and
nZW = Flight path normal load factor (not greater than 1.0)
at VCLMAX.
(c) In lieu of compliance with Sec. 25.107(b), the following
applies: V2MIN, in terms of calibrated airspeed, may not be
less than--
(1) 1.03 VSR;
(2) 1.18 VSRPWR, with the operative engines at the
minimum thrust or power existing at any point in the takeoff path; and
(3) 1.10 times VMC established under Sec. 25.149.
(d) In addition to compliance with Secs. 25.107(c)(1) and (c)(2),
the following also applies: A speed that provides the maneuvering
capability specified in paragraph (k) below.
(e) In addition to compliance with Secs. 25.107(a) through (f), the
following also applies: VFTO, in terms of calibrated
airspeed, must be selected by the applicant to provide at least the
gradient of climb required by paragraph (h) below, but may not be less
than--
(1) 1.18 VSR; and
(2) A speed that provides the maneuvering capability specified in
paragraph (k) below.
(f) In lieu of compliance with Sec. 25.111(a), the following
applies: The takeoff path extends from a standing start to a point in
the takeoff at which the airplane is 1,500 feet above the takeoff
surface, or at which the transition from the takeoff to the en route
configuration is completed and VFTO is reached, whichever
point is higher. In addition--
(1) The takeoff path must be based on the procedures prescribed in
Sec. 25.101(f);
[[Page 26912]]
(2) The airplane must be accelerated on the ground to
VEF, at which point the critical engine must be made
inoperative and remain inoperative for the rest of the takeoff; and
(3) After reaching VEF, the airplane must be accelerated
to V2.
(g) In lieu of compliance with Sec. 25.119 (b), the following
applies: A climb speed of not more than VREF.
(h) In lieu of compliance with Sec. 25.121(c), the following
applies:
Final takeoff. In the en route configuration at the end of the
takeoff path determined in accordance with Sec. 25.111, the steady
gradient of climb may not be less than 1.2 percent for two-engine
airplanes, 1.5 percent for three-engine airplanes, and 1.7 percent for
four engine airplanes, at VFTO and with--
(1) The critical engine inoperative and the remaining engines at
the available maximum continuous power or thrust; and
(2) The weight equal to the weight existing at the end of the
takeoff path, determined under Sec. 25.111.
(i) In lieu of compliance with Sec. 25.121(d), the following
applies:
Approach. In a configuration corresponding to the normal all-
engines-operating procedure in which VSRPWR for this
configuration, with the operative engines at the minimum thrust or
power existing at any point in the go-around, does not exceed 110
percent of the VSRPWR for the related all-engines-operating
landing configuration, with the operative engines at the power or
thrust setting for approach at the reference flight path angle at
VREF, the steady gradient of climb may not be less than 2.7
percent with--
(1) The critical engine inoperative, the remaining engines at the
go-around power or thrust setting;
(2) The maximum landing weight;
(3) A climb speed established in connection with normal landing
procedures, but not more than 1.4 VSRPWR with the operative
engines at the minimum power or thrust setting existing at any point in
the go-around; and
(4) The landing gear retracted.
(j) In lieu of compliance with Sec. 25.125(a)(2), the following
applies: A stabilized approach, with a calibrated airspeed of not less
than VREF or VMCL, whichever is greater, must be
maintained down to the 50 foot height. VREF may not be less
than--
(1) 1.03 VSR0;
(2) 1.20 VSR0PWR with the operative engines at the power
or thrust setting for approach at the reference flight path angle;
(3) The airspeed that provides an angle-of-attack margin to stall
for not less than a 20 knot equivalent airspeed vertical gust with all
engines operating at the power or thrust setting for approach at the
reference flight path angle;
(4) The airspeed that provides an angle-of-attack margin to stall
for not less than a 15 knot equivalent airspeed vertical gust with the
critical engine inoperative at the power or thrust setting for approach
at the reference flight path angle; and
(5) A speed that provides the maneuvering capability specified in
paragraph (k) below.
(k) In addition to compliance with Sec. 25.143, the following
applies: The maneuvering capabilities in a constant speed coordinated
turn, as specified in the table below, must be free of stall warning or
other characteristics that might interfere with normal maneuvering.
BILLING CODE 4910-13-U
[[Page 26913]]
[GRAPHIC] [TIFF OMITTED] TP18MY99.012
BILLING CODE 4910-13-C
(l) In lieu of compliance with Sec. 25.145(a), the following
applies: It must be possible at any speed between the trim speed
prescribed in paragraph (b)(1)(v), or (b)(4)(v), of this special
condition for flaps extended configurations, and the minimum speed
obtained in conducting a stalling maneuver, to pitch the nose downward
so that the acceleration to this selected trim speed is prompt with--
(1) The airplane trimmed at the speed prescribed in paragraph
(b)(1)(v) of this special condition for flaps retracted configurations,
or as prescribed in paragraph (b)(4)(v) of this special condition for
flaps extended configurations;
(2) The landing gear extended;
(3) The wing flaps--
(i) retracted, and
(ii) extended; and
(4) Power--
(i) off with the flaps retracted and, with the flaps extended, with
all engines operating at the minimum power or thrust level consistent
with that used to determine the power-on reference stall speeds; and
(ii) at maximum continuous power on the engines.
(m) In lieu of compliance with Sec. 25.145(b)(2), the following
applies: Repeat paragraph (b)(1) of this section, except begin with the
flaps fully extended and all engines at the minimum power or thrust
level consistent with that used to determine the power-on reference
stall speed for that flap position, and then retract the flaps as
rapidly as possible.
(n) In lieu of compliance with Sec. 25.145(b)(5), the following
applies: Repeat paragraph (b)(4) of this section, except with the flaps
extended and all engines at the minimum power or thrust level
consistent with that used to determine the reference power-on stall
speed.
(o) In lieu of compliance with Sec. 25.145(b)(6), the following
applies: With all engines at the minimum power or thrust level
consistent with that used to determine the reference power-on stall
speed, flaps extended, and the airplane trimmed at 1.3
VSR1PWR, obtain and maintain airspeeds between
VSW, and either 1.6 VSR1PWR or VFE,
whichever is lower.
(p) In lieu of compliance with Sec. 25.161(c)(2), the following
applies: A glide with the landing gear extended, the most unfavorable
center of gravity position approved for landing with the maximum
landing weight, and the most unfavorable center of gravity position
approved for landing, regardless of weight with the wing flaps--
(1) retracted with power off at a speed of 1.3 VSR1, and
(2) extended with all engines at the minimum power or thrust level
consistent with that used to determine the power-on reference stall
speed at a speed of 1.3 VSR1PWR.
(q) In lieu of compliance with Sec. 25.175(d)(4), the following
applies: All engines at the minimum power or thrust level consistent
with that used to determine the power-on reference stall speed.
(r) In lieu of compliance with Sec. 25.175(d)(5), the following
applies: The airplane trimmed at 1.3 VSR0PWR.
[[Page 26914]]
(s) In lieu of the speeds given in the following part 25
requirements, comply with the speeds as follows:
Secs. 25.145(b)(1) and (b)(4), 1.3 VSR1, in lieu of 1.4
VS1.
Sec. 25.145(b)(1), 30 percent, in lieu of 40 percent.
Sec. 25.145(b)(1), power-on reference stall speed, in lieu of
stalling speed.
Sec. 25.145(c), 1.08 VSR1, in lieu of 1.1
VS1.
Sec. 25.145(c), 1.18 VSR1PWR, in lieu of 1.2
VS1.
Sec. 25.147(a), (a)(2), (c), and (d), 1.3 VSR1, in lieu
of 1.4 VS1.
Sec. 25.149(c), 1.13 VSR, in lieu of 1.2 VS.
Sec. 25.161(b), (c)(1), and (c)(2), 1.3 VSR1, or 1.3
VSR1PWR for flaps extended configurations, in lieu of 1.4
VS1.
Sec. 25.161(c)(3), 1.3 VSR1, in lieu of the first
instance of 1.4 VS1, and 1.3 VSR1PWR, in lieu of
the second instance of 1.4 VS1.
Sec. 25.161(d), 1.3 VSR1 in lieu of 1.4 VS1.
Sec. 25.161(e)(3), 0.013 VSR02, in lieu of
0.013 VS02.
Sec. 25.175(a)(2), (b)(1), (b)(2), and (b)(3), 1.3 VSR1,
in lieu of 1.4 VS1.
Sec. 25.175(b)(2)(ii), (VMO + 1.3 VSR1)/2, in
lieu of VMO + 1.4 VS1/2.
Sec. 25.175(c), VSW and 1.7 VSR1PWR, in lieu
of 1.1 VS1 and 1.8 VS1.
Sec. 25.175(c)(4), 1.3 VSR1PWR, in lieu of 1.4
VS1.
Sec. 25.175(d), VSW and 1.7 VSR0PWR, in lieu
of 1.1 VS0 and 1.3 VS0.
Sec. 25.177(c), 1.13 VSR1, or 1.18 VSR1PWR
for flaps extended configurations, in lieu of 1.2 VS1.
Sec. 25.181(a) and (b), 1.13 VSR1, or 1.18
VSR1PWR for flaps extended configurations, in lieu of 1.2
VS1.
Sec. 25.201(a)(2), 1.5 VSR1PWR (where VSR1PWR
corresponds to the power-on reference stall speed with flaps in the
approach position, the landing gear retracted, and maximum landing
weight), in lieu of 1.6 VS1 (where VS1
corresponds to the stalling speed with flaps in the approach position,
the landing gear retracted, and maximum landing weight).
(t) In addition to compliance with Secs. 25.201(a)(1) and (a)(2),
the following also applies: The critical engine inoperative and the
power or thrust setting on the remaining engines at the minimum power
or thrust level appropriate for the flight condition used to show
compliance with a required performance standard.
(u) In lieu of compliance with Sec. 25.207(b), the following
applies: The warning may be furnished either through the inherent
aerodynamic qualities of the airplane or by a device that will give
clearly distinguishable indications under expected conditions of
flight. However, a visual stall warning device that requires the
attention of the crew within the cockpit is not acceptable by itself.
If a warning device is used, it must provide a warning in each of the
airplane configurations prescribed in paragraph (a) of this section at
the speed prescribed in paragraph (v)(1) and (2) below.
(v) In lieu of compliance with Sec. 25.207(c), the following
applies:
(1) In each normal configuration with the flaps retracted, when the
speed is reduced at rates not exceeding one knot per second, stall
warning must begin at a speed, VSW, exceeding the speed at
which the stall is identified in accordance with Sec. 25.201(d) by not
less than five knots or five percent, whichever is greater. Once
initiated, stall warning must continue until the angle of attack is
reduced to approximately that at which stall warning began.
(2) In addition to the requirement of paragraph (v)(1) above, when
the speed is reduced at rates not exceeding one knot per second, in
straight flight with engines idling and at the center of gravity
position specified in paragraph (b)(1)(iv) above, VSW, in
each normal configuration with the flaps retracted, must exceed
VSR by not less than three knots or three percent, whichever
is greater.
(3) In each normal configuration with the flaps extended, when the
speed is reduced at rates not exceeding one knot per second, stall
warning must begin at a speed, VSW, exceeding the speed at
which the stall is identified in accordance with Sec. 25.201(d) by not
less than five knots or five percent, whichever is greater. Once
initiated, stall warning must continue until the angle of attack is
reduced to approximately that at which stall warning began.
(4) In addition to the requirement of paragraph (v)(3) above, when
the speed is reduced at rates not exceeding one knot per second, in
straight flight with the critical engine inoperative and the power or
thrust setting on the remaining engines at the minimum power or thrust
level appropriate for the flight condition used to show compliance with
a required performance standard, and at the center of gravity position
specified in paragraph (b)(4)(i) above, VSW, in each normal
configuration with the flaps extended, must exceed VSRPWR by
not less than three knots or three percent, whichever is greater.
(5) In slow-down turns with at least 1.5g load factor normal to the
flight path and airspeed deceleration rates greater than 2 knots per
second, with the flaps and landing gear in any normal position, the
stall warning margin must be sufficient to allow the pilot to prevent
stalling (as defined in Sec. 25.201(d)) when recovery is initiated not
less than one second after the onset of stall warning. Compliance with
this requirement must be demonstrated with--
(i) The airplane trimmed for straight flight at a speed of 1.3
VSR with the flaps retracted or 1.3 VSRPWR with
the flaps extended; and
(ii) The power or thrust necessary to maintain level flight at 1.3
VSR with the flaps retracted or 1.3 VSRPWR with
the flaps extended.
(w) In addition to compliance with Sec. 25.207(a) and paragraphs
(u) and (v) above, the following applies: Stall warning must also be
provided in each abnormal configuration of the high lift devices likely
to be used in flight following system failures (including all
configurations covered by Airplane Flight Manual procedures).
(x) In lieu of the speeds given in Secs. 25.233(a) and 25.237(a),
comply with speeds as follows: 0.2 VSR0PWR in lieu of 0.2
VS0.
(y) In lieu of the definition of V in Sec. 25.735(f)(2), the
following apply:
V=VREF/1.3
VREF=Airplane steady landing approach speed, in knots,
at the maximum design landing weight and in the landing configuration
at sea level.
(z) In lieu of compliance with Sec. 25.735(g), the following
applies: The minimum speed rating of each main wheel-brake assembly
(that is, the initial speed used in the dynamometer tests) may not be
more than the V used in the determination of kinetic energy in
accordance with paragraph (f) of this section, assuming that the test
procedures for wheel-brake assemblies involve a specified rate of
deceleration, and, therefore, for the same amount of kinetic energy,
the rate of energy absorption (the power absorbing ability of the
brake) varies inversely with the initial speed.
(aa) In lieu of the speeds given in the following part 25
requirements, comply with the speeds as follows:
Sec. 25.773(b)(1)(i), 1.5 VSR1, in lieu of 1.6
VS1.
Sec. 25.1001(c)(1) and (c)(3), 1.3 VSR1, in lieu of 1.4
VS1.
Sec. 25.1323(c)(1), 1.23 VSR1, in lieu of 1.3
VS1.
Sec. 25.1323(c)(2), 1.20 VSR0PWR, in lieu of 1.3
VS0.
Sec. 25.1325(e), 1.20 VSR0PWR, in lieu of 1.3
VS0, and 1.7 VSR1, in lieu of 1.8 VS1.
2. Head-up Display Used as a Primary Flight Display
(a) Display Requirements.
(1) The HUD must provide information necessary to enable rapid
[[Page 26915]]
pilot interpretation of the airplane's flight state and position during
all phases of flight. This information shall enable the flightcrew to
manually control the airplane and monitor the performance of the
automatic flight control system. The HUD display shall enable manual
airplane control and including guidance, if necessary, during an engine
failure during any phase of flight. The monochrome HUD must
equivalently perform the intended function of conventional color
primary flight instruments and utilize display features that compensate
for the lack of color. Operational acceptability of the HUD system for
use while manually controlling the airplane shall be demonstrated and
evaluated by the FAA. This task-oriented demonstration will evaluate
crew workload and pilot compensation for normal, abnormal, and
emergency operations, with single and multiple failures not shown to be
extremely improbable by the system safety analysis, and is extended to
all HUD display formats, unless use of specific formats is prohibited
for specific phases of flight.
(2) The current mode of the flight guidance/automatic flight
control system shall be clearly annunciated in the HUD, unless it is
displayed elsewhere in close proximity to the HUD field of view and
shown to be equivalently conspicuous. Likewise, other essential
information and alerts that are related to displayed information and
may require immediate pilot action must be displayed for instant
recognition. Such information, depending on the phase of flight,
includes malfunctions of primary data sources, guidance and control,
and excessive deviations that require a go-around maneuver.
(3) If a windshear detection system or a traffic alert and
collision avoidance system (TCAS) is installed, the guidance will be
provided on the HUD. When the ground proximity warning system detects
excessive terrain closure, appropriate annunciations are displayed on
the HUD. Additional warnings and annunciations that are required to be
a part of these systems, and are normally required as part of the
approved design to be in the pilot's primary field of view (i.e., the
line of vision when looking forward along the flight path), must remain
in the pilot's primary field of view when utilizing the HUD for flight
information.
(4) Symbols must appear clean-shaped, clear, and explicit. Lines
must be narrow, sharp-edged, and without halo or aliasing. Symbols must
be stable with no discernible flicker or jitter.
(5) The optical qualities (accommodation, luminance, vergence) of
the HUD shall be uniform across the entire field of view. When viewed
by both eyes from any off-center position within the eyebox, non-
uniformities shall not produce perceivable differences in binocular
view.
(6) For all phases of flight, the HUD must update the positions and
motions of primary control symbols with sufficient rates and latencies
to support satisfactory manual control performance.
(7) The HUD display must present all information in a clear and
unambiguous manner. Display clutter must be minimized. The HUD
symbology must not interfere with the pilots' forward view, ability to
visually maneuver the airplane, acquire opposing traffic, and see the
runway environment. Critical and essential data elements of primary
flight displays must not be removed by any declutter function. Changes
in the display format and primary flight data arrangement should be
minimized to prevent confusion and to enhance the pilots' ability to
interpret vital data.
(8) The content, arrangement, and format of the information must be
sufficiently compatible with the head down displays to preclude pilot
confusion, misinterpretation, or excessive cognitive workload.
Immediate transition between the two displays, whether required by
navigation duties, failure conditions, unusual airplane attitudes, or
other reasons, must not present difficulties in data interpretation or
delays/interruptions in the crew's ability to manually control the
airplane or to monitor the automatic flight control system.
(9) The HUD display must enable the flightcrew to immediately
recognize and perform a safe recovery from unusual airplane attitudes.
This capability must be shown in a simulator and on the airplane for
all foreseeable modes of upset. However, ``corner conditions'' (i.e.,
test conditions where more than one attitude parameter is at its
extreme value) may be demonstrated in the simulator. Foreseeable modes
of upset include--
(i) flightcrew mishandling;
(ii) autopilot failure (including ``slowovers'' which are slowly
developing changes in attitude that do not create forces directly felt
by the pilot, and are only detectable by pilot reference to the flight
instruments or automatic alerts); and
(iii) turbulence/gust encounters.
(b) Installation Requirements.
(1) The arrangement of HUD display controls must be visible to and
within reach of the pilot from any normal seated position. The position
and movement of the controls must not lead to inadvertent operation.
The HUD controls must be illuminated to be visible for all normal
cockpit lighting conditions, and must not create any objectionable
reflections on the HUD or other flight instruments.
(2) The HUD combiner brightness must be controllable to ensure
uninterrupted visibility of all displayed information in the presence
of dynamically changing background (ambient) lighting conditions. If
automatic control of HUD brightness is not provided, it must be shown
that a single setting is satisfactory. When the HUD brightness level is
changed, the relative luminance of each displayed symbol, character, or
data shall vary smoothly. In no case shall any selectable brightness
level allow any information to be invisible while other data remains
discernible. There shall be no objectionable brightness transients when
switching between manual and automatic control. The HUD data shall be
visible in lighting conditions from 0 fL to 10,000 fL. If certain
lighting conditions prevent the crew from seeing and interpreting HUD
data (for example, flying directly toward the sun), accommodation must
be provided to permit the crew to make a ready transition to the head
down displays.
(3) To the greatest extent practicable, the HUD controls must be
integrated with other controls, including the flight director, to
minimize the crew workload associated with HUD operation and to ensure
flightcrew awareness of engaged flight guidance modes.
(4) The visibility of the HUD and the primary flight information
displayed is paramount to the HUD's ability to perform its intended
function as a primary flight display. The fundamental requirements for
instrument arrangement and visibility specified in Secs. 25.1321,
25.773, and 25.777 apply to these devices.
The design eyebox should be laterally and vertically centered
around the respective pilot's design eye position, and should be large
enough that the minimum monocular field of view is visible at the
following minimum displacements from the cockpit design eye position:
Lateral: 1.5 inches left and right
Vertical: 1.0 inches up and down
Longitudinal: 2.0 inches fore and aft
The HUD installation must accommodate pilots from 5'2'' to 6'3''
tall, seated with seat belts fastened and positioned at the design eye
position (ref. Sec. 25.777(c)). Larger eyebox
[[Page 26916]]
dimensions may be required for meeting operational requirements for use
as a full time primary flight display. Operational suitability and
compliance with the requirements of the above cited regulations must be
demonstrated and evaluated by the FAA. The design eye position must
comply with the above cited regulations.
(5) Notwithstanding compliance with the minimum eyebox dimensions
given above, the HUD eyebox must be large enough to serve as a primary
flight display without inducing adverse effects on pilot vision and
fatigue. Use of the HUD system shall not place physiologically
burdensome limitations on head position. There must be no adverse
physiological effects of long term use of the HUD system, such as
fatigue or eye strain, that force the pilot to revert to the HDD. Long
term use is considered four hours of continuous use of the HUD, or
multiple flights per day with eight or more hours of use.
(c) System Requirements.
(1) The HUD system must be shown to perform its intended function
as a primary flight display during all phases of flight. The normal
operation of the HUD system cannot adversely affect, or be adversely
affected by, other airplane systems. Malfunctions of the HUD system
that cause loss of all primary flight information, including that
displayed on the HUD and head down instruments, shall be extremely
improbable.
(2) The classification of the HUD system's failure to display
flight information and navigation information, as applicable to the
airplane type design, including the potential to display hazardously
misleading information, must be assessed according to Secs. 25.1309 and
25.1333. All alleviating flightcrew actions that are considered in the
HUD safety analysis must be validated during testing for incorporation
in the airplane flight manual procedures section or for inclusion in
type-specific training. The failure cases discussed below, which
consider the entire suite of cockpit displays of each flight parameter,
hazardously misleading failures are, by definition, not associated with
a suitable warning.
(i) Attitude. Display of attitude in the cockpit is a critical
function. Loss of all attitude display, including standby attitude, is
classified as a catastrophic failure and must be extremely improbable.
Loss of primary attitude display for both pilots is classified as a
major failure and must be improbable. Display of hazardously misleading
roll or pitch attitude simultaneously on the primary attitude displays
for both pilots is classified as a catastrophic failure and must be
extremely improbable. Display of hazardously misleading roll or pitch
attitude on any single primary attitude display is classified as a
major failure and must be improbable.
(ii) Airspeed. Display of airspeed in the cockpit is a critical
function. Loss of all airspeed display, including standby, is
classified as a catastrophic failure and must be extremely improbable.
Loss of primary airspeed display for both pilots is classified as a
major failure and must be improbable. Displaying hazardously misleading
airspeed simultaneously on both pilots' displays, coupled with the loss
of stall warning or overspeed warning functions, is classified as a
catastrophic failure and must be extremely improbable.
(iii) Barometric Altitude. Display of altitude in the cockpit is a
critical function. Loss of all altitude display, including standby, is
classified as a catastrophic failure and must be extremely improbable.
Loss of primary altitude display for both pilots is classified as a
major failure and must be improbable. Displaying hazardously misleading
altitude simultaneously on both pilots' displays is classified as a
catastrophic failure and must be extremely improbable.
(iv) Vertical Speed. Display of vertical speed in the cockpit is an
essential function. Loss of vertical speed display to both pilots is
classified as a major failure and must be improbable.
(v) Slip/Skid Indication. The slip/skid or side slip indication is
an essential function. Loss of this function to both pilots is
classified as a major failure and must be improbable. Simultaneously
misleading slip/skid or side slip information to both pilots is
classified as a major failure and must be improbable.
(vi) Heading. Display of stabilized heading in the cockpit is an
essential function. Displaying hazardously misleading heading
information on both pilots' primary displays is classified as a major
failure and must be improbable. Loss of stabilized heading in the
cockpit is classified as a major failure and must be improbable. Loss
of all heading information in the cockpit is classified as a
catastrophic failure and must be extremely improbable.
(vii) Navigation. Display of navigation information (excluding
heading, airspeed, and clock data) in the cockpit is an essential
function. Loss of all navigation information is classified as a major
failure and must be improbable. Displaying hazardously misleading
navigational or positional information simultaneously on both pilots'
displays is classified as a major failure and must be improbable.
However, the nonrestorable loss of the combination of all navigation
and communication functions is classified as a catastrophic failure and
must be extremely improbable.
(viii) Crew Alerting Displays. Loss of crew alerting for essential
functions is classified as a major failure and must be improbable.
Display of hazardously misleading crew alerting messages is classified
as a major failure and must be improbable.
(3) The display of hazardously misleading information on more than
one primary flight display is classified as a catastrophic failure and
must be extremely improbable; therefore, the HUD system software which
generates, displays, or affects the generation or display of primary
flight information shall be developed to Level A requirements, as
specified by RTCA Document DO-178B, ``Software Considerations in
Airborne Systems and Equipment Certification,'' or similar processes
that provide equivalent product and compliance data. Monitoring
software shown to have no ability to generate, display, or affect the
generation or display of primary flight information, and which has the
capability to command shutdown of the HUD system, shall be developed to
no less rigor than that defined for Level C, or criticality as
determined by a safety assessment of the HUD system.
(4) The HUD system must monitor the position of the combiner and
provide a warning to the crew when the combiner position is such that
conformal symbols will be hazardously misaligned.
(5) The HUD system must be shown to comply with the high intensity
radiated fields certification requirements of Special Condition No. 3.
3. Protection From Unwanted Effects of High Intensity Radiated Fields
(a) Each electrical and electronic system that performs critical
functions must be designed and installed to ensure that the operation
and operational capability of these systems to perform critical
functions are not adversely affected when the airplane is exposed to
high-intensity radiated fields.
(b) For the purpose of this special condition, the following
definition applies:
Critical Functions. Functions whose failure would contribute to or
cause a failure condition that would prevent the continued safe flight
and landing of the airplane.
Discussion: With the trend toward increased power levels from
ground-
[[Page 26917]]
based transmitters, plus the advent of space and satellite
communications, coupled with electronic command and control of the
airplane, the immunity of critical digital avionics systems to HIRF
must be established.
It is not possible to precisely define the HIRF to which the
airplane will be exposed in service. There is also uncertainty
concerning the effectiveness of airframe shielding for HIRF.
Furthermore, coupling of electromagnetic energy to cockpit-installed
equipment through the cockpit window apertures is undefined. Based on
surveys and analysis of existing HIRF emitters, an adequate level of
protection exists when compliance with the HIRF protection special
condition is shown with either paragraph 1 or 2 below:
1. A minimum threat of 100 volts per meter peak electric field
strength from 10 KHz to 18 GHz.
a. The threat must be applied to the system elements and their
associated wiring harnesses without the benefit of airframe shielding.
b. Demonstration of this level of protection is established through
system tests and analysis.
2. A threat external to the airframe of the following field
strengths for the frequency ranges indicated.
------------------------------------------------------------------------
Field strength
(volts per meter)
Frequency ---------------------
Peak Average
------------------------------------------------------------------------
10 KHz-100 KHz.................................... 30 30
100 KHz-500 KHz................................... 40 30
500 KHz-2 MHz..................................... 30 30
2 MHz-30 MHz...................................... 190 190
30 MHz-70 MHz..................................... 20 20
70 MHz-100 MHz.................................... 20 20
100 MHz-200 MHz................................... 30 30
200 MHz-400 MHz................................... 30 30
400 MHz-700 MHz................................... 80 80
700 MHz-1 GHz..................................... 690 240
1 GHz-2 GHz....................................... 970 70
2 GHz-4 GHz....................................... 1570 350
4 GHz-6 GHz....................................... 7200 300
6 GHz-8 GHz....................................... 130 80
8 GHz-12 GHz...................................... 2100 80
12 GHz-18 GHz..................................... 500 330
18 GHz-40 GHz..................................... 780 20
------------------------------------------------------------------------
4. Interaction of Systems and Structures
(a) General. Airplanes equipped with systems that affect structural
performance, either directly or as a result of a failure or
malfunction, must account for the influence of these systems and their
failure conditions in showing compliance with the requirements of
subparts C and D of part 25. The following criteria must be used to
evaluate the structural performance of airplanes equipped with flight
control systems, autopilots, stability augmentation systems, load
alleviation systems, flutter control systems, and fuel management
systems: If these criteria are used for other systems, it may be
necessary to adapt the criteria to the specific system.
(b) System fully operative. With the system fully operative, the
following apply:
(1) Limit loads must be derived in all normal operating
configurations of the systems from all the limit conditions specified
in subpart C, taking into account any special behavior of such systems
or associated functions or any effect on the structural performance of
the airplane that may occur up to the limit loads. In particular, any
significant nonlinearity (rate of displacement of control surface,
thresholds, or any other system nonlinearities) must be accounted for
in a realistic or conservative way when deriving limit loads from limit
conditions.
(2) The airplane must meet the strength requirements of part 25
(static strength, residual strength), using the specified factors to
derive ultimate loads from the limit loads defined in paragraph (b)(1)
above. The effect of nonlinearities must be investigated beyond limit
conditions to ensure the behavior of the systems presents no anomaly
compared to the behavior below limit conditions. However, conditions
beyond limit conditions need not be considered when it can be shown
that the airplane has design features that make it impossible to exceed
those limit conditions.
(3) The airplane must meet the aeroelastic stability requirements
of Sec. 25.629.
(c) System in the Failure Condition. For any system failure
condition not shown to be extremely improbable, the following apply:
(1) At the time of occurrence. Starting from 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. The airplane must be able to
withstand these loads, multiplied by an appropriate factor of safety
that is related to the probability of occurrence of the failure. The
factor of safety (F.S.) is defined in Figure 1.
BILLING CODE 4910-13-U
[GRAPHIC] [TIFF OMITTED] TP18MY99.013
BILLING CODE 4910-13-C
(i) These loads must also be used in the damage tolerance
evaluation required by Sec. 25.571(b) if the failure condition is
probable.
(ii) Freedom from aeroelastic instability must be shown up to the
[[Page 26918]]
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 the increased speeds, so that
the margins intended by Sec. 25.629(b)(2) are maintained.
(iii) Notwithstanding subparagraph (1) of this paragraph, failures
of the system that result in forced structural vibrations (oscillatory
failures) must not produce peak loads that could result in catastrophic
fatigue failure or 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) Static and residual strength must be determined for loads
derived from the following conditions at speeds up to Vc, or
the speed limitation prescribed for the remainder of the flight:
(A) The limit symmetrical maneuvering conditions specified in
Secs. 25.331 and 25.345.
(B) The limit gust conditions specified in Sec. 25.341, but using
the gust velocities for Vc, and in Sec. 25.345.
(C) The limit rolling conditions specified Sec. 25.349 and the
limit unsymmetrical conditions specified in Secs. 25.367 and 25.427 (b)
and (c).
(D) The limit yaw maneuvering conditions specified in Sec. 25.351.
(E) The limit ground loading conditions specified in Secs. 25.473
and 25.491.
(ii) For static strength substantiation, each part of the structure
must be able to withstand the loads specified in subparagraph (2)(i) of
this paragraph, multiplied by a factor of safety depending on the
probability of being in this failure state. The factor of safety is
defined in Figure 2.
BILLING CODE 4910-13-U
[GRAPHIC] [TIFF OMITTED] TP18MY99.014
BILLING CODE 4910-13-C
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 as defined in
Sec. 25.571(b), structures affected by failure of the system and with
damage in combination with the system failure, a reduced factor may be
applied to the loads of subparagraph (2)(i) of this paragraph. However,
the residual strength level must not be less than the 1-g flight load,
combined with the loads introduced by the failure condition, plus two-
thirds of the load increments of the conditions specified in
subparagraph (2)(i) of this paragraph, applied in both positive and
negative directions (if appropriate). The residual strength factor
(R.S.F.) is defined in Figure 3.
BILLING CODE 4910-13-U
[[Page 26919]]
[GRAPHIC] [TIFF OMITTED] TP18MY99.015
BILLING CODE 4910-13-C
Note: If Pj is greater than 10-3 per
flight hour, then a residual strength factor of 1.0 must be used.
(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 the
speeds determined from Figure 4.
BILLING CODE 4910-13-U
[GRAPHIC] [TIFF OMITTED] TP18MY99.016
BILLING CODE 4910-13
Note: If Pj is greater than 10-3 per
flight hour, then the flutter clearance speed must not be less than
V''.
(vi) Freedom from aeroelastic instability must also be shown up to
V' in Figure 4 above, for any probable system failure condition
combined with any damage considered in the evaluation required by
Sec. 25.571(b).
(vii) If the mission analysis method is used to account for
continuous turbulence, all the systems failure conditions associated
with their probability must be accounted for in a rational or
conservative manner in order to ensure that the probability of
exceeding the limit load is not higher than the value prescribed in
appendix G to part 25.
(3) Consideration of certain failure conditions may be required by
other sections of this part, 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.
[[Page 26920]]
(d) Warning Considerations. For system failure detection and
warning, the following apply:
(1) The system must be checked for failure conditions, not shown to
be extremely improbable, that degrade the structural capability of the
airplane below the level required by 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 ensure failure detection. 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 shown to be
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 below 1.25, as determined
by paragraph (c) of this special condition, or flutter clearance speeds
below V'', as determined by paragraph (c) of this special condition,
must be signaled to the flightcrew during flight.
(e) 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 this
special condition must be met for the dispatched condition and for
subsequent failures. Operational and flight limitations may be taken
into account.
(f) The following definitions are applicable to this special
condition:
Structural performance: The capability of the airplane to meet the
structural requirements of part 25.
Flight limitations: Limitations that can be applied to the airplane
flight conditions following an 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 and payload limitations).
Probabilistic terms: The probabilistic terms (probable, improbable,
extremely improbable) used in this special condition are the same as
those used in Advisory Circular (AC) 25.1309-1A.
Failure condition: The term failure condition is the same as that
used in AC 25.1309-1A; however, this special condition applies only to
system failure conditions that affect the structural performance of the
airplane (e.g., failure conditions that induce loads, change the
response of the airplane to inputs such as gusts or pilot actions, or
lower flutter margins).
5. Design Maneuvering Requirements for Fly-by-Wire
(a) Maximum elevator displacement at VA. In lieu of
compliance with Sec. 25.331(c)(1) of the FAR; the airplane is assumed
to be flying in steady level flight (point A1,
Sec. 25.333(b)) and, except as limited by pilot effort in accordance
with Sec. 25.397, 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 that occur subsequent to the normal
acceleration at the center of gravity exceeding the maximum positive
limit maneuvering factor, n, need not be considered.
(b) Pitch maneuver loads. In addition to the requirements of
Sec. 25.331; 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
accounted for.
(c) Roll maneuver loads. In lieu of compliance with Sec. 25.349(a),
the following conditions, speeds, and spoiler and aileron deflections
(except as the deflections may be limited by pilot effort) must be
considered in combination with an airplane load factor of zero and of
two-thirds of the positive maneuvering factor used in design. In
determining the required aileron and spoiler deflections, the torsional
flexibility of the wing must be considered in accordance with
Sec. 25.301(b).
(1) Conditions corresponding to steady rolling velocities must be
investigated. In addition, conditions corresponding to maximum angular
acceleration must be investigated. For the angular acceleration
conditions, zero rolling velocity may be assumed in the absence of a
rational time history investigation of the maneuver.
(2) At VA, sudden deflection of the cockpit roll control
up to the limit is assumed.
(3) At VC, the cockpit roll control must be moved
suddenly and maintained so as to achieve a rate of roll not less than
that obtained in paragraph (2).
(4) At VD, the cockpit roll control must be moved
suddenly and maintained so as to achieve a rate of roll not less than
one third of that obtained in paragraph (2).
(5) It must also be established that roll maneuver loads induced by
the system itself (i.e., abrupt changes in orders made possible by
electrical rather than mechanical combination of different inputs) are
acceptably accounted for.
(d) Yaw maneuver loads. In lieu of compliance with Sec. 25.351, the
airplane must be designed for loads resulting from the conditions
specified in paragraph (e) below. Unbalanced aerodynamic moments about
the center of gravity must be reacted in a rational or conservative
manner considering the principal masses furnishing the reacting inertia
forces. Physical limitations of the airplane from the cockpit yaw
control device to the control surface deflection, such as control stop
position, maximum power and displacement rate of the servo controls, or
control law limiters, may be taken into account.
(e) Maneuvering. At speeds from VMC to VD,
the following maneuvers must be considered. In computing the tail
loads, the yawing velocity may be assumed to be zero.
(1) With the airplane in unaccelerated flight at zero yaw, it is
assumed that the cockpit yaw control device (pedal) is suddenly
displaced (with critical rate) to the maximum deflection, as limited by
the stops.
(2) With the cockpit yaw control device (pedal) deflected as
specified in paragraph (1) above, it is assumed that the airplane yaws
to the resulting side slip angle (beyond the static side slip angle).
(3) With the airplane yawed to the static sideslip angle with the
cockpit yaw control device deflected as specified in paragraph (1)
above, it is assumed that the cockpit yaw control device is returned to
neutral.
6. Limit Engine Torque Loads for Sudden Engine Stoppage
In lieu of showing compliance with Sec. 25.361(b), the following
apply:
(a) For turbine engine and auxiliary power unit installations, the
mounts and local supporting structure must be designed to withstand
each of the following:
(1) The maximum limit torque load imposed by--
(i) A sudden deceleration due to a malfunction that could result in
a
[[Page 26921]]
temporary loss of power or thrust capability, and could cause a
shutdown due to vibrations; and
(ii) The maximum acceleration of the engine and auxiliary power
unit.
(2) The maximum torque load, considered as ultimate, imposed by
sudden engine or auxiliary power unit stoppage due to a structural
failure, including fan blade failure.
(3) The load condition defined in paragraph (a)(2) of this section
is also assumed to act on adjacent airframe structure, such as the wing
and fuselage. This load condition is multiplied by a factor of 1.25 to
obtain ultimate loads when the load is applied to the wing and fuselage
structure.
7. Flight Characteristic Compliance Determination by use of the
Handling Qualities Rating System for EFCS Failure Cases
(a) In lieu of showing compliance with Sec. 25.672(c), a handling
qualities rating system will be used for evaluation of EFCS
configurations resulting from single and multiple failures not shown to
be extremely improbable. The handling qualities ratings are:
(1) Satisfactory: Full performance criteria can be met with routine
pilot effort and attention.
(2) Adequate: Adequate for continued safe flight and landing; full
or specified reduced performance can be met, but with heightened pilot
effort and attention.
(3) Controllable: Inadequate for continued safe flight and landing,
but controllable for return to a safe flight condition, safe flight
envelope, and/or reconfiguration so that the handling qualities are at
least adequate.
(b) Handling qualities will be allowed to progressively degrade
with failure state, atmospheric disturbance level, and flight envelope.
Specifically, within the normal flight envelope, the pilot-rated
handling qualities must be satisfactory/adequate in moderate
atmospheric disturbance for probable failures, and must not be less
than adequate in light atmospheric disturbance for improbable failures.
8. Static Longitudinal Stability
In lieu of compliance with Sec. 25.173, the airplane must be shown
to have suitable static longitudinal stability in any condition
normally encountered in service, including the effects of atmospheric
disturbance. The HQRS may be used to make this assessment.
9. Static Lateral-Directional Stability
In lieu of compliance with Sec. 25.177, the following applies:
(a) The airplane must be shown to have suitable static lateral
directional stability in any condition normally encountered in service,
including the effects of atmospheric disturbance. The HQRS may be used
to make this assessment.
(b) In straight, steady sideslips, the rudder control movements and
forces must be substantially proportional to the angle of sideslip in a
stable sense; and the factor of proportionality must lie between limits
found necessary for safe operation throughout the range of sideslip
angles appropriate to the operation of the airplane. At greater angles,
up to the angle at which full rudder is used or a rudder force of 180
pounds is obtained, the rudder pedal forces may not reverse; and
increased rudder deflection must be needed for increased angles of
sideslip. Compliance with this paragraph must be demonstrated for all
landing gear and flap positions and symmetrical power conditions at
speeds from 1.13 VSR1, or 1.18 VSR1PWR for flaps
extended configurations, to VFE, VLE, or
VFC/MFC, as appropriate.
10. Control Surface Awareness
In addition to compliance with Secs. 25.143, 25.671, and 25.672,
when a flight condition exists where, without being commanded by the
crew, control surfaces are coming so close to their limits that return
to the normal flight envelope and (or) continuation of safe flight
requires a specific crew action, a suitable flight control position
annunciation shall be provided to the crew, unless other existing
indications are found adequate or sufficient to prompt that action.
Note: The term suitable also indicates an appropriate balance
between nuisance and necessary operation.
11. Steep Approach Air Distance
In lieu of compliance with Sec. 25.125(a) for steep approach
landing distances, the following applies:
(a) The horizontal distance necessary to land and to come to a
complete stop, including an airborne distance of no less than the
greater of either 500 feet or the distance resulting from the
combination of the distance between the runway threshold and the
touchdown aim point to be used in operations plus the demonstrated
3 dispersion distance from the touchdown aim point, must be
determined (at each weight for temperature, altitude, and wind within
the operational limits established by the applicant for the airplane)
as follows:
(1) The airplane must be in the landing configuration.
(2) A stabilized approach, with a calibrated airspeed of not less
than VREF or VMCL, whichever is greater, must be
maintained down to the 50 foot height. VREF may not be less
than--
(i) 1.03 VSR0;
(ii) 1.20 VSR0PWR with the operative engines at the
power or thrust setting for approach at the reference flight path
angle;
(iii) The airspeed that provides an angle-of-attack margin to stall
for not less than a 20 knot equivalent airspeed vertical gust with all
engines operating at the power or thrust setting for approach at the
reference flight path angle;
(iv) The airspeed that provides an angle-of-attack margin to stall
for not less than a 15 knot equivalent airspeed vertical gust with the
critical engine inoperative at the power or thrust setting for approach
at the reference flight path angle; and
(v) A speed that provides the maneuvering capability specified in
paragraph (k) of Special Condition No. 1.
(3) Changes in configuration, power or thrust, and speed, must be
made in accordance with the established procedures for service
operation.
(4) The landing must be made without excessive vertical
acceleration, tendency to bounce, nose over, ground loop, porpoise, or
water loop.
(5) The landings may not require exceptional piloting skill or
alertness.
12. Landing Distances for Special Approaches to Short Field Landings
(a) In lieu of compliance with Sec. 25.125(a), the following
applies: The horizontal distance necessary to land and come to a
complete stop from a point 50 feet above the landing surface must be
determined (for each weight, altitude, wind, temperature, and runway
slope within the operational limits established for the airplane) as
follows:
(1) The airplane must be in the landing configuration.
(2) A stabilized approach, with a calibrated airspeed of not less
than VREF or VMCL, whichever is greater, must be
maintained down to the 50 foot height. VREF may not be less
than--
(i) 1.03 VSR0;
(ii) 1.20 VSR0PWR with the operative engines at the
power or thrust setting for approach at the reference flight path
angle;
(iii) The airspeed that provides an angle-of-attack margin to stall
for not less than a 20 knot equivalent airspeed vertical gust with all
engines operating at the power or thrust setting for approach at the
reference flight path angle;
[[Page 26922]]
(iv) The airspeed that provides an angle-of-attack margin to stall
for not less than a 15 knot equivalent airspeed vertical gust with the
critical engine inoperative at the power or thrust setting for approach
at the reference flight path angle; and
(v) A speed that provides the maneuvering capability specified in
paragraph (k) of Special Condition No. 1.
(3) Changes in configuration, power or thrust, and speed, must be
made in accordance with the established procedures for service
operation.
(4) The landing must be made without excessive vertical
acceleration, tendency to bounce, nose over, ground loop, porpoise, or
water loop.
(5) The landings may not require exceptional piloting skill or
alertness.
(b) In lieu of compliance with Sec. 25.125(b), the following
applies: For land planes, the landing distance on land must be
determined on level, smooth, dry and wet, hard-surfaced runways. In
addition--
(1) The pressures on the wheel braking systems may not exceed those
specified by the brake manufacturer;
(2) The brakes may not be used so as to cause excessive wear of
brakes or tires; and
(3) Means other than wheel brakes may be used if that means--
(i) Is safe and reliable;
(ii) Is used so that consistent results can be expected in service;
and
(iii) Is such that exceptional skill is not required to control the
airplane.
(4) The average touchdown rate of descent must not exceed 4 feet
per second and the approach flight path angle must be no greater than
-3 degrees for a normal approach.
(c) Procedures must be established by the applicant for use in
service that are consistent with those used to establish the
performance data under this special condition. These procedures must be
able to be consistently executed in service by crews of average skill,
and must include, as applicable, speed additives for turbulence and
gusts for approaches with all engines operating and with an engine
failure on final approach, and the use of thrust reversers on all
operative engines during the landing rollout.
(d) The procedures and performance data established under this
special condition must be furnished in the Airplane Flight Manual.
13. Thrust for Landing Climb
In lieu of compliance with Sec. 25.119(a), the following applies:
The engines at the power or thrust that is available eight seconds
after initiation of movement of the power or thrust controls to the go-
around power or thrust setting from the thrust level necessary to
maintain a stabilized approach at a flight path angle two degrees
steeper than the desired flight path angle.
Issued in Renton, WA on May 7, 1999.
John J. Hickey,
Acting Manager, Transport Airplane Directorate, Aircraft Certification
Service.
[FR Doc. 99-12361 Filed 5-17-99; 8:45 am]
BILLING CODE 4910-13-U