99-12361. Special Conditions: McDonnell Douglas Corporation (MDC) Model MD- 17 Series Airplanes

[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 [phone]; 
facsimile [phone]; Email: [e-mail].

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