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

DEPARTMENT OF TRANSPORTATION

Federal Aviation Administration

14 CFR Part 25

[Docket No. NM156, Special Conditions No. 25-151-SC]

Special Conditions: McDonnell Douglas Corporation (MDC) Model MD-

17 Series Airplanes

AGENCY: Federal Aviation Administration, DOT.

ACTION: Final special conditions.

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SUMMARY: These special conditions are issued for the McDonnell Douglas

Corporation Model MD-17 airplane. This airplane incorporates 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. These special conditions contain the

additional safety standards that the Administrator considers necessary

to establish a level of safety equivalent to that provided by the

existing airworthiness standards.

EFFECTIVE DATE: December 30, 1999.

FOR FURTHER INFORMATION CONTACT: Gerry Lakin, Project Officer, FAA

Transport Airplane Directorate, Standardization Branch, ANM-113, 1601

Lind Avenue SW., Renton, WA 98055-4056; telephone (425) 227-1187;

facsimile (425) 227-1149; Email: [email protected].

SUPPLEMENTARY INFORMATION:

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 these special conditions.

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 Public Law 92-574, the ``Noise Control

Act of 1972.''

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

[[Page 66724]]

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 electronically 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 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

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 electronically 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 and a throttle-coupled reduction in spoiler deflection 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

[[Page 66725]]

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 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

[[Page 66726]]

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

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 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.

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

[[Page 66727]]

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.

Special Condition 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 adopted 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. . . .'' Section

25.1303 does not adequately address the MD-17 HUD's novel and unique

location for a primary flight display, which is above the instrument

panel and in the field of view of the forward window.

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.

Therefore, in addition to compliance with Sec. 25.1321(b), the

special condition requires 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, the optical characteristics of HUD's 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 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

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

[[Page 66728]]

conformal symbolic information as primary flight information.

Therefore, the 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

display limitations of a monochrome 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. A monochrome 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. 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 2 is adopted 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

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 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

[[Page 66729]]

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 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 3 requires 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 adopted.

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

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 of 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

Use of the EFCS also affects the maneuvering capability of the MD-

17, which is not adequately addressed by the current part 25 design

maneuver requirements. 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, is 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.

[[Page 66730]]

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. Special condition 6 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 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 6 is adopted.

Flight Characteristics

7. Flight Characteristics Compliance via Handling Qualities Rating

Method

The EFCS 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 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 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 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 7 is adopted 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 8

ensures that the MD-17 has suitable static longitudinal stability in

any condition normally encountered in service. The HQRS prescribed by

Special Condition 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

[[Page 66731]]

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 9

is adopted 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 7 may be used to make this

assessment.

10. Control Surface Awareness

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. 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.

As a result of these concerns, Special Condition 10 is adopted to

require 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 adopted 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 approach for the MD-17 is

defined as an approach flight path angle no steeper than -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 502,100 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 electronically linking

spoilers and throttle movement provides much more precise flight path

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.

[[Page 66732]]

There are 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 applicable special

conditions; and (2) special approaches to short field landings that

will be conducted in accordance with existing part 25, a proposed SFAR

(to be published at a later date), and applicable 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 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

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 11 is adopted in conjunction with Special

Condition 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 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.

Crosswinds........................... ....................... Height at threshold....

Speed 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 no steeper than -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 special condition.

Special Condition 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 special

condition does not constitute operational approval

[[Page 66733]]

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 13 is

adopted. The thrust for a stabilized approach, including an appropriate

margin for operational safety, will be used as a basis for determining

the thrust available for the landing climb requirement. The initial

thrust level at the start of the 8-second spool-up time will be the

thrust for a stabilized approach at a flight path angle 2 degrees

steeper than the desired flight path angle. This thrust level will

account for thrust variations during the approach and conservatively

represents the initial thrust level.

This special condition is 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.

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.

Discussion of Comments

Notice of Proposed Special Conditions 25-99-04-SC for the McDonnell

Douglas Corporation Model MD-17 airplane was published in the Federal

Register on May 18, 1999 (64 FR 26900). Two commenters, including the

applicant, responded. Some of the comments were of an editorial or

clarifying nature and have been incorporated where appropriate. A

discussion of the remainder of the comments follows, corresponding to

the special conditions as proposed in Notice 25-99-04-SC.

General Comments

The commenter asks what the military certification basis is for the

Model MD-17 military version (the C-17), and states that it would be

interesting to compare it with the civil basis.

The C-17 was designed for the U.S. Air Force in accordance with the

design standards defined in the C-17 System Specification and C-17 Air

Vehicle Specification documents per the contractual agreement between

the company and the U.S. Air Force.

The specifications for C-17 power-augmented-lift performance speeds

include: (1) criteria for power-on minimum margins from stall speeds;

(2) angle-of-attack margins from stall expressed in terms of vertical

gust margins; and (3) maneuvering capabilities. These C-17 criteria and

the corresponding MD-17 criteria, which meet the applicable

airworthiness standards of part 25 and are discussed in the MD-17

special condition for power-augmented-lift, are similar or identical in

both the nature and magnitude of the required margins.

In the areas of flight controls and flying qualities, previously

existing military standards were invoked as part of the overall C-17

specifications. For instance, the flying qualities specifications were

a tailored revision of Mil-F-8785B. Similarly, for the MD-17, the FAA

adopted previous special conditions issued for other fly-by-wire

airplanes.

In summary, the MD-17 special conditions are similar to the

standards used for contractual acceptance of the C-17 by the U.S. Air

Force, but reflect the part 25 airworthiness standards and do not

include U.S. Air Force mission specific items.

The commenter would like to know more about the assumptions made

when thrust handling techniques were developed, and further states that

the technique proposed for flying the approach on the ``backside of the

drag curve'' is radically different than conventional airplanes, and

from airplanes on which most, if not all, civil pilots will have been

trained. The commenter is concerned that while such pilots may be able

to demonstrate sufficient proficiency during training, there is a real

risk that under certain conditions of high workload they may revert to

conventional flying techniques. The commenter believes that there

should be some safeguarding of the human factors aspects.

The thrust handling techniques for the backside approach for power-

augmented-lift aircraft were developed from flight simulator research

dating back to the 1970's. Test pilots from several regulatory

agencies, including the FAA and the U.K. CAA, participated in these

development tests. Test

[[Page 66734]]

findings are summarized in FAA Report No. FAA-RD-76-100, ``Progress

Toward Development of Civil Airworthiness Criteria for Powered-Lift

Aircraft,'' dated May 1976, a copy of which is in the docket for this

rulemaking. The results of this research indicate that the ease of

flying the backside approach and the capability to accurately hold

airspeed and flight path depend to a great extent on minimizing pitch

and thrust coupling. Minimizing airspeed changes as a result of thrust

changes and minimizing flight path angle changes as a result of pitch

changes not only allows more precise speed and path control, but also

provides better feedback to the pilot on the effect of the use of the

throttle and pitch controls. As noted in the preamble to the proposed

special conditions, the MD-17 design minimizes pitch and thrust

coupling.

Notwithstanding the research results noted above, and the MD-17

adherence to the design principles resulting from that research, the

FAA considered the potential for pilots to revert to the control

techniques used on conventional transport category airplanes to be a

major concern during the development of the special conditions. To

address this concern, the FAA interviewed U.S. Air Force reserve

pilots, flew simulator exercises, and reviewed the C-17 service

history.

The interviews with the U.S. Air Force reserve pilots were

considered to be especially valuable as many of these pilots also fly

as line pilots for major airlines, flying conventional transport

category airplanes ranging from older Boeing 727's to more modern

Boeing MD-11's. These pilots appeared to have little difficulty in

transitioning back and forth between the conventional airplanes and the

MD-17 with its unique characteristics. Training was essential for

introducing the backside technique, but after being exposed to the

differences in techniques in the simulator, reversion has not proven to

be a problem. The piloting cues and airplane response are significantly

different from those of a conventional transport category airplane,

which reinforces the use of the backside technique.

The simulator exercises flown by the FAA reinforced both the

conclusions of the earlier research efforts and the experiences of the

U.S. Air Force reserve pilots. The service history of the C-17 with the

U.S. Air Force has been very good, including experience under high

workload, high stress conditions. It should also be noted that the

Lockheed C-130 airplane, in the short takeoff and landing mode, also

flies on the backside and has had a good safety record.

1. Stall Speeds and Minimum Operating Speeds

The commenter states that credit over and above that already given

in the part 25 requirements is given for a reduced factor to obtain

V2 as a result of increased effects of power on stalling

speed, and asks if adequate stall margins will be available with the

use of reduced thrust/EPR techniques.

The credit for power-augmented lift for stall speed in the takeoff

phase of flight is based on the minimum power that exists at any point

in the takeoff flight path. This requirement includes consideration of

derated/reduced power techniques. The same speed margin will exist

between the power-on stall speed and the minimum takeoff safety speed

for a derated/reduced power takeoff as for a full power takeoff.

The commenter considers the justification for the reduction in

VREF resulting from a lower factor applied to a power-on

stall speed to be insufficient, at least until some operational

experience is gained. For example, one of the reasons given for using

the power-on stall speed is that there is no need to reduce thrust to

idle at any point during the approach. The commenter further states

that while this may be accurate, it is no guarantee that thrust will

never be reduced to idle (unless of course a physical movement

restriction is provided). The commenter asks how the probability of the

airplane being operated, albeit inadvertently, outside of the

certification assumptions has been considered within the special

conditions.

The FAA considered inadvertent speed and flight path excursions not

only due to piloting issues, but also due to environmental conditions

and other reasons. The requirements address each of these concerns by

providing margins for speed, angle-of-attack, thrust, and

maneuverability. Also, certain design features, combined with the

piloting cues and operating characteristics of the airplane, reduce the

probability of inadvertent and excessive thrust reduction, as well as

provide the capability for a quick recovery from both speed and flight

path excursions.

Minimal coupling between pitch and thrust reinforces the proper

operating techniques of using thrust to control flight path and pitch

to control airspeed. Targets for pitch angle, flight path angle, and

thrust level are displayed in the head-up primary flight display, along

with the current values to assist the pilot in making appropriate

control inputs. Large downward flight path changes are enabled through

the use of the Direct Lift Control, and rapid changes in the upward

direction are possible because of the separate spoiler bias design

feature.

The FAA considers the proposed margins provided at the reference

landing speed, VREF, to be adequate considering the specific

design features and operating characteristics of the MD-17.

Given that the probability of engine failure for part 25 airplanes

is generally assumed to be 1.0, the commenter asks what the

justification is for the 5-knot reduction in one-engine-inoperative

vertical gust margin based solely on the short exposure time in that

condition.

Although ensuring safe flight characteristics and performance

capability in the event of an engine failure is a fundamental principle

embodied in part 25, this does not mean that the probability of an

engine failure is generally assumed to be 1.0. The FAA continues to

consider a vertical gust margin of 15 knots is adequate to ensure

safety in the event of engine failure. For the normal all-engines-

operating condition, the FAA considers it appropriate to require a

larger margin, equivalent to the vertical gust margin typical of

conventional transport category airplanes operating at their minimum

landing approach speed. The commenter states that it is not clear how

the power is required to be set for the one-engine-inoperative power-on

stall speed demonstrations, and that in any case the thrust must be set

asymmetrically to simulate a realistic condition, rather than to have

thrust set symmetrically.

The power-on stall speeds for the MD-17 power-augmented-lift design

are influenced by which engines are operating. The distribution of the

engine efflux interacting with the externally blown flaps is different

for the all-engines-operating, outboard-engine-inoperative, and

inboard-engine-inoperative configurations. As a result, the power-on

stall speeds differ between engines-operating configurations for a

given weight and total airplane thrust level. Accordingly, the one-

engine-inoperative stall speeds for the C-17, the military version of

the MD-17, were determined from flight testing with asymmetric thrust.

In addition to the all-engines-operating configuration, the C-17

one-engine-inoperative power-on stall speeds were determined from

flight testing of both the outboard-engine-inoperative and inboard-

engine-inoperative configurations. The power-on stall speeds for these

different engine operating configurations were determined at airplane

thrust levels ranging from idle to takeoff thrust. For

[[Page 66735]]

test safety purposes, the one-engine-inoperative stall speeds were

determined from flight testing with a majority of the one-engine-

inoperative test points flown with the ``inoperative'' engine at idle

thrust and the remaining engines at the thrust level desired for a

particular test point. A smaller number of power-on stall test points

were flown for both the outboard-engine-inoperative and inboard-engine-

inoperative configurations with the ``inoperative'' engine shut down.

These test points provided a basis for correcting the majority of the

power-on stall speed test data, flown with the ``inoperative'' engine

at idle thrust, to a power-on stall speed level for the ``inoperative''

engine shut down. This same technique will be acceptable to the FAA for

showing compliance with Special Condition 1.

Another commenter points out that the last sentence of Special

Condition 1, paragraph (2)(i), ``Approach,'' defines a 2.7 percent

gradient of climb requirement without specifying the number of engines.

The commenter states that for consistency with the takeoff requirements

for gradient of climb [paragraph 1(h)], this should specify the

gradient of climb required based on the number of engines.

The inconsistency identified by the commenter was not intended by

the FAA when developing the special conditions. The FAA has tailored

these special conditions specifically for the MD-17, which is a four-

engine airplane. To correct this inconsistency, paragraph 1(h) of the

special condition has been revised to limit the applicability to a

four-engine airplane.

The commenter states that the preamble description of the spoiler

system may imply that the throttle-to-spoiler coupling is a mechanical

linkage, and believes that wording changes are needed to clarify that

the linkage is not mechanical.

The FAA agrees. The general discussion of the MD-17 design features

has been revised to provide clarification that the linkage is

electronic. Also, the discussion of ``Stall Speeds and Minimum

Operating Speeds'' has been revised to clarify that in addition to a

slight thrust increase to reduce the rate of descent at touchdown, ``a

throttle-coupled reduction in spoiler deflection'' may be used.

2. Head-Up Display (HUD) Used as Primary Flight Display (PFD)

One commenter considers the reliance on dual HUD's for the display

of primary flight information to be radical and in need of careful

attention, and further considers that better guidance is required for

the unusual attitude recovery training using HUD.

The FAA agrees that the use of the HUD as a primary flight display,

which includes its use by the pilots for unusual attitude recognition

and recovery, is a novel design feature. This is one of the key reasons

for the HUD special condition.

The FAA recognizes that unlike conventional primary attitude

displays, the HUD is monochrome and ``stroke-written,'' without the

contrast of color and shading found in conventional headdown attitude

displays. The FAA conducted a multiple expert opinion team study of the

C-17 HUD to explore this and several other factors related to its use

as a primary flight display. In addition, FAA test pilots flew several

unusual attitude recognition and recovery scenarios.

The special condition specifically requires that the HUD perform

the function of conventional color primary flight instruments and that

the flightcrew must be able to immediately recognize and perform a safe

recovery from unusual attitudes. One of many factors that the FAA must

evaluate is the ability of the monochrome HUD symbology to effectively

distinguish positive (sky) and negative (ground) pitch attitudes. The

FAA will carefully determine compliance with these requirements through

the use of flight test and simulation.

The commenter states that the preamble discussion of this special

condition seems to imply that the dual HUD installation is the novel

feature of the MD-17, and that it should emphasize the HUD as the

primary flight display (PFD) for each pilot, not just a dual HUD

installation.

The FAA considers that the current preamble discussion does, in

fact, adequately emphasize that these HUD's will be used as the primary

flight display. The fact that this is a dual-HUD installation is also

potentially significant, due to their location, depending on the

information content displayed and the concurrent use of both HUD's by

the flightcrew. The preamble discussion therefore remains unchanged.

The commenter requests that under the preamble discussion of the

``Arrangement and Visibility'' of the HUD, the second sentence be

revised to read, ``Section 25.1303 does not adequately address the MD-

17 HUD's location, and novel, unique features which allow the pilots to

keep their heads up and eyes out of the cockpit while viewing primary

flight data.'' The commenter states that this revision reinforces that

the MD-17 HUD installation deviates from the strict location

requirements of Sec. 25.1321(b) in order to enhance crew awareness

outside the cockpit.

The purpose of this discussion is to explain what is unique and

novel about the design that requires the special condition, not to

endorse potential advantages of the design. However, to address the

commenter's concern, the sentence in question has been revised to read,

``Section 25.1303 does not adequately address the MD-17 HUD's novel and

unique location for a primary flight display, which is above the

instrument panel and in the field of view of the forward window.''

The commenter requests that the last sentence of the preamble

discussion of the ``Arrangement and Visibility'' of the HUD be deleted,

stating that it is too vague and implies, too generally, that

additional data must be displayed on the HUD.

The FAA disagrees. This portion of the preamble discussion

describes the scope of, and need for, the kind of requirements

specified in the special condition. It is not meant to state the

specific requirements of the special condition that require compliance.

This discussion therefore remains unchanged.

The commenter requests that the first sentence of the preamble

discussion of the ``Compatibility with Other Cockpit Displays,'' be

rewritten as it implies that the MD-17 HUD has monochrome limitations

that other current HUD's would not have.

The FAA agrees and has revised this discussion accordingly.

The commenter further states that because the MD-17 monochrome HUD

represents current state-of-the art, it should not be made to sound as

if it is less than current technology. The commenter adds that this

requirement for HUD's to highlight certain information is important

only if a monochrome HUD is specifically used as a PFD.

This special condition applies only to the MD-17 HUD, not generally

to all HUD's. The FAA did not intend to imply that the MD-17 monochrome

HUD, alone, has limitations due to the lack of color. However, to

address the commenter's concern, the FAA has revised the preamble

discussion referred to by the commenter to state that a ``monochrome

HUD'' must have an equivalent means to unmistakably highlight and

distinguish the same information.

The commenter states that the wording of the last sentence of the

discussion of the ``Compatibility with Other Cockpit Displays'' which

reads,

[[Page 66736]]

``the existing rules do not adequately address the compatibility of

different display formats in the MD-17 cockpit'' implies that the MD-17

cockpit design has a display compatibility problem. The commenter

asserts that the MD-17 display formats were designed using human

factors design principles to be compatible with other cockpit displays,

and recommends that the phrase ``in the MD-17 cockpit'' be removed to

prevent potential misunderstandings of the Boeing display design

philosophy.

The FAA agrees and has revised the discussion accordingly.

The commenter recommends that the discussion of the ``Additional

Recommendations and Supporting Data'' be removed, stating that it

provides no additional or revised requirements, but simply collects

into one location part 25 requirements that the MD-17 must meet.

The FAA does not agree. The regulations and advisory material

referred to by the commenter are not part of the special conditions and

are not additional requirements. They are listed in the preamble

discussion for information only in the context of this special

condition. The FAA considers that they should be given special

attention due to the uniqueness of the HUD.

Further to the above discussion, the commenter states that the

discussion of Sec. 25.1301(a) digresses into a minimum equipment list

set of requirements, dictating which displays must be operative and how

displays must be used. The commenter considers these to be operational

issues that do not belong in this discussion.

The FAA disagrees. Unlike other transport HUD's, the MD-17 HUD's

are used as PFD's. Certain environmental light conditions significantly

affect the pilot's ability to use the HUD compared to headdown

instruments. In some of these conditions, the HUD cannot be relied on

as the PFD, so another PFD must be available. The safety objective is

to ensure that the flight has functional primary flight displays in all

foreseeable conditions. While it may also have MEL implications, this

requirement is stated for the sake of the design and functional

allocation of the display suite of the flight deck in which these HUD's

are installed.

The commenter states that paragraph (a)(2) of the special condition

is confusing. The first sentence allows for guidance to be displayed in

``close proximity'' to the HUD field of view, while the second sentence

begins ``Likewise'' and yet implies that the information must be

displayed on the HUD, not in close proximity. The commenter suggests

that the second sentence be revised to read ``Likewise, other essential

information and alerts that are related to displayed information and

may require pilot action must be displayed for instant recognition,

either on the HUD or in close proximity to the HUD field of view.''

The FAA agrees and has revised the special condition as proposed by

the commenter.

The commenter requests that the wording of the third sentence of

paragraph (a)(7) of the special condition be revised to state that the

HUD symbology must not ``excessively'' interfere with the pilot's

forward view, etc. The commenter's reason for the change is that

without the word ``excessively,'' a strict FAA interpretation might

require all HUD symbology to be removed so as not to interfere with the

pilot's forward view at all, thus defeating the intended purpose of the

HUD.

The word ``excessively'' was removed because the criteria for what

is and is not excessive were undefined at the time. This is a

compliance finding based on FAA flight test pilot judgement. The word

``excessively'' has been restored, as suggested by the commenter, with

an explanation added that ``interference would be considered excessive

if it prevents the pilot from seeing flight hazards, such as airborne

traffic, terrain, and obstacles, or outside visual references required

for safe operation such as approach lights, runway lights, runways, and

runway markings.''

The commenter notes that the term ``slowovers'' in paragraph

(a)(9)(ii) of the special condition is used when discussing autopilot

failures, and points out that the unique MD-17 fly-by-wire control

design is not subject to slowovers in the same way as conventional

designs.

The use of the term ``slowover'' was intended only as an example of

autopilot failures that may cause an upset; the emphasis is actually on

the upset. However, to avoid any confusion in this regard, the

reference to ``slowover'' has been removed and the words ``as

applicable to the MD-17 type design'' have been added in its place.

In paragraph (b)(5), the commenter recommends that the FAA maintain

the portion of the sentence that reads, ``There must be no adverse

physiological effects of long term use of the HUD system, such as

fatigue or eye strain'' and delete the remainder of that sentence and

the sentence that follows. The commenter maintains that the design of

the MD-17 is such that the pilot can always choose to use the head-down

PFD instead of the HUD while seated in a reclined position, and that

the HUD is not intended to be relied on as the sole PFD.

The FAA agrees with the change recommended by the commenter and has

revised the special condition accordingly.

The same commenter recommends that paragraphs (c)(2)(i) through

(viii) be removed, stating that these paragraphs impose a series of

safety requirements interpretations for hazards associated with loss or

erroneous display of parameters on the HUD and/or elsewhere in the

cockpit. The commenter further states that most of these

interpretations are already provided in AC 25-11, ``Transport Category

Airplane Electronic Display Systems.'' The commenter questions the

value of a special condition that applies criteria with which the MD-17

will comply using existing guidance.

The FAA disagrees that this information should be removed. Since

direct reference to the AC cannot be included in the text of the

special condition, the applicable criteria were inserted instead. This

does not change the requirements that were originally agreed to by the

applicant and imposes no additional burden.

The commenter states that the MD-17 HUD, by design, will not

display any data unless the combiners are fully deployed and aligned,

so the warning called out in paragraph (c)(4) of the special condition

is of little value. The commenter suggests revising this paragraph to

say that the HUD system must monitor the position of the combiner and

must not display conformal data that is hazardously aligned due to

combiner position. A suitable warning, alerting the crew of this

condition, is also acceptable.

The FAA agrees with the intent of the comment and has revised

paragraph (c)(4) accordingly.

4. Interaction of Systems and Structure

The commenter points out that a sentence appears to be missing from

the special condition.

The FAA agrees. The sentence the commenter is referring to concerns

the flutter clearance speeds that may be based on the speed limitation

specified for the remainder of the flight. The omission of this

sentence in the proposed special condition was an inadvertent

oversight, which has been corrected.

7. Flight Characteristics Compliance via Handling Qualities Rating

System

The commenter states that in order to determine whether the

airplane has suitable stability, objective requirements

[[Page 66737]]

are necessary against which to make the assessment. The commenter does

not consider the Handling Qualities Rating System to be an acceptable

alternative.

The FAA disagrees. The special conditions for flight

characteristics evaluation of the MD-17 are the same as those used on

other airplanes with similar fly-by-wire flight control systems. The

FAA Handling Qualities Rating System has been used successfully to

evaluate airplanes with fly-by-wire flight control systems since the

early 1980's.

11. Steep Approach Air Distance, and 12. Landing Distances for

Approaches to Short Field Landings

The commenter states that the intention to distinguish between

conventional and special approaches to short field landings is noted

and would be interested in reviewing the complete SFAR, which will

address short field operations, when it becomes available. The

commenter further states that there also needs to be a clear

distinction operationally between the two and asks, ``While there is a

clear upper limit on the steep approach angle (i.e., 5 degrees or 1,000

fpm), what will be the upper limit for a conventional approach?''

The commenter has misunderstood the proposals related to steep

approach operations and special approaches to short field landings.

There are 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. These two types of landing operations would be

distinguished by the additional equipment, training, and operating

requirements associated with approval to conduct special approaches to

short field landings. The applicant intends to provide steep approach

capability (allowing operators to seek steep approach approval) for

both types of landing operations.

The Steep Approach Air Distance special condition, which provides

an alternative methodology for determining the airborne part of the

landing distance, would apply only to those steep approaches flown as

part of a special approach to a short field landing conducted in

accordance with the proposed SFAR. In general, the FAA considers a

steep approach to be any approach conducted at angles steeper than 3.77

degrees. This value is derived from the normal 3 degrees approach path

angle, plus the outside limit for vertical displacement from the 3

degrees glide slope on the Instrument Landing System (ILS), as

established by the FAA Flight Standardization Board.

Another commenter notes that the steep approach air distance

definition in paragraph (a) of Special Condition 11 does not reflect

the specific distance between the runway threshold and the touchdown

aim point to be used in operation.

The FAA has revised the wording of this special condition to

provide the clarification requested by the commenter.

The same commenter notes that Special Conditions 11(a)(4) and

12(a)(4) refer to a ``water loop'' maneuver and questions whether this

maneuver has ever been demonstrated with a land-based airplane.

The FAA has determined that there is no need to consider this

maneuver for a land-based airplane such as the MD-17, and has removed

the reference to the water loop maneuver from both special conditions.

This commenter points out that there are several references to

approach flight path angle in both the preamble discussion of Special

Conditions 11 and 12, and in the text of paragraph 12(b)(4) of Special

Condition 12, that use a negative sign convention that could lead to

confusion.

The FAA agrees and has revised the wording accordingly.

With the exception of the changes discussed above, the special

conditions are adopted as proposed in Notice 25-99-04-SC.

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 Special Conditions

Accordingly, pursuant to the authority delegated to me by the

Administrator, the following special conditions are issued as part of

the type certification basis for 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

[[Page 66738]]

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] TR30NO99.001

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] TR30NO99.002

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 Sec. 25.107(a) and

Secs. 25.107(c) 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);

(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 1.7 percent 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.

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

Maneuvering

Configuration Speed Bank Angle Thrust Representative of

(degrees)

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

Takeoff................................ V2 30 Asymmetric WAT-Limited.\1\

[[Page 66739]]

Takeoff................................ V2+XX \2\ 40 All-engines operating

climb.\3\

En route............................... VFTO 40 Asymmetric WAT-Limited.\1\

Landing................................ VREF 40 Symmetric for approach at

the reference approach

flight path angle.\4\

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

\1\ A combination of Weight, Altitude and Temperature (WAT) such that the thrust or power setting produces the

minimum climb gradient specified in Sec. 25.121 for the flight condition.

\2\ Airspeed approved for all-engines-operating initial climb.

\3\ That thrust or power setting which, in the event of failure of the critical engine and without any crew

action to adjust the thrust or power of the remaining engines, would result in the thrust or power specified

for the takeoff condition at V2, or any lesser thrust or power setting that is used for all-engines-operating

initial climb procedures.

\4\ Thrust may be adjusted during the maneuver to maintain the reference approach flight path angle.

(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.

(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.

[[Page 66740]]

(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

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 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, either on the HUD or in close proximity to the HUD field

of view. 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

[[Page 66741]]

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 excessively interfere with the pilots' forward view,

ability to visually maneuver the airplane, acquire opposing traffic,

and see the runway environment. Interference would be considered

excessive if it prevents the pilot from seeing flight hazards, such as

airborne traffic, terrain, and obstacles, or outside visual references

required for safe operation such as approach lights, runway lights,

runways, and runway markings. 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, as applicable to the MD-17 type design; 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.

(i) 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

(ii) 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 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.

(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.

[[Page 66742]]

(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

must not display conformal data that is hazardously aligned due to

combiner position, without a warning to alert the crew of the

condition.

(5) The HUD system must be shown to comply with the high intensity

radiated fields certification requirements of Special Condition 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-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.

1a. 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.

[[Page 66743]]

(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.

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(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

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.

[[Page 66744]]

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Qj=(Tj)(Pj) where:

Tj=Average time spent in failure condition j (in hours)

Pj=Probability of occurrence of failure mode j (per hour)

Note: If Pj is greater than 10-3 per

flight hour, then a 1.5 factor of safety must be 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.

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BILLING CODE 4910-13-C

Qj=(Tj)(Pj) where:

Tj=Average time spent in failure condition j (in hours)

Pj=Probability of occurrence of failure mode j (per Hour)

Note: If Pj is greater than 10-3 per

flight hour, then a 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. Flutter clearance speeds V' and V''

may be based on the speed limitation specified for the remainder of the

flight, using the margins defined by Sec. 25.629(b).

[[Page 66745]]

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BILLING CODE 4910-13-C

V'=Clearance speed as defined by Sec. 25.629(b)(2).

V''=Clearance speed as defined by Sec. 25.629(b)(1).

Qj=(Tj)(Pj) where:

Tj=Average time spent in failure condition j (in hours)

Pj=Probability of occurrence of failure mode j (per hour)

Note: If Pj is greater than 10-3 per

flight hour, then the flutter clearance speed must not be less than

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.

(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

[[Page 66746]]

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 co

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