265 inoperative, and do not cause a smoke or fire hazard. Flight test recordings may be useful to verify laboratory demonstrations of system transients and electromagnetic interference (EMI) during normal operations.
(6) Section 25.1351(b)(5) - Power Source Disconnection .
(a) Verify that there are means accessible in flight, to the appropriate crewmembers, for the individual and collective disconnection of the electrical power sources from the system.
(b) Demonstrate that the airplane can be flown with one less generator than allowed by the MMEL. Also verify that prior to switching off the non-essential electrical loads specified in the AFM, the remaining generators’ short term (up to 5 minutes) capacity is not exceeded.
(7) Section 25.1351(b)(6) - Power Source Indicators .
(a) Verify that there are means to indicate, to the appropriate crewmembers, the generating system quantities essential for the safe operation of the system, such as the voltage and current supplied by each power source, if appropriate to type of aircraft, i.e., glass cockpit design may provide information on multi-function display (MFD), electronic centralized aircraft monitor, (ECAM), engine indicating and crew alerting system (EICAS), etc.)
(b) Verify proper operation of electrical power system warning, caution, and advisory indications, if any, and that these indications comply with §§ 25.1309(c) and 25.1322.
(8) Section 25.1351(c) - External Power . If provisions are made for connecting external power to the airplane, and that external power can be electrically connected to equipment other than that used for engine starting, verify that a means is provided to ensure that no external power supply having a reverse polarity, a reverse phase sequence, overvoltage, or reverse phase/neutral can supply power to the airplane’s electrical system. This demonstration may be accomplished in a laboratory environment.
(9) Section 25.1351(d) - Operation Without Normal Electrical Power . The applicant must demonstrate that the airplane can be operated safely in visual flight rules (VFR) conditions, for a period of not less than five minutes, with the normal electrical power (electrical power sources excluding the battery) inoperative, with critical type fuel (from the standpoint of flameout and restart capability), and with the airplane initially at the maximum certificated altitude.
(10) Emergency Electrical Power System.
(a) For airplanes with a type certification basis that includes amendment 25-23, the FAA has applied a general policy that following total failure of the electrical power system, if not shown to be extremely improbable, emergency electrical power should be available to power instrument displays, systems, equipment, or parts of the airplane essential for safety of
266 flight during IMC operations for at least 30 minutes. It may also be advisable to check that in the event when normal operation is reestablished following the use of emergency power devices, such devices can withstand, or at least do not adversely affect, the subsequent reversion to “normal” operation.
1 Thirty (30)-Minute Requirement Te st. Accomplish ground and/or flight testing to verify that the main battery/non time-limited power source (e.g., auxiliary power unit (APU), ram air turbine, pneumatic or hydraulic motor, etc.) is sufficient to power the essential loads required for safety of flight during IMC operations for at least 30 minutes. If the emergency power is provided by a non time-limited power source, the main battery capacity should be sufficient for at least 5 minutes to show compliance with § 25.1351(d). Essential loads for safety of flight during IMC operations are considered to include the following:
(aa) Those essential for continued safe flight and landing during visual meteorological conditions (VMC) operations.
(bb) One display of attitude, direction, airspeed and altitude as specifically stated in § 25.1333(b), a free-air temperature indicator, and pitot/static heat capability, if required.
(cc) Communication, intercommunication, and navigation capability (for IMC operations) necessary cockpit and instrument lighting, and any other instrument displays, systems, equipment, or parts of the airplane that are necessary to safely complete the flight.
2 Perform a VMC approach and landing with the airplane powered with emergency electrical power only. Verify acceptable operation of the emergency equipment.
3 Perform a simulated night IMC approach and landing with the airplane powered with emergency electrical power only. Verify acceptable visibility and operation of the emergency equipment. This test can be accomplished on a simulator.
(b) (Time-Limited Emergency Electrical Power Supply.) [Reserved]
187. Electrical Equipment and Installations - § 25.1353.
a. Explanation . None.
b. Procedures .
(1) Section 25.1353(a) - Equipment Interference . Verify by design analysis, laboratory demonstration, or airplane functional tests that the operation of each electrical unit, or system of units, will not adversely affect simultaneous operation of any other unit or system essential to the safe operation of the airplane.
(2) [Reserved]
(3) Section 25.1353(c)(1) - Storage Batteries . Verify, for all storage batteries, that during any probable charging and discharging condition, including when the battery is recharged after previous complete discharge, the cell temperatures and pressures of the batteries are safely maintained:
(a) During maximum regulated voltage or power;
(b) During a flight of maximum duration; and
(c) Under the most adverse cooling condition likely to occur in service.
(4) Section 25.1353(c)(2) - Compliance with § 25.1353(c)(1) . Compliance with § 25.1353(c)(1), above, must be shown by test, unless experience with similar batteries and installations has shown that maintaining safe cell temperatures and pressures is not a problem.
(5) Section 25.1353(c)(3) - Storage Battery Gases . Verify by design analysis, laboratory demonstration, or airplane functional tests that no explosive or toxic gases emitted by any battery in normal operation, or as a result of any probable malfunction in the charging system or battery installation, accumulate in hazardous quantities within the airplane.
(6) Section 25.1353(c)(4) - Equipment Adjacent to Storage Batteries . Verify by design review that no corrosive fluids or gases that may escape from the battery may damage surrounding airplane structures or adjacent essential equipment.
(7) Section 25.1353(c)(5) - Storage Battery Heat . Verify by design analysis, or laboratory demonstration, that each nickel-cadmium battery installation, has provisions to prevent any hazardous effect on structure or essential systems that may be caused by the maximum amount of heat the battery can generate during a short circuit of the battery or its individual cells.
(8) Section 25.1353(c)(6) - Storage Batteries Used for Engine Start . Verify by design analysis, laboratory demonstration, or airplane functional tests that any nickel-cadmium battery installation capable of being used to start an engine or auxiliary power unit has one or more of the following:
(a) A system to control the charging rate of the battery automatically to prevent battery overheating;
(b) A battery temperature sensing and over-temperature warning system with a means for disconnecting the battery from its charging source in the event of an over-temperature condition; or
(c) A battery failure sensing and warning system with a means for disconnecting the battery from its charging source in the event of battery failure.
188. Distribution System - § 25.1355.
a. Explanation . Section 25.1355(a) defines the distribution system, including the distribution busses, their associated feeders, and each control and protective device.
b. Procedures . Section 25.1355(c) - Independent Power Sources. For equipment or systems that are required to have two independent sources of electrical power, verify by airplane demonstration that in the event of a failure of one power source for such equipment or system, another power source (including its separate feeder) is automatically provided or can be manually selected to maintain equipment or system operation.
189. Circuit Protective Devices - § 25.1357. [Reserved]
190. Electrical System Tests - § 25.1363.
a. Explanation . Laboratory tests are conducted to verify that the control, regulation, protection, and automatic operation of the electrical system comply with §§ 25.1351 through
25.1357. Laboratory tests must be conducted on a mockup using the same type and length of feeder wires, and generating equipment that will be installed in the airplane. Additionally, the generator drives must simulate the actual prime movers on the airplane with respect to their reaction to generator loading, including loading due to faults. Section 25.1363(b) states that for each flight condition that cannot be adequately simulated in the laboratory or by ground tests on the airplane, a flight test must be conducted.
b. Procedures . Negative Acceleration Performance - Verify that no hazardous malfunctions of the electrical power system occur when the airplane is operated at the negative accelerations within the flight envelope prescribed in § 25.333. This should be shown for the greatest duration expected for the acceleration. This test is normally conducted in conjunction with the engine negative acceleration test to comply with § 25.943.
Section 4. Lights
191. Instrument Lights - § 25.1381.
a. Explanation . None.
b. Procedure . Under actual or simulated nighttime conditions, evaluate instrument lighting for the following characteristics:
(1) Illumination of adequate intensity is provided for all appropriate cockpit instruments/controls/equipment.
(2) Illumination is of appropriate color, intensity is evenly distributed, and is free of objectionable flicker, glare, or reflection.
(3) Dimming capability allows smoothly adjustable illumination intensity between appropriate limits.
192. Landing Lights - § 25.1383.
a. Explanation . None
b. Procedure . The landing gear lights should be evaluated to determine that they:
(1) Are aimed properly, provide sufficient intensity to facilitate night landings, and are acceptable at different pitch attitudes during landing approach with various cg, flap settings, and airspeed;
(2) Are not a source of objectionable glare or halation;
(3) Are functional in MMEL configurations and during operations in adverse weather.
193. Position Light System Installation - § 25.1385.
a. Explanation . Section 25.1385(a) requires the position light system to comply with detailed technical specifications contained in §§ 25.1387 through 25.1397. These sections define location, color, visibility, and intensity requirements with considerable precision. An in depth examination of the installed position light system with regard to these requirements is not necessarily considered appropriate or within the scope of a flight test evaluation, unless it becomes apparent while accomplishing the following procedures that further examination is warranted.
b. Procedures .
(1) Operate the position light system. Verify that general locations and color are as prescribed in §§ 25.1385(b) and (c).
(2) Verify that position light illumination does not cause objectionable glare to the flightcrew.
[194. - 200.] [Reserved]
201. Anti-Collision Light System - § 25.1401
a. Explanation . Section 25.1401(a)(2) requires the anti-collision light system to comply with the detailed specifications contained in § 25.1401(b) through (f). These sections define coverage, color, flash rate, and intensity requirements with some precision. An in depth examination of the installed anti-collision system, with regard to all of these requirements, is not necessarily considered appropriate or within the scope of a flight test evaluation, unless it becomes apparent while accomplishing the following procedures that further examination is warranted.
b. Procedure .
(1) Operate the anti-collision light system. Verify that one or more red or white lights are installed in locations that appear to provide the required visibility and which flash at the appropriate rates.
(2) Verify that the anti-collision lights do not cause objectionable glare to the flightcrew.
(3) Evaluate anti-collision lights during flight in clouds.
202. Wing Icing Detection Lights - § 25.1403.
a. Explanation . If the airplane is intended to be certificated for night flight into known icing conditions, a means for visually or otherwise determining the extent of icing on critical surfaces must be provided.
b. Procedure . The wing icing detection lights, if required, should be evaluated to determine that they:
(1) Are aimed properly toward the appropriate surfaces, and are of sufficient intensity for required illumination.
(2) Are not a source of objectionable glare, reflection, or halation.
Section 5. Safety Equipment
[203. - 207.] [Reserved]
Section 6. Miscellaneous Equipment
208. Electronic Equipment - § 25.1431.
a. Explanation . The possibility of interaction be tween two systems of communications and navigation equipment should be evaluated. Momentary deflection or flicker should be permitted. (Flicker frequencies above 55 Hz for stroke symbology or non-interlaced raster and 30/60 Hz for interlaced raster are generally satisf actory.) Loss of required function of the usable and assigned frequencies in the national airspace system should be considered unacceptable.
b. Procedures . None.
[209. - 223.] [Reserved]
Chapter 7 – Operating Limitations and Information.
Section 1. General [Reserved]
Section 2. Operating Limitations [Reserved]
Section 3. Markings and Placards [Reserved]
Section 4. Airplane Flight Manual
224. General - § 25.1581.
a. Explanation . The primary purpose of the AFM is to provide an authoritative source of information considered to be necessary for safe operation of the airplane. Since the flightcrew is most directly concerned with operation of the airplane, the language and presentation of the flight manual should be directed principally to the needs and convenience of the flightcrew, but should not ignore the needs of other contributors to safe operation of the airplane in accordance with the applicable operating regulations.
(1) Section 25.1501 requires that the operating limitations specified in §§ 25.1503 through 25.1533, and other information necessary for safe operation, be included in the AFM, be expressed in markings and placards, and also be made available by any other means that will convey the necessary information to the crew members.
(2) Information and data that are mandatory for an acceptable AFM are prescribed in §§ 25.1581 through 25.1587. The material required by 14 CFR parts 25 and 36 must be included in the AFM. At the option of the applicant, the AFM may be expanded to contain additional FAA approved information.
(3) The manufacturer or operator may include other “unapproved” data in a separate and distinctively identified portion of the AFM.
b. Reference . See AC 25.1581-1, “Airplane Flight Manual,” dated July 14, 1997, for detailed guidance on the required content and general structure of the AFM.
[225. - 227.] [Reserved]
272 Chapter 8 - Airworthiness: Miscellaneous Items
228. Design and Function of Artificial Stall Warning and Identification Systems.
a. Applicable Regulations . Sections 25.103, 25.201, 25.203, and 25.207.
b. Explanation . Some airplanes require artificial stall warning systems, such as stick shakers, to compensate for a lack of clearly identifiable natural stall warning to show compliance with the stall warning requirements of § 25.207. Similarly, some airplanes require a stall identification device or system (e.g., stick pusher, automatic inboard slat segment retraction, auto-trim, etc.) to compensate for an inability to meet the stalling definitions of § 25.201 or the stall characteristics requirements of § 25.203. In addition to compliance with the flight test requirements prescribed in paragraph 29 of this AC, certain system design and function criteria should also be addressed during the certification process of these airplanes. Included are system arming and disarming, preflight checks, failure indications and warnings, and system reliability and safety. The reliability of these systems can be evaluated in terms of the probability of the system not operating when required, and the safety aspects in terms of the probability of the system operating inadvertently. The required reliability and safety of stall warning and identification systems should be defined as a function of how critical their respective functioning is to safety of flight.
c. Arming and Disarming .
(1) Stall warning systems should be armed any time the airplane is in flight.
(a) Arming of stall warning systems has typically been accomplished by a ground/air logic circuit, which requires the nose and/or main gear squat switches to sense air mode before the system is armed. A pitch angle threshold during rotation has also been used to arm the stall warning system. These types of system arming schemes provide stall warning protection during liftoff and initial climb, where a stall would most probably have catastrophic consequences. They also provide protection against nuisance warnings during the takeoff roll, where the angle-of-attack (AOA) sensor vanes may be misaligned. Service history, however, has shown that systems armed around the liftoff point have caused pilots to abort takeoffs due to false alerts resulting from stall warning system faults or failures. In some cases, these highenergy rejected takeoffs have resulted in overruns. Therefore, system faults and failures that would lead to a false stall warning near liftoff should be made evident as early in the takeoff as practicable.
(b) In accordance with the requirements of § 25.207(b), if a stall warning system is required for any normal combination of flap and landing gear position, it must be used for all combinations of flap and landing gear positions. The purpose of this requirement is to provide a standard, consistent warning to the flightcrew of an operational flight envelope limit.
(2) Stall identification systems should be armed any time the airplane is in flight.
(a) The arming should take place automatically and may be provided by the same ground/air sensing system used for arming the stall warning system. The stall identification system may be inhibited during the takeoff rotation, but should become functional immediately after main gear liftoff. For airplanes with both stall warning and stall identification systems, it is permissible to have the stall identification system armed by operation of the stall warning system, provided the resulting probability of the stall identification not to operate when required is not greater than that specified in paragraph 228e, below.
(b) Stall identification systems may incorporate automatic disarming in flight regimes where the risk of stalling is extremely remote or where their unwanted operation would pose a threat to continued safe flight; examples of such inhibits would be high airspeed, and “g” cutouts (typically 0.5g), and while the pilot is following windshear recovery Flight Director guidance.
(c) A means to quickly deactivate the stall identification system should be provided and be made readily available to the pilots. It should be effective at all times, and should be capable of preventing the system from making any input to the longitudinal control system. It should also be capable of canceling any input that has already been applied, from either normal operation or from a failure condition.
(d) If a stall identification system is required to show compliance with the stall requirements of part 25 in one (or some) airplane configuration(s), it does not have to be used for stall identification in configurations where compliance can be demonstrated without it. Unlike stall warning, the stall point, be it aerodynamic or artificially induced, represents an end-point outside the in-service operational envelope of the airplane, and, therefore, does not need to be provided by the same means for all flap and landing gear configurations. Additionally, the added system complexity, and increased exposure to malfunctions and failures, would not warrant the use of a stall identification system for configurations where it is not required.
d. Indicating and Warning Devices .
(1) A method should be provided to adequately ascertain the proper operation of the stall warning and stall identification systems prior to takeoff. This method should be described in the operating procedures section of the AFM.
(2) Warning that the associated systems for operating the stall warning or stall identification devices has failed should be provided. As far as is practicable, this warning should cover all system failure modes.
(3) A clear and distinctive cockpit indication should be given when the stall identification system has been deactivated by the flightcrew (see paragraph 228c(2)(c), above).
This indication should be present as long as the system is deactivated.
(4) Any related limitations, and normal and emergency operating procedures, together with any information found necessary for safety during operation of the stall warning and
274 identification systems, should be included in the AFM and supplemented by such markings and placards as deemed necessary.
e. System Reliability and Safety . When stall warning and/or stall identification systems are installed to show compliance with the stalling requirements of §§ 25.201, 25.203, and
25.207, engineering data should be supplied to satisfy the following criteria, determined in accordance with the procedures of AC 25.1309-1A, “System Design Analysis,” dated June 21, 1988, (or later) where appropriate.
(1) Reliability . Probability of artificial stall warning and stall identification systems not operating when required:
(a) If stall warning is not clearly identifiable by natural characteristics, the loss of artificial stall warning should be improbable (not greater than 10-5 per flight hour). This reliability requirement is normally met by using dual, independent stall warning systems.
(b) If the natural stall characteristics are unacceptable, the combination of failure of the stall identification system to operate and entry into a stall should be extremely improbable (not greater than 10-9 per flight hour). A stall identification system with a failure rate not greater than 10-4 per flight hour will satisfy this requirement.
(c) If the natural stall characteristics are demonstrated to be benign, with the stall identification system disabled, a maximum failure rate of 10-3 per flight hour will be acceptable.
(2) Safety . Probability of artificial stall warning and stall identification systems operating inadvertently.
(a) The probability of inadvertent operation of artificial stall warning systems, during critical phases of flight, should not be greater than 10-5 per flight hour.
(b) To ensure that inadvertent operation of the stall identification system does not jeopardize safe flight, and to maintain crew confidence in the system, it should be shown that:
1 No single failure will result in inadvertent operation of the stall identification system; and
2 The probability of inadvertent operation from all causes is improbable (not greater than 10-5 per flight hour).
(c) Stall identification systems should be designed so that flight in turbulence will not result in inadvertent operation.
NOTE: In making the assessments of subparagraphs (d), (e) and (f), below, it should be assumed that in the climb, cruise, and descent flight regimes, corrective pilot action will not be initiated until three seconds after unwanted operation has been recognized. During takeoff and final approach, this time delay may be reduced to one second.
(d) If inadvertent operation of the stall identification system would result in limit loads being exceeded in any part of the airplane structure, the probability of inadvertent operation should not be greater than 10-7 per flight hour.
(e) If inadvertent operation of the stall identification system would result in ultimate loads being exceeded in any part of the airplane structure, the probability of inadvertent operation should be extremely improbable (not greater than 10-9 per flight hour).
(f) Inadvertent operation of the stall identification system should not cause catastrophic ground contact. This should be achieved by limiting the effect of the stall identification system to that necessary for stall identification purposes, without undue flight path deviation (e.g., by limiting the stroke of a stick pusher). Alternatively, if inadvertent operation could result in catastrophic ground contact, per § 25.1309(b)(1), the probability of inadvertent operation must be extremely improbable. Inhibition of the system close to the ground (e.g., for a fixed time after liftoff or below a radar altitude) would not normally be an acceptable means of compliance with this requirement.
f. System Functional Requirements .
(1) Operation of the stall identification system should reduce the airplane’s angle-ofattack far enough below the point for its activation that inadvertent return to the stall angle-ofattack is unlikely.
(2) The characteristics of stall identification systems, which by design are intended to apply an abrupt nose-down control input (e.g., stick pushers), should be such that it is unlikely that a member of the flightcrew will prevent or delay their operation.
(3) Normal operation of the stall identification system should not result in the total normal acceleration of the airplane becoming negative.
(4) The longitudinal maneuvering capability of an airplane equipped with stall identification systems, at all speeds likely to be encountered in normal operations, should be substantially the same as would be expected for an airplane with acceptable aerodynamic stall characteristics.
229. Reduced and Derated Thrust Takeoff Operations.
a. Explanation . The use of derated and reduced thrust for takeoff operations can produce substantial reductions in operating costs due to lower fuel consumption and increased operating margins. With the appropriate limitations and operating procedures applied, these operations can also offer safety benefits. Three methods have been approved by the FAA for derated and reduced thrust takeoff operations. These methods are as follows:
(1) Derated thrust approvals entail the use of completely new takeoff power setting charts and AFM performance information. The new power settings are less than the engine
276 manufacturer’s approved takeoff thrust settings, and the AFM performance is based on the thrust developed at these new, lower thrust settings.
(2) A constant reduced thrust increment can be used for all operating conditions. In this method the engine parameter by which thrust is set is reduced by a constant, such as an engine pressure ratio increment (EPR) of 0.02. A method is supplied to determine the airplane’s takeoff performance, at this reduced thrust level, from the AFM data representative of full takeoff thrust.
(3) The assumed temperature method of reduced thrust takeoff operations entails using takeoff performance information determined for an “assumed” temperature above ambient, but not above the temperature at which the takeoff weight would be limited by the takeoff field length available for a particular airport runway or by FAA climb requirements. Engine thrust settings, takeoff speed schedules, takeoff field lengths, and climb performance are determined at the assumed temperature. This assumed temperature method is the most flexible in its application and the most widely used by transport category airplane manufacturers and operators.
b. Procedures . Because there is a reduction in the takeoff performance level when any of the derated or reduced thrust methods are used, the FAA has certain limitations on the use of these reduced thrust levels. The appropriate guidance for derated and reduced thrust approvals, including limitations and procedures, is presented in AC 25-13, “Reduced and Derated Takeoff Thrust (Power) Procedures,” dated May 4, 1988.
230. Runway Gradients Greater Than ± 2 Percent.
a. Applicable Regulations . Sections 25.105, 25.115, 25.119, 25.121, 25.125, 25.1533, and
25.1587.
b. Explanation . The sections of part 25, referenced above, require accounting for the effects of runway gradient. Typically, performance limitations and information are determined for runway gradients up to ±2 percent in the AFM expansion of test data. Though these gradient extremes are adequate for addressing the majority of runways, there are a number of airports frequented by transport category airplanes that have runway slopes greater than ±2 percent.
Consequently, approvals have been granted for operations on runways with slopes exceeding ±2 percent with specific testing and analysis validation for the effects of the higher slopes.
Additional concerns, beyond runway slope effect on acceleration and braking and proper accounting of elevations during obstacle clearance analysis, include takeoff flare from liftoff to 35 feet, minimum takeoff climb gradients, minimum approach and landing climb gradients, landing flare distances, and unique operating procedures.
c. Procedures .
(1) Takeoff Flare from Liftoff to 35 Feet . The AFM expansion of the takeoff data should account for the effect of the runway slope on the portion of the takeoff distance after liftoff. At climb performance-limiting thrust-to-weight ratios, the average gradient of climb will