(f) If an onboard APU is required to provide power assist for inflight engine restart, the minimum APU start reliability should be determined, and its operation evaluated by flight tests in critical portions of the flight envelope.
(g) If start cartridges are used to provide restart capability, they should be of sufficient capacity (or number) to provide at least two start attempts for each engine.
(h) The inflight engine restart envelope presented in the AFM should identify a “core windmill relight envelope” that will provide engine ignition within 30 seconds and stabilized idle within 90 seconds of restart initiation. A larger envelope, which includes appropriately labeled longer restart times, may be allowed if it can be shown that indication of a clear progression of engine start is provided to the flightcrew. The AFM should also contain the restart procedures appropriate to each phase of flight (i.e. low altitude “hot start,” high altitude “cold soaked” start, etc.).
91. Automatic Takeoff Thrust Control System (ATTCS) - § 25.904.
a. Explanation.
(1) Beginning in the 1970’s, some manufacturers of turbojet powered airplanes elected to equip their airplanes with engine thrust control systems that automatically increased the thrust on the operating engine(s) when any engine failed. A similar system, referred to as an Automatic Takeoff Thrust Control System, was subsequently installed on turbopropeller airplanes.
(2) Takeoff performance credit was granted for ATTCS based upon prescribed system functional and reliability requirements, and performance related restrictions (e.g., initial takeoff power or thrust not less than 90 percent of that set by the ATTCS).
(3) These systems represented “novel or unusual design features” not adequately addressed by the requirements of part 25 at the time. Consequently, the airworthiness requirements for the early ATTCS certifications were prescribed in special conditions in accordance with § 21.16. The regulatory and technical content of those special conditions was added to part 25 as § 25.904 and appendix I by Amendment 25-62.
b. Procedures. Certification of an airplane with an ATTCS requires flight test demonstration of certain performance and functional aspects of the system, as outlined below:
(1) In order to comply with the part 25 airplane performance requirements, as required by Appendix I, I25.3(c), takeoff speeds, as limited by VMCG and VMCA, must reflect the effect of ATTCS operation following failure of the critical engine. It is permissible to publish two sets of takeoff performance data: one for ATTCS unarmed and one for ATTCS armed. In such cases, the AFM limitations, operating procedures, and performance information should clearly differentiate between the two sets of performance data.
(2) Engine operating characteristics should be investigated during operation of the ATTCS, and with the operating engines at the steady-state maximum power or thrust level achieved following operation of the ATTCS. (Refer to the guidance material provided in paragraph 99 of this AC.)
(3) Manual override of the ATTCS should be verified in flight test. This capability has been provided by the ability of the pilot to push the throttle levers to a higher power or thrust setting, for airplanes that use less than “full-throttle” for takeoff, and by activating an override switch for “firewall-type” fuel control systems. The override switch must be located on or forward of the power or thrust levers, and it must be easily accessible to the pilot’s hand that normally controls the power lever or thrust position in accordance with Appendix I, I25.5(b)(2).
It should also be demonstrated that the thrust/power level can be manually decreased at any time following ATTCS operation.
(4) A critical time period must be determined during which the probability of concurrent engine and ATTCS failure must be shown to be extremely improbable (not greater than 10-9 per flight hour). This critical time period is defined in Appendix I to part 25 as being from one second before the airplane attains V1, to a time where the actual takeoff flight path (i.e., no gradient reductions), following a simultaneous engine and ATTCS failure, would intersect the normal (i.e., engine failure at VEF and no ATTCS) one-engine-inoperative actual takeoff flight path at no less than 400 feet above the takeoff surface. The probability of failure of the automatic takeoff thrust control system, itself, must be shown to be improbable (not greater than 10-5 per flight hour).
(5) Performance credit for an operating ATTCS is not to be taken when operations are conducted using reduced takeoff power or thrust methods. If the ATTCS is armed during reduced power or thrust takeoffs, the relevant takeoff speeds should meet the required controllability criteria of part 25 at the power or thrust level provided by operation of the ATTCS. The applicant should demonstrate that the airplane has no adverse handling characteristics and the engine(s) do not exhibit adverse operating characteristics or exceed operating limits when the ATTCS operates.
(a) In accordance with § 25.1585, the AFM must contain information, instructions, and procedures, as required, regarding the peculiarities of normal and abnormal operations when scheduling reduced power or thrust operations with an armed ATTCS.
(b) Takeoff with an armed ATTCS is not restricted when airplane performance is based on an approved “derate” power or thrust rating that has corresponding airplane and engine limitations approved for use under all weight, altitude, and temperature (WAT) conditions.
92. Propellers - § 25.905. [Reserved]
93. Propeller Vibration and Fatigue - 25.907. [Reserved]
94. Propeller Clearance - § 25.925. [Reserved]
95. Propeller Deicing - § 25.929.
a. Explanation. None.
b. Procedures. If the propellers are equipped with fluid type deicers, the flow test should be conducted starting with a full tank of fluid and operated at maximum flow rate for a time period (15 min.) found operationally suitable. The operation should be checked at all engine speeds and powers. The tank should be refilled to determine the amount of fluid used after the airplane has landed.
96. Reversing Systems - § 25.933.
a. Turbojet Reversing Systems - § 25.933(a).
(1) Explanation.
(a) For reversers intended to be operable on the ground only, it must be shown that the airplane can be safely landed and stopped with a critical engine reverser deployed. In addition, if an undamaged reverser inadvertently becomes deployed in flight, it must be shown that it can be safely restored to a forward power or thrust position.
(b) For turbojet reversing systems intended for ground and/or inflight use, it must be shown that unwanted deployment of a critical reverser under normal operating conditions will not prevent continued safe flight and landing. Flight tests may be required to obtain aerodynamic data with the critical reverser deployed, to confirm that its deployment in the normal operating envelope will not be catastrophic to the airplane.
(2) Procedures.
(a) Turbine engine thrust reversers may be approved provided the following basic criteria are met:
1 Exceptional piloting skill is not required in taxiing, or in any condition in which reverse thrust is to be used.
2 Necessary operating procedures, operating limitations, and placards are established.
3 The airplane control characteristics are satisfactory with regard to control forces encountered.
4 The directional control is adequate using normal piloting skill. This is of particular importance for airplanes with aft-mounted engines, which may experience a loss of rudder effectiveness with reverse thrust. This may result in reductions in maximum reverse thrust levels (i.e., lower EPR or N1 settings) and/or increases in reverse thrust minimum
190 operational speeds. The stabilizing effect of the nose gear should also be investigated, particularly for the one-engine inoperative use of reverse thrust on wet runway surfaces.
5 A determination is made that no dangerous condition is encountered in the event of the sudden failure of one engine in any likely reverse thrust operating condition throughout the airplane’s approved operating envelope.
6 The operating procedures and airplane configuration are such as to provide reasonable safeguards against engine foreign object ingestion and serious structural damage to parts of the airplane due to the reverse airflow. This is normally accomplished by specifying in the limitations section of the AFM a minimum airspeed at which the thrust reversers must be retracted.
7 It is determined that the pilot’s vision is not dangerously obscured under normal operating conditions on dusty or wet runways and where light snow is on the runway.
8 For seaplanes, it is determined that the pilot’s vision is not dangerously obscured by spray due to reverse airflow under normal water operating conditions.
9 The procedure and mechanisms for reversing should provide a reverse idle setting such that, without requiring exceptional piloting skill, at least the following conditions are met:
(aa) Sufficient power or thrust is maintained to keep the engine running at an adequate speed to prevent engine stalling during and after the reversing operation.
(bb) The engine does not overspeed or stall during and after the reversing operation.
(cc) The engine cooling characteristics should be satisfactory in any likely operating condition.
10 For airplanes equipped with thrust reversers intended for inflight use, the effect of non-deployment of a reverser (i.e., asymmetric deployment) on airplane controllability should be investigated.
(b) For the failed reverser demonstration tests, the following criteria are provided:
1 The landing with a reverser deployed should be conducted with a flap setting and an airspeed such that a landing can be accomplished safely and consistently. The conditions and operating procedures to use when the landing is made with the deployed reverser must be defined and incorporated in the AFM per § 25.1585(a)(2) or (a)(3), whichever is applicable.
2 The restow test should be conducted at a reasonable and safe altitude and at an airspeed where the airplane can be safely controlled (approximately 200 knots). A
191 procedure should be developed so that the reverser can be restowed (if undamaged) safely, without causing unacceptable airplane controllability problems. The restowing procedure, airspeed, and airplane flight controls configuration should be provided in the AFM.
(3) Aircraft Backing Using Reverse Thrust. Limited operational approvals have been granted for the use of thrust reverser systems to back away from terminal gates in lieu of a tug pushback. These approvals are granted by the cognizant FAA Flight Standards office for each operator, the approval consisting of an amendment to the operator’s operations specifications identifying the airplane type, the airport, and the specific gates at that airport at which reverse thrust backing may be used. Though reverse thrust backing is not specifically an airworthiness approval item, there are certain areas of concern that overlap airworthiness and operations.
(These items are addressed in paragraph 240 of this AC.)
b. Propeller Reversing Systems - § 25.933(b).
(1) Explanation. None.
(2) Procedures. Reverse thrust propeller installations may be approved, provided the following is acceptable:
(a) A reliable means for preventing the inflight selection of a power setting below flight idle is provided.
(b) Exceptional piloting skill is not required in taxiing or in any condition in which reverse thrust is to be used.
(c) Necessary operating procedures, operating limitations, and placards are established.
(d) The airplane control characteristics are satisfactory with regard to control forces encountered.
(e) The directional control is adequate using normal piloting skill.
(f) A determination is made that no dangerous condition is encountered in the event of sudden failure of one engine in any likely operating condition.
(g) The operating procedures and airplane configuration are such as to provide reasonable safeguards against serious structural damage to parts of the airplane due to the direct effects of the reverse airflow or any resultant buffeting.
(h) It is determined that the pilot’s vision is not dangerously obscured under normal operating conditions on dusty or wet runways and where light snow is on the runway.
(i) It is determined that the pilot’s vision is not dangerously obscured by spray due to reverse airflow under normal water operating conditions with seaplanes.
(j) The procedure and mechanisms for reversing should provide a reverse idle setting such that without requiring exceptional piloting skill at least the following conditions are met:
1 Sufficient power is maintained to keep the engine running at an adequate speed to prevent engine stalling during and after the propeller reversing operation.
2 The propeller does not exceed the approved speed limit of 14 CFR part 35 or the airplane-manufacturer-declared propeller speed limit during and after the propeller reversing operation.
3 This idle setting does not exceed 25 percent of the maximum continuous rating.
(k) The engine cooling characteristics should be satisfactory in any likely operating condition.
97. Turbojet Engine Thrust Reverser System Tests - § 25.934. [Reserved]
98. Turbopropeller-Drag Limiting Systems - § 25.937.
a. Explanation. Approval of Automatic Propeller Feathering Systems. All parts of the feathering device that are integral with the propeller, or attached to it in a manner that may affect propeller airworthiness, should be considered from the standpoint of the applicable provisions of part 35. The determination of the continuing eligibility of the propeller under the existing type certificate, when the device is installed or attached, will be made on the following basis:
(1) The automatic propeller feathering system should not adversely affect normal propeller operation and should function properly under all temperature, altitude, airspeed, vibration, acceleration, and other conditions to be expected in normal ground and flight operation.
(2) The automatic device should be demonstrated to be free from malfunctioning which may cause feathering under any conditions other than those under which it is intended to operate. For example, it should not cause feathering during:
(a) Momentary loss of power or thrust; or
(b) Approaches with reduced throttle settings.
(3) The automatic propeller feathering system should be capable of operating in its intended manner whenever the throttle control is in the normal position to provide takeoff power or thrust. No special operations at the time of engine failure should be necessary on the part of the crew in order to make the automatic feathering system operative. (See also §§ 25.101,
25.111, 25.121 and 25.1501.)
(4) The automatic propeller feathering system installation should be such that not more than one engine will be feathered automatically, even if more than one engine fails simultaneously.
(5) The automatic propeller feathering system installation should be such that normal operation may be regained after the propeller has begun to feather automatically.
(6) The automatic propeller feathering system installation should incorporate a switch, or equivalent means, by which to make the system inoperative.
b. Procedures.
(1) Propeller Feathering System Operational Tests.
(a) Tests should be conducted to determine the time required for the propeller to change from windmilling (with the propeller controls set for takeoff) to the feathered position at the takeoff safety speed, V2.
(b) The propeller feathering system should be tested to demonstrate non-rotation or as a minimum, non-hazardous rotation at up to 1.2 times the maximum level flight speed, with one engine inoperative or the speed employed in emergency descents, whichever is higher, with:
1 The critical engine inoperative;
2 Wing flaps retracted;
3 Landing gear retracted; and
4 Cowl flaps (if applicable) closed.
(c) A sufficient speed range should be covered to assure that the propeller feathering angle, established on the basis of the high speed requirement, will not permit hazardous reverse rotation at the lower speeds. In addition, the propeller should not inadvertently unfeather during these tests.
(d) In order to demonstrate that the feathering system operates satisfactorily, the propeller should be feathered and unfeathered at the maximum operating altitude established in accordance with § 25.1527. The following data should be recorded:
1 Time to feather the propeller at the one-engine-inoperative cruising speed;
2 Time to unfeather the propeller to the minimum declared governing speed at maximum operating altitude (note that some driftdown may occur) and one-engine-inoperative cruising speed;
3 Altitude of propeller feathering tests; and
4 Ambient air temperature of propeller feathering tests.
(e) In order to demonstrate that the feathering system operates satisfactorily, the propeller should be feathered at the condition within the airplane operating envelope that is critical for the propeller.
99. Turbine Engine Operating Characteristics - § 25.939.
a. Explanation. The turbine engines of a transport category airplane must continue to operate safely during normal and emergency operation within the range of operating limitations of the airplane. Generally, compliance with § 25.939(a) can be determined to some extent while ascertaining compliance with other part 25 requirements, such as performance, controllability, maneuverability, and stall speed characteristics. Turbine engines should be stable in their operation and run free of adverse characteristics throughout the normal flight envelope. Certain adverse characteristics are allowed in specific flight regimes if they do not present a hazardous condition.
b. Reference. See AC 25.939-1, “Evaluating Turbine Engine Operating Characteristics,” dated March 19, 1986, for comprehensive guidance in the evaluation of turbine engine operating characteristics to show compliance with the requirements of § 25.939(a), only. The referenced AC does not provide guidance for compliance with § 25.939(c), which addresses inlet compatibility.
c. Inlet Compatibility - § 25.939(c).
(1) Explanation. Section 25.939(c) requires substantiation that the engine inlet systems on turbine engines not cause harmful vibration to the engine as a result of airflow distortion.
This should be verified in both static and transient power or thrust conditions.
(2) Procedures. Compliance with § 25.939(c) may require special instrumentation of the inlet itself, or the engine’s most critical component (i.e., fan blades). Inlet rakes permit the applicant to verify that the installed airflow distortion patterns are within the limits established by the engine manufacturer. In addition, accelerometer and/or strain gauge data could be acquired in flight tests to verify that vibration and stress level limits are not exceeded during operation in the normal flight envelope.
100. Inlet, Engine, and Exhaust Compatibility - § 25.941. [Reserved]
101. Negative Acceleration - § 25.943.
a. Explanation. Section 25.943 requires that no hazardous malfunction of an engine, APU, or any component or system associated with their operation, should result from airplane operation at negative accelerations. A hazardous malfunction in this case is considered to be one
195 that causes loss or sustained malfunction of the engine or APU, or improper operation of the engine accessories or systems. This requirement can be satisfied by flight test demonstrations that take into consideration the critical airplane, engine and APU configurations. The range of negative accelerations to be tested is prescribed by the flight envelope for the airplane, as defined in § 25.333. The duration of the negative acceleration excursions is intended to represent anticipated non-normal operational events such as atmospheric upsets, collision avoidance maneuvers, etc.
b. Procedures.
(1) In conducting negative acceleration tests, consideration should be given to engine accessory configurations, and critical levels of fuel and oil.
(2) Accelerations should be measured as close as practicable to the airplane’s c.g.
position.
(3) With the engines operating at maximum continuous power or thrust, and the APU operating with normal load (if flight operable), the airplane should be flown at a critical negative acceleration within the flight envelope. The duration of each test condition should be a minimum of 7 seconds between 0 and -1.0g, with a total accumulation of 20 seconds of negative acceleration operation.
(4) Test data should be analyzed with regard to maintaining adequate fuel flow to the engines and APU, and maintaining lubrication of critical components.
(5) Test planning should consider that sufficient altitude be available to conduct a suction feed relight in the unlikely event of an all-engine flameout, in which case the tank pump feed lines will become uncovered and air will enter each feed line.
102. Thrust or Power Augmentation System § 25.945. [Reserved]
Section 2. Fuel System
[103. - 108.] [Reserved]
109. Unusable Fuel Supply - § 25.959.
a. Explanation.
(1) The purpose of this test is to determine, for each fuel tank, the quantity of fuel that is not available to the engines, as specified in § 25.959. The unusable fuel quantity is the quantity of fuel that can be drained from the fuel tank sump with the airplane in its normal level ground attitude after a fuel tank unusable fuel test has been performed, plus the quantity remaining in the fuel tank (undrainable fuel).
(2) A fuel tank that is not designed to feed the engines under all flight conditions need be tested only for the flight regime for which it is designed to do so (e.g., cruise conditions).
Tanks that are not subject to aeroelastic effects of flight, such as wing bending or tank flexing, may have their unusable fuel quantity determined during a ground test. Suitable instructions on the conditions under which the tank may be used should be provided in the AFM. Other part 25 requirements, such as fuel flow (§ 25.955(a)(2)) and fuel quantity gauge calibration (§ 25.1553), are also related to the unusable fuel quantity. These other requirements must also be considered when the unusable fuel quantity is being determined for each fuel tank.
b. Procedures.
(1) The fuel system and tank geometry should be analyzed to determine the critical conditions for the specific tanks being considered (i.e., main and auxiliary or cruise tanks). The analysis should determine the amount of unusable fuel as a function of airplane pitch and roll attitudes, including those encountered when executing sideslips and dynamic maneuvers such as go-around pitch-up and acceleration. Particular attention should be directed toward the fuel tank or cell geometry and orientation with respect to the longitudinal axis of the airplane and location of the fuel tank outlets (i.e., fuel pump inlets or pickups). Care should be taken in planning how the critical attitude conditions are tested, so that the test procedure does not result in a nonconservative unusable fuel quantity.
(2) The term “most adverse fuel feed condition” is not intended to include radical or extreme conditions not likely to be encountered in operation. Judgment should be used in determining what maneuvers are appropriate to the type of airplane being tested. The test conditions should be selected using good judgment with regard to the kind of conditions the airplane under test will be subjected to in operation.
(3) Airplane attitude limitations may be used as a means of reducing the unusable fuel quantity, provided it is demonstrated that likely operational flight maneuvers can be accomplished with those attitude limits. Nose down pitch attitude should not be less than that for normal descent, approach, and landing maneuvers. Nose up pitch attitude consistent with a normal go-around condition, or a minimum of 10 degrees nose up, whichever is less, should be considered. Roll attitude limitations should not be less than that required to enter a normal traffic pattern, intercept the final approach course, and land with a 10-knot crosswind.
(4) After the most adverse fuel feed condition and the critical flight attitude have been determined for the specific fuel tanks being considered, the appropriate flight tests should be conducted. The flight testing should investigate the effects of the following:
(a) Steady state sideslips anticipated during operation with the airplane in both the approach and landing configurations.
(b) For those airplanes capable of high roll and pitch rates, abrupt maneuvers should be considered.
(c) A go-around condition at maximum acceleration and maximum rotation rate to the maximum pitch attitude should be considered.
(d) Effects of turbulence on unusable fuel quantity should be considered.
(e) If the airplane includes a low fuel quantity warning system, it should be demonstrated that the airplane can complete a go-around, approach, and return to landing, without fuel feed interruption, using the normal go-around pitch attitude; this should include goarounds accomplished with the aid of automated flight guidance systems.
(5) If airplane attitude limitations are employed to reduce the unusable fuel quantity, those attitude limitations must be published in the AFM (per § 25.1581(a)(2)) as limits for flight maneuvers after the low fuel warning light/message illuminates. This will provide assurance that the fuel remaining that is above the unusable quantity can be used without risk of fuel feed interruption to the engines. Flight tests should be conducted to confirm that the proposed pitch attitude limit:
(a) Is practical in terms of airplane flight characteristics for accomplishing a goaround; and
(b) Will not result in lift and drag characteristics that will increase the time and/or fuel quantity necessary to complete the go-around to a point where the fuel remaining is less than the unusable fuel quantity.
(6) If fuel pump failure en route would result in a significant reduction in usable fuel, the unusable fuel supply test should include a determination of this quantity. The effects of pump failure on the unusable fuel quantity should be presented in the AFM so that the flightcrew can take the reduction in usable fuel into account for flight planning purposes.
(7) Auxiliary fuel tanks and fuel transfer tanks designed or restricted for use during cruise flight only (i.e., not suitable for takeoff and landing) should be tested for unusable fuel quantity by appropriate investigation of the cruise environment. This should include reasonable turbulence levels, asymmetrical power or thrust, adverse fuel feed/transfer configuration, etc.
However, the unusable fuel quantity should not be less than fuel tank sump quantity.
110. Fuel System Hot Weather Operation - § 25.961.
a. Explanation.
(1) A flight test is normally necessary to complete the hot weather operation tests required by § 25.961(a). If a ground test is performed, § 25.961(b) requires that the ground test must closely simulate flight conditions. If a flight test is performed, the test should be conducted with hot fuel in the tanks normally used for takeoff and landing, and with the maximum number of engines drawing fuel from each tank, in accordance with the operating procedures provided in the AFM, to obtain the maximum anticipated fuel flow through the lines. In the case of symmetrical fuel tank systems, the tests may be confined to one of each such system. Critical
198 fuel that is unweathered or has not been exposed to long storage periods should be used during the tests. This ensures that the fuel has the maximum Reid vapor pressure (RVP) for which approval is requested. Fuel samples should be taken from the fuel tank prior to the test; for typical JET B (JP-4) type fuels, a minimum RVP of 3.0 has been required. The fuel temperature, just prior to takeoff, should be as close as practical to the maximum value for which operational approval is sought, but not less than 110º F as required by § 25.961(a)(5). If the fuel needs to be heated to this temperature, caution should be taken to prevent overheating during the process.
(2) The desirable outside ambient air temperature at the airport from which the tests are being conducted should be at least 85º F (29º C). If tests are performed in weather cold enough to interfere with test results, § 25.961(b) requires insulating fuel tank surfaces, fuel lines, and other fuel system components from cold air to simulate hot-day conditions. It should not be necessary to provide additional heat to the fuel after the original fuel sample is heated to temperature during the hot weather tests. However, if the fuel is used as the cooling medium for any heat exchangers, the maximum heat available to the fuel should be considered.
(3) If auxiliary pumps are being considered for use as emergency pumps, they should be inoperative. This test may be used to establish the maximum pressure altitude for operation with these pumps off. A fuel pressure failure is considered to occur when the fuel pressure decreases below the minimum prescribed by the engine manufacturer, or the engine does not operate satisfactorily.
b. Test Procedures and Required Data.
(1) The fuel temperature, just prior to takeoff, should be as close as practical to the maximum value for which operational approval is sought. If heating of the fuel is required, a takeoff and climb should be made as soon as possible after the fuel in the tank has been heated to avoid cooling of the fuel.
(2) Power or thrust settings should be maintained at the maximum approved levels for takeoff and climb in accordance with § 25.961(a)(2). Section 25.961(a)(3) requires the weight of the airplane to be the weight with full fuel tanks, minimum crew, and the ballast necessary to maintain the center of gravity within allowable limits. The airspeed during the climb should not exceed that speed used in demonstrating the requirements specified in § 25.961(a)(4). The combination of power or thrust, airplane weight and climb airspeed establishes the climb rates that must be considered when determining the critical conditions for testing. The airplane fuel load and rate of climb are the critical parameters for this test. Although the weight of the airplane must be based on full fuel tanks, experience has shown that full fuel tanks may not always be the most critical fuel load. Generally, full tanks result in the maximum head pressure over the tank pumps. A lesser tank quantity results in a lesser head pressure that may be more critical to the formation of fuel vapor in the pumps and fuel system lines. A greater rate of climb results in more air per unit time being released in the feed lines during the climb. An analysis may be needed to determine the critical combination of fuel quantity state and airplane climb rate that is operationally likely to occur. The test should be conducted at this critical combination of fuel quantity state and airplane climb rate with the airplane loaded to the weight and flown within the range of climb airspeeds at the power or thrust required by the regulation. NOTE: