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 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
(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.
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 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 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 center of gravity position.
(3) With the engines operating at maximum continuous 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 considered
as 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.
(2) A fuel tank that is designed not to feed the engines under all flight conditions need be tested only for the flight regime for which it is designed (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 attitudes (i.e., pitch and roll), 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 non-conservative 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 (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, adverse fuel feed/transfer configuration, etc.
110. Fuel System Hot Weather Operation - § 25.961.
a. Explanation .
(1) A flight test is normally necessary to complete the hot weather operation tests. If a ground test is performed, it should closely simulate flight conditions. If a hot fuel climb test is elected, hot weather fuel systems tests 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, as is allowed in the airplane flight manual, 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 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 (29oC). If tests are performed in weather cold enough to interfere with test results, steps should be taken to minimize the effects of cold temperatures. This may be accomplished by insulating fuel tank surfaces, as appropriate, 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 settings should be maintained at the maximum approved levels for takeoff and climb in accordance with § 25.961(a)(2). The airspeed during the climb should not exceed that speed used in demonstrating the requirements specified in § 25.961(a)(4). While § 25.961(a)(3) requires full fuel tanks for this test, experience has shown that 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. A greater rate of climb results in more air per unit time being released in the feed lines during the climb. Rate of climb is dictated primarily by airplane weight; however, airplane weight is a combination of payload and fuel state, and thus an analysis may be required to determine the critical combination of fuel quantity state and airplane weight that is operationally likely to occur.
(3) If the engines are normally operated with the auxiliary/emergency pumps “off,” they should remain “off” until fuel pressure failure occurs. Restoration of fuel pressure should
be noted and the climb continued to the maximum operating altitude selected by the applicant for certification. If a lower altitude is substantiated, appropriate limitations should be noted in the AFM.
(4) The tests should be conducted with the fuel system operating and configured normally, in accordance with the normal procedures outlined in the AFM. The following data should be recorded at reasonable time intervals:
(a) Fuel temperature in the tank
(b) Engine fuel pressure, measured at the engine/airplane interface, at the start of the test and during climb (note any pressure failure, fluctuation, or variation)
(c) Main and auxiliary/emergency fuel pump operation, as applicable
(d) Pressure altitude
(e) Ambient air temperature
(f) Airspeed
(g) Engine power setting and operating parameters (i.e., engine pressure ratio, gas generator speed, fan speed, exhaust gas temperature, fuel flow, etc.)
(h) Comments on engine operation
(i) Fuel quantities in fuel tanks
(j) Fuel grade or designation determined prior to test
(k) Airplane pitch and roll attitudes
(5) If significant fuel pressure fluctuati ons occur during testing of the critical flight conditions, but pressure failure does not occur, additional testing should be considered to determine that pressure failure may not occur during any expected operating mode. Also, the fuel system should be evaluated for vapor formation when switching from different fuel feed configurations, or at low fuel flow and idling approach and landing.
[111. - 116.] [Reserved]
117. Fuel Tank Vents and Carburetor Vapor Vents - § 25.975.
a. Explanation . Approval of Fuel Tank Vents (§ 25.975(a)).
(1) The tank venting arrangement must prevent siphoning of fuel during normal operation. Also the venting capacity and vent pressure levels must maintain acceptable differences of pressure between the interior and exterior of the tank during:
(a) Normal flight operation;
(b) Maximum rate of ascent and descent; and
(c) Refueling and defueling.
(2) No vent or drainage provision may e nd at any point where the discharge of fuel from the vent outlet would constitute a fire hazard, or where the discharge of fumes could enter personnel compartments.
(3) Each carburetor vent system must have means to avoid stoppage by ice.
b. Procedures .
(1) Tests should be conducted to ensure that no hazardous quantities of fuel will be siphoned overboard during any likely maneuvers encountered during normal operations.
Maneuvers that may require evaluation include, but are not necessarily limited to, the following:
(a) Taxi turns and turning takeoff maneuvers with fuel tanks filled to the maximum volume (below the required 2 percent expansion space) allowed by the type design, including consideration of tolerances in the volumetric shutoff level. Typically, left and righthand turns are conducted in a “figure eight” maneuver, followed by a maximum pitch and ascent rate takeoff;
(b) Maximum climb power ascent at high climb angles;
(c) Simulated turbulent air oscillations at or near the natural yawing and pitching frequency of the airplane;
(d) Rapid descent with high initial pitch-down rate;
(e) In-flight power-up turns; and
(f) Sideslip maneuvers on approach.
(2) The changes in tank secondary barrier cavity pressure during all airplane maneuvers, including emergency descent, should be accounted for in the design of the fuel tank.
Bladder type tanks may be critical under emergency descent conditions, depending on the cavity vent line sizing. Per § 25.975(a)(3), the vent/drain configuration must provide the required positive and negative pressure relief between the outer shell and the bladder or inner wall, to prevent collapse or over-expansion of the inner tank. Depending on the location of the overboard vent/drain exit and the airflow characteristics around the exit or exit mast, a flight test
may be required to evaluate the ascent and/or emergency descent characteristics of the cavity vent system with the airplane in both the “clean” and “wheels and flaps down” configuration.
(3) Verification that liquid discharge from the vent mast will flow clear of the airplane, not attaching itself to any airplane surface or re-entering any compartment of the airplane, may need to be accomplished by impingement tests conducted inflight. Small discharges of fuel from the fuel tank vent outlet have been acceptable, provided the fuel discharges clear of the airplane and does not result in siphoning of fuel from the tank. This can be accomplished using dyed fluid and/or coating the surfaces required to be free of impingement with powder compounds that will be washed away if contacted by liquid. If dyed water or other liquid is used, it may be necessary to add chemicals to prevent freezing during the test. Sufficient test maneuvers should be accomplished to ensure that impingement will not occur during any inadvertent discharge from the venting system.
(4) Carburetor vent systems may require flight testing to ensure against stoppage by freezing. Such tests can be conducted in conjunction with the tests required by § 25.1093 and/or § 25.1101 (see the “Procedures” for those sections in this AC).
[118. - 120.] [Reserved]
Section 3. Fuel System Components
[121. - 126.] [Reserved]
127. Fuel Jettisoning System- § 25.1001.
a. Explanation .
(1) Section 25.1001(a) prescribes the conditions governing the need for installation of fuel jettisoning systems; if an airplane can meet the climb requirements of §§ 25.119 and
25.121(d), at the weight existing after a 15 minute flight consisting of a maximum weight takeoff and immediate return landing, a fuel jettisoning system is not required. Credit is given for the actual or computed weight of fuel consumed in the 15 minute flight using the airplane configurations, power settings, and speeds appropriate to each flight segment.
(2) If a fuel jettisoning system is required, § 25.1001(b) prescribes the conditions that will determine the minimum flow rate of the system. Section 25.1001(b) requires the fuel jettisoning system to be capable of reducing the weight of the airplane, within 15 minutes of operation, from that specified in § 25.1001(a) to a weight at which the airplane will meet the climb requirements of §§ 25.119 and 25.121(d). Since the weight defined in § 25.1001(a) allows credit for a 15 minute fuel burn, a literal interpretation of this rule would result in a 15 minute jettisoning period beginning after a 15 minute takeoff, go-around, and approach flight. In application, the 15 minute jettisoning period will occur during a 30 minute flight in which weight reduction credit will be given for the fuel consumed and jettisoned. The airplane must be able to meet the specified climb requirements at the weight existing at the end of this 30 minute flight.
(3) Airplanes should also be investigated for other elements that may limit their ability to safely accomplish an immediate return landing without a fuel jettisoning system. Advances in wing design and propulsion technology have resulted in transport category airplane designs that can take off at weights considerably above their maximum landing weights. Many of these airplanes are capable of meeting the climb requirements of §§ 25.119 and 25.121(d), following a 15 minute flight, without a fuel jettisoning system. Some of these airplanes, however, may not be capable of landing without exceeding other certification limits such as maximum brake energy, landing distance, and tire speed. This is particularly true when non-normal procedures, implemented as a result of failures that have been shown to be foreseeable events, call for reduced flap settings and increases of as much as 30 knots, for a given weight, over speeds associated with the normal landing flap setting. Margins to flap placard limit speeds and flap load-relief activation speeds should be established and maintained for non-normal configurations that may be used in immediate return landings.
(4) An additional consideration that is representative of actual operating conditions is the ability to perform a go-around from field elevation with the flaps in the approach position and the landing gear down. Through compliance with § 25.1001(b), assurance will be obtained that the airplane can accomplish an all-engines-operating balked landing go-around, with normal landing flaps, followed by a one-engine-inoperative climb-out with approach flaps and landing gear up. However, non-normal procedures generally call for one-engine-inoperative landings to be made with the flaps in the position used to show compliance with the approach climb requirements of § 25.121(d). It should therefore be determined under what combinations of weight, altitude, and temperature the airplane can establish a positive rate-of-climb with oneengine-inoperative and the other operating at go-around thrust, with the flaps in the appropriate go-around position and the landing gear down.
b. Procedures . The basic purpose of these tests is to verify that the minimum jettisoning rate will allow the airplane to safely execute an immediate return landing, and to determine that the required amount of fuel may be safely jettisoned under reasonably anticipated operating conditions within the prescribed time limit, without danger from fire, explosion, or adverse effects on the flying qualities.
(1) Jettisoning Rate
(a) In determining the minimum jettisoning rate, the tanks, tank combinations, or fuel feed configurations that are critical should be selected for demonstrating the flow rate.
(b) It should be determined if airplane attitude or configuration has an effect on the jettisoning rate.
(c) It should be demonstrated that the means to prevent jettisoning of the fuel in the tanks used for takeoff and landing, below the level to meet the requirements of § 25.1001(e) and (f), are effective.