(iv) Effects of turbulence on unusable fuel quantity should be considered.
(v) 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 go- arounds accomplished with the aid of automated flight guidance systems.
(5) If airplane attitude limitations are employed to reduce the u nusable fuel quantity, those attitude limitations must be published in the AFM as limits for flight maneuvers after the low fuel warning light/message illuminates. This will provide insurance that the fuel remaining that is above the unusable quantity can be used without risk of fuel feed interruption to the engines. Flight tests must be conducted to confirm that the proposed pitch attitude limit:
(i) is practical in terms of airplane flight characteristics for accomplishing a go-around; and
(ii) 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 enroute 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 110° F as required by § 25.961(a)(5). If the fuel must 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 85 o F. 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 considere d 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 leve ls 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 w ith 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:
(i) fuel temperature in the tank
(ii) engine fuel pressure, me asured at the engine/airplane interface, at the start of the test and during climb (note any pressure failure, fluctuation, or variation)
(iii) main and auxiliary/emergency fuel pump operation, as applicable
(iv) pressure altitude
(v) ambient air temperature
(vi) airspeed
(vii) engine power setting and operating parameters (i.e., engine pressure ratio, gas generator speed, fan speed, exhaust gas temperature, fuel flow, etc.)
(viii) comments on engine operation
(ix) fuel quantities in fuel tanks
(x) fuel grade or designation determined prior to test
(xi) airplane pitch and roll attitudes
(5) If significant fuel pressure fluctuations 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:
(i) Normal flight operation;
(ii) Maximum rate of ascent a nd descent; and
(iii) Refueling and defueling.
(2) No vent or drainage provision may end 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:
(i) 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 right-hand turns are conducted in a “figure eight” maneuver, followed by a maximum pitch and ascent rate takeoff;
(ii) Maximum c limb power ascent at high climb angles;
(iii) Simulated turbulent air oscillations at or near the natural yawing and pitching frequency of the airplane;
(iv) Rapid descent with high initial pitch-down rate;
(v) In-flight power-up turns; and
(vi) Sideslip maneuvers on approach.
(2) The changes in tank secondary barrier cavity pressure during all airplane maneuvers, including emergency descent, must 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. 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 te sting 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 repres entative 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-enginesoperating 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 one-engine-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
(i) 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.
(ii) It should be determined if airplane attitude or configuration has an effect on the jettisoning rate.
(iii) 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.
(iv) It should be demonstrated that operation of the jettisoning system does not have a detrimental effect on operation of the engines (and the APU if installed and approved for inflight operation).
(2) Fire Hazard
(i) The fuel jettisoning flow pattern should be demonstrated from all normally used tank or tank combinations on both sides of the airplane, whether or not both sides are symmetrical.
(ii) The fuel jettisoning flow pattern should be demonstrated for the flight conditions specified in §§ 25.1001(d)(1), (2), and (3). Steady-state sideslips anticipated during operation should be conducted during flight conditions.
(iii) Fuel in liquid or vapor form s hould not impinge upon any external surface of the airplane during or after jettisoning. Colored fuel, or surface treatment that liquid or vaporous fuel changes the appearance of, may be used on airplane surfaces for detection purposes. Other equivalent methods for detection may be acceptable.
(iv) Fuel in liquid or vapor form should not enter any portion of the airplane during or after jettisoning. The fuel may be detected by its scent, a combustible mixture detector, or by visual inspection. In pressurized airplanes, the check for the presence of liquid or vaporous fuel should be accomplished with the airplane unpressurized.
(v) There should be no evidence of leakage after the fuel jettisoning valve is closed.
(vi) Testing should be conduct ed with the wing flaps in all available positions and during transition from each position to the next. If there is any evidence that wing control surface (flaps, slats, etc.) positions may adversely affect the flow pattern and allow fuel to impinge on the airplane, the airplane should be placarded and a limitation noted in the Airplane Flight Manual.
(3) Control.
(i) Changes in the airplane control qualities during the fuel jettisoning tests should be investigated, including asymmetrical jettisoning.
(ii) Discontinuance of fuel jettisoning should be demonstrated in flight.
(4) Residual Fuel. The quantity of usable fuel that cannot be jettisoned should be determined to meet the requirements of §§ 25.1001(e) or (f), as applicable. One acceptable means to show compliance with the requirement for sufficient fuel remaining to permit continued flight has been to drain the remaining fuel from the test tank(s) after landing; applicants may propose other means of compliance.
Section 4. OIL SYSTEM [RESERVED]
Section 5. COOLING [128. - 130.] [RESERVED]
Section 6. INDUCTION SYSTEM
131. AIR INDUCTION - § 25.1091
a. Explanation - § 25.1091(d)(2) . The turbine engines of transport category airplanes are susceptible to surge, stall, and flameout when excessive amounts of water are ingested. The certification requirements for turbine engines include the demonstration of a capability to operate in simulated rainfall with no adverse operating effects.
The quantities of water spray that may be directed toward the engine inlets, resulting from the airplane passing through standing water on taxi and runways, may exceed that used in the simulated rainfall ingestion testing. This becomes particularly important during takeoff, where the engines are operated at high power settings and the airplane will experience a wide range of speeds. During takeoff and landing ground rolls, the airplane's tires generate bow waves, side spray, and "rooster tails" (spray that is thrown off the tires as they rotate), which can collect into concentrated streams of water that, if ingested into the engines, Auxiliary Power Units (APU), or air conditioning systems, could affect operation sufficiently to cause an unsafe condition. Similarly, water ingestion into the airspeed system during ground operations may cause errant cockpit indications of airspeed. It is often practical to investigate these effects concurrently with the engine water ingestion evaluation.
b. References .
(1) Advisory Circular 20-124, Water Inge stion Testing for Turbine Powered Airplanes, dated September 30, 1985.
(2) Advisory Circular 91-6A, “Performance Information for Operation with Water, Slush, Snow or Ice on the Runway,” dated May 24, 1978.
c. Procedures .
(1) Method of Compliance . The applicant may show compliance with the requirements of §
25.1091(d)(2) by test or by reasonable analysis. The analytical approach may be acceptable in cases where an airplane and/or its engines were modified, but the overall geometry and configuration remained unchanged.
(2) Water Depth . Takeoffs should not be attempted when the depth of standing water, slush, or wet snow is greater than one-half inch over an appreciable part of the runway (Ref. AC 91-6A). Therefore, one-half inch of standing water is the accepted criteria for conducting tests to demonstrate compliance with the water ingestion requirements of § 25.1091(d)(2). Testing may be conducted in specially constructed water test beds, where the selected water depth is maintained over not less than 90 percent of the test bed area. If the airplane successfully completes testing with a one-half inch water depth, no operating limitations will be imposed for wet runway operations.
(3) Airplane Configuration . Testing should be conducted to demonstrate all critical phases of taxi, takeoff, and landing operations. High lift devices and landing gear doors should be in the position appropriate to each phase of operation tested. All portions of the airplane that may affect water spray patterns should be in the production configuration desired for approval (e.g., chine tires, fenders). The most critical powerplant operating configuration (bleed air, electrical load, ignition system, etc.) that is consistent with the applicant's recommended procedures for operation on wet runways should be used.
(4) Test Facility . The test facility must have adequate space for simulated takeoff and landing runs.
Intermediate dams may be used in the water test bed to maintain the selected water depth over not less than 90 percent of the test bed area. The test bed length should not be less than that required to produce a spray pattern of one second duration at critical test speeds. The test bed width need only accommodate one landing gear, if acceptable data are presented to the FAA that show the combined nose and maingear spray is not more hazardous than the spray produced by either gear separately.
(5) Test Procedures . The tests should be conducted with the appropriate thrust or power setting. To establish the critical speed ranges, test runs should be made through the water test bed in speed increments of not more than 20 knots. If the rotated airplane attitude is suspected to be critical, at least one run at V R should be conducted with the airplane rotated to the normal rotation pitch attitude. If thrust reversers or reversing propellers are provided, testing should also be conducted with those systems operating.
(6) Test Data . The applicant is expected to provide personnel and equipment to collect the following data:
(i) Test Site Data. The outside air temperature, wind velocity, and water depth should be recorded before each test run.
(ii) Airplane Performance Data . The airplane velocity should be recorded when it enters and exits the water test bed. Engine rotor speeds and interstage turbine temperature (ITT) (or exhaust gas temperature (EGT), as applicable) should be recorded and any engine abnormal engine sounds noted.
(iii) Water Spray Pattern . Appropriate areas of the airplane should be coated with agents that will permit identification of water spray impingement. Suitable high speed photography or video equipment should be utilized to record the origin, trajectory, and configuration of the water spray.
(iv) Pilot Comments . Pilot comments should be noted to determine if the takeoff would have been affected by abnormal audible engine sounds, airplane instrument anomalies, etc.
(7) Test Results . No hazardous quantities of water are considered to have been ingested when review of the collected test data shows that:
(i) No engine flameout, performance degradation, distress, or airspeed fluctuations occurred that would create a safety hazard.
(ii) No abnormal engine sounds occurred, such as pops or bangs, or cockpi t instrument indications of incipient engine surge or stall, or they were not of sufficient magnitude to cause a pilot in service to abort the planned operation.
132. INDUCTION SYSTEM ICING PROTECTION - § 25.1093 .
a. Reference. Advisory Circular 20-73, Aircraft Ice Protection, dated April 21, 1971.
b. Reciprocating Engines - § 25.1093(a) .
(1) Explanation.
(i) Conditions for tests . The carburetor air temperature measurement has been found to provide satisfactory average readings through the use of a minimum of three thermocouples so arranged as to give an average air temperature. This indicator should be calibrated prior to the test. Operationally, it has been determined that the tests should be conducted at an altitude where the free air temperature is 30 o F, or at two altitudes of different temperatures, one of which is near 30 o F.
(ii) Configuration . The test should be conducted in the configuration that follows:
(A) Weight - optional.
(B) C.G. Position - optional.
(C) Wing Flap Position - optional.
(D) Landing Gear Position - optional.
(E) Engines - 60 percent maximum continuous power.
(F) Cowl Flaps - appropriate for flight condition.
(G) Mixture Setting - normal cruising position.
(2) Procedures . Test procedures and required data.
(i) After all temperatures have stabilized (i.e., when the rate of temperature change is less than 2 o F per minute), and with the airplane in level flight and full cold carburetor at 60% maximum continuous power, the following data should be recorded:
(A) Pressure altitude.
(B) Ambient air temperature.
(C) Indicated airspeed.
(D) Carburetor air temperature.
(E) Engines, R.P.M. and manifold pressure.
(F) Torque pressure.
(G) Mixture setting.
(H) Cowl flap setting.
(ii) Preheat should then be applied slowly (power may be restored to 60 percent maximum continuous at the applicant's option) and the above data recorded again after the temperatures have stabilized. The carburetor heat rise is determined by comparing the results of the data obtained with and without preheat.
c. Turbine Powered Airplanes - § 25.1093(b) .
(1) Explanation.
(i) Section 25.1093(b) requires that each turbine engine operate throughout its flight po wer range without adverse effect on engine operation or serious loss of power or thrust in the icing conditions specified, including those related to falling and blowing snow. This requirement must be met for all operating conditions of the airplane and is not limited to operations where icing can be predicted or when icing penetrations are specifically intended. It is clear that the engines require protection against the possible hazardous effects of ice ingestion for all operating conditions, while the remaining airframe need be protected only if certification for flight in icing conditions is desired. This can also be verified by a review of §§ 25.1093(b), 25.1419, and 33.68.
(ii) One purpose of flight in natural and/or tanker-provided icing condit ions is to demonstrate that chunks of ice discharged from unprotected surfaces do not cause damage to the engine or other critical parts of the airplane, and to demonstrate that ice discharged from protected surfaces after an undetected icing encounter, during which the ice protection system was inoperative, does not cause engine damage or malfunction, or damage to other critical parts of the airplane. In order to accomplish these objectives, the airplane should be exposed to an icing condition of a magnitude and duration sufficient to accumulate enough ice to produce an acceptable demonstration.
If, after installation on the airplane, it is possible for the engines to ingest ice shed from the airframe, which will cause an adverse effect or serious loss of power or thrust on the engine, compliance with § 25.1093(b) has not been achieved, and the required level of airworthiness for the engine installation has not been obtained.
(iii) It has been established from experience that turbine engines can be aff ected seriously and adversely by inadvertent in-flight ice encounters in which only a minor performance loss could be attributable to the amount of ice accumulated on the airframe. Based upon this experience, it was deemed necessary to require a higher level of ice protection for the engines than was found necessary for the airframe. The requirements of §§
25.1093(b), 25.1419, and 33.68 reflect this concept. Satisfactory operation of protected surfaces at the design point condition, as described in Advisory Circular 20-73, should be demonstrated by icing tunnel tests or by analysis supported by tests. The ability to de-ice and anti-ice protected surfaces should be demonstrated during the natural and/or tanker-provided icing tests.
(iv) The secondary effects of falling and blowing snow should also be considered in the evaluation of susceptibility to, and protection from, ice accumulations. The critical ambient temperatures should be defined;
this evaluation should take into consideration not only temperatures in the immediate vicinity of the freezing point, where "wet, sticky snow" is likely, but also colder temperatures where snow may adhere to partially heated interior inlet surfaces, melt, and refreeze in a colder location. Any hardware or ancillary systems (e.g., screens, particle separators, oil coolers) installed in turbine engine inlets may facilitate the accumulation of snow and potential for generation of ice. The effects of falling and blowing snow should be evaluated for both ground and flight conditions. The number of actual airplane tests within the critical snow and temperature environment should be maximized. The test article must be in production configuration with respect to surface finish, texture, and material type, to assure accurate representation of the in-service configuration. Recognizing that the desired atmospheric conditions may be difficult to find, some conditions may be substantiated by analysis, provided that analysis is supported by some form of test data (e.g., temperature survey data).
(v) It would be inconsistent to consider that § 25.1093(b) pertains only to the engine or engine inlet lip. The wording is broad and objective. It should be noted that the icing requirements pertaining specifically to the engine are included in § 33.68. Section 25.1093(b) was included to assure that the engine, as installed in the airplane, did not suffer adverse effects when the airplane was subjected to the icing envelope requirements covered by Appendix C to Part 25. It is not logical to exclude any part of the airplane that might shed ice so as to produce an adverse effect on the engine.
(vi) The foregoing discussions are intended to describe the minimum program to achieve certification for flight in icing conditions.
(2) Procedures . None.
[133. - 136.] [RESERVED]