199 The fuel systems of some airplanes are designed to feed fuel to the engines from auxiliary and main fuel tanks. This results in variations in fuel feed configurations as part of the approved fuel management. Variations in fuel feed configurations, including changes in fuel feed configuration such as transition from feeding fuel to the engine(s) from one tank to feeding from another, should be considered in assessing fuel feed performance in demonstrating compliance to §§ 25.951 or 25.955 and considered in establishing the critical conditions for the hot fuel evaluation.
(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 operating limitations should be established and furnished 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 or thrust 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 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
200 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 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:
(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 or thrust 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 or thrust 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 power or 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.
(d) 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.
(a) 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.
(b) 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.
(c) Fuel in liquid or vapor form should 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.
(d) 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.
(e) There should be no evidence of leakage after the fuel jettisoning valve is closed.
(f) Testing should be conducted 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 AFM.
(3) Control.
(a) Changes in the airplane control qualities during the fuel jettisoning tests should be investigated, including asymmetrical jettisoning.
(b) 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. - 129.] [Reserved]
130. Cooling Test Procedures - § 25.1045.
a. Explanation. The following guidance applies to cooling test procedures for turbine engine powered airplanes.
(1) Purpose. In accordance with § 25.1041, cooling tests must be conducted to determine the ability of the powerplant cooling provisions to maintain the temperatures of powerplant components and engine fluids within the temperature limits for which they have been certificated. These limits will normally be specified on the engine type certificate data sheet (TCDS), qualification specification sheet for the component, and/or in the approved engine installation handbook.
(2) Scope. Cooling tests should be conducted under (or data corrected to) critical ground and flight operating conditions to the maximum altitude for which approval is requested.
b. Flight Test Procedures.
(1) Moisture. The tests should be conducted in air free of visible moisture.
(2) Weight and C.G. Forward c.g. at maximum gross weight is usually the most critical condition as this results in the lowest airspeed and/or vertical speed. However, a lighter initial gross weight or a critical step climb profile may be necessary to get to maximum altitude. In any case, the most critical climb profile must be established by the applicant and agreed to by the FAA prior to commencing testing.
(3) Test Conditions. The critical flight profile(s) should be tested. It may be necessary to fly multiple flights with different flight profiles to assure all components under test are exposed to their most critical anticipated sequence of test conditions. Flight profiles with periods
205 of relatively high heat rejection and relatively low cooling, such as can occur during step climbs and at the top of descent, may be more critical for some components, for example, those in or affected by the engine oil system. The period after engine shutdown on the ground may also be critical for some components. The applicant should identify and obtain the FAA’s approval of the flight test profile(s) prior to beginning certification testing. The following sequence of test conditions is usually adequate to cover the critical case for most components:
(a) Initiate the flight test once critical temperatures have stabilized after engine start;
(b) Use the test engine to perform a 1 mile single engine taxi;
(c) Hold at idle power or thrust in a 10 knot or greater crosswind for 20 minutes or until temperatures stabilize (use ground test definition of stabilization);
(d) Operate at least the test engine at the rated takeoff power or thrust for the maximum approved period (usually either 5 or 10 minutes) during a maximum gross weight/forward c.g. takeoff;
(e) Perform simulated one-engine-inoperative and all-engines-operating climbs, operating the test engine at maximum continuous power or thrust until the engine temperatures stabilize or the airplane reaches maximum operating altitude (or until the airplane is essentially unable to climb further as indicated by a very low climb rate, e.g., 200 feet/minute);
(f) Cruise with the test engine at maximum continuous power or thrust (but a speed no higher than VMO/MMO) at maximum operating altitude until temperatures stabilize;
(g) Conduct a normal descent at VMO/MMO to a typical holding altitude and hold until temperatures stabilize;
(h) Conduct a normal approach to landing, but from not less than 200 feet above the ground:
1 Perform a simulated engine out go-around;
2 Climb to pattern altitude; then
3 Perform a normal approach and landing.
(i) Taxi back to the ramp and shut down the engines; then
(j) Allow the test engine’s heat-soak to peak.
(4) Oil Quantity. The critical condition should be tested.
(5) Thermostat. Airplanes that incorporate a thermostat in the engine oil system may have the thermostat retained, removed, or blocked in such a manner as to pass all engine oil through the oil cooler. If the thermostat is retained, the oil temperature readings obtained on a cooler day corrected to hot day conditions may therefore be greater than those obtained under actual hot day conditions. Caution should be exercised when operating an airplane with the thermostat removed or blocked during cold weather to prevent failure of the lubricating system components.
(6) Instrumentation. The applicant must identify all critical components (electronic components, actuators, etc.), including structural elements that have temperature limits. The limits are typically based on component qualification or certification testing. Each limit may be expressed in terms of a surface temperature or environmental temperature, and may have associated time limits. Instrumentation should be installed to provide data needed to show that each component corrected temperature remains below the identified limit. Accurate and calibrated temperature-measuring devices should be used, along with acceptable thermocouples or temperature pickup devices. The temperature pickup should be located at critical engine positions.
(7) Generator. The alternator/generator should be electrically loaded to the rated capacity for the engine/accessory cooling tests.
(8) Maximum Ambient Atmospheric Temperature. Section 25.1043(b) establishes 100° F (38° C) at sea level as the lowest maximum ambient temperature for cooling tests, except for winterization installations. (See paragraph 130b(12) for guidance on certifying winterization equipment.) Applicants may establish a higher temperature limit if desired. In accordance with § 25.1041, applicants must show that cooling provisions can maintain the temperatures of powerplant components, engine fluids, and auxiliary power unit components and fluids within the established temperature limit. The assumed temperature lapse rate is -3.6° F (-2° C) per thousand feet of altitude above sea level until a temperature of -69.7° F (-56.5° C) is reached, above which altitude the temperature is considered constant at -69.7° F (-56.5° C). The compliance demonstration flight test should be conducted with an ambient temperature as close to the desired maximum ambient atmospheric temperature as practical. If testing is accomplished at lower ambient temperatures, then the test data must be corrected to that which would have resulted from testing on a day with the maximum ambient atmospheric temperature.
(See paragraph 130e(2).) The maximum ambient temperature selected and demonstrated satisfactorily, taking account of correction factors, shall not be less than the minimum hot day conditions prescribed by § 25.1043(b) and shall be an airplane operating limitation per the requirements of § 25.1521(d). The applicant should correct the engine temperatures to as high a value as possible in order to minimize the impact of this limitation.
(9) Temperature Stabilization. For the cooling flight tests, a temperature is usually considered stabilized when its observed rate of change is less than 2° F per minute. However, regardless of the rate of temperature rise, if a component or fluid temperature is still rising and is near the limit for that component or fluid, sufficient test data must be gathered to show that the limit will not be exceeded in that steady state operating condition. A combination of test data and rational analysis may be used to determine the maximum temperature expected for a
207 component or fluid rather than continuing a test condition for a long period of time after the rate of temperature rise has fallen below 2° F per minute.
(10) Temperature inversion. During an inversion, where a layer of colder air is at ground level, component temperatures at the beginning of the cooling test climb will be lower relative to the ambient air temperature encountered during the climb. For components that require significant time to adjust to ambient temperature changes, the artificially low starting temperature may result in erroneous results. If cooling tests are conducted when an inversion exists, the applicant should present a more rational correction method than either the “degree for degree” method described in paragraph 130e(2) or the method defined in § 25.1043. The FAA has accepted a correction method that used the difference between the test day ambient and the hot day atmospheric temperature taken at the beginning of the flight being used as the correction factor for all flight data. Although conservative, this method accounts for the effects of an inversion.
(11) Airport Altitude. The cooling tests should be conducted from the lowest practical airport altitude, usually below 3,000 feet mean sea level (MSL), to provide test data reasonably close to sea level.
(12) Winterization Equipment Procedures. The following procedures should be applied when certificating winterization equipment:
(a) Maximum ambient sea level atmospheric temperature less than 100° F (38° C).
Cooling test results for winterization installations may be corrected to any temperature desired by the applicant rather than the conventional 100° F (38° C) hot day. For example, an applicant may choose to demonstrate cooling to comply with requirements for a 50° F (10° C) or 60° F
(15.5° C) day with winterization equipment installed. This temperature becomes a limitation to be shown in the AFM. In such a case, the sea level temperature for correction purposes should be considered to be the value elected by the applicant with a rate of temperature drop of 3.6° F (2° C) per thousand feet above sea level.
(b) Tests. Cooling tests and temperature correction methods should be the same as for conventional cooling tests.
(c) Limit Temperature. The AFM should clearly indicate that winterization equipment should be removed whenever the temperature reaches the limit for which adequate cooling has been demonstrated. The cockpit should be placarded accordingly.
(d) Equipment Marking. If practical, winterization equipment, such as baffles for oil radiators or for engine cooling air openings, should be marked clearly to indicate the limiting temperature at which this equipment should be removed.
(e) Installation Instructions. Since winterization equipment is often supplied in kit form and accompanied by instructions for its installation, manufacturers should provide suitable information regarding temperature limitations in the installation instructions.
c. Ground Test Procedures.
(1) General. The flight testing guidance in paragraphs 130b(1), (4), (5), (6), (7), and
(11) are equally applicable to ground testing.
(2) Test Conditions. Ground testing should be conducted with the engine operated at idle power or thrust until all critical temperatures have stabilized, followed by operation of the engine at rated takeoff power or thrust for 5 minutes, and then idle power or thrust until all critical temperatures have stabilized.
(3) Temperature Stabilization. During the ground operation portion of the compliance demonstration, the definition of stabilized temperatures defined for the flight test (rate of change less than 2° F per minute (1o C)) should not be used for determining the maximum component temperatures, unless it can be shown that ground operation of the engine is limited to the conditions tested. The reason for using a different definition of stabilized temperatures for ground operation is that during sustained ground operations at a particular condition, such as at idle power or thrust for using the engines as a pneumatic source, may result in temperatures that gradually exceed the defined temperature limits. To address this concern, some manufacturers have used engine operating limitations for ground operations that limit engine operation for discrete time periods. For example, engine operation at takeoff power or thrust may be limited to 2 minutes at ambient temperatures above 110° F (43o C), or engine operation at idle power or thrust is limited to 2 hours at ambient temperatures above 110° F.
(4) Temperature Correction. Recorded ground temperatures should be corrected to the maximum ambient temperature selected, without consideration of the altitude temperature lapse rate. For example, if an auxiliary power unit is being tested for ground cooling margins, the cooling margin should be determined from the recorded ground temperature without regard to the test site altitude.
d. Data Acquisition. The following data should be recorded at the time intervals specified in the particular test program. The data may be manually recorded unless the quantity and frequency necessitate automatic or semi-automatic means:
(1) Outside air temperature (OAT);
(2) Altitude;
(3) Airspeed (knots);
(4) Gas generator r.p.m.;
(5) Engine torque;
(6) Time;
(7) Fan r.p.m.;
(8) Engine oil temperature;
(9) Pertinent engine temperature; and
(10) Pertinent nacelle and component temperatures.
e. Data Reduction.
(1) Purpose. Seldom is testing actually accomplished at the maximum required ambient temperature of at least 100º F at sea level lapsed 3.6º F (2o C) per 1,000 feet pressure altitude. Component and fluid temperatures must therefore be corrected to derive the item temperature that would have been reached if the test day had matched exactly the maximum ambient temperature day. The applicant may select a higher maximum ambient temperature for cooling certification than the 100º F (38o C) sea level hot day prescribed. Provisions are also made for selecting a maximum ambient temperature less than the 100º F sea level hot day for winterization installations not intended to function at the hot day conditions.
(2) Correction Factors for Ambient Conditions. Unless a more rational method applies, a correction factor of 1.0 is applied to the temperature data as follows: corrected temperature = true temperature + 1.0 [100 - 0.0036 (Hp) - true OAT]. A correction factor other than “degreefor-degree” should be based on engineering test data. The corrected temperature is then compared with the maximum permissible temperature to determine compliance with the cooling requirements. No corrected temperatures may exceed established limits.
Sample Calculation:
True Temperature = 300° F (149o C) True OAT = 15° F (-9o C) Hp = 5,000 feet
The corrected temperature = 300 + 1.0 [100 - 0.0036 (5,000) - 15] = 367° F (186o C).
(3) Correction Factor for Minimum Engine. An important correction factor that is not discussed in the regulations, but is frequently necessary to show the cooling adequacy required by § 25.1041, is the minimum engine (i.e., the thermal limit) correction factor. This factor is sometimes required if, at test day conditions, the engine measured temperature does not correspond to the engine temperature that would have occurred on a minimum specification engine in hot day conditions. The correction factor would not apply to those components not affected by changes in exhaust gas temperature (EGT) at a constant power or thrust. Typical items expected to be affected by changes in the EGT at constant power or thrust would be engine oil temperature, thermocouple harnesses, or other fluid, component, or ambient temperatures in the vicinity of the engine hot-section or exhaust gases. Other items remote from the hot section, like the starter-generator or fuel control, would not be expected to be influenced by EGT variations; however, the items affected and the magnitude of the factor to be applied should be established by testing. There are several acceptable methods for establishing the appropriate
210 correction factor during development testing. The general idea is to establish a stabilized flight condition, typically during ground runs, and vary the measured EGT at approximately fixed power or thrust and OAT conditions. This may be accomplished by using engine anti-ice bleed air, customer bleed air, or by ingesting warmer than ambient air (either an external source or the engine bleed air) into the engine inlet. Care should be used when ingesting warmer than ambient air to assure that the warm air is diffused in order to avoid possible engine surge. If it is not possible to attain an adequate variation in EGT by these methods, an acceptable, but conservative correction may be obtained by allowing both power or thrust and EGT to vary at stabilized engine operating conditions and OAT. The component temperature is plotted as a function of EGT, and the correction from test EGT for any flight condition, to the EGT that would have existed with minimum specification engines on a hot day, is then applied to derive the corrected component temperature. Both of these methods assume the inlet air and cooling air sources are essentially independent. Where they are not independent, such as those designs that take cooling air from the fan exit stream (e.g., via a fan box in bifurcation) rather than via a free stream scoop, care needs to be taken in selecting a technique to assure the results are neither overly conservative as may result from use of the hot air technique nor overly optimistic as may result from use of the higher power or thrust technique.
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 or thrust 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, 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) AC 20-124, “Water Ingestion Testing for Turbine Powered Airplanes,” dated September 30, 1985.
(2) AC 91-6A, “Performance Information for Operation with Water, Slush, Snow or Ice on the Runway,” dated May 24, 1978.