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.
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 needs to 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 main gear spray is not more hazardous than the spray produced by either gear separately.
(5) Test Procedures. The tests should be conducted with the appropriate power or thrust 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 VR 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:
(a) Test Site Data. The outside air temperature, wind velocity, and water depth should be recorded before each test run.
(b) 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 EGT, as applicable) should be recorded and any engine abnormal engine sounds noted.
(c) 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 used to record the origin, trajectory, and configuration of the water spray.
(d) 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:
(a) No engine flameout, performance degradation, distress, or airspeed fluctuations occurred that would create a safety hazard.
(b) No abnormal engine sounds occurred, such as pops or bangs, or cockpit 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. References.
(1) AC 20-73A, “Aircraft Ice Protection,” dated August 16, 2006.
(2) AC 20-147, “Turbojet, Turboprop, and Turbofan Engine Induction System Icing and Ice Ingestion,” dated February 2, 2004.
b. Reciprocating Engines - § 25.1093(a).
(1) Explanation.
(a) 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 30º F (-1o C), or at two altitudes of different temperatures, one of which is near 30º F (-1 oC).
(b) Configuration. The test should be conducted in the configuration that follows:
1 Weight - optional.
213 2 C.G. Position - optional.
3 Wing Flap Position - optional.
4 Landing Gear Position - optional.
5 Engines - 60 percent maximum continuous power.
6 Cowl Flaps - appropriate for flight condition.
7 Mixture Setting - normal cruising position.
(2) Procedures. Test procedures and required data.
(a) After all temperatures have stabilized (i.e., when the rate of temperature change is less than 2º F (1 oC) per minute), and with the airplane in level flight and full cold carburetor at 60 percent maximum continuous power, the following data should be recorded:
1 Pressure altitude.
2 Ambient air temperature.
3 Indicated airspeed.
4 Carburetor air temperature.
5 Engines’ r.p.m. and manifold pressure.
6 Torque pressure.
7 Mixture setting.
8 Cowl flap setting.
(b) 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.
(a) Section 25.1093(b) requires that each turbine engine operate throughout its flight power or thrust 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
214 for flight in icing conditions is desired. This can also be verified by a review of §§ 25.1093(b),
25.1419, and 33.68.
(b) One purpose of flight in natural and/or tanker-provided icing conditions 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.
(c) It has been established from experience that turbine engines can be affected 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 AC 20-73A, 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.
(d) 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. In accordance with § 21.35(a)(3), 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).
(e) 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
215 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.
(f) The preceding paragraphs are intended to describe the minimum program to achieve certification for flight in icing conditions.
(2) Procedures. None.
[133. - 136.] [Reserved]
Section 7. Exhaust System
137. General - § 25.1121.
a. Explanation. Section 25.1121(a) - Carbon monoxide contamination. Carbon monoxide detection tests are conducted in accordance with this requirement to determine that the disposal of exhaust gases from each exhaust system does not cause carbon monoxide contamination of any personnel compartment.
b. References. Also see information related to the evacuation of other personnel compartment atmosphere contaminants contained in paragraphs 84 and 165 of this AC addressing the requirements of §§ 25.831 and 25.1197, respectively.
[138. - 140.] [Reserved]
Section 8. Powerplant Controls and Accessories
[141. - 154.] [Reserved]
Section 9. Powerplant Fire Protection
[155. - 158.] [Reserved]
159. Drainage and Ventilation Of Fire Zones - § 25.1187.
a. Explanation. The rule requires that each part of each designated fire zone be drained completely of flammable fluids to minimize hazards resulting from malfunctioning or failing components that contain flammable fluids. The drainage means must be:
(1) Effective under conditions expected to prevail when drainage is needed;
(2) Arranged so that no discharged fluid will cause an additional fire hazard; and
(3) Arranged so that no discharged fluid will enter any other fire zone.
[160. - 164.] [Reserved]
165. Fire Extinguishing Systems - § 25.1197.
a. Explanation. Carbon Dioxide in Flightcrew Compartments. Carbon dioxide has been found to adversely affect flightcrew personnel in the performance of their duties. Therefore, in airplanes equipped with built-in carbon dioxide fuselage compartment fire extinguisher systems, the carbon dioxide concentration occurring at the flightcrew stations as a result of discharging the fire extinguishers should be determined in accordance with the procedures of this section (also see paragraph 84 of this AC), except that such determination is not considered necessary if:
(1) Five pounds or less of carbon dioxide will be discharged into any one such fuselage compartment in accordance with established fire control procedures; or
(2) Protective breathing equipment is provided for each flight crewmember on flight deck duty.
b. Procedures. Flight Test Investigation.
(1) The carbon dioxide concentrations at breathing level at the flightcrew stations should be determined in flight tests during which fuselage compartment fire extinguishers are discharged in accordance with established fire control procedures. Since carbon dioxide is heavier than air, a nose-down attitude is likely to produce the critical concentrations in the crew compartment. Perform the tests described in paragraphs (2) and (3) below.
(2) A rapid descent at the “maximum operating limit speed” of the airplane with the flaps and landing gear up.
(3) A rapid descent with the flaps and landing gear down, at the maximum permissible speed for this configuration. If it appears that any other condition is likely to be critical on a particular airplane, it should also be investigated.
(4) In the flight tests specified above, it will be permissible to institute emergency ventilating procedures immediately prior to or following the discharge of carbon dioxide, provided such procedures can be accomplished easily and quickly by the flightcrew, and do not appreciably reduce the effectiveness of the fire protection system.
(5) If the measured carbon dioxide concentrations exceed three percent by volume (corrected to sea level, standard day conditions), protective breathing equipment should be provided for each flight crewmember on flight deck duty.
(6) Appropriate emergency operating procedures should be entered in the AFM.
[166. - 169.] [Reserved]
218 Chapter 6 - Equipment
Section 1. General
170. Equipment - Function and Installation - § 25.1301.
a. Explanation.
(1) Certification of the installation of modern avionics/electrical systems on airplanes can be summarized, generally, by stating that the systems/equipment must:
(a) Perform its intended function (§ 25.1301);
(b) Be adequately protected for failure conditions (§ 25.1309);
(c) Be arranged to provide proper pilot visibility and utilization (as appropriate) (§ 25.1321);
(d) Be protected by circuit breakers to preclude failure propagation and/or minimize distress to the airplane’s electrical system (§ 25.1357); and
(e) Be installed in a manner such that operation of the system will not adversely affect the simultaneous operation of any other system (§ 25.1431).
(2) Accordingly, the recommended flight test procedures for equipment covered in subpart F of part 25 (excluding powerplant instruments, airspeed calibration, safety equipment, and lights) are grouped together under this section and are organized as follows:
Para. b. Communication Systems.
Para. c. Navigation Systems.
Para. d. Instruments and Displays.
Para. e. Sensors and Warning Systems.
Para. f. Recording Systems.
Para. g. Engine Interfacing Systems (autothrottle, power/thrust rating, ATTCS, etc.).
Para. h. Stability Augmentation Systems.
Para. i. All Weather Operation (Reduced Visibility) Systems.
(3) Section 25.1301 refers to each item of installed equipment. Type certification involving equipment approvals is governed by Title 49 United States Code, section 44704. That section requires the Administrator to make, or require the applicant for a type certificate to make, such tests “necessary in the interest of safety.” Tests may be necessary in order to allow the Administrator to find that the aircraft is properly designed and manufactured, performs properly, and meets the regulations and minimum standards prescribed under section 44701(a) of Title 49.
219 “Optional equipment,” however, is not a term that is meaningful under the Act in connection with the type certification of any given airplane. Where equipment or a system is a part or appurtenance of the airplane and is designed to aid and will obviously be used by the crew, the statutorily required tests and findings are the same, regardless of whether or not it is characterized as optional. Moreover, the regulatory requirement of § 21.21(b)(2) specifies that the Administrator must find that no feature or characteristic of the airplane makes it unsafe for the category in which certification is requested. Therefore, the extent to which that equipment needs to be tested or evaluated, in order that the Administrator may make the necessary finding with respect to the whole airplane, is a technical determination within the engineering and operational expertise of the Administrator. (See AC 20-68, “Certification Guidance for Installation of Non-Essential, Non-Required Aircraft Cabin Systems & Equipment (CS&E),” dated July 22, 2010, for further information.)
(4) Criteria and requirements for flight should consider the applicant’s engineering analysis and laboratory (simulator) test program. The combination of analysis, laboratory, and flight evaluation will form the whole of the certification requirement and, as such, should be in harmony and provide full evaluation. The flight evaluations supplement the analysis and simulation as required in both the representative operating conditions of the airplane and for analysis of the airplane modes of operation and configuration conditions.
(5) The requirement for demonstrating safe operation should normally include induced failures during flight. The requirement for failure demonstrations is also an outgrowth of the analysis and laboratory test results submitted by the applicant and is a result of the particular design being evaluated (e.g., consideration should be made for multiple channel systems specifically designed to be self-adaptive to failure conditions). Performance and malfunctions testing should include those flight conditions and airplane configurations that have been identified to be the most critical by analysis and/or test. Such items as weight, c.g., speed, altitude, flaps, slats, gear, speed brakes, and airplane system degradation should be considered.
(6) The amount of flight testing should be determined through the cooperative efforts of the assigned project personnel. It is recommended that the procedures that follow be used as a guide in preparing for the flight testing of an initial certification program. Follow-on items relevant to system derivative certification may result in considerably reduced flight testing.
However, sufficient testing should be accomplished to assure satisfactory performance. When ground or flight test data, available from similar previously approved installations, are sufficient to properly evaluate a system’s performance, additional testing may not be required. In the absence of such data, additional testing or analysis should be presented to substantiate the areas potentially affected.
(7) Particular attention should be given to those installations where an external piece of gear, such as an antenna, could affect the flight characteristics. All installations of this nature should be evaluated by the flight test pilot.
(8) Installations that can or may change the established limitations, flight characteristics, performance, operating procedures, or any required systems require approval by an FAA ACO flight test branch. New installations of equipment in the cockpit, or modifications
220 that affect existing equipment in the cockpit, should be evaluated through the cooperative efforts of the FAA, the project engineer, and the assigned flight test pilot, and assessed for the need of an FAA flight test demonstration.
(9) Throughout the systems/equipment evaluation, the operation of annunciators should be assessed to determine that proper conspicuity and display are provided to the appropriate flight crewmember. Any mode of operation selected by a manual action, or automatically, should be positively identified. Any submode should be evaluated to determine the need for annunciation.
b. Procedures - Communications.
(1) Very high frequency (VHF) Systems.
(a) Airplanes to be operated above 18,000 ft. Intelligible communications should be provided between the airplane and facilities throughout circle-turns within 160 NM of an FAA approved ground facility and above minimum radio line-of-sight with no intervening terrain. Bank angles up to 10 degrees on all headings should be used. Drop outs that are relieved by a reduction in bank angle at the same relative heading to the station are satisfactory.
It is suggested that the “Long Range Reception Test” of paragraph (d), below, be conducted first.
If this test is successful, the circle-turns (within 160 NM) of this paragraph need not be conducted. Skidding turns may be used to minimize turn radius.)
(b) Airplanes to be operated below 18,000 ft. For airplanes limited to operation below 18,000 ft., intelligible communications should be provided as given in (i) above, except that the distance from the ground facility need not exceed 80 NM.
(c) Antenna Coverage Measurement. If the antenna is located on the airplane centerline, tests may be conducted using only one direction of turn. When antenna radiation pattern data are available, flight testing in a 360 degree turn may not be necessary, if satisfactory communication is achieved during checks in the vicinity of the predicted bearings and bank angles of worst performance.
(d) Long Range Reception. At a distance of at least 160 NM (or 80 NM for airplanes to be operated below 18,000 ft.) from the ground facility antenna on a heading and above radio line-of-sight, perform a right and/or left 360 degree turn at a bank angle of at least 10 degrees. Communicate with the ground facility every 10 degrees of turn to test the intelligibility of the signals received at the ground station and in the airplane. The minimum line-of-sight altitude for 160 NM is approximately 17,000 ft. and for 80 NM is approximately 4,000 ft. Radio line-of-sight distances vs. flight level (or altitude) are given in the following table:
Table 170-1 Radio Line-of-Sight Distance vs. Flight Level or Altitude
Flight Level or Altitude in Feet Radio Line-of- Sight in Nautical Miles Flight Level or Altitude in Feet Radio Line-of-Sight in Nautical Miles FL500 275.0 FL260 198.3 FL490 272.3 FL250 194.5 FL480 269.5 FL240 190.6 FL470 266.7 FL230 186.5 FL460 263.8 FL220 182.4 FL450 260.9 FL210 178.2 FL440 258.0 FL200 173.9 FL430 255.1 FL190 169.5 FL420 252.1 FL180 165.0 FL410 249.1 17,000 160.4 FL400 246.0 16,000 155.6 FL390 242.9 15,000 150.6 FL380 239.8 14,000 145.5 FL370 236.6 13,000 140.2 FL360 233.4 12,000 134.7 FL350 230.1 11,000 129.0 FL340 226.8 10,000 123.0 FL330 223.4 9,000 116.7 FL320 220.0 8,000 110.0 FL310 216.6 7,000 102.9 FL300 213.0 6,000 95.3 FL290 209.5 5,000 87.0 FL280 205.8 4,000 77.8 FL270 202.1 3,000 67.4
(e) High Angle Reception. Repeat the above test at a distance of 50 to 70 NM from the ground station and at an altitude of 35,000 ft., or the maximum operating altitude of the airplane (20 to 30 NM for airplanes to be operated below 18,000 ft.).
(f) Approach Configuration. With the landing gear extended and the flaps in the approach configuration, demonstrate intelligible communications between the airplane and the ground facility.
(g) Electromagnetic Compatibility (EMC). With all systems operating in flight, if practicable, verify, by observation, that no adverse effects are present in the required flight systems.
(2) High Frequency (HF) Systems.
(a) Acceptable communication should be demonstrated by contacting