14 C.F.R. Subpart G—Operating Limitations and Information
Title 14 - Aeronautics and Space
(a) Each operating limitation specified in §§25.1503 through 25.1533 and other limitations and information necessary for safe operation must be established. (b) The operating limitations and other information necessary for safe operation must be made available to the crewmembers as prescribed in §§25.1541 through 25.1587. [Amdt. 25–42, 43 FR 2323, Jan. 16, 1978] When airspeed limitations are a function of weight, weight distribution, altitude, or Mach number, limitations corresponding to each critical combination of these factors must be established. The maximum operating limit speed (V [Amdt. 25–23, 35 FR 5680, Apr. 8, 1970] The maneuvering speed must be established so that it does not exceed the design maneuvering speed V The established flap extended speed V The minimum control speed V (a) The established landing gear operating speed or speeds, V (b) The established landing gear extended speed V [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–38, 41 FR 55468, Dec. 20, 1976] Any other limitation associated with speed must be established. [Doc. No. 2000–8511, 66 FR 34024, June 26, 2001] A rough air speed, VRA, for use as the recommended turbulence penetration airspeed in §25.1585(a)(8), must be established, which— (1) Is not greater than the design airspeed for maximum gust intensity, selected for VB; and (2) Is not less than the minimum value of VB specified in §25.335(d); and (3) Is sufficiently less than VMO to ensure that likely speed variation during rough air encounters will not cause the overspeed warning to operate too frequently. In the absence of a rational investigation substantiating the use of other values, VRA must be less than VMO—35 knots (TAS). [Doc. No. 27902, 61 FR 5222, Feb. 9, 1996] The airplane weight, center of gravity, and weight distribution limitations determined under §§25.23 through 25.27 must be established as operating limitations. (a) General. The powerplant limitations prescribed in this section must be established so that they do not exceed the corresponding limits for which the engines or propellers are type certificated and do not exceed the values on which compliance with any other requirement of this part is based. (b) Reciprocating engine installations. Operating limitations relating to the following must be established for reciprocating engine installations: (1) Horsepower or torque, r.p.m., manifold pressure, and time at critical pressure altitude and sea level pressure altitude for— (i) Maximum continuous power (relating to unsupercharged operation or to operation in each supercharger mode as applicable); and (ii) Takeoff power (relating to unsupercharged operation or to operation in each supercharger mode as applicable). (2) Fuel grade or specification. (3) Cylinder head and oil temperatures. (4) Any other parameter for which a limitation has been established as part of the engine type certificate except that a limitation need not be established for a parameter that cannot be exceeded during normal operation due to the design of the installation or to another established limitation. (c) Turbine engine installations. Operating limitations relating to the following must be established for turbine engine installations: (1) Horsepower, torque or thrust, r.p.m., gas temperature, and time for— (i) Maximum continuous power or thrust (relating to augmented or unaugmented operation as applicable). (ii) Takeoff power or thrust (relating to augmented or unaugmented operation as applicable). (2) Fuel designation or specification. (3) Any other parameter for which a limitation has been established as part of the engine type certificate except that a limitation need not be established for a parameter that cannot be exceeded during normal operation due to the design of the installation or to another established limitation. (d) Ambient temperature. An ambient temperature limitation (including limitations for winterization installations, if applicable) must be established as the maximum ambient atmospheric temperature established in accordance with §25.1043(b). [Amdt. 25–72, 55 FR 29786, July 20, 1990] If an auxiliary power unit is installed in the airplane, limitations established for the auxiliary power unit, including categories of operation, must be specified as operating limitations for the airplane. [Amdt. 25–72, 55 FR 29786, July 20, 1990] The minimum flight crew must be established so that it is sufficient for safe operation, considering— (a) The workload on individual crewmembers; (b) The accessibility and ease of operation of necessary controls by the appropriate crewmember; and (c) The kind of operation authorized under §25.1525. The criteria used in making the determinations required by this section are set forth in appendix D. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–3, 30 FR 6067, Apr. 29, 1965] The kinds of operation to which the airplane is limited are established by the category in which it is eligible for certification and by the installed equipment. The extremes of the ambient air temperature and operating altitude for which operation is allowed, as limited by flight, structural, powerplant, functional, or equipment characteristics, must be established. [Doc. No. 2000–8511, 66 FR 34024, June 26, 2001] The applicant must prepare Instructions for Continued Airworthiness in accordance with appendix H to this part that are acceptable to the Administrator. The instructions may be incomplete at type certification if a program exists to ensure their completion prior to delivery of the first airplane or issuance of a standard certificate of airworthiness, whichever occurs later. [Amdt. 25–54, 45 FR 60173, Sept. 11, 1980] Load factor limitations, not exceeding the positive limit load factors determined from the maneuvering diagram in §25.333(b), must be established. (a) Additional operating limitations must be established as follows: (1) The maximum takeoff weights must be established as the weights at which compliance is shown with the applicable provisions of this part (including the takeoff climb provisions of §25.121(a) through (c), for altitudes and ambient temperatures). (2) The maximum landing weights must be established as the weights at which compliance is shown with the applicable provisions of this part (including the landing and approach climb provisions of §§25.119 and 25.121(d) for altitudes and ambient temperatures). (3) The minimum takeoff distances must be established as the distances at which compliance is shown with the applicable provisions of this part (including the provisions of §§25.109 and 25.113, for weights, altitudes, temperatures, wind components, runway surface conditions (dry and wet), and runway gradients) for smooth, hard-surfaced runways. Additionally, at the option of the applicant, wet runway takeoff distances may be established for runway surfaces that have been grooved or treated with a porous friction course, and may be approved for use on runways where such surfaces have been designed constructed, and maintained in a manner acceptable to the Administrator. (b) The extremes for variable factors (such as altitude, temperature, wind, and runway gradients) are those at which compliance with the applicable provisions of this part is shown. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–38, 41 FR 55468, Dec. 20, 1976; Amdt. 25–72, 55 FR 29786, July 20, 1990; Amdt. 25–92, 63 FR 8321, Feb. 18, 1998]
Title 14: Aeronautics and Space
PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES
Subpart G—Operating Limitations and Information
§ 25.1501 General.
Operating Limitations
§ 25.1503 Airspeed limitations: general.
§ 25.1505 Maximum operating limit speed.
§ 25.1507 Maneuvering speed.
§ 25.1511 Flap extended speed.
§ 25.1513 Minimum control speed.
§ 25.1515 Landing gear speeds.
§ 25.1516 Other speed limitations.
§ 25.1517 Rough air speed, V
§ 25.1519 Weight, center of gravity, and weight distribution.
§ 25.1521 Powerplant limitations.
§ 25.1522 Auxiliary power unit limitations.
§ 25.1523 Minimum flight crew.
§ 25.1525 Kinds of operation.
§ 25.1527 Ambient air temperature and operating altitude.
§ 25.1529 Instructions for Continued Airworthiness.
§ 25.1531 Maneuvering flight load factors.
§ 25.1533 Additional operating limitations.
Markings and Placards
§ 25.1541 General.
(a) The airplane must contain—
(1) The specified markings and placards; and
(2) Any additional information, instrument markings, and placards required for the safe operation if there are unusual design, operating, or handling characteristics.
(b) Each marking and placard prescribed in paragraph (a) of this section—
(1) Must be displayed in a conspicuous place; and
(2) May not be easily erased, disfigured, or obscured.
§ 25.1543 Instrument markings: general.
For each instrument—
(a) When markings are on the cover glass of the instrument, there must be means to maintain the correct alignment of the glass cover with the face of the dial; and
(b) Each instrument marking must be clearly visible to the appropriate crewmember.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–72, 55 FR 29786, July 20, 1990]
§ 25.1545 Airspeed limitation information.
The airspeed limitations required by §25.1583 (a) must be easily read and understood by the flight crew.
§ 25.1547 Magnetic direction indicator.
(a) A placard meeting the requirements of this section must be installed on, or near, the magnetic direction indicator.
(b) The placard must show the calibration of the instrument in level flight with the engines operating.
(c) The placard must state whether the calibration was made with radio receivers on or off.
(d) Each calibration reading must be in terms of magnetic heading in not more than 45 degree increments.
§ 25.1549 Powerplant and auxiliary power unit instruments.
For each required powerplant and auxiliary power unit instrument, as appropriate to the type of instrument—
(a) Each maximum and, if applicable, minimum safe operating limit must be marked with a red radial or a red line;
(b) Each normal operating range must be marked with a green arc or green line, not extending beyond the maximum and minimum safe limits;
(c) Each takeoff and precautionary range must be marked with a yellow arc or a yellow line; and
(d) Each engine, auxiliary power unit, or propeller speed range that is restricted because of excessive vibration stresses must be marked with red arcs or red lines.
[Amdt. 25–40, 42 FR 15044, Mar. 17, 1977]
§ 25.1551 Oil quantity indication.
Each oil quantity indicating means must be marked to indicate the quantity of oil readily and accurately.
[Amdt. 25–72, 55 FR 29786, July 20, 1990]
§ 25.1553 Fuel quantity indicator.
If the unusable fuel supply for any tank exceeds one gallon, or five percent of the tank capacity, whichever is greater, a red arc must be marked on its indicator extending from the calibrated zero reading to the lowest reading obtainable in level flight.
§ 25.1555 Control markings.
(a) Each cockpit control, other than primary flight controls and controls whose function is obvious, must be plainly marked as to its function and method of operation.
(b) Each aerodynamic control must be marked under the requirements of §§25.677 and 25.699.
(c) For powerplant fuel controls—
(1) Each fuel tank selector control must be marked to indicate the position corresponding to each tank and to each existing cross feed position;
(2) If safe operation requires the use of any tanks in a specific sequence, that sequence must be marked on, or adjacent to, the selector for those tanks; and
(3) Each valve control for each engine must be marked to indicate the position corresponding to each engine controlled.
(d) For accessory, auxiliary, and emergency controls—
(1) Each emergency control (including each fuel jettisoning and fluid shutoff must be colored red; and
(2) Each visual indicator required by §25.729(e) must be marked so that the pilot can determine at any time when the wheels are locked in either extreme position, if retractable landing gear is used.
§ 25.1557 Miscellaneous markings and placards.
(a) Baggage and cargo compartments and ballast location. Each baggage and cargo compartment, and each ballast location must have a placard stating any limitations on contents, including weight, that are necessary under the loading requirements. However, underseat compartments designed for the storage of carry-on articles weighing not more than 20 pounds need not have a loading limitation placard.
(b) Powerplant fluid filler openings. The following apply:
(1) Fuel filler openings must be marked at or near the filler cover with—
(i) The word “fuel”;
(ii) For reciprocating engine powered airplanes, the minimum fuel grade;
(iii) For turbine engine powered airplanes, the permissible fuel designations; and
(iv) For pressure fueling systems, the maximum permissible fueling supply pressure and the maximum permissible defueling pressure.
(2) Oil filler openings must be marked at or near the filler cover with the word “oil”.
(3) Augmentation fluid filler openings must be marked at or near the filler cover to identify the required fluid.
(c) Emergency exit placards. Each emergency exit placard must meet the requirements of §25.811.
(d) Doors. Each door that must be used in order to reach any required emergency exit must have a suitable placard stating that the door is to be latched in the open position during takeoff and landing.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–32, 37 FR 3972, Feb. 24, 1972; Amdt. 25–38, 41 FR 55468, Dec. 20, 1976; Amdt. 25–72, 55 FR 29786, July 20, 1990]
§ 25.1561 Safety equipment.
(a) Each safety equipment control to be operated by the crew in emergency, such as controls for automatic liferaft releases, must be plainly marked as to its method of operation.
(b) Each location, such as a locker or compartment, that carries any fire extinguishing, signaling, or other life saving equipment must be marked accordingly.
(c) Stowage provisions for required emergency equipment must be conspicuously marked to identify the contents and facilitate the easy removal of the equipment.
(d) Each liferaft must have obviously marked operating instructions.
(e) Approved survival equipment must be marked for identification and method of operation.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–46, 43 FR 50598, Oct. 30, 1978]
§ 25.1563 Airspeed placard.
A placard showing the maximum airspeeds for flap extension for the takeoff, approach, and landing positions must be installed in clear view of each pilot.
Airplane Flight Manual
§ 25.1581 General.
(a) Furnishing information. An Airplane Flight Manual must be furnished with each airplane, and it must contain the following:
(1) Information required by §§25.1583 through 25.1587.
(2) Other information that is necessary for safe operation because of design, operating, or handling characteristics.
(3) Any limitation, procedure, or other information established as a condition of compliance with the applicable noise standards of part 36 of this chapter.
(b) Approved information. Each part of the manual listed in §§25.1583 through 25.1587, that is appropriate to the airplane, must be furnished, verified, and approved, and must be segregated, identified, and clearly distinguished from each unapproved part of that manual.
(c) [Reserved]
(d) Each Airplane Flight Manual must include a table of contents if the complexity of the manual indicates a need for it.
[Amdt. 25–42, 43 FR 2323, Jan. 16, 1978, as amended by Amdt. 25–72, 55 FR 29786, July 20, 1990]
§ 25.1583 Operating limitations.
(a) Airspeed limitations. The following airspeed limitations and any other airspeed limitations necessary for safe operation must be furnished:
(1) The maximum operating limit speed V
(2) If an airspeed limitation is based upon compressibility effects, a statement to this effect and information as to any symptoms, the probable behavior of the airplane, and the recommended recovery procedures.
(3) The maneuvering speed V
(4) The flap extended speed V
(5) The landing gear operating speed or speeds, and a statement explaining the speeds as defined in §25.1515(a).
(6) The landing gear extended speed V
(b) Powerplant limitations. The following information must be furnished:
(1) Limitations required by §25.1521 and §25.1522.
(2) Explanation of the limitations, when appropriate.
(3) Information necessary for marking the instruments required by §§25.1549 through 25.1553.
(c) Weight and loading distribution. The weight and center of gravity limitations established under §25.1519 must be furnished in the Airplane Flight Manual. All of the following information, including the weight distribution limitations established under §25.1519, must be presented either in the Airplane Flight Manual or in a separate weight and balance control and loading document that is incorporated by reference in the Airplane Flight Manual:
(1) The condition of the airplane and the items included in the empty weight as defined in accordance with §25.29.
(2) Loading instructions necessary to ensure loading of the airplane within the weight and center of gravity limits, and to maintain the loading within these limits in flight.
(3) If certification for more than one center of gravity range is requested, the appropriate limitations, with regard to weight and loading procedures, for each separate center of gravity range.
(d) Flight crew. The number and functions of the minimum flight crew determined under §25.1523 must be furnished.
(e) Kinds of operation. The kinds of operation approved under §25.1525 must be furnished.
(f) Ambient air temperatures and operating altitudes. The extremes of the ambient air temperatures and operating altitudes established under §25.1527 must be furnished.
(g) [Reserved]
(h) Additional operating limitations. The operating limitations established under §25.1533 must be furnished.
(i) Maneuvering flight load factors. The positive maneuvering limit load factors for which the structure is proven, described in terms of accelerations, must be furnished.
[Doc. No. 5066, 29 FR 1891, Dec. 24, 1964, as amended by Amdt. 25–38, 41 FR 55468, Dec, 20, 1976; Amdt. 25–42, 43 FR 2323, Jan. 16, 1978; Amdt. 25–46, 43 FR 50598, Oct. 30, 1978; Amdt. 25–72, 55 FR 29787, July 20, 1990; Amdt. 25–105, 66 FR 34024, June 26, 2001]
§ 25.1585 Operating procedures.
(a) Operating procedures must be furnished for—
(1) Normal procedures peculiar to the particular type or model encountered in connection with routine operations;
(2) Non-normal procedures for malfunction cases and failure conditions involving the use of special systems or the alternative use of regular systems; and
(3) Emergency procedures for foreseeable but unusual situations in which immediate and precise action by the crew may be expected to substantially reduce the risk of catastrophe.
(b) Information or procedures not directly related to airworthiness or not under the control of the crew, must not be included, nor must any procedure that is accepted as basic airmanship.
(c) Information identifying each operating condition in which the fuel system independence prescribed in §25.953 is necessary for safety must be furnished, together with instructions for placing the fuel system in a configuration used to show compliance with that section.
(d) The buffet onset envelopes, determined under §25.251 must be furnished. The buffet onset envelopes presented may reflect the center of gravity at which the airplane is normally loaded during cruise if corrections for the effect of different center of gravity locations are furnished.
(e) Information must be furnished that indicates that when the fuel quantity indicator reads “zero” in level flight, any fuel remaining in the fuel tank cannot be used safely in flight.
(f) Information on the total quantity of usable fuel for each fuel tank must be furnished.
[Doc. No. 2000–8511, 66 FR 34024, June 26, 2001]
§ 25.1587 Performance information.
(a) Each Airplane Flight Manual must contain information to permit conversion of the indicated temperature to free air temperature if other than a free air temperature indicator is used to comply with the requirements of §25.1303(a)(1).
(b) Each Airplane Flight Manual must contain the performance information computed under the applicable provisions of this part (including §§25.115, 25.123, and 25.125 for the weights, altitudes, temperatures, wind components, and runway gradients, as applicable) within the operational limits of the airplane, and must contain the following:
(1) In each case, the conditions of power, configuration, and speeds, and the procedures for handling the airplane and any system having a significant effect on the performance information.
(2) VSR determined in accordance with §25.103.
(3) The following performance information (determined by extrapolation and computed for the range of weights between the maximum landing weight and the maximum takeoff weight):
(i) Climb in the landing configuration.
(ii) Climb in the approach configuration.
(iii) Landing distance.
(4) Procedures established under §25.101(f) and (g) that are related to the limitations and information required by §25.1533 and by this paragraph (b) in the form of guidance material, including any relevant limitations or information.
(5) An explanation of significant or unusual flight or ground handling characteristics of the airplane.
(6) Corrections to indicated values of airspeed, altitude, and outside air temperature.
(7) An explanation of operational landing runway length factors included in the presentation of the landing distance, if appropriate.
[Doc. No. 2000–8511, 66 FR 34024, June 26, 2001, as amended by Amdt. 25–108, 67 FR 70828, Nov. 26, 2002]
Appendix A to Part 25
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Appendix B to Part 25
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Appendix C to Part 25
(a) Continuous maximum icing. The maximum continuous intensity of atmospheric icing conditions (continuous maximum icing) is defined by the variables of the cloud liquid water content, the mean effective diameter of the cloud droplets, the ambient air temperature, and the interrelationship of these three variables as shown in figure 1 of this appendix. The limiting icing envelope in terms of altitude and temperature is given in figure 2 of this appendix. The inter-relationship of cloud liquid water content with drop diameter and altitude is determined from figures 1 and 2. The cloud liquid water content for continuous maximum icing conditions of a horizontal extent, other than 17.4 nautical miles, is determined by the value of liquid water content of figure 1, multiplied by the appropriate factor from figure 3 of this appendix.
(b) Intermittent maximum icing. The intermittent maximum intensity of atmospheric icing conditions (intermittent maximum icing) is defined by the variables of the cloud liquid water content, the mean effective diameter of the cloud droplets, the ambient air temperature, and the interrelationship of these three variables as shown in figure 4 of this appendix. The limiting icing envelope in terms of altitude and temperature is given in figure 5 of this appendix. The inter-relationship of cloud liquid water content with drop diameter and altitude is determined from figures 4 and 5. The cloud liquid water content for intermittent maximum icing conditions of a horizontal extent, other than 2.6 nautical miles, is determined by the value of cloud liquid water content of figure 4 multiplied by the appropriate factor in figure 6 of this appendix.
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Appendix D to Part 25
Criteria for determining minimum flight crew. The following are considered by the Agency in determining the minimum flight crew under §25.1523:
(a) Basic workload functions. The following basic workload functions are considered:
(1) Flight path control.
(2) Collision avoidance.
(3) Navigation.
(4) Communications.
(5) Operation and monitoring of aircraft engines and systems.
(6) Command decisions.
(b) Workload factors. The following workload factors are considered significant when analyzing and demonstrating workload for minimum flight crew determination:
(1) The accessibility, ease, and simplicity of operation of all necessary flight, power, and equipment controls, including emergency fuel shutoff valves, electrical controls, electronic controls, pressurization system controls, and engine controls.
(2) The accessibility and conspicuity of all necessary instruments and failure warning devices such as fire warning, electrical system malfunction, and other failure or caution indicators. The extent to which such instruments or devices direct the proper corrective action is also considered.
(3) The number, urgency, and complexity of operating procedures with particular consideration given to the specific fuel management schedule imposed by center of gravity, structural or other considerations of an airworthiness nature, and to the ability of each engine to operate at all times from a single tank or source which is automatically replenished if fuel is also stored in other tanks.
(4) The degree and duration of concentrated mental and physical effort involved in normal operation and in diagnosing and coping with malfunctions and emergencies.
(5) The extent of required monitoring of the fuel, hydraulic, pressurization, electrical, electronic, deicing, and other systems while en route.
(6) The actions requiring a crewmember to be unavailable at his assigned duty station, including: observation of systems, emergency operation of any control, and emergencies in any compartment.
(7) The degree of automation provided in the aircraft systems to afford (after failures or malfunctions) automatic crossover or isolation of difficulties to minimize the need for flight crew action to guard against loss of hydraulic or electric power to flight controls or to other essential systems.
(8) The communications and navigation workload.
(9) The possibility of increased workload associated with any emergency that may lead to other emergencies.
(10) Incapacitation of a flight crewmember whenever the applicable operating rule requires a minimum flight crew of at least two pilots.
(c) Kind of operation authorized. The determination of the kind of operation authorized requires consideration of the operating rules under which the airplane will be operated. Unless an applicant desires approval for a more limited kind of operation. It is assumed that each airplane certificated under this Part will operate under IFR conditions.
[Amdt. 25–3, 30 FR 6067, Apr. 29, 1965]
Appendix E to Part 25
I—Limited Weight Credit For Airplanes Equipped With Standby Power
(a) Each applicant for an increase in the maximum certificated takeoff and landing weights of an airplane equipped with a type-certificated standby power rocket engine may obtain an increase as specified in paragraph (b) if—
(1) The installation of the rocket engine has been approved and it has been established by flight test that the rocket engine and its controls can be operated safely and reliably at the increase in maximum weight; and
(2) The Airplane Flight Manual, or the placard, markings or manuals required in place thereof, set forth in addition to any other operating limitations the Administrator may require, the increased weight approved under this regulation and a prohibition against the operation of the airplane at the approved increased weight when—
(i) The installed standby power rocket engines have been stored or installed in excess of the time limit established by the manufacturer of the rocket engine (usually stenciled on the engine casing); or
(ii) The rocket engine fuel has been expended or discharged.
(b) The currently approved maximum takeoff and landing weights at which an airplane is certificated without a standby power rocket engine installation may be increased by an amount that does not exceed any of the following:
(1) An amount equal in pounds to 0.014 IN, where I is the maximum usable impulse in pounds-seconds available from each standby power rocket engine and N is the number of rocket engines installed.
(2) An amount equal to 5 percent of the maximum certificated weight approved in accordance with the applicable airworthiness regulations without standby power rocket engines installed.
(3) An amount equal to the weight of the rocket engine installation.
(4) An amount that, together with the currently approved maximum weight, would equal the maximum structural weight established for the airplane without standby rocket engines installed.
II—Performance Credit for Transport Category Airplanes Equipped With Standby Power
The Administrator may grant performance credit for the use of standby power on transport category airplanes. However, the performance credit applies only to the maximum certificated takeoff and landing weights, the takeoff distance, and the takeoff paths, and may not exceed that found by the Administrator to result in an overall level of safety in the takeoff, approach, and landing regimes of flight equivalent to that prescribed in the regulations under which the airplane was originally certificated without standby power. For the purposes of this appendix, “standby power” is power or thrust, or both, obtained from rocket engines for a relatively short period and actuated only in cases of emergency. The following provisions apply:
(1) Takeoff; general. The takeoff data prescribed in paragraphs (2) and (3) of this appendix must be determined at all weights and altitudes, and at ambient temperatures if applicable, at which performance credit is to be applied.
(2) Takeoff path.
(a) The one-engine-inoperative takeoff path with standby power in use must be determined in accordance with the performance requirements of the applicable airworthiness regulations.
(b) The one-engine-inoperative takeoff path (excluding that part where the airplane is on or just above the takeoff surface) determined in accordance with paragraph (a) of this section must lie above the one-engine-inoperative takeoff path without standby power at the maximum takeoff weight at which all of the applicable air-worthiness requirements are met. For the purpose of this comparison, the flight path is considered to extend to at least a height of 400 feet above the takeoff surface.
(c) The takeoff path with all engines operating, but without the use of standby power, must reflect a conservatively greater overall level of performance than the one-engine-inoperative takeoff path established in accordance with paragraph (a) of this section. The margin must be established by the Administrator to insure safe day-to-day operations, but in no case may it be less than 15 percent. The all-engines-operating takeoff path must be determined by a procedure consistent with that established in complying with paragraph (a) of this section.
(d) For reciprocating-engine-powered airplanes, the takeoff path to be scheduled in the Airplane Flight Manual must represent the one-engine-operative takeoff path determined in accordance with paragraph (a) of this section and modified to reflect the procedure (see paragraph (6)) established by the applicant for flap retraction and attainment of the en route speed. The scheduled takeoff path must have a positive slope at all points of the airborne portion and at no point must it lie above the takeoff path specified in paragraph (a) of this section.
(3) Takeoff distance. The takeoff distance must be the horizontal distance along the one-engine-inoperative take off path determined in accordance with paragraph (2)(a) from the start of the takeoff to the point where the airplane attains a height of 50 feet above the takeoff surface for reciprocating-engine-powered airplanes and a height of 35 feet above the takeoff surface for turbine-powered airplanes.
(4) Maximum certificated takeoff weights. The maximum certificated takeoff weights must be determined at all altitudes, and at ambient temperatures, if applicable, at which performance credit is to be applied and may not exceed the weights established in compliance with paragraphs (a) and (b) of this section.
(a) The conditions of paragraphs (2)(b) through (d) must be met at the maximum certificated takeoff weight.
(b) Without the use of standby power, the airplane must meet all of the en route requirements of the applicable airworthiness regulations under which the airplane was originally certificated. In addition, turbine-powered airplanes without the use of standby power must meet the final takeoff climb requirements prescribed in the applicable airworthiness regulations.
(5) Maximum certificated landing weights.
(a) The maximum certificated landing weights (one-engine-inoperative approach and all-engine-operating landing climb) must be determined at all altitudes, and at ambient temperatures if applicable, at which performance credit is to be applied and must not exceed that established in compliance with paragraph (b) of this section.
(b) The flight path, with the engines operating at the power or thrust, or both, appropriate to the airplane configuration and with standby power in use, must lie above the flight path without standby power in use at the maximum weight at which all of the applicable airworthiness requirements are met. In addition, the flight paths must comply with subparagraphs (i) and (ii) of this paragraph.
(i) The flight paths must be established without changing the appropriate airplane configuration.
(ii) The flight paths must be carried out for a minimum height of 400 feet above the point where standby power is actuated.
(6) Airplane configuration, speed, and power and thrust; general. Any change in the airplane's configuration, speed, and power or thrust, or both, must be made in accordance with the procedures established by the applicant for the operation of the airplane in service and must comply with paragraphs (a) through (c) of this section. In addition, procedures must be established for the execution of balked landings and missed approaches.
(a) The Administrator must find that the procedure can be consistently executed in service by crews of average skill.
(b) The procedure may not involve methods or the use of devices which have not been proven to be safe and reliable.
(c) Allowances must be made for such time delays in the execution of the procedures as may be reasonably expected to occur during service.
(7) Installation and operation; standby power. The standby power unit and its installation must comply with paragraphs (a) and (b) of this section.
(a) The standby power unit and its installation must not adversely affect the safety of the airplane.
(b) The operation of the standby power unit and its control must have proven to be safe and reliable.
[Amdt. 25–6, 30 FR 8468, July 2, 1965]
Appendix F to Part 25
Part I—Test Criteria and Procedures for Showing Compliance with §25.853, or §25.855.
(a) Material test criteria—(1) Interior compartments occupied by crew or passengers. (i) Interior ceiling panels, interior wall panels, partitions, galley structure, large cabinet walls, structural flooring, and materials used in the construction of stowage compartments (other than underseat stowage compartments and compartments for stowing small items such as magazines and maps) must be self-extinguishing when tested vertically in accordance with the applicable portions of part I of this appendix. The average burn length may not exceed 6 inches and the average flame time after removal of the flame source may not exceed 15 seconds. Drippings from the test specimen may not continue to flame for more than an average of 3 seconds after falling.
(ii) Floor covering, textiles (including draperies and upholstery), seat cushions, padding, decorative and nondecorative coated fabrics, leather, trays and galley furnishings, electrical conduit, air ducting, joint and edge covering, liners of Class B and E cargo or baggage compartments, floor panels of Class B, C, D, or E cargo or baggage compartments, cargo covers and transparencies, molded and thermoformed parts, air ducting joints, and trim strips (decorative and chafing), that are constructed of materials not covered in subparagraph (iv) below, must be self-extinguishing when tested vertically in accordance with the applicable portions of part I of this appendix or other approved equivalent means. The average burn length may not exceed 8 inches, and the average flame time after removal of the flame source may not exceed 15 seconds. Drippings from the test specimen may not continue to flame for more than an average of 5 seconds after falling.
(iii) Motion picture film must be safety film meeting the Standard Specifications for Safety Photographic Film PHI.25 (available from the American National Standards Institute, 1430 Broadway, New York, NY 10018). If the film travels through ducts, the ducts must meet the requirements of subparagraph (ii) of this paragraph.
(iv) Clear plastic windows and signs, parts constructed in whole or in part of elastomeric materials, edge lighted instrument assemblies consisting of two or more instruments in a common housing, seat belts, shoulder harnesses, and cargo and baggage tiedown equipment, including containers, bins, pallets, etc., used in passenger or crew compartments, may not have an average burn rate greater than 2.5 inches per minute when tested horizontally in accordance with the applicable portions of this appendix.
(v) Except for small parts (such as knobs, handles, rollers, fasteners, clips, grommets, rub strips, pulleys, and small electrical parts) that would not contribute significantly to the propagation of a fire and for electrical wire and cable insulation, materials in items not specified in paragraphs (a)(1)(i), (ii), (iii), or (iv) of part I of this appendix may not have a burn rate greater than 4.0 inches per minute when tested horizontally in accordance with the applicable portions of this appendix.
(2) Cargo and baggage compartments not occupied by crew or passengers.
(i) [Reserved]
(ii) A cargo or baggage compartment defined in §25.857 as Class B or E must have a liner constructed of materials that meet the requirements of paragraph (a)(1)(ii) of part I of this appendix and separated from the airplane structure (except for attachments). In addition, such liners must be subjected to the 45 degree angle test. The flame may not penetrate (pass through) the material during application of the flame or subsequent to its removal. The average flame time after removal of the flame source may not exceed 15 seconds, and the average glow time may not exceed 10 seconds.
(iii) A cargo or baggage compartment defined in §25.857 as Class B, C, D, or E must have floor panels constructed of materials which meet the requirements of paragraph (a)(1)(ii) of part I of this appendix and which are separated from the airplane structure (except for attachments). Such panels must be subjected to the 45 degree angle test. The flame may not penetrate (pass through) the material during application of the flame or subsequent to its removal. The average flame time after removal of the flame source may not exceed 15 seconds, and the average glow time may not exceed 10 seconds.
(iv) Insulation blankets and covers used to protect cargo must be constructed of materials that meet the requirements of paragraph (a)(1)(ii) of part I of this appendix. Tiedown equipment (including containers, bins, and pallets) used in each cargo and baggage compartment must be constructed of materials that meet the requirements of paragraph (a)(1)(v) of part I of this appendix.
(3) Electrical system components. Insulation on electrical wire or cable installed in any area of the fuselage must be self-extinguishing when subjected to the 60 degree test specified in part I of this appendix. The average burn length may not exceed 3 inches, and the average flame time after removal of the flame source may not exceed 30 seconds. Drippings from the test specimen may not continue to flame for more than an average of 3 seconds after falling.
(b) Test Procedures—(1) Conditioning. Specimens must be conditioned to 70 ±5 F., and at 50 percent ±5 percent relative humidity until moisture equilibrium is reached or for 24 hours. Each specimen must remain in the conditioning environment until it is subjected to the flame.
(2) Specimen configuration. Except for small parts and electrical wire and cable insulation, materials must be tested either as section cut from a fabricated part as installed in the airplane or as a specimen simulating a cut section, such as a specimen cut from a flat sheet of the material or a model of the fabricated part. The specimen may be cut from any location in a fabricated part; however, fabricated units, such as sandwich panels, may not be separated for test. Except as noted below, the specimen thickness must be no thicker than the minimum thickness to be qualified for use in the airplane. Test specimens of thick foam parts, such as seat cushions, must be
(3) Apparatus. Except as provided in subparagraph (7) of this paragraph, tests must be conducted in a draft-free cabinet in accordance with Federal Test Method Standard 191 Model 5903 (revised Method 5902) for the vertical test, or Method 5906 for horizontal test (available from the General Services Administration, Business Service Center, Region 3, Seventh & D Streets SW., Washington, DC 20407). Specimens which are too large for the cabinet must be tested in similar draft-free conditions.
(4) Vertical test. A minimum of three specimens must be tested and results averaged. For fabrics, the direction of weave corresponding to the most critical flammability conditions must be parallel to the longest dimension. Each specimen must be supported vertically. The specimen must be exposed to a Bunsen or Tirrill burner with a nominal
(5) Horizontal test. A minimum of three specimens must be tested and the results averaged. Each specimen must be supported horizontally. The exposed surface, when installed in the aircraft, must be face down for the test. The specimen must be exposed to a Bunsen or Tirrill burner with a nominal
(6) Forty-five degree test. A minimum of three specimens must be tested and the results averaged. The specimens must be supported at an angle of 45° to a horizontal surface. The exposed surface when installed in the aircraft must be face down for the test. The specimens must be exposed to a Bunsen or Tirrill burner with a nominal
(7) Sixty degree test. A minimum of three specimens of each wire specification (make and size) must be tested. The specimen of wire or cable (including insulation) must be placed at an angle of 60° with the horizontal in the cabinet specified in subparagraph (3) of this paragraph with the cabinet door open during the test, or must be placed within a chamber approximately 2 feet high by 1 foot by 1 foot, open at the top and at one vertical side (front), and which allows sufficient flow of air for complete combustion, but which is free from drafts. The specimen must be parallel to and approximately 6 inches from the front of the chamber. The lower end of the specimen must be held rigidly clamped. The upper end of the specimen must pass over a pulley or rod and must have an appropriate weight attached to it so that the specimen is held tautly throughout the flammability test. The test specimen span between lower clamp and upper pulley or rod must be 24 inches and must be marked 8 inches from the lower end to indicate the central point for flame application. A flame from a Bunsen or Tirrill burner must be applied for 30 seconds at the test mark. The burner must be mounted underneath the test mark on the specimen, perpendicular to the specimen and at an angle of 30° to the vertical plane of the specimen. The burner must have a nominal bore of
(8) Burn length. Burn length is the distance from the original edge to the farthest evidence of damage to the test specimen due to flame impingement, including areas of partial or complete consumption, charring, or embrittlement, but not including areas sooted, stained, warped, or discolored, nor areas where material has shrunk or melted away from the heat source.
Part II—Flammability of Seat Cushions
(a) Criteria for Acceptance. Each seat cushion must meet the following criteria:
(1) At least three sets of seat bottom and seat back cushion specimens must be tested.
(2) If the cushion is constructed with a fire blocking material, the fire blocking material must completely enclose the cushion foam core material.
(3) Each specimen tested must be fabricated using the principal components (i.e., foam core, flotation material, fire blocking material, if used, and dress covering) and assembly processes (representative seams and closures) intended for use in the production articles. If a different material combination is used for the back cushion than for the bottom cushion, both material combinations must be tested as complete specimen sets, each set consisting of a back cushion specimen and a bottom cushion specimen. If a cushion, including outer dress covering, is demonstrated to meet the requirements of this appendix using the oil burner test, the dress covering of that cushion may be replaced with a similar dress covering provided the burn length of the replacement covering, as determined by the test specified in §25.853(c), does not exceed the corresponding burn length of the dress covering used on the cushion subjected to the oil burner test.
(4) For at least two-thirds of the total number of specimen sets tested, the burn length from the burner must not reach the side of the cushion opposite the burner. The burn length must not exceed 17 inches. Burn length is the perpendicular distance from the inside edge of the seat frame closest to the burner to the farthest evidence of damage to the test specimen due to flame impingement, including areas of partial or complete consumption, charring, or embrittlement, but not including areas sooted, stained, warped, or discolored, or areas where material has shrunk or melted away from the heat source.
(5) The average percentage weight loss must not exceed 10 percent. Also, at least two-thirds of the total number of specimen sets tested must not exceed 10 percent weight loss. All droppings falling from the cushions and mounting stand are to be discarded before the after-test weight is determined. The percentage weight loss for a specimen set is the weight of the specimen set before testing less the weight of the specimen set after testing expressed as the percentage of the weight before testing.
(b) Test Conditions. Vertical air velocity should average 25 fpm±10 fpm at the top of the back seat cushion. Horizontal air velocity should be below 10 fpm just above the bottom seat cushion. Air velocities should be measured with the ventilation hood operating and the burner motor off.
(c) Test Specimens. (1) For each test, one set of cushion specimens representing a seat bottom and seat back cushion must be used.
(2) The seat bottom cushion specimen must be 18 ±
(3) The seat back cushion specimen must be 18 ±
(4) The specimens must be conditioned at 70 ±5 °F (21 ±2 °C) 55%±10% relative humidity for at least 24 hours before testing.
(d) Test Apparatus. The arrangement of the test apparatus is shown in Figures 1 through 5 and must include the components described in this section. Minor details of the apparatus may vary, depending on the model burner used.
(1) Specimen Mounting Stand. The mounting stand for the test specimens consists of steel angles, as shown in Figure 1. The length of the mounting stand legs is 12 ±
(2) Test Burner. The burner to be used in testing must—
(i) Be a modified gun type;
(ii) Have an 80-degree spray angle nozzle nominally rated for 2.25 gallons/hour at 100 psi;
(iii) Have a 12-inch (305 mm) burner cone installed at the end of the draft tube, with an opening 6 inches (152 mm) high and 11 inches (280 mm) wide, as shown in Figure 3; and
(iv) Have a burner fuel pressure regulator that is adjusted to deliver a nominal 2.0 gallon/hour of # 2 Grade kerosene or equivalent required for the test.
Burner models which have been used successfully in testing are the Lennox Model OB–32, Carlin Model 200 CRD, and Park Model DPL 3400. FAA published reports pertinent to this type of burner are: (1) Powerplant Enginering Report No. 3A, Standard Fire Test Apparatus and Procedure for Flexible Hose Assemblies, dated March 1978; and (2) Report No. DOT/FAA/RD/76/213, Reevaluation of Burner Characteristics for Fire Resistance Tests, dated January 1977.
(3) Calorimeter.
(i) The calorimeter to be used in testing must be a (0–15.0 BTU/ft2 -sec. 0–17.0 W/cm2 ) calorimeter, accurate ±3%, mounted in a 6-inch by 12-inch (152 by 305 mm) by
(ii) Because crumbling of the insulating board with service can result in misalignment of the calorimeter, the calorimeter must be monitored and the mounting shimmed, as necessary, to ensure that the calorimeter face is flush with the exposed plane of the insulating board in a plane parallel to the exit of the test burner cone.
(4) Thermocouples. The seven thermocouples to be used for testing must be
(5) Apparatus Arrangement. The test burner must be mounted on a suitable stand to position the exit of the burner cone a distance of 4 ±
(6) Data Recording. A recording potentiometer or other suitable calibrated instrument with an appropriate range must be used to measure and record the outputs of the calorimeter and the thermocouples.
(7) Weight Scale. Weighing Device—A device must be used that with proper procedures may determine the before and after test weights of each set of seat cushion specimens within 0.02 pound (9 grams). A continuous weighing system is preferred.
(8) Timing Device. A stopwatch or other device (calibrated to ±1 second) must be used to measure the time of application of the burner flame and self-extinguishing time or test duration.
(e) Preparation of Apparatus. Before calibration, all equipment must be turned on and the burner fuel must be adjusted as specified in paragraph (d)(2).
(f) Calibration. To ensure the proper thermal output of the burner, the following test must be made:
(1) Place the calorimeter on the test stand as shown in Figure 4 at a distance of 4 ±
(2) Turn on the burner, allow it to run for 2 minutes for warmup, and adjust the burner air intake damper to produce a reading of 10.5 ±0.5 BTU/ft2 -sec. (11.9 ±0.6 w/cm2 ) on the calorimeter to ensure steady state conditions have been achieved. Turn off the burner.
(3) Replace the calorimeter with the thermocouple rake (Figure 5).
(4) Turn on the burner and ensure that the thermocouples are reading 1900 ±100 °F (1038 ±38 °C) to ensure steady state conditions have been achieved.
(5) If the calorimeter and thermocouples do not read within range, repeat steps in paragraphs 1 through 4 and adjust the burner air intake damper until the proper readings are obtained. The thermocouple rake and the calorimeter should be used frequently to maintain and record calibrated test parameters. Until the specific apparatus has demonstrated consistency, each test should be calibrated. After consistency has been confirmed, several tests may be conducted with the pre-test calibration before and a calibration check after the series.
(g) Test Procedure. The flammability of each set of specimens must be tested as follows:
(1) Record the weight of each set of seat bottom and seat back cushion specimens to be tested to the nearest 0.02 pound (9 grams).
(2) Mount the seat bottom and seat back cushion test specimens on the test stand as shown in Figure 2, securing the seat back cushion specimen to the test stand at the top.
(3) Swing the burner into position and ensure that the distance from the exit of the burner cone to the side of the seat bottom cushion specimen is 4 ±
(4) Swing the burner away from the test position. Turn on the burner and allow it to run for 2 minutes to provide adequate warmup of the burner cone and flame stabilization.
(5) To begin the test, swing the burner into the test position and simultaneously start the timing device.
(6) Expose the seat bottom cushion specimen to the burner flame for 2 minutes and then turn off the burner. Immediately swing the burner away from the test position. Terminate test 7 minutes after initiating cushion exposure to the flame by use of a gaseous extinguishing agent (i.e., Halon or CO2).
(7) Determine the weight of the remains of the seat cushion specimen set left on the mounting stand to the nearest 0.02 pound (9 grams) excluding all droppings.
(h) Test Report. With respect to all specimen sets tested for a particular seat cushion for which testing of compliance is performed, the following information must be recorded:
(1) An identification and description of the specimens being tested.
(2) The number of specimen sets tested.
(3) The initial weight and residual weight of each set, the calculated percentage weight loss of each set, and the calculated average percentage weight loss for the total number of sets tested.
(4) The burn length for each set tested.
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Part III—Test Method To Determine Flame Penetration Resistance of Cargo Compartment Liners.
(a) Criteria for Acceptance. (1) At least three specimens of cargo compartment sidewall or ceiling liner panels must be tested.
(2) Each specimen tested must simulate the cargo compartment sidewall or ceiling liner panel, including any design features, such as joints, lamp assemblies, etc., the failure of which would affect the capability of the liner to safely contain a fire.
(3) There must be no flame penetration of any specimen within 5 minutes after application of the flame source, and the peak temperature measured at 4 inches above the upper surface of the horizontal test sample must not exceed 400 °F.
(b) Summary of Method. This method provides a laboratory test procedure for measuring the capability of cargo compartment lining materials to resist flame penetration with a 2 gallon per hour (GPH) #2 Grade kerosene or equivalent burner fire source. Ceiling and sidewall liner panels may be tested individually provided a baffle is used to simulate the missing panel. Any specimen that passes the test as a ceiling liner panel may be used as a sidewall liner panel.
(c) Test Specimens. (1) The specimen to be tested must measure 16 ±
(2) The specimens must be conditioned at 70 °F.±5 °F. (21 °C. ±2 °C.) and 55%±5% humidity for at least 24 hours before testing.
(d) Test Apparatus. The arrangement of the test apparatus, which is shown in Figure 3 of Part II and Figures 1 through 3 of this part of appendix F, must include the components described in this section. Minor details of the apparatus may vary, depending on the model of the burner used.
(1) Specimen Mounting Stand. The mounting stand for the test specimens consists of steel angles as shown in Figure 1.
(2) Test Burner. The burner to be used in tesing must—
(i) Be a modified gun type.
(ii) Use a suitable nozzle and maintain fuel pressure to yield a 2 GPH fuel flow. For example: an 80 degree nozzle nominally rated at 2.25 GPH and operated at 85 pounds per square inch (PSI) gage to deliver 2.03 GPH.
(iii) Have a 12 inch (305 mm) burner extension installed at the end of the draft tube with an opening 6 inches (152 mm) high and 11 inches (280 mm) wide as shown in Figure 3 of Part II of this appendix.
(iv) Have a burner fuel pressure regulator that is adjusted to deliver a nominal 2.0 GPH of #2 Grade kerosene or equivalent.
Burner models which have been used successfully in testing are the Lenox Model OB–32, Carlin Model 200 CRD and Park Model DPL. The basic burner is described in FAA Powerplant Engineering Report No. 3A, Standard Fire Test Apparatus and Procedure for Flexible Hose Assemblies, dated March 1978; however, the test settings specified in this appendix differ in some instances from those specified in the report.
(3) Calorimeter. (i) The calorimeter to be used in testing must be a total heat flux Foil Type Gardon Gage of an appropriate range (approximately 0 to 15.0 British thermal unit (BTU) per ft.2 sec., 0–17.0 watts/cm2 ). The calorimeter must be mounted in a 6 inch by 12 inch (152 by 305 mm) by
(ii) The insulating block must be monitored for deterioration and the mounting shimmed as necessary to ensure that the calorimeter face is parallel to the exit plane of the test burner cone.
(4) Thermocouples. The seven thermocouples to be used for testing must be
(5) Apparatus Arrangement. The test burner must be mounted on a suitable stand to position the exit of the burner cone a distance of 8 inches from the ceiling liner panel and 2 inches from the sidewall liner panel. The burner stand should have the capability of allowing the burner to be swung away from the test specimen during warm-up periods.
(6) Instrumentation. A recording potentiometer or other suitable instrument with an appropriate range must be used to measure and record the outputs of the calorimeter and the thermocouples.
(7) Timing Device. A stopwatch or other device must be used to measure the time of flame application and the time of flame penetration, if it occurs.
(e) Preparation of Apparatus. Before calibration, all equipment must be turned on and allowed to stabilize, and the burner fuel flow must be adjusted as specified in paragraph (d)(2).
(f) Calibration. To ensure the proper thermal output of the burner the following test must be made:
(1) Remove the burner extension from the end of the draft tube. Turn on the blower portion of the burner without turning the fuel or igniters on. Measure the air velocity using a hot wire anemometer in the center of the draft tube across the face of the opening. Adjust the damper such that the air velocity is in the range of 1550 to 1800 ft./min. If tabs are being used at the exit of the draft tube, they must be removed prior to this measurement. Reinstall the draft tube extension cone.
(2) Place the calorimeter on the test stand as shown in Figure 2 at a distance of 8 inches (203 mm) from the exit of the burner cone to simulate the position of the horizontal test specimen.
(3) Turn on the burner, allow it to run for 2 minutes for warm-up, and adjust the damper to produce a calorimeter reading of 8.0 ±0.5 BTU per ft.2 sec. (9.1 ±0.6 Watts/cm2 ).
(4) Replace the calorimeter with the thermocouple rake (see Figure 3).
(5) Turn on the burner and ensure that each of the seven thermocouples reads 1700 °F. ±100 °F. (927 °C. ±38 °C.) to ensure steady state conditions have been achieved. If the temperature is out of this range, repeat steps 2 through 5 until proper readings are obtained.
(6) Turn off the burner and remove the thermocouple rake.
(7) Repeat (1) to ensure that the burner is in the correct range.
(g) Test Procedure. (1) Mount a thermocouple of the same type as that used for calibration at a distance of 4 inches (102 mm) above the horizontal (ceiling) test specimen. The thermocouple should be centered over the burner cone.
(2) Mount the test specimen on the test stand shown in Figure 1 in either the horizontal or vertical position. Mount the insulating material in the other position.
(3) Position the burner so that flames will not impinge on the specimen, turn the burner on, and allow it to run for 2 minutes. Rotate the burner to apply the flame to the specimen and simultaneously start the timing device.
(4) Expose the test specimen to the flame for 5 minutes and then turn off the burner. The test may be terminated earlier if flame penetration is observed.
(5) When testing ceiling liner panels, record the peak temperature measured 4 inches above the sample.
(6) Record the time at which flame penetration occurs if applicable.
(h) Test Report. The test report must include the following:
(1) A complete description of the materials tested including type, manufacturer, thickness, and other appropriate data.
(2) Observations of the behavior of the test specimens during flame exposure such as delamination, resin ignition, smoke, ect., including the time of such occurrence.
(3) The time at which flame penetration occurs, if applicable, for each of the three specimens tested.
(4) Panel orientation (ceiling or sidewall).
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Part IV—Test Method To Determine the Heat Release Rate From Cabin Materials Exposed to Radiant Heat.
(a) Summary of Method. Three or more specimens representing the completed aircraft component are tested. Each test specimen is injected into an environmental chamber through which a constant flow of air passes. The specimen's exposure is determined by a radiant heat source adjusted to produce, on the specimen, the desired total heat flux of 3.5 W/cm2 . The specimen is tested with the exposed surface vertical. Combustion is initiated by piloted ignition. The combustion products leaving the chamber are monitored in order to calculate the release rate of heat.
(b) Apparatus. The Ohio State University (OSU) rate of heat release apparatus, as described below, is used. This is a modified version of the rate of heat release apparatus standardized by the American Society of Testing and Materials (ASTM), ASTM E–906.
(1) This apparatus is shown in Figures 1A and 1B of this part IV. All exterior surfaces of the apparatus, except the holding chamber, must be insulated with 1 inch (25 mm) thick, low density, high temperature, fiberglass board insulation. A gasketed door, through which the sample injection rod slides, must be used to form an airtight closure on the specimen hold chamber.
(2) Thermopile. The temperature difference between the air entering the environmental chamber and that leaving must be monitored by a thermopile having five hot, and five cold, 24-guage Chromel-Alumel junctions. The hot junctions must be spaced across the top of the exhaust stack, .38 inches (10 mm) below the top of the chimney. The thermocouples must have a .050 ±.010 inch (1.3 ±.3mm) diameter, ball-type, welded tip. One thermocouple must be located in the geometric center, with the other four located 1.18 inch (30 mm) from the center along the diagonal toward each of the corners (Figure 5 of this part IV). The cold junctions must be located in the pan below the lower air distribution plate (see paragraph (b)(4) of this part IV). Thermopile hot junctions must be cleared of soot deposits as needed to maintain the calibrated sensitivity.
(3) Radiation Source. A radiant heat source incorporating four Type LL silicon carbide elements, 20 inches (508 mm) long by .63 inch (16 mm) O.D., must be used, as shown in Figures 2A and 2B of this part IV. The heat source must have a nominal resistance of 1.4 ohms and be capable of generating a flux up to 100 kW/m2 . The silicone carbide elements must be mounted in the stainless steel panel box by inserting them through .63 inch (16 mm) holes in .03 inch (1 mm) thick ceramic fiber or calcium-silicate millboard. Locations of the holes in the pads and stainless steel cover plates are shown in Figure 2B of this part IV. The truncated diamond-shaped mask of .042±.002 inch (1.07±.05mm) stainless steel must be added to provide uniform heat flux density over the area occupied by the vertical sample.
(4) Air Distribution System. The air entering the environmental chamber must be distributed by a .25 inch (6.3 mm) thick aluminum plate having eight No. 4 drill-holes, located 2 inches (51 mm) from sides on 4 inch (102 mm) centers, mounted at the base of the environmental chamber. A second plate of 18 guage stainless steel having 120, evenly spaced, No. 28 drill holes must be mounted 6 inches (152 mm) above the aluminum plate. A well-regulated air supply is required. The air-supply manifold at the base of the pyramidal section must have 48, evenly spaced, No. 26 drill holes located .38 inch (10 mm) from the inner edge of the manifold, resulting in an airflow split of approximately three to one within the apparatus.
(5) Exhaust Stack. An exhaust stack, 5.25×2.75 inches (133×70 mm) in cross section, and 10 inches (254 mm) long, fabricated from 28 guage stainless steel must be mounted on the outlet of the pyramidal section. A. 1.0×3.0 inch (25×76 mm) baffle plate of .018±.002 inch (.50±.05 mm) stainless steel must be centered inside the stack, perpendicular to the air flow, 3 inches (76 mm) above the base of the stack.
(6) Specimen Holders. (i) The specimen must be tested in a vertical orientation. The specimen holder (Figure 3 of this part IV) must incorporate a frame that touches the specimen (which is wrapped with aluminum foil as required by paragraph (d)(3) of this Part) along only the .25 inch (6 mm) perimeter. A “V” shaped spring is used to hold the assembly together. A detachable .50×.50×5.91 inch (12×12×150 mm) drip pan and two .020 inch (.5 mm) stainless steel wires (as shown in Figure 3 of this part IV) must be used for testing materials prone to melting and dripping. The positioning of the spring and frame may be changed to accommodate different specimen thicknesses by inserting the retaining rod in different holes on the specimen holder.
(ii) Since the radiation shield described in ASTM E–906 is not used, a guide pin must be added to the injection mechanism. This fits into a slotted metal plate on the injection mechanism outside of the holding chamber. It can be used to provide accurate positioning of the specimen face after injection. The front surface of the specimen must be 3.9 inches (100 mm) from the closed radiation doors after injection.
(iii) The specimen holder clips onto the mounted bracket (Figure 3 of this part IV). The mounting bracket must be attached to the injection rod by three screws that pass through a wide-area washer welded onto a
(7) Calorimeter. A total-flux type calorimeter must be mounted in the center of a
(8) Pilot-Flame Positions. Pilot ignition of the specimen must be accomplished by simultaneously exposing the specimen to a lower pilot burner and an upper pilot burner, as described in paragraph (b)(8)(i) and (b)(8)(ii) or (b)(8)(iii) of this part IV, respectively. Since intermittent pilot flame extinguishment for more than 3 seconds would invalidate the test results, a spark ignitor may be installed to ensure that the lower pilot burner remains lighted.
(i) Lower Pilot Burner. The pilot-flame tubing must be .25 inch (6.3 mm) O.D., .03 inch (0.8mm) wall, stainless steel tubing. A mixture of 120 cm3 /min. of methane and 850 cm3 /min. of air must be fed to the lower pilot flame burner. The normal position of the end of the pilot burner tubing is .40 inch (10 mm) from and perpendicular to the exposed vertical surface of the specimen. The centerline at the outlet of the burner tubing must intersect the vertical centerline of the sample at a point .20 inch (5 mm) above the lower exposed edge of the specimen.
(ii) Standard Three-Hole Upper Pilot Burner. The pilot burner must be a straight length of .25 inch (6.3 mm) O.D., .03 inch (0.8 mm) wall, stainless steel tubing that is 14 inches (360 mm) long. One end of the tubing must be closed, and three No. 40 drill holes must be drilled into the tubing, 2.38 inch (60 mm) apart, for gas ports, all radiating in the same direction. The first hole must be .19 inch (5 mm) from the closed end of the tubing. The tube must be positioned .75 inch (19 mm) above and .75 inch (19 mm) behind the exposed upper edge of the specimen. The middle hole must be in the vertical plane perpendicular to the exposed surface of the specimen which passes through its vertical centerline and must be pointed toward the radiation source. The gas supplied to the burner must be methane and must be adjusted to produce flame lengths of 1 inch (25 mm).
(iii) Optional Fourteen-Hole Upper Pilot Burner. This burner may be used in lieu of the standard three-hole burner described in paragraph (b)(8)(ii) of this part IV. The pilot burner must be a straight length of .25 inch (6.3 mm) O.D., .03 inch (0.8 mm) wall, stainless steel tubing that is 15.75 inches (400 mm) long. One end of the tubing must be closed, and 14 No. 59 drill holes must be drilled into the tubing, .50 inch (13 mm) apart, for gas ports, all radiating in the same direction. The first hole must be .50 inch (13 mm) from the closed end of the tubing. The tube must be positioned above the specimen holder so that the holes are placed above the specimen as shown in Figure 1B of this part IV. The fuel supplied to the burner must be methane mixed with air in a ratio of approximately 50/50 by volume. The total gas flow must be adjusted to produce flame lengths of 1 inch (25 mm). When the gas/air ratio and the flow rate are properly adjusted, approximately .25 inch (6 mm) of the flame length appears yellow in color.
(c) Calibration of Equipment. (1) Heat Release Rate. A calibration burner, as shown in Figure 4, must be placed over the end of the lower pilot flame tubing using a gas tight connection. The flow of gas to the pilot flame must be at least 99 percent methane and must be accurately metered. Prior to usage, the wet test meter must be properly leveled and filled with distilled water to the tip of the internal pointer while no gas is flowing. Ambient temperature and pressure of the water are based on the internal wet test meter temperature. A baseline flow rate of approximately 1 liter/min. must be set and increased to higher preset flows of 4, 6, 8, 6 and 4 liters/min. Immediately prior to recording methane flow rates, a flow rate of 8 liters/min. must be used for 2 minutes to precondition the chamber. This is not recorded as part of calibration. The rate must be determined by using a stopwatch to time a complete revolution of the wet test meter for both the baseline and higher flow, with the flow returned to baseline before changing to the next higher flow. The thermopile baseline voltage must be measured. The gas flow to the burner must be increased to the higher preset flow and allowed to burn for 2.0 minutes, and the thermopile voltage must be measured. The sequence must be repeated until all five values have been determined. The average of the five values must be used as the calibration factor. The procedure must be repeated if the percent relative standard deviation is greater than 5 percent. Calculations are shown in paragraph (f) of this part IV.
(2) Flux Uniformity. Uniformity of flux over the specimen must be checked periodically and after each heating element change to determine if it is within acceptable limits of plus or minus 5 percent.
(3) As noted in paragraph (b)(2) of this part IV, thermopile hot junctions must be cleared of soot deposits as needed to maintain the calibrated sensitivity.
(d) Preparation of Test Specimens. (1) The test specimens must be representative of the aircraft component in regard to materials and construction methods. The standard size for the test specimens is 5.91 ±.03 × 5.91 ±.03 inches (149 ±1 × 149 ±1 mm). The thickness of the specimen must be the same as that of the aircraft component it represents up to a maximum thickness of 1.75 inches (45 mm). Test specimens representing thicker components must be 1.75 inches (45 mm).
(2) Conditioning. Specimens must be conditioned as described in Part 1 of this appendix.
(3) Mounting. Each test specimen must be wrapped tightly on all sides of the specimen, except for the one surface that is exposed with a single layer of .001 inch (.025 mm) aluminum foil.
(e) Procedure. (1) The power supply to the radiant panel must be set to produce a radiant flux of 3.5 ±.05 W/cm2 , as measured at the point the center of the specimen surface will occupy when positioned for the test. The radiant flux must be measured after the air flow through the equipment is adjusted to the desired rate.
(2) After the pilot flames are lighted, their position must be checked as described in paragraph (b)(8) of this part IV.
(3) Air flow through the apparatus must be controlled by a circular plate orifice located in a 1.5 inch (38.1 mm) I.D. pipe with two pressure measuring points, located 1.5 inches (38 mm) upstream and .75 inches (19 mm) downstream of the orifice plate. The pipe must be connected to a manometer set at a pressure differential of 7.87 inches (200 mm) of Hg. (See Figure 1B of this part IV.) The total air flow to the equipment is approximately .04 m3 /seconds. The stop on the vertical specimen holder rod must be adjusted so that the exposed surface of the specimen is positioned 3.9 inches (100 mm) from the entrance when injected into the environmental chamber.
(4) The specimen must be placed in the hold chamber with the radiation doors closed. The airtight outer door must be secured, and the recording devices must be started. The specimen must be retained in the hold chamber for 60 seconds, plus or minus 10 seconds, before injection. The thermopile “zero” value must be determined during the last 20 seconds of the hold period. The sample must not be injected before completion of the “zero” value determination.
(5) When the specimen is to be injected, the radiation doors must be opened. After the specimen is injected into the environmental chamber, the radiation doors must be closed behind the specimen.
(6) [Reserved]
(7) Injection of the specimen and closure of the inner door marks time zero. A record of the thermopile output with at least one data point per second must be made during the time the specimen is in the environmental chamber.
(8) The test duration is five minutes. The lower pilot burner and the upper pilot burner must remain lighted for the entire duration of the test, except that there may be intermittent flame extinguishment for periods that do not exceed 3 seconds. Furthermore, if the optional three-hole upper burner is used, at least two flamelets must remain lighted for the entire duration of the test, except that there may be intermittent flame extinguishment of all three flamelets for periods that do not exceed 3 seconds.
(9) A minimum of three specimens must be tested.
(f) Calculations. (1) The calibration factor is calculated as follows:
F0=flow of methane at baseline (1pm)
F1=higher preset flow of methane (1pm)
V0=thermopile voltage at baseline (mv)
V1=thermopile voltage at higher flow (mv)
Ta=Ambient temperature (K)
P=Ambient pressure (mm Hg)
Pv=Water vapor pressure (mm Hg)
(2) Heat release rates may be calculated from the reading of the thermopile output voltage at any instant of time as:
HRR=heat release rate (kw/m2 )
Vb=baseline voltage (mv)
Vm=measured thermopile voltage (mv)
Kh=calibration factor (kw/mv)
(3) The integral of the heat release rate is the total heat release as a function of time and is calculated by multiplying the rate by the data sampling frequency in minutes and summing the time from zero to two minutes.
(g) Criteria. The total positive heat release over the first two minutes of exposure for each of the three or more samples tested must be averaged, and the peak heat release rate for each of the samples must be averaged. The average total heat release must not exceed 65 kilowatt-minutes per square meter, and the average peak heat release rate must not exceed 65 kilowatts per square meter.
(h) Report. The test report must include the following for each specimen tested:
(1) Description of the specimen.
(2) Radiant heat flux to the specimen, expressed in W/cm2 .
(3) Data giving release rates of heat (in kW/m2 ) as a function of time, either graphically or tabulated at intervals no greater than 10 seconds. The calibration factor (kn) must be recorded.
(4) If melting, sagging, delaminating, or other behavior that affects the exposed surface area or the mode of burning occurs, these behaviors must be reported, together with the time at which such behaviors were observed.
(5) The peak heat release and the 2-minute integrated heat release rate must be reported.
Figures to Part IV of Appendix F
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Part V. Test Method To Determine the Smoke Emission Characteristics of Cabin Materials
(a) Summary of Method. The specimens must be constructed, conditioned, and tested in the flaming mode in accordance with American Society of Testing and Materials (ASTM) Standard Test Method ASTM F814–83.
(b) Acceptance Criteria. The specific optical smoke density (Ds), which is obtained by averaging the reading obtained after 4 minutes with each of the three specimens, shall not exceed 200.
Part VI—Test Method To Determine the Flammability and Flame Propagation Characteristics of Thermal/Acoustic Insulation Materials
Use this test method to evaluate the flammability and flame propagation characteristics of thermal/acoustic insulation when exposed to both a radiant heat source and a flame.
(a) Definitions.
“Flame propagation” means the furthest distance of the propagation of visible flame towards the far end of the test specimen, measured from the midpoint of the ignition source flame. Measure this distance after initially applying the ignition source and before all flame on the test specimen is extinguished. The measurement is not a determination of burn length made after the test.
“Radiant heat source” means an electric or air propane panel.
“Thermal/acoustic insulation” means a material or system of materials used to provide thermal and/or acoustic protection. Examples include fiberglass or other batting material encapsulated by a film covering and foams.
“Zero point” means the point of application of the pilot burner to the test specimen.
(b) Test apparatus.
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(1) Radiant panel test chamber. Conduct tests in a radiant panel test chamber (see figure 1 above). Place the test chamber under an exhaust hood to facilitate clearing the chamber of smoke after each test. The radiant panel test chamber must be an enclosure 55 inches (1397 mm) long by 19.5 (495 mm) deep by 28 (710 mm) to 30 inches (maximum) (762 mm) above the test specimen. Insulate the sides, ends, and top with a fibrous ceramic insulation, such as Kaowool MTM board. On the front side, provide a 52 by 12-inch (1321 by 305 mm) draft-free, high-temperature, glass window for viewing the sample during testing. Place a door below the window to provide access to the movable specimen platform holder. The bottom of the test chamber must be a sliding steel platform that has provision for securing the test specimen holder in a fixed and level position. The chamber must have an internal chimney with exterior dimensions of 5.1 inches (129 mm) wide, by 16.2 inches (411 mm) deep by 13 inches (330 mm) high at the opposite end of the chamber from the radiant energy source. The interior dimensions must be 4.5 inches (114 mm) wide by 15.6 inches (395 mm) deep. The chimney must extend to the top of the chamber (see figure 2).
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(2) Radiant heat source. Mount the radiant heat energy source in a cast iron frame or equivalent. An electric panel must have six, 3-inch wide emitter strips. The emitter strips must be perpendicular to the length of the panel. The panel must have a radiation surface of 12
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(i) Electric radiant panel. The radiant panel must be 3-phase and operate at 208 volts. A single-phase, 240 volt panel is also acceptable. Use a solid-state power controller and microprocessor-based controller to set the electric panel operating parameters.
(ii) Gas radiant panel. Use propane (liquid petroleum gas—2.1 UN 1075) for the radiant panel fuel. The panel fuel system must consist of a venturi-type aspirator for mixing gas and air at approximately atmospheric pressure. Provide suitable instrumentation for monitoring and controlling the flow of fuel and air to the panel. Include an air flow gauge, an air flow regulator, and a gas pressure gauge.
(iii) Radiant panel placement. Mount the panel in the chamber at 30° to the horizontal specimen plane, and 7
(3) Specimen holding system.
(i) The sliding platform serves as the housing for test specimen placement. Brackets may be attached (via wing nuts) to the top lip of the platform in order to accommodate various thicknesses of test specimens. Place the test specimens on a sheet of Kaowool MTM board or 1260 Standard Board (manufactured by Thermal Ceramics and available in Europe), or equivalent, either resting on the bottom lip of the sliding platform or on the base of the brackets. It may be necessary to use multiple sheets of material based on the thickness of the test specimen (to meet the sample height requirement). Typically, these non-combustible sheets of material are available in
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(ii) Attach a
(iii) Place the test specimen horizontally on the non-combustible board(s). Place a steel retaining/securing frame fabricated of mild steel, having a thickness of
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(4) Pilot Burner. The pilot burner used to ignite the specimen must be a BernzomaticTM commercial propane venturi torch with an axially symmetric burner tip and a propane supply tube with an orifice diameter of 0.006 inches (0.15 mm). The length of the burner tube must be 2
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(5) Thermocouples. Install a 24 American Wire Gauge (AWG) Type K (Chromel-Alumel) thermocouple in the test chamber for temperature monitoring. Insert it into the chamber through a small hole drilled through the back of the chamber. Place the thermocouple so that it extends 11 inches (279 mm) out from the back of the chamber wall, 11
(6) Calorimeter. The calorimeter must be a one-inch cylindrical water-cooled, total heat flux density, foil type Gardon Gage that has a range of 0 to 5 BTU/ft2 -second (0 to 5.7 Watts/cm2 ).
(7) Calorimeter calibration specification and procedure.
(i) Calorimeter specification.
(A) Foil diameter must be 0.25 ±0.005 inches (6.35 ±0.13 mm).
(B) Foil thickness must be 0.0005 ±0.0001 inches (0.013 ±0.0025 mm).
(C) Foil material must be thermocouple grade Constantan.
(D) Temperature measurement must be a Copper Constantan thermocouple.
(E) The copper center wire diameter must be 0.0005 inches (0.013 mm).
(F) The entire face of the calorimeter must be lightly coated with “Black Velvet” paint having an emissivity of 96 or greater.
(ii) Calorimeter calibration.
(A) The calibration method must be by comparison to a like standardized transducer.
(B) The standardized transducer must meet the specifications given in paragraph VI(b)(6) of this appendix.
(C) Calibrate the standard transducer against a primary standard traceable to the National Institute of Standards and Technology (NIST).
(D) The method of transfer must be a heated graphite plate.
(E) The graphite plate must be electrically heated, have a clear surface area on each side of the plate of at least 2 by 2 inches (51 by 51 mm), and be
(F) Center the 2 transducers on opposite sides of the plates at equal distances from the plate.
(G) The distance of the calorimeter to the plate must be no less than 0.0625 inches (1.6 mm), nor greater than 0.375 inches (9.5 mm).
(H) The range used in calibration must be at least 0–3.5 BTUs/ft2 second (0–3.9 Watts/cm2 ) and no greater than 0–5.7 BTUs/ft2 second (0–6.4 Watts/cm2 ).
(I) The recording device used must record the 2 transducers simultaneously or at least within
(8) Calorimeter fixture. With the sliding platform pulled out of the chamber, install the calorimeter holding frame and place a sheet of non-combustible material in the bottom of the sliding platform adjacent to the holding frame. This will prevent heat losses during calibration. The frame must be 13
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(9) Instrumentation. Provide a calibrated recording device with an appropriate range or a computerized data acquisition system to measure and record the outputs of the calorimeter and the thermocouple. The data acquisition system must be capable of recording the calorimeter output every second during calibration.
(10) Timing device. Provide a stopwatch or other device, accurate to ±1 second/hour, to measure the time of application of the pilot burner flame.
(c) Test specimens.
(1) Specimen preparation. Prepare and test a minimum of three test specimens. If an oriented film cover material is used, prepare and test both the warp and fill directions.
(2) Construction. Test specimens must include all materials used in construction of the insulation (including batting, film, scrim, tape etc.). Cut a piece of core material such as foam or fiberglass, and cut a piece of film cover material (if used) large enough to cover the core material. Heat sealing is the preferred method of preparing fiberglass samples, since they can be made without compressing the fiberglass (“box sample”). Cover materials that are not heat sealable may be stapled, sewn, or taped as long as the cover material is over-cut enough to be drawn down the sides without compressing the core material. The fastening means should be as continuous as possible along the length of the seams. The specimen thickness must be of the same thickness as installed in the airplane.
(3) Specimen Dimensions. To facilitate proper placement of specimens in the sliding platform housing, cut non-rigid core materials, such as fiberglass, 12
(d) Specimen conditioning. Condition the test specimens at 70 ±5 °F (21 ±2°C) and 55% ±10% relative humidity, for a minimum of 24 hours prior to testing.
(e) Apparatus Calibration.
(1) With the sliding platform out of the chamber, install the calorimeter holding frame. Push the platform back into the chamber and insert the calorimeter into the first hole (“zero” position). See figure 7. Close the bottom door located below the sliding platform. The distance from the centerline of the calorimeter to the radiant panel surface at this point must be 7.
(2) Ignite the panel. Adjust the fuel/air mixture to achieve 1.5 BTUs/ft2 -second ±5% (1.7 Watts/cm2 ±5%) at the “zero” position. If using an electric panel, set the power controller to achieve the proper heat flux. Allow the unit to reach steady state (this may take up to 1 hour). The pilot burner must be off and in the down position during this time.
(3) After steady-state conditions have been reached, move the calorimeter 2 inches (51 mm) from the “zero” position (first hole) to position 1 and record the heat flux. Move the calorimeter to position 2 and record the heat flux. Allow enough time at each position for the calorimeter to stabilize. Table 1 depicts typical calibration values at the three positions.
Table 1_Calibration Table------------------------------------------------------------------------ Position BTU's/ft\2\sec Watts/cm\2\------------------------------------------------------------------------``Zero'' Position........... 1.5 1.7Position 1.................. 1.51-1.50-1.49 1.71-1.70-1.69Position 2.................. 1.43-1.44 1.62-1.63------------------------------------------------------------------------
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(4) Open the bottom door, remove the calorimeter and holder fixture. Use caution as the fixture is very hot. (f) Test Procedure. (1) Ignite the pilot burner. Ensure that it is at least 2 inches (51 mm) above the top of the platform. The burner must not contact the specimen until the test begins. (2) Place the test specimen in the sliding platform holder. Ensure that the test sample surface is level with the top of the platform. At “zero” point, the specimen surface must be 7 (3) Place the retaining/securing frame over the test specimen. It may be necessary (due to compression) to adjust the sample (up or down) in order to maintain the distance from the sample to the radiant panel (7 (4) Immediately push the sliding platform into the chamber and close the bottom door. (5) Bring the pilot burner flame into contact with the center of the specimen at the “zero” point and simultaneously start the timer. The pilot burner must be at a 27° angle with the sample and be approximately 
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(6) Leave the burner in position for 15 seconds and then remove to a position at least 2 inches (51 mm) above the specimen. (g) Report. (1) Identify and describe the test specimen. (2) Report any shrinkage or melting of the test specimen. (3) Report the flame propagation distance. If this distance is less than 2 inches, report this as a pass (no measurement required). (4) Report the after-flame time. (h) Requirements. (1) There must be no flame propagation beyond 2 inches (51 mm) to the left of the centerline of the pilot flame application. (2) The flame time after removal of the pilot burner may not exceed 3 seconds on any specimen. Part VII—Test Method To Determine the Burnthrough Resistance of Thermal/Acoustic Insulation Materials Use the following test method to evaluate the burnthrough resistance characteristics of aircraft thermal/acoustic insulation materials when exposed to a high intensity open flame. (a) Definitions. Burnthrough time means the time, in seconds, for the burner flame to penetrate the test specimen, and/or the time required for the heat flux to reach 2.0 Btu/ft2 sec (2.27 W/cm2 ) on the inboard side, at a distance of 12 inches (30.5 cm) from the front surface of the insulation blanket test frame, whichever is sooner. The burnthrough time is measured at the inboard side of each of the insulation blanket specimens. Insulation blanket specimen means one of two specimens positioned in either side of the test rig, at an angle of 30° with respect to vertical. Specimen set means two insulation blanket specimens. Both specimens must represent the same production insulation blanket construction and materials, proportioned to correspond to the specimen size. (b) Apparatus. (1) The arrangement of the test apparatus is shown in figures 1 and 2 and must include the capability of swinging the burner away from the test specimen during warm-up. 
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(2) Test burner. The test burner must be a modified gun-type such as the Park Model DPL 3400. Flame characteristics are highly dependent on actual burner setup. Parameters such as fuel pressure, nozzle depth, stator position, and intake airflow must be properly adjusted to achieve the correct flame output. 
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(i) Nozzle. A nozzle must maintain the fuel pressure to yield a nominal 6.0 gal/hr (0.378 L/min) fuel flow. A Monarch-manufactured 80° PL (hollow cone) nozzle nominally rated at 6.0 gal/hr at 100 lb/in2 (0.71 MPa) delivers a proper spray pattern. (ii) Fuel Rail. The fuel rail must be adjusted to position the fuel nozzle at a depth of 0.3125 inch (8 mm) from the end plane of the exit stator, which must be mounted in the end of the draft tube. (iii) Internal Stator. The internal stator, located in the middle of the draft tube, must be positioned at a depth of 3.75 inches (95 mm) from the tip of the fuel nozzle. The stator must also be positioned such that the integral igniters are located at an angle midway between the 10 and 11 o'clock position, when viewed looking into the draft tube. Minor deviations to the igniter angle are acceptable if the temperature and heat flux requirements conform to the requirements of paragraph VII(e) of this appendix. (iv) Blower Fan. The cylindrical blower fan used to pump air through the burner must measure 5.25 inches (133 mm) in diameter by 3.5 inches (89 mm) in width. (v) Burner cone. Install a 12 +0.125-inch (305 ±3 mm) burner extension cone at the end of the draft tube. The cone must have an opening 6 ±0.125-inch (152 ±3 mm) high and 11 ±0.125-inch (280 ±3 mm) wide (see figure 3). (vi) Fuel. Use JP–8, Jet A, or their international equivalent, at a flow rate of 6.0 ±0.2 gal/hr (0.378 ±0.0126 L/min). If this fuel is unavailable, ASTM K2 fuel (Number 2 grade kerosene) or ASTM D2 fuel (Number 2 grade fuel oil or Number 2 diesel fuel) are acceptable if the nominal fuel flow rate, temperature, and heat flux measurements conform to the requirements of paragraph VII(e) of this appendix. (vii) Fuel pressure regulator. Provide a fuel pressure regulator, adjusted to deliver a nominal 6.0 gal/hr (0.378 L/min) flow rate. An operating fuel pressure of 100 lb/in2 (0.71 MPa) for a nominally rated 6.0 gal/hr 80° spray angle nozzle (such as a PL type) delivers 6.0 ±0.2 gal/hr (0.378 ±0.0126 L/min). 
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(3) Calibration rig and equipment. (i) Construct individual calibration rigs to incorporate a calorimeter and thermocouple rake for the measurement of heat flux and temperature. Position the calibration rigs to allow movement of the burner from the test rig position to either the heat flux or temperature position with minimal difficulty. (ii) Calorimeter. The calorimeter must be a total heat flux, foil type Gardon Gage of an appropriate range such as 0–20 Btu/ft2 -sec (0–22.7 W/cm2 ), accurate to ±3% of the indicated reading. The heat flux calibration method must be in accordance with paragraph VI(b)(7) of this appendix. (iii) Calorimeter mounting. Mount the calorimeter in a 6- by 12- ±0.125 inch (152- by 305- ±3 mm) by 0.75 ±0.125 inch (19 mm ±3 mm) thick insulating block which is attached to the heat flux calibration rig during calibration (figure 4). Monitor the insulating block for deterioration and replace it when necessary. Adjust the mounting as necessary to ensure that the calorimeter face is parallel to the exit plane of the test burner cone. 
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(iv) Thermocouples. Provide seven (v) Air velocity meter. Use a vane-type air velocity meter to calibrate the velocity of air entering the burner. An Omega Engineering Model HH30A is satisfactory. Use a suitable adapter to attach the measuring device to the inlet side of the burner to prevent air from entering the burner other than through the measuring device, which would produce erroneously low readings. Use a flexible duct, measuring 4 inches wide (102 mm) by 20 feet long (6.1 meters), to supply fresh air to the burner intake to prevent damage to the air velocity meter from ingested soot. An optional airbox permanently mounted to the burner intake area can effectively house the air velocity meter and provide a mounting port for the flexible intake duct. (4) Test specimen mounting frame. Make the mounting frame for the test specimens of (5) Backface calorimeters. Mount two total heat flux Gardon type calorimeters behind the insulation test specimens on the back side (cold) area of the test specimen mounting frame as shown in figure 6. Position the calorimeters along the same plane as the burner cone centerline, at a distance of 4 inches (102 mm) from the vertical centerline of the test frame. 
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(i) The calorimeters must be a total heat flux, foil type Gardon Gage of an appropriate range such as 0–5 Btu/ft2 -sec (0–5.7 W/cm2 ), accurate to ±3% of the indicated reading. The heat flux calibration method must comply with paragraph VI(b)(7) of this appendix. (6) Instrumentation. Provide a recording potentiometer or other suitable calibrated instrument with an appropriate range to measure and record the outputs of the calorimeter and the thermocouples. (7) Timing device. Provide a stopwatch or other device, accurate to ±1%, to measure the time of application of the burner flame and burnthrough time. (8) Test chamber. Perform tests in a suitable chamber to reduce or eliminate the possibility of test fluctuation due to air movement. The chamber must have a minimum floor area of 10 by 10 feet (305 by 305 cm). (i) Ventilation hood. Provide the test chamber with an exhaust system capable of removing the products of combustion expelled during tests. (c) Test Specimens. (1) Specimen preparation. Prepare a minimum of three specimen sets of the same construction and configuration for testing. (2) Insulation blanket test specimen. (i) For batt-type materials such as fiberglass, the constructed, finished blanket specimen assemblies must be 32 inches wide by 36 inches long (81.3 by 91.4 cm), exclusive of heat sealed film edges. (ii) For rigid and other non-conforming types of insulation materials, the finished test specimens must fit into the test rig in such a manner as to replicate the actual in-service installation. (3) Construction. Make each of the specimens tested using the principal components (i.e., insulation, fire barrier material if used, and moisture barrier film) and assembly processes (representative seams and closures). (i) Fire barrier material. If the insulation blanket is constructed with a fire barrier material, place the fire barrier material in a manner reflective of the installed arrangement For example, if the material will be placed on the outboard side of the insulation material, inside the moisture film, place it the same way in the test specimen. (ii) Insulation material. Blankets that utilize more than one variety of insulation (composition, density, etc.) must have specimen sets constructed that reflect the insulation combination used. If, however, several blanket types use similar insulation combinations, it is not necessary to test each combination if it is possible to bracket the various combinations. (iii) Moisture barrier film. If a production blanket construction utilizes more than one type of moisture barrier film, perform separate tests on each combination. For example, if a polyimide film is used in conjunction with an insulation in order to enhance the burnthrough capabilities, also test the same insulation when used with a polyvinyl fluoride film. (iv) Installation on test frame. Attach the blanket test specimens to the test frame using 12 steel spring type clamps as shown in figure 7. Use the clamps to hold the blankets in place in both of the outer vertical formers, as well as the center vertical former (4 clamps per former). The clamp surfaces should measure 1 inch by 2 inches (25 by 51 mm). Place the top and bottom clamps 6 inches (15.2 cm) from the top and bottom of the test frame, respectively. Place the middle clamps 8 inches (20.3 cm) from the top and bottom clamps. 
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(Note: For blanket materials that cannot be installed in accordance with figure 7 above, the blankets must be installed in a manner approved by the FAA.) (v) Conditioning. Condition the specimens at 70° ±5 °F (21° ±2 °C) and 55% ±10% relative humidity for a minimum of 24 hours prior to testing. (d) Preparation of apparatus. (1) Level and center the frame assembly to ensure alignment of the calorimeter and/or thermocouple rake with the burner cone. (2) Turn on the ventilation hood for the test chamber. Do not turn on the burner blower. Measure the airflow of the test chamber using a vane anemometer or equivalent measuring device. The vertical air velocity just behind the top of the upper insulation blanket test specimen must be 100 ±50 ft/min (0.51 ±0.25 m/s). The horizontal air velocity at this point must be less than 50 ft/min (0.25 m/s). (3) If a calibrated flow meter is not available, measure the fuel flow rate using a graduated cylinder of appropriate size. Turn on the burner motor/fuel pump, after insuring that the igniter system is turned off. Collect the fuel via a plastic or rubber tube into the graduated cylinder for a 2-minute period. Determine the flow rate in gallons per hour. The fuel flow rate must be 6.0 ±0.2 gallons per hour (0.378 ±0.0126 L/min). (e) Calibration. (1) Position the burner in front of the calorimeter so that it is centered and the vertical plane of the burner cone exit is 4 ±0.125 inches (102 ±3 mm) from the calorimeter face. Ensure that the horizontal centerline of the burner cone is offset 1 inch below the horizontal centerline of the calorimeter (figure 8). Without disturbing the calorimeter position, rotate the burner in front of the thermocouple rake, such that the middle thermocouple (number 4 of 7) is centered on the burner cone. 
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Ensure that the horizontal centerline of the burner cone is also offset 1 inch below the horizontal centerline of the thermocouple tips. Re-check measurements by rotating the burner to each position to ensure proper alignment between the cone and the calorimeter and thermocouple rake. (Note: The test burner mounting system must incorporate “detents” that ensure proper centering of the burner cone with respect to both the calorimeter and the thermocouple rakes, so that rapid positioning of the burner can be achieved during the calibration procedure.) (2) Position the air velocity meter in the adapter or airbox, making certain that no gaps exist where air could leak around the air velocity measuring device. Turn on the blower/motor while ensuring that the fuel solenoid and igniters are off. Adjust the air intake velocity to a level of 2150 ft/min, (10.92 m/s) then turn off the blower/motor. (Note: The Omega HH30 air velocity meter measures 2.625 inches in diameter. To calculate the intake airflow, multiply the cross-sectional area (0.03758 ft2 ) by the air velocity (2150 ft/min) to obtain 80.80 ft3 /min. An air velocity meter other than the HH30 unit can be used, provided the calculated airflow of 80.80 ft3 /min (2.29 m3 /min) is equivalent.) (3) Rotate the burner from the test position to the warm-up position. Prior to lighting the burner, ensure that the calorimeter face is clean of soot deposits, and there is water running through the calorimeter. Examine and clean the burner cone of any evidence of buildup of products of combustion, soot, etc. Soot buildup inside the burner cone may affect the flame characteristics and cause calibration difficulties. Since the burner cone may distort with time, dimensions should be checked periodically. (4) While the burner is still rotated to the warm-up position, turn on the blower/motor, igniters and fuel flow, and light the burner. Allow it to warm up for a period of 2 minutes. Move the burner into the calibration position and allow 1 minute for calorimeter stabilization, then record the heat flux once every second for a period of 30 seconds. Turn off burner, rotate out of position, and allow to cool. Calculate the average heat flux over this 30-second duration. The average heat flux should be 16.0 ±0.8 Btu/ft2 sec (18.2 ±0.9 W/cm2 ). (5) Position the burner in front of the thermocouple rake. After checking for proper alignment, rotate the burner to the warm-up position, turn on the blower/motor, igniters and fuel flow, and light the burner. Allow it to warm up for a period of 2 minutes. Move the burner into the calibration position and allow 1 minute for thermocouple stabilization, then record the temperature of each of the 7 thermocouples once every second for a period of 30 seconds. Turn off burner, rotate out of position, and allow to cool. Calculate the average temperature of each thermocouple over this 30-second period and record. The average temperature of each of the 7 thermocouples should be 1900°F ±100°F (1038 ±56 °C). (6) If either the heat flux or the temperatures are not within the specified range, adjust the burner intake air velocity and repeat the procedures of paragraphs (4) and (5) above to obtain the proper values. Ensure that the inlet air velocity is within the range of 2150 ft/min ±50 ft/min (10.92 ±0.25 m/s). (7) Calibrate prior to each test until consistency has been demonstrated. After consistency has been confirmed, several tests may be conducted with calibration conducted before and after a series of tests. (f) Test procedure. (1) Secure the two insulation blanket test specimens to the test frame. The insulation blankets should be attached to the test rig center vertical former using four spring clamps positioned as shown in figure 7 (according to the criteria of paragraph (c)(4) or (c)(4)(i) of this part of this appendix). (2) Ensure that the vertical plane of the burner cone is at a distance of 4 ±0.125 inch (102 ±3 mm) from the outer surface of the horizontal stringers of the test specimen frame, and that the burner and test frame are both situated at a 30° angle with respect to vertical. (3) When ready to begin the test, direct the burner away from the test position to the warm-up position so that the flame will not impinge on the specimens prematurely. Turn on and light the burner and allow it to stabilize for 2 minutes. (4) To begin the test, rotate the burner into the test position and simultaneously start the timing device. (5) Expose the test specimens to the burner flame for 4 minutes and then turn off the burner. Immediately rotate the burner out of the test position. (6) Determine (where applicable) the burnthrough time, or the point at which the heat flux exceeds 2.0 Btu/ft2 -sec (2.27 W/cm2 ). (g) Report. (1) Identify and describe the specimen being tested. (2) Report the number of insulation blanket specimens tested. (3) Report the burnthrough time (if any), and the maximum heat flux on the back face of the insulation blanket test specimen, and the time at which the maximum occurred. (h) Requirements. (1) Each of the two insulation blanket test specimens must not allow fire or flame penetration in less than 4 minutes. (2) Each of the two insulation blanket test specimens must not allow more than 2.0 Btu/ft2 -sec (2.27 W/cm2 ) on the cold side of the insulation specimens at a point 12 inches (30.5 cm) from the face of the test rig. [Amdt. 25–32, 37 FR 3972, Feb. 24, 1972] Editorial Note: For Appendix G to Part 25—Continuous Gust Design CriteriaThe continuous gust design criteria in this appendix must be used in establishing the dynamic response of the airplane to vertical and lateral continuous turbulence unless a more rational criteria is used. The following gust load requirements apply to mission analysis and design envelope analysis: (a) The limit gust loads utilizing the continuous turbulence concept must be determined in accordance with the provisions of either paragraph (b) or paragraphs (c) and (d) of this appendix. (b) Design envelope analysis. The limit loads must be determined in accordance with the following: (1) All critical altitudes, weights, and weight distributions, as specified in §25.321(b), and all critical speeds within the ranges indicated in paragraph (b)(3) of this appendix must be considered. (2) Values of Ā (ratio of root-mean-square incremental load root-mean-square gust velocity) must be determined by dynamic analysis. The power spectral density of the atmospheric turbulence must be as given by the equation— where: φ=power-spectral density (ft./sec.)2 /rad./ft. σ=root-mean-square gust velocity, ft./sec. Ω=reduced frequency, radians per foot. L=2,500 ft. (3) The limit loads must be obtained by multiplying the Ā values determined by the dynamic analysis by the following values of the gust velocity U (i) At speed Vc: U (1) The transfer function of the new design should exhibit no unusual characteristics as compared to the similar design which will significantly affect response to turbulence; e.g., coalescence of modal response in the frequency regime which can result in a significant increase of loads. (2) The typical mission of the new airplane is substantially equivalent to that of the similar design. (3) The similar design should demonstrate the adequacy of the U (ii) At speed VB: U (iii) At speed VD: U (iv) At speeds between VB and Vc and between Vc and VD: U (4) When a stability augmentation system is included in the analysis, the effect of system nonlinearities on loads at the limit load level must be realistically or conservatively accounted for. (c) Mission analysis. Limit loads must be determined in accordance with the following: (1) The expected utilization of the airplane must be represented by one or more flight profiles in which the load distribution and the variation with time of speed, altitude, gross weight, and center of gravity position are defined. These profiles must be divided into mission segments or blocks, for analysis, and average or effective values of the pertinent parameters defined for each segment. (2) For each of the mission segments defined under paragraph (c)(1) of this appendix, values of Ā and No must be determined by analysis. Ā is defined as the ratio of root-mean-square incremental load to root-mean-square gust velocity and No is the radius of gyration of the load power spectral density function about zero frequency. The power spectral density of the atmospheric turbulence must be given by the equation set forth in paragraph (b)(2) of this appendix. (3) For each of the load and stress quantities selected, the frequency of exceedance must be determined as a function of load level by means of the equation— where— t=selected time interval. y=net value of the load or stress. Y N(y)=average number of exceedances of the indicated value of the load or stress in unit time. No, Ā=parameters determined by dynamic analysis as defined in paragraph (c)(2) of this appendix. P1, P2, b1, b2=parameters defining the probability distributions of root-mean-square gust velocity, to be read from Figures 1 and 2 of this appendix. The limit gust loads must be read from the frequency of exceedance curves at a frequency of exceedance of 2×10−5 exceedances per hour. Both positive and negative load directions must be considered in determining the limit loads. (4) If a stability augmentation system is utilized to reduce the gust loads, consideration must be given to the fraction of flight time that the system may be inoperative. The flight profiles of paragraph (c)(1) of this appendix must include flight with the system inoperative for this fraction of the flight time. When a stability augmentation system is included in the analysis, the effect of system nonlinearities on loads at the limit load level must be conservatively accounted for. (d) Supplementary design envelope analysis. In addition to the limit loads defined by paragraph (c) of this appendix, limit loads must also be determined in accordance with paragraph (b) of this appendix, except that— (1) In paragraph (b)(3)(i) of this appendix, the value of U (2) In paragraph (b) of this appendix, the reference to paragraphs (b)(3)(i) through (b)(3)(iii) of this appendix is to be understood as referring to the paragraph as modified by paragraph (d)(1). 
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[Amdt. 25–54, 45 FR 60173, Sept. 11, 1980] Appendix H to Part 25—Instructions for Continued AirworthinessH25.1 General. (a) This appendix specifies requirements for the preparation of Instructions for Continued Airworthiness as required by §25.1529. (b) The Instructions for Continued Airworthiness for each airplane must include the Instructions for Continued Airworthiness for each engine and propeller (hereinafter designated “products”), for each appliance required by this chapter, and any required information relating to the interface of those appliances and products with the airplane. If Instructions for Continued Airworthiness are not supplied by the manufacturer of an appliance or product installed in the airplane, the Instructions for Continued Airworthiness for the airplane must include the information essential to the continued airworthiness of the airplane. (c) The applicant must submit to the FAA a program to show how changes to the Instructions for Continued Airworthiness made by the applicant or by the manufacturers or products and appliances installed in the airplane will be distributed. H25.2 Format. (a) The Instructions for Continued Airworthiness must be in the form of a manual or manuals as appropriate for the quantity of data to be provided. (b) The format of the manual or manuals must provide for a practical arrangement. H25.3 Content. The contents of the manual or manuals must be prepared in the English language. The Instructions for Continued Airworthiness must contain the following manuals or sections, as appropriate, and information: (a) Airplane maintenance manual or section. (1) Introduction information that includes an explanation of the airplane's features and data to the extent necessary for maintenance or preventive maintenance. (2) A description of the airplane and its systems and installations including its engines, propellers, and appliances. (3) Basic control and operation information describing how the airplane components and systems are controlled and how they operate, including any special procedures and limitations that apply. (4) Servicing information that covers details regarding servicing points, capacities of tanks, reservoirs, types of fluids to be used, pressures applicable to the various systems, location of access panels for inspection and servicing, locations of lubrication points, lubricants to be used, equipment required for servicing, tow instructions and limitations, mooring, jacking, and leveling information. (b) Maintenance instructions. (1) Scheduling information for each part of the airplane and its engines, auxiliary power units, propellers, accessories, instruments, and equipment that provides the recommended periods at which they should be cleaned, inspected, adjusted, tested, and lubricated, and the degree of inspection, the applicable wear tolerances, and work recommended at these periods. However, the applicant may refer to an accessory, instrument, or equipment manufacturer as the source of this information if the applicant shows that the item has an exceptionally high degree of complexity requiring specialized maintenance techniques, test equipment, or expertise. The recommended overhaul periods and necessary cross references to the Airworthiness Limitations section of the manual must also be included. In addition, the applicant must include an inspection program that includes the frequency and extent of the inspections necessary to provide for the continued airworthiness of the airplane. (2) Troubleshooting information describing probable malfunctions, how to recognize those malfunctions, and the remedial action for those malfunctions. (3) Information describing the order and method of removing and replacing products and parts with any necessary precautions to be taken. (4) Other general procedural instructions including procedures for system testing during ground running, symmetry checks, weighing and determining the center of gravity, lifting and shoring, and storage limitations. (c) Diagrams of structural access plates and information needed to gain access for inspections when access plates are not provided. (d) Details for the application of special inspection techniques including radiographic and ultrasonic testing where such processes are specified. (e) Information needed to apply protective treatments to the structure after inspection. (f) All data relative to structural fasteners such as identification, discard recommendations, and torque values. (g) A list of special tools needed. H25.4 Airworthiness Limitations section. (a) The Instructions for Continued Airworthiness must contain a section titled Airworthiness Limitations that is segregated and clearly distinguishable from the rest of the document. This section must set forth— (1) Each mandatory replacement time, structural inspection interval, and related structural inspection procedures approved under §25.571; and (2) Each mandatory replacement time, inspection interval, related inspection procedure, and all critical design configuration control limitations approved under §25.981 for the fuel tank system. (b) If the Instructions for Continued Airworthiness consist of multiple documents, the section required by this paragraph must be included in the principal manual. This section must contain a legible statement in a prominent location that reads: “The Airworthiness Limitations section is FAA-approved and specifies maintenance required under §§43.16 and 91.403 of the Federal Aviation Regulations, unless an alternative program has been FAA approved.” [Amdt. 25–54, 45 FR 60177, Sept. 11, 1980, as amended by Amdt. 25–68, 54 FR 34329, Aug. 18, 1989; Amdt. 25–102, 66 FR 23130, May 7, 2001] Appendix I to Part 25—Installation of an Automatic Takeoff Thrust Control System (ATTCS)I25.1 General. (a) This appendix specifies additional requirements for installation of an engine power control system that automatically resets thrust or power on operating engine(s) in the event of any one engine failure during takeoff. (b) With the ATTCS and associated systems functioning normally as designed, all applicable requirements of Part 25, except as provided in this appendix, must be met without requiring any action by the crew to increase thrust or power. I25.2 Definitions. (a) Automatic Takeoff Thrust Control System (ATTCS). An ATTCS is defined as the entire automatic system used on takeoff, including all devices, both mechanical and electrical, that sense engine failure, transmit signals, actuate fuel controls or power levers or increase engine power by other means on operating engines to achieve scheduled thrust or power increases, and furnish cockpit information on system operation. (b) Critical Time Interval. When conducting an ATTCS takeoff, the critical time interval is between V1 minus 1 second and a point on the minimum performance, all-engine flight path where, assuming a simultaneous occurrence of an engine and ATTCS failure, the resulting minimum flight path thereafter intersects the Part 25 required actual flight path at no less than 400 feet above the takeoff surface. This time interval is shown in the following illustration: 
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I25.3 Performance and System Reliability Requirements. The applicant must comply with the performance and ATTCS reliability requirements as follows: (a) An ATTCS failure or a combination of failures in the ATTCS during the critical time interval: (1) Shall not prevent the insertion of the maximum approved takeoff thrust or power, or must be shown to be an improbable event. (2) Shall not result in a significant loss or reduction in thrust or power, or must be shown to be an extremely improbable event. (b) The concurrent existence of an ATTCS failure and an engine failure during the critical time interval must be shown to be extremely improbable. (c) All applicable performance requirements of Part 25 must be met with an engine failure occurring at the most critical point during takeoff with the ATTCS system functioning. I25.4 Thrust Setting. The initial takeoff thrust or power setting on each engine at the beginning of the takeoff roll may not be less than any of the following: (a) Ninety (90) percent of the thrust or power set by the ATTCS (the maximum takeoff thrust or power approved for the airplane under existing ambient conditions); (b) That required to permit normal operation of all safety-related systems and equipment dependent upon engine thrust or power lever position; or (c) That shown to be free of hazardous engine response characteristics when thrust or power is advanced from the initial takeoff thrust or power to the maximum approved takeoff thrust or power. I25.5 Powerplant Controls. (a) In addition to the requirements of §25.1141, no single failure or malfunction, or probable combination thereof, of the ATTCS, including associated systems, may cause the failure of any powerplant function necessary for safety. (b) The ATTCS must be designed to: (1) Apply thrust or power on the operating engine(s), following any one engine failure during takeoff, to achieve the maximum approved takeoff thrust or power without exceeding engine operating limits; (2) Permit manual decrease or increase in thrust or power up to the maximum takeoff thrust or power approved for the airplane under existing conditions through the use of the power lever. For airplanes equipped with limiters that automatically prevent engine operating limits from being exceeded under existing ambient conditions, other means may be used to increase the thrust or power in the event of an ATTCS failure provided the means is located on or forward of the power levers; is easily identified and operated under all operating conditions by a single action of either pilot with the hand that is normally used to actuate the power levers; and meets the requirements of §25.777 (a), (b), and (c); (3) Provide a means to verify to the flightcrew before takeoff that the ATTCS is in a condition to operate; and (4) Provide a means for the flightcrew to deactivate the automatic function. This means must be designed to prevent inadvertent deactivation. I25.6 Powerplant Instruments. In addition to the requirements of §25.1305: (a) A means must be provided to indicate when the ATTCS is in the armed or ready condition; and (b) If the inherent flight characteristics of the airplane do not provide adequate warning that an engine has failed, a warning system that is independent of the ATTCS must be provided to give the pilot a clear warning of any engine failure during takeoff. [Amdt. 25–62, 52 FR 43156, Nov. 9, 1987] Appendix J to Part 25—Emergency EvacuationThe following test criteria and procedures must be used for showing compliance with §25.803: (a) The emergency evacuation must be conducted with exterior ambient light levels of no greater than 0.3 foot-candles prior to the activation of the airplane emergency lighting system. The source(s) of the initial exterior ambient light level may remain active or illuminated during the actual demonstration. There must, however, be no increase in the exterior ambient light level except for that due to activation of the airplane emergency lighting system. (b) The airplane must be in a normal attitude with landing gear extended. (c) Unless the airplane is equipped with an off-wing descent means, stands or ramps may be used for descent from the wing to the ground. Safety equipment such as mats or inverted life rafts may be placed on the floor or ground to protect participants. No other equipment that is not part of the emergency evacuation equipment of the airplane may be used to aid the participants in reaching the ground. (d) Except as provided in paragraph (a) of this appendix, only the airplane's emergency lighting system may provide illumination. (e) All emergency equipment required for the planned operation of the airplane must be installed. (f) Each internal door or curtain must be in the takeoff configuration. (g) Each crewmember must be seated in the normally assigned seat for takeoff and must remain in the seat until receiving the signal for commencement of the demonstration. Each crewmember must be a person having knowledge of the operation of exits and emergency equipment and, if compliance with §121.291 is also being demonstrated, each flight attendant must be a member of a regularly scheduled line crew. (h) A representative passenger load of persons in normal health must be used as follows: (1) At least 40 percent of the passenger load must be female. (2) At least 35 percent of the passenger load must be over 50 years of age. (3) At least 15 percent of the passenger load must be female and over 50 years of age. (4) Three life-size dolls, not included as part of the total passenger load, must be carried by passengers to simulate live infants 2 years old or younger. (5) Crewmembers, mechanics, and training personnel, who maintain or operate the airplane in the normal course of their duties, may not be used as passengers. (i) No passenger may be assigned a specific seat except as the Administrator may require. Except as required by subparagraph (g) of this paragraph, no employee of the applicant may be seated next to an emergency exit. (j) Seat belts and shoulder harnesses (as required) must be fastened. (k) Before the start of the demonstration, approximately one-half of the total average amount of carry-on baggage, blankets, pillows, and other similar articles must be distributed at several locations in aisles and emergency exit access ways to create minor obstructions. (l) No prior indication may be given to any crewmember or passenger of the particular exits to be used in the demonstration. (m) The applicant may not practice, rehearse, or describe the demonstration for the participants nor may any participant have taken part in this type of demonstration within the preceding 6 months. (n) Prior to entering the demonstration aircraft, the passengers may also be advised to follow directions of crewmembers but may not be instructed on the procedures to be followed in the demonstration, except with respect to safety procedures in place for the demonstration or which have to do with the demonstration site. Prior to the start of the demonstration, the pre-takeoff passenger briefing required by §121.571 may be given. Flight attendants may assign demonstration subjects to assist persons from the bottom of a slide, consistent with their approved training program. (o) The airplane must be configured to prevent disclosure of the active emergency exits to demonstration participants in the airplane until the start of the demonstration. (p) Exits used in the demonstration must consist of one exit from each exit pair. The demonstration may be conducted with the escape slides, if provided, inflated and the exits open at the beginning of the demonstration. In this case, all exits must be configured such that the active exits are not disclosed to the occupants. If this method is used, the exit preparation time for each exit utilized must be accounted for, and exits that are not to be used in the demonstration must not be indicated before the demonstration has started. The exits to be used must be representative of all of the emergency exits on the airplane and must be designated by the applicant, subject to approval by the Administrator. At least one floor level exit must be used. (q) Except as provided in paragraph (c) of this section, all evacuees must leave the airplane by a means provided as part of the airplane's equipment. (r) The applicant's approved procedures must be fully utilized, except the flightcrew must take no active role in assisting others inside the cabin during the demonstration. (s) The evacuation time period is completed when the last occupant has evacuated the airplane and is on the ground. Provided that the acceptance rate of the stand or ramp is no greater than the acceptance rate of the means available on the airplane for descent from the wing during an actual crash situation, evacuees using stands or ramps allowed by paragraph (c) of this appendix are considered to be on the ground when they are on the stand or ramp. [Amdt. 25–72, 55 FR 29788, July 20, 1990, as amended by Amdt. 25–79, Aug. 26, 1993; Amdt. 25–117, 69 FR 67499, Nov. 17, 2004] Browse Previous |
