Title 14—Aeronautics and Space--Volume 1
PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT
§ 27.1309   Equipment, systems, and installations.
The equipment, systems, and installations whose functioning is required by this subchapter must be designed and installed to ensure that they perform their intended functions under any foreseeable operating condition. For any item of equipment or system whose failure has not been specifically addressed by another requirement in this chapter, the following requirements also apply:
(a) The design of each item of equipment, system, and installation must be analyzed separately and in relation to other rotorcraft systems and installations to determine and identify any failure that would affect the capability of the rotorcraft or the ability of the crew to perform their duties in all operating conditions.
(b) Each item of equipment, system, and installation must be designed and installed so that:
(1) The occurrence of any catastrophic failure condition is extremely improbable;
(2) The occurrence of any major failure condition is no more than improbable; and
(3) For the occurrence of any other failure condition between major and catastrophic, the probability of the failure condition must be inversely proportional to its consequences.
(c) A means to alert the crew in the event of a failure must be provided when an unsafe system operating condition exists and to enable them to take corrective action. Systems, controls, and associated monitoring and crew alerting means must be designed to minimize crew errors that could create additional hazards.
(d) Compliance with the requirements of this section must be shown by analysis and, where necessary, by ground, flight, or simulator tests. The analysis must account for:
(1) Possible modes of failure, including malfunctions and misleading data and input from external sources;
(2) The effect of multiple failures and latent failures;
(3) The resulting effects on the rotorcraft and occupants, considering the stage of flight and operating conditions; and
(4) The crew alerting cues and the corrective action required.
§ 27.1316   Electrical and electronic system lightning protection.
(a) Each electrical and electronic system that performs a function, for which failure would prevent the continued safe flight and landing of the rotorcraft, must be designed and installed so that—
(1) The function is not adversely affected during and after the time the rotorcraft is exposed to lightning; and
(2) The system automatically recovers normal operation of that function in a timely manner after the rotorcraft is exposed to lightning.
(b) For rotorcraft approved for instrument flight rules operation, each electrical and electronic system that performs a function, for which failure would reduce the capability of the rotorcraft or the ability of the flightcrew to respond to an adverse operating condition, must be designed and installed so that the function recovers normal operation in a timely manner after the rotorcraft is exposed to lightning.
§ 27.1317   High-intensity Radiated Fields (HIRF) Protection.
(a) Except as provided in paragraph (d) of this section, each electrical and electronic system that performs a function whose failure would prevent the continued safe flight and landing of the rotorcraft must be designed and installed so that—
(1) The function is not adversely affected during and after the time the rotorcraft is exposed to HIRF environment I, as described in appendix D to this part;
(2) The system automatically recovers normal operation of that function, in a timely manner, after the rotorcraft is exposed to HIRF environment I, as described in appendix D to this part, unless this conflicts with other operational or functional requirements of that system;
(3) The system is not adversely affected during and after the time the rotorcraft is exposed to HIRF environment II, as described in appendix D to this part; and
(4) Each function required during operation under visual flight rules is not adversely affected during and after the time the rotorcraft is exposed to HIRF environment III, as described in appendix D to this part.
(b) Each electrical and electronic system that performs a function whose failure would significantly reduce the capability of the rotorcraft or the ability of the flightcrew to respond to an adverse operating condition must be designed and installed so the system is not adversely affected when the equipment providing these functions is exposed to equipment HIRF test level 1 or 2, as described in appendix D to this part.
(c) Each electrical and electronic system that performs a function whose failure would reduce the capability of the rotorcraft or the ability of the flightcrew to respond to an adverse operating condition, must be designed and installed so the system is not adversely affected when the equipment providing these functions is exposed to equipment HIRF test level 3, as described in appendix D to this part.
(d) Before December 1, 2012, an electrical or electronic system that performs a function whose failure would prevent the continued safe flight and landing of a rotorcraft may be designed and installed without meeting the provisions of paragraph (a) provided—
(1) The system has previously been shown to comply with special conditions for HIRF, prescribed under § 21.16, issued before December 1, 2007;
(2) The HIRF immunity characteristics of the system have not changed since compliance with the special conditions was demonstrated; and
(3) The data used to demonstrate compliance with the special conditions is provided.
§ 27.1321   Arrangement and visibility.
(a) Each flight, navigation, and powerplant instrument for use by any pilot must be easily visible to him.
(b) For each multiengine rotorcraft, identical powerplant instruments must be located so as to prevent confusion as to which engine each instrument relates.
(c) Instrument panel vibration may not damage, or impair the readability or accuracy of, any instrument.
(d) If a visual indicator is provided to indicate malfunction of an instrument, it must be effective under all probable cockpit lighting conditions.
§ 27.1322   Warning, caution, and advisory lights.
If warning, caution or advisory lights are installed in the cockpit, they must, unless otherwise approved by the Administrator, be—
(a) Red, for warning lights (lights indicating a hazard which may require immediate corrective action):
(b) Amber, for caution lights (lights indicating the possible need for future corrective action);
(c) Green, for safe operation lights; and
(d) Any other color, including white, for lights not described in paragraphs (a) through (c) of this section, provided the color differs sufficiently from the colors prescribed in paragraphs (a) through (c) of this section to avoid possible confusion.
§ 27.1323   Airspeed indicating system.
(a) Each airspeed indicating instrument must be calibrated to indicate true airspeed (at sea level with a standard atmosphere) with a minimum practicable instrument calibration error when the corresponding pitot and static pressures are applied.
(b) The airspeed indicating system must be calibrated in flight at forward speeds of 20 knots and over.
(c) At each forward speed above 80 percent of the climbout speed, the airspeed indicator must indicate true airspeed, at sea level with a standard atmosphere, to within an allowable installation error of not more than the greater of—
(1) ±3 percent of the calibrated airspeed; or
(2) Five knots.
§ 27.1325   Static pressure systems.
(a) Each instrument with static air case connections must be vented so that the influence of rotorcraft speed, the opening and closing of windows, airflow variation, and moisture or other foreign matter does not seriously affect its accuracy.
(b) Each static pressure port must be designed and located in such manner that the correlation between air pressure in the static pressure system and true ambient atmospheric static pressure is not altered when the rotorcraft encounters icing conditions. An anti-icing means or an alternate source of static pressure may be used in showing compliance with this requirement. If the reading of the altimeter, when on the alternate static pressure system, differs from the reading of the altimeter when on the primary static system by more than 50 feet, a correction card must be provided for the alternate static system.
(c) Except as provided in paragraph (d) of this section, if the static pressure system incorporates both a primary and an alternate static pressure source, the means for selecting one or the other source must be designed so that—
(1) When either source is selected, the other is blocked off; and
(2) Both sources cannot be blocked off simultaneously.
(d) For unpressurized rotorcraft, paragraph (c)(1) of this section does not apply if it can be demonstrated that the static pressure system calibration, when either static pressure source is selected is not changed by the other static pressure source being open or blocked.
§ 27.1327   Magnetic direction indicator.
(a) Except as provided in paragraph (b) of this section—
(1) Each magnetic direction indicator must be installed so that its accuracy is not excessively affected by the rotorcraft's vibration or magnetic fields; and
(2) The compensated installation may not have a deviation, in level flight, greater than 10 degrees on any heading.
(b) A magnetic nonstabilized direction indicator may deviate more than 10 degrees due to the operation of electrically powered systems such as electrically heated windshields if either a magnetic stabilized direction indicator, which does not have a deviation in level flight greater than 10 degrees on any heading, or a gyroscopic direction indicator, is installed. Deviations of a magnetic nonstabilized direction indicator of more than 10 degrees must be placarded in accordance with § 27.1547(e).
§ 27.1329   Automatic pilot and flight guidance system.
For the purpose of this subpart, an automatic pilot and flight guidance system may consist of an autopilot, flight director, or a component that interacts with stability augmentation or trim.
(a) Each automatic pilot and flight guidance system must be designed so that it:
(1) Can be overpowered by one pilot to allow control of the rotorcraft;
(2) Provides a means to disengage the system, or any malfunctioning component of the system, by each pilot to prevent it from interfering with the control of the rotorcraft; and
(3) Provides a means to indicate to the flight crew its current mode of operation. Selector switch position is not acceptable as a means of indication.
(b) Unless there is automatic synchronization, each system must have a means to readily indicate to the pilot the alignment of the actuating device in relation to the control system it operates.
(c) Each manually operated control for the system's operation must be readily accessible to the pilots.
(d) The system must be designed so that, within the range of adjustment available to the pilot, it cannot produce hazardous loads on the rotorcraft, or create hazardous deviations in the flight path, under any flight condition appropriate to its use or in the event of a malfunction, assuming that corrective action begins within a reasonable period of time.
(e) If the automatic pilot and flight guidance system integrates signals from auxiliary controls or furnishes signals for operation of other equipment, there must be a means to prevent improper operation.
(f) If the automatic pilot system can be coupled to airborne navigation equipment, means must be provided to indicate to the pilots the current mode of operation. Selector switch position is not acceptable as a means of indication.
§ 27.1337   Powerplant instruments.
(a) Instruments and instrument lines. (1) Each powerplant instrument line must meet the requirements of §§ 27.- 961 and 27.993.
(2) Each line carrying flammable fluids under pressure must—
(i) Have restricting orifices or other safety devices at the source of pressure to prevent the escape of excessive fluid if the line fails; and
(ii) Be installed and located so that the escape of fluids would not create a hazard.
(3) Each powerplant instrument that utilizes flammable fluids must be installed and located so that the escape of fluid would not create a hazard.
(b) Fuel quantity indicator. Each fuel quantity indicator must be installed to clearly indicate to the flight crew the quantity of fuel in each tank in flight. In addition—
(1) Each fuel quantity indicator must be calibrated to read “zero” during level flight when the quantity of fuel remaining in the tank is equal to the unusable fuel supply determined under § 27.959;
(2) When two or more tanks are closely interconnected by a gravity feed system and vented, and when it is impossible to feed from each tank separately, at least one fuel quantity indicator must be installed; and
(3) Each exposed sight gauge used as a fuel quantity indicator must be protected against damage.
(c) Fuel flowmeter system. If a fuel flowmeter system is installed, each metering component must have a means for bypassing the fuel supply if malfunction of that component severely restricts fuel flow.
(d) Oil quantity indicator. There must be means to indicate the quantity of oil in each tank—
(1) On the ground (including during the filling of each tank); and
(2) In flight, if there is an oil transfer system or reserve oil supply system.
(e) Rotor drive system transmissions and gearboxes utilizing ferromagnetic materials must be equipped with chip detectors designed to indicate the presence of ferromagnetic particles resulting from damage or excessive wear. Chip detectors must—
(1) Be designed to provide a signal to the device required by § 27.1305(v) and be provided with a means to allow crewmembers to check, in flight, the function of each detector electrical circuit and signal.
(2) [Reserved]
§ 27.1351   General.
(a) Electrical system capacity. Electrical equipment must be adequate for its intended use. In addition—
(1) Electric power sources, their transmission cables, and their associated control and protective devices must be able to furnish the required power at the proper voltage to each load circuit essential for safe operation; and
(2) Compliance with paragraph (a)(1) of this section must be shown by an electrical load analysis, or by electrical measurements that take into account the electrical loads applied to the electrical system, in probable combinations and for probable durations.
(b) Function. For each electrical system, the following apply:
(1) Each system, when installed, must be—
(i) Free from hazards in itself, in its method of operation, and in its effects on other parts of the rotorcraft; and
(ii) Protected from fuel, oil, water, other detrimental substances, and mechanical damage.
(2) Electric power sources must function properly when connected in combination or independently.
(3) No failure or malfunction of any source may impair the ability of any remaining source to supply load circuits essential for safe operation.
(4) Each electric power source control must allow the independent operation of each source.
(c) Generating system. There must be at least one generator if the system supplies power to load circuits essential for safe operation. In addition—
(1) Each generator must be able to deliver its continuous rated power;
(2) Generator voltage control equipment must be able to dependably regulate each generator output within rated limits;
(3) Each generator must have a reverse current cutout designed to disconnect the generator from the battery and from the other generators when enough reverse current exists to damage that generator; and
(4) Each generator must have an overvoltage control designed and installed to prevent damage to the electrical system, or to equipment supplied by the electrical system, that could result if that generator were to develop an overvoltage condition.
(d) Instruments. There must be means to indicate to appropriate crewmembers the electric power system quantities essential for safe operation of the system. In addition—
(1) For direct current systems, an ammeter that can be switched into each generator feeder may be used; and
(2) If there is only one generator, the ammeter may be in the battery feeder.
(e) External power. If provisions are made for connecting external power to the rotorcraft, and that external power can be electrically connected to equipment other than that used for engine starting, means must be provided to ensure that no external power supply having a reverse polarity, or a reverse phase sequence, can supply power to the rotorcraft's electrical system.
§ 27.1353   Energy storage systems.
Energy storage systems must be designed and installed as follows:
(a) Energy storage systems must provide automatic protective features for any conditions that could prevent continued safe flight and landing.
(b) Energy storage systems must not emit any flammable, explosive, or toxic gases, smoke, or fluids that could accumulate in hazardous quantities within the rotorcraft.
(c) Corrosive fluids or gases that escape from the system must not damage surrounding structures, adjacent equipment, or systems necessary for continued safe flight and landing.
(d) The maximum amount of heat and pressure that can be generated during any operation or under any failure condition of the energy storage system or its individual components must not result in any hazardous effect on rotorcraft structure, equipment, or systems necessary for continued safe flight and landing.
(e) Energy storage system installations required for continued safe flight and landing of the rotorcraft must have monitoring features and a means to indicate to the pilot the status of all critical system parameters.
§ 27.1357   Circuit protective devices.
(a) Protective devices, such as fuses or circuit breakers, must be installed in each electrical circuit other than—
(1) The main circuits of starter motors; and
(2) Circuits in which no hazard is presented by their omission.
(b) A protective device for a circuit essential to flight safety may not be used to protect any other circuit.
(c) Each resettable circuit protective device (“trip free” device in which the tripping mechanism cannot be overridden by the operating control) must be designed so that—
(1) A manual operation is required to restore service after trippling; and
(2) If an overload or circuit fault exists, the device will open the circuit regardless of the position of the operating control.
(d) If the ability to reset a circuit breaker or replace a fuse is essential to safety in flight, that circuit breaker or fuse must be located and identified so that it can be readily reset or replaced in flight.
(e) If fuses are used, there must be one spare of each rating, or 50 percent spare fuses of each rating, whichever is greater.
§ 27.1361   Master switch.
(a) There must be a master switch arrangement to allow ready disconnection of each electric power source from the main bus. The point of disconnection must be adjacent to the sources controlled by the switch.
(b) Load circuits may be connected so that they remain energized after the switch is opened, if they are protected by circuit protective devices, rated at five amperes or less, adjacent to the electric power source.
(c) The master switch or its controls must be installed so that the switch is easily discernible and accessible to a crewmember in flight.
§ 27.1365   Electric cables.
(a) Each electric connecting cable must be of adequate capacity.
(b) Each cable that would overheat in the event of circuit overload or fault must be at least flame resistant and may not emit dangerous quantities of toxic fumes.
(c) Insulation on electrical wire and cable installed in the rotorcraft must be self-extinguishing when tested in accordance with appendix F, part I(a)(3), of part 25 of this chapter.
§ 27.1367   Switches.
Each switch must be—
(a) Able to carry its rated current;
(b) Accessible to the crew; and
(c) Labeled as to operation and the circuit controlled.
§ 27.1381   Instrument lights.
The instrument lights must—
(a) Make each instrument, switch, and other devices for which they are provided easily readable; and
(b) Be installed so that—
(1) Their direct rays are shielded from the pilot's eyes; and
(2) No objectionable reflections are visible to the pilot.
§ 27.1383   Landing lights.
(a) Each required landing or hovering light must be approved.
(b) Each landing light must be installed so that—
(1) No objectionable glare is visible to the pilot;
(2) The pilot is not adversely affected by halation; and
(3) It provides enough light for night operation, including hovering and landing.
(c) At least one separate switch must be provided, as applicable—
(1) For each separately installed landing light; and
(2) For each group of landing lights installed at a common location.
§ 27.1385   Position light system installation.
(a) General. Each part of each position light system must meet the applicable requirements of this section, and each system as a whole must meet the requirements of §§ 27.1387 through 27.1397.
(b) Forward position lights. Forward position lights must consist of a red and a green light spaced laterally as far apart as practicable and installed forward on the rotorcraft so that, with the rotorcraft in the normal flying position, the red light is on the left side and the green light is on the right side. Each light must be approved.
(c) Rear position light. The rear position light must be a white light mounted as far aft as practicable, and must be approved.
(d) Circuit. The two forward position lights and the rear position light must make a single circuit.
(e) Light covers and color filters. Each light cover or color filter must be at least flame resistant and may not change color or shape or lose any appreciable light transmission during normal use.
§ 27.1387   Position light system dihedral angles.
(a) Except as provided in paragraph (e) of this section, each forward and rear position light must, as installed, show unbroken light within the dihedral angles described in this section.
(b) Dihedral angle L (left) is formed by two intersecting vertical planes, the first parallel to the longitudinal axis of the rotorcraft, and the other at 110 degrees to the left of the first, as viewed when looking forward along the longitudinal axis.
(c) Dihedral angle R (right) is formed by two intersecting vertical planes, the first parallel to the longitudinal axis of the rotorcraft, and the other at 110 degrees to the right of the first, as viewed when looking forward along the longitudinal axis.
(d) Dihedral angle A (aft) is formed by two intersecting vertical planes making angles of 70 degrees to the right and to the left, respectively, to a vertical plane passing through the longitudinal axis, as viewed when looking aft along the longitudinal axis.
(e) If the rear position light, when mounted as far aft as practicable in accordance with § 25.1385(c), cannot show unbroken light within dihedral angle A (as defined in paragraph (d) of this section), a solid angle or angles of obstructed visibility totaling not more than 0.04 steradians is allowable within that dihedral angle, if such solid angle is within a cone whose apex is at the rear position light and whose elements make an angle of 30° with a vertical line passing through the rear position light.
§ 27.1389   Position light distribution and intensities.
(a) General. the intensities prescribed in this section must be provided by new equipment with light covers and color filters in place. Intensities must be determined with the light source operating at a steady value equal to the average luminous output of the source at the normal operating voltage of the rotorcraft. The light distribution and intensity of each position light must meet the requirements of paragraph (b) of this section.
(b) Forward and rear position lights. The light distribution and intensities of forward and rear position lights must be expressed in terms of minimum intensities in the horizontal plane, minimum intensities in any vertical plane, and maximum intensities in overlapping beams, within dihedral angles L, R, and A, and must meet the following requirements:
(1) Intensities in the horizontal plane. Each intensity in the horizontal plane (the plane containing the longitudinal axis of the rotorcraft and perpendicular to the plane of symmetry of the rotorcraft) must equal or exceed the values in § 27.1391.
(2) Intensities in any vertical plane. Each intensity in any vertical plane (the plane perpendicular to the horizontal plane) must equal or exceed the appropriate value in § 27.1393, where I is the minimum intensity prescribed in § 27.1391 for the corresponding angles in the horizontal plane.
(3) Intensities in overlaps between adjacent signals. No intensity in any overlap between adjacent signals may exceed the values in § 27.1395, except that higher intensities in overlaps may be used with main beam intensities substantially greater than the minima specified in §§ 27.1391 and 27.1393, if the overlap intensities in relation to the main beam intensities do not adversely affect signal clarity. When the peak intensity of the forward position lights is greater than 100 candles, the maximum overlap intensities between them may exceed the values in § 27.1395 if the overlap intensity in Area A is not more than 10 percent of peak position light intensity and the overlap intensity in Area B is not more than 2.5 percent of peak position light intensity.
§ 27.1391   Minimum intensities in the horizontal plane of forward and rear position lights.
Each position light intensity must equal or exceed the applicable values in the following table:
§ 27.1393   Minimum intensities in any vertical plane of forward and rear position lights.
Each position light intensity must equal or exceed the applicable values in the following table:
§ 27.1395   Maximum intensities in overlapping beams of forward and rear position lights.
No position light intensity may exceed the applicable values in the following table, except as provided in § 27.1389(b)(3).
(a) Area A includes all directions in the adjacent dihedral angle that pass through the light source and intersect the common boundary plane at more than 10 degrees but less than 20 degrees, and
(b) Area B includes all directions in the adjacent dihedral angle that pass through the light source and intersect the common boundary plane at more than 20 degrees.
§ 27.1397   Color specifications.
Each position light color must have the applicable International Commission on Illumination chromaticity coordinates as follows:
(a) Aviation red —
y is not greater than 0.335; and
z is not greater than 0.002.
(b) Aviation green —
x is not greater than 0.440−0.320 y ;
x is not greater than y −0.170; and
y is not less than 0.390−0.170 x.
(c) Aviation white —
x is not less than 0.300 and not greater than 0.540;
y is not less than x −0.040” or y c −0.010, whichever is the smaller; and
y is not greater than x + 0.020 nor 0.636−0.400 x ;
Where y c is the y coordinate of the Planckian radiator for the value of x considered.
§ 27.1399   Riding light.
(a) Each riding light required for water operation must be installed so that it can—
(1) Show a white light for at least two nautical miles at night under clear atmospheric conditions; and
(2) Show a maximum practicable unbroken light with the rotorcraft on the water.
(b) Externally hung lights may be used.
§ 27.1401   Anticollision light system.
(a) General. If certification for night operation is requested, the rotorcraft must have an anticollision light system that—
(1) Consists of one or more approved anticollision lights located so that their emitted light will not impair the crew's vision or detract from the conspicuity of the position lights; and
(2) Meets the requirements of paragraphs (b) through (f) of this section.
(b) Field of coverage. The system must consist of enough lights to illuminate the vital areas around the rotorcraft, considering the physical configuration and flight characteristics of the rotorcraft. The field of coverage must extend in each direction within at least 30 degrees below the horizontal plane of the rotorcraft, except that there may be solid angles of obstructed visibility totaling not more than 0.5 steradians.
(c) Flashing characteristics. The arrangement of the system, that is, the number of light sources, beam width, speed of rotation, and other characteristics, must give an effective flash frequency of not less than 40, nor more than 100, cycles per minute. The effective flash frequency is the frequency at which the rotorcraft's complete anticollision light system is observed from a distance, and applies to each sector of light including any overlaps that exist when the system consists of more than one light source. In overlaps, flash frequencies may exceed 100, but not 180, cycles per minute.
(d) Color. Each anticollision light must be aviation red and must meet the applicable requirements of § 27.1397.
(e) Light intensity. The minimum light intensities in any vertical plane, measured with the red filter (if used) and expressed in terms of “effective” intensities, must meet the requirements of paragraph (f) of this section. The following relation must be assumed:
(f) Minimum effective intensities for anticollision light. Each anticollision light effective intensity must equal or exceed the applicable values in the following table:
§ 27.1411   General.
(a) Required safety equipment to be used by the crew in an emergency, such as flares and automatic liferaft releases, must be readily accessible.
(b) Stowage provisions for required safety equipment must be furnished and must—
(1) Be arranged so that the equipment is directly accessible and its location is obvious; and
(2) Protect the safety equipment from damage caused by being subjected to the inertia loads specified in § 27.561.
§ 27.1413   Safety belts.
Each safety belt must be equipped with a metal to metal latching device.
§ 27.1415   Ditching equipment.
(a) Emergency flotation and signaling equipment required by any operating rule in this chapter must meet the requirements of this section.
(b) Each raft and each life preserver must be approved and must be installed so that it is readily available to the crew and passengers. The storage provisions for life preservers must accommodate one life preserver for each occupant for which certification for ditching is requested.
(c) Each raft released automatically or by the pilot must be attached to the rotorcraft by a line to keep it alongside the rotorcraft. This line must be weak enough to break before submerging the empty raft to which it is attached.
(d) Each signaling device must be free from hazard in its operation and must be installed in an accessible location.
§ 27.1419   Ice protection.
(a) To obtain certification for flight into icing conditions, compliance with this section must be shown.
(b) It must be demonstrated that the rotorcraft can be safely operated in the continuous maximum and intermittent maximum icing conditions determined under appendix C of Part 29 of this chapter within the rotorcraft altitude envelope. An analysis must be performed to establish, on the basis of the rotorcraft's operational needs, the adequacy of the ice protection system for the various components of the rotorcraft.
(c) In addition to the analysis and physical evaluation prescribed in paragraph (b) of this section, the effectiveness of the ice protection system and its components must be shown by flight tests of the rotorcraft or its components in measured natural atmospheric icing conditions and by one or more of the following tests as found necessary to determine the adequacy of the ice protection system:
(1) Laboratory dry air or simulated icing tests, or a combination of both, of the components or models of the components.
(2) Flight dry air tests of the ice protection system as a whole, or its individual components.
(3) Flight tests of the rotorcraft or its components in measured simulated icing conditions.
(d) The ice protection provisions of this section are considered to be applicable primarily to the airframe. Powerplant installation requirements are contained in Subpart E of this part.
(e) A means must be indentified or provided for determining the formation of ice on critical parts of the rotorcraft. Unless otherwise restricted, the means must be available for nighttime as well as daytime operation. The rotorcraft flight manual must describe the means of determining ice formation and must contain information necessary for safe operation of the rotorcraft in icing conditions.
§ 27.1435   Hydraulic systems.
(a) Design. Each hydraulic system and its elements must withstand, without yielding, any structural loads expected in addition to hydraulic loads.
(b) Tests. Each system must be substantiated by proof pressure tests. When proof tested, no part of any system may fail, malfunction, or experience a permanent set. The proof load of each system must be at least 1.5 times the maximum operating pressure of that system.
(c) Accumulators. No hydraulic accumulator or pressurized reservoir may be installed on the engine side of any firewall unless it is an integral part of an engine.
§ 27.1457   Cockpit voice recorders.
(a) Each cockpit voice recorder required by the operating rules of this chapter must be approved, and must be installed so that it will record the following:
(1) Voice communications transmitted from or received in the rotorcraft by radio.
(2) Voice communications of flight crewmembers on the flight deck.
(3) Voice communications of flight crewmembers on the flight deck, using the rotorcraft's interphone system.
(4) Voice or audio signals identifying navigation or approach aids introduced into a headset or speaker.
(5) Voice communications of flight crewmembers using the passenger loudspeaker system, if there is such a system, and if the fourth channel is available in accordance with the requirements of paragraph (c)(4)(ii) of this section.
(6) If datalink communication equipment is installed, all datalink communications, using an approved data message set. Datalink messages must be recorded as the output signal from the communications unit that translates the signal into usable data.
(b) The recording requirements of paragraph (a)(2) of this section may be met:
(1) By installing a cockpit-mounted area microphone located in the best position for recording voice communications originating at the first and second pilot stations and voice communications of other crewmembers on the flight deck when directed to those stations; or
(2) By installing a continually energized or voice-actuated lip microphone at the first and second pilot stations.
The microphone specified in this paragraph must be so located and, if necessary, the preamplifiers and filters of the recorder must be adjusted or supplemented so that the recorded communications are intelligible when recorded under flight cockpit noise conditions and played back. The level of intelligibility must be approved by the Administrator. Repeated aural or visual playback of the record may be used in evaluating intelligibility.
(c) Each cockpit voice recorder must be installed so that the part of the communication or audio signals specified in paragraph (a) of this section obtained from each of the following sources is recorded on a separate channel:
(1) For the first channel, from each microphone, headset, or speaker used at the first pilot station.
(2) For the second channel, from each microphone, headset, or speaker used at the second pilot station.
(3) For the third channel, from the cockpit-mounted area microphone, or the continually energized or voice-actuated lip microphone at the first and second pilot stations.
(4) For the fourth channel, from:
(i) Each microphone, headset, or speaker used at the stations for the third and fourth crewmembers; or
(ii) If the stations specified in paragraph (c)(4)(i) of this section are not required or if the signal at such a station is picked up by another channel, each microphone on the flight deck that is used with the passenger loudspeaker system if its signals are not picked up by another channel.
(iii) Each microphone on the flight deck that is used with the rotorcraft's loudspeaker system if its signals are not picked up by another channel.
(d) Each cockpit voice recorder must be installed so that:
(1)(i) It receives its electrical power from the bus that provides the maximum reliability for operation of the cockpit voice recorder without jeopardizing service to essential or emergency loads.
(ii) It remains powered for as long as possible without jeopardizing emergency operation of the rotorcraft.
(2) There is an automatic means to simultaneously stop the recorder and prevent each erasure feature from functioning, within 10 minutes after crash impact;
(3) There is an aural or visual means for preflight checking of the recorder for proper operation;
(4) Whether the cockpit voice recorder and digital flight data recorder are installed in separate boxes or in a combination unit, no single electrical failure external to the recorder may disable both the cockpit voice recorder and the digital flight data recorder; and
(5) It has an independent power source—
(i) That provides 10 ±1 minutes of electrical power to operate both the cockpit voice recorder and cockpit-mounted area microphone;
(ii) That is located as close as practicable to the cockpit voice recorder; and
(iii) To which the cockpit voice recorder and cockpit-mounted area microphone are switched automatically in the event that all other power to the cockpit voice recorder is interrupted either by normal shutdown or by any other loss of power to the electrical power bus.
(e) The record container must be located and mounted to minimize the probability of rupture of the container as a result of crash impact and consequent heat damage to the record from fire.
(f) If the cockpit voice recorder has a bulk erasure device, the installation must be designed to minimize the probability of inadvertent operation and actuation of the device during crash impact.
(g) Each recorder container must be either bright orange or bright yellow.
(h) When both a cockpit voice recorder and a flight data recorder are required by the operating rules, one combination unit may be installed, provided that all other requirements of this section and the requirements for flight data recorders under this part are met.
§ 27.1459   Flight data recorders.
(a) Each flight recorder required by the operating rules of Subchapter G of this chapter must be installed so that:
(1) It is supplied with airspeed, altitude, and directional data obtained from sources that meet the accuracy requirements of §§ 27.1323, 27.1325, and 27.1327 of this part, as applicable;
(2) The vertical acceleration sensor is rigidly attached, and located longitudinally within the approved center of gravity limits of the rotorcraft;
(3)(i) It receives its electrical power from the bus that provides the maximum reliability for operation of the flight data recorder without jeopardizing service to essential or emergency loads.
(ii) It remains powered for as long as possible without jeopardizing emergency operation of the rotorcraft.
(4) There is an aural or visual means for preflight checking of the recorder for proper recording of data in the storage medium;
(5) Except for recorders powered solely by the engine-driven electrical generator system, there is an automatic means to simultaneously stop a recorder that has a data erasure feature and prevent each erasure feature from functioning, within 10 minutes after any crash impact; and
(6) Whether the cockpit voice recorder and digital flight data recorder are installed in separate boxes or in a combination unit, no single electrical failure external to the recorder may disable both the cockpit voice recorder and the digital flight data recorder.
(b) Each nonejectable recorder container must be located and mounted so as to minimize the probability of container rupture resulting from crash impact and subsequent damage to the record from fire.
(c) A correlation must be established between the flight recorder readings of airspeed, altitude, and heading and the corresponding readings (taking into account correction factors) of the first pilot's instruments. This correlation must cover the airspeed range over which the aircraft is to be operated, the range of altitude to which the aircraft is limited, and 360 degrees of heading. Correlation may be established on the ground as appropriate.
(d) Each recorder container must:
(1) Be either bright orange or bright yellow;
(2) Have a reflective tape affixed to its external surface to facilitate its location under water; and
(3) Have an underwater locating device, when required by the operating rules of this chapter, on or adjacent to the container which is secured in such a manner that they are not likely to be separated during crash impact.
(e) When both a cockpit voice recorder and a flight data recorder are required by the operating rules, one combination unit may be installed, provided that all other requirements of this section and the requirements for cockpit voice recorders under this part are met.
§ 27.1461   Equipment containing high energy rotors.
(a) Equipment containing high energy rotors must meet paragraph (b), (c), or (d) of this section.
(b) High energy rotors contained in equipment must be able to withstand damage caused by malfunctions, vibration, abnormal speeds, and abnormal temperatures. In addition—
(1) Auxiliary rotor cases must be able to contain damage caused by the failure of high energy rotor blades; and
(2) Equipment control devices, systems, and instrumentation must reasonably ensure that no operating limitations affecting the integrity of high energy rotors will be exceeded in service.
(c) It must be shown by test that equipment containing high energy rotors can contain any failure of a high energy rotor that occurs at the highest speed obtainable with the normal speed control devices inoperative.
(d) Equipment containing high energy rotors must be located where rotor failure will neither endanger the occupants nor adversely affect continued safe flight.
§ 27.1501   General.
(a) Each operating limitation specified in §§ 27.1503 through 27.1525 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 §§ 27.1541 through 27.1589.
§ 27.1503   Airspeed limitations: general.
(a) An operating speed range must be established.
(b) When airspeed limitations are a function of weight, weight distribution, altitude, rotor speed, power, or other factors, airspeed limitations corresponding with the critical combinations of these factors must be established.
§ 27.1505   Never-exceed speed.
(a) The never-exceed speed, V NE, must be established so that it is—
(1) Not less than 40 knots (CAS); and
(2) Not more than the lesser of—
(i) 0.9 times the maximum forward speeds established under § 27.309;
(ii) 0.9 times the maximum speed shown under §§ 27.251 and 27.629; or
(iii) 0.9 times the maximum speed substantiated for advancing blade tip mach number effects.
(b) V NE may vary with altitude, r.p.m., temperature, and weight, if—
(1) No more than two of these variables (or no more than two instruments integrating more than one of these variables) are used at one time; and
(2) The ranges of these variables (or of the indications on instruments integrating more than one of these variables) are large enough to allow an operationally practical and safe variation of V NE .
(c) For helicopters, a stabilized power-off V NE denoted as V NE (power-off) may be established at a speed less than V NE established pursuant to paragraph (a) of this section, if the following conditions are met:
(1) V NE (power-off) is not less than a speed midway between the power-on V NE and the speed used in meeting the requirements of—
(i) § 27.65(b) for single engine helicopters; and
(ii) § 27.67 for multiengine helicopters.
(2) V NE (power-off) is—
(i) A constant airspeed;
(ii) A constant amount less than power-on V NE ; or
(iii) A constant airspeed for a portion of the altitude range for which certification is requested, and a constant amount less than power-on V NE for the remainder of the altitude range.
§ 27.1509   Rotor speed.
(a) Maximum power-off (autorotation). The maximum power-off rotor speed must be established so that it does not exceed 95 percent of the lesser of—
(1) The maximum design r.p.m. determined under § 27.309(b); and
(2) The maximum r.p.m. shown during the type tests.
(b) Minimum power off. The minimum power-off rotor speed must be established so that it is not less than 105 percent of the greater of—
(1) The minimum shown during the type tests; and
(2) The minimum determined by design substantiation.
(c) Minimum power on. The minimum power-on rotor speed must be established so that it is—
(1) Not less than the greater of—
(i) The minimum shown during the type tests; and
(ii) The minimum determined by design substantiation; and
(2) Not more than a value determined under § 27.33(a)(1) and (b)(1).
§ 27.1519   Weight and center of gravity.
The weight and center of gravity limitations determined under §§ 27.25 and 27.27, respectively, must be established as operating limitations.
§ 27.1521   Powerplant 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 are type certificated.
(b) Takeoff operation. The powerplant takeoff operation must be limited by—
(1) The maximum rotational speed, which may not be greater than—
(i) The maximum value determined by the rotor design; or
(ii) The maximum value shown during the type tests;
(2) The maximum allowable manifold pressure (for reciprocating engines);
(3) The time limit for the use of the power corresponding to the limitations established in paragraphs (b)(1) and (2) of this section;
(4) If the time limit in paragraph (b)(3) of this section exceeds two minutes, the maximum allowable cylinder head, coolant outlet, or oil temperatures;
(5) The gas temperature limits for turbine engines over the range of operating and atmospheric conditions for which certification is requested.
(c) Continuous operation. The continuous operation must be limited by—
(1) The maximum rotational speed which may not be greater than—
(i) The maximum value determined by the rotor design; or
(ii) The maximum value shown during the type tests;
(2) The minimum rotational speed shown under the rotor speed requirements in § 27.1509(c); and
(3) The gas temperature limits for turbine engines over the range of operating and atmospheric conditions for which certification is requested.
(d) Fuel grade or designation. The minimum fuel grade (for reciprocating engines), or fuel designation (for turbine engines), must be established so that it is not less than that required for the operation of the engines within the limitations in paragraphs (b) and (c) of this section.
(e) Turboshaft engine torque. For rotorcraft with main rotors driven by turboshaft engines, and that do not have a torque limiting device in the transmission system, the following apply:
(1) A limit engine torque must be established if the maximum torque that the engine can exert is greater than—
(i) The torque that the rotor drive system is designed to transmit; or
(ii) The torque that the main rotor assembly is designed to withstand in showing compliance with § 27.547(e).
(2) The limit engine torque established under paragraph (e)(1) of this section may not exceed either torque specified in paragraph (e)(1)(i) or (ii) of this section.
(f) Ambient temperature. For turbine engines, ambient temperature limitations (including limitations for winterization installations, if applicable) must be established as the maximum ambient atmospheric temperature at which compliance with the cooling provisions of §§ 27.1041 through 27.1045 is shown.
(g) Two and one-half-minute OEI power operation. Unless otherwise authorized, the use of 2 1/2 -minute OEI power must be limited to engine failure operation of multiengine, turbine-powered rotorcraft for not longer than 2 1/2 minutes after failure of an engine. The use of 2 1/2 -minute OEI power must also be limited by—
(1) The maximum rotational speed, which may not be greater than—
(i) The maximum value determined by the rotor design; or
(ii) The maximum demonstrated during the type tests;
(2) The maximum allowable gas temperature; and
(3) The maximum allowable torque.
(h) Thirty-minute OEI power operation. Unless otherwise authorized, the use of 30-minute OEI power must be limited to multiengine, turbine-powered rotorcraft for not longer than 30 minutes after failure of an engine. The use of 30-minute OEI power must also be limited by—
(1) The maximum rotational speed, which may not be greater than—
(i) The maximum value determined by the rotor design; or
(ii) The maximum value demonstrated during the type tests;
(2) The maximum allowable gas temperature; and
(3) The maximum allowable torque.
(i) Continuous OEI power operation. Unless otherwise authorized, the use of continuous OEI power must be limited to multiengine, turbine-powered rotorcraft for continued flight after failure of an engine. The use of continuous OEI power must also be limited by—
(1) The maximum rotational speed, which may not be greater than—
(i) The maximum value determined by the rotor design; or
(ii) The maximum value demonstrated during the type tests;
(2) The maximum allowable gas temperature; and
(3) The maximum allowable torque.
(j) Rated 30-second OEI power operation. Rated 30-second OEI power is permitted only on multiengine, turbine-powered rotorcraft, also certificated for the use of rated 2-minute OEI power, and can only be used for continued operation of the remaining engine(s) after a failure or precautionary shutdown of an engine. It must be shown that following application of 30-second OEI power, any damage will be readily detectable by the applicable inspections and other related procedures furnished in accordance with Section A27.4 of appendix A of this part and Section A33.4 of appendix A of part 33. The use of 30-second OEI power must be limited to not more than 30 seconds for any period in which that power is used, and by—
(1) The maximum rotational speed, which may not be greater than—
(i) The maximum value determined by the rotor design; or
(ii) The maximum value demonstrated during the type tests;
(2) The maximum allowable gas temperature; and
(3) The maximum allowable torque.
(k) Rated 2-minute OEI power operation. Rated 2-minute OEI power is permitted only on multiengine, turbine-powered rotorcraft, also certificated for the use of rated 30-second OEI power, and can only be used for continued operation of the remaining engine(s) after a failure or precautionary shutdown of an engine. It must be shown that following application of 2-minute OEI power, any damage will be readily detectable by the applicable inspections and other related procedures furnished in accordance with Section A27.4 of appendix A of this part and Section A33.4 of appendix A of part 33. The use of 2-minute OEI power must be limited to not more than 2 minutes for any period in which that power is used, and by—
(1) The maximum rotational speed, which may not be greater than—
(i) The maximum value determined by the rotor design; or
(ii) The maximum value demonstrated during the type tests;
(2) The maximum allowable gas temperature; and
(3) The maximum allowable torque.
§ 27.1523   Minimum flight crew.
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 kinds of operation authorized under § 27.1525.
§ 27.1525   Kinds of operations.
The kinds of operations (such as VFR, IFR, day, night, or icing) for which the rotorcraft is approved are established by demonstrated compliance with the applicable certification requirements and by the installed equipment.
§ 27.1527   Maximum operating altitude.
The maximum altitude up to which operation is allowed, as limited by flight, structural, powerplant, functional, or equipment characteristics, must be established.
§ 27.1529   Instructions for Continued Airworthiness.
The applicant must prepare Instructions for Continued Airworthiness in accordance with appendix A 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 rotorcraft or issuance of a standard certificate of airworthiness, whichever occurs later.
§ 27.1541   General.
(a) The rotorcraft must contain—
(1) The markings and placards specified in §§ 27.1545 through 27.1565, and
(2) Any additional information, instrument markings, and placards required for the safe operation of rotorcraft with 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.
§ 27.1543   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 arc and line must be wide enough, and located, to be clearly visible to the pilot.
§ 27.1545   Airspeed indicator.
(a) Each airspeed indicator must be marked as specified in paragraph (b) of this section, with the marks located at the corresponding indicated airspeeds.
(b) The following markings must be made:
(1) A red line—
(i) For rotorcraft other than helicopters, at V NE .
(ii) For helicopters, at V NE (power-on).
(iii) For helicopters, at V NE (power-off). If V NE (power-off) is less than V NE (power-on) and both are simultaneously displayed, the red line at V NE (power-off) must be clearly distinguishable from the red line at V NE (power-on).
(2) [Reserved]
(3) For the caution range, a yellow range.
(4) For the normal operating range, a green or unmarked range.
§ 27.1547   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.
(e) If a magnetic nonstabilized direction indicator can have a deviation of more than 10 degrees caused by the operation of electrical equipment, the placard must state which electrical loads, or combination of loads, would cause a deviation of more than 10 degrees when turned on.
§ 27.1549   Powerplant instruments.
For each required powerplant instrument, as appropriate to the type of instrument—
(a) Each maximum and, if applicable, minimum safe operating limit must be marked with a red line;
(b) Each normal operating range must be marked as a green or unmarked range;
(c) Each takeoff and precautionary range must be marked with a yellow range or yellow line;
(d) Each engine or rotor range that is restricted because of excessive vibration stresses must be marked with red ranges or red lines; and
(e) Each OEI limit or approved operating range must be marked to be clearly differentiated from the markings of paragraphs (a) through (d) of this section except that no marking is normally required for the 30-second OEI limit.
§ 27.1551   Oil quantity indicator.
Each oil quantity indicator must be marked with enough increments to indicate readily and accurately the quantity of oil.
§ 27.1553   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.
§ 27.1555   Control markings.
(a) Each cockpit control, other than primary flight controls or control whose function is obvious, must be plainly marked as to its function and method of operation.
(b) 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 any engine of a multiengine rotorcraft must be marked to indicate the position corresponding to each engine controlled.
(c) Usable fuel capacity must be marked as follows:
(1) For fuel systems having no selector controls, the usable fuel capacity of the system must be indicated at the fuel quantity indicator unless it is:
(i) Provided by another system or equipment readily accessible to the pilot; and
(ii) Contained in the limitations section of the rotorcraft flight manual.
(2) For fuel systems having selector controls, the usable fuel capacity available at each selector control position must be indicated near the selector control.
(d) For accessory, auxiliary, and emergency controls—
(1) Each essential visual position indicator, such as those showing rotor pitch or landing gear position, must be marked so that each crewmember can determine at any time the position of the unit to which it relates; and
(2) Each emergency control must be red and must be marked as to method of operation.
(e) For rotorcraft incorporating retractable landing gear, the maximum landing gear operating speed must be displayed in clear view of the pilot.
§ 27.1557   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.
(b) Seats. If the maximum allowable weight to be carried in a seat is less than 170 pounds, a placard stating the lesser weight must be permanently attached to the seat structure.
(c) Fuel and oil 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 rotorcraft, the minimum fuel grade;
(iii) For turbine engine powered rotorcraft, 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”.
(d) Emergency exit placards. Each placard and operating control for each emergency exit must be red. A placard must be near each emergency exit control and must clearly indicate the location of that exit and its method of operation.
§ 27.1559   Limitations placard.
There must be a placard in clear view of the pilot that specifies the kinds of operations (such as VFR, IFR, day, night, or icing) for which the rotorcraft is approved.
§ 27.1561   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 so marked.
§ 27.1565   Tail rotor.
Each tail rotor must be marked so that its disc is conspicuous under normal daylight ground conditions.
§ 27.1581   General.
(a) Furnishing information. A Rotorcraft Flight Manual must be furnished with each rotorcraft, and it must contain the following:
(1) Information required by §§ 27.1583 through 27.1589.
(2) Other information that is necessary for safe operation because of design, operating, or handling characteristics.
(b) Approved information. Each part of the manual listed in §§ 27.1583 through 27.1589, that is appropriate to the rotorcraft, must be furnished, verified, and approved, and must be segregated, identified, and clearly distinguished from each unapproved part of that manual.
(c) [Reserved]
(d) Table of contents. Each Rotorcraft Flight Manual must include a table of contents if the complexity of the manual indicates a need for it.
§ 27.1583   Operating limitations.
(a) Airspeed and rotor limitations. Information necessary for the marking of airspeed and rotor limitations on, or near, their respective indicators must be furnished. The significance of each limitation and of the color coding must be explained.
(b) Powerplant limitations. The following information must be furnished:
(1) Limitations required by § 27.1521.
(2) Explanation of the limitations, when appropriate.
(3) Information necessary for marking the instruments required by §§ 27.1549 through 27.1553.
(c) Weight and loading distribution. The weight and center of gravity limits required by §§ 27.25 and 27.27, respectively, must be furnished. If the variety of possible loading conditions warrants, instructions must be included to allow ready observance of the limitations.
(d) Flight crew. When a flight crew of more than one is required, the number and functions of the minimum flight crew determined under § 27.1523 must be furnished.
(e) Kinds of operation. Each kind of operation for which the rotorcraft and its equipment installations are approved must be listed.
(f) [Reserved]
(g) Altitude. The altitude established under § 27.1527 and an explanation of the limiting factors must be furnished.
§ 27.1585   Operating procedures.
(a) Parts of the manual containing operating procedures must have information concerning any normal and emergency procedures and other information necessary for safe operation, including takeoff and landing procedures and associated airspeeds. The manual must contain any pertinent information including—
(1) The kind of takeoff surface used in the tests and each appropriate climbout speed; and
(2) The kind of landing surface used in the tests and appropriate approach and glide airspeeds.
(b) For multiengine rotorcraft, information identifying each operating condition in which the fuel system independence prescribed in § 27.953 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.
(c) For helicopters for which a V NE (power-off) is established under § 27.1505(c), information must be furnished to explain the V NE (power-off) and the procedures for reducing airspeed to not more than the V NE (power-off) following failure of all engines.
(d) For each rotorcraft showing compliance with § 27.1353 (g)(2) or (g)(3), the operating procedures for disconnecting the battery from its charging source must be furnished.
(e) If the unusable fuel supply in any tank exceeds five percent of the tank capacity, or one gallon, whichever is greater, information must be furnished which 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.
(g) The airspeeds and rotor speeds for minimum rate of descent and best glide angle as prescribed in § 27.71 must be provided.
§ 27.1587   Performance information.
(a) The Rotorcraft Flight Manual must contain the following information, determined in accordance with §§ 27.49 through 27.87 and 27.143(c) and (d):
(1) Enough information to determine the limiting height-velocity envelope.
(2) Information relative to—
(i) The steady rates of climb and descent, in-ground effect and out-of-ground effect hovering ceilings, together with the corresponding airspeeds and other pertinent information including the calculated effects of altitude and temperatures;
(ii) The maximum weight for each altitude and temperature condition at which the rotorcraft can safely hover in-ground effect and out-of-ground effect in winds of not less than 17 knots from all azimuths. These data must be clearly referenced to the appropriate hover charts. In addition, if there are other combinations of weight, altitude and temperature for which performance information is provided and at which the rotorcraft cannot land and take off safely with the maximum wind value, those portions of the operating envelope and the appropriate safe wind conditions must be stated in the Rotorcraft Flight Manual;
(iii) For reciprocating engine-powered rotorcraft, the maximum atmospheric temperature at which compliance with the cooling provisions of §§ 27.1041 through 27.1045 is shown; and
(iv) Glide distance as a function of altitude when autorotating at the speeds and conditions for minimum rate of descent and best glide as determined in § 27.71.
(b) The Rotorcraft Flight Manual must contain—
(1) In its performance information section any pertinent information concerning the takeoff weights and altitudes used in compliance with § 27.51; and
(2) The horizontal takeoff distance determined in accordance with § 27.65(a)(2)(i).
§ 27.1589   Loading information.
There must be loading instructions for each possible loading condition between the maximum and minimum weights determined under § 27.25 that can result in a center of gravity beyond any extreme prescribed in § 27.27, assuming any probable occupant weights.