Part V - Airworthiness Manual Chapter 527 - Normal Category Aircraft

Canadian Aviation Regulations (CARs) 2017-2

Preamble

SUBCHAPTERS

  • A (527.1-527.2), 
  • B (527.21-527.251), 
  • C (527.301-527.571), 
  • D (527.601-527.873), 
  • E (527.901-527.1195), 
  • F (527.1301-527.1461), 
  • G (527.1501-527.1589)

APPENDICES

ABCD

(2002/06/01; no previous version)

SUBCHAPTER F EQUIPMENT - GENERAL

527.1301 Function and Installation

Each item of installed equipment must:

  1. (a) Be of a kind and design appropriate to its intended function;
  2. (b) Be labelled as to its identification, function, or operating limitations, or any applicable combination of these factors;
  3. (c) Be installed according to limitations specified for that equipment; and
  4. (d) Function properly when installed.

527.1301-1 Rotorcraft Operations After Ground Cold Soak

Substantiation of satisfactory operation of the rotorcraft as a total system, by cold weather testing or by documented evidence of satisfactory operation at low temperature is required after the rotorcraft has experienced a prolonged exposure to ground ambient temperatures equal to or less than -35°C unless an alternative minimum ground ambient temperature has been proposed by the applicant and accepted by the Minister.

FAR: No equivalent text.

(amended 2004/12/01; previous version)

527.1303 Flight and Navigation Instruments

The following are the required flight and navigation instruments:

  1. (a) An airspeed indicator.
  2. (b) An altimeter.
  3. (c) A magnetic direction indicator.

527.1305 Powerplant Instruments

The following are the required powerplant instruments:

  1. (a) A carburetor air temperature indicator, for each engine having a preheater that can provide a heat rise in excess of 60°F.
  2. (b) A cylinder head temperature indicator, for each:
    1. (1) Air cooled engine;
    2. (2) Rotorcraft with cooling shutters; and
    3. (3) Rotorcraft for which compliance with 527.1043 is shown in any condition other than the most critical flight condition with respect to cooling.
  3. (c) A fuel pressure indicator, for each pump-fed engine.
  4. (d) A fuel quantity indicator, for each fuel tank.
  5. (e) A manifold pressure indicator, for each altitude engine.
  6. (f) An oil temperature warning device to indicate when the temperature exceeds a safe value in each main rotor drive gearbox (including any gearboxes essential to rotor phasing) having an oil system independent of the engine oil system.
  7. (g) An oil pressure warning device to indicate when the pressure falls below a safe value in each pressure-lubricated main rotor drive gearbox (including any gearboxes essential to rotor phasing) having an oil system independent of the engine oil system.
  8. (h) An oil pressure indicator for each engine.
  9. (i) An oil quantity indicator for each oil tank.
  10. (j) An oil temperature indicator for each engine.
  11. (k) At least one tachometer to indicate the r.p.m. of each engine and, as applicable:
    1. (1) The r.p.m. of the single main rotor;
    2. (2) The common r.p.m. of any main rotors whose speeds cannot vary appreciably with respect to each other; or
    3. (3) The r.p.m. of each main rotor whose speed can vary appreciably with respect to that of another main rotor.
  12. (l) A low fuel warning device for each fuel tank which feeds an engine. This device must:
    1. (1) Provide a warning to the flight crew when approximately 10 minutes of usable fuel remains in the tank; and
    2. (2) Be independent of the normal fuel quantity indicating system.
  13. (m) Means to indicate to the flight crew the failure of any fuel pump installed to show compliance with 527.955.
  14. (n) A gas temperature indicator for each turbine engine.
  15. (o) Means to enable the pilot to determine the torque of each turbo shaft engine, if a torque limitation is established for that engine under 527.1521 (e).
  16. (p) For each turbine engine, an indicator to indicate the functioning of the powerplant ice protection system.
  17. (q) An indicator for the fuel filter required by 527.997 to indicate the occurrence of contamination of the filter at the degree established by the applicant in compliance with 527.955.
  18. (r) For each turbine engine, a warning means for the oil strainer or filter required by 527.1019, if it has no bypass, to warn the pilot of the occurrence of contamination of the strainer or filter before it reaches the capacity established in accordance with 527.1019 (a)(2).
  19. (s) An indicator to indicate the functioning of any selectable or controllable heater used to prevent ice clogging of fuel system components.
  20. [(t) For rotorcraft for which a 30-second/2 minute OEI power rating is requested, a means must be provided to alert the pilot when the engine is at the 30-second and the 2-minute OEI power levels, when the event begins, and when the time interval expires.
  21. [(u) For each turbine engine utilizing 30-second/2 minute OEI power, a device or system must be provided for use by ground personnel which:
    1. [(1) Automatically records each usage and duration of power at the 30-second and 2-minute OEI levels;
    2. [(2) Permits retrieval of the recorded data;
    3. [(3) Can be reset only by ground maintenance personnel; and
    4. [(4) Has a means to verify proper operation of the system or device.
  22. [(v) Warning or caution devices to signal to the flight crew when ferromagnetic particles are detected by the chip detector required by 527.1337 (e).]

    (Change 527-1 (89-01-01))

    (Change 527-4)

527.1307Miscellaneous Equipment

The following is the required miscellaneous equipment:

  1. (a) An approved seat for each occupant.
  2. (b) An approved safety belt for each occupant.
  3. (c) A master switch arrangement.
  4. (d) An adequate source of electrical energy, where electrical energy is necessary for operation of the rotorcraft.
  5. (e) Electrical protective devices.

527.1309 Equipment, Systems, and Installations

  1. (a) The equipment, systems, and installations whose functioning is required by this Manual must be designed and installed to ensure that they perform their intended functions under any foreseeable operating condition.
  2. (b) The equipment, systems, and installations of a multi-engine rotorcraft must be designed to prevent hazards to the rotorcraft in the event of a probable malfunction or failure.
  3. (c) The equipment, systems, and installations of single-engine rotorcraft must be designed to minimize hazards to the rotorcraft in the event of a probable malfunction or failure.

527.1316 Electrical and Electronic System Lightning Protection

(effective 2014/07/08)

  1. (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. (1) the function is not adversely affected during and after the time the rotorcraft is exposed to lightning; and
    2. (2) the system automatically recovers normal operation of that function in a timely manner after the rotorcraft is exposed to lightning.
  2. (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 flight crew members 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.

527.1317 High-intensity Radiated Fields (HIRF) Protection

(amended 2008/10/30; no previous version)

  1. (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 shall be designed and installed so that:
    (amended 2008/10/30; no previous version)
    1. (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 of this chapter;
      (amended 2008/10/30; no previous version)
    2. (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 of this chapter, unless this conflicts with other operational or functional requirements of that system;
      (amended 2008/10/30; no previous version)
    3. (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 of this chapter; and
      (amended 2008/10/30; no previous version)
    4. (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 of this chapter.
      (amended 2008/10/30; no previous version)
  2. (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 flight crew to respond to an adverse operating condition shall 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 of this chapter.
    (amended 2008/10/30; no previous version)
  3. (c) Each electrical and electronic system that performs a function whose failure would reduce the capability of the rotorcraft or the ability of the flight crew to respond to an adverse operating condition, shall 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 of this chapter.
    (amended 2008/10/30; no previous version)
  4. (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:
    (amended 2008/10/30; no previous version)
    1. (1) The system has previously been shown to comply with Special Conditions - Airworthiness for HIRF, specified by the Minister pursuant to Part V of the Canadian Aviation Regulations (CARs).
      (amended 2008/10/30; no previous version)
    2. (2) The HIRF immunity characteristics of the system have not changed since compliance with the special conditions was demonstrated; and
      (amended 2008/10/30; no previous version)
    3. (3) The data used to demonstrate compliance with the Special Conditions - Airworthiness for HIRF is provided.
      (amended 2008/10/30; no previous version)

Instruments: Installation

527.1321 Arrangement and Visibility

  1. (a) Each flight, navigation, and powerplant instrument for use by any pilot must be easily visible to him.
  2. (b) For each multi-engine rotorcraft, identical powerplant instruments must be located so as to prevent confusion as to which engine each instrument relates.
  3. (c) Instrument panel vibration may not damage, or impair the readability or accuracy of, any instrument.
  4. (d) If a visual indicator is provided to indicate malfunction of an instrument, it must be effective under all probable cockpit lighting conditions.

527.1322 Warning, Caution, and Advisory Lights

If warning, caution or advisory lights are installed in the cockpit, they must, unless otherwise approved by the Minister, be:

  1. (a) Red, for warning lights (lights indicating a hazard which may require immediate corrective action);
  2. (b) Amber, for caution lights (lights indicating the possible need for future corrective action);
  3. (c) Green, for safe operation lights; and
  4. (d) Any other colour, including white, for lights not described in paragraphs (a) through (c) of this section, provided the colour differs sufficiently from the colours prescribed in paragraphs (a) through (c) of this section to avoid possible confusion.

527.1323 Airspeed Indicating System

  1. (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.
  2. (b) The airspeed indicating system must be calibrated in flight at forward speeds of 20 knots and over.
  3. (c) At each forward speed above 80 percent of the climb out 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. (1) +3 percent of the calibrated airspeed; or
    2. (2) 5 knots.

527.1325 Static Pressure Systems

  1. (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.
  2. (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 the 50 feet, a correction card must be provided for the alternate static system.
  3. (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. (1) When either source is selected, the other is blocked off; and
    2. (2) Both sources cannot be blocked off simultaneously.
  4. (d) For unpressurized rotorcraft, paragraph (c)(1) of the 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.

527.1327 Magnetic Direction Indicator

  1. (a) Except as provided in paragraph (b) of this section:
    1. (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. (2) The compensated installation may not have a deviation, in level flight, greater than 10 degrees on any heading.
  2. (b) A magnetic non-stabilized 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 non-stabilized direction indicator of more than 10 degrees must be placarded in accordance with 527.1547 (e).

527.1329 Automatic Pilot System

  1. (a) Each automatic pilot system must be designed so that the automatic pilot can:
    1. (1) Be sufficiently overpowered by one pilot to allow control of the rotorcraft; and
    2. (2) Be readily and positively disengaged by each pilot to prevent it from interfering with control of the rotorcraft.
  2. (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.
  3. (c) Each manually operated control for the system's operation must be readily accessible to the pilots.
  4. (d) The system must be designed and adjusted 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, either during normal operation or in the event of a malfunction, assuming that corrective action begins within a reasonable period of time.
  5. (e) If the automatic pilot integrates signals from auxiliary controls or furnishes signals for operation of other equipment, there must be positive interlocks and sequencing of engagement to prevent improper operation.
  6. [(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.]

    (Change 527-4)

527.1335 Flight Director Systems

If a flight director system is installed, means must be provided to indicate to the flight crew its current mode of operation. Selector switch position is not acceptable as a means of indication.

527.1337 Powerplant Instruments

  1. (a) Instruments and instrument lines.
    1. (1) Each powerplant instrument line must meet the requirements of 527.961 and 527.993.
    2. (2) Each line carrying flammable fluids under pressure must:
      1. (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
      2. (ii) Be installed and located so that the escape of fluids would not create a hazard.
    3. (3) Each powerplant instrument that utilizes flammable fluids must be installed and located so that the escape of fluid would not create a hazard.
  2. (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. (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 527.959;
    2. (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. (3) Each exposed sight gauge used as a fuel quantity indicator must be protected against damage.
  3. (c) Fuel flow metre system. If a fuel flow metre system is installed, each metering component must have a means for bypassing the fuel supply if malfunction of that component severely restricts fuel flow.
  4. (d) Oil quantity indicator. There must be means to indicate the quantity of oil in each tank:
    1. (1) On the ground (including during the filling of each tank); and
    2. (2) In flight, if there is an oil transfer system or reserve oil supply system.
  5. (e) [Rotor drive system transmissions and gearboxes utilizing ferromagnetic materials must be equipped with chip detectors designed to indicate or reveal the presence of ferromagnetic particles resulting from damage or excessive wear. Chip detectors must:
    1. (1) [Be designed to provide a signal to the device required by 527.1305 (v) and be provided with a means to allow crew members to check, in flight, the function of each detector electrical circuit and signal.
    2. (2) [Reserved]

    (Change 527-1 (89-01-01))

    (Change 527-4)

Electrical Systems and Equipment

527.1351 General

  1. (a) Electrical system capacity. Electrical equipment must be adequate for its intended use. In addition:
    1. (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. (2) Compliance with subparagraph (1) of this paragraph 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.
  2. (b) Function. For each electrical system, the following apply:
    1. (1) Each system, when installed, must be:
      1. (i) Free from hazards in itself, in its method of operation, and in its effects on other parts of the rotorcraft; and
      2. (ii) Protected from fuel, oil, water, other detrimental substances, and mechanical damage.
    2. (2) Electric power sources must function properly when connected in combination or independently.
    3. (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. (4) Each electric power source control must allow the independent operation of each source.
  3. (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. (1) Each generator must be able to deliver its continuous rated power;
    2. (2) Generator voltage control equipment must be able to dependably regulate each generator output within rated limits;
    3. (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. (4) Each generator must have an over-voltage 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 over-voltage condition.
  4. (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. (1) For direct current systems, an ammeter that can be switched into each generator feeder may be used; and
    2. (2) If there is only one generator, the ammeter may be in the battery feeder.
  5. (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.

    (Change 527-1 (89-01-01))

527.1353 Storage Battery Design and Installation

  1. (a) Each storage battery must be designed and installed as prescribed in this section.
  2. (b) Safe cell temperatures and pressures must be maintained during any probable charging and discharging condition. No uncontrolled increase in cell temperature may result when the battery is recharged (after previous complete discharge):
    1. (1) At maximum regulated voltage or power;
    2. (2) During a flight of maximum duration; and
    3. (3) Under the most adverse cooling condition likely to occur in service.
  3. (c) Compliance with paragraph (b) of this section must be shown by test unless experience with similar batteries and installations has shown that maintaining safe cell temperatures and pressures presents no problem.
  4. (d) No explosive or toxic gases emitted by any battery in normal operation, or as the result of any probable malfunction in the charging system or battery installation, may accumulate in hazardous quantities within the rotorcraft.
  5. (e) No corrosive fluids or gases that may escape from the battery may damage surrounding structures or adjacent essential equipment.
  6. (f) Each nickel cadmium battery installation capable of being used to start an engine or auxiliary power unit must have provisions to prevent any hazardous effect on structure or essential systems that may be caused by the maximum amount of heat the battery can generate during a short circuit of the battery or of its individual cells.
  7. (g) Nickel cadmium battery installations capable of being used to start an engine or auxiliary power unit must have:
    1. (1) A system to control the charging rate of the battery automatically so as to prevent battery overheating;
    2. (2) A battery temperature sensing and over-temperature warning system with a means for disconnecting the battery from its charging source in the event of an over-temperature condition; or
    3. (3) A battery failure sensing and warning system with a means for disconnecting the battery from its charging source in the event of battery failure.

527.1357 Circuit Protective Devices

  1. (a) Protective devices, such as fuses or circuit breakers, must be installed in each electrical circuit other than:
    1. (1) The main circuits of starter motors; and
    2. (2) Circuits in which no hazard is presented by their omission.
  2. (b) A protective device for a circuit essential to flight safety may not be used to protect any other circuit.
  3. (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. (1) A manual operation is required to restore service after trippling; and
    2. (2) If an overload or circuit fault exists, the device will open the circuit regardless of the position of the operating control.
  4. (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.
  5. (e) If fuses are used, there must be one spare of each rating, or 50 percent spare fuses of each rating, whichever is greater.

527.1361 Master Switch

  1. (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.
  2. (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 5 amperes or less, adjacent to the electric power source.
  3. (c) The master switch or its controls must be installed so that the switch is easily discernible and accessible to a crewmember in flight.

527.1365 Electric Cables

  1. (a) Each electric connecting cable must be of adequate capacity.
  2. (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.
  3. [(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 Chapter 525 of this Manual.]

    (Change 527-4)

527.1367 Switches

Each switch must be:

  1. (a) Able to carry its rated current;
  2. (b) Accessible to the crew; and
  3. (c) Labelled as to operation and the circuit controlled.

Lights

527.1381 Instrument Lights

The instrument lights must:

  1. (a) Make each instrument, switch, and other devices for which they are provided easily readable; and
  2. (b) Be installed so that:
    1. (1) Their direct rays are shielded from the pilot's eyes; and
    2. (2) No objectionable reflections are visible to the pilot.

527.1383 Landing Lights

  1. (a) Each required landing or hovering light must be approved.
  2. (b) Each landing light must be installed so that:
    1. (1) No objectionable glare is visible to the pilot;
    2. (2) The pilot is not adversely affected by halation; and
    3. (3) It provides enough light for night operation, including hovering and landing.
  3. (c) At least one separate switch must be provided, as applicable:
    1. (1) For each separately installed landing light; and
    2. (2) For each group of landing lights installed at a common location.

527.1385 Position Light System Installation

  1. (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 527.1387 through 527.1397.
  2. (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.
  3. (c) Rear position light. The rear position light must be a white light mounted as far aft as practicable, and must be approved.
  4. (d) Circuit. The two forward position lights and the rear position light must make a single circuit.
  5. (e) Light covers and colour filters. Each light cover or colour filter must be at least flame resistant and may not change colour or shape or lose any appreciable light transmission during normal use.

527.1387 Position Light System Dihedral Angles

  1. (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.
  2. (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° to the left of the first, as viewed when looking forward along the longitudinal axis.
  3. (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° to the right of the first, as viewed when looking forward along the longitudinal axis.
  4. (d) Dihedral angle A (aft) is formed by two intersecting vertical planes making angles of 70° 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.
  5. (e) If the rear position light, when mounted as far aft as practicable in accordance with 527.1385 (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 totalling 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.

527.1389 Position Light Distribution and Intensities

  1. (a) General. The intensities prescribed in this section must be provided by new equipment with light covers and colour 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.
  2. (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. (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 527.1391.
    2. (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 527.1393, where I is the minimum intensity prescribed in 527.1391 for the corresponding angles in the horizontal plane.
    3. (3) Intensities in overlaps between adjacent signals. No intensity in any overlap between adjacent signals may exceed the values in 527.1395, except that higher intensities in overlaps may be used with main beam intensities substantially greater than the minima specified in 527.1391 and 527.1393, 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 527.1395 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.

527.1391 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:

Dihedral angle
(light included)
Angle from right or left of longitudinal axis, measured from dead ahead Intensity
(candles)
L and R
(forward red and green)
0o to 10o 40
  10o to 20o 30
  20o to 110o 5
A (rear white) 110o to 180o 20

527.1393 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:

Angle above or below the horizontal plane Intensity, I
0o 1.00
0o to 5o 0.90
5o to 10o 0.80
10o to 15o 0.70
15o to 20o 0.50
20o to 30o 0.30
30o to 40o 0.10
40o to 90o 0.05

527.1395 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 527.1389 (b)(3).

Overlaps Maximum Intensity
  Area A (candles) Area B (candles)
Green in dihedral angle L 10 4
Red in dihedral angle R 10 1
Green in dihedral angle A 5 1
Red in dihedral angle A 5 1
Rear white in dihedral angle L 5 1
Rear white in dihedral angle R 5 1

Where:

  1. (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° but less than 20°; and
  2. (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°.

527.1397 Colour Specifications

Each position light colour must have the applicable International Commission on Illumination chromaticity co-ordinates as follows:

  1. (a) Aviation red:
    1. "y" is not greater than 0.335; and
    2. "z" is not greater than 0.002.
  2. (b) Aviation green:
    1. "x" is not greater than 0.440-0.320y;
    2. "x" is not greater than y-0.170 ; and
    3. "y" is not less than 0.390-0.170x .
  3. (c) Aviation white:
    1. "x" is not less than 0.300 and not greater than 0.540;
    2. "y" is not less than "x--0.040" or "yc 0-0.010", whichever is the smaller; and
    3. "y" is not greater than "x+0.020" nor "0.636-0.400x";
  4. Where " yc is the "y" co-ordinate of the Planckian radiator for the value of "x" considered.

527.1399 Riding Light

  1. (a) Each riding light required for water operation must be installed so that it can:
    1. (1) Show a white light for at least two nautical miles at night under clear atmospheric conditions; and
    2. (2) Show a maximum practicable unbroken light with the rotorcraft on the water.
  2. (b) Externally hung lights may be used.

527.1401 Anti-collision Light System

  1. (a) General. If certification for night operation is requested, the rotorcraft must have an anti-collision light system that:
    1. (1) Consists of one or more approved anti-collision 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. (2) Meets the requirements of paragraphs (b) through (f) of this section.
  2. (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° above and 30° below the horizontal plane of the rotorcraft, except that there may be solid angles of obstructed visibility totalling not more than 0.5 steradians.
  3. (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 anti-collision 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.
  4. (d) Colour. Each anti-collision light must be aviation red and must meet the applicable requirements of 527.1397.
  5. (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:
     
  6. where:
  7. Ie= effective intensity (candles).
  8. I(t)= instantaneous intensity as a function of time.
  9. t2 - t1 = flash time interval (seconds).
  10. Normally, the maximum value of effective intensity is obtained when t2 and t1 are chosen so that the effective intensity is equal to the instantaneous intensity at t2 and t1.
  11. (f) Minimum effective intensities for anti-collision light. Each anti-collision light effective intensity must equal or exceed the applicable values in the following table:
    Angle above or below the horizontal plane Effective intensity (candles)
    0o to 5o 150
    5o to 10o 90
    10o to 20o 30
    20o to 30o 15

Safety Equipment

527.1411 General

  1. (a) Required safety equipment to be used by the crew in an emergency, such as flares and automatic life raft releases, must be readily accessible.
  2. (b) Stowage provisions for required safety equipment must be furnished and must:
    1. (1) Be arranged so that the equipment is directly accessible and its location is obvious; and
    2. (2) Protect the safety equipment from damage caused by being subjected to the inertia loads specified in 527.561.

527.1413 Safety Belts

Each safety belt must be equipped with a metal to metal latching device.

527.1415 Ditching Equipment

  1. (a) Emergency flotation and signalling equipment required by any applicable operating rule must meet the requirements of this section.
  2. (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.
  3. (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.
  4. (d) Each signalling device must be free from hazard in its operation and must be installed in an accessible location.

527.1419 Ice Protection

  1. (a) To obtain certification for flight into icing conditions, compliance with this section must be shown.
  2. (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 Chapter 529 of this Manual 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.
  3. (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 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. (1) Laboratory dry air or simulated icing tests, or a combination of both, of the components or models of the components.
    2. (2) Flight dry air tests of the ice protection system as a whole, or its individual components.
    3. (3) Flight tests of the rotorcraft or its components in measured simulated icing conditions.
  4. (d) The ice protection provisions of this section are considered to be applicable primarily to the airframe. Powerplant installation requirements are contained in Subchapter E of this chapter.
  5. (e) A means must be identified 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.

527.1435 Hydraulic Systems

  1. (a) Design. Each hydraulic system and its elements must withstand, without yielding, any structural loads expected in addition to hydraulic loads.
  2. (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.
  3. (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.

527.1457 Cockpit Voice Recorders

  1. (a) Each cockpit voice recorder required by the applicable operating rules must be approved and installed so that it will record the following:
    (effective 2014/07/08)
    1. (1) Voice communications transmitted from or received in the rotorcraft by radio.
    2. (2) Voice communications of flight crew members on the flight deck.
    3. (3) Voice communications of flight crew members on the flight deck, using the rotorcraft's interphone system.
    4. (4) Voice or audio signals identifying navigation or approach aids introduced into a headset or speaker.
    5. (5) Voice communications of flight crew members 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. (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.
      (amended 2009/05/11; no previous version)
  2. (b) The recording requirements of paragraph (a)(2) of this section may be met:
    1. (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 crew members on the flight deck when directed to those stations; or
    2. (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 Minister. Repeated aural or visual playback of the record may be used in evaluating intelligibility.
  3. (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. (1) For the first channel, from each microphone, headset, or speaker used at the first pilot station.
    2. (2) For the second channel, from each microphone, headset, or speaker used at the second pilot station.
    3. (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. (4) For the fourth channel, from:
      1. (i) Each microphone, headset, or speaker used at the stations for the third and fourth crew members; or
      2. (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.
      3. (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.
  4. (d) Each cockpit voice recorder must be installed so that:
    1. (1) 
      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,
        (effective 2014/07/08)
      2. (ii) it remains powered for as long as possible without jeopardizing emergency operation of the rotorcraft;
        (effective 2014/07/08)
    2. (2) There is an automatic means to simultaneously stop the recorder and prevent each erasure feature from functioning, within 10 minutes after crash impact; and
    3. (3) There is an aural or visual means for pre-flight checking of the recorder for proper operation.
    4. (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
      (amended 2009/05/11; no previous version)
    5. (5) It has an independent power source:
      (amended 2009/05/11; no previous version)
      1. (i) That provides 10 ± 1 minutes of electrical power to operate both the cockpit voice recorder and cockpit-mounted area microphone;
        (amended 2009/05/11; no previous version)
      2. (ii) That is located as close as practicable to the cockpit voice recorder; and
        (amended 2009/05/11; no previous version)
      3. (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.
        (amended 2009/05/11; no previous version)
  5. (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.
  6. (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.
  7. (g) Each recorder container must be either bright orange or bright yellow.

    (Change 527-1 (89-01-01))

  8. (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 chapter are met.
    (amended 2009/05/11; no previous version)

527.1459 Flight Data Recorders

(effective 2014/07/08)

  1. (a) Each flight data recorder required by the applicable operating rules must be installed so that:
    (effective 2014/07/08)
    1. (1) It is supplied with airspeed, altitude, and directional data obtained from sources that meet the accuracy requirements of 527.1323527.1325, and 527.1327 of this chapter, as applicable;
    2. (2) The vertical acceleration sensor is rigidly attached, and located longitudinally within the approved centre of gravity limits of the rotorcraft;
    3. (3) 
      1. (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; and
        (effective 2014/07/08)
      2. (ii) it remains powered for as long as possible without jeopardizing emergency operation of the rotorcraft;
        (effective 2014/07/08)
    4. (4) There is an aural or visual means for pre-flight checking of the recorder for proper recording of data in the storage medium;
    5. (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. (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.
      (amended 2009/05/11; no previous version)
  2. (b) Each non-ejectable 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.
  3. (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.
  4. (d) Each recorder container must:
    1. (1) Be either bright orange or bright yellow;
    2. (2) Have a reflective tape affixed to its external surface to facilitate its location under water; and
    3. (3) Have an underwater locating device, when required by the operating rules of this manual, on or adjacent to the container which is secured in such a manner that they are not likely to be separated during crash impact.

    (Change 527-1 (89-01-01))

  5. (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 chapter are met.
    (amended 2009/05/11; no previous version)

527.1461 Equipment Containing High Energy Rotors

  1. (a) Equipment containing high energy rotors must meet paragraph (b), (c), or (d) of this section.
  2. (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. (1) Auxiliary rotor cases must be able to contain damage caused by the failure of high energy rotor blades; and
    2. (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.
  3. (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.
  4. (d) Equipment containing high energy rotors must be located where rotor failure will neither endanger the occupants nor adversely affect continued safe flight.

    (Change 527-1 (89-01-01))

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