Part V - Airworthiness Manual Chapter 529 - Transport Category Rotorcraft
On August 31, 2013, the Canadian Aviation Regulations (CARS) will only be available on the Department of Justice website. Standards, incorporated by reference, will continue to be published on the Transport Canada website. Transport Canada will continue to make the index of the CARS and Standards available from its site for the convenience of its users.
Please be advised that there was no scheduled amendment for December 2012 (2012-2). The next amendment is planned for the end of 2013 (2013-1).
Canadian Aviation Regulations (CARs) 2012-1
Content last revised: 2009/12/01
- A (529.1-529.20),
- B (529.21-529.300),
- C (529.301-529.600),
- D (529.601-529.900),
- E (529.901-529.1300),
- F (529.1301-529.1500),
- G (529.1501-529.1589)
SUBCHAPTER D DESIGN AND CONSTRUCTION - GENERAL
(a) The rotorcraft shall have no design features or details that experience has demonstrated to be hazardous or unreliable.
(b) The suitability of each questionable design detail and part shall be established by tests.
529.602 Critical Parts
(amended 2003/06/23; previous version)
(a) Critical part. A critical part is a part, the failure of which could have a catastrophic effect upon the rotorcraft and for which critical characteristics have been identified which, in turn, must be controlled to ensure the required level of integrity.
(amended 2009/12/01; previous version)
(b) If the type design includes critical parts, a critical parts list shall be established. Procedures are established to define the critical design characteristics, identify processes that affect those characteristics and identify the design change and process change controls necessary for showing compliance with the quality assurance requirements of Part V, Subparts 21, 61, 71 and Part VI, Subpart 5 of the Canadian Aviation Regulations.
(amended 2009/12/01; previous version)
The suitability and durability of materials used for parts, the failure of which could adversely affect safety, shall:
(a) be established on the basis of experience or tests;
(b) meet approved specifications that ensure their having the strength and other properties assumed in the design data; and
(c) take into account the effects of environmental conditions, such as temperature and humidity, expected in service.
(a) The methods of fabrication used shall produce consistently sound structures. If a fabrication process (such as gluing, spot welding or heat-treating) requires close control to reach this objective, the process shall be performed according to an approved process specification.
(b) Each new aircraft fabrication method shall be substantiated by a test program.
(a) Each removable bolt, screw, nut, pin or other fastener whose loss could jeopardize the safe operation of the rotorcraft shall incorporate two separate locking devices. The fastener and its locking devices may not be adversely affected by the environmental conditions associated with the particular installation.
(b) No self-locking nut shall be used on any bolt subject to rotation in operation unless a non-friction locking device is used in addition to the self-locking device.
Each part of the structure shall:
(a) be suitably protected against deterioration or loss of strength in service due to any cause, including:
(2) corrosion, and
(3) abrasion; and
(b) have provisions for ventilation and drainage where necessary to prevent the accumulation of corrosive, flammable or noxious fluids.
(a) The rotorcraft structure shall be protected against catastrophic effects from lightning.
(b) For metallic components, compliance with (a) of this section may be demonstrated by:
(1) electrically bonding the components properly to the airframe; or
(2) designing the components so that a strike will not endanger the rotorcraft.
(c) For non-metallic components, compliance with (a) of this section may be demonstrated by:
(1) designing the components to minimize the effect of a strike; or
(2) incorporating acceptable means of diverting the resulting electrical current to not endanger the rotorcraft.
(d) The electric bonding and protection against lightning and static electricity shall:
(1) minimize the accumulation of electrostatic charge;
(2) minimize the risk of electric shock to crew, passengers, and service and maintenance personnel using normal precautions;
(3) provide an electrical return path, under both normal and fault conditions, on rotorcraft having grounded electrical systems; and
(4) reduce to an acceptable level the effects of lightning and static electricity on the functioning of essential electrical and electronic equipment.
There shall be means to allow close examination of each part that requires:
(a) recurring inspection;
(b) adjustment for proper alignment and functioning; or
(a) Material strength properties shall be based on enough tests of material meeting specifications to establish design values on a statistical basis.
(b) Design values shall be chosen to minimize the probability of structural failure due to material variability. Except as provided in (d) and (e) of this section, compliance with (a) of this section shall be demonstrated by selecting design values that assure material strength with the following probability:
(1) where applied loads are eventually distributed through a single member within an assembly, the failure of which would result in loss of structural integrity of the component, 99 percent probability with 95 percent confidence; and
(2) for redundant structures, those in which the failure of individual elements would result in applied loads being safely distributed to other load-carrying members, 90 percent probability with 95 percent confidence.
(c) The strength, detail design, and fabrication of the structure shall minimize the probability of disastrous fatigue failure, particularly at points of stress concentration.
(d) Design values may be those contained in the following publications (available from the Naval Publications and Forms Centre, 5801 Tabor Avenue, Philadelphia, Pennsylvania 19120):
(1) MIL-HDBK-5, "Metallic Materials and Elements for Flight Vehicle Structure";
(2) MIL-HDBK-17, "Plastics for Flight Vehicles";
(3) ANC-18, "Design of Wood Aircraft Structures";
(4) MIL-HDBK-23, "Composite Construction for Flight Vehicles".
(e) Other design values may be used if a selection of the material is made in which a specimen of each individual item is tested before use and it is determined that the actual strength properties of that particular item will equal or exceed those used in design.
(2) likely to deteriorate in service before normal replacement; or
(3) subject to appreciable variability due to:
(i) uncertainties in manufacturing processes, or
(ii) uncertainties in inspection methods.
(2) any other factor great enough to ensure that the probability of the part being understrength because of the uncertainties specified in (a) of this section is extremely remote.
(a) General. The factors, tests and inspections specified in (b) and (c) of this section shall be applied in addition to those necessary to establish foundry quality control. The inspections shall meet approved specifications. The dispositions (c) and (d) of this section apply to structural castings except castings that are pressure tested as parts of hydraulic or other fluid systems and do not support structural loads.
(b) Bearing stresses and surfaces. The casting factors specified in (c) and (d) of this section:
(1) need not exceed 1.25 with respect to bearing stresses regardless of the method of inspection used; and
(2) need not be used with respect to the bearing surfaces of a part whose bearing factor is larger than the applicable casting factor.
(c) Critical castings. For each casting whose failure would preclude continued safe flight and landing of the rotorcraft or result in serious injury to any occupant, the following apply:
(1) each critical casting shall:
(i) have a casting factor of not less than 1.25, and
(ii) receive 100 percent inspection by visual, radiographic and magnetic particle (for ferromagnetic materials) or penetrant (for non-ferromagnetic materials) inspection methods or approved equivalent inspection methods;
(2) for each critical casting with a casting factor less than 1.50, three sample castings shall be static tested and demonstrated to meet:
(i) the strength requirements of section 529.305 at an ultimate load corresponding to a casting factor of 1.25, and
(ii) the deformation requirements of section 529.305 at a load of 1.15 times the limit load.
(d) Non-critical castings. For each casting other than those specified in (c) of this section, the following apply:
(1) except as provided in (d)(2) and (d)(3) of this section, the casting factors and corresponding inspections shall meet the following table:
|2.0 or greater||100% visual|
|Less than 2.0, greater than 1.5||100% visual, and magnetic particle (ferromagnetic materials), penetrant (non-ferromagnetic materials), or approved equivalent inspection methods.|
|1.25 through 1.50||100% visual, and magnetic particle (ferromagnetic materials), penetrant (non-ferromagnetic materials), and radiographic or approved equivalent inspection methods.|
(2) the percentage of castings inspected by non-visual methods may be reduced below that specified in (d)(1) of this section when an approved quality control procedure is established;
(3) for castings procured to a specification that guarantees the mechanical properties of the material in the casting and provides for demonstration of these properties by test of coupons cut from the castings on a sampling basis:
(i) a casting factor of 1.0 may be used, and
(ii) the castings shall be inspected as provided in (d)(1) of this section for casting factors of "1.25 through 1.50" and tested under (c)(2) of this section.
(a) Except as provided in (b) of this section, each part that has clearance (free fit), and that is subject to pounding or vibration, shall have a bearing factor large enough to provide for the effects of normal relative motion.
(b) No bearing factor is needed to be used on a part for which any larger special factor is prescribed.
For each fitting (part or terminal used to join one structural member to another) the following apply:
(a) for each fitting whose strength is not proven by limit and ultimate load tests in which actual stress conditions are simulated in the fitting and surrounding structures, a fitting factor of at least 1.15 shall be applied to each part of:
(1) the fitting,
(2) the means of attachment, and
(3) the bearing on the joined members;
(b) no fitting factor is needed to be used:
(1) for joints made under approved practices and based on comprehensive test data (such as continuous joints in metal plating, welded joints, and scarf joints in wood); and
(2) with respect to any bearing surface for which a larger special factor is used.
(c) for each integral fitting, the part shall be treated as a fitting up to the point at which the section properties become typical of the member;
(d) each seat, berth, litter, safety belt and harness attachment to the structure shall be demonstrated by analysis, tests, or both, to be able to withstand the inertia forces prescribed in section 529.561 (b)(3) multiplied by a factor of 1.33.
Each aerodynamic surface of the rotorcraft shall be free from flutter and divergence under each appropriate speed and power condition.
The rotorcraft shall be designed to ensure capability of continued safe flight and landing (for Category A) or safe landing (for Category B) after impact with a 2.2-lbs. (1.0 kg) bird when the velocity of the rotorcraft (relative to the bird along the flight path of the rotorcraft) is equal to VNE or VH (whichever is the lesser) at altitudes up to 8,000 feet. Compliance shall be demonstrated by tests or by analysis based on tests carried out on sufficiently representative structures of similar design.
(a) For each rotor blade:
(1) there shall be means for venting the internal pressure of the blade;
(2) drainage holes shall be provided for the blade; and
(3) the blade shall be designed to prevent water from becoming trapped in it.
(b) The conditions prescribed under (a) (1) and (2) of this section do not apply to sealed rotor blades capable of withstanding the maximum pressure differentials expected in service.
(a) The rotor and blades shall be mass balanced as necessary to:
(1) prevent excessive vibration; and
(2) prevent flutter at any speed up to the maximum forward speed.
(b) The structural integrity of the mass balance installation shall be substantiated.
There shall be enough clearance between the rotor blades and other parts of the structure to prevent the blades from striking any part of the structure during any operating condition.
(a) The reliability of the means for preventing ground resonance shall be demonstrated either by analysis and tests, or reliable service experience, or by demonstrating through analysis or tests that malfunction or failure of a single means will not cause ground resonance.
(b) The probable range of variations, during service, of the damping action of the ground resonance prevention means shall be established and shall be investigated during the test required by section 529.241.
(a) Each control and control system shall operate with the ease, smoothness and positiveness appropriate to its function.
(b) Each element of each flight control system shall be designed, or distinctively and permanently marked, to minimize the probability of any incorrect assembly that could result in the malfunction of the system.
(c) A means shall be provided to allow full control movement of all primary flight controls prior to flight, or a means shall be provided that will allow the pilot to determine that full control authority is available prior to flight.
If the functioning of stability augmentation or other automatic or power-operated systems is necessary to demonstrate compliance with the flight characteristics requirements of this Chapter, the system shall comply with section 529.671 of this Chapter and the following:
(a) a warning which is clearly distinguishable to the pilot under expected flight conditions without requiring the pilot’s attention shall be provided for any failure in the stability augmentation system or in any other automatic or power-operated system which could result in an unsafe condition if the pilot is unaware of the failure. Warning systems shall not activate the control systems;
(b) the design of the stability augmentation system or of any other automatic or power-operated system shall allow initial counteraction of failures without requiring exceptional pilot skill or strength, by overriding the failure by moving the flight controls in the normal sense, and by deactivating the failed system;
(c) it shall be demonstrated that after any single failure of the stability augmentation system or any other automatic or power-operated system:
(1) the rotorcraft is safely controllable when the failure or malfunction occurs at any speed or altitude within the approved operating limitations;
(2) the controllability and manoeuvrability requirements of this Chapter are met within a practical operational flight envelope (for example, speed, altitude, normal acceleration and rotorcraft configurations) which is described in the Rotorcraft Flight Manual; and
(3) the trim and stability characteristics are not impaired below a level needed to allow continued safe flight and landing.
Primary flight controls are those used by the pilot for immediate control of pitch, roll, yaw and vertical motion of the rotorcraft.
Each primary flight control system shall provide for safe flight and landing and operate independently after a malfunction, failure or jam of any auxiliary interconnected control.
(a) Each control system shall have stops that positively limit the range of motion of the pilot’s controls.
(b) Each stop shall be located in the system so that the range of travel of its control is not appreciably affected by:
(2) slackness; or
(3) take-up adjustments.
(c) Each stop shall be able to withstand the loads corresponding to the design conditions for the system.
(d) For each main rotor blade:
(1) stops that are appropriate to the blade design shall be provided to limit travel of the blade about its hinge points; and
(2) there shall be means to keep the blade from hitting the droop stops during any operation other than starting and stopping the rotor.
If there is a device to lock the control system with the rotorcraft on the ground or water, there shall be means to:
(a) automatically disengage the lock when the pilot operates the controls in a normal manner, or limit the operation of the rotorcraft so as to give unmistakable warning to the pilot before take-off; and
(b) prevent the lock from engaging in flight.
(a) Compliance with the limit load requirements of this Chapter shall be demonstrated by tests in which:
(1) the direction of the test loads produces the most severe loading in the control system; and
(2) each fitting, pulley and bracket used in attaching the system to the main structure is included.
(b) Compliance shall be demonstrated (by analyses or individual load tests) with the special factor requirements for control system joints subject to angular motion.
It shall be demonstrated by operation tests that, when the controls are operated form the pilot compartment with the control system loaded to correspond with loads specified for the system, the system is free from:
(b) excessive friction; and
(c) excessive deflection.
(a) Each detail of each control system shall be designed to prevent jamming, chafing and interference from cargo, passengers, loose objects, or the freezing of moisture.
(b) There shall be means in the cockpit to prevent the entry of foreign objects into places where they would jam the system.
(c) There shall be means to prevent the slapping of cables or tubes against other parts.
(d) Cable systems shall be designed as follows:
(1) cables, cable fittings, turnbuckles, splices and pulleys shall be of an acceptable kind;
(2) the design of cable systems shall prevent any hazardous change in cable tension throughout the range of travel under any operating conditions and temperature variations;
(3) no cable smaller than one eight inch diameter shall be used in any primary control system;
(4) pulley kinds and sizes shall correspond to the cables with which they are used. The pulley-cable combinations and strength values specified in MIL-HDBK-5 shall be used unless they are inapplicable;
(5) pulleys shall have close fitting guards to prevent the cables from being displaced or fouled;
(6) pulleys shall lie close enough to the plane passing through the cable to prevent the cable from rubbing against the pulley flange;
(7) no fairlead shall cause a change in cable direction of more than 3°;
(8) no clevis pin, subject to load or motion and retained only by cotter pins, shall be used in the control system;
(9) turnbuckles, attached to parts having angular motion, shall be installed to prevent binding throughout the range of travel; and
(10) there shall be means for visual inspection at each fairlead, pulley, terminal, and turnbuckle.
(e) Control system joints subject to angular motion shall incorporate the following special factors with respect to the ultimate bearing strength of the softest material used as a bearing:
(1) 3.33 for push-pull systems other than ball and roller bearing systems;
(2) 2.0 for cable systems.
(f) For control system joints, the manufacturer’s static, non-Brinell rating of ball and roller bearings shall not be exceeded.
(a) Each control system spring device whose failure could cause flutter or other unsafe characteristics shall be reliable.
(b) Compliance with (a) of this section shall be demonstrated by tests simulating service conditions.
Each main rotor blade pitch control mechanism shall allow rapid entry into autorotation after power failure.
(a) If a power boost or power-operated control system is used, an alternate system shall be immediately available that allows continued safe flight and landing in the event of:
(1) any single failure in the power portion of the system; or
(2) the failure of all engines.
(b) Each alternate system may be a duplicate power portion or a manually operated mechanical system. The power portion includes the power source (such as hydraulic pumps), and such items as valves, lines and actuators.
(c) The failure of mechanical parts (such as piston rods and links), and the jamming of power cylinders, shall be considered unless they are extremely improbable.
The landing inertia load factor and the reserve energy absorption capacity of the landing gear shall be substantiated by the tests prescribed in sections 529.725 and 529.727 respectively. These tests shall be conducted on the complete rotorcraft or on units consisting of wheel, tire and shock absorber in their proper relation.
The limit drop test shall be conducted as follows:
(a) the drop height shall be at least 8 inches (203 mm);
(b) if considered, the rotor lift specified in section 529.473 (a) shall be introduced into the drop test by appropriate energy absorbing devices or by the use of an effective mass;
(c) each landing gear unit shall be tested in the attitude simulating the landing condition that is most critical from the standpoint of the energy to be absorbed by it;
(d) when an effective mass is used in demonstrating compliance with (b) of this section, the following formula may be used instead of more rational computations:
We = the effective weight to be used in the drop test (lbs);
W = WM for main gear units (lbs), equal to the static reaction on the particular unit with the rotorcraft in the most critical attitude. A rational method may be used in computing a main gear static reaction, taking into consideration the moment arm between the main wheel reaction and the rotorcraft centre of gravity;
W = WN for nose gear units (lbs), equal to the vertical component of the static reaction that would exist at the nose wheel, assuming that the mass of the rotorcraft acts at the centre of gravity and exerts a force of 1.0g downward and 0.25g forward;
W = WT for tailwheel units (lbs), equal to whichever of the following is critical:
(1) the static weight on the tailwheel with the rotorcraft resting on all wheels; or
(2) the vertical component of the ground reaction that would occur at the tailwheel assuming that the mass of the rotorcraft acts at the centre of gravity and exerts a force of 1g downward with the rotorcraft in the maximum nose-up attitude considered in the nose-up landing conditions;
h = specified free drop height (inches);
L = ratio of assumed rotor lift to the rotorcraft weight;
d = deflection under impact of the tire (at the proper inflation pressure) plus the vertical component of the axle travel (inches) relative to the drop mass;
n = limit inertia load factor;
nj = the load factor developed during impact on the mass used in the drop test (i.e., the acceleration dv/dt in g’s recorded in the drop test plus 1.0).
The reserve energy absorption drop test shall be conducted as follows:
(a) the drop height shall be 1.5 times that specified in section 529.725 (a);
(b) rotor lift, where considered in a manner similar to that prescribed in section 529.725 (b), shall not exceed 1.5 times the lift allowed under section 529.725 (b); and
(c) the landing gear shall withstand this test without collapsing. Collapse of the landing gear occurs when a member of the nose, tail or main gear will not support the rotorcraft in the proper attitude or allows the rotorcraft structure, other than landing gear and external accessories, to impact the landing surface.
For rotorcraft with retractable landing gear, the following apply:
(a) Loads. The landing gear, retracting mechanism, wheel well doors and supporting structure shall be designed for:
(1) the loads occurring in any manoeuvring condition with the gear retracted;
(2) the combined friction, inertia and air loads occurring during retraction and extension at any airspeed up to the design maximum landing gear operating speed; and
(3) the flight loads, including those in yawed flight, occurring with the gear extended at any airspeed up to the design maximum landing gear extended speed;
(b) Landing gear lock. A positive means shall be provided to keep the gear extended;
(c) Emergency operation. When other than manual power is used to operate the gear, emergency means shall be provided for extending the gear in the event of:
(1) any reasonably probable failure in the normal retraction system; or
(2) the failure of any single source of hydraulic, electric or equivalent energy;
(d) Operation tests. The proper functioning of the retracting mechanism shall be demonstrated by operation tests;
(e) Position indicator. There shall be means to indicate to the pilot when the gear is secured in the extreme positions;
(g) Landing gear warning. An aural or equally effective landing gear warning device shall be provided that functions continuously when the rotorcraft is in a normal landing mode and the landing gear is not fully extended and locked. A manual shut-off capability shall be provided for the warning device and the warning system shall automatically reset when the rotorcraft is no longer in the landing mode.
(a) Each landing gear wheel shall be approved.
(b) The maximum static load rating of each wheel shall not be less than the corresponding static ground reaction with:
(1) maximum weight; and
(2) critical centre of gravity.
(c) The maximum limit load rating of each wheel shall equal or exceed the maximum radial limit load determined under the applicable ground load requirements of this Chapter.
Each landing gear wheel shall have a tire:
(a) that is a proper fit on the rim of the wheel;
(b) of a rating that is not exceeded under:
(1) the design maximum weight,
(2) a load on each main wheel tire equal to the static ground reaction corresponding to the critical centre of gravity, and
(3) a load on nose wheel tires (to be compared with the dynamic rating established for those tires) equal to the reaction obtained at the nose wheel, assuming that the mass of the rotorcraft acts at the most critical centre of gravity and exerts a force of 1.0 g. downward and 0.25 g. forward, the reactions being distributed to the nose and main wheels according to the principles of statics with the drag reaction at the ground applied only at wheels with brakes; and
(c) each tire installed on a retractable landing gear system shall, at the maximum size of the tire type expected in service, have a clearance to surrounding structure and systems that is adequate to prevent contact between the tire and any part of the structure or systems.
For rotorcraft with wheel-type landing gear, a braking device shall be installed that is:
(a) controllable by the pilot;
(b) useable during power-off landings; and
(c) adequate to:
(1) counteract any normal unbalanced torque when starting or stopping the rotor;and
(2) hold the rotorcraft parked on a 10 degree slope on a dry, smooth pavement.
(a) The maximum limit load rating of each ski shall equal or exceed the maximum limit load determined under the applicable ground load requirements of this Chapter.
(b) There shall be a stabilizing means to maintain the ski in an appropriate position during flight. This means shall have enough strength to withstand the maximum aerodynamic and inertia loads on the ski.
Floats and Hulls
(a) For main floats, the buoyancy necessary to support the maximum weight of the rotorcraft in fresh water shall be exceeded by:
(1) 50 percent, for single floats; and
(2) 60 percent, for multiple floats.
(b) Each main float shall have enough watertight compartments so that, with any single main float compartment flooded, the main floats will provide a margin of positive stability great enough to minimize the probability of capsizing.
(a) Bag floats. Each bag float shall be designed to withstand:
(1) the maximum pressure differential that might be developed at the maximum altitude for which certification with the float is requested; and
(2) the vertical loads prescribed in section 529.521 (a), distributed along the length of the bag over three quarters of its projected area.
(b) Rigid floats. Each rigid float shall be able to withstand the vertical, horizontal and side loads prescribed in section 529.521. An appropriate load distribution under critical conditions shall be used.
Water-based and amphibian rotorcraft. The hull and auxiliary floats, if used, shall have enough watertight compartments so that, with any single compartment of the hull or auxiliary floats flooded, the buoyancy of the hull and auxiliary floats, and wheel tires if used, provides a margin of positive water stability great enough to minimize the probability of capsizing the rotorcraft for the worst combination of wave heights and surface winds for which approval is desired.
The hull, and auxiliary floats if used, shall withstand the water loads prescribed in section 529.519 with a rational and conservative distribution of local and distributed water pressures over the hull and float bottom.
Personnel and Cargo Accommodations
For each pilot compartment:
(a) the compartment and its equipment shall allow each pilot to perform his duties without unreasonable concentration or fatigue;
(b) if there is provision for a second pilot, the rotorcraft shall be controllable with equal safety from either pilot position. Flight and powerplant controls shall be designed to prevent confusion or inadvertent operation when the rotorcraft is piloted from either position;
(c) the vibration and noise characteristics of cockpit appurtenances shall not interfere with safe operation;
(d) in-flight leakage of rain or snow that could distract the crew or harm the structure shall be prevented.
(a) Nonprecipitation conditions. For nonprecipitation conditions, the following apply:
(1) each pilot compartment shall be arranged to give the pilots a sufficiently extensive, clear, and undistorted view for safe operation;
(2) each pilot compartment shall be free of glare and reflection that could interfere with the pilot’s view. If certification for night operation is requested, this shall be demonstrated by night flight tests.
(b) Precipitation conditions. For precipitation conditions, the following apply:
(1) each pilot shall have a sufficiently extensive view for safe operation:
(i) in heavy rain at forward speeds up to VH, and
(ii) in the most severe icing condition for which certification is requested;
(2) the first pilot shall have a window that:
(i) is openable under the conditions prescribed in (b)(1) of this section; and
(ii) provides the view prescribed in (b)(1) of this section.
Windshields and windows shall be made of material that will not break into dangerous fragments.
Cockpit controls shall be:
(a) located to provide convenient operation and to prevent confusion and inadvertent operation; and
(b) located and arranged with respect to the pilot’s seats so that there is full and unrestricted movement of each control without interference from the cockpit structure or the pilot’s clothing when pilots from 5’2" to 6’0" in height are seated.
Cockpit controls shall be designed so that they operate in accordance with the following movements and actuation:
(a) flight controls, including the collective pitch control, shall operate with a sense of motion which corresponds to the effect on the rotorcraft;
(b) twist-grip engine power controls shall be designed so that, for left-hand operation, the motion of the pilot’s hand is clockwise to increase power when the hand is viewed from the edge containing the index finger. Other engine power controls, excluding the collective control, shall operate with a forward motion to increase power; and
(c) normal landing gear controls shall operate downward to extend the landing gear.
(a) Each closed cabin shall have at least one adequate and easily accessible external door.
(b) Each external door shall be located, and appropriate operating procedures shall be established, to ensure that, persons using the door will not be endangered by the rotors, propellers, engine intakes and exhausts when the operating procedures are used.
(c) There shall be means for locking crew and external passenger doors and for preventing their opening in flight inadvertently or as a result of mechanical failure. It shall be possible to open external doors from inside and outside the cabin with the rotorcraft on the ground even though persons may be crowded against the door on the inside of the rotorcraft. The means of opening shall be simple and obvious and so arranged and marked that it can be readily located and operated.
(d) There shall be reasonable provisions to prevent the jamming of any external doors in a minor crash as a result of fuselage deformation under the following ultimate inertial forces except for cargo or service doors not suitable for use as an exit in an emergency:
(1) upward - 1.5g;
(2) forward - 4.0g;
(3) sideward - 2.0g; and
(4) downward - 4.0g.
(e) There shall be means for direct visual inspection of the locking mechanism by crew members to determine whether the external doors (including passenger, crew, service and cargo doors) are fully locked. There shall be visual means to signal to appropriate crew members when normally used external doors are closed and fully locked.
(f) For outward opening external doors useable for entrance or egress, there shall be an auxiliary safety latching device to prevent the door from opening when the primary latching mechanism fails. If the door does not meet the requirements prescribed in (c) of this section with this device in place, suitable operating procedures shall be established to prevent the use of the device during take-off and landing.
(g) If an integral stair is installed in a passenger entry door that is qualified as a passenger emergency exit, the stair shall be designed so that under the following conditions the effectiveness of passenger emergency egress will not be impaired:
(1) the door, integral stair and operating mechanism have been subjected to the inertial forces specified in (d) of this section, acting separately relative to the surrounding structure;
(2) the rotorcraft is in the normal ground attitude and in each of the attitudes corresponding to collapse of one or more legs, or primary members, as applicable, of the landing gear.
(h) Non-jettisonable doors used as ditching emergency exits shall have means to enable them to be secured in the open position and remain secure for emergency egress in sea state conditions prescribed for ditching.
(a) Each seat, safety belt, harness and adjacent part of the rotorcraft at each station designated for occupancy during take-off and landing shall be free of potentially injurious objects, sharp edges, protuberances and hard surfaces and shall be designed so that a person making proper use of these facilities will not suffer serious injury in an emergency landing as a result of the inertial factors specified in section 529.561 (b) and dynamic conditions specified in section 529.562.
(b) Each occupant shall be protected from serious head injury by a safety belt plus a shoulder harness that will prevent the head from contacting any injurious object except as provided for in section 529.562 (c)(5). A shoulder harness (upper torso restraint), in combination with the safety belt, constitutes a torso restraint system as described in TSO-C114.
(c) Each occupant’s seat shall have a combined safety belt and shoulder harness with a single point release. Each pilot’s combined safety belt and shoulder harness shall allow each pilot when seated with safety belt and shoulder harness fastened to perform all functions necessary for flight operations. There shall be a means to secure belts and harnesses when not in use to prevent interference with the operation of the rotorcraft and with rapid egress in an emergency.
(d) If seat backs do not have a firm handhold, there shall be hand grips or rails along each aisle to let the occupants steady themselves while using the aisle in moderately rough air.
(e) Each projecting object that would injure persons seated or moving about in the rotorcraft in normal flight shall be padded.
(f) Each seat and its supporting structure shall be designed for an occupant weight of 170 pounds, considering the maximum load factors, inertia forces and reactions between the occupant, seat, and safety belt or harness corresponding with the applicable flight and ground-load conditions, including the emergency landing conditions prescribed in section 529.561 (b). In addition:
(1) each pilot seat shall be designed for the reactions resulting from the application of the pilot forces prescribed in section 529.397; and
(2) the inertial forces prescribed in section 529.561 (b) shall be multiplied by a factor of 1.33 in determining the strength of the attachment of:
(i) each seat to the structure, and
(ii) each safety belt or harness to the seat or structure.
(g) When the safety belt and shoulder harness are combined, the rated strength of the safety belt and shoulder harness shall not be less than that corresponding to the inertial forces specified in section 529.561(b), considering an occupant weight of at least 170 pounds, considering the dimensional characteristics of the restraint system installation, and using a distribution of at least 60 percent load to the safety belt and at least 40 percent load to the shoulder harness. If the safety belt is capable of being used with out the shoulder harness, the inertial forces specified shall be met by the safety belt alone.
(h) When a headrest is used, the headrest and its supporting structure shall be designed to resist the inertia forces specified in section 529.561, with a 1.33 fitting factor and a head weight of at least 13 pounds.
(i) Each seating device system includes the device such as the seat, the cushions, the occupant restraint system and attachment devices.
(j) Each seating device system may use design features such as crushing or separation of certain parts of the seat in the design to reduce occupant loads for the emergency landing dynamic conditions of section 529.562; otherwise, the system shall remain intact and shall not interfere with rapid evacuation of the rotorcraft.
(k) For the purpose of this section, a litter is defined as a device designed to carry a non-ambulatory person, primarily in a recumbent position, into and on the rotorcraft. Each berth or litter shall be designed to withstand the load reaction of an occupant weight of at least 170 pounds when the occupant is subjected to the forward inertial factors specified in section 529.561 (b). A berth or litter installed within 15° or less of the longitudinal axis of the rotorcraft shall be provided with a padded end-board, cloth diaphragm or equivalent means that can withstand the forward load reaction. A berth or litter oriented greater than 15° with the longitudinal axis of the rotorcraft shall be equipped with appropriate restraints, such as straps or safety belts, to withstand the forward load reaction. In addition:
(1) the berth or litter shall have a restraint system and shall not have corners or other protuberances likely to cause serious injury to a person occupying it during emergency landing conditions; and
(2) the berth or litter attachment and the occupant restraint system attachments to the structure shall be designed to withstand the critical loads resulting from flight and ground load conditions and from the conditions prescribed in section 529.561 (b). The fitting factor required by section 529.625 (d) shall be applied.
(a) Each cargo and baggage compartment shall be designed with the prescribed conditions of maximum weight of contents and for the critical load distributions at the appropriate maximum load factors corresponding to the specified flight and ground load conditions, except the emergency landing conditions of section 529.561.
(b) There shall be means to prevent the contents of any compartment from becoming a hazard by shifting under the loads specified in (a) of this section.
(c) Under the emergency landing conditions of section 529.561, cargo and baggage compartments shall:
(1) be positioned so that if the contents break loose, they are unlikely to cause injury to the occupants or restrict any of the escape facilities provided for use after an emergency landing; or
(2) have sufficient strength to withstand the conditions specified in section 529.561 including the means of restraint, and their attachments, required by (b) of this section. Sufficient strength shall be provided for the maximum authorized weight of cargo and baggage at the critical loading distribution.
(d) If cargo compartment lamps are installed, each lamp shall be installed so as to prevent contact between lamp bulb and cargo.
(b) Each practicable design measure, compatible with the general characteristics of the rotorcraft, shall be taken to minimize the probability that in an emergency landing on water, the behaviour of the rotorcraft would cause immediate injury to the occupants or would make it impossible for them to escape.
(c) The probable behaviour of the rotorcraft in a water landing shall be investigated by model tests or by comparison with rotorcraft of similar configuration for which the ditching characteristics are known. Scoops, flaps, projections and any other factors likely to affect the hydrodynamic characteristics of the rotorcraft shall be considered.
(d) It shall be demonstrated that, under reasonably probable water conditions, the flotation time and trim of the rotorcraft will allow the occupants to leave the rotorcraft and enter the life rafts required by section 529.1415. If compliance with this provision is demonstrated by buoyancy and trim computations, appropriate allowances shall be made for probable structural damage and leakage. If the rotorcraft has fuel tanks (with fuel jettisoning provisions) that can reasonably be expected to withstand a ditching without leakage, the jettisonable volume of fuel may be considered as buoyancy volume.
(e) Unless the effects of the collapse of external doors and windows are accounted for in the investigation of the probable behaviour of the rotorcraft in a water landing (as prescribed in (c) and (d) of this section), the external doors and windows shall be designed to withstand the probable maximum local pressures.
(a) Each crew and passenger area shall have means for rapid evacuation in a crash landing, with the landing gear (1) extended and (2) retracted, considering the possibility of fire.
(d) Except as provided in (e) of this section, the following categories of rotorcraft shall be tested in accordance with the requirements of Appendix D of this Chapter to demonstrate that the maximum seating capacity, including the crew members required by the operating rules, can be evacuated from the rotorcraft to the ground within 90 seconds:
(1) rotorcraft with a seating capacity of more than 44 passengers;
(2) rotorcraft with all of the following:
(i) ten or more passengers per passenger exit as determined under section 529.807 (b),
(ii) no main aisle, as described in section 529.815, for each row of passenger seats,
(iii) access to each passenger exit for each passenger by virtue of design features of seats, such as folding or break-over seat backs or folding seats.
(e) A combination of analysis and tests may be used to demonstrate that the rotorcraft is capable of being evacuated within 90 seconds under the conditions specified in section 529.803 (d) if the Minister finds that the combination of analysis and tests will provide data, with respect to the emergency evacuation capability of the rotorcraft, equivalent to that which would be obtained by actual demonstration.
(a) For rotorcraft with passenger emergency exits that are not convenient to the flight crew, there shall be flight crew emergency exits, on both sides of the rotorcraft or as a top hatch, in the flight crew area.
(b) Each flight crew emergency exit shall be of sufficient size and shall be located so as to allow rapid evacuation of the flight crew. This shall be demonstrated by test.
(c) Each exit shall not be obstructed by water or flotation devices after a ditching. This shall be demonstrated by test, demonstration or analysis.
(a) Type. For the purpose of this Chapter, the types of passenger emergency exits are as follows:
(1) Type I. This type shall have a rectangular opening of not less than 24 inches (610 mm) wide by 48 inches (1 220 mm) high, with corner radii not greater than one-third the width of the exit, in the passenger area in the side of the fuselage at floor level and as far away as practicable from areas that might become potential fire hazards in a crash;
(2) Type II. This type is the same as Type I, except that the opening shall be at least 20 inches (508 mm) wide by 44 inches (1 117 mm) high;
(3) Type III. This type is the same as Type I, except that:
(i) the opening shall be at least 20 inches (508 mm) wide by 36 inches (915 mm) high, and
(ii) the exits need not be at floor level;
(4) Type IV. This type shall have a rectangular opening of not less than 19 inches (483 mm) wide by 26 inches (660 mm) high, with corner radii not greater than one-third the width of the exit, in the side of the fuselage with a step-up inside the rotorcraft of not more than 29 inches (736 mm).
Openings with dimensions larger than those specified in this section may be used, regardless of shape, if the base of the opening has a flat surface of not less than the specified width.
(b) Passenger emergency exits; side of fuselage. Emergency exits shall be accessible to the passengers and, except as provided in (d) of this section, shall be provided in accordance with the following table:
|Passenger Seating Capacity||Emergency Exits For Each Side Of The Fuselage|
|Type I||Type II||Type III||Type IV|
|1 through 10||.........||.........||.........||1|
|11 through 19||.........||.........||1 or||2|
|20 through 39||.........||1||.........||1|
|40 through 59||1||.........||.........||1|
|60 through 79||1||.........||1 or||2|
Refer to Preamble Change 529-4.
(c) Passenger emergency exits with the rotorcraft resting on its side. In addition to the requirements prescribed in (b) of this section:
(1) there shall be enough openings in the fuselage to allow egress with the rotorcraft on its side;
FAR: (c) Passenger emergency exits with the rotorcraft resting on its side. In addition to the requirements of paragraph (b) of this section:
(1) There must be enough openings in the top, bottom, or ends of the fuselage to allow evacuation with the rotorcraft on its side; or
(2) The probability of the rotorcraft coming to rest on its side in a crash landing must be extremely remote.
Refer to Preamble Change 529-4.
(d) Ditching emergency exits for passengers. If certification with ditching provisions is requested, ditching emergency exits shall be provided in accordance with the following requirements and shall be proven by test, demonstration or analysis unless the emergency exits required by (b) of this section already meet them:
(1) for rotorcraft that have a passenger seating configuration, excluding pilot seats, of nine seats or less, one exit above the waterline in each side of the rotorcraft, meeting at least the dimensions of a Type IV exit;
(2) for rotorcraft that have a passenger seating configuration, excluding pilot seats, of 10 seats or more, one exit above the waterline in a side of the rotorcraft meeting at least the dimensions of a Type III exit, for each unit (or part of a unit) of 35 passenger seats, but no less than two such exits in the passenger cabin, with one on each side of the rotorcraft. However, where it has been demonstrated through analysis, ditching demonstrations, or any other tests found necessary by the Minister, that the evacuation capability of the rotorcraft during ditching is improved by the use of larger exits, or by other means, the passenger seat to exit ratio may be increased;
(3) flotation devices, whether stowed or deployed, shall not interfere with or obstruct the exits.
(e) Ramp exits. One Type I exit only, or one Type II exit only, that is required in the side of the fuselage under (b) of this section, may be installed instead in the ramp of floor ramp rotorcraft if:
(1) its installation in the side of the fuselage is impractical; and
(2) its installation in the ramp meets the requirements prescribed in section 529.813.
(f) Tests. The proper functioning of each emergency exit shall be demonstrated by test.
(a) Each emergency exit shall consist of a movable door or hatch in the external walls of the fuselage and shall provide an unobstructed opening to the outside.
(b) Each emergency exit shall be openable from the inside and from the outside.
(c) The means of opening each emergency exit shall be simple and obvious and shall not require exceptional effort.
(d) There shall be means for locking each emergency exit and for preventing opening in flight inadvertently or as a result of mechanical failure.
(e) There shall be means to minimize the probability of the jamming of any emergency exit in a minor crash landing as a result of fuselage deformation under the ultimate inertial forces prescribed in section 529.783(d).
(f) Except as provided in (h) of this section, each land-based rotorcraft emergency exit shall have an approved slide as stated in (g) of this section, or its equivalent, to assist occupants in descending to the ground from each floor level exit and an approved rope, or its equivalent, for all other exits, if the exit threshold is more that 6 feet above the ground:
(1) with the rotorcraft on the ground and with the landing gear extended;
(2) with one or more legs or part of the landing gear collapsed, broken or not extended; and
(3) with the rotorcraft resting on its side, if required by section 529.803 (d).
(g) The slide for each passenger emergency exit shall be a self-supporting slide or equivalent, and shall be designed to meet the following requirements:
(1) it shall be automatically deployed, and deployment shall begin during the interval between the time the exit opening means is actuated from inside the rotorcraft and the time the exit is fully opened. However, each passenger emergency exit which is also a passenger entrance door or a service door shall be provided with means to prevent deployment of the slide when the exit is opened from either the inside or the outside under non emergency conditions for normal use;
(2) it shall be automatically erected within 10 seconds after deployment is begun;
(3) it shall be of such length after full deployment that the lower end is self-supporting on the ground and provides safe evacuation of occupants to the ground after collapse of one or more legs or part of the landing gear;
(4) it shall have the capability, in 25-knot winds directed from the most critical angle, to deploy and, with the assistance of only one person, to remain useable after full deployment to evacuate occupants safely to the ground;
(5) each slide installation shall be qualified by five consecutive deployment and inflation tests conducted (per exit) without failure, and at least three tests of each such five-test series shall be conducted using a single representative sample of the device. The sample devices shall be deployed and inflated by the system’s primary means after being subjected to the inertia forces specified in section 529.561 (b). If any part of the system fails or does not function properly during the required tests, the cause of the failure or malfunction shall be corrected by positive means and after that, the full series of five consecutive deployment and inflation tests shall be conducted without failure.
(h) For rotorcraft having 30 or fewer passenger seats and having an exit threshold more than 6 feet above the ground, a rope or other assist means may be used in place of the slide specified in (f) of this section, provided an evacuation demonstration is accomplished as prescribed in section 529.803 (d) or (e).
(i) If a rope, with its attachment, is used for compliance with (f), (g), or (h) of this section, it shall:
(1) withstand a 400-pound static load; and
(2) attach to the fuselage structure at or above the top of the emergency exit opening, or at another approved location if the stowed rope would reduce the pilot’s view in flight.
(a) Each passenger emergency exit, its means of access, and its means of opening shall be conspicuously marked for the guidance of occupants using the exits in daylight or in the dark. Such markings shall be designed to remain visible for rotorcraft equipped for overwater flights if the rotorcraft is capsized and the cabin is submerged.
(b) The identity and location of each passenger emergency exit shall be recognizable from a distance equal to the width of the cabin.
(c) The location of each passenger emergency exit shall be indicated by a sign visible to occupants approaching along the main passenger aisle. There shall be a locating sign:
(1) next to or above the aisle near each floor emergency exit, except that one sign may serve two exits if both exits can be seen readily from that sign; and
(2) on each bulkhead or divider that prevents fore and aft vision along the passenger cabin, to indicate emergency exits beyond and obscured by it, except that if this is not possible the sign may be placed at another appropriate location.
(d) Each passenger emergency exit marking and each locating sign shall have white letters 1 inch (25.4 mm) high on a red background 2 inches (50.8 mm) high, be self or electrically illuminated, and have a minimum luminescence (brightness) of at least 160 microlamberts. The colours may be reversed if this will increase the emergency illumination of the passenger compartment.
(e) The location of each passenger emergency exit operating handle and instructions for opening shall be identified:
(1) for each emergency exit, by a marking on or near the exit that is readable from a distance of 30 inches (760 mm); and
(2) for each Type I or Type II emergency exit with a locking mechanism released by rotary motion of the handle, by:
(i) a red arrow, with a shaft at least three-fourths inch (19 mm) wide and a head twice the width of the shaft, extending along at least 70° of arc at a radius approximately equal to three-fourths of the handle length, and
(ii) the word "open" in red letters 1 inch (25.4 mm) high, placed horizontally near the head of the arrow.
(f) Each emergency exit, and its means of opening, shall be marked on the outside of the rotorcraft. In addition, the following apply:
(1) there shall be a 2-inch (50.8 mm) coloured band outlining each passenger emergency exit, except small rotorcraft with a maximum weight or 12,500 pounds or less may have a 2-inch coloured band outlining each exit release lever or device of passenger emergency exits which are normally used doors; and
(2) each outside marking, including the band, shall have colon contrast to be readily distinguishable from the surrounding fuselage surface. The contrast shall be such that, if the reflectance of the darker colon is 15 percent or less, the reflectance of the lighter colon shall be at least 45 percent. "Reflectance" is the ratio of the luminous flux reflected by a body to the luminous flux it receives. When the reflectance of the darker colon is greater than 15 percent, at least a 30 percent difference between its reflectance and the reflectance of the lighter colon shall be provided.
(g) Exits marked as such, though in excess of the required number of exits, shall meet the requirements for emergency exits of the particular type. Emergency exits need only be marked with the word "Exit."
For transport Category A rotorcraft, the following apply:
(a) A source of light with its power supply independent of the main lighting system shall be installed to:
(1) illuminate each passenger emergency exit marking and locating sign; and
(2) provide enough general lighting in the passenger cabin so that the average illumination, when measured at 40-inch (1.02 m) intervals at seat armrest height on the centre line of the main passenger aisle, is at least 0.05 foot-candles (0.538 lux);
(b) Exterior emergency lighting shall be provided at each emergency exit. The illumination shall not be less than 0.05 foot-candles (0.538 lux) (measured normal to the direction of incident light) for minimum width on the ground surface, with landing gear extended, equal to the width of the emergency exit where an evacuee is likely to make first contact with the ground outside the cabin. The exterior emergency lighting may be provided by either interior or exterior sources with light intensity measurements made with the emergency exits open;
(c) Each light required by (a) or (b) of this section, shall be operable manually from the cockpit station and from a point in the passenger compartment that is readily accessible. The cockpit control device shall have an "on", "off", and "armed" position so that when turned on at the cockpit or passenger compartment station or when armed at the cockpit station, the emergency lights will either illuminate or remain illuminated upon interruption of the rotorcraft’s normal electric power;
(d) Any means required to assist the occupants in descending to the ground shall be illuminated so that the erected assist means is visible from the rotorcraft. The following apply:
(1) the assist means shall be provided with an illumination of not less than 0.03 foot-candles (0.324 lux) (measured normal to the direction of the incident light) at the ground end of the erected assist means where an evacuee using the established escape route would normally make first contact with the ground, with the rotorcraft in each of the attitudes corresponding to the collapse of one or more legs of the landing gear;
(2) if the emergency lighting subsystem illuminating the assist means is independent of the rotorcraft’s main emergency lighting system, it:
(i) shall automatically be activated when the assist means is erected,
(ii) shall provide the illumination required by (d)(1) of this section, and
(iii) shall not be adversely affected by stowage;
(e) The energy supply to each emergency lighting unit shall provide the required level of illumination for at least 10 minutes at the critical ambient conditions after an emergency landing;
(f) If storage batteries are used as the energy supply for the emergency lighting system, they may be recharged from the rotorcraft’s main electrical power system provided the charging circuit is designed to preclude inadvertent battery discharge into charging circuit faults.
(a) Each passageway between passenger compartments, and each passageway leading to Type I and Type II emergency exits, shall be:
(1) unobstructed; and
(2) at least 20 inches (508 mm) wide.
(b) For each emergency exit covered by section 529.809 (f), there shall be enough space adjacent to that exit to allow a crew member to assist in the evacuation of passengers without reducing the unobstructed width of the passageway below that required for that exit.
(c) There shall be access from each aisle to each Type III and Type IV exit, and:
(1) for rotorcraft that have a passenger seating configuration, excluding pilot seats, of 20 or more, the projected opening of the exit provided shall not be obstructed by seats, berths, or other protrusions (including seatbacks in any position) for a distance from that exit of not less than the width of the narrowest passenger seat installed on the rotorcraft;
(2) for rotorcraft that have a passenger seating configuration, excluding pilot seats, of 19 or less, there may be minor obstructions in the region described in (c)(1) of this section, if there are compensating factors to maintain the effectiveness of the exit.
(d) It shall be demonstrated through the design of the rotorcraft that there is easy access to each usable emergency exit when the rotorcraft is resting on its side.
No equivalent text is in the FAR.
(1) Refer to Preamble Change 529-4.
The main passenger aisle width between seats shall equal or exceed the values in the following table:
|Passenger Seating Capacity||Minimum Main Passenger Aisle Width|
|Less than 25 inches (635 mm) from floor||25 inches (635 mm) and more from floor|
|10 or less||12* inches (305 mm)||15 inches (457 mm)|
|11 through 19||12 inches (305 mm)||20 inches (508 mm)|
|20 or more||15 inches (381 mm)||20 inches (508 mm)|
|* A narrow width not less than 9 inches (229 mm) may be approved when substantiated by tests found necessary by the Minister.|
(a) Each passenger and crew compartment shall be ventilated, and each crew compartment shall have enough fresh air (but not less than 10 cu. ft. per minute per crew member) to let crew members perform their duties without undue discomfort or fatigue.
(b) Crew and passenger compartment air shall be free from harmful or hazardous concentrations of gases or vapours.
(c) The concentration of carbon monoxide shall not exceed one part in 20,000 parts of air during forward flight. If the concentration exceeds this value under other conditions, there shall be suitable operating restrictions.
(d) There shall be means to ensure compliance with (b) and (c) of this section under any reasonably probable failure of any ventilating, heating, or other system or equipment.
Each combustion heater shall be approved.
(a) Hand fire extinguishers. For hand fire extinguishers the following apply:
(1) each hand fire extinguisher shall be approved;
(2) the kinds and quantities of each extinguishing agent used shall be appropriate to the kinds of fires likely to occur where that agent is used; and
(3) each extinguisher for use in a personnel compartment shall be designed to minimize the hazard of toxic gas concentrations.
(b) Built-in fire extinguishers. If a built-in fire extinguishing system is required:
(1) the capacity of each system, in relation to the volume of the compartment where used and the ventilation rate, shall be adequate for any fire likely to occur in that compartment;
(2) each system shall be installed so that:
(i) no extinguishing agent likely to enter personnel compartments will be present in a quantity that is hazardous to the occupants, and
(ii) no discharge of the extinguisher can cause structural damage.
For each compartment to be used by the crew or passengers, the requirements of this section shall be met.
(a) The materials (including finishes or decorative surfaces applied to the materials) shall meet the following test criteria as applicable:
(1) interior ceiling panels, interior wall panels, partitions, galley structure, large cabinet walls, structural flooring, and materials used in the construction of stowage compartments (other than underseat stowage compartments and compartments for stowing small items such as magazines and maps) shall be self-extinguishing when tested vertically in accordance with the applicable portions of Appendix F of Chapter 525 of this Manual, or other approved equivalent methods. The average burn length shall not exceed 6 inches (152 mm) and the average flame time after removal of the flame source shall not exceed 15 seconds. Drippings from the test specimen shall not continue to flame for more than an average of 3 seconds after falling;
(2) floor covering, textiles (including draperies and upholstery), seat cushions, padding, decorative and non-decorative coated fabrics, leather, trays and galley furnishings, electrical conduit, thermal and acoustical insulation and insulation covering, air ducting joint and edge covering, cargo compartment liners, insulation blankets, cargo cover, and transparencies, moulded and thermoformed parts, air ducting joints, and trim strips (decorative and chafing) that are constructed of materials not covered in (a)(3) of this section, shall be self extinguishing when tested vertically in accordance with the applicable portion of Appendix F of Chapter 525 of this Manual, or other approved equivalent methods. The average burn length shall not exceed 8 inches (203 mm) and the average flame time after removal of the flame source shall not exceed 15 seconds. Drippings from the test specimen shall not continue to flame for more than an average of 5 seconds after falling;
(3) acrylic windows and signs, parts constructed in whole or in part of elastometric materials, edge lighted instrument assemblies consisting of two or more instruments in a common housing, seat belts, shoulder harnesses, and cargo and baggage tie-down equipment, including containers, bins, pallets, etc., used in passenger or crew compartments, shall not have an average burn rate greater than 2.5 inches (63 mm) per minute when tested horizontally in accordance with the applicable portions of Appendix F of Chapter 525 of this Manual, or other approved equivalent methods;
(4) except for electrical wire and cable insulation, and for small parts (such as knobs, handles, rollers, fasteners, clips, grommets, rub strips, pulleys, and small electrical parts) that the Minister finds would not contribute significantly to the propagation of a fire, materials and items not specified in (a)(1), (a)(2), or (a)(3) of this section shall not have a burn rate greater than 4 inches (102 mm) per minute when tested horizontally in accordance with the applicable portions of Appendix F of Chapter 525 of this Manual, or other approved equivalent methods.
(b) In addition to meeting the requirements prescribed in (a)(2) of this section, seat cushions, except those on flight crew member seats, shall meet the test requirements of Part II of Appendix F of Chapter 525 of this Manual, or equivalent.
(c) If smoking is to be prohibited, there shall be a placard so stating, and if smoking is to be allowed:
(1) where shall be an adequate number of self-contained removable ashtrays, and
(2) where the crew compartment is separated from the passenger compartment, there shall be at least one illuminated sign (using either letters or symbols) notifying all passengers when smoking is prohibited. Signs, which notify when smoking is prohibited, shall:
(i) when illuminated, be legible to each passenger seated in the passenger cabin under all probable lighting conditions, and
(ii) be so constructed that the crew can turn the illumination on and off;
(d) Each receptacle for towels, paper, or waste shall be at least fire-resistant and shall have means for containing possible fires;
(e) There shall be a hand fire extinguisher for the flight crew members;
(f) At least the following number of hand fire extinguishers shall be conveniently located in passenger compartments:
|Passenger Capacity||Fire Extinguishers|
|7 through 30||1|
|31 through 60||2|
|61 or more||3|
(a) Each cargo and baggage compartment shall be constructed of, or lined with materials in accordance with the following:
(1) for accessible and inaccessible compartments not occupied by passengers or crew, the material shall be at least fire resistant;
(2) materials shall meet the requirements prescribed in section 529.853 (a)(1), (a)(2), and (a)(3) for cargo or baggage compartments in which:
(i) the presence of a compartment fire would be easily discovered by a crew member while at the crew member’s station,
(ii) each part of the compartment is easily accessible in flight,
(iii) the compartment has a volume of 200 cubic feet or less, and
(iv) notwithstanding section 529.1439 (a), protective breathing equipment is not required.
(b) No compartment shall contain any controls, wiring, lines, equipment, or accessories whose damage or failure would affect safe operation, unless those items are protected so that:
(1) they cannot be damaged by the movement of cargo in the compartment; and
(2) their breakage or failure will not create a fire hazard.
(c) The design and sealing of inaccessible compartments shall be adequate to contain compartment fires until a landing and safe evacuation can be made.
(d) Each cargo and baggage compartment that is not sealed so as to contain cargo compartment fires completely without endangering the safety of a rotorcraft or its occupants shall be designed, or shall have a device, to ensure detection of fires or smoke by a crew member while at his station and to prevent the accumulation of harmful quantities of smoke, flame, extinguishing agents, and other noxious gases in any crew or passenger compartment. This shall be demonstrated in flight.
(e) For rotorcraft used for the carriage of cargo only, the cabin area may be considered a cargo compartment and, in addition to (a) through (d) of this section, the following apply:
(1) there shall be means to shut off the ventilating airflow to or within the compartment. Controls for this purpose shall be accessible to the flight crew in the crew compartment.
(2) required crew emergency exits shall be accessible under all cargo loading conditions.
(3) sources of heat within each compartment shall be shielded and insulated to prevent igniting the cargo.
(a) Combustion heater fire zones. The following combustion heater fire zones shall be protected against fire under the applicable provisions of sections 529.1181 through 529.1191, and sections 529.1195 through 529.1203:
(1) the region surrounding any heater, if that region contains any flammable fluid system components (including the heater fuel system), that could:
(i) be damaged by heater malfunctioning, or
(ii) allow flammable fluids or vapours to reach the heater in case of leakage;
(2) each part of any ventilating air passage that:
(i) surround the combustion chamber, and
(ii) would not contain (without damage to other rotorcraft components) any fire that may occur within the passage.
(b) Ventilating air ducts. Each ventilating air duct passing through any fire zone shall be fireproof. In addition:
(1) unless isolation is provided by fireproof valves or by equally effective means, the ventilating air duct downstream of each heater shall be fireproof for a distance great enough to ensure that any fire originating in the heater can be contained in the duct; and
(2) each part of any ventilating duct passing through any region having a flammable fluid system shall be so constructed or isolated from that system that the malfunctioning of any component of that system cannot introduce flammable fluids or vapours into the ventilating airstream.
(c) Combustion air ducts. Each combustion air duct shall be fireproof for a distance great enough to prevent damage from backfiring or reverse flame propagation. In addition:
(1) no combustion air duct shall communicate with the ventilating airstream unless flames from backfires or reverse burning cannot enter the ventilating airstream under any operating condition, including reverse flow or malfunction of the heater or its associated components; and
(2) no combustion air duct shall restrict the prompt relief of any backfire that, if so restricted, could cause heater failure.
(d) Heater controls; general. There shall be means to prevent the hazardous accumulation of water or ice on or in any heater control component, control system tubing, or safety control.
(e) Heater safety controls. For each combustion heater, safety control means shall be provided as follows:
(1) means independent of the components provide for the normal continuous control of air temperature, airflow, and fuel flow shall be provided, for each heater, to automatically shut off the ignition and fuel supply of that heater at a point remote from that heater when any of the following occurs:
(i) the heat exchanger temperature exceeds safe limits,
(ii) the ventilating air temperature exceeds safe limits,
(iii) the combustion airflow becomes inadequate for safe operation, and
(iv) the ventilating airflow becomes inadequate for safe operation;
(2) the means of complying with (e)(1) of this section for any individual heater shall:
(i) be independent of components serving any other heater whose heat output is essential for safe operation, and
(ii) keep the heater off until restarted by the crew;
(3) there shall be means to warn the crew when any heater whose heat output is essential for safe operation has been shut off by the automatic means prescribed in (e)(1) of this section.
(f) Air intakes. Each combustion and ventilating air intake shall be where no flammable fluids or vapours can enter the heater system under any operating condition:
(1) during normal operation; or
(2) as a result of the malfunction of any other component.
(1) each exhaust shroud shall be sealed so that no flammable fluids or hazardous quantities of vapours can reach the exhaust systems through joints; and
(2) no exhaust system shall restrict the prompt relief of any backfire that, if so restricted, could cause heater failure.
(h) Heater fuel systems. Each heater fuel system shall meet the powerplant fuel system requirements affecting safe heater operation. Each heater fuel system component in the ventilating airstream shall be protected by shrouds so that no leakage from those components can enter the ventilating airstream.
(i) Drains. There shall be means for safe drainage of any fuel that might accumulate in the combustion chamber or the heat exchanger. In addition:
(1) each part of any drain that operates at high temperatures shall be protected in the same manner as heater exhausts; and
(2) each drain shall be protected against hazardous ice accumulation under any operating condition.
Each part of the structure, controls and the rotor mechanism, and other parts essential to controlled landing and (for Category A) flight that would be affected by powerplant fires, shall be isolated under section 529.1191, or shall be:
(a) for Category A rotorcraft, fireproof; and
(b) for Category B rotorcraft, fireproof or protected so that they can perform their essential functions for at least 5 minutes under any foreseeable powerplant fire condition.
(a) In each area where flammable fluids or vapours might escape by leakage of a fluid system, there shall be means to minimize the probability of ignition of the fluids and vapours, and the resultant hazards if ignition does occur.
(b) Compliance with (a) of this section shall be demonstrated by analysis or tests, and the following factors shall be considered:
(1) possible sources and paths of fluid leakage, and means of detecting leakage;
(2) flammability characteristics of fluids, including effects of any combustible or absorbing materials;
(3) possible ignition sources, including electrical faults, overheating of equipment, and malfunctioning of protective devices;
(4) means available for controlling or extinguishing a fire, such as stopping flow of fluids, shutting down equipment, fireproof containment, or use of extinguishing agents; and
(5) ability of rotorcraft components that are critical to safety of flight to withstand fire and heat.
(c) If action by the flight crew is required to prevent or counteract a fluid fire (e.g. equipment shutdown or actuation of a fire extinguisher), quick acting means shall be provided to alert the crew.
(d) Each area where flammable fluids or vapours might escape by leakage of a fluid system shall be identified and defined.
(a) It shall be demonstrated by analysis, test, or both, that the rotorcraft external load attaching means for rotorcraft-load combinations to be used for nonhuman external cargo applications can withstand a limit static load equal to 2.5, or some lower load factor approved under sections 529.337 through 529.341, multiplied by the maximum external load for which authorization is requested. It shall be demonstrated by analysis, test, or both that the rotorcraft external load attaching means and corresponding personnel carrying device system for rotorcraft-load combinations to be used for human external cargo applications can withstand a limit static load equal to 3.5 or some lower load factor, not less than 2.5, approved under sections 529.337 through 529.341, multiplied by the maximum external load for which authorization is requested. The load for any rotorcraft-load combination class, for any external cargo type, shall be applied in the vertical direction. For jettisonable external loads of any applicable external cargo type, the load shall also be applied in any direction making the maximum angle with the vertical that can be achieved in service but not less than 30°. However, the 30° angle may be reduced to a lesser angle if:
(1) an operating limitation is established limiting external load operations to such angles for which compliance with this has been demonstrated; or
(2) it is demonstrated that the lesser angle shall not be exceeded in service.
(b) The external load attaching means, for jettisonable rotorcraft-load combinations, shall include a quick-release system to enable the pilot to release the external load quickly during flight. The quick-release system shall consist of a primary quick-release subsystem and a backup quick-release subsystem that are isolated from one another. The quick-release system, and the means by which it is controlled, shall comply with the following:
(1) a control for the primary quick-release subsystem shall be installed either on one of the pilot’s primary controls or in an equivalently accessible location and shall be designed and located so that it may be operated by either the pilot or a crew member without hazardously limiting the ability to control the rotorcraft during an emergency situation;
(2) a control for the backup quick-release subsystem, readily accessible to either the pilot or another crew member, shall be provided;
(3) both the primary and backup quick release subsystems shall:
(i) be reliable, durable, and function properly with all external loads up to and including the maximum external limit load for which authorization is requested,
(ii) be protected against electromagnetic interference (EMI) from external and internal sources and against lightning to prevent inadvertent load release;
(a) the minimum level of protection required for jettisonable rotorcraft-load combinations used for nonhuman external cargo is a radio frequency field strength of 20 volts per metre, and
(b) the minimum level of protection required for jettisonable rotorcraft-load combinations used for human external cargo is a radio frequency field strength of 200 volts per metre, and
(iii) be protected against any failure that could be induced be a failure mode of any other electrical or mechanical rotorcraft system.
(c) For rotorcraft-load combinations to be used for human external cargo applications, the rotorcraft shall:
(1) for jettisonable external loads, have a quick-release systems that meets the requirements of (b) of this section and that:
(i) provides a dual actuation device for the primary quick-release subsystem, and
(ii) provides a separate dual actuation device for the backup quick-release subsystem;
(2) have a reliable, approved personnel carrying device system that has the structural capability and personnel safety features essential for external occupant safety;
(3) have placards and markings at all appropriate locations that clearly state the essential system operating instructions and, for the personnel carrying device system, ingress and egress instructions;
(4) have equipment to allow direct intercommunication among required crew members and external occupants;
(5) have the appropriate limitations and procedures incorporated in the flight manual for conducting human external cargo operations; and
(6) for human external cargo applications requiring use of Category A rotorcraft, have one-engine-inoperative hover performance data and procedures in the flight manual for the weights, altitudes, and temperatures for which external load approval is requested.
(d) The critically configured jettisonable external loads shall be demonstrated by a combination of analysis, ground tests and flight tests to be both transportable and releasable throughout the approved operational envelope without hazard to the rotorcraft during normal flight conditions. In addition, these external loads shall be demonstrated to be releasable without hazard to the rotorcraft during emergency flight conditions.
(e) A placard or marking shall be installed next to the external load attaching means clearly stating any operational limitations and the maximum authorized external load as demonstrated under section 529.25 and this section.
(f) The fatigue evaluation of section 529.571 of this chapter does not apply to rotorcraft-load combinations to be used for nonhuman external cargo except for the failure of critical structural elements that would result in a hazard to the rotorcraft. For rotorcraft-load combinations to be used for human external cargo, the fatigue evaluation of section 529.571 of this chapter applies to the entire quick-release and personnel carrying device structural systems and their attachments.
There shall be reference marks for levelling the rotorcraft on the ground.
Ballast provisions shall be designed and constructed to prevent inadvertent shifting of ballast in flight.
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